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tan Moir and Allan Sea bridge 

Military Avionics 

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Aerospace Series 

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Military Avionics Systems 

Military Avionics Systems Ian Moir and Allan G. Seabridge 
© 2006 John Wiley & Sons, Ltd. ISBN: 0-470-01632-9 

Military Avionics Systems 

Ian Moir 

Allan G. Seabridge 

Displays chapter contributed by Malcolm Jukes 

John Wiley & Sons, Ltd 

Copyright © 2006 John Wiley & Sons Ltd, The Atrium, Southern Gate, Chichester, 
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Library of Congress Cataloging-in-Publication Data 

Moir, I. (Ian) 

Military avionics systems/Ian Moir. 

p. cm. 
"Displays chapter contributed by Malcom Jukes." 
Includes bibliographical references and index. 
ISBN 0-470-01632-9 (cloth : alk. paper) 
1. Avionics. 2. Airplanes, Military-Electronic equipment. 
3. Electronics in military engineering. I. Jukes, Malcom. II. Title. 
UG1420.M565 2006 
623.74'6049-dc22 2005031935 

British Library Cataloguing in Publication Data 

A catalogue record for this book is available from the British Library 

ISBN-13 978-0-470-01632-9 (HB) 
ISBN-10 0-470-01632-9 (HB) 

Typeset in 10/12pt Times by Thomson Press (India) Limited, New Delhi. 

Printed and bound in Great Britain by Antony Rowe Ltd, Chippenham, Wiltshire 

This book is printed on acid-free paper responsibly manufactured from sustainable forestry 

in which at least two trees are planted for each one used for paper production. 


Series preface 

/\I^KI1U WlctlgCIIltiiia 

About the authors 


1 Military roles 




Air superiority 


Role description 


Key performance characteristics 


Crew complement 


Systems architecture 


Air superiority - aircraft types 


Ground attack 


Role description 


Key performance characteristics 


Crew complement 


Systems architecture 


Ground attack - aircraft types 



lc bomber 


Role description 


Key performance characteristics 


Crew complement 


Systems architecture 


Strategic bomber - aircraft types 


Maritime patrol 


Role description 


Anti- surface unit warfare (ASuW) role 


Anti-submarine warfare (ASW) role 


Search and rescue (SAR) role Datum searches 






19 Area searches Scene-of-action commander 


Exclusive economic zone protection Oil and gas rig patrols Anti-pollution Fishery protection Customs and excise cooperation 


Key performance characteristics 


Crew complement 


Systems architecture 


MPA aircraft types 

.6 Battlefield surveillance 


Role description 


Key performance characteristics 


Crew complement 


Systems architecture 


Battlefield surveillance aircraft types 

.7 Airborne early warning 


Role description 


Key performance characteristics 


Crew complement 


Systems architecture 


AEW aircraft types 

.8 Electronic warfare 


Role description 


Electronic countermeasures 


Electronic support measures 


Signals intelligence (SIGINT) 


Key performance characteristics 


Crew complement 


Systems architecture 


Example aircraft types 

.9 Photog 

raphic reconnaissance 


Role description 


Key performance characteristics 


Crew complement 


Systems architecture 


Typical aircraft types 

.10 Air-to-; 

air refuelling 


Role description 


Key performance characteristics 


Crew complement 


Systems architecture 


Aircraft types 

.11 Troop/materiel transport 


Role description 


Key performance characteristics 




1.11.3 Crew complement 39 

1.11.4 Systems architecture 39 

1.11.5 Typical aircraft types 40 

1.12 Unmanned air vehicles 41 

1.13 Training 43 
1.13.1 Typical aircraft types 45 

1.14 Special roles 45 
1.14.1 Examples of special roles 45 

1.15 Summary 45 
Further reading 46 

2 Technology and architectures 47 

2.1 Evolution of avionics architectures 47 

2.1.1 Distributed analogue architecture 49 

2.1.2 Distributed digital architecture 50 

2.1.3 Federated digital architecture 52 

2.1.4 Integrated modular architecture 54 

2.1.5 Open architecture issues 54 

2.1.6 Impact of digital systems 56 

2.1.7 Response of the services to digital avionics systems issues 57 

2.1.8 Need to embrace COTS 58 

2.2 Aerospace-specific data buses 60 

2.2.1 Tornado serial 60 

2.2.2 ARINC 429 62 

2.2.3 MIL-STD-1553B 64 

2.2.4 STANAG 3910 67 

2.3 JIAWG architecture 70 

2.3.1 Generic JIAWG architecture 70 

2.3.2 High-speed data bus 72 

2.3.3 PI bus 73 

2.3.4 TM bus 73 

2.3.5 Obsolescence issues 73 

2.4 COTS data buses 74 

2.4.1 Fibre channel 75 

2.4.2 Fibre channel options 77 

2.4.3 IEEE 1394 firewire 77 

2.5 Real-time operating systems 78 

2.5.1 Key attributes 78 

2.5.2 Safety 79 

2.5.3 Software partitioning 80 

2.5.4 Software languages 82 

2.5.5 Security 82 

2.6 RF integration 83 
2.6.1 Primary radar evolution 84 Independent systems of the 1950s 84 Integrated systems of the 1960s and 1970s 85 Integrated modular architecture of the 1990s 86 


2.6.2 JIAWG RF subsystem integration 87 

2.7 Pave Pace/F-35 shared aperture architecture 94 

References 97 

3 Basic radar systems 99 

3.1 Basic principles of radar 99 

3.2 Radar antenna characteristics 104 

3.3 Major radar modes 107 

3.3.1 Air-to-air search 107 

3.3.2 Air-to-air tracking 108 

3.3.3 Air-to-air track- while- scan 109 

3.3.4 Ground mapping 110 

3.4 Antenna directional properties 111 

3.5 Pulsed radar architecture 112 

3.5.1 Pulsed radar components 112 Modulator 113 Transmitter 113 Antenna 113 Receiver 114 Video processor 114 

3.5.2 Pulsed modulation 114 

3.5.3 Receiver characteristics 116 Noise 116 Clutter 122 

3.5.4 Radar range equation 122 

3.6 Doppler radar 123 

3.7 Other uses of radar 124 

3.7.1 Frequency modulation ranging 124 

3.7.2 Terrain-following radar 125 

3.7.3 Continuous wave illumination 126 

3.7.4 Multimode operation 127 

3.8 Target tracking 128 

3.8.1 Range tracking 128 

3.8.2 Angle tracking 129 Sequential lobing 129 Conical scan 130 Monopulse 132 
References 134 

4 Advanced radar systems 135 

4.1 Pulse compression 135 

4.1.1 Coherent transmission 137 

4.1.2 Fourier transform 140 

4.2 Pulsed Doppler operation 140 

4.2.1 Range ambiguities 143 

4.2.2 Effect of the PRF on the frequency spectrum - Doppler ambiguities 144 

4.2.3 Range and Doppler ambiguities 145 


4.3 Pulsed Doppler radar implementation 149 

4.3.1 Receiver 150 

4.3.2 Signal processor 150 

4.3.3 Radar data processor 154 

4.4 Advanced antennas 156 

4.4.1 Principle of the phased array 156 

4.4.2 Planar arrays 157 

4.4.3 Electronically scanned array 158 

4.4.4 Active electronically steered array (AESA) 160 

4.5 Synthetic aperture radar 164 

4.6 Low observability 170 

4.6.1 Factors affecting the radar cross-section 173 

4.6.2 Reducing the RCS 178 

4.6.3 Comparative RCS values 179 

4.6.4 Low probability of intercept operation 180 
References 181 

5 Electrooptics 183 

5 . 1 Introduction 183 


Electronic warfare 229 

6.1 Introduction 229 

6.2 Signals intelligence (SIGINT) 233 
6.2.1 Electronic intelligence (ELINT) 234 





vision goggles 


IR ima 



IR imaging device 


Rotating scanner 


Planar image 


Focal plane array or 'staring array' 


IR detector technology 


IR tracking 


IR seeker heads 


Image tracking 


IR search and track systems 




Principles of operation 


Laser sensor applications 


US Air Force airborne laser (ABL) 


Laser safety 


Integrated systems 


Electrooptic sensor fusion 


Pod installations 


Turret installations 


Internal installations 



6.2.2 Communications intelligence (COMINT) 236 

6.3 Electronic support measures 238 

6.4 Electronic countermeasures and counter-countermeasures 241 

6.4.1 Noise jamming 241 Burnthrough 243 

6.4.2 Deception jamming 244 

6.4.3 Deployment of the jamming platform 245 

6.4.4 Low probability of intercept (LPI) radar 246 

6.5 Defensive aids 247 

6.5.1 Radar warning receiver 248 

6.5.2 Missile warning receiver 249 

6.5.3 Laser warning systems 249 

6.5.4 Countermeasure dispensers 250 Chaff and flares 251 Towed decoy 252 

6.5.5 Integrated defensive aids systems 253 AN/APG-79 AESA radar 253 AN/ALR-67 radar warning receiver 253 AN/ALQ-214 integrated defensive electronic 
countermeasures (IDECM) 254 AN/ALE-47 countermeasure dispenser 255 
References 256 

7 Communications and identification 257 

7.1 Definition of CNI 257 

7.1.1 RF spectrum 258 

7.1.2 Communications control systems 259 

7.2 RF propagation 259 

7.2.1 High frequency 261 

7.2.2 Very high frequency 262 

7.2.3 Satellite communications 264 

7.3 Transponders 266 

7.3.1 Air traffic control (ATC) transponder - mode S 267 

7.3.2 Traffic collision and avoidance system 269 

7.3.3 Automatic dependent surveillance - address mode (ADS- A) 272 

7.3.4 Automatic dependent surveillance - broadcast mode (ADS-B) 272 

7.3.5 Indentification friend or foe (IFF) 272 

7.4 Data links 273 

7.4.1 JTIDS operation 275 

7.4.2 Other data links 277 

7.5 Network-centric operations 277 
References 279 

8 Navigation 281 
8.1 Navigation principles 281 

8.1.1 Introduction 281 

8.1.2 Basic navigation 282 


8.2 Radio navigation 284 
8.2.1 Oceanic crossings 285 

8.3 Inertial navigation fundamentals 286 

8.4 Satellite navigation 287 

8.4.1 Differential GPS 288 

8.4.2 Wide-area augmentation system (WAAS) 288 

8.4.3 Local-area augmentation system (LAAS) 289 

8.5 Integrated navigation 290 

8.5.1 Sensor usage - phases of flight 291 

8.5.2 GPS overlay programme 292 

8.5.3 Categories of GPS receiver 292 

8.6 Flight management system 292 

8.6.1 FMS CDU 294 

8.6.2 FMS functions 296 

8.6.3 LNAV 297 

8.6.4 Airway navigation 298 

8.6.5 Area navigation 299 

8.6.6 VNAV 299 

8.6.7 Four-dimensional navigation 299 

8.6.8 Full performance based navigation 299 

8.6.9 FMS procedures 300 

8.6.10 Traffic collision and avoidance system (TCAS) 300 

8.6.11 GPWS and EGPWS 300 

8.7 Navigation aids 302 

8.7.1 Automatic direction finding 303 

8.7.2 Very high-frequency omnirange (VOR) 303 

8.7.3 Distance-measuring equipment (DME) 303 

8.7.4 TACAN 304 

8.7.5 VORTAC 305 

8.7.6 Hyperbolic navigation systems - LORAN-C 305 

8.7.7 Instrument landing system 307 

8.7.8 Microwave landing system (MLS) 309 

8.8 Inertial navigation 310 

8.8.1 Principles of operation 310 

8.8.2 Stand-alone inertial navigation system 311 

8.8.3 Air data and inertial reference systems (ADIRS) 313 

8.8.4 Inertial platform implementations 315 

8.8.5 Space axes to the Greenwich meridian 316 

8.8.6 Earth axes to geographic axes 316 

8.8.7 Geographic to great circle (navigation) 318 

8.8.8 Great circle/navigation axes to body axes (strapdown) 319 

8.8.9 Platform alignment 319 Platform levelling 320 Gyrocompass alignment 320 

8.8.10 Historical perspective - use of inertial platforms 321 

8.9 Global navigation satellite systems 322 
8.9.1 Introduction to GNSS 322 


8.9.2 Principles of operation 323 

8.9.3 Integrity features 324 

8.9.4 GPS satellite geometry 325 
8.10 Global air transport management (GATM) 325 

8.10.1 Communications 326 

8. 10. 1 . 1 Air-to-ground VHF data link 326 Air-to-ground SATCOM communications 327 HF data link 327 8.33 kHz VHF voice communications 327 Protected ILS 328 

8.10.2 Navigation 328 Area navigation (RNAV) 328 RNP RNAV and actual navigation performance 329 Required navigation performance (RNP) 329 RNAV standards within Europe 329 RVSM 330 RVSM implementation 331 Differential GPS enhancements 331 

8.10.3 Surveillance 332 
References 333 

9 Weapons carriage and guidance 335 

9.1 Introduction 335 

9.2 F-16 Fighting Falcon 336 

9.2.1 F-16 evolution 337 

9.2.2 F-16 mid-life update 340 

9.2.3 F-16 E/F (F-16 block 60) 343 

9.3 AH-64 C/D Longbow Apache 345 

9.3.1 Baseline system 345 

9.3.2 Longbow Apache 345 

9.3.3 Modernisation of TADS/PNVS 348 

9.3.4 Weapons 349 

9.4 Eurofighter Typhoon 349 
9.4.1 Sensors and navigation 352 


9.5 F/A-22 Raptor 357 



Displays and controls 


Flight control 


Utilities control 


Systems integration 










AESA radar 


Electronic warfare and electronic support measures 






Vehicle management system 


9.5.7 Weapons 361 

9.6 Nimrod MRA4 361 

9.6.1 Navigation and displays 363 


10 Vehicle Management Systems 379 

10.1 Introduction 379 

10.2 Historical development of control of utility systems 381 

10.3 Summary of utility systems 383 

10.3.1 Mechanical systems 383 

10.3.2 Crew systems 383 

10.3.3 Power systems 383 

10.3.4 Fuel system 384 

10.3.5 Air systems 384 

10.3.6 Electrical power distribution systems 384 

10.3.7 Vehicle management system (VMS) 384 

10.3.8 Prognostics and health management 385 

10.4 Control of utility systems 385 

10.5 Subsystem descriptions 389 
10.5.1 Mechanical systems 390 Primary flight controls 390 Secondary flight controls 392 


Utilities control 




Mission and sensor systems 




Weapons and stores 


F-35 joint strike fighter 


Integrated common processors 


AESA radar 


Integrated EW/CNI and EO systems 


Displays suite 


Vehicle management system 






MIL-STD-1760 standard stores interface 



air missiles 


AIM-9 Sidewinder 






' Air-to-, 

ground ordnance 


Wind-corrected munition dispenser 


Joint direct attack munition 






Storm Shadow/SCALP EN 



xiv CONTENTS Landing gear 392 Wheels, brakes and tyres 393 Arrestor hook/brake parachute 393 Actuation mechanisms 394 

10.5.2 Crew systems 394 Crew escape 394 Aircrew clothing 394 Life support 395 Oxygen/OBOGS 395 Canopy jettison 397 

10.5.3 Power systems 397 Propulsion system 397 Secondary power 397 Emergency power 398 

10.5.4 Electrical power generation and distribution 399 F/A-18E/F Super Hornet 399 F/A-22 Raptor 399 

10.5.5 Hydraulic power generation and distribution 400 

10.5.6 Fuel systems 401 

10.5.7 Air systems 403 

10.5.8 Electrical utilisation systems 405 

10.5.9 Prognostics and health management 405 
10.6 Design considerations 406 

10.6.1 General 406 

10.6.2 Processor and memory 407 

10.6.3 Interfacing 407 

10.6.4 Software 408 

10.6.5 Obsolescence 408 

10.6.6 Human-machine interface 409 
References 410 
Further reading 410 

11 Displays 411 

11.1 Introduction 411 

11.2 Crew station 412 

11.2.1 Hawker Siddley (BAe) Harrier GR.Mkl and GR.Mk3 

(RAF) and AV-8A (USMC) 412 

11.2.2 McDonnell Douglas F/A- 1 8 Hornet 4 1 3 

11.2.3 Eurofighter Typhoon 415 

11.2.4 Lockheed Martin F-22 Raptor 416 

11.2.5 Boeing (McDonnell Douglas/Hughes) AH-64D Longbow Apache 417 

11.3 Head-up display 418 

11.3.1 HUD principles 419 

11.3.2 Collimating (refractive optics) head-up display 420 

11.3.3 Field of view 421 

11.3.4 Collimating (refractive) HUD - examples 421 British Aerospace Harrier GR.Mkl and GR.Mk3 421 

CONTENTS xv McDonnell Douglas/British Aerospace Night- Attack 

Harrier II (GR-7 and AV-8B) " 422 

11.3.5 Pupil-forming (reflective/diffractive) head-up displays 423 

11.3.6 Pupil-forming (reflective/diffractive) HUD - examples 425 F16 LANTIRN HUD - multibounce quasi-axial 
configuration 425 Eurofighter Typhoon HUD - single-element 

off-axis configuration 426 

11.3.7 Head-up display functional description 427 

11.3.8 Image generation 429 

11.3.9 HUD symbology and principles of use 430 Primary flight data 431 Navigation symbology 432 Air-to- surface weapon aiming 432 Air-to-air weapon aiming 433 

11.4 Helmet-mounted displays 434 

11.4.1 HMD physiological and environmental aspects 436 

11.4.2 Head tracker 437 

1 1 .4.3 Optical head tracker 438 

11.4.4 Electromagnetic head tracker 438 

11.4.5 HMD accuracy and dynamic performance 438 

11.4.6 HMD optical configurations 439 

11.4.7 Helmet-mounted displays - examples 442 Integrated helmet and display sight system (IHADS) 442 

1 1 .4.7.2 Helmet-mounted sight 442 Joint helmet-mounted cueing system (JHMCS) 443 

1 1 .4.7.4 Eurofighter Typhoon HMD 444 Joint strike fighter HMD 446 

11.4.8 Helmet-mounted display functional description 447 

11.4.9 Binocular day/night HMD architectures 448 

11.4.10 HMD symbology 448 

11.4.11 HMD as a primary flight reference 449 

11.5 Head-down displays 450 

11.5.1 CRT multifunction head-down display 450 Shadow-mask CRT 450 X/Y deflection amplifier 452 Shadow-mask CRT characteristics 452 CRT MFD: principles of operation 453 F/A-18 and AV-8B multipurpose colour display (MPCD) 453 

11.5.2 AMLCD multifunction head-down display 455 AMLCD display head assembly 455 Backlight 457 AMLCD characteristics 458 AMLCD sourcing 459 AMLCD MFD: principles of operation 459 Integrated display unit 460 

11.6 Emerging display technologies 461 


11.6.1 Microdi splay technologies 461 

11.6.2 High-intensity light sources 462 

11.6.3 Transmissive LCD 462 

11.6.4 Reflective LCD 462 

11.6.5 Digital micromirror device 463 

11.6.6 Rear-projection 'big picture' head-down display 466 

11.6.7 Solid-state helmet-mounted display 467 

11.6.8 Organic light-emitting diodes (OLEDs) 469 

11.6.9 Virtual retinal displays 470 
11.7 Visibility requirements 471 

11.7.1 Military requirements 47 1 Head-down display: high ambient - sun rear 472 Head-down display: high ambient - sun forward 472 Head-up display and helmet-mounted display: 

high ambient - sun forward 473 Low ambient - dusk/dawn transition 474 Night 474 

11.7.2 US DoD definitions and requirements 475 

11.7.3 European (Eurofighter Typhoon) definitions and requirements 475 

11.7.4 Viewability examples 476 

1 1 .7.4. 1 AMLCD head-down display 476 Head-up display 477 Night- vision imaging system compatibility 477 
References 480 

Bibliography 483 

Commercial handbooks, standards and specifications 484 

Military handbooks, standards and specifications 485 

Advisory circulars 485 

Standards 486 

Technical standing orders 486 

Useful websites 486 

Glossary 487 

Units 499 

Aircraft types 500 

Index 503 

Series Preface 

The field of aerospace is wide ranging and covers a variety of products, disciplines and 
domains, not merely in engineering but in many related supporting activities. These combine 
to enable the aerospace industry to produce exciting and technologically challenging 
products. A wealth of knowledge is contained by practitioners and professionals in the 
aerospace fields that is of benefit to other practitioners in the industry, and to those entering 
the industry from University. 

The Aerospace Series aims to be a practical and topical series of books aimed at 
engineering professionals, operators, users and allied professions such as commercial and 
legal executives in the aerospace industry. The range of topics spans design and develop- 
ment, manufacture, operation and support of aircraft as well as infrastructure operations, and 
developments in research and technology. The intention is to provide a source of relevant 
information that will be of interest and benefit to all those people working in aerospace. 

Ian Moir and Allan Seabridge 


This book has taken a long time to prepare and we would not have completed it without the 
help and support of colleagues and organisations who gave their time and provided 
information with enthusiasm. 

We would especially like to thank Malcolm Jukes, Kevin Burke and Keith Atkin who 
reviewed a number of chapters, and to Leon Skorczewski who bravely reviewed the entire 
manuscript and provided valuable comments. 

The following organisations kindly provided information and images: 


Brilliant Technology 
Cambridge Display Technology 
Eurofighter GmbH 
Federation of American Scientists 

for UAV pictures and Nimrod R 

Mk 1 

Honeywell Aerospace Yeovil 
InfraRed 1 
Kaiser Electronics 

Korry Electronics 

Lockheed Martin 

Lockheed Martin Fire Control 

Martin Baker 

Micro vision Inc 

Northrop Grumman Corporation 


Rockwell Collins 

Royal Aeronautical Society 

SAAB Avitronics 

Smiths Group 

Texas Instruments 


Thales Optronics 


Special thanks to Marc Abshire, Randy Anderson, Myrna Buddemeyer, Ron Colman, 
Francesca De Florio, Joan Ferguson, Alleace Gibbs, Charlotte Haensel-Hohenhausen, Karen 
Hager, Dexter Henson, Clive Marrison, Katelyn Mileshosky, Ian Milne, Marianne Murphy, 
Shelley Northcott, Beth Seen, Kevin Skelton and Greg Siegel for their kind help in securing 
high quality images for inclusion in the book. 

Aircraft pictures were obtained from the US Department of Defence Air Force Link 
website at : B-2 Spirit - US Air Force Photo by Tech Sgt Cecilio Ricardo; B-1B 
Lancer - US Air Force Photo by Senior Airman Michel B.Keller; A- 10 Thunderbolt - US 
Air Force Photo by Senior Airman Stephen Otero; T-38 Talon US Air Force Photo by Staff 
Sgt Steve Thurow; C-5 US Air Force Photo by Master Sgt Clancy Pence, C-17 US Air Force 


Photo by Airman 1st Class Aldric Bowers; Predator US Air Force Photo by Master Sgt Deb 
Smith, Global Hawk by George Rohlmaller; E-JSTARS US Air Photo by Staff Sgt Shane 
Cuomo; F-l 17A Nighthawk US Air Force Photo by Staff Sgt Aaron D Allmon II; B-2 Spirit 
Bomber US Air Force Photo by Master Sgt Val Gempis; 

VP-45 US Navy Squadron is acknowledged for the photograph of a MX-20 turret on P-3 

We would like to thank the staff at John Wiley who took on this project part way through 
its progress and guided us to a satisfactory conclusion. 

About the Authors 

After 20 years in the Royal Air Force, Ian Moir went on to Smiths Industries in the UK 
where he was involved in a number of advanced projects. Since retiring from Smiths he is 
now in demand as a highly respected consultant. Ian has a broad and detailed experience 
working in aircraft avionics systems in both military and civil aircraft. From the RAF 
Tornado and Apache helicopter to the Boeing 777, Ian's work has kept him at the forefront of 
new system developments and integrated systems implementations. He has a special interest 
in fostering training and education in aerospace engineering. 

Allan Seabridge is the Chief Flight Systems Engineer at BAE SYSTEMS at Warton in 
Lancashire in the UK. In over 30 years in the aerospace industry his work has included 
avionics on the Nimrod MRA 4 and Joint Strike Fighter as well as a the development of a 
range of flight and avionics systems on a wide range of fast jets, training aircraft and ground 
and maritime surveillance projects. Spending much of his time between Europe and the US, 
Allan is fully aware of systems developments worldwide. He is also keen to encourage a 
further understanding of integrated engineering systems. 

Plate 1 F/A 18 C Cockpit (McDonnell Douglas). 

Plate 2 Eurofighter Typhoon Instrument Panel Layout (BAE SYSTEMS). 

Military Avionics Systems Ian Moir and Allan G. Seabridge 
© 2006 John Wiley & Sons, Ltd. ISBN: 0-470-01632-9 

Plate 3 F16 Cockpit (Lockheed Martin). 

Plate 4 Eurofighter Typhoon (BAE SYSTEMS). 

Plate 5 Colour Shadow Mask CRT. 

Plate 6 CIE 1976 UCS Chromaticity Diagram. 

Plate 7 Eurofighter Typhoon MHDD (Smiths). 

Plate 8 JSF 'Big Picture' Projection Head-Down Display (Rockwell Collins). 


Evolution of Avionics 

Avionics is a word coined in the late 1930s to provide a generic name for the increasingly 
diverse functions being provided by AVIation electrONICS. World War II and subsequent 
Cold War years provided the stimulus for much scientific research and technology devel- 
opment which, in turn, led to enormous growth in the avionic content of military aircraft. 
Today, avionics systems account for up to 50% of the fly-away cost of an airborne military 
platform and are key components of manned aircraft, unmanned aircraft, missiles and 
weapons. It is the military avionics of an aircraft that allow it to perform defensive, offensive 
and surveillance missions. 

A brief chronology of military avionics development illustrates the advances that have 
been made from the first airborne radio experiments in 1910 and the first autopilot 
experiments a few years later. The 1930s saw the introduction of the first electronic aids 
to assure good operational reliability such as blind flying panels, radio ranging, non- 
directional beacons, ground-based surveillance radar, and the single-axis autopilot. The 
1940s saw developments in VHF communications, identification friend or foe (IFF), gyro 
compass, attitude and heading reference systems, airborne intercept radar, early electronic 
warfare systems, military long-range precision radio navigation aids, and the two-axis 
autopilot. Many of these development were stimulated by events leading up to World War II 
and during the war years. 

The 1950s saw the introduction of tactical air navigation (TACAN), airborne intercept 
radar with tracking capability and Doppler radar, medium pulse repetition frequency (PRF) 
airborne intercept radar, digital mission computers and inertial navigation systems. The 
1960s saw the introduction of integrated electronic warfare systems, fully automated weapon 
release, terrain-following radar, automatic terrain following, the head-up display laser target 
marketing technology and the early digital mission computer. 

Over the years, as specialist military operational roles and missions have evolved, they 
have often driven the development of role-specific platforms and avionics. Looking across 
the range of today's airborne military platforms, it is possible to identify categories of 

Military Avionics Systems Ian Moir and Allan G. Seabridge 
© 2006 John Wiley & Sons, Ltd. ISBN: 0-470-01632-9 


avionics at system, subsystem and equipment levels that perform functions common to all 
platforms, or indeed perform unique mission- specific functions. 

Technology improvements in domestic markets have driven development in both com- 
mercial and military systems, and the modern military aircraft is likely to contain avionic 
systems that have gained benefit from domestic computing applications, especially in the IT 
world, and from the commercial aircraft field. This has brought its own challenges in 
qualifying such development for use in the harsh military environment, and the challenge of 
meeting the rapid turnround of technology which leads to early obsolescence. 

Avionics as a Total System 

An avionics system is a collection of subsystems that display the typical characteristics of 
any system as shown in Figure LI. The total system may be considered to comprise a number 
of major subsystems, each of which interacts to provide the overall system function. Major 
subsystems themselves may be divided into minor subsystems or equipment which in turn 
need to operate and interact to support the overall system. Each of these minor subsystems is 
supported by components or modules whose correct functional operation supports the overall 
system. The overall effect may be likened to a pyramid where the total system depends upon 
all the lower tiers. 

Avionics systems may be represented at a number of different levels as described below: 

1. A major military task force may comprise a large number of differing cooperating 
platforms, each of which contributes to the successful accomplishment of the task force 





A system 
typically is 

comprised of 
a series of 

inter- related 

Major Task Group 

Aircraft Equipment 

\* — >SS3 

Aircraft System J. J. -L 

Elecfronic Module 


Figure 1.1 Avionics as a 'system of systems'. 


A military avionics system 


Nav Aids 

Landing Aids 


Flight Management 
— Air data 

Radar Altimeter 


Accident data recorder 



— Data Link 


- Radar 

- Electro-Optics 

- Defensive Aids 

Mission Computing 
Data Fusion 
Data Loading 
Weapons Aiming 

— Weapons Management 

— Mission recorders 
Intelligence data base 
Electronic countermeasures 

Cockpit displays 
Mission Displays 
Head Up Display 
Helmet Mounted Display 

Figure 1.2 Product breakdown structure of a military aircraft system. 

mission. Within this context an individual strike aircraft or surveillance platform avionics 
system may represent one component of many within the task force. 

2. At the individual platform or aircraft level, a collection of subsystems and components or 
modules operate to support the successful completion of the primary role of the platform, 
be it reconnaissance, strike, support or surveillance. 

3. The individual equipment that supports the overall system of the platform is a collection 
of units or modules, control panels and displays, each of which has to operate correctly to 
support subsystem and overall system operation. 

4. Finally, the electronic modules that form the individual components of the aircraft 
avionics systems may be regarded as systems within their own right, with their own 
performance requirements and hardware and software elements. 

In general within this book, most discussion is centred upon the aircraft-level avionics 
system and upon the major subsystems and minor subsystems or equipment that support it. 
Passing reference to the higher-level system is made during brief coverage of network centric 
operations. In some cases the detailed operation of some components such as data buses is 
addressed in order that the reader may understand the contribution that these elements have 
made to advances in the overall integration of platform avionics assets. 

The product breakdown structure of a military aircraft system is shown in Figure 1.2. 

Increasing Complexity of Functional Integration 

As avionics systems have evolved, particularly over the past two or three decades, the level 
of functional integration has increased dramatically. The nature of this increase and the 
accompanying increase in complexity is portrayed in Figure 1.3. 

In the early stages, the major avionics subsystems such as radar, communications, 
navigation and identification (CNI), displays, weapons and the platform vehicle could be 






i 1 1- 

l I I 

I l l 

-1- r 1 i 

|cni | ; ; 

_ — , r _ 


• , 


s / / / 


Figure 1.3 Increase in functional integration over time. 









Supportability (Re-use) 

Software Programs 

Memory Requirements 


Data Handling 

Data Links 




Power Consumption 

Technology Window 

considered as discrete subsystems, the function of which could be easily understood. The 
performance requirements could be relatively easily specified and captured, and, although 
there were information interchanges between them, each could stand alone and the 
boundaries of each subsystem was 'hard' in the sense that it was unlikely to be affected 
by the performance of a neighbouring subsystem. 

As time progressed, the functionality of each subsystem increased and some boundaries 
blurred and functions began to overlap. Also, the number of subsystems began to increase 
owing to the imposition of more complex mission requirements and because of the 
technology developments that furnished new sensors. Improved data processing and higher 
bandwidth data buses also contributed to providing much higher data processing capabilities 
and the means to allow the whole system to become more integrated. 

Further technology developments added another spiral to this trend, resulting in greater 
functionality, further increasing integration and with a blurring of functional boundaries as 
subsystems became able to share ever greater quantities of data. This evolution has been a 
continual process, although it is portrayed in three stages in Figure 1.3 for reasons of 

The outcome of this evolution has been to increase: performance; sensor types; 
functionality; cost; integration; complexity; supportability (reuse); software programs in 
terms of executable code; memory requirements; throughput; reliability; data handling; data 
links; and obsolescence. 

The result has been to decrease: size; weight; power consumption; and technology 


Organisation of the Book 

This book is organised to help the reader to comprehend the overarching avionics systems 
issues but also to focus on specific functional areas. Figure 1.4 shows how the various 
chapters relate to the various different major functional subsystems. 

The book provides a military avionics overview aimed at students and practitioners in the 
field of military avionics. 

Chapter 1 lists and describes the roles that military air forces typically need to perform. It 
is the understanding of these roles that defines the requirements for a particular suite of 
avionics, sensors and weapons for different platforms. 

Chapter 2 examines the technology that has led to different types of system architecture. 
This technology has resulted in sophisticated information processing structures to transfer 
high volumes of data at high rates, and has resulted in greatly increased functional 

The subject of radar is covered in Chapter 3 which describes radar basic principles, while 
Chapter 4 explains some of the advanced features that characterise different types of radar 
used for specific tasks. 

Chapter 5 deals with electrooptical (EO) sensors and their use in passive search, detection 
and tracking applications. This includes a description of the integration of EO sensor 
applications in turrets and pods, as well as personal night vision goggles. 

Chapter 6 looks at the sensitive, and often highly classified, field of electronic warfare and 
the gathering of intelligence by aircraft using sophisticated receiving equipment and 
processing techniques. 


Chapters 3 & 4 

L i 


Chapter 6 ^— ^— 






Chapters 7 & 8 


Chapter 5 



\ Chapter 1 / 


Chapter 2 

-High Speed Data Bus* 


Chapter 11 
i ' 




Chapter 10 



Chapter 9 

Figure 1.4 Military avionics functional subsystems. 


Chapter 7 is concerned with communications and identification; this describes the 
mechanisms by which an aircraft is identified to other stakeholders such as air traffic control 
and to friendly forces, as well as the different form of communications available for speech 
and encrypted data. 

Chapter 8 covers the subject of navigation and the means by which pilots are able to 
navigate precisely to their engagement zones, understand their location during and after an 
engagement and return safely to home base. This makes maximum use of military and 
civilian navigation aids, using state-of-the-art on-board systems. 

Chapter 9 addresses the subject of weapons carriage and guidance to give an under- 
standing of the integrated weapon system. Individual weapons types are described, together 
with the systems required to ensure that they can be aimed and released to maximum effect. 

Chapter 10 deals with the vehicle management system; those systems that provide the 
platform with power, energy and management of basic platform control functions. Although 
provided as a separate control system today, it is inevitable that these functions will be 
absorbed into mission system processing in the future. 

Chapter 1 1 covers part of the human-machine interface - the displays in the cockpit and 
the mission crew areas that enable the crew to prosecute the operational mission. This 
chapter deals with the technology of displays and provides numerous examples of display 
systems in military applications. 

The authors believe that this volume will complete the set of companion volumes that 
describe the aircraft general, avionic and mission systems, as well as the way in which they 
are developed. This series provides a guide to the interested public, to students and to 
practitioners in the aerospace field. It should be recognised that this book, like its companion 
volumes, only scratches the surface of a series of complex topics. Within the book we have 
provided a comprehensive bibliography as a guide to specialised volumes dealing in detail 
with the topics outlined here, and it is to be hoped that the reader will continue to read on to 
understand aviation electronic systems. 

1 Military Roles 

1.1 Introduction 

The military were quick to seize upon the opportunities offered to them by an ability to leave 
the ground and gain an advantage of height. The initial attempts to make use of this advantage 
were by using tethered balloons as observation posts, and then as positions from which to 
direct artillery. The advent of a moving and powered platform allowed guns and, later, bombs 
to be carried, which led to air war between aircraft, and upon ground troops. Thus, fairly early 
in the history of the aircraft the main military roles of observation, interception and ground 
attack had been firmly established. These initial roles increased in sophistication and led to 
the development of more capable aircraft weapons, aircrew and tactics. 

Today the military are called upon to perform a wide variety of aviation roles using fixed- 
wing and rotary-wing aircraft. The roles largely define the type of aircraft because of the 
specialist nature of the task; however, there are a number of aircraft types that have been 
designed as multirole aircraft, or designed to change roles during the prosecution of a 
mission, the so-called swing-role type. 

The military roles that are in place today have emerged over many years of aerial combat 
experience. The long development timescales of the complex military aircraft have resulted 
in many types remaining in service long after their original introduction. Consequently, 
aircraft have adopted new roles as a result of role-fit weapons or mid-life updates. Many of 
the roles, particularly the intelligence gathering roles, have persisted after combat into the 
post-war stabilisation period and peacekeeping operations. 

The flexibility of weapons and methods of carrying weapons and the adaptability of 
sensors and avionic systems are what enables this situation to persist. Although many of the 
'traditional' roles still exist, there are signs that the changing nature of conflict may lead to 
new roles or alternative solutions. 

To a large extent these new roles and alternative solutions are being driven by advances in 
the technology of sensors and avionics. Ever more sensitive and effective sensor systems are 
capable of detecting targets, the use of stealth techniques increases the effectiveness of 
delivery platforms and the increased capability of on-board computing systems is extending 

Military Avionics Systems Ian Moir and Allan G. Seabridge 
© 2006 John Wiley & Sons, Ltd. ISBN: 0-470-01632-9 


Figure 1.1 Typical battlefield scenario and the major players. 

and speeding up the processing of data. The existence of these advances in the hands of 
enemies spurs on further development. 

This chapter will describe the roles that are required in the military defence environment. 
Some examples of avionic architectures will be described, along with examples of the types 
of aircraft in service today that perform the various roles. Other chapters in this book will 
deal with the detail of a number of military avionic systems. 

1.2 Air Superiority 

1.2.1 Role Description 

The primary aim of this role is to deny to an enemy the airspace over the battlefield, thus 
allowing ground attack aircraft a free rein in destroying ground targets and assisting ground 
forces, secure in the knowledge that the airborne threat has been suppressed. 

The air superiority aircraft is typically designed to enable the pilot to respond rapidly to a 
deployment call, climb to intercept or loiter on combat air patrol (CAP) and then to engage 
enemy targets, preferably beyond visual range. The aircraft should also have the capability to 
engage in close combat, or dogfight, with other aircraft should this prove to be necessary. For this 
to be successful, an extremely agile machine is necessary with 'carefree handling' capability. 

The systems must allow for accurate navigation, accurate identification of targets, 
prioritisation of targets, accurate weapon aiming capability and the ability to join the tactical 
communications network. 

A typical mission profile is shown in Figure 1.2. 

1.2.2 Key Performance Characteristics 

The air superiority aircraft is usually a highly manoeuvrable aircraft with a high Mach 
number capability and rapid climb rate. Many fighters are equipped with afterburning to 
allow Mach 2 capability, a power to weight ratio greater than 1 , allowing acceleration in a 
climb, and the ability to climb to beyond 60 000 ft. Some types are designed to operate from 



Figure 1.2 Air superiority mission profile. 



carriers and will be equipped for catapult launch and for steep approaches and arrestor wire 

Many modern fighters are unstable and have full authority flight control systems that are 
designed to allow the pilot to execute manoeuvres to envelope limits without fear of losing 
control or damaging the aircraft. This is known as 'carefree handling' capability. 

1.2.3 Crew Complement 

Usually single pilot, but some types employ a pilot and a rear-seat air electronics officer or 
navigator depending on the role. Trainers or conversion aircraft will have two seats for 
instructor and student. 

1.2.4 Systems Architecture 

A typical air superiority platform architecture is shown in Figure 1.3. Typical air superiority 
systems are listed in Table 1.1. 





Display Computing 


Nav Aids 






Mission Computing 

Data Link 


1 — Vehicle Management Computing 






FCS | | Fuel | | Electrics | | Hydraulics| 



Weapon Computing 

Air - Air Weapon & Stores 

Figure 1.3 Typical air superiority platform architecture. 



Table 1.1 Typical air superiority systems 



Mission system 










SHF SatCom 



Link 16 



Mission computer 


Data loader 

Landing aids 


LPI RadAlt 

Air data 

Digital map 


Displays and controls 


Helmet-mounted display 


Avionics data bus 

Mission system 

data bus 


missile - ASRAAM, 

Internal gun 

Weapons bus 

F*1 17 (Lockheed Martin) 

F/A-18 Hornet (US Department of Defence) 
Figure 1.4 Air superiority aircraft types. 

JSF & F-16 (Lockheed Martin) 


1.2.5 Air Superiority - Aircraft Types 

The various types of air superiority aircraft are as follows (Figure 1.4): 

• McDonnell Douglas F-4 Phantom; 

• English Electric Lightning; 

• Euro fighter Typhoon; 

• Panavia Tornado F-3; 

• Dassault Rafale; 

• Dassault Mirage 2000; 

• SAAB Gripen; 

• F-15; 

• F-16; 

• F-18; 

• Mig-21 Fishbed; 

• Mig-23 Flogger; 

• F-117. 

1.3 Ground Attack 

1.3. 1 Roie Description 

The ground attack role has been developed to assist the tactical situation on the battlefield. 
The pilot must be able to identify the right target among the ground clutter and multiplicity 
of targets and friendly units on the battlefield. The ability to designate targets by laser has 
enabled precision bombing to be adopted by the use of laser-guided bombs or 'smart' bombs. 
The role must enable fixed targets such as buildings, radar installations and missile sites, as 
well as mobile targets such as tanks, guns, convoys, ships and troop formations, to be 
detected, positively identified and engaged. 

This role includes close air support (CAS), where support is given to ground forces, often 
under their direction, where weapons will be deployed in close proximity to friendly forces. 

1.3.2 Key Performance Characteristics 

Depending on the target and the on-going military situation, the ground attack role may be 
performed by either fixed-wing or rotary- wing aircraft. A fixed-wing aircraft usually needs 
very fast, low-level performance with good ride qualities. It should also be reasonably agile 
to perform attack manoeuvres and take evasive action. Rotary-wing aircraft benefit from 
extreme low-level nap of the earth penetration, and the ability to loiter in natural ground 
cover - popping up when required to deliver a weapon. 


Take-off Land 

Figure 1.5 Ground attack mission profile. 








Display Computing 


Nav Aids 







Mission Computing 


Data Link 


1 — Vehicle Management Computing 


FCS | | Fuel | | Electrics | | Hydraulics| 



Weapon Computing 

Air - Ground Weapon & Stores 

Figure 1.6 Typical ground attack platform architecture. 

1.3.3 Crew Complement 

This role is usually conducted by two crew members, a pilot and a crew member to operate 
the sensors and weapons systems. The advent of smart weapons or cooperative target 
designation means that the mission can be conducted by a single crew, often a role 
designated to a fighter aircraft as a secondary role. 

1.3.4 Systems Architecture 

A typical ground attack platform architecture is shown in Figure 1.6. Typical ground attack 
systems are listed in Table 1.2. 

1.3.5 Ground Attack - Aircraft Types 

The various types of ground attack aircraft are as follows (Figure 1.7): 

• Sepecat Jaguar; 

• Panavia Tornado GR4; 

• Fairchild A- 10 Thunderbolt; 

• Apache; 

• Sukhoi Su-24 Fencer. 

1.4 Strategic Bomber 

1.4.1 Role Description 

The role of the strategic bomber is to penetrate deep into enemy territory and to carry out 
strikes that will weaken defences and undermine the morale of the troops. The strategic 



Table 1.2 Typical ground attack systems 



Mission system 




Air-to-ground missiles 







Free-fall bombs 


SHF SatCom 


Laser-guided bombs 


Link 16 

Laser designator 

Airfield denial 


Internal gun 


Mission recording 

gun pod 


Data loader 


Landing aids 



LPI RadAlt 

Air data 

Digital map 


Displays and controls 

Head-Up display 

Helmet-mounted display 


Avionics data bus 

Mission system data bus 

Weapons bus 

bomber was usually a very high-flying aircraft capable of carrying a large load of bombs 
which were released in a 'carpet bombing' pattern. The modern aircraft may choose to fly 
low and fast and rely on stealth to evade enemy radar defences. Different weapons may also 
be employed such as Cruise missiles and joint direct attack munition (JDAM). 

Lockheed F-16 (Lockheed Martin) 


;p T*!^ 

A-10 (US Department of Defence) 

Harrier GR7 (BAE SYSTEMS) 
Figure 1.7 Ground attack aircraft types. 




Figure 1.8 Strategic bomber mission profile. 



1.4.2 Key Performance Characteristics 

Strategic bomber aircraft attributes include high altitude cruise, long range and high payload 

1.4.3 Crew Complement 

The crew includes pilots, a navigator, an engineer and specialist mission crew. For very long 
missions a relief crew may be provided. 

1.4.4 Systems Architecture 

A typical strategic bomber platform architecture is shown in Figure 1.9. Typical strategic 
bomber systems are listed in Table 1.3. 


Display Computing 


Nav Aids 






Mission Computing 


Data Link 


1 — Vehicle Management Computing 


FCS | | Fuel | | Electrics | | Hydraulics | 



Weapon Computing 

Weapon & Stores 

Figure 1.9 Typical strategic bomber platform architecture. 



Table 1.3 Typical strategic bomber systems 



Mission system 








Free-fall bombs 




Laser-guided bombs 


SHF SatCom 


Airfield denial 


Link 16 


Cruise missiles 



Mission recording 


Data loader 

Landing aids 



LPI RadAlt 

Air data 

Digital map 


Displays and Controls 


Avionics data bus 

Mission system 
data bus 

Weapons bus 

1.4.5 Strategic Bomber - Aircraft Types 

The various types of strategic bomber aircraft are as follows (Figure 1.10): 

• Boeing B-52; 

• AVRO Vulcan; 

B-1B Lancer (US Department of Defence) 

B-2 Spirit (US Department of DeScncc) 

Figure 1.10 Strategic bomber aircraft types. 


• Northrop B-2; 

• Tupolev Tu-22M Backfire; 

• Tupolev TU-160 Blackjack; 

• General Dynamics F-lll. 

1.5 Maritime Patrol 

Over 60% of the earth's surface is covered by oceans - a natural resource that is exploited by 
many means: as a medium for transportation of cargo, as a source of food, as a means of 
deploying naval assets such as capital ships and submarines and for movement of men and 
materiel. It is also used for pleasure and for criminal purposes such as the smuggling of 
drugs, liquor, tobacco and illegal aliens. It is not surprising, therefore, that surveillance of the 
ocean's surface is of importance to military and paramilitary forces. 

The most practical way of carrying out surveillance or reconnaissance is by air, and the 
flexibility of the fixed-wing aircraft with its comparatively high speed, long range and 
excellent detection capability from high altitude made it an excellent complement to surface 
vessels in carrying out naval or policing duties. 

Over 90 years of development have led to the emergence of the maritime patrol aircraft 
(MPA) as one of the most complex of systems aircraft with a demanding role embracing a 
broad spectrum of tactical and strategic tasks, as well as support for civilian and 
humanitarian activities. 

The general MPA specification that has evolved calls for the ability to transit at high speed 
to a distant patrol area of interest, and then to remain in that area for a long time, carrying out 
searches for surface, subsurface or both types of activity. Operational requirements typically 
ask for an ability to fly over 800 miles to an area, remain on task for over 7 h, return to base 
and have sufficient fuel remaining to carry out a bad weather diversion. To perform such a 
task requires an aircraft weighing up to 100 1, a crew of 12 or more and a suite of electronic 
sensors and communication systems. 

1.5.1 Role Description 

The typical tasks that an MPA is called upon to perform include: 

Anti-surface unit warfare (ASuW) 

• Reconnaissance; 

• Shadowing; 

• Strike against surface vessels; 

• Tactical support of maritime strike aircraft; 

• Over-the-horizon targeting for friendly vessels; 

• Intelligence collection; 

• Communications relay; 

• Limited airborne early warning capability. 

Anti-submarine warfare (ASW) 

• Close air support to task forces and convoys; 

• Open ocean searches; 

• Extended tracking of submerged targets; 


• Deterrence of hostile submarines; 

• Cooperation with friendly submarines; 

• Intelligence collection. 

Search and rescue (SAR) 

• Location of survivors; 

• Dropping of survival equipment; 

• Scene-of- action commander for rescue operations; 

• Escort to rescue helicopters; 

• Cooperation with rescue services; 

• Escort of aircraft in difficulties. 

Exclusive economic zone protection 

• Oil rig surveillance; 

• Fishery protection; 

• Pollution detection and dispersal. 

Customs and excise cooperation 

• Anti-illegal immigration; 

• Anti-gun running; 

• Anti-terrorist operations; 

• Anti-drug smuggling. 

1.5.2 Anti-surface Unit Warfare (ASuW) Role 

MPAs take part in all aspects of the war at sea. In the role of anti-surface unit warfare, the 
MPA may carry out autonomous strikes against surface targets using free-fall bombs or 
stand-off weapons. Alternatively, it may be used to search, identify and shadow surface 
forces, remaining in contact but out of range of a surface ship's weapons for long periods of 
time. This can be performed overtly using the integral radar of the MPA to detect, classify 
and track targets, or covertly using passive electronic support measure systems to detect the 
ship's radars while remaining outside the ship's maximum radar detection range. 

Frequently it is necessary to shadow naval forces for days using a number of MPAs, each 
handing over the task to a relief aircraft at the end of its endurance on task. If there is then a 
requirement for specialist strike aircraft to carry out attacks against the ships under 
surveillance, then a cooperative attack is planned. The MPA can guide the attacking force 
accurately to suitable attack positions using its own radar, while the attacking aircraft can 
approach covertly under any defensive radar screen, only being detected by the air defence 
radar of the target at very short range. During the attack the MPA can carry out jamming of 
radar and communications to distract the surface ship's defensive tactics. 

The MPA can also use similar tactics to cooperate with attack helicopters, or to provide 
over-the-horizon targeting for surface missiles launched from friendly naval ships. 

1.5.3 Anti-submarine Warfare (ASW) Role 

Traditionally, submarines have waged strategic warfare by effectively blockading enemy 
countries, preventing military supplies, reinforcements and essential food and medical 


supplies from arriving by sea. The counter to the submarine campaigns was to regulate 
shipping by organising it into strictly disciplined convoys and concentrating naval forces to 
protect the convoys. However, as submarines became more effective and became organised 
into 'wolf packs', the escort ships found themselves outperformed and sometimes out- 

Furthermore, as detection ranges from improved sensors increased, surface ships did not 
have the speed necessary to exploit the detections and to supply a secure cordon around the 

The use of the aircraft in general reconnaissance of the sea surface was a natural evolution 
of its role in war. However, it was not until significant performance improvements in sensors, 
weapons, aircraft range and endurance and communications that the MPA could play its full 
part in integrated close support of surface forces. 

The MPA can often put itself at risk of friendly fire when joining a force. An unexpected 
aircraft contact in the war situation of jammed communications and strict emission control 
policies is often seen as a potential threat. As a result, complex joining procedures are 
adopted before closing within range of the defensive missile engagement zone of a friendly 

While on task, the tactics of the MPA are likely to involve searching at low altitude using a 
constant radar policy to force down submarines that may have closed the force and are 
attempting to get periscope ranging for an attack solution. The maritime radar is optimised to 
detect small contacts against a backgound of reflections or 'clutter' from the sea. Although 
the submarine will prefer to operate submerged, there are situations in which it must expose 
itself above the surface. For example, there are surface-to-surface anti-ship missiles that may 
require the submarine to surface partially in order to fire the missile, providing opportunities 
for detection by the MPA radar, leading to engagement by mines, torpedo or anti-ship 

Sonobuoys are also used to create barriers across a perceived threat axis, allowing the 
MPA to listen to noises that are characteristic of different submarine types. The experienced 
acoustic operator can distinguish between different types and between different operating 
states, and can detect noises from a submarine at rest on the sea bottom. Electronic support 
measures are used to detect the slightest transmission from an extended communications 
mast, and the maritime radar can detect a fleeting extension of a communications mast or a 
diesel air inlet mast. A contact is confirmed by overflying the contact and using the magnetic 
anomaly detector to distinguish a metallic mass from a shoal of fish or other natural 

The MPA crew work closely together with their individual sensors, while the tactical 
commander uses fused sensor data to view the whole surface picture. The MPA works closely 
with other assets to detect, locate, track and prosecute an attack over many days or weeks of 
continuous operation. Search patterns similar to those described in the next section are 
routinely used to conduct an efficient open ocean search. 

1.5.4 Search and Rescue (SAR) Role 

Search and rescue is the public and humanitarian aspect of military maritime patrol. The 
extension of the lifeboat service has grown from the original requirement for the military to 
provide a rescue capability for military aircrew who are forced to make emergency landings 



in the sea. The task fell to maritime patrol aircraft, rather than a specialist aircraft, for 
economic reasons, particularly in peacetime when the number of available aircraft is low. 
The requirements for an effective SAR aircraft have much in common with a maritime patrol 
aircraft and include: 

• Long range and endurance; 

• High transit speed and long loiter time; 

• All-weather operations capability; 

• Precise navigation system; 

• Comprehensive communications; 

• Extensive sensor suite; 

• Good visual platform; 

• Large crew complement; 

• Displays for tactical control for scene-of-action command; 

• Ability to carry large quantities of air-dropped survival equipment. 

A military base may keep at least one aircraft and crew on SAR standby 24 h a day, 365 days 
of the year. The aircraft is normally capable of taking to the air in response to a call for 
assistance within 2 h. As soon as the SAR aircraft is airborne, a new aircraft and crew are 
readied to provide cover for any subsequent calls for help. Datum Searches 

Where the distressed person or vessel has been able to pass a message to an emergency 
service, it is likely that a reasonably accurate datum is available and a number of datum 
search patterns can be used. Some example patterns are shown in Figure 1.11. 

If a very accurate position is obtained and the searching aircraft is able to arrive on the 
scene quickly after the incident, it is possible that the cloverleaf search pattern can be used. 
This is often the case when military aircrew are forced to eject from their aircraft which 

Creeping line 
ahead search 

~ n : > 

Clover leaf search 

Expanding square 

Figure 1.11 Search pattern examples. 


has been tracked by a surveillance radar. The fact that the SAR aircraft has arrived shortly 
after the ejection means that the survivors will not have drifted far from the last known 

However, if the SAR aircraft is delayed and the position of the datum is not so accurately 
defined, then there will be a need for the MPA to search out to a greater range. A more 
suitable search pattern would be the expanding square search, covering an area calculated by 
the SAR crew to enclose any inaccuracies in position or drift as a result of prevailing weather 
conditions. If the survivors are equipped with a personal locator beacon (PLB), then the 
aircraft will use its direction finding (DF) and homing systems to fly straight to the survivor's 
position. Area Searches 

It often happens that notification of an incident is received hours or even days after its 
occurrence. This usually happens when an aircraft or ship fails to make its scheduled 
position report. In these circumstances the SAR aircraft is likely to have to search a very 
large area for survivors or wreckage. Frequently the MPA would be one of a number of 
cooperating aircraft, helicopters, lifeboats, naval vessels and surface ships involved in the 
search. The procedure used for organisation, command and control are well tried and tested, 
and subject to international agreement and standardisation. Specific SAR radio frequencies 
are available, as are call signs and communication procedures that allow effective integration 
of civilian and military resources. 

The tactics employed by an MPA to search a very large area depend on all the factors 
described above. If the incident involves a missing military aircraft, the MPA will normally 
make a number of medium- to high-altitude passes over the area allocated for search. This is 
firstly to establish a shipping density using radar, and secondly to determine whether any 
survivors are using PLBs. In the event that no PLB signals are received, the MPA will 
descend to an altitude suitable for visual and radar search. The type of search will frequently 
be creeping line ahead (CLA), with track spacing determined by the calculation 
1.5 x estimated visibility. Where the track spacing is very short, in poor visibility, for 
example, the MPA will have difficulty in maintaining the integrity of the search because of 
the tight turning circles at the end of each loop. A modified creeping line ahead pattern can 
be used to compensate for this. Scene-of-Action Commander 

One aspect of the SAR operation where a modern MPA excels is as a scene-of-action 
commander (SAC). A typical example of this was the Piper Alpha oil platform disaster in the 
North Sea. A serious explosion and fire occurred and rescue forces were mobilised. There 
was no shortage of ships and helicopters to carry out rescue work, but, owing to poor 
visibility, burning oil on the sea surface and poor communications, there was a distinct 
danger of the rescue craft hampering or colliding with each other. A Nimrod MR2 from RAF 
Kinloss was directed to the scene and was able to establish firm control of all rescue forces 
using radar to deconflict the various helicopters, direct firefighting ships and keep the rescue 
control centre fully informed of developments. Each ship and helicopter was electronically 
tagged and displayed on the tactical display, while the area around the rig was divided into 
small search boxes and allocated to specific ships or helicopters. 


1.5.5 Exclusive Economic Zone Protection Oil and Gas Rig Patrols 

Many countries have a requirement to police their territorial waters, particularly those 
declared to be an economic exclusion zone (EEZ) and containing vital national resources 
such as oil, natural gas and fishing grounds. This task requires regular patrols over large areas 
of coastal waters by specialist aircraft cooperating with surface vessels to ensure the security 
of oil and gas installations which may be potential targets for terrorist action. Anti-pollution 

There is a requirement for early detection of pollution of the sea, whether by accidental 
discharge from ships or installations, or by illegal washing of tanks and bilges by merchant 
vessels. This is an ideal task for aerial surveillance with specially designed sideways looking 
radar (SLAR) and electrooptical devices using ultraviolet (UV) and infrared (IR) techniques 
to detect and measure the density and area of oil slicks at sea. Fishery Protection 

Fishing represents an increasingly important element of national economies, and there has 
been a growing tendency for fishing fleets to ignore international agreements for control and 
licensing of fishing within territorial waters. If a nation is to protect its own fishing rights, it 
must be capable of demonstrating a capability of detecting and apprehending any vessels 
fishing illegally within its EEZ. To achieve the very large-scale surveillance task effectively 
and in an acceptable timeframe dictates the use of an aircraft. There are obvious difficulties if 
an illegal fishing vessel detected by a fixed-wing aircraft needs to be arrested and brought 
before a court of law. 

Current tactics involve very close cooperation between aircraft and surface vessels, with 
the aircraft locating the offender, and the ships performing the arrest. However, there is often 
a delay of some hours before the surface ship can transit to the offender's position, who will 
no doubt have got rid of any evidence of illegal fishing and possible even have sailed out of 
the area. It is, therefore, incumbent on the EEZ patrol aircraft to obtain sufficient evidence to 
allow a reliable case to be brought before an international court, if the offender escapes 
immediate arrest. 

To have a good chance of winning a case in court, the aircraft must catch the offender in 
the act and obtain high-quality photography of the time and position at which each offence 
took place. The film needs to have a superimposed image of latitude, longitude, date and 
time to be used as secure evidence. Customs and Excise Cooperation 

Customs and excise operations are usually inshore, and a large military MPA may not be best 
suited to this kind of role. An alternative is a small twin- turboprop aircraft fitted with a 
minimum sensor set operating from a civilian airfield by a police or customs crew. This type 
of aircraft would not normally be armed, its role being surveillance and recording. However, 
the long-range aircraft will be called upon if a target vessel needs to be tracked over the 
high seas. 


High level transit 

Climb/ ;; —A / \ 

Take-off c p « cnrfarA - nn tn spa state & Land 

Sea surface - up to sea state 4 

Up to 1,000 nm 
Figure 1.12 Maritime patrol aircraft mission profile. 

1.5.6 Key Performance Characteristics 

The key performance characteristics are: 

• Long endurance; 

• Long range. 

1.5.7 Crew Complement 

The flight deck crew consists of two pilots who may alternate the roles of flying pilot 
and second officer throughout a long-duration mission in order to maintain vigilance. 
Some types may carry an engineer who will operate the general systems and usually acts 
as a monitor for height. On types expected to perform very long-duration missions, for 
example, with air-to-air refuelling this may be in excess of 20 h, a supernumerary pilot may 
be carried. 

The mission crew will be sized to operate the sensors and conduct the tactical 
mission. Crew sizes for a long-range maritime patrol and anti-submarine aircraft may 
exceed 10. 

1.5.8 Systems Architecture 

A typical maritime patrol aircraft platform architecture is shown in Figure 1.13. Typical 
MPA systems are listed in Table 1.4. 

1.5.9 MPA Aircraft Types 

The various types of MPA aircraft are as follows (Figure 1.14): 

• Shackleton; 

• BAE SYSTEMS Nimrod MR2; 






Display Computing 

■ i =p 

Nav Nav Aids Comms 




■— Vehicle Management Computing Propulsion 

FCS | | Fuel | | Electrics | | Hydraulics| 


I ■ ■ ■ I 


Mission Computing 


Data Link 



Weapon Computing 

Weapon & Stores 

Figure 1.13 Typical maritime patrol aircraft platform architecture. 

• Lockheed P-3C; 

• Lockheed S-3 Viking; 

• Dassault Atlantic; 

• Tupolev Tu-20; 

• Westland Sea King. 

Table 1.4 Typical MPA systems 



Mission system 



Maritime radar 

Anti-ship missiles 



Electrooptics turret 





Free-fall bombs 


SHF SatCom 


ASR kit 


Link 16 




Link 11 

Acoustic system 

Smoke markers 


Marine band 

Mission recording 




Data loader 


Landing aids 



Oceanographic database 

LPI RadAlt 

Mission computing 

Air data 

Mission crew workstations 

Digital map 

Intelligence databases 


Direction finding 


Displays and 



Avionics data bus 

Mission system data bus 

Weapons bus 




P-3 C Orion (US Department of Defence) 
Figure 1.14 Maritime patrol aircraft types. 

1.6 Battlefield Surveillance 

Detailed knowledge of the tactical scenario on the battlefield is of importance to military 
commanders and planners who need real-time intelligence of enemy and friendly force 
disposition, size and movement. Many commercial aircraft types have been converted to 
perform this role to complement specifically designed military types. The aircraft are 
equipped with a radar located on the upper or lower surface of the airframe that is designed 
to look obliquely at the ground. The aircraft flies a fixed pattern at a distance outside the 
range of enemy defences and detects fixed and moving contacts. These contacts are 
confirmed by using intelligence from other sensors or from remote intelligence databases 
to build up a picture of the battlefield and the disposition of enemy and friendly forces. The 
mission crew operate as a team to build up a surface picture, and can operate as an airborne 
command centre to direct operations such as air or ground strikes. 

1.6.1 Role Description 

A battlefield surveillance mission profile is shown in Figure 1.15. 

1.6.2 Key Performance Characteristics 

The key performance characteristics are high altitude, long range and a stable platform often 
based on a commercial airliner airframe. 



Search pattern 



Figure 1.15 Battlefield surveillance mission profile. 


1.6.3 Crew Complement 

The flight deck crew consists of two pilots who may alternate the roles of flying pilot 
and second officer throughout a long-duration mission in order to maintain vigilance. Some 
types may carry an engineer who will operate the general systems and usually acts as a 
monitor for height. On types expected to perform very long-duration missions, for example, 
with air-to-air refuelling this may be in excess of 20 h, a supernumerary pilot may be carried. 
The mission crew will be sized to operate the sensors and conduct the tactical mission. 
Crew sizes for a long-range, long-duration mission may exceed 10. 

1.6.4 Systems Architecture 

A typical battlefield surveillance platform architecture is shown in Figure 1.16. Typical 
battlefield surveillance systems are listed in Table 1.5. 

1.6.5 Battlefield Surveillance Aircraft Types 

The various types of battlefield surveillance aircraft are shown in Figure 1.17. 



Display Computing 


Nav Aids 







Mission Computing 


Data Link 



1 — Vehicle Management Computing Propulsion 

FCS | | Fuel | | Electrics | | Hydraulics| 

Figure 1.16 Typical battlefield surveillance platform architecture. 



Table 1.5 Typical battlefield surveillance systems 



Mission system 











SHF SatCom 



Link 16 

Moving target indicator 



Mission recording 


Data loader 

Landing aids 



Mission computer 

LPI RadAlt 

Mission crew workstations 

Air data 

Intelligence databases 

Digital map 


Displays and controls 


Avionics data bus 

Mission system data bus 

1.7 Airborne Early Warning 

1.7.1 Role Description 

Early detection and warning of airborne attack is important to give air superiority and 
defensive forces sufficient time to prepare a sound defence. It is also important to alert 
ground and naval forces of impending attack to allow for suitable defence, evasion or 
countermeasures action. 

E-B J STARS (US Department of Defence} 

ASTOR (Raytheon) 

Figure 1.17 Battlefield surveillance aircraft types. 



/ \ Descent 

Climb/ \ 

Take-off Land 

Figure 1.18 Airborne early warning mission profile. 

Operating from high altitude gives the airborne early warning (AEW) aircraft an 
advantage of detecting hostile aircraft at longer range than surface radar, which gives vital 
seconds for ground defence forces. 

1.7.2 Key Performance Characteristics 

A long-range, long-endurance aircraft enables a patrol pattern to be set up to cover a 
widesector area from which attack is most likely. A radar with a 360° scan, and a capability 
to look down and look up, provides detection of incoming low-level and high-altitude 
attack. The radar will usually be integrated with an interrogator to enable friendly aircraft 
to be positively identified. The aircraft will also act as an airborne command post, 
controlling all airborne movements in the tactical area, compiling intelligence and 
providing near real-time displays of the tactical situation to both local forces and remote 

1.7.3 Crew Complement 

The flight deck crew consists of two pilots who may alternate the roles of flying pilot 
and second officer throughout a long-duration mission in order to maintain vigilance. 
Somevtypes may carry an engineer who will operate the general systems and usually acts 
as a monitor for height. On types expected to perform very long-duration missions, for 
example, with air-to-air refuelling this may be in excess of 20 h, a supernumerary pilot may 
be carried. 

The mission crew will be sized to operate the sensors and conduct the tactical mission. 
Crew sizes for a long-range, long-duration mission may exceed 10. 

1 . 7.4 Systems Architecture 

A typical airborne early warning platform architecture is shown in Figure 1.19. Typical 
AEW systems are listed in Table 1.6. 




Display Computing 


Nav Aids 







Mission Computing 


Data Link 


■— Vehicle Management Computing 


FCS | | Fuel | | Electrics | | Hydraulics| 

Figure 1.19 Typical airborne early warning platform architecture. 

1. 7.5 AEW Aircraft Types 

The various types of AEW aircraft are as follows (Figure 1.20): 

• Grumman E-2 Hawkey e; 

• Boeing E-3 Sentry; 

• Lockheed P-3 AEW; 

• Tupolev Tu-126 AEW; 

• Westland Sea King. 


Table 1.6 Typical AEW systems 



Mission system 



AEW radar 







SHF SatCom 



Link 16 



Mission recording 


Data loader 

Landing aids 

Mission computer 


Mission crew workstations 

LPI RadAlt 

Intelligence databases 

Air data 

Digital map 


Displays and controls 


Avionics data bus 

Mission system data bus 



E-3 Sentry (US Department of Defence) 
Figure 1.20 Airborne early warning aircraft types. 

P-3 AEW (Lockheed Martin) 

1.8 Electronic Warfare 

1.8. 1 Role Description 

Electronic warfare (EW) refers to a number of related subjects across a wide spectrum of 

• Electronic countermeasures (ECM); 

• Electronic support measures (ESM); 

• Signals intelligence (SIGINT); 

• Electronic counter-countermeasures (ECCM). 

1.8.2 Electronic Countermeasures 

Electronic countermeasures or jamming are a commonly used form of electronic warfare 
used to disrupt communications or defence radars. In noise jamming, radio frequency at the 
same frequency as a target emitter/receiver is modulated and transmitted at the target. 
Depending on the transmitted power level, it is capable of denying range information to the 
target or degrading communications to an unacceptable level. As the jamming power 
increases, or the range between the jammer and the receiver reduces, the jamming can 
become sufficiently strong to break down the directional properties of the target antenna. In 
this case both range and directional information can be denied. 

Deception jamming is a more subtle form of countermeasure, where the intention is to 
confuse the enemy as to the correct bearing, range and number of targets. It has major 
implications in the countering of weapon guidance systems, where the technique of range 
gate or velocity gate stealing can be particularly effective. 

Chaff is a passive ECM application, in which the transmitter energy of a threat radar is 
reflected to create false targets. Chaff can be used in a distraction sense by dispensing small 
discrete bundles to create an impression of specific small targets to confuse a radar or seduce 


a missile guidance system. Large clouds of chaff can be dispensed in a confusion sense 
completely to obscure one's own position. 

Jamming is a key tactical role on the battlefield and is often carried out by fast jets 
equipped with jamming equipment. These aircraft are known in the United States as Wild 
Weasel squadrons. 

1.8.3 Electronic Support Measures 

Electronic support measures (ESM) comprise the division of electronic warfare involving 
actions taken to intercept, locate, record and analyse radiated electromagnetic energy for the 
purpose of gaining tactical advantage. An important advantage of ESM when used as a 
sensor is that they are completely passive. They also provide the potential for detecting 
enemy radars at much greater ranges than the detection ranges of these radars. Signals of 
ESM interest are usually radar systems. However, they can also include communications, 
guidance or navigational emissions in the radio-frequency spectrum, as well as laser 
emissions and infrared radiation in the electrooptics field. 

Electronic intelligence information is required for both short- and medium- term planning 
and also for immediate tactical use in support of offensive and defensive EW operations. 
ESM are primarily used to support activities such as: 

1. Threat warning - the short-term or tactical activity of ESM concerned with detecting 
transmissions that pose a physical threat. A typical example is the use of radar warning 
receivers (RWR) to provide an indication of impending attack by fighter or surface-to-air 

2. Target acquisition - the presence of radar systems can indicate the existence of a target, or 
can assist in the identification of a radar-defended target. An example of this is a maritime 
patrol aircraft detecting and classifying surface ships by ESM. 

3. Homing - an attack radar homing passively onto land or naval base radar-defended 

1.8.4 Signals Intelligence (SIGINT) 

Signals intelligence consists of a number of different but related activities that are usually 
complementary in their employment and results. SIGINT is acknowledged as being used by 
most military forces and governments. Security surrounds its exact operational deployment 
and the degree of capability available to governments. SIGINT consists of three major 

1. Communications intelligence (COMINT) which is achieved by the interception of 
communication signals of all types - telegraphy, voice or data - and obtaining 
intelligence on a prospective enemy's intentions, capabilities and military preparedness. 
Frequently, the text of messages is enciphered and cannot be read immediately. However, 
there is still a great deal of intelligence to be derived from signals traffic analysis and 
direction finding which can provide both tactical and strategic advantages. Patterns of 
communication can be used to identify the state of readiness and location of participants. 


Sudden communication activity may indicate battle readiness or changes of plan. Silence 
can indicate departure of forces from an area, or deliberate radio silence prior to an attack. 
A database of such communications activity is useful in establishing a potential enemy's 
radio discipline and movements. Selective or broad band jamming can be deployed to 
restrict useful communications. 

2. Radar intelligence (RADINT) which uses standard or special-purpose radar systems to 
obtain intelligence on an enemy's capabilities, deployments and intentions. These radar 
systems may be space, air, sea or land based. RADINT is collected by an electronic 
support measures (ESM) system employing a number of sensitive antennas that are able 
to detect radar signals in different bands. As well as detecting the signal, the ESM also 
establishes a precise direction of arrival of the signal. Analysis of the signal character- 
istics such as frequency, pulse duration, amplitude and the spacing of main power lobes 
and side lobes will identify a particular type of radar. Continued collection of analysed 
data mapped onto the types of platform carrying different radar systems enables an 
experienced EW operator to identify a particular ship, aircraft or land-based system type. 
There are claims that experienced operators can even identify an individual platform by 
its radar transmitter characteristics. A successfully managed EW campaign can identify 
radar types and their exact locations for subsequent database update or for selective 
jamming. This technique is known as 'fingerprinting' and it requires some very clever 
waveform analysis equipment. 

3. ELINT Electronic Intelligence involves the interception and analysis of non- 
communication radio-frequency signals, usually radar, to obtain many aspects of a 
nation's intentions and capabilities such as technological progress, military preparedness, 
orders of battle, military competence, intentions etc. 

1.8.5 Key Performance Characteristics 

For high-altitude, long-duration intelligence gathering, a high-altitude stable platform is 
required. For cooperative 'Wild Weasel' support, a fast, low-level aircraft is required. 

1.8.6 Crew Complement 

The flight deck crew consists of two pilots who may alternate the roles of flying pilot 
and second officer throughout a long-duration mission in order to maintain vigilance. Some 


C ^^S^ — J- 



Take-off Land 

Figure 1.21 Electronic warfare mission profile. 




Display Computing 

Nav NavAids Comms 






Mission Computing 


Data Link 



■—Vehicle Management Computing Propulsion 

FCS | | Fuel | | Electrics | | Hydraulics| 

Figure 1.22 Typical electronic warfare platform architecture. 

types may carry an engineer who will operate the general systems and usually acts as a 
monitor for height. On types expected to perform very long-duration missions, for 
example, with air-to-air refuelling this may be in excess of 20 h, a supernumerary pilot 
may be carried. 

The mission crew will be sized to operate the sensors and conduct the tactical mission. 
Crew sizes for a long-range, long-duration mission may exceed 10. A typical wild weasel 
aircraft will have a one- or two-person crew. 

1.8. 7 Systems Architecture 

A typical electronic warfare platform architecture is shown in Figure 1.22. Typical EW 
systems are listed in Table 1.7. 

1.8.8 Example Aircraft Types 

Various types of electronic warfare aircraft are shown in Figure 1.23. 

1.9 Photographic Reconnaissance 

1.9.1 Role Description 

Photographic imagery (IMINT) can be used to confirm SIGINT intelligence by providing a 
high-resolution permanent image using ground-mapping cameras. Such images can be 
analysed by specialists and used to confirm the types identified by SIGINT. The images 
may provide further intelligence by providing numbers, groups and battalion identification 
marks, as well as the deployment of troops. 

Although often obtained from satellite imaging systems, IMINT can also be collected by 
high- and low-altitude fixed-wing aircraft overflying the battlespace in wartime, or by 



Table 1.7 Typical EW systems 



Mission system 



Maritime radar 



Electrooptics turret 





SHF SatCom 



Link 16 

Active jamming 



Mission recording 


Data loader 

Landing aids 

Mission computer 


Mission crew workstations 

LPI RadAlt 

Intelligence databases 

Air data 

Digital map 

Displays and controls 


Avionics data bus 

Mission system data bus 

overflying territory in peacetime. This activity is risky, leading to loss of aircraft from missile 
attack, or leading to diplomatic incidents. An example of this is the loss of the US U-2 
aircraft in 1960 by surface-to-air missile over the Soviet Union, and over Cuba in 1962. 

Cameras mounted in the fuselage or in a pod provide forward, rearward, downward and 
oblique looking views. Oblique cameras provide the opportunity to film territory without the 
need to conduct direct overflights - a necessity to prevent diplomatic incidents. The mission 
computer determines the best rate of height and speed to optimise the film frame rate and 
varies lens aperture to match light conditions. High-speed focal plane shutters reduce the 

EA-6 Prowler (US Department 
of Defence) 

EF-111A Raven (US Department 
of Defence) 

Nimrod R Mk1 (Federation of 
American Scientists) 

RC-1 35 (US Department of Defence) 

Figure 1.23 Electronic warfare aircraft types. 





Take-off Land 

Figure 1.24 Photographic reconnaissance mission profile. 

effect of vibration. A window in the airframe or pod is heated to prevent icing or 
condensation. Wet-film and digital cameras are often complemented by sideways looking 
synthetic aperture radar and infrared cameras. The reconnaissance pod fitted to Jaguar 
contained both wet-film and IR cameras, and the Tornado Raptor pod provides real-time 
high-resolution images from its CCD camera. 

Low-altitude, high-speed passes over terrain by fast jets using cameras installed in 
underwing pods can obtain tactical information at very short notice, and this can be used 
in conjunction with gun and head-up display (HUD) cameras to provide an instant battlefield 
picture and confirmation of inflicted damage. 

1.9.2 Key Performance Characteristics 

Often a very high-altitude aircraft to allow enemy territory overflights beyond radar detection 
and missile engagement range. To enable large areas of terrain to be covered, extremely high 
speed is an advantage. To obtain pictures of battlefield damage, low-level high-speed flights 
may be used. 

1.9.3 Crew Complement 

Usually a single crew, although aircraft carrying out long-duration precision mapping type 
photography may carry camera operators and film processors. 

1.9.4 Systems Architecture 

A typical photographic reconnaissance platform architecture is shown in Figure 1.25. 
Typical photographic reconnaissance systems are listed in Table 1.8. 

1.9.5 Typical Aircraft Types 

Typical photographic reconnaissance aircraft types are as follows (Figure 1.26): 

• Lockheed U2; 

• Lockheed SR71 Blackbird; 




Display Computing 

Nav NavAids Comms 





Mission Computing 




Data Link 


■—Vehicle Management Computing Propulsion 

FCS | | Fuel | | Electrics | | Hydraulics| 

Figure 1.25 Typical photographic reconnaissance platform architecture. 

• Canberra PR9; 

• Jaguar GR1; 

• Tornado GR4. 

1.10 Air-to-Air Refuelling 

1. 10. 1 Role Description 

Military aircraft of nearly all types find it necessary to extend their range or endurance. This 
may be the result of the global nature of conflict which is giving rise to long-range missions 

Table 1.8 Typical photo reconnaissance systems 



Mission system 



High-resolution cameras 



Film storage 



Film processing 


SHF SatCom 

Infrared cameras 


Link 16 

Optional FLIR 


Optional SAR 



Landing aids 


Air data 

Digital map 


Displays and controls 


Avionics data bus 



Lockheed SR-71 (Lockheed Martin) 

Lockheed U-2 (Lockheed Martin) 

Canberra PR-9 {BAE SYSTEMS) 
Figure 1.26 Photographic reconnaissance aircraft types. 

or long ferry flights over large oceanic distances, such as the United Kingdom to Falklands 
flights. This is especially true for large military aircraft which may be sent to an area to 
establish a presence or force projection as a diplomatic move. If there are no intermediate 
airfields available, or no airfields readily accessible in friendly nations, then air-to-air 
refuelling becomes a necessity. 

Air superiority aircraft or aircraft on combat air patrol are better refuelled close to their CAP, 
rather than returning to base to refuel. Not only are time and fuel saved, but so is the need 
constantly to send replacement aircraft to ensure a continuous presence on patrol. 

Incidentally, aircraft companies have extended the duration of their test missions by the 
use of air-to-air refuelling, extending test flights from a nominal 1 h to up to 8 h. 

Most tankers in operation today are conversions of well established commercial aircraft 
types such as the Boeing 707, Lockheed L1011 Tristar and the BAE SYSTEMS (Vickers) 
VC10. Military types have also been used, such as the Boeing C-17 and the Handley Page 




Figure 1.27 Air-to-air tanker mission profile. 




There are two types of refuelling method in widespread use: 

1. The probe and drogue method is widely used by most air forces. The tanker aircraft 
deploys and trails one or more hoses, each with a 'basket' at the end. The receiving 
aircraft homes on to the tanker, extends its refuelling probe and engages the probe tip in 
the basket. When positive contact is made, refuelling commences, the aircraft matching 
their speed - a tricky manoeuvre with the tanker becoming lighter and the receiver 
heavier as fuel is transferred. It is possible to refuel up to three receivers simultane- 

2. Most US Air Force tanker aircraft are fitted with a boom that is deployed from the tanker. 
The receiver aircraft homes on to the tanker and flies close. The tanker boom operator 
engages the boom tip into a receptacle on the aircraft. 

1. 10.2 Key Performance Characteristics 

Long range, capable of carrying a large fuel load and the ability to fly a stable refuelling 
pattern. Often a converted commercial airliner equipped with additional fuel tanks and 
loading equipment. 

1 . 10.3 Crew Complement 

Flight deck crew and boom operators. 

1. 10.4 Systems Architecture 

A typical air-to-air tanker platform architecture is shown in Figure 1.28. Typical tanker 
systems are listed in Table 1.9. 



Display Computing 
■ i =p 

Nav Nav Aids Comms 


Refuelling workstation 


Mission Computing 


Data Link 



1 — Vehicle Management Computing Propulsion 

FCS | | Fuel | | Electrics | | Hydraulics| 

Figure 1.28 Typical air-to-air tanker platform architecture. 



Table 1.9 Typical tanker 




Mission system Weapons 









SHF SatCom 



Link 16 



Mission recording 


Data loader 

Landing aids 


Refuelling equipment 

LPI RadAlt 

Air data 

Digital map 


Displays and controls 


Avionics data bus 

Mission system data bus 

1.10.5 Aircraft Types 

The various types of air-to-air tanker aircraft are as follows (Figure 1.29): 

• KC 135; 

• KC-10 (DC- 10 militarised version); 

• Lockheed L- 1011 Tristar; 

• VC10. 


Figure 1.29 Air-to-air tanker aircraft types. 

KC-135 (US Department of Defence) 


1.11 Troop/Materiel Transport 

1.11.1 Role Description 

The global nature of conflict and peacekeeping operations demands the movement of troops 
and materiel to remote theatres of operation. While the bulk of the task is performed by 
marine vessels, rapid deployment in order to establish a military position requires fast aerial 
transport. This is often seen by the military as force projection - the ability to establish a 
rapid presence in times of tension. 

Troops can be carried with their personal arms and equipment by fixed-wing or rotary- 
wing aircraft, disembarking from a landed platform or by parachute. Stores, ammunition and 
light vehicles can be carried by the same types and either unloaded on the ground or dropped 
on pallets during a low-speed, very low-height transit. They may also be dropped by 
parachute if the terrain or the defences do not permit a low-speed pass. 

Larger and heavier items of equipment such as artillery, trucks, personnel carriers or 
helicopters are unloaded from large short take-off and landing (STOL) transport aircraft. 

Transport aircraft are used to assist in evacuations from battlefields and also to assist in 
humanitarian operations by carrying foodstuffs and relief supplies, as well as evacuating 

1.11.2 Key Performance Characteristics 

A large volume of cargo-carrying space with secure stowage and a rapid deployment ramp is 
required, combined with the ability to lift heavy loads from short poorly prepared strips. 
Long-range transit with air-to-air refuelling if necessary. Good accurate navigation and the 
ability to operate with mobile landing aids into new designated landing areas. The ability to 
operate in poor weather conditions and from ill-prepared airstrips. 

1. 1 1.3 Crew Complement 

Flight deck crew of pilots and flight engineer, with crew to provide loading, unloading and 
paratroop deployment services. 

1. 1 1.4 Systems Architecture 

A typical troop/material transport platform architecture is shown in Figure 1.31. Typical 
transport systems are listed in Table 1.10. 

/ \ Descent 


Take-off Land 

Figure 1.30 Troop/materiel transport mission profile. 




Display Computing 

Nav Nav Aids Comms 


1 — Vehicle Management Computing Propulsion 

FCS | | Fuel | | Electrics | | Hydraulics | 

Figure 1.31 Typical troop/materiel transport platform architecture. 

1. 1 1.5 Typical Aircraft Types 

Typical types of troop/materiel transport aircraft are as follows (Figure 1.32): 

• Lockheed C- 130; 

• Boeing C-17; 

• Lockheed C-5 Galaxy; 

• Antonov An- 124 Condor; 

Table 1.10 Typical transport systems 


Mission system 






Electrooptics turret 





SHF SatCom 



Link 16 




Data loader 

Landing aids 


Station keeping 

LPI RadAlt 

Materiel deployment 

Air data 

Paratroop deployment 

Digital map 


Displays and controls 


Avionics data bus 



Lockheed C-130 (Lockheed Martin) 

C-5 Galaxy (US Department 

of Defence) 

C-17 (US Department of Defence) 
Figure 1.32 Troop/materiel transport aircraft types. 

• Transall C-16; 

• FiatG-222; 

• Chinook; 

• Puma; 

1.12 Unmanned Air Vehicles 

Many aircraft have evolved to perform the roles described in the previous paragraphs, with 
ever-increasing performance and sophistication. Aircraft, sensors and systems have become 
more capable, as have the crews who perform the flying and sensor management tasks. This 
has led to the situation where the cost of an individual aircraft is very high - tens or even 
hundreds of millions of dollars per copy, and the value of the crew in terms of training time 
and experience is also very high. A need to operate over the battlefield without risking 
human life or expensive assets has led to the development of unmanned aircraft. 

Unmanned air vehicles [or uninhabited aerial vehicles (UAVs)] have been used for some 
time as pilotless target drones. Often these were obsolete aircraft converted for use as an 
aerial target. Aircraft such as Meteor, Sea Vixen and Canberrra have been used in the 
United Kingdom and F4 Phantom in the United States. Special-purpose drones were 
constructed using aircraft parts or designed specifically for the task. Most UAVs today 
are designed specifically to fulfil a particular role. Without the need to carry a crew, the 
vehicle can be relatively small and simple, hence reducing cost. The reduced mass also 
leads to vehicles that can operate for very long duration and at high altitudes. There are 
examples of UAVs that can operate for over 50 h continuously at altitudes in excess of 
70 000 ft. 

Some UAVs are small, and careful low-observability design makes them extremely 
difficult to detect on radar. 

An ideal role for the unmanned air vehicle is aerial observation and intelligence gathering. 
Optical sensors using infrared or TV are used to transmit information to a relay for visual 



image intelligence (IMINT) to record enemy asset deployment or bomb damage assessment, 
while antennas are used to obtain communications intelligence (COMINT) or signals 
intelligence (SIGINT). Some UAVs carry wetfilm cameras, the film being processed on 
return to base. 

Some simple UAVs are used as decoys to entice defence radars to illuminate them as 
targets. This enables a force to test the alertness of the defensive screen, and also positively 
to identify the radar type, and hence defence weapon, in order to deploy suitable counter- 

All UAVs require a ground station for operators and mission planners, with associated data 
links for control and data communication. The human factor elements of the traditional 
cockpit design must be incorporated into the ground station control panels. 

Some UAVs can operate autonomously for large portions of a mission without ground 
station intervention. It is anticipated, as technology evolves, that many of the roles and 
missions described in the previous paragraphs will be subsumed by role-specific UAVs. 
Current-generation cruise missiles are an example of a capable UAV, although they are 

Some examples of UAVs are listed in Table 1.11 and shown in Figure 1.33. Unmanned 
combat air vehicles (UCAVs) are now emerging that are capable of carrying out offensive 
operations. This covers the delivery of air-launched weapons to attack ground targets, or a 
capability of engaging fighter aircraft and launching air-to-air missiles. 

Table 1.11 Examples of unmanned air vehicles 



Country of origin 

Typical use 



Target drone 


Teledyne Ryan 








Israeli Aircraft 



General Atomics 



Mirach 26 





South Africa 


Flight Refuelling 


Target drone 

CL227 Sentinel 




Yak 60 




TU141 Strizh 





Xian ASN 



Skoja III 


Czech Republic/ 


Fox ATI/4 





STN Atlas 



Global Hawk 

Teledyne Ryan 




Flight Refuelling 



Storm Shadow 



Global Hawk (US Department of Defence} 

Predator (US Department of Defence) 

Dark Star 

(Federation of 

American Scientists) 


(Federation of American 



(Federation of 



Figure 1.33 Examples of unmanned air vehicles. 

Further developments include the design of micro-UAVs that act in flocks to provide a 
multiple redundant sensor system capability. These devices can be programmed to mimic 
bird flocking behaviour so that a number of UAVs will follow a leader. 

1.13 Training 

Training of aircrew is an important task, from primary training through conversion to type to 
refresher training to maintain combat readiness. Training is accomplished by a variety of 
means. Initial training (ab initio) is usually carried out using a dynamic simulator to gain 
familiarisation before transferring to a single-engine light aircraft and then to a single-engine 
jet trainer. Conversion to an operational type again uses a simulator before transferring to a 
two-seat version, and then to solo. This is usually carried out on an operational conversion 
unit (OCU), a squadron staffed by qualified flying instructors. 

Routine mission and tactics training can be performed on dynamic simulators, designed to 
have a high degree of fidelity in six degrees of motion and an outside world real-time picture. 
However, realism is an important aspect of training, and nations have designated areas of 
their territory for low-level or combat training. While this includes ranges in restricted areas 
for very high-speed flight, live weapons delivery and electronic warfare training, much flying 
takes place in airspace that conflicts with other users such as the general public, private 
flying, commercial helicopters and leisure flying of balloons, hang-gliding and microlights. 
Training in these circumstances is essential for combat readiness but requires delicate 
handling to maintain public sympathy. For this reason, many air forces have resident public 
liaison officers to deal with public complaints. 



T-38 Talon (US Department of Defence) 

Typhoon simulator (BAE SYSTEMS) 


Figure 1.34 Training system aircraft types. 

Simpler simulations of parts of the mission task are often undertaken in classrooms, 
especially for cockpit familiarisation or for mission systems crews to practise team tasks 
such as MPA or AEW. These are known as part task trainers. In this situation, real missions 
can be replayed and unexpected scenarios generated. Occasionally, flying classrooms using 
modified business aircraft to seat a number of trainees at workstations are used. 

The training role also includes the provision of aerial targets. These may be implemented 
as remotely piloted drones such as converted obsolete aircraft or specially designed target 
vehicles. Banner targets can be towed by converted operational aircraft for air-to-air gunnery 


Special Forces Chinook (Ian Moir) 

Figure 1.35 Examples of special role types. 


practice. Converted small business jets have been used to carry underwing pods equipped 
with radio-frequency transmitters for use in a simulated EW environment. 

1. 13. 1 Typical Aircraft Types 

Typical types of aircraft as follows: 


• Aermacchi MB339; 

• Aermacchi MB326; 

• Dassault/Dornier Alphajet; 

• Embraer EMB-312Tucano; 

• Pilatus/BAE SYSTEMS PC-9; 

• Fairchild-Republic T-46; 

• -endNorthrop T-38 Talon. 

1.14 Special Roles 

Military aircraft are often called on to perform roles beyond their original design intention. 
This may be for research and development of sensors and systems, development of new 
tactics, for intelligence gathering or for peaceful information-gathering missions. Conversion 
to these roles may be by major modification to the type or by adding payload internally or 
externally - such as under- wing pods. 

Many of these special roles are, of course, so special that their existence is not revealed to 
the general public. 

1. 14. 1 Examples of Special Roles 

Examples of special roles are as follows: 

• Gunship; 

• Air ambulance; 

• Arctic survey; 

• Nuclear contamination detection; 

• Biological/chemical sensing; 

• Meteorological research; 

• Covert troop deployments; 

• Remotely piloted vehicle (RPV) aerial launch. 

1.15 Summary 

This list of roles is by no means exhaustive, the roles may also be known by other titles in 
different nations and different air forces. The nature of the roles and the system solutions to 
enable them to be fulfilled will also vary according to the perceived threat, the prevailing 
defence policy, local politics and the national air force inventory. 


What they have in common is a comprehensive system and sensor solution that will 
include the following elementes: 

• Communications and navigation; 

• Mission sensors; 

• Mission computing and data communications; 

• Weapons and stores carriage and release; 

• Man-machine interface. 

The following chapters will describe the elements and will provide some examples of typical 
practical implementations. 

Further Reading 

Adamy, D.A. (2003) EW 101 A First Course in Electronic Warfare, Artech House. 

Airey, T.E. and Berlin, G.L. (1985) Fundamentals of Remote Sensing and Airphoto Interpretation, 

Bamford, J. (2001) Body of Secrets, Century. 

Beaver, P. (ed.) (1989) The Encyclopaedia of Aviation, Octopus Books Ltd. 
Budiansky, S. (2004) Air Power, Viking. 

Burberry, R.A. (1992) VHF and UHF Antennas, Peter Pergrinus. 
Gardner, W.J.R. (1996) Anti-submarine Warfare, Brassey's. 

Hughes-Wilson, J. (1999) Military Intelligence Blunders, Robinson Publishing Ltd. 
Jukes, M. (2004) Aircraft Display Systems, Professional Engineering Publishing. 
Kayton, M. and Fried, W.R. (1997) Avionics Navigation Systems, John Wiley. 
McDaid, H. and Oliver, D. (1997) Robot Warriors - The Top Secret History of the Pilotless Plane, Orion 

Oxlee, G.J. (1997) Aerospace Reconnaissance, Brassey's. 

Poisel, R.A. (2003) Introduction to Communication Electronic Warfare Systems. Artech House. 
Schleher, CD. (1978) MTI Radar, Artech House. 

Schleher, C. (1999) Electronic Warfare in the Information Age, Artech House. 
Skolnik, M.I. (1980) Introduction to Radar Systems, McGraw-Hill. 
Stimson, G.W. (1998) Introduction to Airborne Radar, 2nd edn, SciTech Publishing Inc. 
Thornborough, A.M. and Mormillo, F.B. (2002) Iron Hand - Smashing the Enemy's Air Defences, 

Patrick Stephens. 
Urick, R.J. (1983) Sound Propagation in the Sea, Peninsula Publishers. 
Urick, R.J. (1982) Principles of Underwater Sound, Peninsula Publishers 
Van Brunt, L.B. (1995) Applied ECM, EW Engineering Inc. 
Walton, J.D. (1970) Radome Engineering Handbook, Marcel Dekker. 

2 Technology and 

2.1 Evolution of Avionics Architectures 

The introduction gave examples of how an avionics architecture may be structured and 
explained that, in the main, aircraft level and equipment level architectures would be 
addressed within the book. 

The application of avionics technology to military aircraft has occurred rapidly as aircraft 
performance has increased. The availability of reliable turbojet engines gave a huge 
performance boost to both military and civil operators alike. New and powerful sensors 
such as multimode radars, electrooptics and other advanced sensors have provided an 
immense capability to modern military aircraft that enables them better to perform their 
roles. At the same time, the advances in digital avionics technology in the areas of processing 
and digital communication - by means of data buses - enabled the new systems to be 
integrated on a much higher scale. 

This chapter addresses some of the developments and technology drivers that have led to 
many of today's advanced platforms; in many cases significant barriers have defeated the 
attainment of the original aims. These may be summarised as: 

• The evolution of avionics architectures from analogue to totally integrated digital 
implementations (section 2.1). 

• Aerospace- specific data buses - the 'electronic string' that binds avionics systems 
together (section 2.2). 

• A description of the joint industrial avionics working group (JIAWG) architecture; 
originally conceived as a US triservice standard (section 2.3). 

• An overview of commerical off-the-shelf (COTS) data buses which are providing a new 
cost-effective means of integration for the latest avionics systems (section 2.4). 

• An overview of the latest developments in software real-time operating systems (RTOS); 
these developments are leading to increasing portability of software while also providing 
integrity and security partitioning (section 2.5). 

Military Avionics Systems Ian Moir and Allan G. Seabridge 
© 2006 John Wiley & Sons, Ltd. ISBN: 0-470-01632-9 





Power, Cost, 




□ □ 



Decreasing Weight 

Volume, Power 








Figure 2.1 Evolution of avionics architectures. 

• The increasing integration of radio-frequency subsystems (section 2.6). 

• The evolution of the Pave Pace/JSF shared aperture radio-frequency architecture (section 2.7). 

To utilise these improvements, the aircraft avionics systems rapidly grew in terms of capability 
and complexity. Technology brought improvements in terms of increased performance, 
computing power, complexity and reliability, although all at an increase in cost. Other benefits 
included a decrease in weight, volume, power consumption and wiring (Figure 2.1). 

Figure 2.1 portrays how avionics architectures for modern fighter aircraft have evolved 
from the 1960s to the present day. The key architectural steps during this time have been: 

• Distributed analogue architecture; 

• Distributed digital architecture; 

• Federated digital architecture; 

• Integrated modular architecture; also digital. 

The evolution of these different architectures has been shaped in the main by aircraft-level 
design drivers (Moir and Seabridge, 2004). Their capabilities and performance have been both 
enabled and constrained by the avionics technology building blocks available at the time. As 
shown in Figure 2.1, there have been changes in many characteristics throughout the period. 
Prior to the 1960s, military aircraft had been manufactured in a similar way to their World 
War II (WWII) forebears. Avionics units were analogue in nature and interconnected via a 
considerable quantity of aircraft wiring. Key advances were enabled by the advent of digital 
computing technology in the 1960s which first found application in the architectures reaching 
fruition during the 1970s. The availability of digital computers that could be adopted for the 
rugged and demanding environment of the aerospace application brought computing power 
and accuracies that had not been available during the analogue era. The development of serial 



digital data buses greatly eased the interconnection and transfer of data between the major 
system units. In the early days this was achieved by means of fairly slow half-duplex 
(unidirectional), point-to-point digital links such as ARINC 429 and Tornado serial data link. 

The arrival of microelectronics technology and the first integrated circuits (ICs) enabled 
digital computing techniques to be applied to many more systems around the aircraft. At 
the same time, more powerful data buses such as MIL-STD-1553B provided a full-duplex 
(bidirectional), multidrop capability at higher data rates, of up to 1 Mbit/s. This enabled the 
federated architectures that evolved during the 1980s, when multiple data bus architectures 
were developed to cater for increased data flow and system segregation requirements. At this 
stage the aerospace electronic components were mainly bespoke, being dedicated solutions 
with few, if any, applications outside aerospace. 

The final advance occurred when electronic components and techniques were developed, 
driven mainly by the demands of the communications and IT industry which yielded a far 
higher capability than that which aerospace could sustain. This heralded the use of COTS 
technology, the use of which became more prevalent, and integrated modular avionics 
architectures began to follow and adapt to technology developed elsewhere. 

The key attributes of each of the evolutionary stages of architectural development are 
described below. 

2. 1. 1 Distributed Analogue Architecture 

The distributed analogue architecture is shown in Figure 2.2. In this type of system the major 
units are interconnected by hardwiring; no data buses are employed. This results in a huge 
amount of aircraft wiring and the system is extremely difficult to modify if change is 
necessary. This wiring is associated with power supplies, sensor excitation, sensor signal 
voltage and system discrete mode selection and status signals. These characteristics are 
evident in those aircraft conceived and designed throughout the 1950s and 1960s, many of 
which are still in service today. 

Figure 2.2 Distributed analogue architecture. 


These systems have dedicated subsystems, controls and displays. The displays are electro- 
mechanical and often extremely intricate in their operation, requiring instrument maker skills 
for assembly and repair. 

The use of analogue computing techniques does not provide the accuracy and stability 
offered by the later digital systems. Analogue systems are prone to bias or drift, and 
these characteristics are often more pronounced when the aircraft and equipment are subject 
to a hot or cold soak over a prolonged operating period. The only means of signalling rotary 
position in an analogue system is by means of synchro angular transmission systems. The 
older analogue aircraft - termed classic in the industry - therefore contain a huge quantity of 
synchros and other systems to transmit heading, attitude and other rotary parameters. Pallet 
(1992) is an excellent source of information on many of the older analogue techniques. 

The older equipment is very bulky and heavy and tends to be unreliable as there are many 
moving parts. This is not a criticism; the designers of the time did their best with the 
technology available, and many very elegant engineering solutions can be found in this type of 
equipment. Futhermore, the skills required to maintain some of the intricate instruments and 
sensors are gradually becoming scarcer, and consequently the cost of repair continues to rise 
even assuming spare parts are available. Many educational and training institutions no longer 
teach at this technology level, giving rise to a knowledge gap, which in turn has implications 
for organisations wishing to refurbish or maintain legacy aircraft and products. 

As has already been mentioned, these systems are very difficult to modify, which leads to 
significant problems when new equipment such as a flight management system has to be 
retrofitted to a classic aircraft. This is required when military aircraft are upgraded to comply 
with modern Air Traffic Control (ATC) procedures or a global air transport system (GATM) 
which are described in Chapter 7. 

Typical aircraft in this category are: Boeing 707, VC10, BAC 1-11, DC-9 and early Boeing 
737s. Many of these types are still flying, and some such as the VC-10 and the KC-135 and 
E-3/E-4/E-6 (Boeing 707 derivatives) are fulfilling military roles. They will continue to do so 
for several years, but gradually their numbers are dwindling as aircraft structural problems are 
manifested and the increasing cost of maintaining the older systems takes its toll. 

2. 1.2 Distributed Digital Architecture 

The maturity of digital computing devices suitable for airborne use led to the adoption of 
digital computers, allowing greater speed of computation, greater accuracy and removal of 
bias and drift problems. The digital computers as installed on these early systems were a far 
cry from today, being heavy, slow in computing terms, housing very limited memory and 
being difficult to reprogram - requiring removal from the aircraft in order that modifications 
could be embodied. 

A simplified version of the distributed digital architecture is shown in Figure 2.3. The key 
characteristics of this system are described below. 

Major functional units contained their own digital computer and memory. In the early days 
of military applications, memory was comprised of magnetic core elements which were very 
heavy and which in some cases could only be reprogrammed off-aircraft in a maintenance 
shop. This combined with the lack of experience in programming real-time computers with 
limited memory and the almost total lack of effective software development tools resulted in 
heavy maintenance penalties. 

As electrically reprogrammable memory became available, this was used in preference to 
magnetic memory. A significant development accompanying the emergence of digital 



Figure 2.3 Distributed digital architecture. 

processing was the adoption of serial half-duplex (unidirectional) digital data buses - 
ARINC 429 (civil aircraft) and Tornado serial (UK military) - which allowed important 
system data to be passed in digital form between the major processing centres on the aircraft. 
Although slow by today's standards (llOkbit/s for ARINC 429 and 64kbit/s for Tornado 
serial), the introduction of these data buses represented a major step forward, giving 
navigation and weapon-aiming systems major performance improvements by adopting this 

At this stage, individual system components and equipment were still dedicated in 
function, although clearly the ability to transfer data between the units had significantly 
improved. The adoption of data buses, particularly ARINC 429, spawned a series of ARINC 
standards which standardised the digital interfaces for different types of equipment. The 
uptake of this standardisation led manufacturers producing inertial navigation systems (INS) 
to prepare standard interfaces for these systems. This eventually led to the standardisation 
between systems of different manufacturers, potentially easing the prospect of system 
modification or upgrade. The ARINC data bus is still important in military systems since 
many civil platforms adopted for military use rely upon the bus for baseline avionics system 
integration. The Boeing 737 [maritime patrol and airborne early warning and command 
system (AWACS)], Boeing 767 (AWACS and tanker) and A330 (tanker) are all recent 
examples. The Nimrod MRA4 uses an Airbus-based architecture for the avionics and flight deck 
display system which was based on ARINC 429. This is a modern example of a new military 
aircraft successfully blending a commercial system with military standard systems. 

Displays in the cockpit were dedicated to their function as for the analogue architecture 
already described. The displays were still the intricate electromechanical devices used 
previously, with the accompanying problems. In later implementations the displays become 
multifunctional and multicolour, and the following display systems were developed in the 
civil field: 

• Electronic flight instrument system (EFIS); 

• Engine indication and crew alerting system (EICAS) 

Electronic checkout and maintenance (ECAM) 

Boeing and others; 



Data buses offered a great deal of flexibility in the way that signals were transferred from 
unit to unit. They also allowed architectures to be constructed with a considerable reduction 
in interunit wiring and multipin connectors. This led to a reduction in weight and cost, and 
also eased the task of introducing large and inflexible wiring harnesses into the airframe. 
This, in turn, led to reductions in the non-recurring cost of producing harness drawings, and 
the recurring cost of manufacturing and installing harnesses. 

Although data buses did remove a great deal of aircraft wiring, the question of adding an 
additional unit to the system at a later stage was still difficult. In ARINC 429 implementa- 
tions, data buses were replicated so that the failure of a single link between equipment did 
not render the system inoperable. 

Overall the adoption of even the early digital technology brought great advantages in 
system accuracy and performance, although the development and maintenance of these early 
digital systems was far from easy. 

Aircraft of this system vintage are: 

• Military - Jaguar, Nimrod MR2, Tornado and Sea Harrier; 

• Civil - Boeing 737 and 767 and Bombardier Global Express; these aircraft are relevant as 
many military platforms in the tanker, AWACs and intelligence gathering roles use these 
baseline civilian platforms. 

2. 1.3 Federated Digital Architecture 

The next development - the federated digital architecture - is shown in Figure 2.4. The 
federated architecture - from now on all architectures described are digital - relied 
principally upon the availability of the extremely widely used MIL-STD-1553B data bus. 
Originally conceived by the US Air Force Wright Patterson development laboratories, as 
they were called at the time, it evolved through two iterations from a basic standard, finally 
ending up with the 1553B standard, for which there are also UK Def-Stan equivalents (UK 
Def Stan 00-18 series). 

The adoption of the 1553B data bus standard offered significant advantages and some 
drawbacks. One advantage was that this was a standard that could be applied across all North 

1553B Mission Bus 

1553B Displays Bus | 


1553B Weapons Bus 

Figure 2.4 Federated avionics architecture. 


Atlantic Treaty Organisation (NATO) members, offering a data bus standard across a huge 
military market, and beyond. The 1553B data bus has been an exceptionally successful 
application, and the resulting vast electronic component market meant that prices of data bus 
interface devices could be reduced as the volume could be maintained. It also turned out, as 
had been the case with earlier data bus implementations, that the interface devices and hence 
the data buses were far more reliable that anyone could have reasonably expected. 
Consequently, the resulting system architectures were more robust and reliable than 
preceding architectures and more reliable than the designers had expected. 

Federated architectures generally use dedicated 1553B-interfaced line replaceable units 
(LRUs) and subsystems, but the wide availability of so much system data meant that 
significant advances could be made in the displays and other aircraft systems such as utilities 
or aircraft systems where avionics technology had not previously been applied. 

Although the higher data rates were most welcome - approximately 10 times that of the 
civil ARINC 429 and about 15 times that of the earlier Tornado serial data link - this 
standard was a victim of its own success in another way. The full-duplex (bidirectional), 
multidrop protocol meant that it was rapidly seized upon as being a huge advance in terms of 
digital data transfer (which it was). However, system designers soon began to realise that in a 
practical system perhaps only 10-12 of the 31 possible remote terminals (RTs) could 
generally be used owing to data bus loading considerations. At the time of the introduction of 
1553B, it was the policy of government procurement agencies to insist that, at system entry 
into service for a military system, only 50% of the available bandwidth could be utilised to 
allow growth for future system expansion. Similar capacity constraints applied to processor 
throughput and memory. Therefore, system designers were prevented from using the last 
ounce of system capability either in terms of data transfer or computing capability. This led 
to the use of subsystem dedicated data buses, as shown in Figure 2.4, in which each major 
subsystem such as avionics, general systems and mission systems had its own bus, complete 
with a dual-redundant bus controller. 

It was also recognised that it was not necessary to have every single piece of data bus 
equipment talking to every other across the aircraft. Indeed, there were sound systems 
reasons for partitioning subsystems traffic by data bus to enable all similar task-oriented 
systems to interchange information on the same bus. The provision of interbus bridges or 
links between different buses allowed data to be exchanged between functional subsystems. 
Therefore, during the late 1980s/early 1990s, many multibus architectures similar to the one 
portrayed in Figure 2.4 were evolved. With minor variations, this architecture is representa- 
tive of most military avionics systems flying today, including the F-16 mid-life update, 
SAAB Gripen and Boeing AH-64 C/D. 

The civil aircraft community was less eager to adapt to the federal approach, having 
collectively invested heavily in the single- source-multiple sink ARINC 429 standard that 
was already widely established and proving its worth in the civil fleets. Furthermore, this 
group did not like some of the detailed implementation/protocol issues associated with 
1553B and accordingly decided to derive a new civil standard that eventually became 
ARINC 629. To date, ARINC 629 is not envisaged as having a military application. 

MIL-STD-1553B utilises a 'command-response' protocol that requires a central control 
entity called a bus controller (BC), and the civil community voiced concerns regarding 
this centralised control philosophy. The civil-oriented ARINC 629 is a 2 Mbit/s system that 
uses a collision avoidance protocol providing each terminal with its own time slot during 


which it may transmit data on to the bus. This represents a distributed control approach. To 
be fair to both parties in the debate, they operate in differing environments. Military systems 
are subject to continuous modification as the Armed Forces need to respond to a continually 
evolving threat scenario requiring new or improved sensors or weapons. In general, the civil 
operating environment is more stable and requires far fewer system modifications. 

ARINC 629 has only been employed on the Boeing 777 aircraft where it is used in a 
federated architecture. The pace of aerospace and the gestation time required for technology 
developments to achieve maturity probably mean that the Boeing 777 will be the sole user of 
the ARINC 629 implementation. 

Along with the developing maturity of electronic memory ICs, in particular non- volatile 
memory, the federated architecture enabled software reprogramming in the various system 
LRUs and systems via the aircraft-level data buses. This is a significant improvement in 
maintainability terms upon the constraints that previously applied. For military systems it 
confers the ability to reprogram essential mission equipment on a mission-by -mission basis. For 
the civil market it also allows operational improvements/updates to be speedily incorporated. 

The more highly integrated federated system provides a huge data capture capability by 
virtue of extensive high-bandwidth fibre-optic networks. 

2. 1.4 Integrated Modular Architecture 

The commercial pressures of the aerospace industry have resulted in other solutions being 
examined for military aircraft avionics systems. The most impressive is the wholesale 
embracement of integrated modular architectures, as evidenced by US Air Force initiatives 
such as the Pave Pillar and JIAWG architectures which will be described later. 

The resulting architectures use open standards, ruggedized commercial technology to 
provide the data bus interconnections between the major aircraft systems and integrated 
computing resources (Figure 2.5). 

2.1.5 Open Architecture Issues 

As the use of digital technology became increasingly widespread across the aircraft 
throughout the 1970s, a number of issues became apparent. The use of digital technology 
introduced new and difficult issues that the system designer and end-user had to embrace in 
order successfully to develop and support digital avionics systems. 

The use of computers and, later, microprocessors, the functionality of which lay in the 
software applications that were downloaded, introduced a new and far-reaching discipline: 
that of software tools and software languages. Early processors were slow and cumbersome 
without the benefit of structured and standardised tools, instruction sets and languages. The 
memory available for the operational software was strictly limited and was a major 
constraint upon how the program was able to fulfil its task. A further compounding factor 
is that the avionics processor is required to perform its operational tasks in real time. At that 
point in the evolution of computer systems, such operational 'real-time' design experience as 
was available was centred on the use of mainframe applications where size and memory are 
less of an impediment. This stimulated a separate offshoot of the computing community that 
had new issues to address and precious few tools to help in its endeavour. 


grated Cabinets 
ular Architecture 
om Module Useage 





Data Bus 

High Speed 

















Another factor had an impact upon the adoption of digital technology. In the main the 
aerospace community is and was conservative and generally prefers to adopt bespoke 
solutions more suited to its unique operational tasks, particularly where high integrity and 
safety are often primary considerations. Furthermore, airborne applications, particularly those 
on military aircraft designed for worldwide use, have to survive severe operational environ- 
ments compared with a mainframe computer operating in a comfortable air-conditioned 
environment. As a consequence, the development timescales, particularly of large or multi- 
national projects requiring extensive software designs, became much longer than those of 
corresponding commercial projects. Therefore, while the military aerospace designers 
conservatively used technology that was proven and that would work within their severe 
environments, the commercial computer world was able to adopt more flexible approaches. 
As both worlds developed independently, the military tended to become a much smaller 
player in terms of component development and procurement while the commercial fields of 
IT and telecommunications became the major driving force. In particular, in this environment, 
the commercial computing world were early to adopt open architectures, whereas the military 
community did so much later. 

2. 1.6 Impact of Digital Systems 

Although early implementations of digital technology followed the distributed digital 
architectures shown in Figure 2.3, the availability of dedicated specifications such as MIL- 
STD-1553B permitted the aircraft wide interconnection of functional aircraft and avionic 
functions. While somewhat limited in capacity at 1 Mbit/s, the standard proved to be reliable 
and the multidrop, full-duplex architecture was central to the adoption of the federated 
digital architecture described above. Digital avionics technology was used not only to 
implement navigation and weapon-aiming functions but also became embedded in sensors, 
communications, flight and engine control, stores management systems, displays and utility 
control functions. 

As functionality and the use of digital technology increased, so the proliferation of 
processor types and software languages also increased. Therefore, while an aircraft system 
was integrated in the sense that the data buses were able readily to interchange data between 
major systems and subsystems, the individual lower-level equipment implementations were 
beginning to diverge. 

One factor that became a major issue was the non-availability of a standardised software 
language. In many cases the time constraints imposed by the real-time application meant that 
the program had to be written in assembler or even machine code. The ability to modify and 
support such implementations once the original design team had dispersed became 
impossible. Reliable and detailed software documentation was generally sparse or non- 
existent, with comment fields to explain the design philosophy. As the scope of digital 
technology expanded, the need for widespread and standardised software tools, languages 
and procedures became recognised. The military community addressed these issues by 
developing more military specifications to enforce standardisation and impose disciplines to 
improve the visibility, documentation and supportability of software application programs. 

Therefore, while the widespread application of digital technology in distributed digital and 
federated digital architectures brought considerable performance advantages, there was a 
significant downside that needed to be addressed. 


2.1.7 Response of the Services to Digital Avionics Systems Issues 

The problems resulting from widespread application of digital technology were addressed - 
at the risk of considerable oversimplification - by the adoption of the following measures: 

• Adoption of federated multibus architectures using a range of line replaceable units 
(LRUs), commonly known as 'black boxes' to the layman; 

• Standardisation of processors; 

• Standardisation of high-order language (HOL). 

These measures were accompanied with a number of standards imposing defined hardware 
and software development methodologies. 

The adoption of the federated bus architecture was generally a success inspite of the 
limited bandwidth of MIL-STD-1553B, and many so-called fourth-generation fighter aircraft 
are flying and enjoying the benefits of that approach. The embedded dual-redundant nature 
of the data bus standard coupled with the inherent reliability of the data bus interface devices 
has provided a robust and durable solution. The real problems existed at a lower level in the 
subsystems and LRUs that comprised the overall avionics system. 

The US military took the approach of standardising upon a common processor in the hope 
that major benefits would result in terms of lower costs and enhanced supportability. It was 
hoped that the use of a standard machine used force- wide would provide these benefits across 
a range of platforms, and the standard adopted by the US Air Force was the MIL-STD- 
1750A limited instruction set processor. As often happens as a result of ambitious 
standardisation initiatives, there is a tendency to 'round up' the performance of the common 
device to meet the more demanding (and more risky) applications; this tends to result in a 
'fit- all' product that is overly complex for many mundane applications. The standardisation 
focus within the US Navy was the AN/AYK-14 computer which suffered similarly. There- 
fore, while the Air Force and Navy were both pursuing standardisation initiatives, their 
chosen solutions differed. 

The US Air Force also attempted to standardise the software language, and during the 
mid-1980s the JOVIAL language was commonly specified. However, by the 1970s, the US 
Department of Defense (DoD) was using more than 2000 languages for its mission-critical 
programming. Most of these were languages that were developed for one specific applica- 
tion. Finally, in 1975, the DoD formed the US Department of Defense High-Order Language 
Working Group (HOLWG) to find a solution to what was often called the 'software crisis'. 

The HOLWG group members decided that they needed to create a language that could be 
used for just about anything, whether it be systems programming, artificial intelligence or, 
most important of all, real-time programming and embedded systems. Rather than create this 
new language themselves, they decided to hold a contest. Coincidentally, all of the 
competing teams created Pascal-based languages. In the end, the winner was the French- 
led CII Honeywell-Bull. Eventually, the language was christened 'Ada', in honour of Lady 
Ada Lovelace, daughter of the poet Lord Byron and assistant to mathematician Charles 
Babbage, who invented the analytical machine. Lady Ada is often considered to be the 
world's first programmer, and more details of her life can be found in Woolley (1999). 

These initiatives were essentially aimed at producing a military off-the-shelf (MOTS) 
suite of products whereby the military system designer could select known products for the 
implementation of the design. The idea was that these military products would be supported 


by the commercial industry computer base, suitably funded, to provide the military 
community with the technology and devices that they desired. The drawback to this 
approach was that the commercial community was moving ahead faster than the military 
could keep pace, not only with regard to devices and software but also in the adoption of 
open system architectures. The military were therefore forced to consider open architecture 
modular integrated cabinet approaches that the commercial world was developing with 

The adoption of modular architectures was not totally new to the services: the US Air 
Force in particular had initiatives under way throughout the 1970s and 1980s. The integrated 
communication navigation and identification architecture (ICNIA) was a serious attempt to 
use common radio-frequency (RF) building blocks to address the proliferation of RF 
equipment. A later initiative called Pave Pillar also addressed modular architecture issues. 
This programme questioned the black box approach to avionics. Pave Pillar architecture 
physically comprised a number of building blocks called common modules. Each module 
contained the circuitry to perform a complete digital processing function including interface 
control and health diagnosis. The common modules were developed from a limited very 
high-speed integrated circuit (VHSIC) chip set. A number of common module types could 
then be built up from a small family of VHSIC chips. Modules could then, in turn, be built up 
to form the basis of any of several avionic subsystems. Some uncommon modules are still 
required for the odd specific function; however, reduction in the types of spare required as a 
result of common module usage provided a significant cost improvement. Common modules 
result in increased production runs for specific modules, which reduces initial purchase cost, 
while the associated reduction in spares reduces maintenance and support costs. A modular 
concept allows maintenance engineers to remove and replace components while allowing 
system designers to adapt the avionics suite to new requirements, and at reduced risk. This 
concept eventually developed into the integrated modular avionics architecture discussed 
later in this chapter. 

By the late 1980s the Joint Integrated Avionics Working Group (JIAWG) had adopted an 
approach that was mandated by the US Congress to be used on the three major aircraft 
developments. These were the US Air Force advanced tactical fighter (ATF), now the F-22 
Raptor; the US Navy advanced tactical aircraft (ATA), or A- 12 Avenger, which was 
cancelled in 1990; and the US Army LHX helicopter which became the RAH-66 Comanche 
and which was cancelled in early 2004. The JIAWG architecture as implemented on the 
F-22 and later IMA/open architectures will be described later in this chapter. Mean- 
while in Europe, the Allied Standards Avionics Architecture Council (ASAAC) developed 
a set of hardware and software architecture standards around the IMA/open architecture 

2.1.8 Need to Embrace COTS 

As it became clear that the military aerospace community was a follower rather than a leader 
in terms of component technologies and architectures, a number of key issues had to be 
addressed if COTS were to embraced. These included: 

• The ability to obtain devices suitable for operation in specified military operating 
temperature ranges: — 55°C to + 125°C; other environmental issues such as vibration 
and humidity would also be important; 


• Achieving support for the military requirements of integrity /safety, security and certification: 
issues to which the donor technology is not necessarily exposed in the commercial field; 

• Determining the heritage of the hardware device or software program - integrity of 
development tools, documentation, design traceability and assurance; 

• Lack of control of the service authorities over the design standards used; 

• The relative short life span and volatility of commercial products, leading to huge lifetime 
support issues. 

The federated architecture already described has proved to be successful but not without a 
number of drawbacks: the relatively limited bandwidth of MIL-STD-1553B has already been 
mentioned. In such a system, units are loosely coupled, and this may lead to non-optimum 
use of data owing to latency issues and overall system performance degradation. While system 
upgrades may be relatively easily accomplished at the data bus level simply by adding another 
remote terminal, this is only part of the story. Any change in data transfer leads to changes in 
the bus controller transaction tables. More importantly, system- wide upgrades usually affect a 
number of different units, each of which may have its own issues relating to the ease (or 
difficulty) of modification. Furthermore, if the units involved in the upgrade are provided by 
different vendors, as is often the case, the modification or upgrade process becomes even more 
complicated as all vendors need to be managed throughout the programme. 

The fundamental advantage offered by an integrated modular avionics (IMA) approach is 
that, from the outset, the system is conceived using standard building blocks that may be 
used throughout the aircraft level system. Therefore, common processor modules, common 
memory modules and, where possible, common input/output modules offer the means of 
rapidly conceiving and constructing quite extensive system architectures. This approach 
reduces risk during the development phase, as well as offering significant supportability 
advantages. The IMA philosophy readily adapts to redundancy implementation in a most 
cost-effective manner so that economies of scale are easily achieved. 

The adoption of COTS-based IMA architectures provides another significant advantage, 
that is, rapid prototyping. As the baseline modules are off-the-shelf produced in a 
commercial format, these may be readily purchased in order that a candidate architecture 
may be built and prototyped using mature commercial boards. Previously, prototyping had to 
wait until early development hardware was available for all the contributing subsystems; this 
hardware tended to be immature, and often featured development bugs. 

A true comparison of the benefits to be attained by utilising COTS can only be reached 
after reviewing all the areas where the commercial technology offers benefits. In virtually 
every respect, COTS offers advantages: 

1. Data buses and networks. Evolution from the initial 1 Mbit/s MIL-STD-1553B data bus to 
high-speed COTS 1 Gbit/s, plus solutions such as the scalable coherent interface (SCI), the 
asynchronous transfer mode (ATM), fibre channel technology and the gigabit Ethernet (GE). 

2. Software. The adoption of a multilayer software model that segregates the application 
software from the functional hardware, thereby allowing software to be portable across a 
range of differing and evolving hardware implementations. 

3. Processing: 

• Signal processing. This offers significant advances in some of the more challenging 
tasks in the area of radar, electronic warfare (EW) and electroopics (EO); many 
technological advances have resulted from the development of the Internet. 


• Data processing. There is often a significant task involved in processing data for 
transmission or display. The pace at which commercial office IT processors are moving 
in terms of clock speed far outweighs any advance that could have been achieved using 
military drivers alone. 

These issues and the competing technologies are discussed in detail by Wilcock et al. (2001). 

In summary, there are many advantages that the military community may gain by the 
adoption of COTS technology. Apart from the obvious advantage of cost, much greater 
capability and functionality benefits result, and COTS solutions have been vigorously sought 
since around the mid-1990s. While not totally without drawbacks, COTS has been adopted 
for many of the new build or upgrade programmes launched in the last decade. 

A few of the many examples of COTS technology in military systems are as follows: 

1. The US JSTARS aircraft is fitted with standard DEC Alpha operator workstations. 

2. The Israeli Python 4 air-to-air missile avionics system is powered by an Intel 486 

3. The Chinese F-7MG has been fitted with a Garmin GPS 150 receiver (as available at most 
good high-street electronics dealers!). 

2.2. Aerospace-specific Data Buses 

As has already been described, digital data buses have been one of the main enablers in the 
use of digital electronics in aircraft avionics systems. Early data buses were single-source, 
single-sink (point-to-point), half-duplex (unidirectional) buses with relatively low data rates 
of the order of a few tens to lOOkbit/s data rates. The next generation of data buses as 
typified by MIL-STD-1553B were multiple- source, multiple-sink (bidirectional) buses with 
data rates of around 1-10 Mbit/s. Later fibre channel buses achieve 1 Gbit/s data rates, with 
the prospect of expanding to several Gbit/s. 

Figure 2.6 depicts a comparative illustration of the data bus transmission rates of buses 
used onboard military avionics platforms. 

The dedicated aerospace data buses used within the military aerospace community are: 

• Tornado serial data bus; 

• ARINC429; 

• MIL-STD-1553B and derivatives; 

• STANAG3910. 

Other bus standards such as the JIAWG high-speed data bus (HSDB), the IEEE 1394b and 
fibre channel buses are all commercial standards that have been adopted for military use. 

2.2.1 Tornado Serial 

The Tornado serial data bus - more correctly referred to by its PAN standard (Panavia 
standard) - was the first to be used on a UK fighter aircraft. The bus was adopted for the 
Tornado avionics system and also used on the Sea Harrier integrated head-up display/ 


















Fiber Channel 
FC-AE 1 - 2G bit/sec 

IEEE 1394b 
800M bit/sec 

80M bit/sec 

20M bit/sec 

1 Mbit/sec 

ARINC 429 


Tornado Serial 



F-35: F/A-18E/F 


(Block 60) 




Very widely used 

in Military 


Civil Bus 

Tornado & Sea 

Figure 2.6 Comparative data bus transmission rates. 

weapon-aiming computer (HUD/WAC) system. This is a half-duplex serial bus operating at a 
rate of 64 kbit/s and is used to pass data between the avionics main computer and other 
sensors, computers and displays within the Tornado nav attack and weapon-aiming system, 
as shown in Figure 2.7. 

The bus comprised four wires implemented as a twisted screened quad format. The lines 
carried clock and complement and data and complement respectively. 

The Tornado system architecture utilising this bus is depicted in the lower part of 
Figure 2.7. It shows the main computer interfacing via the Tornado serial bus with major 
avionics subsystems: 

• Doppler radar; 

• Radar (ground-mapping radar (GMR) and terrain- following radar (TFR); 

• Laser range finder/marked target receiver (MTR); 

• Attitude Sources - inertial navigation system (INS) and secondary attitude and heading 
reference system (SAHRS); 

• Autopilot/flight director system (AFDS); 

• Stores management system (SMS); 

• Front cockpit: 




- Data — 

- Data — 

- Cloc k- 

- Clock- 


Tornado Core 
Avionics Architecture! 
Showing Serial Data 

Figure 2.7 Tornado serial data bus and application. 

- head-up display (HUD), 

- pilot's navigation display, 

- interface unit 1 (IFU 1) for front cockpit I/O; 

• Rear cockpit: 

- navigator's displays (2), 

- interface unit 2 (IFU 2) for rear cockpit I/O. 

This system was designed in the early 1970s and entered service in 1980. 

2.2.2 ARINC429 

ARINC 429 is a single-source, multiple- sink, half-duplex bus that operates at two transmission 
rates; most commonly the higher rate of 100 kbit/s is used. Although the data bus has its origins 
in the civil marketplace, it is also used extensively on civil platforms that have been adopted 
for military use, such as the Boeing 737, Boeing 767 and A330. High-performance business 
jets such as the Bombardier Global Express and Gulfstream GV that are frequently modified as 
electronic intelligence (ELINT) or reconnaissance platforms also employ A429. 

The characteristics of ARINC 429 were agreed among the airlines in 1977/78, and it was 
first used throughout the B757/B767 and Airbus A300 and A3 10 aircraft. ARINC, short for 
Aeronautical Radio Inc., is a corporation in the United States whose stockholders comprise 
US and foreign airlines and aircraft manufacturers. As such it is a powerful organisation 




Node 2 

Node 2 


I I 



Node 2 

Node 3 

L_ J 

Node 4 

Figure 2.8 A429 topology and the effect of adding units. 

central to the specification of equipment standards for known and perceived technical 

The ARINC 429 (A429) bus operates in a single-source-multiple sink mode so that a 
source may transmit to a number of different terminals or sinks, each of which may receive 
the data message. However, if any of the sink equipment needs to reply, then each piece of 
equipment will require its own transmitter and a separate physical bus to do so, and cannot 
reply down the same wire pair. This half-duplex mode of operation has certain disadvan- 
tages. If it is desired to add additional equipment as shown in Figure 2.8, a new set of buses 
may be required - up to a maximum of eight new buses in this example if each new link 
needs to operate in bidirectional mode. 

The physical implementation of the A429 data bus is a screened, twisted wire pair with the 
screen earthed at both ends and at all intermediate breaks. The transmitting element shown 
on the left in Figure 2.9 is embedded in the source equipment and may interface with up to 


Source LRU 

1 n 1 



i ^ 

1 n 

i ^ 

1 "-I 1 



To other 
(up to 20 


1 Bit Period 

Figure 2.9 A429 data bus and encoding format. 



Figure 2.10 A429 data word format. 

20 receiving terminals in the sink equipment. Information may be transmitted at a low rate of 
12-14 kbit/s or a higher rate of 100 kbit/s; the higher rate is by far the most commonly used. 
The modulation technique is bipolar return to zero (RTZ), as shown in the box in the figure. 
The RTZ modulation technique has three signal levels: high, null and low. A logic state 1 is 
represented by a high state returning to zero; a logic state is represented by a low state 
returning to null. Information is transmitted down the bus as 32 bit words, as shown in 
Figure 2.10. 

The standard embraces many fixed labels and formats, so that a particular type of 
equipment always transmits data in a particular way. This standardisation has the advantage 
that all manufacturers of particular equipment know what data to expect. Where necessary, 
additions to the standard may also be implemented. Further reading for A429 may be found 
in Moir and Seabridge (2004). 

2.2.3 MIL-STD-1553B 

MIL-STD-1553B has evolved since the original publication of MIL-STD-1553 in 1973. The 
standard has developed through 1553A standard issued in 1975 to the present 1553B 
standard issued in September 1978. The basic layout of a MIL-STD-1553B data bus is 
shown in Figure 2.11. The data bus comprises a screened twisted wire pair along which 
data combined with clock information are passed. The standard generally supports multiple- 
redundant operation with dual-redundant operation being by far the most common 
configuration actually used. This allows physical separation of the data buses within the 
aircraft, therby permitting a degree of battle damage resistance. 

Control of the bus is performed by a bus controller (BC) which communicates with a 
number of remote terminals (RTs) (up to a maximum of 31) via the data bus. RTs only 
perform the data bus related functions and interface with the host (user) equipment they 
support. In early systems the RT comprised one or more circuit cards, whereas nowadays it is 



Bus Bus 
A B 

44 44 




^- Up to 30 ■► 


■ ----^ Data 
!+••► Data 




•*- Up to 30 •► 


Figure 2.11 MIL-STD-1553B data bus. 

usually an embedded chip or hybrid module within the host equipment. Data are transmitted 
at 1 MHz using a self-clocked Manchester biphase digital format. The transmission of data in 
true and complement form down a twisted screened pair offers an error detection capability. 
Words may be formatted as data words, command words or status words, as shown in 
Figure 2.12. Data words encompass a 16 bit digital word, while the command and status 

20 bits = 20 microseconds 

Data Word 


Status Word 

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20, 

I I I I I I I I l__l I I I I I I I l_ * 

SYNC 10 110 


RT Address T/D Sub-address 

l/K Mode 

Data word count 
/mode code 

Terminal address Message 





Instrumention Reserved 
Figure 2.12 MIL-STD-1553B data bus word formats. 




Busy Parity 



Remote Terminal A to Bus Controller Transfer 


Terminal A 


Terminal A 

Terminal B 


~70 microseconds - 

Status Word 


Data Word 

Remote Terminal A to Remote Terminal B Transfer 



-120 microseconds - 

Status Word Data Word 

Status Word 


Figure 2.13 MIL-STD-1553B typical data transactions. 

words are associated with the data bus transmission protocol. Command and status words are 
compartmented to include various address, subaddress and control functions, as shown in 
Figure 2.12. 

MIL-STD-1553B is a command-response system in which transmissions are conducted 
under the control of a single bus controller at any one time; although only one bus controller 
is shown in these examples, a practical system will employ two bus controllers to provide 
control redundancy. 

Two typical transactions are shown in Figure 2.13. In a simple transfer of data from RT A 
to the BC, the BC sends a transmit command to RT A, which replies after a short interval 
known as the response time with a status word, followed immediately by one or more data 
words up to a maximum of 32 data words. In the example shown in the upper part of the 
figure, transfer of one data word from RT A to the BC will take approximately 70 |is 
(depending upon the exact value of the response time plus propagation time down the bus 
cable). For the direct transfer of data between two RTs as shown from RT A to RT B, the BC 
sends a receive command to RT B followed by a transmit command to RT A. RT A will send 
its status word plus the data (up to a maximum of 32 words) to RT B which then responds by 
sending its status word to the BC, thereby concluding the transaction. In the simple RT to RT 
transaction shown in Figure 2.13, the total elapsed time is around 120 |is for the transmission 
of a single data word, which appears to be rather expensive on account of the overhead of 
having to transmit two command words and two status words as well. However, if the 
maximum number of data words had been transmitted (32), the same overhead of two 
command and two status words would represent a much lower percentage of the overall 
message time. For further reading, see MIL-STD-1553B (1986). 



MIL-STD-1553B has proved to be a very reliable and robust data bus and is very well 
established as a legacy system. Attempts have been made to increase the data rate which is 
the only major shortcoming. A modification of 1553 called 1553 enhanced bit rate (EBR) 
running at 10 Mbit/s has been adopted for bomb carriage on the JSF/F-35 using the miniature 
munitions/store interface (MM/SI). Other vendors have run laboratory demonstrators at 
100 Mbit/s and above, and a feasibility program has been initiated to demonstrate 1553 bit 
rates of 100 Mbit/s with the aim of extending data rates to 500 Mbit/s. This possible 
derivative is termed enhanced 1553 (EB-1553), and the US Air Force recently hosted a 
workshop to investigate the possibilities. In October 2003 the Society of Automotive 
Engineers (SAE) formed an 'Enhanced Performance 1553' task group to address the 
prospect of launching applications with throughputs of 200-500 Mbit/s using existing cables 
and couplers. A further standard using 1553 is MIL-STD-1760 - a standard weapons 
interface which is described in Chapter 9. 

2.2A STANAG3910 

The evolution of STANAG 3910 was motivated by a desire to increase the data rate above the 
1 Mbit/s rate provided by MIL-STD-1553B. The basic architecture is shown in Figure 2.14. 
The high-speed fibre-optic data terminals pass data at 20 Mbit/s and are connected using a 
star coupler. Control is exercised by MIL-STD-1553B using electrical connections. The 
encoding method is Manchester biphase, as for 1553, and data transactions are controlled by 
means of a bus controller, as is also the case for 1553. 

The use of fibre optics passing data at 20 Mbit/s offers a significant improvement over 
1553. Furthermore, the ability to transfer messages of up to 132 blocks of 32 words (a total 
of 4096 data words) is a huge advance over the 32 word blocks permissible in 1553. A total 
of 31 nodes (terminals) may be addressed, which is the same as 1553. 

There are four possible implementations of STANAG 3910 (Table 2.1). The standard also 
makes provision for the high-speed channel to be implemented as an optical transmissive star 
coupler, a reflective star coupler or a linear Tee coupled optical bus. Eurofighter Typhoon 
utilises the type A network with an optical reflective star coupler in a federated architecture. 

MIL-STD-1553 Bus (Electrical) 

20 Mbit/sec Channel (Fibre-Optic) 
Figure 2.14 STANAG 3910 architecture. 



Table 2.1 STANG 3910 implementations 

Type A Low-speed channel 

High-speed channel 

Type B High-speed channel 

Low-speed channel 

Type C High-speed channel 

Low-speed channel 

Type D High-speed channel 

Low-speed channel 

1553B data bus 
Fibre-optic data bus 
Fibre-optic equivalent of 1553B 
Physically separate fibre-optic data bus 
Fibre-optic equivalent of 1553B 
Wavelength division multiplexed with 

low- speed channel 
1553B data bus 
Physically separate wire data bus 

Note: Both the low-speed and high-speed buses use Manchester biphase encoding. 

The STANAG bus is used for the avionics bus and the attack bus while standard MIL-STD- 
1553B is used for the weapons bus. The Typhoon architecture is discussed in Chapter 9 of 
this book. 

The similarity of transactions to 1553 may be seen by referring to the remote terminal 
(RT) to bus controller (BC) transaction on the low-speed bus as shown in Figure 2.15. The 
BC issues a standard 1553 command word followed by a high-speed (HS) action word 
(taking the place of a 1553 data word). After a suitable interval the RT issues a 1553 status 
word that completes the transaction. After an intermessage gap, the next transaction is 
initiated. This cues a high-speed message frame on the high-speed bus which enables the 
transmission of up to 132 blocks of 32 data words, up to a maximum of 4096 words, as has 
already been stated. A full description of all the STANAG data transactions may be found in 
Wooley (1999) which is also a useful source on many other data buses that may be used in 
avionics applications. 

3910 Remote Terminal A to Bus Controller Transfer 
(initiated after one RT to BC message) 


Terminal A 


- 70 ji seconds - 

HS Action 

(1553 Data 



Electrical Bus 

20 Mbps Fibre 
Optic Bus 

HS Message Frame 

[Maximum Transfer of 4096 16 bit words] 

Figure 2.15 STANAG 3910 RT to BC data transaction. 



- 20 bits = 20 microseconds - - 

HS Action 

Status Word 

■ 12 3 4 

1 ' I 



7 8 9 10 11 12 


14 15 16 17 

i i i i 


19 20, 
i 1 


i i i i 

• A 

i or 
1 B 


HS Message Identify/ 
HS Mode 

HS Word Count/ 
HS Mode 

i p i 


i i 


; Reserved 

— i 

HS Transmitter 

HS Receiver 


TF; P ; 


I I 

Figure 2.16 STANAG 3910 HS word formats (low-speed bus). 

I I 

The formats of the low- speed bus HS action word and HS status word are shown in 
Figure 2.16. It can be seen that the general format is similar to the 1553 words, except that 
these protocol words have positive-going synchronisation pulses as they are replacing the 
standard 1553 data word [which also has a positive-going synchronisation pulse 
(Figure 2.12)]. Also, the word content, instead of containing a 16 bit data word, contains 
message fields that relate to the high-speed bus message content or transmitter/receiver 
status. As for other 1553 words, the final bit (bit 20) is reserved for parity. 

The format of the STANAG 3910 high-speed bus data structure is shown in Figure 2.17. 
The data content is preceded by 56 bits associated with the message protocol and followed 

| 4| 4| 8 | 8 | 








Protocol Data Units (PDU) 














CS - 

Control Sequence 

PR - 


SD - 

Start Delimiter 

FC - 

Frame Control 

PA - 

Physical Address 

DA - 

Destination Address 


Word Count 


- Frame Check Sequence 

ED - 

End Delimiter 


- Data — ►! 

3 + WC 

4 + WC 

Protocol Data Units (PDU) 










Figure 2.17 STANAG 3910 data structure. 



by a further 20 bits, giving a total overhead of 76 bits in all. Therefore, a maximum 4096 data 
word transfer will take (56 + 4096 x 16 + 20) = 65 612 bits. At a the high-speed bus data 
rate of 20 Mbit/s, the total elapsed time is 3280.6 lis. 

The application of STANAG 3910 is confined to two European fighter aircraft programmes. 
Eurofighter Typhoon uses a variant called the EFA bus, which is a type A implementation, 
while the French Rafale employs type D, using the electrical high-speed bus version. 

2.3 JIAWG Architecture 

The Joint Integrated Avionics Working Group (JIAWG) was a body charged with developing 
an avionics architecture based upon the Pave Pillar principles of a modular integrated 
architecture. The resulting advanced avionics architecture (A 3 ) was mandated by the US 
Congress to apply to the following major projects of the late 1980s: 

US Air Force Advanced tactical fighter/F-22 Raptor 

US Navy Advanced tactical aircraft (ATA)/A-12 

(cancelled in the early 1990s) 
US Army LH helicopter/RAH-66 Comanche 

(cancelled in early 2004) 

2.3. 1 Generic JIA WG Architecture 

A generic avionics architecture embracing the main features of the JIAWG architecture is 
illustrated in Figure 2.18. The major functional elements are: 








— * I 

i r-h EW/ESM 

V__— '| 







Signal Data 

Processors Processors 



Parallel Buses 

Serial Buses 







■ <High Speed Data Bus- - - - ^* 






p-|— Electrical 






Figure 2.18 Generic JIAWG avionics architecture. 





-PI Bus #1 ■ 


-PI Bus #2 - 
TM Bus- 

Figure 2.19 Generic JIAWG integrated cabinet architecture. 

• Radio-frequency (RF) apertures and sensor front ends associated with electrooptic (EO), 
missile warning, radar, electronic warfare/electronic support measures (EW/ESM), and 
communications, navigation and identification (CNI) systems; 

• A fibre-optic switched network handling incoming preprocessed sensor data; 

• Integrated avionic racks encompassing signal and data processing and interconnected 
using switched networks, parallel buses and serial buses; 

• A fibre-optic switched network handling video data destined for displays; 

• Aircraft and weapons systems; 

• A high-speed data bus (HSDB) (fibre-optic bus) interconnecting the avionics major 
systems to the integrated cabinets. 

The integrated cabinet architecture selected for the JIAWG architecture is portrayed in 
generic form in Figure 2.19. The main processor is called the common integrated processor 
(CIP) and two cabinets are provided in the F-22 architecture with space provision for a third. 
In fact the term processor is a misnomer as the CIP function contains a cluster of processors 
(up to seven different types) working together to perform the aircraft-level signal processing 
and data processing tasks. It has been reported that the combined throughput of the CIP is of 
the order of 700 million operations per second (MOPS). The F-22 CIP is the responsibility of 

The F-22 cabinet has space to accommodate up to 66 SEM-E size modules, but in fact 
only 47 and 44 modules are used respectively in CIP 1 and CIP 2; the remaining 19/22 
modules are growth. The density of the packaging and the high power density mean that 
liquid cooling is used throughout for the cabinets. 

To achieve the necessary communications to interchange data with the donor and 
recipient subsystems, the cabinet interfaces internally and externally using the following 
data buses: 

• A high-speed data bus (HSDB) - a fibre-optic (FO) bus connecting the cabinet with other 
aircraft subsystems, as shown in Figure 2.18; 

• A dual-redundant parallel interface bus (PI bus) (backplane bus) used to interconnect the 
modules within the rack; 

• A serial test and maintenance bus (TM bus) for test and diagnosis. 

The FO buses and interconnects are provided by the Harris Corporation. 






BIU (3) 


BIU (18) 

Protocol ..... .^ 

BIU (2) 

BIU (6) 






BIU (30) 


Figure 2.20 FO linear token passing topology (HSDB). 

2.3.2 High-speed Data Bus 

The high-speed data bus is a linear token ring FO bus operating at 80 Mbit/s. The protocol is 
in accordance with SAE AS4074.1, for a linear token passing multiplexed data bus. The key 
characteristics of the HSDB are: 

• Manchester biphase encoding; 

• Message length up to 4096 16 bit words; 

• Ability to service up to 128 terminals. 

Figure 2.20 shows a FO liner token passing topology typical of that used for the HSDB. 

The SAE linear token passing bus (LTPB) employs a broadcast technique whereby all the 
terminals can receive all the traffic passing on the bus. However, the transactions are 
structured such that only those stations whose address matches the destination address within 
the message header can copy the message. A token passing path is superimposed upon the 
linear transmission media, and it is this that provides the control logic for each bus interface 
unit (BIU). 

A token is passed around the ring from the lowest address to the highest and then to the 
lowest again, forming a continuous loop. Stations that have data to transmit claim the token 
and send the message. The philosophy used is that known as a 'brother's keeper' where every 
terminal is responsible for providing a valid token to pass to the next terminal. Therefore, in 
the example given in Figure 2.20, the token and therefore control of the bus would be passed 
as follows: 

BIU(2) > BIU(3) > BIU(6) > BIU(8) > BIU(18) > BIU(30) > BIU(2) 

and the sequence would repeat. 


The media access control (MAC) protocol adopted for the SAE TTPB is one in which 
timers are used to bound the time a station is allowed to pass data on the bus. As well as 
having bounded time constraints, each BIU has four message priority levels such that the 
highest-priority messages are transmitted first and those of lowest priority last. In the event 
that lower-priority messages cannot be passed owing to insufficient time, these messages can 
be deferred until the next period. 

AIR4288, AIR4271, AS4290 and RTCA DO-178B in the references section of this chapter 
provide further information upon the operation, use and validation of SAE AS4074. 

Unfortunately, this data bus standard, although adopted by the SAE, has not attracted the 
general support and take-up that the system designers had hoped for. Consequently, it will 
need to be replaced with a more recent alternative such as the fibre channel (FC) bus very 
early in the service life of the aircraft. Before its recent cancellation, the RAH-66 Comanche 
had already adopted a fibre channel implementation in place of the HSDB. 

2.3.3 PI Bus 

The PI bus is a fault- tolerant parallel backplane bus very similar in structure to VME. The 
bus operates using a 32 bit wide bus which can be expanded to 64 bit wide. The transfer rate 
is 50 Mbit/s and the JIAWG architecture employs a dual-redundant implementation. The PI 
bus is supported by the SAE 4710 and STANAG 3997 standards. 

2.3.4 TMBus 

The TM bus is a serial bus comprising five wires that supports test and diagnostics according 
to the IEEE Std 1149.5-1995 which allows a master controller to interface with up to 250 
slave units. Depending upon the precise implementation of the TM bus, diagnosis may be 
achieved at the board (module) or chip (component) level. 

2.3.5 Obsolescence Issues 

While the JIAWG architecture was a valiant attempt to achieve cross-service avionics 
standardisation, it has not succeeded owing to a number of factors, none of which could have 
been foreseen. At the outset, each of the services was expecting to procure several hundred 
ATFs, ATAs and LH helicopters. The Cold War had not ended and the USSR/Warsaw Pact 
was seen as the prime threat. The Navy ATA/A-12 was cancelled early for other reasons, but 
the remaining ATF/F-22 and LH/RAH-66 programmes had to contend with a post-Cold War 
and post-Desert Storm world. Inevitably, some of the momentum was lost once the USSR 
disintegrated and contempory weapons systems performed so well in the Iraq conflict. 
Finally, it would have been a visionary who could have forecast the true scale and pace of the 
impact of commercial technology upon military platforms. 

Before cancellation of the RAH-66 Comanche, the following problems had been 
recognised with the JIAWG architecture: 

1. The JIAWG-driven processor (Intel 80960 MX) product line was closed down and the 
processor changed to the i86. 

2. The JIAWG-driven backplane (PI bus) product line was closed down. 





, I 


1 " ' 1 

fCC : | I 




; 4-| EW/ESM 





CNI v 




Signal Data 

PWCi :,',:;.r, Pi .*.-.-■.■:;: r ■■ 



Parallel Buses 

Serial Buses 

l — i '■" ;■ ■? 1 

(1 Gbil/sec or more) 







[ VMS 








Figure 2.21 JIAWG architecture upgraded using fibre channel buses. 

3. The JIAWG-driven test and maintenance (TM bus) is non-standard. 

4. The JIAWG/SAE-driven 80Mbit/s HSDB failed to achieve commercial or military 

5. The JIAWG-driven SEM-E module format is too small for some COTS technology. 

These problems will probably be addressed on the F-22 by the adoption of a fibre channel 
(FC) for the aircraft-level and major subsystem interconnects, permitting the existing 
original architecture to be 'grown' in bandwidth while addressing the obsolescence issues. 
Figure 2.21 shows how the baseline JIAWG generic architecture could be modified to accept 
FC buses in place of the original HSDB, PI and TM buses. The overall impression is of 
simplification, as the original 80 Mbit/s linear token ring HSDB is replaced by FC buses 
capable of carrying 1 Gbit/s or more. The higher bandwidth provided is more suited to 'data 
stream' the sensor data to the signal and data processing areas within the integrated cabinets. 
Similar architectural upgrades have been examined for the Apache AH-64 C/D (which 
predates the JIAWG architecture) and RAH-66/Comanche helicopters. 

Developments on more recent projects have provided a forward path for the integration of 
portable legacy software and new processor types without the need to invest further 
development work on the operational software packages. This approach is outlined later 
in the chapter. 

2.4 COTS Data Buses 

The lack of progress in sponsoring a sustainable high-speed military data bus has led systems 
designers towards developments in the commercial field, namely COTS. There are a number 


of possibilities, including the asynchronous transfer mode (ATM) and the fibre distributed 
data interface (FDDI), but those viewed with the most favour at the time of writing are: 

• The fibre channel bus; 

• IEEE 1394 firewire. 

2.4.1 Fibre Channel 

The fibre channel (FC) is a high-throughput, low-latency, packet-switched or connection- 
oriented network technology; in aerospace applications the latter configuration is usually 
employed. Data rates are presently available up to 1 Gbit/s, although 10 Gbit/s versions are 
under development. A whole series of standards are evolving, with one dedicated to avionics 
applications - the avionics environment project (FC-AE). 

The FC bus is designed to operate in an open environment that can accommodate multiple 
commercial protocols such as the small computer system interface (SCSI) and transport 
control protocol/Internet protocol (TCP/IP). For avionics applications the FC-AE standard 
specifies a number of upper-level protocols (ULPs) that align more directly with aerospace 
applications. These are: 

• FC-AE-1553; 

FC-AE- ASM (anonymous subscriber messaging); 
FC-AE-RDMA (remote direct memory access); 
FC-AE-LP (lightweight protocol). 

FC-1553 is very useful as it allows the MIL-STD-1553 protocol to be mapped on to the high- 
bandwidth FC network, creating a low-overhead highly deterministic protocol with a high 
bandwidth capability. FC-AE-1553 allows a large number of nodes to communicate: 
increasing from 32 for the baseline 1553 implementation to 2 24 , while the number of 
possible subaddresses increases from 32 to 2 32 . Likewise, the maximum word count 
increases from 32 to 2 32 . While these features offer an enormous increase in capability, a 
further benefit is that FC-AE-1553 provides a bridge between legacy 1553 networks and the 
much higher-bandwidth FC networks. Therefore, an upgrade introducing an FC network to 
provide additional bandwidth in certain parts of an avionics architecture can be readily 
achieved while maintaining those parts that do not require a bandwidth increase intact. The 
FC-AE- ASM, RDMA and LP options are lightweight protocols that can be variously adopted 
for specific avionics applications, depending upon the exact requirement. 

SCSI is a widely used commercial mass storage standard that allows FC nodes easily to access 
SCSI-controlled disc space, while TCP/IP is a widely used networking protocol that allows 
modern computer peripherals and software to communicate. This allows the avionics system 
designer to utilise commercially available packages. The compatibility of SCSI and TCP/IP with 
the FC-AE specifications also allows rapid prototyping early in the development cycle. 

The FC standards define a range of topologies, the basic four of which are shown in 
Figure 2.22. These are as follows: 

1. Point-to-point communication is used to provide a dedicated link between two computers. 
This is the least expensive option but one that may find application connecting a radar or 
EO sensor with its associated processor. 






|NL| ► [NL| 

Arbitrated Loop 


Hubbed Loop 
Figure 2.22 Fibre channel topologies. 



2. Arbitrated loop is a ring topology providing a low-cost solution to connect up to 128 
nodes on a shared bandwidth network, where each node is able to gain control of the loop 
and establish a connection with another node and generate traffic. Arbitrated loops are 
limited in their ability to support simultaneous operation and do not include fault 
tolerance to enable the network to withstand node or media failure. The available 
bandwidth is shared between the network nodes. 

3. Hubbed loop is a variation of the arbitrated loop in which the node connections are made by 
means of a central hub. In the event of failure, a port bypass circuit enables the failed or 
inoperative node to be bypassed and allows traffic to continue between the remaining nodes. 

4. Switched fabric is a network capable of providing multiple simultaneous full-bandwidth 
transfers by utilising a 'fabric' of one or more interconnected switches: this provides the 
highest level of performance. An offshoot of the switched fabric topology is the provision 
of arbitrated loops (called a public loop) enabling connection between low-bandwidth 
nodes (connected to the public loop) and high-bandwidth nodes connected directly to 
the fabric. Switched fabric represents the most powerful but also the most expensive 

To enable these topologies, a number of different node types are defined and used, as shown 
in Figure 2.22. These node types are as follows: 

1. Node port (N_Port). This provides a means of transporting data to and from another port. 
It is used in point-to-point and switched topologies. 

2. Fabric port (F_Port). This provides the access for an N_Port to the switched fabric. 

3. Loop port (L_Port). This port is similar to an N_Port but has additional functionality to 
provide the arbitrated loop and hubbed loop functions. It can also be used to connect loops 
and the switched fabric. 

4. Expansion port (E_Port). The E_Port is used to provide interconnection between multiple 
switches within the fabric. 















Bus Bus 

Figure 2.23 Redundant switched fabric topology with bridges. 

The switched fabric topology offers a high-capacity interconnect between a number of 
nodes which can in turn can link to their own shared or dedicated networks (Figure 2.23). 
Figure 2.23 depicts a six-port configuration that is dual redundant, the lower 'ghosted' 
network being a replica of the upper network. Although functionally this appears as a 
network in the figure, in practical terms it will probably be packaged as a module which can 
be located in an integrated modular cabinet with the other buses or systems being hosted 
within the same cabinet. 

2.4.2 Fibre Channel Options 

There are many ways in which the FC may be implemented and products are being offered 
by a range of vendors. At the time of writing, four main implementations appear to be 
attracting the most interest: 

• StarFabric; 

• Rapid 10; 

• Infiniband; 

• 100/1000 Gbit/s Ethernet. 

Of these, StarFabric has been selected for applications on the joint strike fighter (F-35). 
Figure 2.24 shows an eight-port StarFabric switched fabric incorporating Quad PC processing. 

2.4.3 IEEE 1394 Firewire 

IEEE 1394 firewire is a widely used data bus scaleable in its original form from 50 to 
400 Mbit/s. It has an extremely wide market capture, being commonly used in the electronic 
domestic consumer market: video cameras, etc. This marketplace has also paved the way for 
IEEE 1394 to be widely applied in civil aircraft in-flight entertainment (IFE) systems. 


Figure 2.24 Eight-port StarFabric switched fabric module. 

A significant disadvantage of the baseline standard is that it utilises a daisy-chain 
architecture to connect the network devices together. Like the arbitrated loop configuration 
already described, it is therefore unable to tolerate failures by the nodes or the transmission 
media and is not an attractive option for avionics applications. 

Later versions developed under the IEEE 1394b standard are able to work in a network 
form up to 800Mbit/s, and there are reports that this standard has been adopted for 
interconnecting portions of the vehicle management system (VMS) on the F-35. 

2.5 Real-time Operating Systems 

As for a home computer or laptop, the operating system (OS) of an avionics system and its 
major subsystems is key to correct operation and achievement of performance goals. The OS 
needs to be reliable and robust, not prone to malfunctions or crashes. On an aircraft there are 
many functions that may be flight critical or mission critical, and therefore the software must 
be reliable and safe. Furthermore, the aircraft systems perform their functions in real time, 
and therefore the operational programs depend upon a real-time operating system (RTOS) to 
provide the means for them to execute their programs. 

For many years individual companies developed their own operating systems which were 
used in their own products, often associated directly with a particular processor or computer 
application. Such dedicated or proprietary systems are still being developed today, but 
increasingly designers are seeking commercial solutions that are created, maintained and 
supported by technology specialists. This releases high-grade software resources to work on 
more specific application software that executes the aircraft or weapons systems functions. 
An RTOS will be smaller, more modular and more focused in application than the 
commercial OS used on laptops and PCs. In particular, for high-integrity real-time 
applications the RTOS should be kept as compact as possible, featuring only essential 
functions. It also needs to provide a guaranteed level of service, always responding within 
specified time constraints. 

2.5. 1 Key Attributes 

The key attributes expected of a COTS RTOS are: 

1. Versatility. The RTOS must be applicable to multiple systems such that a minimum 
number of RTOSs are required across the aircraft. 


2. Safety. The system must be capable of being partitioned such that software integrity and 
therefore aircraft safety are maintained. 

3. Security. The RTOS must be capable of separating multiple levels of classified and 
sensitive data. 

4. Supportability. The OS must be maintained and enhanced throughout its operational 

2.5.2 Safety 

The software developed for high-integrity or safety-critical systems used in aerospace 
applications is subject to intense scrutiny. All processes associated with developing the 
software are clearly identified, and extensive plans and documentation are produced to 
ensure that the correct validation and verification processes are undertaken. Within the civil 
aerospace community, the specification is issued by the Radio Technical Commission for 
Aerospace (RTCA). This body has issued DO-178B, entitled 'Software considerations in 
airborne systems and equipment certification', which was developed by the avionics industry 
to establish software considerations for developers, installers and users when aircraft 
equipment design is implemented using microcomputer techniques. DO-178B is recognised 
'as an acceptable means to secure FAA approval of digital computer software', RTCA 
DO-178B. In Europe an equivalent specification is issued by the European Organisation for 
Civil Aviation Equipment (EUROCAE) under EUROCAE ED-12B. A joint RTCA/EURO- 
CAE committee is expected to commence work on a third revision standard, DO-178C. 

DO-178B acknowledges that not all computer failures or software malfunctions affect an 
aircraft to the same degree. A malfunction in the aircraft flight control system is clearly more 
hazardous than the failure of a reading light. Accordingly, software is divided into five 
different categories, level A through to E, where level A represents the highest level of 
approval and level E the lowest. 

During the initial aircraft design, all those failures that can cause the various levels of 
failure severity are identified and used to modify the aircraft systems design accordingly. 
Therefore, long before an aircraft is built, all these conditions are identified and 
appropriate design steps taken and quality of design assured. This process helps to define 
the system architecture, the number of control and power channels, the level of redun- 
dancy, etc. It also specifies a design assurance level according to what the effects of a 
failure might be; these design assurance levels are reflected in the RTOS software 
certification levels (Table 2.2). 

There are five main categories of failure severity. The most serious is a catastrophic failure 
which would result in the loss of the aircraft and passengers. The probability of such an event 
occurring is specified as extremely improbable, and in analytical or qualitative terms it is 
directed that a catastrophic failure should occur less than 1 x 10 -9 per flight hour. That is 
less than once per 1000 million flying hours. Other less significant failures are 'hazardous', 
'major', 'minor' and 'no-effect'; in each case the level of risk is reduced and the probability 
of the event occurring is correspondingly increased. Therefore, a minor failure - perhaps the 
failure of a reading light - can be expected to be reasonably probable, with the event 
occurring less than 1 x 10 -3 per flight hour or less than once every 1000 flying hours. A 
brief summary of the applicable failure severities is shown in Table 2.3. 

DO-178B is not mandated for military aircraft use, indeed its use is not mandated within 
the civil community. However, as it is a recognised method of successfully achieving 


Table 2.2 RTOS software certification levels 


certification level Definition 

A Software whose anomalous behaviour would cause or contribute to 

a failure of a system function resulting in a catastrophic failure 

condition for the aircraft 
B Software whose anomalous behaviour would cause or contribute to a 

failure of a system function resulting in a hazardous failure condition 

for the aircraft 
C Software whose anomalous behaviour would cause or contribute to a 

failure of a system function resulting in a major failure condition 

for the aircraft 
D Software whose anomalous behaviour would cause or contribute to a 

failure of a system function resulting in a minor failure condition 

for the aircraft 
E Software whose anomalous behaviour would cause or contribute to a 

failure of a system function resulting in a no-effect failure 

condition for the aircraft 

certification, the industry effectively uses it as a de facto requirement. There are several 
reasons why the military avionics community should adopt the standard: 

1. It should do so where dual-use software exists, for example flight management system 
(FMS) software developed for a civil application and adopted for a military program. 

2. The adoption of 'commercial best practices' is encouraged, and the fact that DO-178B is 
the accepted best practice in the civil community offers some reassurance. 

3. There is a need to adopt DO-178B in areas where military aircraft have to operate in 
situations governed by the civil regulations. FMS software designed to satisfy global air 
transport management (GATM) requirements will also be subject to communications, 
navigation, surveillance/air transport management (CNS/ATM) regulations imposed by 
the civil authorities. 

4. The standard should be adopted to provide robust software products capable of future reuse. 

2.5.3 Software Partitioning 

In recent years there has been a tendency for software to be co-hosted on shared processor 
assets, and this has been given added impetus by integrated modular avionics architectures. 

Table 2.3 Summary of the applicable failure severities 

Failure severity Probability Analytical 

Catastrophic Extremely improbable Less than 1 x 10~ 9 per flight hour 

Hazardous Extremely remote Less than 1 x 10~ 7 per flight hour 

Major Remote Less than 1 x 10~ 5 per flight hour 

Minor Reasonably improbable Less than 1 x 10~ 3 per flight hour 




,.''"' Weapons 

















Figure 2.25 Software and hardware integration. 

In federated architectures of the type already described earlier in this chapter there is little 
opportunity to rationalise processor usage in this way. 

However, systems integrators have over the past 10 years embraced the modular concept 
and have confronted the new software issues that result. Multitasking of processors requires 
software partitioning such that the software applications cannot interfere with each other. In 
particular, there is a pressing need to ensure that high-integrity applications cannot be 
adversely affected by lower integrity functions co-hosted on the same processor. 

ARINC 653 is an industry specification - again originating from the civil community - 
that defines partitioning of software and addresses these issues. Software partitioning is an 
important issue in its own right, but this methodology has helped to address other important 
issues that have proved very difficult to overcome in the past, including the obsolescence of 
hardware. This has always been a problem but has been greatly accentuated by the adoption 
of COTS technology with its rapid hardware development cycles and consequent early 

The philosophy adopted is typified by the tiered approach illustrated in Figure 2.25. These 
tiers or layers provide the following: 

1. At the top level, an ARINC 653 compliant software infrastructure or application program- 
ming interface (API) provides the partitioning of weapons systems functions that may 
have been developed in a previous program (legacy software) or may be new software 
developed specifically for the platform. These will generally, but not exclusively, be 
military- specific applications (the dual use of software has already been discussed). 

2. The API infrastructure interfaces with the RTOS which will generally be a commercial 
package. In many cases the RTOS will be DO-178B compliant. 

3. A board-level (i.e. sub-LRU-level) support package will provide the necessary software 
support to enable the COTS hardware to interface with the commercial RTOS and API layers. 

4. The hardware layer based upon COTS technology will be the most dynamic and rapidly 
varying component of this implementation as the rapid evolution of processor and FO/FC 
network technology continues. The fact that the most rapidly varying hardware content is 
decoupled from the application software means that hardware obsolescence can be 
contained and technology advances enjoyed while the investment in software applications 
and system functionality is protected. 


Recent applications of partitioned and certifiable RTOS have included: 

• Green Hills software with level A, INTEGRITY- 1 7 8B RTOS on the Sikorsky S-92 
helicopter avionics management system (AMS); 

• Lynux Works Lynx-OS level operating system in association with Rockwell Collins on 
the adaptive flight display system on the Bombardier Challenger 300 business jet; 

• Wind River systems with AE653 on the Boeing 767 tanker transport and C-130 avionics 
modification program (AMP); 

• CsLEOS RTOS developed by BAE SYSTEMS and certified to DO-178B level A for a fly- 
by-wire flight control system upgrade to the Sikorsky S-92 helicopter. 

2.5.4 Software Languages 

An added complication to the portability of software application packages relates to the 
software language used. At an early stage in the use of digital technology it was recognised that 
the software burden in terms of initial programming effort and lifetime support was in many 
ways more difficult to quantify and manage than the development of the aerospace-specific 
hardware. By the mid-1970s the US services were in crisis owing to the vast proliferation in 
software languages: reportedly in 1976 there were more than 450, some as dialects of standard 
languages but generally all of which were low in interoperability and reliability and high in 
maintenance costs. This led to the development of Ada as a high-order language. 

By the late 1970s/early 1980s the US services and others specified Ada as the preferred 
language, and Ada 83 was later upgraded to Ada 95 with additional features and capability. 
As the number of compilers available, available expertise and development tools increased, 
Ada became the language of preference in the late 1980s and early 1990s, representing a 
large investment for the Pentagon and their supplier base. 

During the 1990s the armed services withdrew their support for Ada, and new languages 
such as C and C++ were permitted. There have been and are on-going technical and 
business debates of great intensity about the merits and demerits of Ada versus these 
commercial languages. It is not the place of this publication to pass judgement on either 
approach. However, from the avionics systems designer's viewpoint, the issue is how to 
integrate legacy software packages that might largely be written in Ada. In extreme cases, 
software may need to be recoded in order to be compatible with the new environment. The 
benefits accruing from the adoption of COTS in terms of increased performance and 
longevity therefore have to weighed against these issues. 

2.5.5 Security 

As well as the need to partition software functions for reasons of integrity, a more recent 
requirement is the need to partition for reasons of security. The US authorities now want the 
RTOS in some applications to be able to operate at multiple levels while maintaining security 
between them. The evolution of network-centric warfare doctrines means that the same pro- 
cessor that is handling flight management functions in the civil air traffic domain may also have 
to handle highly classified target data, weapons engagement profiles, rules of engagement, etc. 
The rules for protecting sensitive data have evolved over the past three decades but 
originally resides in a publication issued in the United States in 1983 and known as the 
'orange book'. More recently, an international IT security project called 'Common Criteria' 




! Low/ 
-►I Medium 

I Integrity 





High | 
Security | 














Figure 2.26 Integrity and security partitioning. 

has replaced the original publication. 'Common Criteria' specifies seven evaluation assur- 
ance levels (EALs) that correspond to earlier classifications. The highest level, EAL-7, 
requires the system to be multilevel secure, and is able to separate three or more levels of 
data while processing them on shared hardware. 

Usually, the necessary assurance levels are achieved using a combination of hardware 
[perhaps using the memory management unit (MMU) of the processor] and software means 
by using the core of the RTOS or 'microkernel'. This concept, referred to as multiple 
independent levels of security (MILS), permits different levels of classified data to be 
accommodated. The MILS approach is under active consideration for a number of programs 
including C-130 AMP, F-22, F-35, the global positioning system (GPS) and the joint tactical 
radio system (JTRS). MILS can co-exist with ARINC 653/DO-178B implementations, as 
shown in Figure 2.26. 

2.6 RF Integration 

Earlier in this chapter the integration of aircraft at the top level using FC networks was 
examined. Another area where the aircraft can benefit from integration is in the area of the 
radio-frequency (RF) subsystems. Modern fighters are fitted with a plethora of RF systems, 
some of which are listed below: 

• Radar; 

• Electronic warfare (EW); 

• Identification friend or foe (IFF); 

• Radar warning in several RF bands; 

• Navigation aids: TACAN, ILS, MLS, GPS; 

• Communications: VHF, UHF, HF SatCom joint tactical information distribution system 
(JTIDS), Link 11; secure radio. 

These systems each have their own antenna, RF sections, signal and data processing: the result 
is a vast collection of non-standard, heavy and sometimes unreliable hardware modules. 



It has been recognised for some time that this is an area ripe for rationalisation and 
functional integration using a common suite of modules. In the past, initiatives such as 
integrated communications, navigation and identification architecture (ICNIA) and the 
integrated electronic warfare suite (INEWS) have attempted to address this problem. For 
whatever reason, these failed to attain the benefits sought, but advances in RF processing 
technology are now offering the prospect of fuller integration. 

2.6. 1 Primary Radar Evolution 

2.6. 1. 1 Independent Systems of the 1950s 

Integration of the radar, communications and navigation functions has been occurring over 
the past four decades, albeit at a slow pace. Figure 2.27 shows a 1950s era system with 
stand-alone radar, communications and navigation functions. By the standards of today the 
radar was quite rudimentary, with airborne search and tracking modes. Ground-mapping 
capabilities were incidental and a pulse Doppler facility with a look-down capability was not 
yet available. This system was analogue in nature and quite similar to the distributed 
analogue system described earlier. Systems were interconnected using a great deal of 
dedicated wiring; and computer and displays would be dedicated to the relevant system. 
Typical aircraft in this category would be the North American F-86D, Lockheed F-94 Starfire 
and Northrop F-89 Scorpion. 






Figure 2.27 First-generation analogue system - 1950s. 



Figure 2.28 Integrated system - 1960s and 1970s. 

2.6. 1.2 Integrated Systems of the 1960s and 1970s 

The next stage is typical of systems entering service during the 1960s, of which the 
McDonnell Douglas F/A-4 Phantom is a good example (Figure 2.28). In this system the 
radar, communications and navigation systems are integrated into a mission system with a 
display providing mission rather than subsystem data. Although these systems were largely 
analogue, later variants did introduce some digital subsystems. Most system interconnection 
was still very much undertaken by hardwiring as data buses were only just becoming mature 
at this stage. These systems introduced further complexity brought about by additional- 
functionality pulse Doppler radars and inertial navigation systems (INS) and by systems 
integration. The addition of other, new, more capable navigation aids and integrated radar 
warning radar (RWR) suites added a further twist of complexity. This integration had to be 
achieved without the advantage of mature and reliable data buses to enable a federated 
system to be constructed. These aircraft were very challenging to maintain, suffering from 
unreliable equipment, high power consumption, vast amounts of wiring and LRUs that were 
often buried deep in the aircraft as space was very much at a premium. The wiring-intensive 
nature of the system meant that modifications were always difficult and very expensive to 
embody. Nevertheless, in spite of the maintenance penalties, these systems brought a step 
change in aircraft functional capability. 



During the 1980s the availability of mature and cost-effective data buses such as 1553 
eased the integration task and removed much of the interconnecting wiring, leading to the 
multibus federated system seen in the F-16 Fighting Falcon and F/A-18 Hornet in the United 
States and in the Eurofighter Typhoon, SAAB Gripen and Dassault Rafale in Europe. The 
AH-64 Apache was one of the first helicopters to use a multiple 1553 bus network to 
integrate its weapons system. At this stage the need for standardisation and modularisation of 
hardware and software were recognised as the initial adoption of digital technology brought 
with it many teething problems as well as performance improvements. 

2.6. 1.3 Integrated Modular Architecture of the 1990s 

The 1990s federated architecture built upon the lessons learned from the previous generation, 
and modular implementations were sought (Figure 2.29). The JIAWG architecture adopted a 
modular approach but ran into component obsolescence problems, as has already been 
described. However, in this generation of avionics systems, the performance explosion came 
in the form of additional RF and electrooptic (EO) sensors. By now, radar antennas had 
evolved from parabolic dishes into flat plates and used limited 'beam- shaping' techniques, 
but the antenna still needed to be mechanically scanned. Digital signal processing had 


Common Integrated 




Figure 2.29 Integrated avionics architecture. 



evolved to offer true multimode functionality, i.e. the ability to use the same radar for 
airborne intercepts, ground mapping and missile guidance, for example. Later radars such as 
those fitted in the F-15E, A/F-18E/F upgrade and block 60 F-16 upgrade (F-16E/F) included 
an active electronically scanned array (AES A) in place of the flat plate antenna. This antenna 
is fixed, and radar beams are shaped and steered entirely electronically, without the need 
for any moving parts. This also brought increased range performance, and highly reliable 
multimode radar capable of operating in several modes simultaneously. Common integrated 
processor cabinets provided a modular computing resource for all the mission functions. 

However, as will be seen, the RF content of the avionics system was increasing and the 
need for true integration of the RF system, sharing receiver, signal processor and transmitter 
resources, became more pressing. Consequently, technology studies preceding the joint 
strike fighter (JSF) were conducted to consider the route map for future development. The 
joint advanced strike technology (JAST) program embraced a number of study and 
technology demonstrator programs to identify technologies suitable for the JSF and reduce 
risk where possible. 

The JAST architecture highlighted the adoption of a number of technologies to aid systems 
integration, but perhaps the most innovative was the concept of using shared apertures and 
antenna, as shown in Figure 2.30, as well as a modular approach to the rest of the RF 
architecture. Much of the material gained from the JAST studies, combined with experience 
gained from the JIAWG and the F-22 Raptor, will be embodied in the F-35 joint strike fighter 
program. For more information, see the Joint Advanced Strike Technology Program 
Avionics Architecture Definition of 8 August and 9 August 1994. 

2.6.2 JIAWG RF Subsystem Integration 

The RF integration on the F-22 using the JIAWG architecture is indicative of the 
conservative state of the art 10 years ago when the F-22 program entered the engineering 


Common Analogue 

Common Digital 


Figure 2.30 2000 + integrated architecture with shared apertures. 


manufacturing and development (EMD) phase. The aim was to have a highly capable radar, 
EW and CNI suite that would enable the aircraft to survive and prosecute its mission in a 
high-threat environment, and to do so using stealth techniques. The resulting architecture and 
equipment suite, driven particularly by the need to carry all sensors and weapons internally 
to preserve stealth, led to a more complex system than other contemporary 'fourth- 
generation' fighters which, although being very capable in their own right, could not satisfy 
all the requirements of the F-22 mission. These aircraft do not have as comprehensive an 
avionics system, often needing to add additional 'podded' sensors for specific roles and carry 
weapons externally. 

The major RF subsystems on the F-22 are regarded as separate entities containing 
dedicated RF processing, although signal processing and mission data processing is 
integrated within the CIPs. These subsystems are: 

• Active electronically scanned array (AESA) radar; 

• Electronic warfare (EW) suite; 

• Communications, navigation and identification (CNI) equipment. 

Figure 2.31 depicts the F-22/JIAWG architecture, with these sensor subsystems and 
centralised processing functions highlighted. 

Table 2.4 has been compiled with the help of the Joint Advanced Strike Technology 
Program Avionics Architecture Definition of 8 August 1994 and 9 August 1994. It is 
representative of the JIAWG/F-22 implementation and serves the purpose of demonstrating 
the RF architecture integration advances over the past decade or so. 










Common Integrated 
Processor #1 

Common Integrated 
Processor #2 

-High Speed Data Bus 






& & A & & 

Figure 2.31 F-22/JIAWG top-level avionics architecture. 


Table 2.4 Summary of JIAWG apertures 
















Radar, active and 
passive targeting 






FWD Port and 
and STBD 

Radar warning 
receiver (RWR) 




























Situational awareness 
(SA); Fwd 
sector - two arrays, 
each uses 1 RWR 
element: Az and El DF 







Microwave landing 
system (MLS) 







SA; fwd sector - two 
arrays, each uses RWR 
element: Az and El DF 







IFF interrogater 







IFF transponder 





















ILS glideslope 
ILS localiser 







ILS marker 














Special communications 







Common high-band 
data link (CHBDL) 








Cooperative engagement 
capability (CEC) 







HF Comm; Link 11 

Note: CNI omnidirectional antenna in italics are not shown in Figure 2.31. 

Table 2.4 identifies all the apertures - 64 in all - and gives a brief summary of the 
subsystem, type of antenna, number of active elements, frequency band (GHz) and 
location, and a brief summary of the associated function. A diagrammatic portrayal 
of these apertures - ignoring the omnidirectional antenna for clarity - is shown in 
Figure 2.32. 

O E 
Q. O 

0) <D 


2> « -E .E 

re -o — > 

.c re z ^ 













Q. d) 


1 B 

a> o> 



w Nois 




5 £ jo 


-" i- i- 


> < cc £ 


0) Zh5 

* JXh 







The capabilities that this assembly of apertures offer, together with the appropriate 
electronics units, are summarised below: 

1. The AESA radar provides active and passive targeting. This includes ~2000 active 
transmit/receive (TR) elements. The AESA radar is addressed in Chapter 4. 

2. The electronic warfare (EW) suite uses arrays to provide: 

• Radar warning and situational awareness (SA) around the aircraft using a combination 
of spiral antennas over the 0.5-18 GHz frequency bands. A total of 24 spiral antennas 
are located around the aircraft to provide full spherical coverage for any RF source and 
provide direction- finding capabilities for the detected signals. 

• Jamming directed at the aircraft is sensed by a further two spiral antennas. The aircraft 
can, if desired, respond by transmitting electronic countermeasures in the 2-12 GHz 
bands using a total of 12 log periodic (LP) antennas located on the water-line. 

The principles of EW are described in Chapter 6. 

3. The CNI functions are addressed by a total of 21 slot, linear array, LP, ferrite and phased 
array antennas that provide the aircraft CNI capability: 

• microwave landing system (MLS) - 2; 

• IFF interrogator - 2; 


• GPS - 1 array with 4 elements; 

• ILS glideslope, localiser and marker - 2; 

• UHF SatCom - 1; 

• Common high-band data link (CHBDL) - 3; 

• Cooperative engagement capability (data link) - 3; 

• HF communications and Link 11-2. 

The principles of CNI equipment are described in Chapter 7. 

4. The electronics units associated with these apertures are: 

• Rf integrated racks to provide the RF amplification, detection and signal demodulation 
for incoming signals and the modulation and power amplification for outgoing signals. 
In this architecture the RF 'front end' is dedicated to each subsystem. 

• Low-noise amplifiers (LNA) located throughout the aircraft to amplify signals before 
transmission to the RF racks. 

• Travelling wave tube (TWT) amplifiers to provide power for EW countermeasure 
transmissions and IFF transmitters for the IFF interrogator. 

Figure 2.33 does not portray the full picture. In order to secure the best unimpeded fields 
of view, many of the apertures need to be located on or near the aircraft extremities. To 
illustrate this fact, Figure 2.33 shows the provision of the CNI apertures (upper aspect) for 
the F-22. This gives a true impression of the complexity of installing a highly capable system 
while abiding by the constraining factors that are imposed by the need for the aircraft to 
remain stealthy commensurate with its mission. 

The situation with regard to the EW apertures is equally if not more complex. Figure 2.34 
indicates a further aspect of distributing apertures in this manner. Many apertures need 










■i ! 


I L I 




dedicated power supplies provided locally and the provision of LNAs for many functions. 
The RF signals that are gathered by the arrays have to feed through the aircraft to the 
integrated RF racks using coaxial wiring to avoid undue signal attenuation. The fact that 
the EW suite also embodies wide-band transmitters to provide radiated electronic counter- 
measures is a further complicating factor. 

2.7 Pave Pace/F-35 Shared Aperture Architecture 

One of the objectives in developing the Pave Pillar architectures was to address the RF 
functional area and seek rationalisation of the receiver and demodulation and modulation and 
amplification/transmitting functions. In the JIAWG architecture above these are handled on a 
subsystem basis, and the aim of Pave Pace is to provide an integrated RF sensor system - 
sometimes called an integrated sensor system (ISS). The sharing of resources between the 
functional system can enable significant savings in cost, weight, volume and reliability. 
Studies have quantified these savings by comparing a third-generation ISS (JIAWG) with a 
fourth-generation version (Pave Pace) as shown in Table 2.5. 

The basis of a fourth-generation RF ISS is shown in Figure 2.35. The primary arrays may 
typically comprise a large active array: multiarm spiral arrays (MAS As), slot arrays and 
multiturn loops (MTLs). These arrays are connected via an RF interconnect to a collection of 
receive frequency converters that convert the signal to intermediate frequency (IF). The IF 
receive signals are fed through an IF interconnect to the receiver modules. After detection, 
the baseband in-phase (I) and quadrature (Q) components are fed through the fibre-optic 
interconnect to the integrated core processing. 

For transmission the reverse occurs, signals are passed to the multifunction modulators 
and through a separate IF interconnect to the transmit frequency converters. After modula- 
tion and power amplification the output signals are passed via the RF interconnect to the 
appropriate array (s). It is the sharing of these functions within a common RF host that 
enables the major savings to be made. In the comparison given above it is estimated that 
about a third was achieved by the use of more advanced components and improved 
packaging while the remainder came from the integration process. 

The JAST documentation already mentioned produced the Pave Pace equivalent of the 
JIAWG RF architecture shown in Figure 2.32 and expanded in Table 2.4 above. 

Figure 2.36 shows the rationalisation of apertures that can be gained from a fourth- 
generation integrated system, totalling 22 apertures as opposed to 64 apertures for its 
predecessor. Table 2.6 lists and groups these apertures by function. 

Table 2.5 Comparison of third- and fourth-generation RF integrated sensor 

Fourth-generation RF ISS Third-generation RF ISS 

7.9 million 




' 1989 US dollars. 

Source: Reference 2 in JAST report. 

Cost* ($) 

3.9 million 

Weight (lb) 


Volume (ft 3 ) 


Reliability (h) 






REAR Starboard 




Figure 2.36 Pave Pace - shared aperture RF architecture. 

Table 2.6 Pave Pace aperture listing 


Shared resources in Black 
CNI in Blue 
EW In Green 

Number System 

Type Elements (GHz) 



1 Radar, EW, WBSA 


3000 6-12 FWD A/A and A/G radar, RWR, 

ECM, SA, passive targeting, 
CHBDL, weapon data link 








RWR, ECM, SA,passive 
targeting, CHBDL 









RWR, ECM, SA, data 
link, MLS 







RWR, ECM receive 






2 TOP 

UHF radio, GPS, have 
quick, glideslope, JTIDS, 
TACAN, IFF transponder, 
TCAS (functions spread 
among four apertures to 
match coverage and 
functional mix) 







IFF interrogator 


Radar, EW 



















VOR, localiser, marker beacon, 
self-protect, SINCGARS, 
VHF radio 







HF Comm, Link 11 

Note: CNI HF Comms/Link 11 antenna are not shown in Figure 2.36. 


It can be seen that there is considerable rationalisation compared with the previous 
architecture, particularly in terms of rationalising functions: Radar + EW, EW + CNI, etc. 
The AESA has been replaced by a wide-band synthetic array (WBSA) (number 1 in Table 2.6) 
containing ~3000 elements that services radar, EW and CNI functions. Five similar but 
smaller arrays provide forward (numbers 4 and 5), rear (numbers 2 and 3) and aft (number 6) 
coverage for EW and CNI usage. A variety of spiral, MAS A and MTL antennas provide the 
entire gamut of EW and CNI equipment coverage as described in Table 2.6. 


ARINC Specification 429: Mk 33 digital information transfer system, Aeronautical Radio, Inc., 1977. 

ARINC 653, Avionics application software standard interface, Aeronautical Radio Inc., January 1997. 

AIR427 1 - Handbook of system data communication. 

AIR4288 - Linear token passing multiple data bus user's handbook. 

AS4290 - Validation test plan for AS4074. 

Guide to digital interface standards for military avionic applications, Avionic Systems Standardisation 

Committee, ASSC/1 10/6/2, Issue 2, September 2003. 
Joint Advanced Strike Technology Program, avionics architecture definition - issues/decisions/rationale 

document, Version 1, 8 August 1994. 
Joint Advanced Strike Technology Program, avionics architecture definition - appendices, Version 1, 9 

August 1994. 
MIL-STD-1553B digital time division command/response multiplex data bus, Notice 2, 8 September 

Moir, I. and Seabridge, A.G. (2004) Design and Development of Aircraft Systems - An Introduction, 

Professional Engineering Publishing, ISBN 1-86058-437-3. 
Pallet, E.H.J. (1992) Aircraft Instruments and Integrated System, Longman, ISBN 0-582-08627-2. 
RTCA DO-178B, Software considerations in airborne systems and equipment certification. 
UKDef Stan 18-00 Series. 
Wilcock, G., Totten, T. and Wilson, R. (2001) The application of COTS technology in future modular 

avionics systems. Electronics and Communications Engineering Journal, August 2001. 
Wooley, B. (1999). The Bride of Science - Romance, Reason and Ada Lovelace, Macmillan. 

3 Basic Radar Systems 

3.1 Basic Principles of Radar 

The original concept of radar was demonstrated by laboratory experiments carried out 
by Heinrich Hertz in the 1880s. The term RADAR stands for Radio Aid to Detection 
And Ranging. Hertz demonstrated that radio waves had the same properties as light (apart 
from the difference in frequency). He also showed that the radio waves could be reflected 
from a metal object and could be refracted by a dielectric prism mimicking the behaviour of 

The concept of radar was known and was being investigated in the 1930s by a number 
of nations, and the British introduced a ground-based early warning system called Chain 
Home. In the late 1930s, as part of the world's first integrated air defence, this system 
has been credited with the winning of the Battle of Britain in 1940. The invention of the 
magnetron in 1940 gave the ability to produce power at higher frequencies and allowed 
radar to be adopted for airborne use. The first application was to airborne interception (AI) 
radars fitted to fighter aircraft to improve the air defence of Great Britain when used in 
conjunction with the Chain Home system. By the end of WWII, rudimentary ground- 
mapping radars had also been introduced under the dubious name of H 2 S. Echoes of War 
(Lovell, 1992) gives a fascinating account of the development of radar during the war. Since 
that time, radar has evolved to become the primary sensor on military aircraft and is widely 
used in civil aviation as a weather radar able to warn the flight crew of impending heavy 
precipitation or turbulence. For further information, see Pilot's Handbook - Honeywell 
Radar RDR-4B. 

Since that time, enormous advances have been made in airborne radars. Fighter aircraft 
carry multimode radars with advanced pulse Doppler (PD), track- while- scan (TWS) and 
synthetic aperture (SA) modes that impart an awesome capability. Larger aircraft with an 
airborne early warning (AEW), such as the E-3, carry large surveillance radars aloft with 
aerial dishes in excess of 20 ft in diameter. At the other end of the scale, attack helicopters 
such as the AH-64 C/D Longbow Apache deploy near-millimetric radars in a 'doughnut' on 
top of the rotor, measuring no more than 3 ft across (Figure 3.1). 

Military Avionics Systems Ian Moir and Allan G. Seabridge 
© 2006 John Wiley & Sons, Ltd. ISBN: 0-470-01632-9 



Figure 3.1 Contrasting airborne radar applications. (I. Moir) 

The performance and application of radar is highly dependent upon the frequency of 
operation. Figure 3.2 shows the range of electromagnetic (em) applications used in modern 
military avionics systems. The applications may be grouped into three categories in 
ascending order of frequency: 

1. Communications and Navaids, more correctly referred to as Communications, Navigation 
and Identification (CNI), operating in the band from 100 kHz to just over 1 GHz. CNI 
systems are addressed in Chapter 7. 

2. Airborne radar from ~400MHz to a little under 100 GHz. This is the subject of this 
chapter and of Chapter 4. 

3. Electrooptics (EO) including visible light in the band from a little over 10 000 GHz 
(lOTHz) extending to just over 1 000 000 GHz (1000 THz). The frequency numbers are 
so high at this end of the spectrum that wavelength tends to be used instead. The EO band 
encompasses visible light, infrared (IR) and laser systems which are described in Chapter 5. 

Focusing on the airborne radar systems that are the subject of this and the next chapter, these 
cover the frequency range from ~400MHz to 94 GHz, as shown in Figure 3.3. This 
illustrates some of the major areas of the spectrum as used by airborne platforms. In 
ascending order of frequency, typical applications are: 

• E-2C Hawkey e US Navy surveillance radar operating at ~400 MHz; 

• US Air Force E-3 airborne warning aircraft command system (AWACS) employing a 
surveillance radar operating at ~3 GHz; 

• Radar altimeters operating at ~4 GHz, commonly used on civil and military aircraft; 

• Fighter aircraft operating in the 10-18 GHz range; 

• US Army AH-64 C/D Apache attack helicopter with Longbow radar (AH-64 D variant) 
operating at ~35 GHz; 

• Active radar-guided, air-launched or ground-launched antiarmour missiles: either Hellfire 
or Brimstone operating at ~94 GHz. 

The entire frequency range used by radar and other radio applications is categorised by the 
letter identification scheme shown in Table 3.1. However, only those frequencies assigned by 




















— Lasers 
Visible Light 


- Infra-Red 

-Hellfire Missile 

-Longbow Apache Radar 

Most Airborne 


-Radar Altimeter 




-HF Comms 


Figure 3.2 Range of electromagnetic applications in military avionics 
NATO Classification 








Hellfi re/Brimstone 

(~ 94GHz) 

Longbow Apache 

(~ 35 GHz) 



Airborne WX Radar 

(~ 9 GHz) 

Radar Altimeter 

(~ 4 GHz) 

— Surveillance 

(~ 3GHz) 

E2-C Radar 

(~ 400 MHz) 

Figure 3.3 Airborne radar frequency coverage. 



Table 3.1 Designation of radar bands [source: Skolnik, M.L (1980) 
Introduction to Radar Systems, McGraw-Hill] 

Band designator a 

Nominal frequency range 

ITU assignment 


3-30 MHz 


30-300 MHz 

138-144 MHz 
216-225 MHz 


300-1000 MHz 

420-450 MHz 
850-942 MHz 


1-2 GHz 

1215-1400 MHz 


2-4 GHz 

2300-2500 MHz 
2700-3700 MHz 


4-8 GHz 

5250-5925 MHz 


8-12 GHz 

8500-10 680 MHz 


12-18 GHz 

13.4-14.0 GHz 
15.7-17.7 GHz 


18-27 GHz 

24.05-24.25 GHz 


27-40 GHz 

33.4-36 GHz 


40-75 GHz 

59-64 GHz 


75-1 10 GHz 

76-81 GHz 
92-100 GHz 


110-300 GHz 

126-142 GHz 
144-149 GHz 
231-235 GHz 
238-248 GHz 

a IEEE Std 521-1984. 

the International Telecommunications Union (ITU) are available for use. This categorisation 
does not mandate the use of a particular band or frequency but merely indicates that it is 
available to be used. Other factors decide which band to be used in a particular application: 
most notable are the effects of atmospheric absorption and the size of antenna that the 
platform can reasonably accommodate. 

The effect of atmospheric absorption is a constraint depending upon physics that is totally 
outside the control of the designer. Physical antenna size is to some extent under the control 
of the designer, although the platform dimensions will be determined by factors relating to 
its airborne performance, range and so on. As for many systems, the design of a radar system 
is subject to many considerations and trade-offs as the designer attempts to reconcile all the 
relevant drivers to obtain an optimum solution. 

The effects of atmospheric absorption are shown in Figure 3.4. The diagram illustrates the 
loss in dB per kilometre across the frequency spectrum from 1 to 300 GHz. This curve varies 
at various altitudes - the particular characteristic shown is for sea level. A 10 dB loss is 
equivalent to a tenfold loss of signal, so the loss per kilometre at 60 GHz is almost a 1000 
times worse than the loss at around 80 GHz. These peaks of atmospheric absorption occur at 
the resonant frequency of various molecules in the atmosphere: H 2 at 22 and 185 GHz and 
2 at 60 and 120 GHz, with the resonance at 60 GHz being particularly severe. 

Also shown on the diagram are four key frequency bands used by some of the weapons 
systems of today: 






1 - 


2 - 


3 - 

Attack Helicopter 

4 - 

Anti-tank Missile 


1 10 50 100 200 300 

Frequency (GHz) 

Figure 3.4 Effects of atmospheric attenuation. 

• Surveillance radar operating at ~3 GHz; 

• Fighter radar radiating from 10 to 18 GHz; 

• Attack helicopter operating at 35 GHz; 

• Anti-armour missile transmitting at 94 GHz. 

It can be seen that the atmospheric absorption effects have a significant impact upon the 
portions of the spectrum that the radar designer can reasonably utilise. 

The basic principle used by radar is portrayed in Figure 3.5. The energy radiating from a 
radar transmitter propagates in a similar fashion to the way ripples spread from an object 
dropped in water. If the radiated energy strikes an object - such as an aircraft - a small 



Figure 3.5 Basic principles of radar. 





Figure 3.6 Pulse and continuous waveA 

proportion of that energy is reflected back towards the radar. The transmitted energy 
effectively has a double journey: out to the target and back again. Radar uses this principle 
to measure the distance to the target; knowing that the speed of light is ^3 x 10 8 m/s, and 
by measuring the time taken for the reflection to arrive, makes it possible to calculate the 
target range: 

R = cxAt 

where R is the range of the target, c is the speed of light (3 x 10 8 m/s) and A^ is the time 
taken for the radar energy to perform the round trip. 

Radar energy may also be transmitted in a number of ways. Figure 3.6 shows two 
situations; one where the RF energy is sent in pulses and the other where RF energy is 
radiated continuously - also known as a continuous wave. 

Pulsed radar transmission is useful when information is required regarding the range 
of a target. Clearly, by transmitting a pulse of radar energy it is easy to measure when 
the reflected pulse returns and hence determine the target range using the formula given 

Using a continuous wave transmission allows the closing (or receding) velocity of the 
target to be determined. This is achieved by using the Doppler effect. The Doppler effect is 
one by which the frequency of radiation is affected if a target is moving in the radial 
direction between radar and target (see Figure 3.7 which depicts a point radiating source 
travelling with a velocity from left to right). 

If a radiating (or reflecting) target is receding from an observer, the frequency will appear 
to reduce as far as the observer is concerned. Conversely, if the target is approaching then the 
frequency will appear to increase. The classic illustration is of a train approaching, passing 
and receding from a stationary observer: as the train approaches, the sound pitch will be 
higher than when it recedes. As will be seen, the Doppler effect is a very useful property that 
is extensively used in various radar applications. 

3.2 Radar Antenna Characteristics 

In Figure 3.5 it is implied that the radar energy is directed in some manner in the direction of 
the target, and this is in fact the case. In order to achieve such directed energy, early systems 
used parabolic reflectors in which radiated energy is directed towards the reflector from a 
radiating horn at the focal point. Reflected energy returning from the target is 'gathered' by 
the reflector and concentrated at the focal point where the horn feed is situated. In this way 
the radiated energy is directed towards the target and the reflected energy is gathered from it, 
as illustrated in Figure 3.8. 



Frequency Decreasing 






Figure 3.7 Doppler effect due to motion. 

Frequency Increasing 




In recent years, radars have adopted the planar array shown in the lower part of Figure 3.8. 
Whereas the parabolic reflector achieves beam shaping by means of its physical parabolic 
shape, the planar array achieves a similar effect by careful phasing within the RF feeds at the 
rear of the planar array. This principle is described in more detail in Chapter 4. 

This radar antenna directional property is extremely important to the radar as it focuses the 
energy into a beam on transmission and effectively 'gathers' the reflected energy during recep- 
tion. This directional property enhances the operation of the radar and is known as the antenna 
'gain'. The gain depends upon the size of the antenna and the frequency of the radiated energy. 

The beamwidth of an antenna is actually quite a complex function, for, during the 
formation of the main beam, which is at the heart of the performance of the radar (owing to 
the antenna gain), a number of sublobes are also created which are distinctly unhelpful. 


Figure 3.8 Parabolic reflector and planar array. 



I / 

Second Sidelobe 

First Sidelobe 


Main Beam 

\ \ n , „. , First Sidelobe 

Second Sidelobe 





3 Decibels 














Figure 3.9 Antenna pattern showing main beam and sidelobes. 

These sublobes are known as sidelobes. Whereas the main beam is central to the 
performance of the radar, the sidelobes detract from the radar performance since they 
effectively waste energy and have other adverse effects. 

As shown in Figure 3.9, the upper part of the diagram portrays the main beam and 
sidelobes generated by a typical directional antenna. The area of interest to the radar is the 
main beam. The sidelobes are characterised as the first sidelobe, the second sidelobe, etc., 
and for the antenna pattern shown there are five sidelobes, each becoming progressively 
weaker the further off-boresight (the radar centre-line) they are. Not only do the sidelobes 
waste energy during transmission by directing energy away from the target, they also allow 
stray and unwanted energy to enter the antenna and therefore the radar receiver during 
reception. Stray energy may be noise or 'clutter' produced by spurious reflections from the 
ground, alternatively it may be energy being transmitted by an enemy jammer who is 
attempting to confuse the radar. The actual beam pattern is determined by a specific 
mathematical relationship known as a sinx/x waveform. The beamwidth of the main 
beam is defined as the point at which the signal strength of the sinx/x waveform has 
dropped to 3 dB below (— 3 dB) the peak value. In numerical terms this relates to a signal 
level \/y/2 below the peak signal which equates to 0.707 of the peak level. 

The beamwidth varies according to the mode in which the radar is operating and the 
information it is trying to gather. The beamwidth also does not necessarily have to be the 
same in both axes (azimuth and elevation), as will be described. For an air-to-air mode 
the beamwidth will be narrow and be equal in azimuth and elevation, whereas for a ground- 
mapping mode it will be narrow in azimuth and broad in elevation. 

The term decibel relates to a measurement unit that is used extensively within the radar 
community to describe relative signal levels in a short-hand logarithmic form according to a 
base of 10. In radar calculations, dynamic range between two signals may be several orders 



1 Decreasing 




Increasing I 


S 1 ?' 3 

10" 2 

10" 1 


10 1 

10 2 

1 ? 3 S 





Signal Ratio 




Figure 3.10 Decibels compared with numerical format. 

of magnitude, and continuously using many noughts is tiresome and confusing. The principle 
is shown in Figure 3.10. 

It can be seen that, as the signal strength increases to the right, so the decibels increase. At 
a positive signal ratio of 10 3 or 1000, the decibel figure is 30 dB [log 10 ( 1000) is 30], whereas 
for a decreasing signal ratio, a negative ratio of 10~ 3 or 1/1000, the decibel figure is —30 dB 
[log 10 (l/1000) is —30]. It is therefore easier to say that in a good antenna design the first 
sidelobe is 30 dB down (or — 30 dB) on the main lobe signal than saying that the signal 
ratio is l/1000th or 10~ 3 below the main lobe (this is a realistic figure in practice). Therefore, 
within radar terminology the decibel notation is used liberally to describe gains or losses 
when considering system performance. 

3.3 Major Radar Modes 

Examining the operation of some basic radar modes helps to understand how beamwidth and 
other factors such as pulse width, scan patterns, dwell time and pulse repetition frequency 
(PRF) are important to radar operation. The main modes that are described are: 

• Air-to-air search; 

• Air-to-air tracking; 

• Air-to-air track- while-scan (TWS); 

• Ground mapping. 

3.3. 1 Air-to-Air Search 

One of the functions of a fighter aircraft is to be able to search large volumes of air space to 
detect targets. Many scan patterns are able to accomplish this function, but perhaps the most 
common is the four-bar scan shown in Figure 3.11. This scan comprises four bars stacked in 
elevation, and the radar mechanically scans from side to side in azimuth while following the 
four-bar pattern. The pattern shown begins in the top left-hand corner and finishes in the 
bottom left-hand corner before recommencing another cycle. The scan might typically cover 
±30° in azimuth centred about the aircraft centre-line and about 10-12° in elevation. 
Alternatively, sector scans may be used, say ±10° skewed left or right off the centre-line if 
that is where the targets are located. The beamwidth in the air-to-air search mode will 
probably be ~3°, and the scan bars will usually to be positioned one beamwidth or 3 dB 
apart to ensure that no target falls between bars. The search pattern is organised such that a 
target may be illuminated several times during each pass, as indicated by the overlapping 



Azimuth - 

Figure 3.11 Typical air-to-air search pattern. 

antenna coverage shown in the figure. This allows the target to be detected with certainty and 
avoids nuisance detections or false alarms. 

More importantly, integration techniques (summing the return from several successive 
pulses) allow the signal return to be enhanced and therefore the ability to detect the target in 
noise or clutter to be significantly improved. For similar reasons, integration techniques can 
enhance the range at which the target is initially detected. In earlier-generation radars, all the 
targets detected would be shown on the radar display and could lead to a confusing picture, 
particularly when trying to separate friend from foe. More modern radars with digital 
processors are able to categorise multiple targets more easily, thereby simplifying the 
engagement procedure. 

It should be recognised that, when the radar is operating in this air-to-air search mode, 
enemy targets fitted with a radar warning receiver (RWR) or other detection equipment will 
know that they are being illuminated or 'painted' by the searching radar. Furthermore, by 
categorising radar signal parameters such as radiated frequency, pulse width and PRF, the 
enemy target will be able to identify what type of radar and what aircraft type is being 
encountered. The RWR will also give a bearing to the illuminating radar, and the 'blip/scan' 
ratio will indicate whereabouts in the radar scan pattern it is located. It will also be obvious 
that the radar energy only has to travel a single path to reach the potential target, whereas the 
energy has to travel out and back to the radar to produce a return. This means that the target 
aircraft will be receiving a much stronger signal than the radar. Therefore, while air-to-air 
search is a useful mode, radar operators should also be aware that at the same time they are 
also giving potentially useful information to their adversary. 

3.3.2 Air-to-Air Tracking 

On occasions the radar may need to obtain more pertinent data regarding the target, perhaps 
in order to prepare to launch an air-to-air missile. To attain this more specific target 



Target Data Set: 
Range Range Rate 
Azimuth Elevation 
Angular Rates 

Figure 3.12 Air-to-air tracking. 

information, the radar needs to 'lock on' to the target. When this occurs the scan pattern 
changes and the radar antenna tracks the target in azimuth and elevation. The target is also 
locked in terms of range using a range gate. The radar is now able to track the precise 
movements of the target. In some tracking modes the PRF may be switched to higher 
frequency to increase the target data update rate (Figure 3.12). 
The target dataset will include the following data: 

• Range; 

• Azimuth; 

• Azimuth rate; 

• Target identification; 

• Range rate; 

• Elevation; 

• Elevation rate; 

• Target classification. 

The accompanying changes in the radar characteristics detected by the potential target 
following lock-on is a warning that the engagement is becoming more serious. At this point 
the target may attempt evasive tactics - deploy countermeasures or chaff or jam the target 

The means by which the radar achieves target angle tracking and target range is described 
later in the chapter. 

3.3.3 Air-to- Air Track-While-Scan 

The disadvantage of locking on to the target and thereby signifying engagement intentions 
has already been described, but the advent of digital signal/data processing has enabled an 




Figure 3.13 Air-to-air TWS. 

elegant solution to be developed. Track- while- scan automates the process of deciding which 
target to engage (Figure 3.13). As TWS is under way, the radar processor progressively 
builds up a history of the flight path of targets within the scan. If successive measurements 
disagree, then the track is rejected; if the data agree, then the track is maintained. Gates are 
initiated that assign angular information, range and range rate to each track and predict 
where the target will be at the time of the next observation. If the track is stable, then the 
forecast gates will become more accurate and statistical filters will establish that the 
predicted fit is good. Techniques are used to arbitrate when gates overlap or where more 
than one target appears in the same gate perimeter. 
The advantages of TWS are as follows: 

1. Accurate digitised tracking data are established on each track within the antenna scan 
pattern without alerting potential targets that they are being tracked. 

2. The automation process allows many targets to be tracked accurately and independently. 

3. A typical radar using TWS will be able to track 20 or more targets in three-dimensional 

3.3.4 Ground Mapping 

From the early days of radar it was known that the radar could be used to map the terrain 
ahead of the aircraft. Using the different reflective characteristics of land, water, buildings, 
etc., it was possible to paint a representative map of the terrain ahead of the aircraft where 
major features could be identified. With the application of digital processing and advanced 
signal processing techniques, the ability to resolve smaller features increased and high- 
resolution mapping became possible. 





Figure 3.14 Ground mapping. 

While using the ground-mapping mode, the antenna sweeps from side to side as shown in 
Figure 3.14. The area illuminated by the mapping beam equates to the dotted boundary 
shown in the figure. 

Whereas the air-to-air modes use a narrow pencil beam, a fan beam is used for ground 
mapping. That is, a beam where one dimension is narrow — 2 or 3° - while the other is 
relatively broad, say 10 to 15°. The figure shows that the ground-mapping beam is narrow in 
azimuth and wide in elevation; this represents the optimum shape for the mapping function. 

This describes the operation of a basic ground-mapping mode. In the past 20 years, 
improved capability and flexibility have been achieved by the use of digital computing in the 
radar data processor (RDP) and presignal processor (PSP). By the use of fast Fourier 
transform (FFT) for signal processing and aircraft motion compensation, increasingly 
sophisticated radar modes have been developed. These include: 

• Doppler beam sharpening (DBS); 

• Synthetic aperture radar (SAR); 

• Inverse Synthetic Aperture Radar (ISAR). 

These modes will be described in Chapter 4. 

3.4 Antenna Directional Properties 

Earlier it was stated that the directional properties of an antenna were determined by the 
radiated frequency and the size of the antenna. There are simple formulae that help to 
estimate the beamwidth and gain of an antenna if these parameters are known. 


The frequency and wavelength of an electromagnetic wave are related to each other and 
the speed of light, c, by the equation 


where c is the speed of light (3 x 10 8 m/s),/ is the frequency (Hz) and A is the wavelength 

Therefore, for the airborne fighter radar described earlier, operating at a frequency 
of 10 GHz (10 10 Hz), X = c/f = 3x 10 8 /10 10 = 3/100 or 0.03 m or 3 cm. 

Another formula is a ready approximation to determine the beamwidth of an antenna 
knowing the frequency and the antenna size: 

„ 65 x A 


where is the beamwidth (deg), A is the wavelength (m) and D is the antenna dimension 

Using again the example of the airborne fighter radar, and assuming an antenna dimension 
of 0.6 m (-24 in), « 65 x 0.03/0.6 « 3.5°. 

Using a similar approximation, it is possible to estimate the antenna gain: 

_ 4 X 7T 

Ob x <£b 

where Gd is the antenna gain, B is the beamwidth (rad) in one axis and (/?b is the beamwidth 
(rad) in the orthogonal axis (one radian ^57.3°). 

Again using the fighter radar example, Gd = 4 x 7r x (57.3) /(3.5) w 3368. This gives 
an idea of the advantage that the antenna gain confers. Expressed in decibels, the antenna has 
a gain of log 10 (3368) or around 35 dB. 

3.5 Pulsed Radar Architecture 

The basic principles of radar operation have already been outlined. The detailed operation of 
pulse radar is described in this section. A top-level diagram of a pulsed radar system is 
shown in Figure 3.15. 

3.5. 1 Pulsed Radar Components 

The diagram shows the major elements which are: 

• Modulator; 

• Transmitter; 

• Antenna; 

• Receiver; 

• Video processor. 
















*n t rJ 

Power Supply 

Figure 3.15 Top-level pulsed radar architecture. Modulator 

The modulator determines the pulse shape and the nature of the radar modulation. Although 
pulsed transmission is the most elementary form of radar operation, the modulation in a 
modern multimode radar may take many forms depending upon the nature of information 
being sought. The operation of the modulator is controlled by the synchroniser which 
dictates when a pulse should be initiated. The modulator uses the superheterodyne ('super- 
het') principle of modulation to superimpose the modulating signal upon the high-frequency 
carrier to provide a composite waveform. Transmitter 

The transmitter amplifies the modulated carrier signal and feeds it to the antenna via a 
duplexer. This serves the function of directing the transmitter energy to the antenna 
waveguide system to be fed by the antenna elements for transmission into the atmosphere. 
It also routes the reflected target energy to the receiver. Antenna 

The antenna, as has been described, directs the radar energy towards the target and receives 
the reflected energy from the target. Along with the target echo, a substantial amount of 
clutter from ground returns is also received. The antenna beam is focused according to the 
shape of the antenna and the nature of the beam required. Unwanted radar energy enters 
through the antenna sidelobes as well as the main beam. The antenna also receives noise 
from a variety of external sources that can help mask the true target signal. 



Returning energy is passed through a receiver protective device which blocks the large 
amounts of transmitted power that would cause severe damage to the receiver, but also at the 
appropriate time allows the reflected target energy to pass through. Receiver 

The receiver amplifies the reflected target signal and performs the demodulation process to 
extract the target data from the surrounding noise, and the resulting target video data are 
passed to the video processor. Video Processor 

The video processor is also controlled by the synchroniser in order that transmitted pulse and 
target return pulses are coordinated and that a range measurement may be made. The 
resulting data are coordinated and displayed on the radar display. 

3.5.2 Pulsed Modulation 

The nature of the pulse modulation in terms of pulse width and frequency of repetition is 
highly interactive with a number of important radar characteristics and has a significant 
impact upon the performance of the radar. The basic parameters of a pulsed radar signal are 
described in Figure 3.16. 

In pulsed radar operation, the carrier frequency is modulated by the envelope of a single 
rectangular pulse; in this case the pulse embraces a fixed carrier frequency. As will 
subsequently be discovered, in sophisticated radar operations there are advanced forms of 
modulation/transmission in which the pulse is not rectangular nor the carrier fixed in 
frequency. The pulse width is denoted by the symbol r and is usually fairly narrow, perhaps 






-Pulse Period- 



Frequency (fo) 


A f M 

fo - 1/ —I I fo + 1 



Figure 3.16 Pulsed radar transmission. 


~1 |is in an air-to-air mode. After a time interval called the pulse period, a second pulse is 
transmitted and the sequence is repeated. The rate at which the pulses are repeated is called 
the pulse repetition frequency (PRF), and both the pulse width r and the PRF are key radar 

It will be noted on the diagram that the terms 'time domain' and 'frequency domain' are 
mentioned. The time domain is familiar in everyday life as it is the domain in which we live; 
the frequency domain is more abstract but is of great importance to the radar designer. In 
fact, the time and frequency domains are interdependent and interwoven, and this has a 
significant impact upon the operation of radar systems. What happens in the time domain 
affects the frequency domain, and vice versa. 

The rectangular pulse r results in a response in the frequency domain that has a sinx/x 
response, the same generic response that determines the pattern of the antenna main beam 
and sidelobes. However, in this case the response is occurring on an axis relating to 
frequency rather than angle off-boresight as is the case in the antenna pattern. When the 
incoming pulse is received, it results in a frequency response of received power portrayed by 
the sin x/x response and centred upon the radiated frequency /q. The practical limits of the 
main sin x/x response are ±/i centred onf , that is,/o + l/r;/o — 1/r (/o being the carrier 
frequency and r the pulse width), and this determines the bandwidth required of the receiver 
in order to be able to pass all the components of the target return. Therefore, for a 1 (is pulse 
the receiver bandwidth would need to be 2/r = 2/(1 x 10~ 6 ) = 2 x 10 6 or 2 MHz (see the 
lower part of Figure 3.16). The narrower the transmitted pulse, the wider will be the 
bandwidth; the converse also applies. 

In basic pulsed radar operation the pulse width also determines the range resolution of 
which the radar is capable. The radar can only resolve to half the pulse width, as at a lower 
interval than this part of the pulse has been reflected while part has not yet reached the target. 
A pulse of 1 |is duration will be approximately 1000 ft long (as light travels at 3.3 x 3 x 10 8 
or ~ 10 9 ft/s and the duration of the pulse is 1 x 10 -6 ; distance = velocity x time). There- 
fore, a 1 (is pulse will be able to resolve the target range to no less than 500 ft. In fact, by 
using more complex modulation and demodulation methods called pulse compression, it is 
possible to achieve much better resolution than this; pulse compression is discussed in 
Chapter 4. 

The pulse period - and therefore PRF - also has an impact upon the radar design that 
affects target ambiguity, as shown in Figure 3.17. The figure shows an aircraft illuminating 
two targets and compares the effect of the returns from these targets for two different 
pulse periods. In pulse period T\ (PRF : = \IT\), the returns from both targets are received 
before the successive pulse is transmitted, and the range is unambiguous. In the case of 
the shorter pulse period T 2 (PRF 2 = 1/72), the return from the most distant target occurs 
after the transmission of the successive pulse and to the radar appears as a relatively 
close target within the second period. In this case the target range is ambiguous and 
misleading. The selection of PRF is one of the most difficult choices the radar designer 
has to make, and some of the effects of range ambiguity are discussed in more detail in 
Chapter 4. 

The fact that the radar only transmits for a portion of the time means that the average 
power is relative low. The average power is given by the following expression: 


X T 

where P pQi ± is the peak power, r is the pulse width and T is the pulse period. 



Period, T 


1000 - 



H h 



PRF (kHz) 

12 3 4 5 

Figure 3.17 Effect of pulse period on target ambiguity. 

For a peak power of 10 kW, a pulse width of 1 u.s and a pulse period of 250 (is, we have 

10 4 x 1 x 10 6 

-* av — 

10 6 x 250 

40 W 

3.5.3 Receiver Characteristics 

In order to be able to detect the target, the radar receiver has to able to discriminate from 
unwanted effects. The main adverse affects are as follows: 

1. Noise is either internally generated or radar transmitter induced or externally sourced. 
Noise is random and can only be minimised by good design. 

2. Clutter due to unwanted returns from the ground and other sources is usually more 
systematic and can be countered by filtering and processing techniques. 

3.5.3. 1 Noise 

The sources of noise that can affect the ability of the radar receiver to detect a target 
signal are shown in Figure 3.18. The total system noise includes noise from the following 

1. Antenna noise T a . The antenna noise includes all those sources of noise that are external 
to the radar, including radiation from the sun, terrain, emissions from man-made objects 




Te Equipment Noise 


Tr Transmission 
Line Noise 

Ta Antenna Noise 


Ts System Noise 
Figure 3.18 Sources of noise affecting radar signal. 

and the weather. Noise from jamming may also be included in this category. The radome 
and the antenna itself may also generate noise. While external noise will be most 
troublesome when it enters the system via the antenna main beam, it should also be 
remembered that noise can also enter via the antenna sidelobes. 

2. Transmission line noise T Y . This includes noise originating within the waveguide couplers, 
duplexer and the receiver protection device. 

3. Equipment noise T e . The equipment noise is generated within the receiver itself and is the 
most difficult to counter. 

The total system noise, T s , is the sum of these individual components: 

t s = r a + T r + r e 

The problem with the noise in a receiver is that, once present, it is there to stay. Signal 
amplification in subsequent stages will only amplify the noise as well as the signal and 
accentuate the problem of target detection. One technique commonly used is to insert a low- 
noise amplifier (LNA) at the front of the receiver to amplify the signal proportionately more 
than the noise. LNAs are also commonly used where antennas (or apertures) are mounted 
remotely throughout the airframe and where transmission losses might be relatively high. 
The receiver noise is defined as noise per unit of receiver gain: 

Receiver noise 

noise at output of receiver 
receiver gain 


The receiver gain can be easily measured using laboratory techniques. 

The receiver noise may be characterised by a figure of merit or noise figure F n . This is 
defined as the ratio of the noise figure of the actual (imperfect) receiver to the hypothetical 
ideal receiver providing equal gain. Therefore: 

noise output of actual receiver 
noise output of ideal receiver 

An ideal receiver would produce no noise; the only noise that would exist would be that from 
external sources. This external noise can be represented as though resulting from thermal 
agitation in a conductor (resistor) since the two have similar spectral characteristics. 
Therefore, in the derivation of F n , for both ideal and actual receivers, the thermal noise 
can be portrayed as the voltage across a resistor. Thermal noise is governed by the random 
motion of the free electrons within the conductor and is uniformly spread across the entire 
spectrum. This motion is determined by the absolute temperature of the notional resistor, 
denoted by 7b. Also, the noise depends upon the receiver bandwidth B. Thus, to derive the 
mean noise power for an ideal receiver, the expression 

mean noise power = k x Tq x B (W) 

may be used for an ideal receiver, where k is Boltzmann's constant =1.38 x 10~ 23 W s/K, 
To is the absolute temperature of the resistor representing the external noise (K) and B is the 
receiver bandwidth (Hz). 

The external noise is the same for both receivers, and by convention To is taken to be 
290 K which is close to room temperature. Where the external noise is small by comparison 
with that generated by the receiver, as is usually the case, the mean noise figure for an actual 
receiver may be determined by the following: 

Mean noise power = F n x k x r x B (W) 

As was shown earlier, the total noise may represented by T s where the mean noise power 
(all sources) = k x T s x B. 

The nature of the modulation used also has an impact upon receiver noise. This is shown 
in Figure 3.19. The figure shows the simple comparison of narrow and broad rectangular 
pulse modulation. It was shown earlier in Figure 3.16 that the bandwidth needed to 
accommodate all the frequency components of a rectangular pulse was governed by the 
sinx/x waveform, and that the theoretical bandwidth was 2/r. The narrow (sharper) pulse t\ 
needs a greater bandwidth than the broader pulse ti. The narrow pulse gives an improved 
range resolution and, for a given pulse period (PRF), a reduced mean power, so it can be seen 
that there are performance trade-offs to consider that affect bandwidth and hence receiver 

In practical systems a compromise is allowed and generally a bandwidth of 1/r is 
regarded as sufficient. Therefore, it is common practice to narrow the IF bandpass filter until 
it is 1/r wide, just wide enough to pass the bulk of the target-related energy but reject the 
unwanted noise. This design is called a matched filter, and the mean noise energy per pulse 
is kTo/r. 

In the Doppler radars addressed in Chapter 4 the Doppler filters downstream of the IF filter 
are much finer, and greater noise and clutter rejection result. 



Figure 3.19 Effect of different pulses on the receiver bandwidth. 

The detection and extraction of a target echo from a background of noise depends upon the 
four factors outlined below: 

• The average power radiated in the direction of the target; 

• The proportion of the radiated energy reflected back in the direction of the radar; 

• The proportion of power recaptured by the radar antenna; 

• The length of time the antenna beam is trained upon the target. 

Average power is determined by the relationship of the peak power, P pea io transmitted by the 
radar and the modulation characteristics of pulse width, r, and pulse period, T 9 as shown in 
the previous section. The antenna gain, Gd, also increases the power density related to the 
beamwidth(s) and beam geometry. 

As the radiated signal is directed towards the target, it spreads out an increasing area, 
proportional to R 2 , where R is the range from the radar. This means that the power density 
reduces by a factor of 1 /R 2 as the energy is propagated in the direction of the target. 

A fraction of the energy incident upon the target will be reflected back in the direction of 
the radar. In the simplest form the target may be considered to be a simple sphere with a 
specific cross-sectional area, denoted by the symbol a and specified in square metres. The 
reality is much more complicated than that, and other factors such as reflectivity and 
directivity play a great part, as will be seen in the discussion on low observability or stealth 
in Chapter 4. 

As the energy is reflected back to the target, the \/R 2 effect applies in terms of the 
reduction in received power density. The impact of this effect means that the energy received 
at the radar has been reduced by a total factor of 1 /R 4 in its outward and return path to and 
from the target. This has an impact upon the ability of the target signal to be detected above 
the noise, as shown in Figure 3.20. The figure shows how the returning signal (not to scale) 




Target Range 

Figure 3.20 Effect of range upon the target echo. 

decreases with increasing range to the point where the signal is not detectable against the 
noise background. 

As will be seen, the equation governing the strength of the return signal is a fourth-power 
law, and this means that the receiver has to accommodate a very large dynamic excursion 
in terms of variation in target signal strength as the range varies. In certain modes this is 
addressed by a technique called sensitivity time control (STC) in which the receiver gain 
is reduced at very short ranges and increased progressively during the range sweep. This 
technique is sometimes referred to as swept gain and to some extent mitigates the problem of 
extremely high signal returns at short range. 

Another technique is often used to counter this effect and prevent the receiver amplifiers 
from saturating: if the receivers saturate, then both signal and noise will merge as the amplifiers 
clip both noise and target signal returns. In this case, automatic gain control (AGC), as the 
name suggests, automatically reduces amplifier gain to prevent saturation occurring. 

The actual detection of the target signal is determined by the setting of a target detection 
threshold as shown in Figure 3.21. This shows two targets, A and B, against a background of 
noise on a time axis: A and B are obviously at different ranges from the radar. The figure 
shows the importance of setting the target threshold correctly with respect to the mean noise 
level. If the threshold is set low, then it may be anticipated that more targets may be detected. 
However, as the diagram shows, setting a low target threshold has the accompanying risk of 
detecting a spurious target - called a false alarm. For the low threshold setting shown, the 
radar would detect three targets: genuine targets A and B and the false alarm. 

Conversely, there are problems with setting the threshold too high to avoid false alarms. In 
this case the return from genuine target A is lost and only target B is detected. 

One of the major factors affecting target detection was antenna time on the target. So far, 
only the detection of a target using a single pulse has been considered. In fact, as the radar 
beam sweeps through the target, a number of successive pulses will illuminate the target in a 
short period of time. Most radars have the capability of integrating the detected output over a 
number of pulses, and this has significant advantages, as can be seen from Figure 3.22. 



Target B 

Target A 

High Threshold 

Low Threshold 


Figure 3.21 Receiver threshold setting. 

Noise is generally random in terms of amplitude and phase. The target return is more 
systematic and repetitive in nature, at least over a range of successive pulses in an antenna 
scan. The effect of integrating noise over a series of pulses is to end up with noise at more or 
less the mean noise level before integration. The converse is true for a real target return. The 
target return is aggregated during the integration process and the result is a much stronger 
target return. The figure shows that as an example - integration over 12 pulses produces an 


Target Detection Threshold 

i i i ■ i " ' 

Time on Target 
(Integration Period) 



Figure 3.22 Effect of integration over several pulses. 


integrated signal that comfortably exceeds the target detection threshold, whereas the 
integrated noise does not. This occurs in spite of the fact that each of the individual target 
signals are well below the target detection threshold and without pulse integration would 
each be subsumed by noise. This shows the powerful capability of 'extracting' a signal from 
noise using integration techniques. 

The actual antenna time on target depends upon a combination of three factors: 

• The antenna scan or slew rate; 

• The antenna beam width; 

• The PRE 

Taking some simple figures by way of illustration, if the antenna scan rate is 60 deg/s and the 
3 dB beam width is 3°, then the antenna will dwell upon a target for l/20th of a second. If a 
medium PRF of 1000 Hz is assumed, then the antenna will theoretically have a total of 50 
'hits 'on the target during every pass across the target. Clutter 

The effects of clutter, particularly from ground returns or precipitation, can cause large 
amounts of unwanted signal being returned to the receiver. Clutter can enter the receiver 
channel through the main beam or via the sidelobes. It can depend upon the nature of the 
terrain, terrain geometry and the aspect (depression angle) of the antenna boresight. If the 
clutter is from the water, then it may depend upon sea state (the height of the waves or 
the smoothness of the water surface). In some ways, clutter may be systematic in terms of the 
effect that it has on the radar, and in these cases it is easier to counter or filter. 

Moving targets, or targets with a significant radial velocity with respect to the radar, may 
have a Doppler shift component that may enable the target to be distinguished against a 
stationery background. The use of Doppler filters and target velocity techniques is described 
in Chapter 4. 

3.5.4 Radar Range Equation 

The foregoing discussion leads us to the equation that is the most powerful and commonly 
used when examining the performance of radar systems, that is, the radar range equation. 
The radar range equation takes many forms depending upon those factors that need to be 
taken into account and the type of transmission being considered. In the simplest form, the 
maximum range for a single radar pulse is determined by the following equation: 

(47r) 2 X Smin 


where R is the radar range (m), P pea k is the peak power (W), G is the antenna gain (m ) (this 
may also be expressed in decibels for ease of calculation, as explained earlier), a is the target 
cross-sectional area (m 2 ), r is the transmitted pulse width (s), and S m m is the minimum 
detectable signal energy (W-s). This equation does not take account of pulse integration. 


There are some interesting observations to make regarding this formula: 

1. Peak power P pea k. As the peak power only affects the radar range by the inverse fourth 
power, doubling the peak power of the radar only increases the range by the inverse fourth 
power of 2 « 1.19 or 19%. 

2. Antenna gain G. If the antenna is circular, doubling the size of the antenna will increase 
the gain of the antenna by 4, and the overall range by a factor 1/2 or by about 71%. 
However, commensurate with the antenna gain, the beamwidth would halve, which may 
make target acquisition more difficult. Dwell time might also have to increase to improve 
target integration. Altering the wavelength of the radiated transmission would have an 
effect upon radar range as the range alters by the inverse square of the wavelength. 
Decreasing wavelength or increasing frequency can therefore increase the range. The 
atmospheric absorption outlined in Figure 3.4 earlier in the chapter will be an important 
factor, as at certain parts of the spectrum absorption rates are punitive, more than 
cancelling out any benefit that increasing radiated frequency may confer. 

3. Target cross-sectional area a. Reducing the target cross- sectional area by a factor of 60 dB 
(equivalent to 1 x 10~ 6 ) by using extensive low observability (LO) techniques reduces the 
range by a factor of ~30. 

4. Pulse width r. Maintaining mean power but decreasing the pulse width increases the peak 
power but also increases the receiver bandwidth, allowing more noise into the receiver. 

5. Minimum detectable signal S m { n . Decreasing the minimum detectable signal increases the 
radar range, but the risk of false alarms may increase. 

As more factors are taken into account, so more trade-offs need to be made. However, as will 
be seen later, the adoption of sophisticated modulation and signal processing techniques can 
gain significant performance enhancements in modern digital radars. 

3.6 Doppler Radar 

In the early part of the chapter the Doppler effect was described, that is, the effect 
upon radiated frequency when a moving source approaches or recedes from an observer. 
The same effect occurs when radar energy is reflected by ground clutter, except that the 
Doppler frequency shift is doubled as the radio energy has to travel out and back to the 
radar. Normally, ground returns are a nuisance as far as the radar is concerned, and all 
means are used to reject the ground clutter. However, there is one radar application where 
the ground clutter Doppler frequency shift is utilised, and that is the Doppler radar, some- 
times called the Doppler navigator. A typical configuration for a Doppler radar is shown in 
Figure 3.23. 

The Doppler radar comprises three or four narrow, continuous wave radar beams angled 
down from the horizontal and skewed to the left and right of the centre-line. The three-beam 
layout shown in the figure is called a lambda configuration for obvious reasons. The diagram 
shows a situation where the aircraft is flying straight ahead, i.e. heading equals track and 
there is no angle of drift because of crosswind. The forward beams 2 and 3 will experience a 
positive Doppler shift as the ground is advancing towards the aircraft. The Doppler shift Af 
is proportional to 2V/X, where V is the aircraft forward velocity and A is the wavelength of 
the radiated frequency. The aft beam 1 will experience a negative Doppler shift proportional 



f + Af 

Figure 3.23 Doppler radar. 

to 2V/X as the ground is receding from the aircraft. There are several scaling factors 
including direction cosines associated with the beam, but subtracting forward and aft beams 
yields a signal proportional to 4V/X. Therefore, by manipulating and scaling the Doppler 
shifted returns from all three beams, the aircraft horizontal velocity with respect to the 
ground (i.e. ground speed), V x , may be calculated. 

If the aircraft is drifting left or right owing to a cross-wind, then, by using the cross-track 
Doppler shift components and a similar manipulation process, the cross-track velocity, V y , may 
be calculated. The vertical velocity component, V z , may also be calculated. The vector sum of 
V x , V y and V z enables the total aircraft velocity, V, to be established. Doppler radars do have 
one disadvantage: if the terrain is very flat with a low reflectivity coefficient, then insufficient 
energy may be reflected back to the radar and the Doppler shift cannot be measured. Such 
effects can be achieved when travelling over very calm water or ice-covered expanses of water. 
Doppler radars were very commonly used before inertial navigation systems (INS) became the 
norm about 30 years ago; more recently, INS has been augmented by the satellite-based global 
positioning system (GPS). The initial avionics configuration of Tornado included a Doppler 
radar, and they are still frequently used on helicopters as air data becomes very unreliable at 
low airspeeds. For further data on Doppler radars, see Kayton and Freid (1997). 

As will be explained in Chapter 4, sophisticated radars in use today combine the use of 
pulse techniques and Doppler to produce pulsed Doppler (PD) modes of operation. 

3.7 Other Uses of Radar 

3.7.1 Frequency Modulation Ranging 

The use of pulsed radar techniques has hitherto been described to measure target range. 
However, frequency modulation may also be used to determine range as depicted in 
Figure 3.24. 







Transmit Signal 

Receive Signal 


f 2 ^ 

I A 

/ / 

^v \ 

/ / 

fl ^ 

Af / 


/ / 

t / 

/ / 



1 t 




Figure 3.24 Frequency modulation ranging. 

The transmitted signal consists of a triangular wave modulation, as shown, that sweeps 
across the frequency spectrum, completing one cycle in 0.01 s in the example given. The 
received frequency will lag the transmitted frequency by an amount Af owing to the time 
taken to complete the out and return journey. The example shows a measurement taken when 
the reflected received frequency, /i, is compared with the current frequency at the transmitter, 
/2, with the difference in frequency being Af. The associated time difference signal, At, is 
proportional to the range of the target. 

The figures shown on the diagram relate to the use of this technique in a radar altimeter, 
where the radar returns are used to calculate the instantaneous altitude of the aircraft above 
the terrain over which the aircraft is flying. In this example, the transmitter is sweeping in a 
linear manner over a frequency range of 4250-4350 MHz in 0.01s. The use of radar 
altimeters is described in Chapter 7. 

3.7.2 Terrain-following Radar 

Whereas the radar altimeter is useful in informing pilots where they are in relation to the 
terrain underneath the aircraft, it does not tell them where the terrain is in front of the 
aircraft. To do this, the pilot needs to use a terrain avoidance (TA) mode or, better still, a 
dedicated terrain- following radar (TFR). The TA function can be crudely achieved by using a 
normal pulsed radar in a single-bar scan mode with a fixed depression angle. This will tell 
the pilot where he is in relation to the terrain ahead of the aircraft, but it is not a sophisticated 
mode and does not readily lend itself to coupling into the autopilot (Figure 3.25). 

The TFR is a dedicated radar coupling into a dedicated functional system and autopilot 
that allows the pilot much greater performance and flexibility when penetrating at low level 
at night. The TFR scans the terrain ahead of the aircraft and receives ground returns that are 



Scan 1 




Figure-of- 1 
Eight Scanl 

3 1 7 5 13 

Figure 3.25 Terrain-following radar operation. 

used for guidance. Normally, a simple box scan is used where the active sweeps are those in 
the vertical direction (sections 1 and 3). In some circumstances a figure-of-eight scan is used 
which provides broader lateral coverage than the simple box scan. The TFR therefore builds 
up a range/elevation picture of the terrain ahead of the aircraft and calculates an imaginary 
'ski-toe' profile that reaches out ahead of the aircraft. This profile is calculated taking into 
account such factors as aircraft speed, manoeuvrability, etc., and provides an envelope within 
which the aircraft will not be able to avoid the terrain ahead. The system is configured so 
that, whenever the terrain ahead broaches the ski-toe envelope, the aircraft pitches up to 
rectify the situation. Similarly, if the terrain drops away in front of the aircraft, the aircraft 
pitches down until just operating outside the profile. The system operates just like the toe of a 
ski, moving up or down to follow the terrain ahead of the aircraft but always ensuring the 
aircraft can safely manoeuvre. 

The measurements from the radar altimeter are also fed into the terrain- following system 
which calculates the 'most nose-up command' provided by either TFR or radar altimeter. 
This has the advantage of providing the pilot with an additional altitude safety buffer directly 
beneath the aircraft as the TFR is looking several miles ahead. 

The TFR/radar altimeter commands may be coupled into the autopilot to provide an 
auto-TF mode while the aircraft is approaching the target area, thereby enabling the 
aircraft to fly at low level automatically while the pilot performs other mission-related 
tasks. The TFR may be an embedded system forming part of the aircraft primary radar, 
alternatively it may be provided in a pod that is loaded on to the aircraft. The AN/AAQ-13 
LANTIRN navigation pod fitted to F-15 and F-16 aircraft performs a TFR function for 
these aircraft. 

3.7.3 Continuous Wave Illumination 

On some weapons systems a continuous wave (CW) illumination mode is provided. This 
mode is used when aircraft are fitted with semi-active air-to-air missiles; that is, missiles that 
can receive incoming RF energy and once fired can track and engage the target. As the 



Target Data Set : 
Range Range Rate 

Azimuth Elevation 
Angular Rates 

Figure 3.26 CW illumination. 

missiles are unable to transmit, the aircraft radar has to provide the target illumination and 
it does this by using a CW illuminator co-boresighted with the aircraft radar antenna. 
Therefore, when the aircraft radar is locked on to the aircraft it can simultaneously illuminate 
the target (Figure 3.26). The disadvantage of this technique is that the aircraft radar has to 
remain locked on to the target and transmitting CW illumination until the engagement is 
complete. In high-density air-to-air combat this may not always be possible. 

3.7 A Multimode Operation 

Modern radars such as those on the F-15E and F-22 have the capability of operating 
simultaneously in a number of modes, an example of which is shown in Figure 3.27. In this 
hypothetical example, three simultaneous modes are depicted: 

Synthetic i 

Aperture \ \ 
Radar (SAR)U-J 

Target 1 

Track While 
Scan (TWS) 

Ground \ 
Mapping y 

Figure 3.27 Simultaneous multimode operation. 



• Sector ground mapping; 

• Synthetic aperture (SA) spot mode; 

• Track- while- scan (TWS) mode engaging three separate targets. 

The radar achieves this capability by interleafing the radar modulation required for each 
mode on a pulse-by-pulse basis and effectively operating as several radars in one. This offers 
immense flexibility to the aircraft as a weapons platform. 

3.8 Target Tracking 

During the pulsed radar tracking mode when the radar is locked on, it follows and 
automatically maintains key data with respect to the target: 

• Tracking in range; 

• Angle tracking in azimuth and elevation. 

Tracking is maintained and the radar is said to have 'target lock' when all these loops are 

3.8.1 Range Tracking 

Tracking in range is usually accomplished using a technique called range gating which 
automatically tracks the target as its range increases or decreases. The concept of the range 
gate is shown in Figure 3.28. 

The radar return in the region of the target return will comprise noise and the target return. 
The range gating technique uses two gates, an 'early gate' and a 'late gate'. The early gate 

Target Return 


Early Gate 

Late Gate 


Figure 3.28 Range gate tracking. 


-Tracking Gate Adjustment - 





is positioned near the leading edge of the target echo and detects and captures energy from 
the early part of the target return. Conversely, the late gate is positioned near the trailing edge 
of the target echo and detects and captures the energy from the trailing edge of the target 

The detected signals from the early and late gate are compared and the result is used to 
position the tracking gate so that it is coincident with the target return. In the example shown, 
both the early and late gates are positioned early (to the left) of the target return and the 
tracking gate is also incorrectly positioned. Consequently, the early gate detects less energy 
than the late gate. Identifying this discrepancy will cause the energy from early and late 
gates to be equalised and the tracking gate to be moved to the right (down range) so that it 
correctly coincides with the target echo. While the radar maintains target lock this process 
will be continued, maintaining the tracking gate at the same range as the target echo. 

3.8.2 Angle Tracking 

During the radar tracking mode the radar tracks the angle to the target in azimuth and 
elevation. In other words, the line-of-sight (LOS) to the target and the radar boresight are 
kept as close as possible. The LOS needs to be established within a frame of reference and 
usually the radar is stabilised in roll and pitch using attitude data from the aircraft attitude 
sources: inertial reference system (IRS) or secondary attitude and heading reference system 
(SAHRS). The final axis in the orthogonal reference set is usually the aircraft centre-line/ 

There are three main methods of angle tracking that are commonly used, these are: 

• Sequential lobing; 

• Conical scan (conscan); 

• monopulse. 

3.8.2. 1 Sequential Lobing 

One of the first tracking radar principles adopted was sequential lobing, which in its earliest 
form was used in a US Army angle tracking radar air defence radar. The principle of 
operation of sequential lobing is shown in Figures 3.29a and 3.29b. 

To track a target in one axis, two lobes are required; each lobe squints off the radar 
boresight. The centre point of where the two lobes overlap represents the boresight of 
the antenna and this is the LOS that the radar antenna is trying to maintain. It can be seen 
that, when the signal return from the target is the same in both beams, the LOS to the target 
has been achieved. As the target moves, continual error signals will be sensed and the 
antenna servo system responds by nulling the error and maintaining LOS to the target. If four 
lobes, A and B and C and D, are positioned as shown in Figure 3.29a, then lobes A and B 
provide tracking in elevation and lobes C and D provide tracking in azimuth. The reflected 
signal from the target received in each of the four lobes is routed via a channel switching 
assembly sequentially switched into the receiver. In this way, each of the four lobe returns is 
measured and error signals are derived to drive the antenna elevation and azimuth drive 

Figure 3.30 illustrates how each of the four lobes is switched in turn into the receiver. 
In practice, the waveguide switching arrangement is cumbersome and prone to losses; 



Two lobes are required to 
track in each axis. As each 
lobe must be sequentially 
switched a total of four 
pulses are required for 
tracking in both axes 



Beam A 

Beam B 

Figure 3.29 Sequential lobing - principle of operation. 

therefore, radar performance is compromised. The other significant disadvantage suffered by 
this method relates to the time taken for the sequencing to occur. The radar PRF will 
determine the maximum time that the receiver will be switched to a particular lobe. Only 
when the radar has completed the range sweep for a particular PRF can the receiver sequence 
to the next lobe. Furthermore, the elevation and azimuth error can only be updated once per 
cycle, and this adversely affects update rate and tracking error. 

Sequential lobing can be detected by a target and transmissions can be devised that will 
cause the radar to break lock. Transmitting on all four beams and receiving only on one may 
counter this. This technique is called lobe on receive only (LORO). Conical Scan 

Conical scan (conscan) is the logical development of the sequential lobbing scheme already 
described. Conscan is another example of the earliest form of angle tracking - used because 






A - B = Elevation Error 


A + B + C + Din sequence 



C - D = Azimuth Error 

Figure 3.30 Sequential lobing - tracking configuration. 

it was the easiest to use with the technology available at the time. The concept of conscan is 
depicted in Figure 3.31. 

In conscan, one lobe is used that squints off-boresight. The lobe is then rotated such that 
the target return is enclosed within the imaginary cone that is swept out around the antenna 
boresight. In a parabolic antenna the rotating conscan beam is achieved by rotating the 
antenna feed at the desired rate. In more sophisticated arrangements the feed may be nutated, 
and this can achieve a better tracking performance than the straightforward rotating feed at 
the expense of a more complex feed mechanism. Typical conscan rates may be in the range 
up to 50 Hz. 

Circular movement of 
radar beam allows 

target to be tracked by 
maintaining it at the 

centre of the conscan 

Figure 3.31 Conical scan - principle of operation. 








Figure 3.32 Conical scan - tracking configuration. 

If the target is located off-boresight, the receiver will receive an amplitude-modulated 
signal at the conscan frequency. By detecting the signal and resolving the error components, 
drive signals are fed to the azimuth and elevation servo motors to move the antenna such that 
the target is back on boresight. The architecture of a conscan tracker is shown in Figure 3.32. 
As for the sequential lobing technique, there is only one receiver, but this is continuously fed 
with the conscan signal and the sequencing delays experienced with sequential lobing are 

Conscan does offer these performance advantages over the sequential lobe technique but 
does itself suffer from a major disadvantage. Potential foes can identify the conscan 
frequency and can radiate a signal modulated with the conscan frequency that can cause 
the radar to break lock. Conscan is therefore susceptible to electronic countermeasures. This 
deficiency may be overcome if the target is illuminated with a non-scanning beam and 
conscan is used only for the receive channel. In this way, adversaries do not know they are 
being tracked by conscan means since the tracking operation is opaque to them. This 
technique is known as conical scan on receive only (COSRO). Monopulse 

Monopulse is the preferred tracking method, and most tracking modern radars use it out of 
choice. The term monopulse means that a tracking solution may be determined on the basis 
of a single pulse rather than the beam sequence (sequential lobing) or a complete conical 
scan. The tracking data rate is therefore much higher and therefore potentially more accurate. 
Another advantage is that the tracking is based upon the simultaneous reception of the target 
return in all four channels and any variation in the echo in time can be readily accom- 
modated. This is not the case with the other techniques. 

Monopulse uses four simultaneous beams as shown in Figure 3.33, in which the beams are 
stacked in elevation and side by side. All four beams squint away from the antenna boresight 
by a small amount. Comparison of the target returns in all four channels is undertaken and 
error signals are derived to drive the antenna azimuth and elevation drive servo motors 
as appropriate. Monopulse techniques may use either phase or amplitude comparison to 





Figure 3.33 Monopulse - principle of operation. 

perform the tracking task. Of the two, amplitude comparison is generally preferred. A former 
UK AI radar (AI23b) used in the Lightning aircraft employed amplitude comparison in the 
elevation channel and phase comparison in the azimuth channel. 

All four channels transmit the same signal. The target return is received in each part of the 
monopulse array and fed into waveform junctions called hybrids which perform the sum and 
differencing function. A simplified portrayal of this arrangement is shown in Figure 3.34. 
Downstream of the summing and differencing process, three RF channels are formed: the 


Figure 3.34 Monopulse - tracking configuration. 


sum channel (A + B + C + D); the elevation difference channel (A + B) — (C + D); the 
azimuth difference channel (A + C) — (B + D). Each channel is fed into the receiver which 
has three corresponding channels. The sum channel is used to measure range, and the 
elevation and azimuth difference channels are used to drive the antenna elevation and 
azimuth servo drives respectively. 

In the early days of monopulse radar the provision of three identical receiver channels 
caused problems and some compromises were sought that multiplexed two channels. The 
radars of today do not experience this problem. 


Kayton, M. and Freid, W.R. (1997) Avionics Navigation Systems, 2nd edn, Wiley-Interscience. 
Lovell, B. (1992) Echoes of War - The Story of H 2 S Radar, Adam Hilger. 
Pilot's Handbook - Honeywell Radar RDR-4B. 

4 Advanced Radar 

4.1 Pulse Compression 

In Chapter 3 the determination of range resolution for a simple pulsed radar was shown as 
being dependent upon the pulse width, r. In fact, the expression for minimum range 
resolution is given by the following: 

C X T 

Minimum range R YQS = — - — 

where c is the speed of light (3 x 10 8 m/s) and r is the pulse width (s). 

There are practical limits as to how small the pulse width may be made. As was seen in 
Chapter 3, the theoretical receiver bandwidth required to pass all the components of a pulse 
of width r is 2/r, or 1/r if a matched filter is used. Therefore, narrow pulses need a wider 
receiver bandwidth which leads in turn to more noise and a greater risk of interference. 
However, perhaps more troublesome is the fact that, as the pulse width reduces, so peak 
power must increase to keep the average power constant. There are clearly definite physical 
limits as to how high the peak power can be. Therefore, to reduce the range resolution, a 
solution has to be sought that does not lie in the direction of ever-reducing pulse widths. 

In fact, techniques exist that permit the range resolution to be determined to a much finer 
degree, although sophisticated modulation and signal processing has to be employed. The 
technique called pulse compression (sometimes colloquially known as 'chirp') is able to both 
improve range resolution and help in extracting target echoes from noise. 

In pulse compression the RF carrier is not a fixed frequency modulated by the pulse 
envelope, rather the RF carrier is modulated with a particular characteristic. In the simplest 
form of pulse compression, the carrier is frequency modulated according to a linear law, in 
fact the carrier frequency increases in linear fashion for the duration of the transmitted pulse. 

Military Avionics Systems Ian Moir and Allan G. Seabridge 
© 2006 John Wiley & Sons, Ltd. ISBN: 0-470-01632-9 



'-Return A + B- 

Figure 4.1 Principle of pulse compression. 

In Figure 4.1 the frequency increases from/i at the pulse leading edge to/2 at the trailing 
edge. The target return contains virtually the same modulation within the target echo. In the 
radar receiver the signal is passed through a filter that has the property of speeding up the 
higher frequencies at the trailing edge of the pulse so that they catch up with the lower- 
frequency components at the leading edge. The overall effect is to compress the signal to a 
width of l/B, where B is the bandwidth of the transmitted pulse, equal to fc —f\. This 
narrow, processed signal in fact has the form sinx/x and is also increased in power by a 
factor equivalent to the pulse compression ratio. Therefore, pulse compression, as well as 
greatly improving range resolution, also greatly increases the signal power and hence target 
detection. Pulse compression ratios for a practical radar system can easily be achieved in the 
region of 100-300 and therefore the advantages can be considerable. 

It has been mentioned that the compressed pulse has the form sinx/x, the same format as 
for the radar antenna main beam and sidelobes. Therefore, pulse compression does have the 
disadvantage that it produces range sidelobes as well, and these may cause difficulties in 
some applications. A further potential problem can occur if the target echo contains a 
significant Doppler shift when range errors may be experienced. 

The lower part of Figure 4.1 shows another advantage of pulse compression. In this 
example, two target returns are overlaid, and, using conventional signal processing, it would 
not be possible to discriminate between them. However, using pulse compression and the 
accompanying signal processing, both echoes are enhanced and may be separated. 

Although linear FM pulse compression as described above is the most widely used, there 
are several other techniques that achieve a similar outcome. In recent years the viability and 
application of surface acoustic wave (SAW) technology has allowed the pulse compression 
process to be implemented in a cost-effective manner. 





Pulses are of same duration and 
frequency but different in phase 







Pulses are of same duration and 

frequency and are in phase as if having 

been 'cut' from the same continuous 


Figure 4.2 Comparison of non-coherent and coherent transmission. 

Pulse compression is used as part of the signal processing associated with synthetic 
aperture radar (SAR) and some advanced forms are used in active electronically scanned 
arrays (AESAs) where compression ratios up to 1000 may be attained using sophisticated 
signal processing methods. 

4. 1. 1 Coherent Transmission 

The description of the pulsed radar in Chapter 3 related to the transmission of non-coherent 
carrier waveforms. That is, the successive pulses of energy that were transmitted were 
unrelated in phase from one pulse to the next. Pulse Doppler techniques using coherent radar 
transmissions enable much more data to be extracted. Coherent transmission can be likened 
to a transmitter that is transmitting continuously while being switched in and out of the 
antenna (Figure 4.2). 

In Figure 4.2, non-coherent transmission is shown at the top of the diagram. The 
transmitter is controlled by the modulator, and a series of pulses of the same duration but 
differing in phase is transmitted towards the target. The reflections from the target will 
likewise be unrelated in phase. 

By contrast, the coherent transmission example shown in the lower part of the picture 
shows a stable local oscillator or 'STALO' being modulated to transmit the same series of 
pulses. The STALO runs continuously and an associated power amplifier is switched on and 
off to produce pulses of the appropriate pulse width and pulse repetition frequency (PRF). 
Therefore, within these pulses, the carrier is in phase, almost as though sections of the carrier 
have been 'cut' from the same continuous wave. The reflected energy from the target largely 
preserves this phasing in the received target echo, and this property provides very useful 
features for the radar designer. 

The properties and composition of a non-coherent and coherent pulse train for a carrier of 
frequency /o are quite different, as shown on Figure 4.3. Each pulse train comprises a series 
of pulses with pulse width r and a pulse period of l// r , equivalent to a PRF of f r . 





nl llr iin 

fr f, 

Figure 4.3 Characteristics of non-coherent and coherent pulse trains. 

The non-coherent pulse train transforms into a continuous sinx/x spectral waveform with 
bandwidth 2/r, represented as a width ±l/r centred on the carrier frequency /o- 

The coherent pulse train spectral response is bounded by the same sinx/x envelope of 
width ±l/r centred on the carrier frequency. However, in this case the energy is reflected by 
spectral lines each separated by f T , the pulse repetition frequency. In fact these spectral lines 
are minispectra rather than spectral lines since the length of the pulse has an important effect, 
as will be seen. 

Figure 4.4 illustrates the effect of varying the length of a series of coherent pulses. The 
longer the pulse train, the narrower are the resulting frequency spectra. In each case the 
frequency spectrum is represented by the sinx/x profile, but, as the length of the pulse train 
increases, this merges to form an apparent single spectral line. In the figure above, an 
infinitely long coherent pulse transforms into a single spectral line. Clearly, this is not 
practical in reality. A long coherent pulse train transforms into a narrow sinx/x response 
and the short coherent pulse train into a broader response. There is obviously a trade-off 
needed in the design to enable the most desirable performance to be achieved, bearing in 
mind practical constraints that apply. In a radar system where a target echo needs to be 
detected, the pulse train has to be of a sensible length. On the other hand, the longer the pulse 
train can be made within reason, then the narrower the frequency spectrum and therefore the 
greater the ability to reject unwanted signals. 

Figure 4.5 depicts the difference between a single pulse and a pulse train comprising eight 
pulses. The single pulse of length r transforms into a single sinx/x spectrum of width 2/r 
(ignoring sidelobes). The short pulse train transforms into a series of sinx/x spectra whose 
width is determined by the pulse width of the pulses spaced at intervals equal to the PRF,/ r . 
The individual spectra are bounded by a sinx/x envelope, this envelope being of width 2/r, 
where r is the length of the transmitted pulse train. 

It may be seen that PRF is an important parameter in the pulse train spectrum. The lower 
the PRF, / r , the closer together each of the spectra will be and after a certain point the 
sidelobes will begin to interfere with the adjacent spectra, distorting the signal content. 



| Infinite Pulse Train | 

■ +oc Time ► -oc 


Long Pulse 

Frequency - 


Short Pulse "b 




v 5 - 

Figure 4.4 Effect of pulse train length. 





Pulse Train 


iiiiiii ' i 



"fr 2/ T fr 

Figure 4.5 Comparison of a single pulse and a pulse train. 


The use of pulsed transmissions in this manner can therefore be used in airborne radars 
provided a number of criteria are satisfied. These are as follows: 

1. The radar is coherent. 

2. The PRF is high enough to spread the spectral lines sufficiently far apart. 

3. The duration of the pulse train is sufficiently short to make the spectral lines reasonably 

4. Doppler filters are devised to filter out the spectral sidelobes. 

4.1.2 Fourier Transform 

The Fourier series, or Fourier transform as it is often called, is a mathematical relationship 
that identifies all those frequency components necessary to synthesise a particular waveform. 
The use of frequency-related analytical techniques is of immense importance in manipulat- 
ing the signals that are frequency dependent, or that comprise important target-related 
frequency components. The mathematical theory is described in virtually every pure 
mathematics textbook, or for that matter most radar textbooks, and therefore will not be 
expounded in this publication. Rather, the subject will be addressed in sufficient detail such 
that readers can comprehend the importance of Fourier techniques as they apply to radar 
signal processing, particularly in Doppler radar and the associated applications. 

The technique called the fast Fourier transform (FFT) is a useful method for creating a 
series of Doppler filters to discriminate a velocity return from ground clutter. The FFT 
process lends itself readily to implementation using digital computing techniques and 
therefore is commonly used in modern digital radars for Doppler filtering, Doppler beam 
sharpening (DBS) and synthetic aperture radar (SAR) modes. The FFT technique is also used 
in electronic warfare (EW) to analyse the characteristics of the opponent's radar systems in 
real time. 

4.2 Pulsed Doppler Operation 

The operation of a pulsed Doppler radar opens totally new areas that need to be considered. 
The derivation of target velocity using the Doppler shift has significant advantages and also 
permits low-level targets to be detected and tracked when flying in ground clutter regions. 
However, the radar sidelobes also collect a significant amount of ground clutter over a range 
of Doppler frequencies, and the characteristics of this clutter need to be fully understood for 
the best performance to be achieved. 

Figure 4.6 shows the principal clutter returns for a pulse Doppler radar. These are as 

1 . Mainlobe clutter (MLC) in the direction of the main beam. This is related to the velocity 
of the aircraft/radar, as has already been described in the Doppler navigator in Chapter 3. 
The velocity component of the mainlobe clutter reduces as the antenna boresight is moved 
to the left or right of the aircraft track. This happens as the component of forward velocity 
in the Doppler shift is reduced by the cosine of the angle on the antenna boresight with 
respect to aircraft track. Similarly, the size of the MLC Doppler shift is modified 
according to the cosine of the antenna look angle (depression angle). 


1 »ll T 


Side Lobe Clutter 

Figure 4.6 Principal returns for a pulse Doppler radar. 

2. An altitude or ground return resulting from stray energy being reflected from the terrain 
directly underneath the aircraft. Since the terrain producing this return is directly below 
the aircraft, this return includes no Doppler component when the aircraft is flying straight 
and level over flat terrain. It therefore usually represents the zero Doppler shift position of 
the spectrum. However, there will be a Doppler bias if the aircraft is ascending or 
descending or the terrain below the aircraft is not level. 

3. Returns are experienced across a whole area extending from ahead of the main beam 
clutter to well behind the aircraft ground return. This return is due to energy entering the 
system via the antenna sidelobes and is therefore known as sidelobe clutter. The sidelobe 
clutter region extends over a region that approximates to ±2 x V T /X. 

The typical Doppler spectrum resulting from an aircraft flying at velocity V r , with pulsed 
Doppler antenna transmitting over terrain, is shown in Figure 4.7. On the diagram, Doppler 
frequency is increasing from left to right, starting with the extreme negative frequencies on 
the left, through the altitude line with zero Doppler shift to the extreme positive frequencies 
on the right. On the example shown there are the three main components of clutter already 

The beginning of the sidelobe occurs at a point called the opening rate, below which 
targets flying much slower than the radar may become disengaged from the most negative 
aspect of the sidelobe clutter (SLC). At this point the Doppler shift from the receding 
terrain is negative and is approximately equal to the velocity of the radar (= —2 x V r /A). 
The other extent of the SLC is called the closing rate, which is a positive Doppler shift 
(= +2 x V r /A). The altitude line approximates to zero Doppler shift, while the main lobe 
clutter (MLC) has a positive shift but less than the closing rate (depending upon the antenna 
look angle). 

The diagram shows a total of five targets from left to right (with the magnitude of the 
target echoes greatly exaggerated): 




High Opening 
Rate > - 2 x Vr 

Low Zero Closing Low Vt Normal to High 

Opening Rate [Target in Closing LOS [Target in Closing 

Rate Altitude Return] Rate MLC] Rate> + 2xVy 



Doppler Frequency 

Figure 4.7 Typical Doppler spectrum. 

1. Target 1 has a high opening rate (Doppler shift > — 2 x V r /A) and appears to the left of 
the negative SLC. 

2. Target 2 has a low opening rate and appears through the negative SLC. 

3. Target 3 has a low closing rate and appears through the positive SLC. 

4. Target 4 is flying at the same velocity or tangentially to the radar and is masked by the 

5. Target 5 is approaching at high speed with a high closing rate and appears in the clutter- 
free zone to the right of the SLC (Doppler shift > +2 x V r /A). 

It should be noted that target echoes appearing outside the SLC, altitude line and MLC 
regions will still have to be detected among the receiver noise, as for standard pulsed radar. 

It can be seen how many variables affect where clutter and the target appear with respect 
to each other: radar velocity and radar and target relative velocities, antenna position, target 
geometry, etc. As many of these variables alter, as they rapidly will in a dynamic combat 
situation, the shape and relative positions of the target echoes and clutter regions will change 
quickly with respect to one another. 

In certain situations, especially when operating at medium PRF, unwanted returns can be 
received in the sidelobes from very large ground targets. Industrial plants such as refineries 
and chemical processing facilities can produce significant returns that can be captured in this 
way, even if the sidelobe gain is 30 dB below the main beam. A solution to this problem is to 
use a guard horn and guard receiver channel, (Figure 4.8). 

The guard horn has a broad low gain response and the gain is selected such that it is 
positioned below the antenna main beam response but just above the response of the first 



Main Lobe of 

Output of Main 

Receiver is 
inhibited when 

target is 


detected in both 


Main Lobe of 
Guard Horn 









-Azimuth - 

Figure 4.8 Use of a guard channel to reduce sidelobe clutter. 

sidelobes. The receiver architecture features two parallel channels for the main antenna and 
the guard horn. Normally, the main returns from targets will be detected via the main beam 
and usual signal processing will occur. Target returns will be received in the guard channel 
but, owing to the low gain of the guard horn, will not be detected. When a target is detected 
simultaneously through the guard horn and the main antenna channels, decision logic will 
cause the main receiver output to be inhibited. Therefore, any targets that would have been 
detected from the sidelobes may be suppressed from the display. This technique may also be 
used as an electronics countermeasure tool to negate unwanted jamming entering via the 
antenna sidelobes. 

4.2. 1 Range Ambiguities 

The effect of range ambiguities in a basic pulsed radar was discussed in Chapter 3 and shown 
in Figure 3.17. However, that diagram portrayed an air-to-air engagement where no ground 
clutter was present. The situation becomes more complex in pulsed Doppler when significant 
ground clutter has to be taken into account. The problem is simply stated in Figure 4.9. 

This diagram shows a situation where a pulsed Doppler radar is looking down at three 
targets: two aircraft and a moving ground vehicle. The range return comprises three main 

• The altitude return on the left; 

• The first target, clear of the MLC; 

• The second and third targets obscured by the MLC; 

If a high PRF is being used, then the range sweep may be less than the total range shown in 
the diagram. 



Figure 4.9 Representative flight profile. 

It may be assumed that, owing to the use of the high PRF, the range is split into effectively 
three unambiguous sectors or zones. The overall effect is to superimpose the range zones on 
top of each other as far as the receiver is concerned. This situation is more complex than the 
example given in Chapter 3 as there is also the altitude return and MLC to consider. The 
overall effect is shown in Figure 4.10. 

The effect of superimposing the range zones leads to the composite return at the lower 
right; this is far more difficult to unscramble than for the simple air-to-air non-clutter case. 
The one target that was detectable outside clutter has now been totally subsumed. This 
extremely simple example shows how altering the PRF - in this case increasing it, probably 
for good reason - has had the effect of losing the target in a combination of the altitude 
return and MLC. 

4.2.2 Effect of the PRF on the Frequency Spectrum - Doppler Ambiguities 

The effect of the PRF on the frequency spectrum also needs to be considered. In an earlier 
description the frequency spectrum of a short coherent pulse train was shown in Figure 3.4 to 

-Zone 1 ► 

«* Zone 2 

Zone 1 

Zone 2 

Zone 3 

Figure 4.10 Effect of superimposing range zones. 


i!"JS envelope 


-f, ►- fr- 

L Doppler J 

Figure 4.11 Effect of a high PRF on the frequency spectrum. 

be a series of sin x/x frequency responses repeated at an interval determined by the PRF (/ 2 ). 
This was an ideal case and represented the ideal pulse response. As we have seen in Figure 
4.7, the true Doppler frequency response is in reality far more complex. Figure 4.11 shows 
the real situation for a high PRF. 

The spectrum comprises replica sets of the true Doppler frequency response repeated at 
intervals determined by the PRF and modulated by a sin x/x envelope of width 2/r, where r 
is the width of the pulse train. In this case the full Doppler passband, containing all the 
components of the Doppler spectrum, is clearly seen, as the high PRF spaces out the Doppler 
spectra so that there is no mutual interference with the sidebands (which are also replicas of 
the Doppler frequency response). 

If the PRF is reduced, then the situation portrayed in Figure 4.12 can occur. The sin x/x 
envelope is unaltered since the length of the pulse train is unchanged. However, the true 
Doppler return is now overlapping with upper and lower sidebands, giving the very 
confusing and ambiguous composite Doppler profile shown at the bottom of the figure. 

It is clear that, when considering the operation, of pulsed Doppler in a look-down mode of 
operation, the system design characteristics need to be chosen with care. In particular, the 
selection of PRF is particularly crucial. 

4.2.3 Range and Doppler Ambiguities 

The operation of a pulse Doppler radar can be affected by both range and Doppler 
ambiguities. The range ambiguity is determined by the 1/x relationship already described 
in Chapter 3 (Figure 3.17) and is a fairly straightforward relationship. The determination of 
Doppler ambiguity is more involved since it depends on the PRF, the velocity of the radar 
and the carrier frequency (wavelength) being used (Figure 4.13). The gap between two 
adjacent Doppler spectra is determined by the value of the radar PRF such that the altitude 



sin x 


rm i p 


+ 1st Lower Sideband : 

+ 2nd Lower Sideband : 

True Doppler Profile 

Composite Doppler Profile 
2fo -fo 6 +fo +2fo 

:+ 1st Upper Sideband 

+ 2nd Upper 

Figure 4.12 Effect of a low PRF on the frequency spectrum. 

Maximum Unambiguous_ 
Positive Doppler 

Figure 4.13 Definition of maximum unambiguous Doppler. 




Ambiguity 1 





0) 200 - 









\ Unambiguous 

100 - 


T 1 1 1" 


Figure 4.14 Areas prone to range and Doppler ambiguity. 

line of the second spectrum will occur at frequency / r after the altitude line for the first 
spectra (the altitude line also represents zero Doppler shift). However, the negative SLC 
extends back from the second altitude line. The maximum unambiguous positive Doppler is 
defined as/ r — 2 x V r /A, as shown in the figure. The reason can be seen from the diagram. It 
shows four closing targets with increasing closing velocity from top to bottom. As the target 
velocity increases, so the target Doppler shift increases in frequency and moves further to the 
right until finally, in the last example, the target return has been subsumed by the negative 
SLC of the previous pulse. 

The effect of ambiguous range and ambiguous Doppler is shown in Figure 4.14. In the top 
left of the diagram the effect of range ambiguity with respect to range and PRF is shown. The 
range area in which no ambiguity occurs is represented by the shaded portion shown at the 
bottom left part of the plot, adjoining the axis. Everything to the right of the curve represents 
areas where the range is ambiguous. This simple 1/x relationship is unaffected by changes in 
V T or A. 

The areas affected by Doppler ambiguity are shown in the top right picture. The 
unambiguous Doppler areas are shown as the shaded portion at the bottom right. Every 
point on the diagram to the left of the diagram represents ambiguous Doppler. 

At the bottom of Figure 4.14, both diagrams are combined to give a total picture of range 
and Doppler ambiguous and unambiguous areas. The shape of this diagram depends upon the 
carrier frequency and therefore wavelength and radar velocity. For the example shown, A is 
3 cm, which is equivalent to a carrier frequency of lOGhz, typical of a fighter radar; V r is 
assumed to be 1000 knots. Therefore, the diagram is based upon realistic figures relating to a 
supersonic engagement. 





PRF (kHz) 





Low (250 to 4kHz) - 

Medium (10 to 20kHz) 

High (200 to 300kHz) 


<= ^ = 10 cm; f= 3Ghz 

X =3 cm; f = 10Ghz 

PRF (kHz) 
Figure 4.15 Factors affecting the unambiguous Doppler zone. 

Both range and Doppler ambiguities may be resolved by changing the PRF at which 
the radar is operating, and the use of staggered or multiple PRFs is often used for this 

Increasing A increases the size of the unambiguous Doppler envelope, while decreasing A 
has the reverse effect. Therefore, an AWACS radar operating at —3 GHz/10 cm will have a 
much bigger unambiguous Doppler envelope than that shown in Figure 4.14. Decreasing 
radar velocity has the reverse effect, and therefore a fighter radar closing at only 500 knots 
will have a much smaller unambiguous Doppler envelope than the one shown in Figure 4.14. 

Figure 4.15, shows the difference between the unambiguous Doppler zones for an 
AWACS and a fighter aircraft. The choice of PRF is crucial to obtaining the optimum 
performance of the radar. Normally, the three PRF bands shown at the top of the figure are 

• Low PRF -250-4000 Hz; 

• Medium PRF -10-20 kHz; 

• High PRF -100-300 kHz. 

These figures are indicative; precise figures may vary from radar to radar, depending upon 
the design drivers and the precise performance being sought. 

The advantages and disadvantages of each of the PRF bands depend in large measure upon 
the type of radar mode being used and the nature of the target engagement. For an extensive 



review of the benefits and drawbacks of each PRF type, see Stimson (1998) and Skolnik 

4.3 Pulsed Doppler Radar Implementation 

The pulsed Doppler radar has a similar layout to the pulsed radar, but there are significant 
differences (Figure 4.16). The key differences are as follows: 

1. A computer called the radar data processor (RDP) has been added and is central to the 
operation of the radar. 

2. The exciter performs the function of stimulating the transmitter in order that coherent 
transmissions may be maintained. 

3. The synchroniser function has been integrated into the exciter and the RDP. The exciter 
provides reference signals to the receiver local oscillators and the synchronous detection 
function. The RDP performs the control functions for the radar and interfaces with the 
radar controls and other on-board avionics systems and sensors. 

4. The modulator task has been included in the transmitter. 

5. A digital signal processor has been added that provides the processing power to undertake 
the necessary filtering and data manipulation tasks. 

6. The indicator function has been moved to a multifunction display. 

The transmitter and receiver paths are similar to the basic pulsed radar shown in Figure 3.15. 
The major functional differences are: 

• The inclusion of the exciter, which performs the STALO function for the transmitter and 
provides the necessary local oscillator (LO) and reference signals for the receiver; 

Figure 4.16 Pulsed Doppler radar. 





• The use of the RDP to perform a control and synchronisation function with the other 
major units: exciter, transmitter, receiver and signal processor; 

• The replacement of a video processor with a digital signal processor. 

The key to understanding how the pulsed Doppler radar works lies in the interaction of the 
receiver, RDP and signal processor, and each are described more comprehensively below. 

4.3.1 Receiver 

The receiver block diagram is shown in Figure 4.17. The receiver takes the raw target 
reflected return received by the antenna and passed via the receiver protection device. 
Usually the signal is passed through a low-noise amplifier (LNA) which improves the signal- 
to-noise ratio before the signal enters the receiver by amplifying the target signal 
proportionately more than the noise. 

The resulting signal is mixed with the LOl reference signal, /loi, from the exciter, 
producing the first intermediate frequency, fyi , which is amplified by the first IF amplifier. 
The amplified IF1 signal is mixed with the L02 reference signal, /lo2» fr° m me exciter, 
producing the second intermediate frequency, fi$2, which enters the second IF amplifier. The 
resulting signal is fed into the synchroniser video detector which also receives reference 
signals from the exciter. 

The synchronised video detector performs the function of detecting the signal and 
resolving it into in-phase (I) and quadrature (Q) components, thereby preserving the phase 
relationship of the incoming signal (Figure 4.18). The synchronous detector compares the 
incoming IF signal with two reference signals to determine the magnitude of the I and Q 
components. The Q component is determined by comparing the incoming IF signal with the 
reference signal from the exciter. This enables the magnitude of the Q (y axis) component to 
be measured and passed into the Q channel A to D converter. Similarly, the incoming IF 
signal is compared with the exciter reference signal, phase shifted (delayed) by 90°, which 
enables the magnitude of the I (x axis) component to be measured and passed into the I 
channel A to D converter. Both A to D converters sample the Q/I channel video at an interval 
approximately equal to the pulse width. 

The magnitude of the received signal is determined by the vector sum of the Q and I 
components. The phase angle ip is determined by tan -1 (Q/I). The digitised Q and I video 
components are passed from the receiver to the signal processor. 

If the radar is required to perform a monopulse angle tracking function, at least two and 
possibly three identical receiver channels are required depending upon the tracking scheme 

4.3.2 Signal Processor 

The signal processor architecture is shown in Figure 4.19. The signal processor sorts the 
digitised Q and I information by time of arrival and hence by range. This information is 
stored in range increment locations called range bins. The signal processor is then able to 
sort out the majority of the unwanted ground clutter on the basis of the Doppler frequency 
content. The processor forms a series or bank of narrowband Doppler filters for each range 
bin; this enables the integration of the energy from successive pulses of the same Doppler 
frequency. An example of clutter rejection is shown in Figure 4.20. 

4° 8^ 


o o 2 
^ co 

- °_ o 

o g 

< o 


CO to 






O) CO 




c 0) o 

o > o 

o o ,_ 

2 co ■■= 


- ;= * 













(0 ±| 


TO x 











Reference signal 
from Synchroniser 

signal If) 







Quadrature Channel 

Q video 

Total ] 
Signal | 

! " / 


A . 


I video 




- Q 

Digitised video 
frequency signals 
- Q ^channels 


- I 

Reference signal 

shifted (delayed) by 

90 degrees 

Figure 4.18 Synchronous video detector. 

In this example, successive return signals from range R are fed into the clutter rejection 
system. Pulse n is delayed by a period T (1/PRF) and subtracted from the succeeding pulse 
(ft + 1) at the summation device. In the case of the central MLC and other unwanted clutter 
at repetitions of f T , the effect is to cancel out the clutter. However, in the case of frequency 
components around / r /2, the effect will be a summation. This has the effect of rejecting or 
cancelling clutter and reinforcing the desired signal. 

The use of fast Fourier transform (FFT) techniques using the digital computation capabilities 
of the signal processor greatly facilitates the clutter rejection and Doppler filtering functions. 

In the same way that the basic pulsed radar may track a target in range and antenna 
LOS, the Doppler radar is in addition able to track the velocity of a target using a velocity 
gate technique. This is very similar to the range gate principle except that, instead of an 
early and late gate, a pair of filters is used to track the target velocity instead, as shown in 
Figure 4.21. 

The velocity gate comprises a low-frequency and a high-frequency filter, and the exact 
tracking Doppler frequency is positioned where the voltage from both filters is equal. In 
the example given, the target has a slightly higher Doppler frequency than the velocity gate, 
with a resultant error e. Accordingly, an error voltage, AV, is generated and used to 
reposition the filters such that the error is minimised and the velocity gate represents the 
correct value. 

The signal processor identifies the level of noise and clutter and sets a threshold such that 
the genuine target echoes may be detected above the background. The resultant detected 
target positions are fed to a scan converter and stored. The scan converter is able to identify 






Returns from 
Range R 

Processing rejects 
clutter while detecting 
the doppler component I 
of moving ground targetl 



^ Pulse (n + 1) J -*\ X ) ► 

To Doppler 

^ Pulse (n) 

Delay = T 

Figure 4.20 Simple clutter cancellation scheme. 

the targets in memory and present them on a multifunction display in the appropriate 

The advantage of the scan converter is that targets are presented as if on a bright raster 
(TV) screen without the fading target echoes that would feature on an analogue display. 

4.3.3 Radar Data Processor 

The radar data processor is shown in Figure 4.22 and is the heart of the radar. The RDP 
serves the following functions: 

1. It receives commands from the radar control panel and avionics systems to exercise 
control over the radar modes and submodes, as determined by the combat situation. 

Velocity Gate 



Figure 4.21 Principle of the velocity gate. 




















2. It accepts aircraft attitude and stabilisation data in pitch roll and yaw axes to maintain the 
correct antenna position; it also receives velocity and acceleration data from the aircraft 
Inertial Navigation System (INS). 

3. The RDP provides antenna drive commands to align the antenna axis with the horizontal 
axis, vertical axis and lateral axis (usually heading) as an orthogonal axis set. The RDP 
receives antenna positional data to close the antenna stabilisation loops such that the radar 
antenna boresight is stabilised and aligned to aircraft/geodetic axes to facilitate weapon- 
aiming functions accounting for the attitude, velocity and acceleration of the aircraft. 

4. The RDP provides control outputs to the transmitter and exciter to determine the nature of 
the radar transmissions: carrier frequency, modulation, PRF, etc. It also provides data to 
the receiver and signal processor to enable the target reflections to be processed, filtered, 
detected and displayed. 

4.4 Advanced Antennas 

In Chapter 3 the characteristics of the radar antenna were described. The description of the 
antenna implemented the beam shaping, and the 'gathering' function was not explored in 
depth. In this section it is intended to move beyond a simple comparison of the parabolic 
reflector and planar array antenna to examine how technology assists in the development of 
new antennas that can offer significant benefits beyond merely increasing the antenna 
beamwidth and minimising the sidelobes. 

The following antenna types will be described: 

• Further examination of the planar array; 

• The electronically steered array (ESA); 

• The active electronically scanned array (AESA). 

For each of these antenna types the principles of operation will be described; examples of 
deployment given and an overall comparison made. 

All of these antenna types use some form of phasing within the antenna and may be 
generically termed 'phased arrays'. 

4.4. 1 Principle of the Phased Array 

The principle of the electronically steered phased array is shown in Figure 4.23. A phased 
array comprises a number of radiating elements - effectively miniature antennas - each of 
which is able to radiate independently. The relative phase of each element decides the 
direction in which the array radiates. Figure 4.23 represents a radiating array in which the 
elements are all radiating at the same frequency but each equally displaced in phase from 
the others. In detail A, each element of radiated energy will be in phase with the other 
elements once a distance A has been travelled. Therefore, at angle A off-boresight all the 
radiating elements will be in phase and a phase front or radiated waveform will be directed 
perpendicular to the phase front. In detail B the distance has changed to distance B and the 
angle to B. In this case phase front B will be formed. 

By altering the phase of the respective radiating elements within the antenna feed, the 
antenna beam may be shaped and electronically steered in any desired direction in elevation 



Phasefront A 

Phasefront B 

Figure 4.23 Principle of the phased array. 

or azimuth. This is achieved by placing phase- shifting elements in the transmit line of each 
radiating element, and electronic control of these devices can be very rapidly executed and 
the ESA can therefore achieve scan rates far exceeding that which a mechanically scanned 
array can achieve. 

4.4.2 Planar Arrays 

The planar array was briefly described in Chapter 3, and most radars produced over the past 
30 years or so have adopted this format. The planar array represents the simplest 
implementation whereby physical phasing of the individual RF distribution system shapes 
the antenna radiating pattern, analogous to the parabolic antenna previously described. The 
radiating surface of the planar array comprises several hundred radiating slots that are clearly 
visible on the front face of the antenna. These slots radiate the transmitted radar energy and 
receive the energy returning as the target echo. These slots behave as dipoles and at the 
relatively high frequencies commonly used are easier to fabricate than a dipole. Whereas 
the parabolic reflector used the physical distance between the antenna focal point and the 
parabolic reflector to shape the beam, the planar array uses a more subtle technique. This can 
be best understood by referring to Figure 4.24. A planar array is mechanically formed (via 
phasing) and mechanically steered by means of conventional antenna scanning servo 

On the left are shown the radiating slots on the front face that have already been described. 
Behind the slots and within the array itself is a carefully arranged series of feeders that each 
feed sections of the slot array. The diagram at the top right shows how these feeders connect 



Typical Slot 

Output to 

Figure 4.24 Physical characteristics of a planar array. 

into the H- shaped power dividers, each of which feeds a group of four radiating slot 
elements. The H- shaped power dividers in turn connect to feeders so that eventually the 
whole planar array slot matrix is connected to the radar summing network. 

The system described above distributes the radiated energy across the antenna face and 
gathers the incident energy of the reflected target echo. The power dividers maintain the 
necessary phasing to ensure that the beam is shaped or focused, and thus provides a narrow 
beam with low sidelobes as required by the radar. 

The planar array does not use electronic switching means to scan the beam as is the case in 
the ESA. Rather, the beam is formed and shaped using phasing techniques embedded within 
the design of the antenna. To scan targets in the outside world, the planar array must be 
mechanically scanned and therefore cannot meet the high slew rate scanning capabilities of 
the ESA. Nevertheless, the planar array is likely to be less expensive than the ESA, not 
requiring the expensive phase- shifting elements, and is still used for many airborne radar 

4.4.3 Electronically Scanned Array 

The principle of the phased array and the electronically scanned array (ESA) has already 
been outlined in the description of the phased array, and a top-level diagram showing an ESA 
radar is presented in Figure 4.25. 

The key point about the ESA is that the remainder of the radar is relatively conventional in 
terms of the transmitter and receiver. It is the individual phase shifters and radiating elements 
that provide the enhanced capability, especially in terms of scan rate. The arrangement 
shown above is called a passive ESA as there are no active or transmitting elements within 



| Conventional Transmitter & , 







o y 

% v. 

0. ^^*. 


j | 




Figure 4.25 Electronically scanned array. 

the antenna; the phase shifters are purely passive devices. This passive configuration will be 
compared with that of an active ESA in a later section. 

Some of the advantages conferred by the beam agility are listed below: 

1. Tracking can be initiated as soon as the target is detected. 

2. The antenna can scan 100° in a matter of milliseconds, whereas a mechanically scanned 
array would take 1 s or more. 

3. Targets may be illuminated even when outside the search volume. 

4. The time on target or dwell time may be optimised according to target type and 

5. Specialised detection and modulation techniques may be used to assist in extracting target 
signals from noise. 

6. Terrain-following techniques may be improved owing to the flexibility in adapting scan 
patterns and beam shaping. 

7. It permits electronic countermeasures (ECM) techniques to be employed anywhere within 
the field of regard of the antenna. 

The ESA removes many of the components that contribute to failures in a conventional 
mechanically scanned antenna. Rotating waveguide joints, gimbals, drive motors, etc., are all 
removed. Consequently, the reliability of the ESA is improved. Failures of the phase shifters 
may be easily accommodated as the antenna can stand up to perhaps 5% of these failing 
before the radar performance is adversely affected. 

The B-1B Lancer employs a passive ESA radar that uses ~1500 phase control devices. 
The antenna is movable and is capable of being locked in a detent such that the antenna 
points forwards, sideways or vertically downwards. The AN/APQ-164 employs a totally 
dual-redundant architecture - with the exception of the antenna - to improve mission 
availability. The B-2 Spirit stealth bomber also uses a passive ESA, in this case the 







Duplexerf J 










Figure 4.26 Comparison of passive and active ESA. 

4.4.4 Active Electronically Steered Array (AESA) 

The passive and active ESA top-level architectures are compared in Figure 4.26. As has 
already been seen, the passive ESA is fairly conventional, with the exception of the antenna 
which includes the many passive phase-shift elements. The active ESA, referred to as an 
active electronically steered array (AESA), is quite different. 

The AESA includes multiple individual active transmit/receive (T/R) elements within the 
antenna. Depending upon the precise implementation, there may be anywhere between 1000 
and 2000 of these individual T/R elements which, together with the RF feed, comprise the 
AESA antenna. As for the passive ESA, these elements are highly redundant and the radar 
can continue to operate with a sizeable percentage of the devices inoperative. This graceful 
redundancy feature means that the radar antenna is extremely reliable; it has been claimed 
that an AESA antenna will outlast the host aircraft. 

The fact that the transmitter elements reside in the antenna itself means there is no stand- 
alone transmitter - there is an exciter but that is all. As before, there is clearly a need for a 
receiver as well as an RDP and signal processor. The active T/R elements are controlled in 
the same way as the phase shifters on the passive ESA, either by using a beam-steering 
computer (BSC) or by embedding the beam- steering function in the RDR 

The T/R elements are very small but encompass significant functionality. The architecture 
of a typical T/R module is shown in Figure 4.27. An early generation T/R module is also 
pictured in the lop left-hand part of the diagram. The T/R module is quite small, measuring 
in the region of 0.7 x 2.0 in and ~0.2in deep, and is very reliable, having an extremely high 
mean time between failure (MTBF). Latest implementations fabricate the T/R module as 
part of a 'tile', where a tile comprises a module of perhaps four T/R devices. 



Figure 4.27 TR module architecture. 

It can be seen from the figure that the T/R module is a miniature RF subsystem in its own 
right. The beam-steering computer or RDP controls the operation of the upper and lower T/R 
switches as follows: 

1 . Transmit. During transmit, the exciter is switched via the upper switch in the T position 
and passes through the variable gain amplifier and variable phase shifter. The transmit 
signal is routed through the lower switch, also in the T position, and passed to the high- 
power amplifier (HPA). The transmitted power is passed through the duplexer to the 
radiator element. 

2. Receive. Received energy from the target is passed via the duplexer and a protection 
device to the low-noise amplifier (LNA). It is then routed through the upper switch, now 
in the R position, and through the variable gain amplifier and variable phase shifter to the 
lower switch which is also now in the R position. This routes the received signal to the 
receiver where the demodulation, detection and signal processing can be undertaken. 



The T/R module architecture described above is implemented today in a hybrid chip 
integrating a series of monolithic microwave integrated circuits (MMICs) to perform the 
RF functions. 

The ability to control many individual T/R modules by software means confers the AESA 
with immense flexibility of which only a few examples are listed below: 

1. Each radiating element may be controlled in terms of amplitude and phase, and this 
provides superior beam-shaping capabilities for advanced radar modes such as terrain- 
following, synthetic aperture radar (SAR) and inverse SAR (ISAR) modes. 

2. Multiple independently steered beams may be configured using partitioned parts of the 
multidevice array. 

3. If suitable care is taken in the design of the T/R module, independent steerable beams 
operating on different frequencies may be accommodated. 

4. The signal losses experienced by the individual T/R cell approach used in the AESA also 
bring considerable advantages in noise reduction, and this is reflected in improved radar 

Figure 4.28 shows the comparative losses for a passive ESA on the left versus an active ESA 
on the right. In each case the losses per device in dB are notified on the left with the device 



Loss Element 



-0.70 dB C J Phase Shifter 

-0.15 dB 

( j Duplexer 
I Low Powe 

-0.80 dB ^ Level 1 Feed 

-0.10 dB 

I " I Receiver 


Noise Figure: 

Fn + 0.25 dB 







Noise Figure: 

Fn + 3.05 dB 

Source: Introduction to Aircraft Radar, Second Edition 
George W Stimson 

Figure 4.28 ESA and AESA front-end loss comparison. 



Table 4.1 US fighter aircraft being fitted or retrofitted with AESA radars 

Number of TR 





F-22 Raptor 



Entering service 

F-18E/F Upgrade 



Entering service 

F-16E/F (Block 60) 



Entering service 




In service 




In development 

named on the right of the RF 'chain'. The total noise figures for each case are totalled at the 
bottom. The results are: 

Passive ESA : LNA noise figure F n + 3.05 dB 
Active ESA : LNA noise figure F n + 0.25 dB 

This is a dramatic improvement in the noise figure; it is especially significant achieving such 
an improvement so early in the RF front end. This results in a remarkable range 
improvement for the AESA radar. 

A number of US fighter aircraft are being fitted or retrofitted with AESA radars. These are 
shown in Table 4.1. 

The increase in AESA radar performance versus conventional radars has been alluded to 
earlier in the section. Figure 4.29, portrays the comparative ranges against a target where the 
cross-sectional area, cr, is normalised at 1 m 2 , the size normally ascribed to a small airborne 
target such as a cruise missile target. The aircraft to the left of the dotted line are fitted with 

Radar Ranges 

Radar detection 

ranges calculated 

against 1sq meter 

target (cruise missile) 

Range - Miles 

Target Detection Radar Ranges - Active 
Electronically Scanned Arrays (AESA) 

Figure 4.29 Comparison of fighter radar ranges. 



conventional radars, while those to the right are AESA equipped. All the AESA-equipped 
aircraft show a significant range advantage. 

Of particular interest are those data points that show a comparison of two models of the 
same aircraft: one fitted with conventional radar and the other with an AESA radar. The F-16 
Block 60 (now the F-16E/F) shows an improvement from 45 to 70 nm (+55%), while the 
F-15C range has increased from 60 to 90 nm (+50%). Apart from the obvious improvement 
in range, it has been stated by a highly authentic source that AESA radar confers 10-30 times 
more in radar operational capability compared with a conventional radar (Report of the 
Defense Science Board Task Force, 2001). 

4.5 Synthetic Aperture Radar 

One of the important requirements of a modern airborne radar is that of providing high- 
quality radar mapping such that topographical features and potential targets may be identified 
and classified. The quality of the radar map depends upon its ability to resolve very closely 
spaced objects on the ground - in the most demanding case down to within a few feet. 

The problem surrounding the need to resolve to such high resolution may be understood 
by reviewing the ground footprint of a radar beam radiated directly abeam the aircraft as 
shown in Figure 4.30. The oval footprint may be resolved into smaller cells or pixels. Along 
the radar boresight (perpendicular to the aircraft track) the range swath may be resolved into 
finer range resolution cells. Across the radar beam (along the aircraft track) the increments 
are azimuth increments as far as the radar is concerned. 

The range resolution of a pulsed radar was described in Chapter 3. It will be recalled that, 
for a pulse of width r, the range resolution increases as the pulse width becomes smaller. For 
a 1 |is pulse, the range may be resolved to 500 ft (^150 m); for a 0.1 |is pulse this would be 
reduced to 50 ft (~15 m); for a 0.01 |is pulse the resolution would reduce to 5 ft (~1.5 m) and 
so on. The major limitation to this approach to increasing range resolution relates to the 
frequency spectrum associated with a very short pulse. In Chapter 3 the receiver bandwidth 
required to pass all the frequency components is determined by the 3 dB bandwidth which in 

Along Track 

Along Track 

Figure 4.30 Principle of synthetic aperture radar. 



turn is related to the pulse width, where B = 1/r. For the 0.01 |is pulse the bandwidth is 
100 MHz. The impact of bandwidth depends upon the value of the carrier frequency; for an 
AI radar operating at 10 GHz, a 100 MHz bandwidth is not excessive, but, as frequency 
decreases, it may become more of a problem. Pulse compression techniques as described 
earlier in the chapter can achieve significant improvements in range resolution by a factor in 
the hundreds. This does not improve the bandwidth, however, as the bandwidth must still be 
sufficient to accommodate the compressed pulse. 

Azimuth resolution is more problematical. The along-track resolution is determined by the 
azimuth bandwidth using the following formula: 

#3dB=A/L (rad) 

wherev A is the wavelength and L is the antenna length. The range resolution cell is 
R x #3 dB > where R is the range at which the resolution cell is being considered. 

For a for sideways looking airborne radar (SLAR) with a 5m (16.5ft) long antenna 
operating at 3 GHz (A = 10 cm) and at a range of 50 nautical miles (90.9 km) the azimuth 
range resolution cell is ~60ft (18.2 m), still at least an order of magnitude away from the 
desired performance. 

The answer to this problem is to use a technique called synthetic aperture radar (SAR). In 
this solution the aircraft forward movement is used to create a large synthetic or artificial 
aperture, as can be seen in Figure 4.30. The aircraft shown on the left of the diagram is 
transmitting a series of pulses at constant intervals, which equates in turn to constant along- 
track increments. A series of patches of ground will be illuminated abeam the aircraft and the 
return from pulses l,2,3,4,...,n may each be detected and collected in a series of range 
bins. The principle of this technique, easily implemented in a digital computer, is shown in 
Figure 4.31. 



Range Bins 

1 1 1 






Pulse 4 
Pulse 3 
Pulse 2 
Pulse 1 

Memory scanned 

rapidly to refresh 


Figure 4.31 Collection and storage of successive pulse returns. 


After a series of pulse returns has been stored, the oldest are eventually discarded; the 
number of pulse returns retained before discarding is a function of the algorithm being 
used. This memory bank of the returns of the previous n pulses may be rapidly scanned to 
present a strip-line picture of the area of the terrain being mapped. For reasons that will 
be described shortly, this fairly crude SAR processing technique is referred to as an 
unfocused array. 

The key point about the forward movement of the aircraft is that it allows the signal 
processing to synthesise an aperture much larger than the real aperture (5 m or 16.5 ft) in the 
previous example. In a typical SAR application, an array length equivalent to 50 m or more 
may be synthesised, and in this case the azimuth resolution may be determined by the 
following approximate formula: 


SAR azimuth resolution 

2 xL 

where A is the wavelength (m), R is the range (m) and L is the synthetic array aperture (m). 

Using the figures in the former illustration, and using the synthetic array aperture figure, a 
SAR azimuth resolution of 3 ft or about 0.9 m may be achieved at 50 nm (90.9 km). This is an 
order of resolution that allows roads and, in some cases, road vehicles to be discriminated. 
Also, for range resolution - receiver bandwidth considerations apart - it was shown earlier 
that similar resolutions may be achieved by using pulse compression techniques. In this case, 
using a pulse width of 1 |is and pulse compression ratios of 100 or more, 3 ft range resolution 
may also be achieved. A combination of pulse compression to improve range resolution 
(across track) and synthetic aperture techniques to improve azimuth resolution (along track) 
means that very high resolutions in both directions may be achieved at long range. 

As an aside, working numbers to determine the minimum resolution requirements for 
ground features are: 

Roads and map details : 30-50 ft(10-15 m) 

Shapes/objects : --— - of the major dimension 

Therefore, an SAR radar providing 3 ft resolution at 50 nm would be able to distinguish 
freeways and trucks with ease, although compact cars may escape resolution. 

The point was made earlier that the array configuration described in Figures 4.30 and 4.28 
represented an unfocused array; the term 'unfocused array' will now be explained, with 
reference to Figure 4.32. The figure shows an unfocused array at the top of the diagram. If a 
linear array of length L is radiating abeam the aircraft, all the successive pulses will be 
radiated at right angles to the array and the pulse paths will effectively be in parallel. Another 
way of viewing this, considering an optical simile, is that the array will be focused at infinity. 
Therefore, point P will not be at the focus of the array; this effect is accentuated the closer P 
is to the array, or the shorter the range. These effects become more pronounced the larger the 
array and limit the effectiveness of an unfocused array. As a rule of thumb, the most effective 
azimuth resolution yielded is approximately equal to 0.4 times the array length. 

The focused array is depicted in the lower part of the diagram - this has a slightly curved 
configuration to emulate a parabolic reflector. In this situation all the successive pulses are 
effectively focused at point P and the unfocusing errors will be negated. To achieve this 
effect, two adjustments need to be made. Allowance has to be made for the fact that pulses at 




— P 

Focused at 

R + AR 

p Focused at 
Point P 

R + AR 

Figure 4.32 Unfocused versus Focused SAR Array. 

the extremity of the array have to travel a small but finite distance, AR, further to reach point 
P than pulses originating from the centre of the array. Also, pulses originating from the 
leading edge of the array have to be biased by a minute angular increment to the rear 
(clockwise), while those from the trailing edge have to be biased by a similar angular 
increment forwards (counterclockwise). If these adjustments can be made during the digital 
processing operation by taking account of the motion of the aircraft, the SAR array can be 
focused at point P. 

The computations undertaken to formulate the focused array are outlined in 
Figure 4.33. The top matrix comprises range bins that are populated during each 
successive range sweep. The contents are incremented until eventually the oldest return 
is discarded. The data in each column are read every time the array advances by the 
minimum azimuth resolution distance. Each array is focused by applying the necessary 
corrections and summed individually for each bin. The magnitude of each range bin is fed 
into the display memory. Once the last memory location has been filled, the display 
memory is decremented and the oldest data discarded. The display memory contents are 
fed to a display or recording media device. 

In effect, this process is similar to the unfocused array data manipulation, except the 
focused array is continually updated. The process of summing the columns in the range bin 
array is called azimuth compression and the effect is to synthesise a new array every time the 
radar advances by the incremental along-track distance. 

As for all these transformations, a number of trade-offs need to be made. However, after 
taking into account a range of factors, the ultimate azimuth resolution achieved by using a 
focused array is given by the following formula: 

Minimum azimuth resolution distance 

length of the real antenna 











1 Returns from 
each pulse 










Data in each 
column read 
when array 
advances by da 

Focussing & 


focussed and 
each bin 


Range Bins 
When last 

received shift 
data down one 


Display Memory 
2 3 4 5 6 

Magnitude of 
each bin 
entered into 
display memory 

Data shifted 
when last 
position filled 

Figure 4.33 Computation associated with formulating a focused array. 

Therefore, in real practical terms, even with a focused array, the minimum azimuth 
resolution distance is limited; in the example previously quoted with a 5m (16.5ft) real 
array, the azimuth resolution is limited to 2.5 m or ~8 ft. 

The key to azimuth resolution is the Doppler processing and signal integration process. 
Figure 4.34 shows the Doppler shift history as a radar approaches, passes abeam and then 
recedes from the target. 

As the aircraft flies by the target, the Doppler shift starts at +2 x V r /A, reduces through 
zero as the aircraft passes abeam and then decreases to —2 x V r /A. As the target passes 
abeam, the rate of change in frequency is maximum, and immediately adjacent to zero 
frequency the lines are virtually straight. Also, for a number of evenly spaced points 
positioned near to each other the frequency difference between them, A/, will be constant. 
Therefore, each of the returns will have a different frequency, determined by its azimuth 
position, and by using Doppler filtering techniques that azimuth position may be measured. 
The way in which this processing is achieved is shown in Figure 4.35. 

The incoming video returns are modified by applying the necessary phase corrections as 
described above to focus the array. Then each point has a constant Doppler frequency which 
can be discriminated from the others and which relates to its azimuth position. Each time the 
radar travels a distance equal to the array length to be synthesised, the phase-corrected 
signals that have been gathered in the range bin row are fed to the array of Doppler filters in 
the columns. The integration time for the filters is the same time it takes the radar to transit 
the array length, and the number of Doppler filters in each column thus depends upon the 
length of the array. The greater the array length, the longer is the integration time and the 



+ 2VR 

Point abeam of 


oc distance 
azimuth points 

Five points 

Figure 4.34 Doppler history of radar passing a target. 

greater the number of Doppler filters formed for a given frequency coverage. Therefore, as 
the array length increases, the filters become narrower and azimuth resolution becomes finer. 
As in this region abeam the target the points on the ground are evenly spaced, and the 
frequencies are evenly spaced, a fast Fourier transform (FFT) may be used to form the filters 
which minimises the amount of computing required. 

-► Focussing - 

1 IF V | <F V V U 

Range Range Range Range Range Range Range 
Bin Bin Bin Bin Bin Bin Bin 




















Filters """ 




















Outputs of 

all filters 

passed to 





Figure 4.35 Example of Doppler processing. 


Figure 4.36 SAR picture of tanks with 1 ft resolution (Sandia National Laboratories) 

The use of synthetic aperture imaging is an important asset in today's airborne platforms. 
Figure 4.36 is an example of an SAR picture that shows two rows of tanks with a resolution 
equivalent to 1 ft. The picture shows one of the curious features of SAR images: each tank 
has a large shadow in the six o'clock area, showing that the targets were illuminated from the 
twelve o'clock position. Therefore, although sophisticated data processing enables the target 
to be viewed in high resolution at long range, it cannot negate the fact that the target was 
illuminated at a low grazing angle. 

The SAR principles have been described for a fixed antenna scanning terrain abeam of the 
aircraft. There are other modes that are commonly used: 

1. Spotlight mode. In the spotlight mode the look angle of the real antenna is altered so that it 
always illuminates the target. This has a number of advantages. Since the real antenna is 
always trained on the target, the length of the synthetic array is not limited by the 
beamwidth of the real antenna. Also, the fact that the target is viewed from different 
aspects helps to reduce the graininess of the response. 

2. Doppler beam sharpening. Doppler beam sharpening (DBS) is a subset of SAR operation, 
lacking several signal processing refinements, and as a result it lacks the performance of 
the more sophisticated SAR mode. Nevertheless it provides a significant improvement in 
performance over the real-beam mapping that is the baseline mapping mode in many 

3. Inverse SAR. Inverse SAR (ISAR) is a variant of SAR that is used against moving targets 
that have a rotational component; normally, SAR is used against fixed ground targets. 
ISAR effectively uses the minor Doppler shifts caused as a result of target movement. 
ISAR may be used against aircraft or moving ground targets such as ships. 

The difference between the main modes of ground mapping, including real beam, DBS, SAR 
and ISAR, are difficult to envisage without a direct comparison. For an air-to-ground radar 
with the basic parameters listed in Table 4.2, a top-level comparison is given in Table 4.3. 

4.6 Low Observability 

The importance of a radar system in detecting, identifying and engaging a target is undeniable. 
However, for years the aircraft designer has been attempting to minimise the ability of a radar 



Table 4.2 Baseline air-to-ground mapping radar specification 





A/D conversion 

Signal Processor 

Radar data processor 




10 GHz 

Flate plate; 3° x 3° beamwidth; 30 dB gain; -30 dB sidelobes 
Travelling wave tube (TWT); lOkW peak power; 5% duty cycle 
5 dB noise figure; dual conversion; RF preamplifier; STC; AGC 
I/Q 8 bits; 120 MHz maximum rate 
100 MFLOPS/s; 8 Mbytes RAM 
2.5 MIPS; 64kBytesRAM 
1553 data bus 

512 x 512 x 8 bits; monochrome 

Up to 250:1 pulse compression; variable CHIRP rate; 0.01 (is 
minimum to lOjas maximum 

to detect the aircraft. This art is known as low observability or, more colloquially, stealth, and 
has been utilised for many years, although perhaps with more prominence with the 
introduction into service of the US Air Force stealth aircraft: F-117 Nighthawk, B-2 Spirit 
and F-22 Raptor (Figure 4.37). 

Table 4.3 Comparison of ground-mapping modes 










Selectable 60° 







±60° (max.) 


2.5 nm x 2.5 nm 


Scan rate 

60 deg/s 

Varies with 
azimuth 60-5 


Spotlight mode 

Spotlight mode 


Real beam 

20:1 beam 

25 ft cross-range 

Target motion 




512 azimuth bins 

512 azimuth bins 

128 doppler 
bins, 0.25-2.0 
Hz resolution 



Selectable 5- 

Selectable 5- 




40 nm 

centre is 

centre is 


R m J256 :950 ft 

AR = R x Aa z x 

25 ft, 

5 ft, 512 range 

at 40 nm 

325 ft for 
30 nm 

512 range bins 



PRF = 2 kHz, 

Variable PRF: 

Variable PRF: 

PRF = 800 Hz 


2.5 kHz to 

1.5-3 kHz, 

at lOOnm 



frequency agility 500 Hz 


32-point FFT 

PRF = 1700 Hz 
at 30°, 25 nm, 
300 knots 
5 1 2-point FFT 1 28-point CFT 



El 17 A Nighthawk 
(US Air Force photo 
by Staff Sgb 
D Allmen 11) 

B-2 Spirit (US Air Force 
photo by Master Sergeant 
Val Gempis) 

F.22 (US Air Force photo) 

Figure 4.37 US Air Force stealth aircraft. 

The history of the use of stealth techniques stretches back before these aircraft were 
developed (Figure 4.38). This shows that the development of stealthy techniques stretches 
back to the mid-1950s. The evolution of stealthy aircraft can be characterised by three 
distinct generations: 

1. First generation. The original strategic reconnaissance aircraft developed by the Lock- 
heed 'Skunk Works' - the U-2 and the SR-71 Blackbird. 

Stealth Generations & Milestones 

Discovery of Full Capability Gulf 
Faceting RCS Prediction War 

Kosovo war 

▼ T 

U-2 SR-71 

Spyplane Blackbird 




US Navy 









1950 1960 1970 

Source : AIAA & AW&ST 

Figure 4.38 History of the development of stealth aircraft. 






2. Second generation. In the early 1970s the concept of 'faceting' was invoked, first on the 
two Have Blue demonstrator aircraft and then applied in production on the F-117. This 
technique involved the use of angular faceted areas designed to redirect or deflect the 
radar energy away from the emitter. 

3. Third generation. This generation was enabled when computational techniques had 
developed such that the entire aircraft could be predicted and evaluated using numerical 
means. These aircraft are soft blended shapes quite unlike the faceted approach. The Tacit 
Blue demonstrator proved these techniques which were subsequently adopted on the US 
Air Force advanced tactical fighter (ATF) fly-off aircraft: the Lockheed Martin YF-22A 
and the Northrop YF-23A. Subsequently, these techniques have been applied to the 
Northrop B-2 Spirit stealth bomber which is in service and the Lockheed Martin F-22A 
Raptor which is just entering service. The F-35 [previously the joint strike fighter (JSF)] 
also applies similar techniques, looking very similar in appearance to the F-22. 

Figure 4.38 also shows the milestones of the campaigns in which the stealth aircraft have 
been successfully deployed: the F-117 in the first and second Gulf Wars and Kosovo and the 
B-2 during Kosovo and the second Gulf War. Despite flying thousands of missions, the only 
casualty was an F-117 shot down near Belgrade during the Kosovo campaign. 

Recalling the basic radar range equation discussed in Chapter 3, the radar range is 
proportional to a 1 / 4 where a is the radar cross-section (m 2 ). Reducing the cross-sectional 
area therefore affects radar range, although only according to the fourth root. However, by 
carefully designing an aircraft, the value of a may be reduced by many decades (and the 
reflected signal by many dB), and therefore reduction in the radar cross-section is an 
attractive and effective approach to reducing the detection range of an aircraft by radar. 

4.6.1 Factors Affecting the Radar Cross-section 

The radar cross-section may be thought of as comprising three elements: 

• Geometric cross-section; 

• Directivity; 

• Reflectivity. 

There is a limit to what can be done regarding the first component since the size of an aircraft 
will be dictated by the role, weapons pay load and range. However, a combination of low- 
directivity and low-reflectivity techniques has been successfully developed, as the degree of 
stealth offered by some of the aircraft testifies. 

The major areas on conventional fighter aircraft that contribute to a high radar cross- 
section are shown in Figure 4.39. The radar, cockpit, engine intakes, drop-tanks, engine 
exhaust and rudder/elevator combination can all produce large returns, primarily because 
they can act as radar reflectors. External carried ordnance and the wing leading edge can 
produce large reflections on occasions. The aircraft fuselage and small blended inlets like a 
gun muzzle produce relatively low radar reflections, as do minor air inlets or blended inlets 
on top of the aircraft. 

For a stealthy shape to be achieved, these factors need to be taken into account at the 
design stage. The design features needed to ensure a low-observability aircraft are shown in 
Figure 4.40. These include: 



Air Inlet 

Intake Cavity/ 
Engine (large^ 


Drop Tank / 

(large if straight) 

Wing -when 

Reheat (large) 


(seeker may 
be large) 

Figure 4.39 Areas contributing to a high radar cross-section. 

1. Blended wings and fuselage, and the use of blended chines at the side of the fuselage. 
These features were originally incorporated on the SR-71 and also utilised on later 
aircraft, particularly on the Northrop YF-23A and B-2, minimising all necessary bumps 
and protuberances including external antenna. 

2. Swept leading edges, not necessarily linear, to reduce 'end-fire' array effects or to ensure 
that constant angles are maintained. 

3. A low-profile 'blended cockpit' to avoid the cockpit and pilot acting as a corner reflector. 

Leading v 

Low Profile Edges \ 
Cockpit \ \ 

_ \ ^ 

Radome \ \ 

^ / 

/ Inlets 

^^ Nozzles 

Minimise / / / 
Breaks & / / / 
Corners / 1 

Blend / / 
Components / 


\ Canted 

Canted / 
Fuselage ' 

Carry Ordnance inside Aircraft 
Eliminate Bumps & Protrusions 

- Performance 

Trade Off 
Figure 4.40 Low-observability design objectives. 



4. Ordnance. On the most stealthy aircraft the weapons carriage is internal. 

5. Bandpass radome and other radar RCS reduction techniques which will be described later. 

6. Canted rudders such as those used on the YF-22A, YF-23A, F-22A and F-35A/B/C assist 
in not providing a corner reflector effect. The Northrop YF-23 A was an extreme example 
of this feature, and its stablemate, the B-2, being a flying wing, has no rudders at all. 

7. Shielded engine nozzles. Note the B-2 overfuselage engine intakes and exhausts. 

Unfortunately, many if not all of these features mitigate against aircraft performance, so in 
reality the RCS and performance have to be the subject of trade-offs. Many of the stealthy 
aircraft are unstable and require high-integrity sophisticated flight control or fly-by-wire 
(FBW) systems. In most other respects the aircraft systems are fairly conventional and have 
often borrowed and adapted major subsystems from non-stealthy combat aircraft. 

The benefits conferred by the design principles outlined above are augmented by the use 
of radar-absorbent material (RAM) to improve the low- observability features. An example of 
the combination of both techniques may be gained by examining the intake design on the 
F-117 and F-35 fighters. 

The F-117 intake uses a combination of single reflections, multiple reflections and RAM to 
lower the engine intake RCS (see Figure 4.41 which shows the intake grill and a diagram 

Reflect . 

Absorb > 
Penetrate ■ 

Radar Energy 

Radar-Absorbing Material 





Energy absorbed 

by multiple 


Hard Reflect 

Soft Absorb 

Grid Element 

Figure 4.41 F-117 intake grill & RCS reduction. 



portraying the intake construction). The intake grill is actually coated with RAM and is 
triangular shaped. Energy incident upon the intake grill may be reflected from one of the 
triangular grid elements, being dissipated and reflected away from the threat radar. This is 
indeed how faceting is used across the entire external surface of the F-117 with considerable 
success. Some incident energy will impinge directly upon the grid element and will be 
absorbed. A proportion of the energy will pass through the grill and enter the intake. After 
being reflected within the intake, it will undergo multiple bounces against the RAM-coated rear 
of the grid element and will mostly be dissipated within the engine intake. Such minor amounts 
of energy that do escape the intake grill will be of very low power and randomly scattered. 

The F-35 intake is reported to be more sophisticated, being a serpentine duct rather than a 
direct, more conventional intake. Although at first glance the relatively open intake would 
appear to suffer from a high RCS, sophisticated techniques are used that lower the cross- 
section over a range of frequencies. The intake is designed to counter radar threats at three 
wavelengths loosely termed long (~30cm), medium (~ 10-20 cm) and short (~3cm), 
equating to 1 GHz (long-range surveillance radar), 1.5-3 GHz (AWACS radar) and 10 GHz 
(fighter radar) respectively. 

At long wavelengths the duct behaves as indicated in Figure 4.42. The wavelength is too 
large to propagate effectively down the inlet and a minimal amount of energy reaches the 
engine/blocker. Most of the incident energy is attenuated by the RAM coating around the inlet 
lip and the remainder reflected away from the threat radar. The RAM coating around the inlet 
lip is optimised to have a maximum effect at ~1 GHz. 

For medium wavelengths the wave is able to propagate effectively down the inlet and 
proceeds down the serpentine duct unimpeded. As it reaches the engine, it impinges upon the 
RAM-coated blocker which is tuned to absorb the maximum amount of energy at this frequency 
and to attenuate subsequent reflections or bounces. Most of the energy is dissipated at this 
point; the very small amount of residual energy that does remain is reflected out of the duct 
(Figure 4.43). 

Radar Energy 

Radar-Absorbing Material 

Long Wavelength 



Wave is too big to enter inlet and reflects diffusely. 
Inlet is angled to reflect away from tactical targets 

RAM coating of inlet lip is tuned to absorb long wavelength 
and reduce reflection 

Very little energy reaches blocker 

Figure 4.42 F-35 engine inlet duct response to long wavelengths. 



Radar Energy 

Radar-Absorbing Material 

Medium Wavelength 
~ 10 -20cm 



Wave is proper size for duct to act as a waveguide and 
carry energy unimpeded 

Wave is too big to enter spacing of blocker vanes. 

RAM coating is tuned to this frequency and absorbs most 

of the energy 

Small amount of energy is reflected back out of the duct 

Figure 4.43 F-35 engine inlet duct response to medium wavelengths. 

Short-wavelength energy incident upon the duct travels readily down the duct, reflect- 
ing off the RAM coating on the inside of the duct walls. This RAM is tuned to achieve 
maximum attenuation at 10 GHz and therefore heavily attenuates the energy. The energy that 
does reach the engine becomes trapped and is mostly dissipated in the blocker/engine 
labyrinth. A small portion of the energy is reflected out of the front of the duct, as shown in 
Figure 4.44. 

Radar Energy 

Radar-Absorbing Material 

Medium Wavelength 

~ 3cm 


RAM Lining 

Wave bounces of walls of duct and is absorbed by RAM 
lining tuned to high frequency 

Small amount of energy enters blocker which acts as a 

Wave exits blocker and hits fan where energy is lost 
in fan/blocker labyrinth 

Energy escaping the blocker/fan loses more energy as it 
is reflected back out of the duct 




Figure 4.44 F-35 engine inlet duct response to short wavelengths. 



From the foregoing it can be seen that the shape and size of the inlet, the nature of the 
serpentine duct, the use of the blocker fan and the judicious use of specialised RAM coating 
material greatly reduce the intake RCS, thereby enabling an apparently normal engine inlet 
to have a low RCS response to different types of radar. 

The maintenance and preservation of the radar- absorbing coatings on a stealth aircraft 
pose a major servicing penalty. Every time an aircraft panel is removed for maintenance it 
has to be resealed - effectively recaulked - in order to preserve the low-observability 
signature of the aircraft. This takes time enough, but in many cases the sealant used has a 
long curing time so it can take some time before the aircraft may be used on a mission. Also, 
certainly during the F-117 production run and possibly with the B-2 bomber, different 
aircraft were coated with different coatings or sealants as production proceeded and 
improved treatments became available. This created a configuration control issue with 
several different variants of coating standard across the fleet. It is understood that modifica- 
tion programmes are in place to ameliorate this problem. 

The stealthy aircraft creates a problem for the radar and radio designers as a vast 
proliferation of antennas have to be fitted on to the aircraft in order for it to fulfil its 
mission, each with its own particular system performance goals. In Chapter 2, the wide array 
of antennas performing the primary radar, CNI, and EW functions on a JIAWG/F-22 RF 
architecture was outlined, and the trend towards shared apertures for JAST/F-35 applications 
was described. The radar, being at the front of the aircraft, faces particular problems that 
need to be addressed to maintain the overall RCS signature of the aircraft. 

4.6.2 Reducing the RCS 

Some of the particular problems that the radar confronts in minimising the RCS are as 

1. Antenna mode reflections. The antenna mode reflections mimic the antenna main beam 
and sidelobes. Therefore, merely positioning the antenna such that the main beam is not 
pointing towards the threat radar is insufficient. 

2. Random scattering. This is caused if the antenna characteristics are not uniform across the 
antenna. The solution is to maintain close production tolerances and hence uniformity 
across the array. 

3. Radar antenna edge diffraction. Mismatches of impedances at the perimeter of the 
antenna can cause reflections called edge diffraction. In effect the outer perimeter of the 
antenna acts as a loop and reflections tend to be abeam of the antenna rather than fore and 

4. Coupling of structural modes into the antenna. In this case the antenna can effectively 
'mirror' the incident energy from the threat radar, thereby compromising low RCS. In an 
ESA, active or passive, this can be readily overcome by rotating the array such that it 
points slightly down and the reflected energy is directed away from the threat radar. This 
technique, although effective, has the disadvantage of reducing antenna effective area by 
the cosine of the depression angle. 

The subtle and different effects of the antenna reflection effects of antenna mode, random 
scattering and edge diffraction are shown in Figure 4.45. Antenna mode reflection is 
minimised by ensuring correct RF matching of the feed elements to avoid reflections. 



Antenna k 

Mode I 



1 1 1 



Random i 


Energy from 
Threat Radar 



Improved RF 




4 x k at the 
Threat Frequency 

Figure 4.45 Factors affecting the antenna RCS. 

Random scattering is minimised by maintaining close production manufacturing tolerances 
to maintain uniformity across the array. Edge diffraction effects may be reduced by placing 
RAM material around the perimeter of the array to minimise the loop reflection effect. 
This should be a width identically equal to 4 x A at the threat radar operating frequency, and 
so a significant penalty may apply as the antenna size and therefore gain is reduced 

4.6.3 Comparative RCS Values 

A crude comparison of the RCS value for a range of objects is shown in Figure 4.46. The 
comparison is shown with two scales: the upper scale depicts the RCS in square metres while 
the lower scale depicts dB/m 2 , that is, dB power reduction of a return related to a reference 
square meter of target. The 1 m 2 target is the reference point for a small airborne target such 
as a cruise missile. 

The stealth aircraft B-2, F-22, F-117 and F-35 appear on the left-hand side of the diagram 
in the —30 to — 40 dB region. Conventional aircraft range from the F-18E/F at OdB to 
bombers and transport aircraft at +30 dB. Ships are a huge target at 10 4 m 2 or more. Normal 
everyday objects range from insects at — 30 dB, through birds to humans at —10 dB. On this 
diagram it could be deduced that an F-22 has an RCS of 1 x 10~ 7 (or one ten-millionth) of 
that of a B-52 bomber. 

Although these comparisons serve some purpose in indicating some of the relative 
magnitudes involved, they do not present the whole picture. In fact, no aircraft may be 
made invisible, and stealth aircraft may be detectable at lower than conventional radar 
frequencies. They may, however, be very difficult to track or engage even if detected. 




Ke Y : A 2nd/3rd Generation 

A Stealth Aircraft 

| Conventional Aircraft 

O Conventional Objects 





















































Figure 4.46 Comparative RCS values. 

4.6.4 Low Probability of Intercept Operation 

While the reduction of the antenna RCS is an important consideration, thought also has to be 
given to the energy radiated when the radar is operating. An important feature is low 
probability of intercept (LPI) which minimises the possibility that the radar is detectable by a 
foe using a radar warning receiver (RWR). The easiest way of avoiding detection is to use 
some type of emission control (EMCON) restricting the radar emissions. However, if this 
strategy were to be carried to extremes, then there would be little point in fitting the radar. 
Since the radar needs to be used to detect, identify and engage targets, the question of LPI 
has to be seriously addressed. 

There are a number of ways by which probability of intercept may be minimised: 

1. Reduce peak power by increasing integration time. The RWR is normally a broad-beam 
device, usually with four sets of antennas giving 360° panoramic cover in azimuth (a 
typical RWR antenna with 90° angle cone coverage has a gain of only ~7 dB). The RWR 
also has to receive a number of successive pulses to determine parameters such as 
frequency, pulse width, PRF, etc., and be able to quantify the characteristics of the radar 
and identify the likely platform. RWR operators have the advantage that the radar energy 
only has to travel one way to them. However, they are using an antenna with low gain and 
have to have a relatively long sample time to make intelligent deductions about the pulses 
of energy they are receiving. The emitting radar is not so constrained. 

2. Balance bandwidth against lower peak power. To be able to discriminate between signals 
in frequency, the RWR has to have a relatively narrow bandwidth. For the emitting radar, 
bandwidth is a function of radar design, and therefore, by using a broad-bandwidth 
receiver and spread spectrum modulation techniques, detection may be avoided. 

3. Maximise antenna gain. Although this aids the RWR during transmission, it helps the 
radar during reception. Moreover, the antenna sidelobes should be reduced as far as 


possible - for a LPI radar design the sidelobes should be reduced to 55 dB below the main 

4. Minimise system losses and noise. For the reception of the target echo, range will be 
maximised when front-end losses are kept to a minimum and the receiver noise figure is 
reduced. The advantage of using an AESA over a passive ESA design was demonstrated 
in Figure 4.25, where the use of the individual TVR modules provided a 3dB noise 

5. Use different forms of modulation. The modulation types described in this book are 
fairly well understood and widely used. However, there a plethora of more advanced 
modulation techniques that may be used either to extract low-level target signals from 
noise or use transmission patterns that a RWR may not be able to detect, let alone 

It has to be assumed that the modern AESA radars being deployed on a range of US fighters 
use these and many other techniques to maximise range while avoiding detection. It has been 
stated publicly that the F-22 AN/APG-77 radar can operate with varying modulation and 
frequency on a pulse-by-pulse basis. It can also operate using multiple beams in multiple 
radar modes using the concept shown in Chapter 3 (Figure 3.27). It is claimed that the AN/ 
APG-77 is able to detect, identify and engage a target without the foe ever being aware that 
an enemy is present. For further information on stealth RF design, see Lynch (2004). 


Lynch Jr, D. (2004) Introduction to RF Stealth, SciTech Publishing Inc. 

Report of the Defense Science Board Task Force (April 2001) Future DoD airborne high-frequency radar 


Skolnik, M.E. (2001) Introduction to Radar, 3rd edn, SciTech Publishing Inc. 

Stimson, G.W. (1998) Introduction to Radar Systems, 2nd edn, SciTech Publishing Inc. 

5 Electrooptics 

5.1 Introduction 

The use of electrooptic sensors in military avionics systems has steadily evolved over the 
past three decades. Infrared (IR) missiles were originally produced during the 1950s with 
missiles such as Sidewinder in the United States and Firestreak and Red Top in the United 
Kingdom. Television (TV) guidance was used on guided missiles such as TV-Martel 
developed jointly by the United Kingdom and France during the 1960s and AGM-62 
Walleye in the United States. Lasers were used for target illumination during the latter stages 
of the Vietnam War. Forward looking IR (FLIR) imaging systems were developed and 
deployed during the 1970s, and third-generation systems are now taking the field. Infrared 
track and scan (IRTS) systems followed. Now, integrated systems are in operation that 
combine a number of sensor types to offer a complete suite of capabilities. 

This chapter describes the following electrooptic technologies that are to be found on a 
range of modern military, law and order and drug enforcement platforms: 

• Television (TV) - day, low-light and colour (section 5.2); 

• Night- vision goggles (NVG) (section 5.3); 

• IR imaging including forward looking infrared (FLIR) (section 5.4); 

• IR tracking systems including IR-guided missiles and infrared track and scan (IRTS) 
(section 5.5); 

• Lasers - target illumination, range-finding and smart bomb guidance (section 5.6); 

• Integrated systems (section 5.7) (usually carried in external pods or multiaxis swivelling 
turrets; on stealth aircraft, carried internally to preserve aircraft low-observability 

The characteristics and range of all the electromagnetic sensors used in modern military 
avionics systems was described in Chapter 2 (Figure 2.2), spanning 10 decades of the 
electromagnetic spectrum. Electrooptic sensors and systems cover the top two decades from 
10 4 GHz to 10 6 GHz in frequency (Figure 5.1). In fact the categorisation of devices in Hz at 

Military Avionics Systems Ian Moir and Allan G. Seabridge 
© 2006 John Wiley & Sons, Ltd. ISBN: 0-470-01632-9 









100nm > 

193 _ 

"lOOOnm > 


10000nm > 


_ — J~Violet 



-750— ■ 

„ 1064 
- YAG 



Green ^ Visible Light 

•'-i Red 

— r 1000 N 


— ! -2500- 




1 micron 


2.5 micron 
3 micron 



i micron 


14 micron 

— J-14000 — ' 
1000 nanometer = 1 micrometer = 1 micron 


Figure 5.1 Electromagnetic Spectrum for electrooptic systems. 

this point in the spectrum becomes meaningless because of the huge numbers involved, so 
that wavelength is more usually used. Therefore, wavelength is referred to in microns, where 
1 jim = 1 x 10 -6 m or one-millionth of a metre. In the visible light portion of the spectrum 
that the human eye uses, angstrom units are sometimes used, especially in the scientific 
community, where 1 A = 10 -2 m or 10 4 urn. 

The three specific bands of interest in Figure 5.1 are: 

1. The visible light spectrum from 750 to 400 nanometres covering red/orange/yellow/green/ 

2. The IR spectrum from 1 to 14 urn which itself is subdivided into three regions: 

• Shortwave IR: 1-2.5 urn; 

• Medium- wave IR: 3-5 urn; 

• Long-wave IR: 8-14 urn. 



3. The spectrum within which airborne lasers operate from 193 nanometres near the 
ultraviolet part of the spectrum to 10 600 nanometres approaching the far IR spec- 
trum. Of particular interest is the frequency at which yttrium aluminium garnet 
(YAG) operates, namely 1064 nanometres, this being one of the most common lasers 

As the area of the spectrum in which electrooptic sensors operate is close to and includes 
visible light, so they experience many of the same shortcomings to a greater or lesser extent: 
obscuration due to water vapour or other gases, scattering due to haze and smoke, etc. The 
near-IR part of the spectrum such as the shortwave IR region suffers most, although sensors 
operating in the long-wave IR band are less affected. 

Therefore, a major problem confronting the use of sensors in the IR region in particular is 
the severe attenuation that occurs in certain parts of the spectrum, allowing only certain 
windows to be used. This is similar to the radar attenuation described in Figure 3.4 of 
Chapter 3. The IR transmission characteristics in the atmosphere are presented in Figure 5.2. 
This shows the percentage transmission over the IR band from 1 to 14 urn at sea level. As 
shown in the figure, at sea level there are a number of areas where the attenuation is 
significant, particularly in a region between 6 and about 7.6 urn where there is no 
transmission at all - mainly owing to water vapour and C0 2 . At altitude the situation is 
much improved, although there are one or two unfavourable attenuation 'notches', again 
owing to H 2 or C0 2 . Therefore, atmospheric attenuation is most likely to affect systems 
being used at low level, while air-to-air missiles being used at medium or high altitude will 
be less affected. 

In fact, IR systems tend to use the IR windows or bands already described to avoid the 
worst of the problem using the SWIR (1-2.5 urn), MWIR (3-5 urn) and LWIR (8-14 urn) 
regions, and all the sensors utilised in avionics systems operate in these bands. 


k Short Wave Mid Wave 

10 (1 to 2.5 micron) (3 to 5 micron) 

Long Wave 

(8 to 12 micron) 

1 2 3 4 5 6 7 8 9 10 11 12 13 14 

Figure 5.2 Transmission characteristics of infrared (sea level). 




5.2 Television 

The visible light sensors used in electrooptics are available to assist and augment the 
platform operator's 'Mk 1 eyeball'. The main categories used are: 

• Direct vision optics, effectively a direct optics system offering image magnification in the 
same wave as a conventional telescope or binoculars; 

• Television (TV) imaging in monochrome and, more recently, colour; 

• Low-light TV (LLTV) for viewing in low-light conditions. 

TV camera sensors used in military applications existed for many years prior to the 1980s 
when the sensing technology was imaging tubes. Charge coupled devices (CCDs) have now 
largely replaced the imaging tube, and these are described below. 

The CCD imaging device comprises a number of resolution cells or pixels that are 
sensitive to incident light. Each pixel comprises elements of polysilicon placed upon a layer 
of silicon dioxide. Below the silicon dioxide is a silicon substrate layer. As the incident light 
arrives at the polysilicon, a 'well' of charge is built up according to the level of the incident 
light. The pixels are arranged in X rows and Y columns to provide the two-dimensional CCD 
array as shown in Figure 5.3. In each pixel the charge according to the amount of incident 
light is captured by a potential barrier that surrounds the well and maintains the charge in 
place. The pixels are separated by an interelement gap and respond to light below a 
wavelength of HOOnm (visible light lies between 400 and 750 nm). Therefore, visible 
spectrum energy incident upon the array will provide a pattern of charge within each pixel of 
the CCD array that corresponds to the image in view. The CCD array is an optical plane that 

Incident Light 
-V +V -V 

J i L 

< — Polysilicon 

< Silicon Dioxide 

Blocking N ^^^g^^^ / Blocking 

Potential Potential Potential 


- Silicon Substrate 

Detector ►! 

-X Array- 




Y Array 

Figure 5.3 Principle of operation of the charge coupled device. 




of Parallel 


i // 









Serial Register 




] Clocks 


Figure 5.4 Full frame CCD. 

preserves the image in the same way as a frame of film, and developments of this technology 
are now to be found in commercial digital cameras - replacing wet-film technology. 

Once the array of retained charge has been established, the image needs to be scanned. 
There are a number of ways of accomplishing this, two of which are described below: 

• Full frame devices; 

• Frame transfer devices. 

A full frame CCD device is shown in Figure 5.4. The charge associated with a row of pixels 
is shifted sequentially to the serial register, whereupon the row is read out as a stream of data. 
This is repeated on a row-by-row basis until the complete array of pixels has been read off 
the imaging device or chip. As the parallel register is used for both imaging and read-out, a 
mechanical shutter or illuminated strobe must be used to preserve the image. The simplicity 
of this method means that operation is simple and provides images of the highest resolution 
and density. 

The frame transfer architecture is shown in Figure 5.5 and is similar to that of the full 
frame transfer except that a separate identical parallel register, called a storage array, is 
provided that is not light sensitive. The captured image is quickly read into the storage array 
from whence the image may be read out as before while the next image is being formed. The 
advantage of this approach is that a shutterless or strobeless imaging process is possible, 
resulting in higher frame rates. Performance is sacrificed as imaging is still occurring while 
the image is being read to the storage array, resulting in smearing of the image. Twice the 
area of silicon is required to fabricate an imaging device of comparable imaging coverage, so 
the frame transfer array provides lower performance at a higher cost. 

The precise optical configuration of a CCD depends entirely upon the intended operational 
use. CCD devices may be used in a tactical platform such as a battlefield helicopter or as a 




of Parallel 


Image Array 

Storage Array 

Serial Register 

Image Array 

Storeage Array 



Figure 5.5 Frame transfer CCD. 

high-resolution imaging system for a theatre reconnaissance aircraft or UAV. The imaging 
requirements and field of view (FOV) will be different in each case. Looking at the 
capabilities of two CCD imaging systems illustrates this fact. The examples chosen are 
the systems on board the AH-64 C/D Longbow Apache and RQ-4 Global Hawk: 

1. AH-64 C/D Apache. The target acquisition designation sight (TADS) or AN/ASQ-170 is 
part of the overall turreted EO sensor system on the Apache battlefield helicopter. The TV 
element is part of the lower turret assembly which is described more fully in Chapter 9. 
The display is presented to the pilot/gunner by means of a TV raster display, originally on 
a CRT, but more recently this has been upgraded to a colour active matrix liquid crystal 
display (AMLCD). Military display technology is described in Chapter 11. The FOV 
options available for the TV are: 







The TADS turret is capable of traversing ± 120° in azimuth and + 30 to -60° in elevation 
with respect to the aircraft axes. Aircraft manoeuvring may reduce the turret FOV in 
certain circumstances. 

2. RQ-4 Global Hawk. The CCD imaging system is a theatre reconnaissance high-resolution 
imaging system designed to provide high-quality images for intelligence purposes. The 
field of view is specified in milliradians (mrad), where a radian is equivalent to 57.3°, 


therefore 1 mrad is equivalent to 0.0573° (1 mrad is the angle subtended by an object of 
1 m length at a range of 1000 m). The specified FOV for the CCD imaging device, part of 
the integrated sensor system (ISS) is 5.1 x 5.2 mrad (0.3° x 0.3°). The platform is 
capable of providing the following coverage in a 24 h period: 

Mode Coverage/quality 

Wide area survey 138 400 km 2 /24h to NIIRS 6 quality 

Spotlight 19 002 km x 2 km spot images/24 h at NIIRS 6.5 quality 

The national imagery interpretability rating scale (NIIRS) is an imaging classification for 
radar, IR and visual imaging systems. The scale is from to 9, where represents 
unusable and 9 represents the highest quality. The scale equates to qualitative criteria to 
which the image interpreter can easily relate. For example: NIIRS 6 quality allows the 
spare tyre on a medium- sized truck to be identified; NIIRS 7 quality allows individual 
railway sleepers (rail ties) to be identified. 

Low-light TV is accomplished by the use of image intensifier tubes or devices that amplify 
the image incident from the scene in view. Today, the use of night- vision goggles (NVGs) 
provides operators with a look-up, look-out capability that is far more flexible than the use 
of a dedicated low-light TV (LLTV) sensor, and this technology is described in the next 

5.3 Night-vision Goggles 

Night- vision devices gather ambient light and intensify the image by using a photocathode 
device to produce electrons from the photons in the incident light. The electrons are 
multiplied and then used to bombard a phosphor screen that changes the electrons back to 
visible light. The sensors are normally configured as a monocular or binocular 'goggle' 
which is attached to the operator's helmet, and through which he views the external low-light 
scene. However, as will be described in Chapter 11, care has to be exercised to ensure that 
the NVGs are compatible with the cockpit displays and night lighting. 

Figure 5.6 shows the principle of operation of the image intensifier. The unenhanced 
image being viewed passes through a conventional objective lens to focus the incoming 
photons that represent the image upon the photocathode. The gallium arsenide (GaAs) 
photocathode produces a stream of electrons that are accelerated with the assistance of 
an applied voltage and fired towards the microchannel plate. Electrons pass through the 
microchannel plate - a small glass disc that has many microscopic holes (or channels) - and 
the effect is to multiply the electrons by a process called cascaded secondary emission. This 
multiplies the electrons by a factor of thousands while registering each of the channels 
accurately with the original image. 

The electrons then impact upon a phosphor screen that reconstitutes the electrons into 
photons, thereby reproducing the original image but much brighter than before. The image 
created on the phosphor screen passes through an ocular lens that allows the operator to 
focus and view the image. As the image is recreated using a green phosphor, the intensified 
image possesses a characteristic green-hued outlook of the object being viewed. 



Photons Electrons Electrons Photons 


Image Intensifier 
Figure 5.6 Principle of image intensification. 

Night- vision devices (NVDs) were originally used over 40 years ago and employed an IR 
source to illuminate the target which was then viewed using an image intensifier. Apart from 
the low reliability of the early imaging systems, the use of an IR illumination source clearly 
provided the enemy with information about the user's whereabouts, and later systems 
obviated the need for IR illumination. This technology is sometimes termed generation 0. 
This led to the development of passive imaging devices which today are classified by four 
generations, each generation providing significantly improved performance over the pre- 
ceding generation. A summary of these generations and the technology advancements 
associated with each is given below: 

1. Generation 1. This generation used the ambient illumination provided by the moon and 
stars. The disadvantage was that the devices did not work well on cloudy or moonless 
nights. Essentially the same technology was used as for generation so that poor 
reliability was also an issue. 

2. Generation 2. Major improvements in image intensifier technology enhanced the perfor- 
mance of the image intensifier and improved the reliability. The biggest improvement 
was the ability to see images in very low ambient light conditions such as may be found 
on a moonless night. This technology introduced the microchannel plate which not 
only magnified the electrons but created an image with less distortion because of the 
'channelling' effect. 

3. Generation 3. This has minor technology additions by way of manufacturing the photo- 
cathode using gallium arsenide (GaAs), a material that is very efficient at converting 
photons into electrons and makes the image sharper and brighter. Another improvement 
was coating the microchannel plate with an ion barrier to improve reliability. Generation 3 
technology is typical of that in use with the US military today. 

4. Generation 4. Generation 4 is referred to as 'filmless or gateless' technology. The ion 
film was removed to allow more electrons to pass through the microchannel plate, thereby 
improving the tube response. Gating of the power supply allows the tube to respond more 
quickly to rapid changes in lighting conditions. 

The use of NVGs is usually achieved by clamping them on to the operator's helmet as 
shown in Figure 5.7. 



Figure 5.7 Helmet with mounted night- vision goggles. (Infrared 1) 

Figure 5.8 shows a typical NVG image of a AV-10 in flight. Therefore, as well as 
providing images to survey and attack an enemy, NVGs can be of considerable use in 
allowing friendly forces to operate covertly, such as air-to-air refuelling at night without 

More recently, the use of helmet-mounted displays (HMDs) incorporating night-vision 
devices has been adopted in the military combat aircraft community, and these are discussed 
in detail in Chapter 11. 

5.4 IR Imaging 

IR imaging by day or by night has become one of the most important sensing technologies 
over the past 30 years. In that time, technology has advanced appreciably up to the point 
where the quality of IR imaging and visible light imaging is virtually indistinguishable. 

As has already been described, the visible light spectrum from 400 to 750 nanometres, 
SWIR from 1000 to 2500 nanometres (1-2.5 urn), MWIR from 3000 to 5000 nanometres 
(3-5 urn) and LWIR from 8000 to 14 000 (8-14 urn) are very close to one another in the 

Figure 5.8 Typical NVG image (AV-10 in flight). (Infrared 1) 




( 3 to 5) 

Expanded Area 

o o o o 
o o o o 
<* m co i^ 

Wavelength (Jim) 

Figure 5.9 Black body radiation. 

electromagnetic spectrum. It is therefore hardly suprising that many of the characteristics are 
very similar. 

Sensing in IR wavelengths is basically about sensing the radiation of heat. All objects 
radiate heat primarily depending upon their temperature but also to some extent upon the 
material and the nature of the surface. In classical physics the emission of thermal energy is 
referenced to that from a black body which is the ideal thermal emitter. Figure 5.9 shows 
typical plots of radiated energy versus wavelength for a black body whose temperature is 
900 K (627°C) and 1000 K (727°C) respectively. The higher the temperature, the higher are 
the levels of radiated energy. It can also be seen that the peak value of the radiated energy 
moves to the left (decreases in wavelength) the hotter the object becomes. This characteristic 
is called Wien's law and will be examined in more detail shortly. 

The problem with this model is that not all objects are black and are perfect radiators. This 
can be accounted for by applying an emissivity coefficient that corrects for an imperfect 
radiator. Table 5.1 tabulates the emissivity coefficient for some common materials. 

It can be seen that most building materials have fairly high emissivity coefficients. On the 
other hand, metals have a relatively low value when polished, but this increases appreciably 
when the surface oxidises. Aluminium is slightly different as it has a higher value when 
anodised which decreases if the surface is oxidised. 

The other effect that Figure 5.9 illustrates is the fact that, if an object gets sufficiently hot, 
it emits visible light. It can be seen that an object at 1000 K is beginning to radiate energy in 
the red portion of the visible light spectrum. If the object were to be heated further, this area 
would encroach to the left. Eventually, if the object were heated to a sufficiently high 
temperature, then it would emit energy right across the visible light spectrum in which case it 
would appear white. This tallies with what everyone knows: if you heat an object it will first 
begin to appear red ('red hot'), and if the object is heated to a high enough temperature it will 
eventually appear white ('white hot'). 


Table 5.1 Emissivity coefficients for some common materials 

Surface material Emissivity coefficient 

Black body (matt) 1.00 

Brick 0.90 

Concrete 0.85 

Glass 0.92 

Plaster 0.98 

Paint 0.96 

Water 0.95 

Wood (oak) 0.90 

Plastics (average) 0.91 

Aluminium (oxidised) 0.11 

Aluminium (anodised) 0.77 

Copper (polished) 0.04 

Copper (oxidised) 0.87 

Stainless steel (polished) 0.15 

Stainless steel (weathered) 0.85 

The effect of Wien's law is presented in Figure 5.10 which shows power density versus 
wavelength for different body temperatures. As can be seen, the wavelength of the power 
density peak decreases as the temperature of an object increases. Summarising the data: 

Wavelength of power 

Temperature (K) 

density peak (nanometres) 









The reverse side of this law is that the peak value falls off very rapidly as the temperature of 
the object decreases. The peak power density wavelength for an object at 373 K (or 100°C) is 
7774 nanometres (7.77 urn); for an object at 290 K (~17°C or room temperature) it is 
10 000 nanometres (10 urn). 
Wien's law provides a formula for the peak wavelength as follows: 

^peak — 


where A peak is the peak wavelength (urn) and T is the temperature of the object (K). 

This suggests that, theoretically, to obtain a maximum response from the IR sensor in the 
region where people and vehicles radiate, the wavelengths calculated above should be 
chosen, i.e. -8000-10 000 nanometres (8-10 urn) at the lower end of the LWIR band. All 
things being equal, this should be the case. However, other factors such as availability and 
maturity of sensors to operate in the band and the effect of scattering due to haze or smoke 
may also have an impact. Also, as the wavelength increases, the size of the optics should also 




T = 6000°K; X peak = 483 i 



724 nm 

peak = 966 nm [IR 

► Wavelength nm 

100 500 

Figure 5.10 Wien's law. 





increase (similarly to radar since both IR and radar emissions are electromagnetic waves), 
and therefore angular resolution will reduce with increasing wavelength unless the optics are 
scaled proportionately. As ever, space and volume are at a premium in a military avionics 
installation, and in some cases the increase in volume to accommodate larger optics is 
unlikely to be acceptable. Medium-wave operation is generally preferred both in high- 
temperature humid (tropical) and in arid (desert) conditions owing to the 3-5 urn window. 
The US Army generally has preference for LWIR operation which is better with haze and 
smoke (being at the radar end of the IR spectrum, this band has characteristics closer to that 
of radar). Some sophisticated systems provide dual-band operation - MWIR and LWIR - to 
enjoy the best of both worlds. 

5.4.1 IR Imaging Device 

A generic IR imaging device is shown in Figure 5.11. The target is shown emitting radiation 
on the left side of the diagram, with the radiation spectral energy determined by a 
combination of absolute temperature and emissivity. The radiated power has to compete 
with a number of extraneous sources: background radiation, sky radiance, the sun, reflections 
from clouds and other unwanted signal sources that generate clutter against which the target 
has to be detected. As well as the clutter, the energy radiated by the target is subject to atmos- 
pheric attenuation which can be particularly acute at low level and at certain frequency bands. 
The incoming energy is focused by an appropriate set of optics, and in most cases some 
scanning arrangement is necessary to scan the target on to the detector array. Some arrays 
called 'staring arrays' do not need the optical scanning, and these will be described later in 
the chapter. Once the detector has formulated the IR image, the result is read out in a similar 




Scanner Detector & 




Figure 5.11 Generic IR imaging system. 

Pre-amp & 



way to the CCD sensor and the resulting data are amplified, processed and displayed. Most 
sensors need cooling in order to operate - usually to around 77 K - and special cooling 
systems are needed to perform the cooling task. A range of sensor materials can be used, all 
of which have their own particular advantages and bands of operation. 

Three typical detector configurations are shown in Figure 5.12, and each type is used for 
different types of IR operation. These are: 

1 . Linear array. The linear array is used to form an image strip, and a scene image may be 
generated by successively adding the strips together. The 1024 x 8 array illustrated is one 
used by BAE SYSTEMS. 



Figure 5.12 Typical IR detector configurations. 



2. Two-dimensional array. The two-dimensional array forms an X—Y matrix that lends itself 
readily to generating a rectangular image in a similar way to the CCD array described 
earlier. The 640 x 480 and 320 x 240 two-dimensional arrays portrayed are typical of 
state-of-the-art third-generation systems in service today. 

3. Cruciform array. This array is used to accomplish an IR tracking function and will be 
described later in the chapter. 

The scanning configuration adopted augments the detector configuration used to provide an 
IR image that may be examined for strategic reconnaissance, intelligence gathering, battle 
damage assessment or for a platform operator to prosecute an engagement. Three basic IR 
scanning techniques for imaging will be described. These are the rotating scanner, planar 
array and focal plane array (FPA). 

5.4.2 Rotating Scanner 

One of the first techniques to be employed was the rotating optics method which was also 
known as linescan. In Figure 5.13 the platform is flying from left to right as the scanner 
rotates about an axis parallel with the aircraft heading. In this case the scanner is rotating in a 
clockwise direction looking forwards, and successive strips of ground are imaged as the 
imaging mirror sweeps from right to left. Therefore, as the platform flies forwards, the series 
of image strips may be recorded and an area image may be constructed of the ground that has 
been scanned. 

This technique was one of the first to be used for area IR imaging and was evolved using 
very small arrays as the technology was not available to produce large linear arrays. The 
image suffers from distortion towards the horizontal limits of the scan as the sightline moves 
appreciably away from the vertical. Furthermore, for clear images, the relationship between 
aircraft velocity, V, and height above the terrain, h, or V/h, has to be closely controlled or 
successive imaging strips will not be contiguous or correctly focused. Another disadvantage 
of this method when it was first introduced into service was that there were no high-density 
digital storage devices available. The images were therefore stored on film which had to be 
developed after the sortie before analysis could begin. Early IR linescan systems such as 

Rotating f\ 
Optics \j 

Direction of 

Figure 5.13 Rotating scanner (IR linescan). 



those carried by the UK F-4K Phantom carried the system in a large pod beneath the aircraft 
centre-line. This technique has been likened to a whisk broom where the brush strokes are 
sequential right to left movements. 

The Royal Air Force Tornado GR4a and Jaguar GR3 reconnaissance variants use an 
embedded IR linescan VIGIL system produced by Thales/Vinten. This system has the 
following attributes: 


Scan rate 

Angular resolution 

Single cadmium mercury telluride (CMT) (CdHgTe) 

detector operating from 8 to 14 urn 

600 lines/s 

<0.67 mrad 

8192/line or -4.9 Mpixels/s 

23 lbs 

5.4.3 Planar Image 

The planar image technique is shown in Figure 5.14. By comparison with the rotating 
scanner system just described, this is called a 'push broom' system since it is analogous to a 
broom being pushed forwards. This system uses line detector arrays as outlined above. As 
the aircraft moves forwards, the optics allow the strip detector to image the area of interest as 
a series of strips that can then be formed into a continuous area image. This type of scanning 
arrangement lends itself to high-altitude imaging systems on platforms such as the Global 
Hawk. The main operational capabilities of the Global Hawk EO system are outlined below 
(NIIRS is the national image interpretability rating scale): 

Detector Indium antimonide (InSb) detector operating from 3.5 to 7 jam (MWIR) 

Field of view Wide area scan: 5.5 x 7.3 jirad 
Performance Wide area scan: NIIRS 5 

Spotlight: NIIRS 5.5 

Direction of 

Figure 5.14 Planar image. 



Figure 5.15 Focal plane array. 

The performance of the IR imaging is almost as good as the CCD visual imaging system 
for the Global Hawk described earlier in the chapter, where the corresponding figures were 
NIIRS 6 and 6.5. 

5.4.4 Focal Plane Array or 'Staring Array' 

The focal plane array (FPA), often referred to as a 'staring array', is portrayed in Figure 5.15. 
The FPA provides an image on to a focal plane that coincides with the sensing array, most 
usually a two-dimensional array whose dimensions scale easily to a standard rectangular 
display format: NTSC, PAL and, more recently, VGA and XVGA and above, greatly 
simplifying the optics. Although the figure depicts the focal plane array with a vertical axis, 
in tactical systems the array face is usually facing directly towards the target. In most cases 
the forward looking IR (FLIR) sensor is looking forwards, the term being relative as it is 
usually mounted upon a gimballed assembly that has extreme angular agility and slew rates 
in order to be able to track targets while the platform is manoeuvring. As will be seen later 
in the chapter, several EO sensor systems are commonly physically integrated into the 
co-boresighted sensor set to aid sensor fusion and allow target data to be readily handed off 
from one sensor type to another. 

In an array the entire surface is not given over to IR energy sensing. There is a certain 
overhead involved with interconnecting the array which prevents this from being the case. In 
a practical array the overhead is represented by a term called the fill factor which describes 
the useful portion of the array as a percentage. On modern state-of-the-art arrays, the fill 
factor is usually around 90 %. 

The array is effectively read in a sequence of frames in the same way as any other 
real-time imaging device. Therefore, the time between successive read-outs of the array 
image is the time available for the array to 'capture' the image, and this is referred to as the 
integration time and permits successive images of the target to be generated. 

The key element in the performance of any IR imaging device lies in the performance of 
detectors and the read-out of the imaged data in a timely fashion. There are many sensor 
types and technology issues to be considered, and some of the detector technology issues are 
outlined briefly below. 



Table 5.2 Overview of IR FPA detector technologies 


Wavelength (urn) Typical array (FPA) Cooling (K) Application 

Lead silicide 



Indium antimonide 



Cadmium mercury 


telluride (CMT) 


Lead tin telluride 


(LTT) (PbSnTe) 

Quantum well 


infrared photodetector 

(QWIP)(GaAs; AlGaAs) 

Not generally used for military applications 

640 x 480 
512 x 512 
320 x 240 
640 x 480 

320 x 240 






Apache M-TADS 
Litening II pod 

Experimental for 

Al, aluminium; As, arsenide; Ga, gallium; Hg, mercury; In, indium; Pb, lead; Sb, antinomy; Si, silicon; Sn, tin; 
Te, tellurium. 

5.4.5 IR Detector Technology 

The technology of the IR imaging detectors is rapidly moving in terms of materials and array 
size. Table 5.2 gives a brief overview of some of the key technologies for the FPA 
implementation in aerospace applications. Many of the materials developed for medical 
and industrial use may not be suitable for aerospace applications. This is a rapidly evolving 
area of technology and the details of new technologies are not always available in the public 

For reasons indicated earlier, most applications today are based in the MWIR and LWIR 
bands, although the band chosen will be dependent upon detailed specification requirements. 
There is a desire to move towards dual-band operation where the optimum wavelength may 
be chosen for the imaging task in hand. There is also an aspiration to introduce multispectral 
imaging technology to aerospace applications because of the increase in operational 
capability that would bring. At the moment, contemporary technology may find it difficult 
to discriminate targets hidden beneath camouflage nets or foliage. Multispectral sensing will 
provide battlefield commanders with sensors that would be able to overcome this deficiency. 
The typical desired capabilities of a modern sensor are summarised below: 

Pixel pitch 
Frame rate 

Maximum integration 
Data rate 
Array size 

~20-40 urn 

50 Hz (PAL); 60 Hz (NTSC) with a desire 

to go to 100 Hz and above 
99% of frame time 

10 MHz upwards 

640 x 480 (VGA resolution), heading towards 
1000 x 1000 (1 Mpixels) or above in next 
generations: F-35 and space applications 



- Displacer 


Figure 5.16 Stirling cycle cooler. 

It can be seen that all the sensor detector types require cooling, and there are two ways of 
doing this. Originally, cooling was achieved using a Dewar flask together with a liquid 
nitrogen cryogenic coolant. More recently, miniature refrigerator devices have been devel- 
oped that work on a Stirling cycle principle. The Stirling machine and the associated cycle is 
shown in Figure 5.16. The Stirling machine comprises a compressor cylinder with two 
moving pistons on the right; these pistons can be moved by means of linear electrical motors. 
This cylinder has finned heat exchangers to assist in dumping heat overboard. In the second 
cylinder on the right, the heat load is mounted on a 'cold finger', at the top of the cylinder. 
This cooling cylinder contains a regenerator device which is free to move up and down but 
which is normally biased to the top of the cylinder by means of a spring. The regenerator 
device has the ability temporarily to store heat, accepting heat from the cycle and donating it 
back to the cycle during different phases. The Stirling machine operates in four discrete 
changes in pressure/volume (P/V) during one cycle, and the cycle has the overall effect of 
extracting heat from the cold finger abutting the sensor and rejecting it from the machine by 
means of heat exchangers on the compressor and regenerative cylinders. The linear motors 
are powered by an aircraft or pod power supplies which draw relatively small amounts of power. 
The principle of operation of the Stirling cycle is described below. At the start of the cycle 
the pistons PI and P2 are at the top and bottom of the compression cylinder respectively and 
the regenerator is at the top of the cooling cylinder: 

1. Phase 1. The linear motors compress the gas, and the heat so generated is dissipated in 
the heat exchangers. The black arrow on the subdiagram at top right shows that heat is 
rejected from the cooler. 

2. Phase 2. The pistons remain in their compressed position so that the volume of the shared 
gas is constant. The gas above the regenerator expands while moving the regenerator 
down, compressing the spring and releasing heat into the regenerator (white arrow). 

3. Phase 3. The pistons are returned to their original positions at the top and bottom 
of the compression cylinder by the linear motors, increasing the volume of the 



shared gas. Heat is rejected from the heat load/seeker assembly into the cycle (black 
4. Phase 4. The regenerator releases heat into the shared gas (white arrow) at constant 
volume and therefore pressure increases. The spring biases the regenerator to the top of 
the cooling cylinder. 

The cycle is repeated continuously and a heat load is withdrawn from the seeker assembly, 
causing it to cool down rapidly. The characteristics of a typical Stirling cooler are: 

• Input power —30-50 W; 

• Heat load (seeker) -0.5-1.5 W; 

• Seeker operating temperature —77 K; 

• Cool down time 5-10 min; 

• No seals, no lubricants, no maintenance, sealed for life; 

• MTBF -5000-10 000 h. 

IR detector packaging for second-generation arrays is now possible in a number of forms that 
are illustrated in Figure 5.17. These are direct hybrid, indirect hybrid, monolithic and Z 
technology (in all cases except the monolithic method, the electrical connection for the 
detector chip is made by means of indium 'bumps' which provide a soft metal interconnect 
for each pixel): 

1 . Direct hybrid. In this configuration the chip is connected to an array of preamplifiers and 
row and column multiplexers to facilitate the read-out. 




|- Derecto- j 

■ m mi- 


Jndium Metal_ 
"Bump Bond - 

f Sector | i Readout I 

| Da:sctor-Rea€cui;| Signal Processing 


m m ! 

— - 

Figure 5.17 IR detector packaging schemes. 


2. Indirect hybrid. This is similar to the direct method except that the detector and read-out 
electronics are interconnected by fan-out which connects the two chips electrically by 
means of a metal bus on a fan-out substrate. This has advantages in testing the detector 
array and allows the size of the preamplifiers to be increased to improve dynamic 

3. Monolithic. In this method both detector and signal processor are combined in the same 
chip which in turn is mounted on the same substrate as the signal processing. In fact, the 
two do not have to be packaged on to the same substrate but can be segregated in terms of 
substrate and operating temperature, thereby possibly reducing the cooling load by 
cooling the detector alone. 

4. Z technology. This provides additional signal processing space on a pixel-by-pixel basis 
in the Z direction (as opposed to the x-y array direction). This is used when the detected 
output of every pixel is to be individually processed, as is the case in multispectral and 
hyperspectral applications (Chan et al., 2002; Bannen and Milner, 2004; Carrano et al., 

5.5 IR Tracking 

The use of IR seeker heads to track and engage targets has been in use in military systems for 
many years. The Raytheon AIM-9 Sidewinder missile was one of the first of many such 
systems to be deployed. It is still in service today with many air forces around the world, and 
the latest version AIM9-X is about to enter service with the US Armed Forces. Petrie and 
Buruksolik (2001) give an interesting perspective on the history of the Sidewinder. The 
introduction of simple man-launched surface-to-air missile (SAM) such as Stinger has been a 
feature of the use of IR technology. The threat of such weapons is still very much with the 
aviation community today when used by renegade or terrorist organisations to attack 
unarmed military or civil transport aircraft. IR search, track and scan (IRSTS) systems are 
used as a primary sensor system on many fourth-generation fighter aircraft. 

5.5. 1 IR Seeker Heads 

To illustrate some of the capabilities and limitations of IR tracking devices, the use of IR 
seekers in an air-to-air missile context will first be examined. 

Reticle tracking is achieved by rotating a small disc or reticle with clear and opaque 
segments in front of the seeker detector cell. In early IR tracking heads the detector would 
have been a simple arrangement; later versions used more complex detector arrays, and the 
very latest missiles use an FPA array with a better FOV. 

A simple example of a tracking reticle is shown in Figure 5.18. Either the disc rotates or 
the IR image is rotated by means of a rotating mirror. Whichever method is used, the 
objective is to scan the IR image with relative rotary movement of the reticle and modulate 
the IR return. Figure 5.18 shows a simple reticle that is translucent and allows 50% 
transmission on one half while alternately chopping the image on the other between clear 
and opaque sectors. This modulation technique yields a series of pulses of IR energy 
that is detected by the detector cell. By carefully choosing the characteristics of the reticle 
and therefore the resulting modulation, an error signal may be derived which allows the 
seeker head to track the target by suitable servo drive systems. The reticle scan rate of early 





50% Transmission 











50% Transmission 



Figure 5.18 Simple reticle tracker. 

seeker heads was ~50-70 Hz, not dissimilar from the radar conscan tracking described in 
Chapter 3. 

The choice of the reticle type determines the kind of modulation that is employed to track 
the target - most seeker heads use either amplitude modulation (AM) or frequency 
modulation (FM). A commonly used technique employs a wagon wheel stationary reticle 
with nutating optics scanning in a rosette pattern, rather like the petals of a flower. The type 
of reticle, type of modulation and frequency of rotation/nutation for a given application are 
usually not advertised, as to do so would reveal key characteristics of the head and its 
performance. Unlike radar tracking techniques such as conscan, where the characteristics of 
the radiated power reveal the angular scan rate, IR scanning is passive and therefore more 
difficult to counter by deception means. 

Earlier, a cruciform detector was shown in Figure 5.12. The operation of a cruciform or 
cross-configured seeker is shown in Figure 5.19. This system uses a stationary element with 
nutating optics which scans the image in a circular fashion over the arms of the cruciform. If 
the target is located on the boresight of the seeker, as shown on the left, the time between 
pulses received from elements 4 and 1 will equal that between pulses received from elements 
3 and 4. If the target drifts off boresight in a 2 o'clock direction as shown on the right, the 
pulses will be an unequal distance apart. Successful tracking is achieved by using pulse 
period measuring techniques with the appropriate servomechanisms to maintain the target 

Early seeker heads possessed a limited capability, only able to engage the target from the 
rear aspect where the IR tracker had clear sight of the engine jet-pipe and exhaust plume. 
Part of the limited performance was that early missile detectors were uncooled lead sulphide 
(PbS) elements, so the sensitivity was very low. The instantaneous field of view was ~4° 
with a seeker head FOVof ~25°. Sightline tracking rates were also low (~11 deg/s), so the 
engagement of manoeuvring targets was not possible. 



Target On- 


Target Off- 

Figure 5.19 Tracking using a cruciform detector array. 

With developments introduced in the 1960s, cooled arrays were introduced using a 'one- 
shot' liquid nitrogen bottle located in the missile launcher. The coolant bottle was renewed 
before each sortie and contained enough coolant to allow the missile head to operate for 
~2.5 h. Modern systems are capable of full-aspect engagements; that is, they are sufficiently 
sensitive to acquire and track the target aircraft from any position. 

In close encounter engagements the sightline spinrate of the target may be high as it 
crosses rapidly in front of the aircraft/seeker head. This can occur so quickly that the seeker 
head is unable to acquire the target. The solution to this problem is to use one of the aircraft 
sensors to track the prospective target and slave the missile to that sightline. The Royal Air 
Force used this technique to slave the Sidewinder missile sightline to the AWG 11/12 radar 
sightline in the F-4M Phantom in a mode called Sidewinder expanded acquisition mode 
(SEAM). The technique is still used today, except that the missile seeker sightline is slaved 
in a more sophisticated manner to the system or pilot cues. Most fighter aircraft entering 
service in the last 10 years are apt to have a means of slaving the seeker boresight to steering 
cues given by a helmet-mounted sight (HMS) projecting directly on to the pilot's visor/sight. 
See the discussion on this topic in Chapter 1 1 . 

5.5.2 Image Tracking 

The high fidelity of IR imaging systems as already described opens up the possibility of 
image tracking and also image recognition, although the algorithms involved with the latter 
function can be quite complex. The resolution available with imaging systems is now close 
to or approaching that available with visible range sensors, and therefore specific objects 
may be easily tracked once identified and designated by the operator. As will be seen, this is 
an important feature in engaging a target, as sensor fusion using a combination of sensors, 
trading off relative strengths and weaknesses, is in many cases an important feature in the 
successful prosecution of a target. 

TV and IR imaging provides good resolution in an angular sense but not in range. Radar 
and lasers offer good range resolution but poor angular resolution. Using the right 
combination of sensors provides the best of both. 


Typical tracking algorithms include the following: 

1 . Centroid tracking, where the sensor tracks the centre of the target as it perceives it. This is 
particularly useful for small targets. 

2. Correlation techniques that use pattern matching techniques. This is useful to engage 
medium to large targets but can be difficult if the target profile alters drastically with 
aspect, for example an aircraft. 

3. Boundary location or edge tracking can be used where the target can be split into 
segments, the arrangement of the segments providing recognition features. 

The use of a human operator is most useful in ensuring that the correct target is identified 
and tracked. Target recognition is also vital under most rules of engagement where it 
is essential to have the ability to fire without positive identification in order that no 'blue- 
on-blue' or friendly fire incidents occur. Again, correlation of imagery with other sources/ 
sensors can be of great assistance. However, in high-density dynamic target situa- 
tions the human operator will soon reach saturation and automatic target tracking will be 

5.5.3 IR Search and Track Systems 

IR search and track systems (IRSTS) have been used for air-to-air engagements for some 
time. The US Navy F-14 Tomcat has such a system, and the Soviet-designed aircraft MIG 
29, SU27 and SU35 all used first-generation systems. The function of IRSTS is to perform a 
function analogous to the airborne radar TWS mode where a large volume of sky is searched 
and targets encountered within the large search volume are characterised and tracked. The 
major difference is that, whereas the radar TWS mode is active, IRSTS is purely passive. 
The key requirements of an IRSTS are: 

• Large search volume; 

• Autonomous and designated tracking of distant targets; 

• Highly accurate multiple-target tracking; 

• Passive range estimation or kinematic ranging where sightline spin rates are high; 

• Full integration with other on-board systems; 

• FLIR imaging; 

• High-definition TV imaging. 

A state-of-the-art implementation of IRSTS is the passive infrared airborne tracking 
equipment (PIRATE) developed by the EUROFIRST consortium which will be fitted to 
the Eurofighter Typhoon. Figure 5.20 shows the PIRATE unit and the installation on 
Typhoon of the left side of the fuselage. The equipment uses dual-band sensing operating 
in the 3-5 and 8-11 urn bands. The MWIR sensor offers greater sensitivity against hot 
targets such as jet engine efflux, while the LWIR sensor is suited to lower temperatures 
associated with frontal engagements. The unit uses linear 760 x 10 arrays with scan motors 
driving optics such that large volumes of sky may be rapidly scanned. The field of regard 
(FOR) is stated to be almost hemispherical in coverage. The detection range is believed to 
be ~40nm. 



Figure 5.20 PIRATE seeker Courtesy Trailers phonics and installation on Eurofighter Typhoon 
(Eurofighter GmbH). 

The operational modes of PIRATE are: 

1. Air-to-air: 

• Multiple-target tracking (MTT) over a hemispherical FOR - the ability to track in 
excess of 200 individual targets, with a tracking accuracy better than 0.25 mrad; 

• Single-target track (STT) mode for individual targets for missile cueing and launch; 

• Single-target track and identification (STTI) for target identification prior to launch, 
providing a high-resolution image and a back-up to identification friend or foe (IFF). 

2. Air-to-ground: 

• Ability to cue ground targets from C 3 data; 

• Landing aid in poor weather; 

• Navigation aid in FLIR mode, allowing low-level penetration. 

The sensor data may be displayed at 50 Hz rates on the head-down display (HDD), head-up 
display (HUD) or helmet-mounted display (HMD), as appropriate. 

5.6 Lasers 

Lasers - the term stands for Light Amplification by Stimulated Emission of Radiation - have 
been used in military systems for almost four decades. The US Air Force used laser-guided 
bombs (LGBs) during the later stages of the Vietnam War, and European avionics systems 
such as Jaguar and Tornado were adopting laser systems during the late 1960s and early 
1970s. These systems are now commonly used as range finders, target designators and 
missile/bomb guidance. Laser systems may be fitted internally within the aircraft such as in 
aircraft like the Tornado GR4 and F-18. They may be housed in pods for external carriage on 
weapons/stores stations on fighter aircraft, or they may be housed in highly mobile 
swivelling turrets for helicopter and fixed- wing use. 



High Voltage 
Power Supply 

Flash Lamp 



Laser Medium 

Efficiency ~ 2 - 








Figure 5.21 Principles of operation of a laser. 

The major advantage of using lasers is the fact that they can provide range information 
that passive systems such as visible light and IR radiation cannot. Lasers are therefore 
particularly useful when used in conjunction with these other technologies to provide sensor 
fusion, blending and merging the advantages and disadvantages of the different capabilities. 
Some of these integrated systems are described later in the chapter. 

5.6. 1 Principles of Operation 

The principles of operation of a laser depend upon exciting the energy levels of electrons 
within specifically 'doped' materials. Electrons within the material are stimulated to higher 
energy levels by an external source; when the electrons revert to a lower energy level, energy 
of specific wavelength is emitted, depending upon the material and the energy supplied. 

A laser works on this principle but has other unique properties. Specifically, the energy 
that is emitted is coherent; i.e. the radiated energy is all in phase rather than being randomly 
related as may be the case during light emission. Figure 5.21 shows a diagrammatic 
representation of a laser device. 

The laser medium may be liquid or solid; most of the lasers used in military systems are 
based upon glass-like compounds. At one end of the medium is a reflecting mirror, at the 
other a partly reflecting mirror. The laser medium is stimulated by an input of energy from a 
flashlamp or other source of energy, which raises the energy levels of the electrons. 

Figure 5.22 shows the various stages that occur for a laser to 'strike': 

1. Stage 1. This is the initial quiescent condition with the electrons all at a natural low- 
energy state. 

2. Stage 2. The flash tube is illuminated, stimulating the electrons and exciting them to 
a higher energy state - this phenomenon is known as population inversion and is an 
unstable state for the electrons. 





• • 


L, , 

Jlas^Tube., „ ] 




V vu 



• • 



^r^K©^ ^>^r^^t 


Stage 1 

Stage 2 

Stage 3 

Stage 4 

Stage 5 


Figure 5.22 Stages leading to a laser strike. 

3. Stage 3. The electrons remain in this state for a short time before decaying down to their 
original, lower and stable energy state. This decay occurs in two ways: 

• Spontaneous decay in which electrons fall down to the lower state while randomly 
emitting photons; 

• Stimulated decay in which photons released from spontaneously decaying electrons 
strike other electrons, causing them to revert to the decayed state - in these cases 
photons are emitted in the direction of the incident photon and with the same phase and 

4. Stage 4. Where the direction of these photons is parallel to the optical axis of the laser, 
the photons will bounce back to and fro between the totally and partially reflecting 
mirrors. This causes an avalanche effect in which the photons are amplified. 

5. Stage 5. Eventually, sufficient energy is built up within the tube for the laser to strike, 
causing a high-energy burst of coherent light to exit the tube. 

The wavelength of the emitted light is dependent upon the nature of the material being used 
in the laser medium since the energy released is specific to the energy levels within the atoms 
of that material. Also, since the amount of energy released is in the same discrete bundles, 
the emitted light is very stable in wavelength and frequency as well as being coherent 
(Figure 5.23). 

Typical pulsed solid-state lasers used in aerospace use the following compounds: 

1. Ruby is chromium-doped sapphire, while sapphire itself is a compound of aluminium 
and oxygen atoms. The formula for ruby is Al20sCr +++ , where Cr +++ indicates the 



Weak Pump 

(730 - 760 nm) 

Strong Pump 

(790 - 820 nm) 


Spontaneous Decay 
\ f (Radiationless Transition) 

1.064 Jim 

Spontaneous & 
Stimulated Emmision 



Spontaneous Decay 
] f (Radiationless Transition) 

Ground State 

Figure 5.23 Energy levels of a YAG laser. 

triply-ionised state of the chromium atom. The ruby laser mode radiates primarily 
at 694.3 nm. The characteristics of this material make it suitable only for pulsed 

2. Neodymium:YAG lasers, where YAG stands for yttrium aluminium garnet (Y3AI5O12). 
This is a popular choice for airborne systems as it may operate in both pulsed and CW 
modes. The YAG laser radiates at 1060 nm. 

3. Neodymium: glass lasers may sometimes be used, but, glass being a poor conductor of 
heat, they are not suitable for continuous operation. Nd:glass operates on a wavelength of 
1060 nm, the same as for YAG. 

Figure 5.23 shows the energy transition levels for a YAG laser. During the excitation state, 
electrons are raised to two energy bands. These are the weak pump band (730-760 nm) and 
the strong pump band (790-820 nm). Electrons in both bands spontaneously decay with a 
radiation-less transition to the upper lasing level. A combination of both spontaneous and 
stimulated emissions occurs as the electrons decay to the lower lasing level. During this 
phase the device radiates energy at a wavelength of 1060 nm or 1.06 urn. Thereafter all 
electrons spontaneously decay to the ground state. 

The stimulation source for lasers is usually a xenon or krypton flash tube. Xenon lamps are 
the best option for ruby lasers, and krypton lamps are a better match for Nd:YAG and 
Nd:glass lasers but are more expensive and so are seldom used. The problem with using a 
flash lamp as the excitation source is that it is very inefficient. Lasers that are lamp pumped 
are very inefficient (~2-3% efficient). The rest of the energy can only be dissipated as heat, 
which causes real problems for the aerospace designer. The reason for this can be seen from 
Figure 5.24. 

The reason for these very low efficiencies is that the flash lamp spectrum is wide compared 
with the narrow band in which the desired spectrum lies. Therefore, the lamp energy is 




YAG Laser 











Figure 5.24 Flash lamp spectral characteristics. 

poorly matched to the band of interest. Modifications can be carried out to shift the lamp 
spectrum more into the red region, but this still presents problems. 

The solution to this problem is to use laser diodes rather than flash lamps to excite the 
laser medium. Laser diodes lend themselves to be more easily tuned to the frequency of 
interest. Figure 5.25 depicts a configuration in which laser diodes are used instead of a flash 
lamp, and this results in higher efficiencies. The higher efficiencies result in a lower 
unwanted heat load that a design has to dissipate, and therefore reliability may improve at 
the same time as performance. Diode-pumped lasers are now used, and this has allowed 
greater usable power output, permitting a designating aircraft to fly much higher while 
illuminating the target, thereby allowing greater stand-off ranges. 


( High Voltage 
Power Supply 

Diode Lasers 


Laser Medium 

Efficiency- 15 -40% 







Figure 5.25 Diode-pumped laser. 



Conventional ^J- 




1 Aperture 






Laser Dispertion ~ 


Figure 5.26 Properties of laser emissions. 

As well as the properties of having a very stable, discrete wavelength and coherent 
transmission, laser emissions possess another important property, that is, the property of low 
dispersion. Figure 5.26 shows a comparison between a light-emitting source and a laser 
emission. The conventional lamp light emits light in all directions, rather like ripples in a 
pool. Even after passing through an aperture, the light diverges into a relatively wide 
wavefront. The laser source has a much narrower beam after passing through the aperture 
and therefore has low divergence. As a result the laser beam still has relatively high beam 
intensities far away from the emitter. Therefore, the laser is able to transmit coherent energy, 
at a fixed stable frequency and with much lower beam divergence than conventional high- 
intensity light sources. 

5. 6.2 Laser Sensor Applications 

The beamwidth of a typical laser is ~0.25 mrad, and this means it is very useful for 
illuminating targets of interest that laser-tuned seeker heads can follow. For example, a laser 
designator illuminating a target at 10 000 ft will have a spot of 2.5 ft in diameter. The first 
deployment of laser-guided bombs (LGBs) used this technique. As the laser designator 
illuminates the target, energy is scattered in all directions in a process called specular 
reflection. A proportion of this energy will be reflected in the director of an observer who 
may wish to identify or engage the target. 

The laser can operate as a range finder when operating in the pulsed mode. Pulsed 
operation has the capability of delivering high power densities at the target, for example, a 
laser delivering a 100 mJ pulse in a 20 ns pulse has a peak instantaneous power of 5 MW. In 
this sense the laser operation is analogous to radar pulsed modes of operation and mean 
power, duty cycle and peak power are equally as important as they are in radar design. 
Even allowing for atmospheric attenuation, a laser can deliver reasonable power densities at a 
target, albeit for very short periods. The narrow pulse width allows accurate range 






Figure 5.27 Laser guidance of an LGB. 

measurements to be made. The 20 ns pulse mentioned above allows range resolutions to 
within ~10ft. 

Therefore, the laser offers a number of options to enhance the aircraft avionics system 
weapon-aiming performance in the air-to-ground mode: 

• Laser designation in CW mode to guide a missile; 

• Laser reception of a target position marked by a third party; 

• Laser ranging to within a few feet. 

Examples of these engagements are shown in Figures 5.27 and 5.28. Figure 5.27 illustrates a 
system where the aircraft has a self-designation capability. The aircraft can designate the 




Figure 5.28 Third-party laser designation by air or ground means. 



target and launch the LGB to engage the target. The designation/launch aircraft illuminates 
the target until the LGB destroys it. This type of engagement is used when the attacking 
force has air supremacy, the launch aircraft is free to fly without fear from counterattack by 
surface-to-air missiles (SAMs) and the target is one that is easy to identify - such as a bridge. 
Third-generation laser systems are capable of engaging from a height of 40 000-50 000 ft 
and perhaps 30 nm from the target. 

In other situations, third-party designation may be easier and more effective. If the target 
is one that is difficult to detect and identify from the air, it may be preferable to use some 
ground forces such as the Special Forces to illuminate the target of interest. The laser 
designator signal structure allows for codes to be set by both ground or air designator and 
launch aircraft so the LGB is used against the correct target. In other cases the designator 
aircraft may possess a higher-quality avionics system than the launch aircraft and may act as 
a force multiplier serving a number of launch aircraft. This technique was used during Desert 
Storm when F-lll aircraft designated targets for F-16s and the RAF Buccaneers designated 
targets for the Tornado GRls. 

The LGBs are not always special bombs but in some cases free-fall bombs fitted with a 
laser guidance kit. This adds a laser seeker head and some guidance equipment to the bomb, 
allowing it to be guided to the target within specified limits, known as a 'footprint'. 
Therefore, provided the LGB is launched within a specified speed and altitude launch 
envelope with respect to the target, it should be able to hit the target if all the systems work 
correctly. The operation of an LGB guidance system is illustrated in Figure 5.29. 

The reflected laser energy passes through optics that are arranged to produce a defocused 
spot on a detector plane. Detector elements sense the energy from the spot and feed four 
channels of amplifiers, one associated with each detector. In an arrangement very similar to 
the radar monopulse tracking described in Chapter 3, various sum and difference channels 
are formed using the outputs of amplifiers A to D. These are multiplexed such that elevation 
and azimuth error signals are produced and then fed to the guidance system which nulls the 


Figure 5.29 LGB Guidance. 


Figure 5.30 F-18 internally fitted laser units. 

The deployment of LGBs was graphically illustrated by media coverage of Desert Storm, 
although there was by no means a 100% success rate and a relatively small proportion of 
bombs dropped during that campaign were laser guided. By the time of the second Gulf War, 
a much higher proportion of laser-guided weaponry was used. However, as in all systems, 
there are drawbacks. Airborne tactical laser systems, by the very nature of their operating 
band very close to visible light, suffer degradation from the same sources: haze, smoke, 
precipitation and, in the desert, sandstorms. In other operating theatres, laser-guided systems 
may suffer more from weather limitations compared with operating in relatively clear 
conditions in the desert. 

As was stated earlier, lasers may be fitted internally in some aircraft and in pods on others. 
Figure 5.30 shows the three units that comprise the laser system for the F-18. 

5.6.3 US Air Force Airborne Laser (ABL) 

The use of the lasers described so far relate to the tactical use of lasers to mark targets, 
determine target range and designate the target so that it may be engaged by a variety of air- 
launched weaponry. These lasers are not 'death rays' and, although they exhibit reasonably 
high energy levels, they are not sufficiently powerful to destroy the target by energy alone. 
They can, however, cause serious damage to the human eye, and laser safety issues are 
discussed later in the chapter. 

The US Air Force airborne laser (ABL), designated as YAL-1, is a high-energy laser 
weapon system designed to destroy tactical theatre ballistic missiles. It is being developed by 
the Air Force in conjunction with a team comprising Boeing, Northrop Grumman and 
Lockheed Martin. The laser system is carried on board a converted Boeing 747F freighter 
and is curently undergoing test and evaluation. If successful, it is intended to procure several 
platforms with an initial operational capability of three aircraft by 2006 with a total 
capability of seven aircraft by 2010. 

The ABL system actually carries a total of three laser systems: 

1. A low-power, multiple set of laser target-illuminating beams comprising the target 
iluminating laser (TILL) to determine the range of the target and provide information 
on the atmosphere through which the primary beam will travel. The TILL provides the 
aiming data for the primary beam. 

2. A beacon iluminating laser (BILL), producing power in kilowatts, reflects energy from 
the target to provide data about the rapidly changing nature of the atmosphere along the 
sightline to the target. This information is used to bias a set of deformation control mirrors 


in the primary laser beam control system such that corrections are applied to the COIL 
laser beam as it engages the target. 
3. The chemical oxygen iodine laser (COIL) is the primary beam generating the killer beam 
to destroy the target. This beam power is in the megawatt region and operates on a 
wavelength of 1315 urn. When a missile launch is detected, either by satellite or by 
AWACS, the target information is passed via data links to the ABL aircraft. The COIL 
beam is directed at the target by means of a large 1.5 m telescope mirror system at the 
nose of the aircraft which focuses the primary beam on the missile, destroying it shortly 
after launch. 

5.6.4 Laser Safety 

Pulsed solid-state lasers of the types commonly used in avionics applications have eye safety 
implications, as do many laser types. The peak powers involved are so high and the beams so 
narrow that direct viewing of the beam or reflections is an eye hazard even at great distances. 
Nd:YAG and Nd: glass lasers are particularly dangerous because their output wavelength 
(1064 nm) is transmitted through the eye and focused on the retina, yet it is not visible to the 
human eye. The wavelengths at which the human eye is most susceptible are between 400 
and 1400 nm where the eye passes the energy and for the greater part the retina absorbs it. 
Above and below this band, the eye tissue absorbs rather than passes the energy. Lasers can 
be made to operate in a safe mode using a technique called Raman shift; in this way, 
Nd:YAG lasers can operate on a 'shifted' wavelength of 1540 nm, outside the hazardous 
band. In fact, lasers operating on this wavelength may be tolerated with power ~10 5 times 
that of the 1064 nm wavelength. 

Military forces using lasers are bound by the same safety code as everyone else and 
therefore have to take precautions, especially when realistic training is required. The solution 
is that many military lasers are designated as being 'eye safe' or utilise dual-band operation. 
This allows personnel to train realistically in peacetime while using the main system in times 
of conflict. 

To ensure the safe operation of lasers, four primary categories have been specified: 

1. Class I. These lasers cannot emit laser radiation at known hazard levels. 

2. Class IA. This is a special designation that applies to lasers 'not intended for viewing', 
such as supermarket scanners. The upper power limit is 4.0 mW. 

3. Class II. These are low-power visible lasers that emit light above class I levels but at a 
radiant power level not greater than 1 mW. The idea is that human aversion to bright light 
will provide an instinctive reaction. 

4. Class III A. These are intermediate-power lasers (CW - 1-5 mW) that are hazardous only 
for direct beam viewing. Laser pointers are in this category. 

5. Class IIIB. These are for moderate-power lasers. 

6. Class IV. These are high-power lasers (CW - 500 mW; pulsed - 10 J/cm 2 ) or diffuse 
reflection conditions that are hazardous under any conditions (either directly or diversely 
scattered) and are a potential fire and skin hazard. Significant controls are required of 
class IV facilities. 

Many military lasers operating in their primary (rather than eye-safe) mode are class IV 
devices and must be handled accordingly. 


5.7 Integrated Systems 

As the electrooptic technologies used in avionics systems have matured, the benefits of 
integrating or fusing the different sensors have become apparent. Mounting the sensor on a 
rotating gimbal assembly allows the sensor(s) to track targets with high slew rates and with 
several degrees of freedom. Co-boresighting the sensors on this common gimbal assembly 
provides the ability to recognise targets and hand off target data from one sensor to another, 
thereby improving the overall capability of the integrated sensor set. Typical sensors that 
might be arranged in this fashion include: 

• FLIR imaging using FPAs with several FOV options and, in some cases, dual-band 

• CCD-TV with two or more FOV options using monochrome or colour displays and often a 
camera to record results; 

• Laser target markers to illuminate targets, laser range finders to determine range and laser 
spot markers for target hand-off. 

Such sensor clusters have to be carefully aligned or 'harmonised' to ensure maximum 
weapon-aiming accuracy. Also, given the high levels of resolution, the sensor cluster has to 
be stabilised to avoid 'jitter' and provide stable imagery. In some cases this stabilisation will 
be within ~ 15-30 urad. 

These EO integrated systems may take any of the following forms: 

• Installation in a pod to be mounted on a fighter aircraft weapon station; 

• Installation in a turret for use on a helicopter or fixed- wing airborne surveillance vehicle; 

• In stealthy aircraft: internal carriage to maintain low observability. 

5. 7. 1 Electrooptic Sensor Fusion 

The wavelength and frequency of the electromagnetic radiation comprising EO systems are 
contained within a relatively narrow portion of the spectrum: 

• Visible light 0.400 urn (V) to 0.750 urn (R); 

• IR bands 1.0 urn (lower SWIR) to 14.0 um (upper LWIR); 

• Airborne laser 1.064 um and 1.54 um (eye safe). 

Therefore the total sensor set covers the relatively narrow band from 0.4 to 14.0 um or a 
dynamic range of around 35:1. 

Inspite of this relatively low dynamic range or coverage compared with radar and CNI, the 
properties of transmission of some of the sensing technologies are, however, quite different 
for the different bands/wavelengths being used. For example, a laser that would be extremely 
hazardous to the human eye when operated at 1.064 um is eye safe when operated at 
1 .54 um. Perhaps even more important, from the point of view of acquiring and engaging the 
target, the different fields of view (FOV) are quite different between laser and visible/IR 
sensors. The laser beam has very low divergence, typically of the order of 0.25 mrad 





Laser Spot 

Figure 5.31 Typical EO sensor fields of view - Harmonisation. 

(2.5 x 10 -4 rad or 0.014°), whereas a navigation FLIR may have an FOV around 20° x 15° 
(~1400 x 1050 times more). For target engagement activities, narrower FOV modes such as 
4° x 4° (MFOV) or 1° x 1° (NFOV) may be used. The alignment or co-boresighting of the 
EO sensors must be carefully controlled (see Figure 5.31 which illustrates the principle of 
harmonisation but obviously does not show the respective fields of view to scale). 

There are a number of systems issues that are very important and that must be taken into 
account if successful EO sensor fusion and weapons launch are to be accomplished. These 
factors include: 

1. A relatively wide FOV for navigation and target acquisition (^20° x 20°) - in the case of 
the Apache TADS/PNVS, 40° x 30°. 

2. A relatively narrow FOV for target identification and lock on (~4° x 4°) MFOV or 
(1° x 1°) NFOV. 

3. A high target line of sightline slew rates, especially at short range, possibly >90deg/s. 

4. Relatively small angles subtended by the target and the need to stabilise the sensor 
package boresight within very tight limits (especially for long-range targets). A small 
bridge may represent a subtended angle of ~0.024° at 40 nm, and a tank about the same 
angle at 10 nm. The problem may be likened to using binoculars when holding them in an 
unsteady manner. The enlarged image may be visible, but jitter renders the magnified 
image virtually useless unless the glasses can be steadied. Typical head stabilisation 
accuracies on third-generation sensor packages are of the order 15-30 Lirad. 

5. For a variety of reasons it is necessary to provide accurate inertial as well as sightline 
stabilisation for GPS/inertially guided weapons such as JDAM - often referred to as the J- 
series weapons. Accordingly, mainly advanced EO packages have a dedicated strapdown 
inertial navigation unit (INU) directly fitted on to the head assembly to improve pointing 
and positional accuracy and reduce data latency. 

These technical issues are illustrated in Figure 5.32. These demanding technical require- 
ments have to be met within a pod mounted under wing or under fuselage in a hostile 




Sensor Head 





Co-bo res ighted 



Figure 5.32 Boresighting, stabilisation and EO package sightline axes. 

environment while the launch platform could be flying a supersonic speed. As will be seen, 
the space and weight available to satisfy these requirements are not unlimited, and modern 
EO sensor attack or surveillance packages represent very sophisticated solutions to very 
difficult engineering problems. 

Figure 5.33 shows a typical integrated EO sensor for carriage within a 41 cm/16 in 
diameter pod, in this case the Northrop Grumman Litening II AT pod. This sensor pod 
contains the following: 

• Strapdown INS; 

• Wide FOV CCD camera/laser spot detector; 

• Narrow FOV CCD camera; 

• Laser designator/range finder; 

• FLIR. 

The entire sensor package is mounted on a gimbal assembly that is free to move in roll, 
elevation and azimuth. 

5.7.2 Pod Installations 

Podded installations are usually carried on fighter aircraft, although in certain circum- 
stances they could be fitted to other aircraft. For example, targeting pods were Mailed on 
the B-52 during the recent Iraq War, and plans are being reviewed to fit them to B-52 and 
B-l bombers. As opposed to the turreted implementations described later in the chapter, 
EO pods lend themselves to be carried on certain weapons stations of fighter aircraft such 
that the aircraft can be reroled to perform a specific mission merely by fitting the pod. 



(Courtesy Northrop Goumman) 


Camera/Laser Spot 






Figure 5.33 Typical pod-mounted sensor package (Litening II AT). 

Apart from tailoring the mission-specific software to the requirements of the mission, no 
aircraft modifications are required. All the hardware and software modifications to adapt 
the pod to the aircraft baseline avionics system are in place such that all the subsystems 
and display symbology will be compatible to enable the mission to be performed. 
Therefore, EO pods give the battlefield commander additional flexibility in discharging 
his overall battle plan. 

The first pods to be developed were the low-altitude navigation and targeting infrared for 
night (LANTIRN) system introduced into the US Air Force in the late 1980s. In truth, this 
system comprised two pods: 

• The AN/AAQ-13 navigation pod containing a terrain-following radar (TFR) and an FLIR 
system to aid low-level navigation at night; 


• The AN/AAQ-14 targeting pod comprising a targeting FLIR and laser designator/range 

The introduction of these pods in the later stages of the Cold War boosted the capability 
of the US Air Force accurately to attack ground targets in all weather conditions. These 
first-generation pods were fitted to a variety of aircraft including F-14, F-16C/D and F-15E 
in the United States and subsequently to a number of allied Air Forces. The almost 
simultaneous demise of the Soviet Union and the successful deployment and execution of 
Desert Storm in Iraq made military planners realise that most attacks using laser-guided 
munitions would need to be performed from 30 000ft or above if reasonable aircraft 
survivability rates were to be achieved. The focus for the use of EO pods therefore shifted 
from low-level attack to attack from medium level. 

To accomplish this new mission, performance improvements were needed to the targeting 
pod. Low-level ingress and egress to/from the target were not required as it was assumed 
that successful weapons launch could be made from over 30 000 ft above and beyond 
short-range and most medium surface-to-air missile (SAM) threats. The emphasis was 
on long-range target detection and identification and the increased deployment of weapons 
from long range with INS/GPS guidance. These drivers led to some of the performance 
improvements mentioned earlier, and with the 'third-generation' pods now entering service 
these requirements are largely satisfied. 

Table 5.3 is a top-level comparison and summary of the most common targeting pods 
developed by contractors in the United States, United Kingdom, France and Israel and 
deployed from the late 1980s onwards. The latest pods have been designed with COTS in 
mind and are modular in construction so that technology and performance improvements 
may be readily inserted. The modular construction leads to easier maintenance, with faulty 
modules replaced on aircraft in a matter of minutes; high levels of built-in test (BIT) are 
provided readily to check out the system with high levels of confidence. Most of these pods 
have mean time between failure (MTBF) rates of a few hundred hours, roughly equivalent to 
one failure per year at peacetime flying rates. 

Examples of EO targeting pods are shown in Figure 5.34. 

5.7.3 Turret Installations 

Whereas podded installations are useful for loading on to the weapon stations of a fast jet, 
turreted installations are more suited to permanent installations. For clear line of sight to the 
targets, they need to be located at the front of the aircraft, and for these reasons they are 
particularly well suited for installation on helicopters. Some turreted installations have been 
on fixed- wing aircraft such as the B-52H, the US Navy Orion P3-C and the Nimrod MR2 and 
have also been considered for the S-2 Viking. The last three aircraft have anti-submarine 
warfare (ASW) roles. 

The first application of EO turrets on helicopters was with systems like the AH-64 
Apache. Many different helicopters are now fitted with EO turrets, including special forces, 
coastguard and law and drug enforcement. In general, these systems are used in low-level 
and short-range engagements rather than the high-level long-range operation of podded 
fighter aircraft. The atmospheric conditions at low level, combined with other conditions 
such as smoke and haze, mean that LWIR systems often fare better than shorter-wavelength 
systems. However, MWIR with a shorter wavelength offers greater resolution. In some 
recent applications, dual-band MWIR and LWIR sensors are accommodated. 



Table 5.3 Summary and comparison of EO targeting pods 





Carriage aircraft 



L: 199 cm/78 in 





D: 31 cm/12 in 

radar (TFR) 



W: 211 kg/470 lb 

Fixed FLIR: 



L: 251 cm/98 in 

FOV 21° x 28° 

+ 12 


D: 38cm/15 in 

(640 x 512) 



W: 236 kg/524 lb 


Air Forces 


4° x4° 
(640 x 512); 
NFOV 1° x 1° 

Laser designator/ 
range finder 

Many upgrades 
during service, 
including third- 
generation FLIR; 

Over 700 pods 
in service 

40 000 ft laser; laser 

spot tracker; 

CCD-TV sensor; 

digital data recorder; 


generation for 

J- series weapons 



L: 290 cm/1 14 in 


Harrier GR 7 



D: 35 cm/14 in 

10° x 10° 

Jaguar GR1/ 

and laser 

W: 210 kg/462 lb 


GR3 Tornado 


3.6° x 3.6° 




zoom x2; x 4 

added a CCD-TV 



L: 230 cm/87 in 


US ANG F-16 


Litening II 

D: 40.6 cm/16 in 

and NFOV 


Variants: Litening 

W: 200 kg/440 lb 

640 x 480 FLIR 


II - 1999 


Spanish Navy; 

Litening II 

on MWIR: 

Italian Navy 

ER - 2001 



Litening II 

18.4° x 24.1° 

Spanish F/A-18; 

AT - 2003 

(Nav); MFOV 

3.5° x 3.5°; 

NFOV 1° x 1° 
Laser spot 


finder Laser 

designator > 

50 000 ft/40 

Litening III has 
dual mode (including 

eye safe) 

Israeli F-15I; 
Israeli- 16C/D/I; 
German Navy 

and Air Force 

Total of 14 

Air Forces 



Table 5.3 (Continued) 





Carriage aircraft 


Sniper XR 

L: 239 cm/87 in 


USAF F-16 


(extended rang* 

s) D: 30 cm/12 in 

4° x 4°; 

block 50 

targeting pod 

W: 200 kg/440 lb 

NFOV: 1° x 1° 

ANG F-16 

Export version 

640 x 480 FPA 

block 30 

known as 

FLIR operating 



on MWIR 
Laser - diode 
pumped; laser 
>40 000 ft; laser 
range finder/spot 
tracker; dual- 
mode laser 
eye safe); geo 
for J- series 




L: 183 cm/72 in 

640 x 480 

F/A-18A + , 


D: 33 cm/13 in 




W: 191kg/ 

operating in 





for AN/AAS- 


6° x 6°; 

2.8° x 2.8°; 

0.7° x 0.7° 

38 Nite-Hawk 



L: 250 cm/98 in 


Super Entendard 

D: not quoted 


Mirage 2000 

W: 265 kg/ 

sensor: WFOV 

replaces Atlis, 


24° x 18° 
(Nav); MFOV 
4° x 3°; NFOV 
1° x 0.75° 

Laser range 
finder: 1 .5 urn 

Laser designator/ 
range finder/spot 
tracker: 1.06 jam 

Expected to 
be fitted to 

The increasing use of unmanned air vehicles (UAVs) in reconnaissance and combat roles 
has given significant impetus to the production of smaller, lighter systems suitable to be used 
as a UAV payload. 

Figure 5.35 shows typical turrets with imagery examples. 




LITENING Pod (Courtesy Northrop Grumman) 

SNIPER XR Pod (Lockheed Martin) 

SNIPER XR Pod as file (Lockheed Martin) 
Figure 5.34 Examples of EO targeting pods. 



AH-64 C/D Apache TADS/PNVS 
(Lockheed Martin) 


TADS/PNVS Arrowhead (Lockheed Martin) 

MX-20 US Navy P-3 
(Courtesy of UP-45 Squadron) 

es/ei/ae man 




ieee n 

UL AUTO 9149 




, _ \s 








; V 

£¥ ' ' £5" 

■ ■ ■ i 


r " 35 






66:l?:2@W T"GT 


Figure 5.35 Typical EO turrets and imagery examples. 

5.7.4 Internal Installations 

Stealthy aircraft such as the F-117 and F-35 incorporate EO sensor suites to assist in 
engaging ground targets. The F-35 in particular will incorporate an interesting internally 
carried system called the electrooptic sensing system (EOSS). This comprises two major 
functional elements: 



Table 5.4 Summary of typical EO turreted systems. 

Manufacturer Product 


Capabilities Carriage aircraft 

Lockheed AN/AAQ-11 


Direct vision optics: 


Martin target acquisition 


WFOV 18° x 18°; 


and designator 

NFOV 3.5° x 3.5° 


sight (TADS) 

TV camera: WFOV 

Over 1000 


4° x 4°; NFOV 


pilot's night- 

0.9° x 0.9°; 


vision sight 



0.45° x 0.45° 

50° x 50°; MFOV 

10.2° x 10.2°; 

NFOV 3.1° x 3.1°; 

underscan 1.6° x 1.6° 
Laser range 




program commenced 

in 2000: M-TADS and 

M-PNVS utilising 

RAH-66 Comanche 

technology. Features: 

LWIR FLIR using 

640 x 480 FPA; 



eye- safe laser; 

FLIR and IR 

sensor fusion; 

colour CCD-TV 


Raytheon AN/AAQ-27 

Surveillance and 

-16B variant: LWIR 


Follow on to 


FLIR, dual FOV 


the AN/AAQ- 


-16C variant: LWIR 



FLIR, dual FOV 


-16D variant: LWIR 


FLIR, three FOV 

version on 

with laser range 

F-18 as AN/ 





using 640 x 480 


FPA. Dual and three 

RAN Super 

[AN/AAQ - 27 


(3 FOV)] versions 




Table 5.4 (Continued) 

Manufacturer Product 



Carriage aircraft 




Features: high-quality 
and picture quality; 
step zoom 
capability on FLIR 
and TV sensors; 
for pinpointing 
ground target 
location; MWIR 
FLIR with 
640 x 480 FPA: 
WFOV 12.0° x 9.3°; 
NFOV 2.9° x 2.3° 

P-3C Orion 
S-2 Viking 

DAS Coverage 

F-35 Fur Installation (Lockheed Martin) 

F-35 Fur Window (Lockheed Martin) 
Figure 5.36 F-35 EO sensor vertical coverage and EOTS installation. 




IR Warning 



in 6 segments 




Figure 5.37 F-35 horizontal coverage using DAS sensors. 

1. The electro-optic targeting system (EOTS) being developed by Lockheed Martin and 
BAE SYSTEMS. This is an internally carried EO targeting system that shares many 
common modules with the SNIPER XR pod already mentioned. The EOTS looks 
downwards and forwards with respect to the aircraft centre-line, as shown in Figure 5.36. 
The EOTS installation and EO sensor window are shown. 

2. The distributed aperture system (DAS) being developed by Northrop Grumman together 
with BAE SYSTEMS comprises six EO sensors located around the aircraft to provide 
the pilot with 360° situational awareness information that is detected by passive means. 
The concept of horizontal coverage of the DAS is depicted in Figure 5.37. The six 
DAS sensors provide a complete lateral coverage and are based upon technology 
developed for the BAE SYSTEMS Sigma package (shown in the inset). Key attributes 
are dual-band MWIR (3-5 urn) and LWIR (8-10 um) using a 640 x 512 FPA. Each 
sensor measures ~7 x 5 x 4 in, weighs ~9 lb and consumes less than 20 W. Sensor 
devices with megapixel capability (1000 x 1000) are under development and will be 



Atkin, K. (ed.) (2002-2003) Jane's Electro-Optic Systems, 8th edn. 

Bannen, D. and Milner, D. (2004) Information across the spectrum. SPIE Optical Engineering Magazine, 

Capper, P. and Elliott, C.T. (eds) (2001) Infrared Detectors and Emitters: Materials and Devices, Kluwer 

Academic Publishers. 
Carrano, J., Perconti, P. and Bannard, K. (2004) Tuning in to detection. SPIE Optical Engineering 

Magazine, April. 
Chan, Goldberg, Der and Nasrabadi (2002) Dual band imagery improves detection of military target. 

SPIE Optical Engineering Magazine, April. 
Kopp, C. (1994) The Sidewinder story, the evolution of the AIM-9 Sidewinder. Australian Aviation, 

Petrie, G. and Buruksolik, G. (2001) Recent developments in airborne infra-red sensors. Geo Informatics, 


6 Electronic Warfare 

6.1 Introduction 

Warfare has always been conducted by adversaries who have been at great pains to 
understand their enemy's strengths and weaknesses in order to minimise the risk to their 
own forces and territory. The detection and interception of messages and the efforts to 
deceive the enemy have long been the task of the 'secret service'. The military aircraft in its 
infancy in World War I was used to detect troop movements and observe enemy movements, 
while on the ground the use of radio interception confirmed the aerial observations. As 
methods of communication developed, so too did methods of interception become more 
effective. Radar has developed from a mere detection mechanism to a means of surveillance 
and guidance. 

Modern warfare is conducted in a rich electromagnetic environment with radio commu- 
nications and radar signals from many sources. Figure 6.1 shows an example military 
situation with combined land, sea and air forces operating against an enemy territory which 
is, in turn, being defended by similar forces. The key players in this example include the 

1 . Military planning maintains communications with all forces either from the battlefield or 
from staff headquarters. Communications needs to be swift and secure at all times to 
include information from tactical units, from cooperating forces and from analysis 
databases. This communications network is vital to build an understanding of the tactical 
situation and to ensure that orders are received and placed, in the secure knowledge that 
the status and disposition of own forces is not disclosed to the enemy. 

2. Air defences will be using radar to detect incoming airborne threats and will be making 
full use of available intelligence received by land line or data link. They will also be 
issuing orders by radio to fighter, missile and artillery defence systems. 

3. Air superiority aircraft will be on quick reaction alert on dispersal or loitering on combat 
air patrol (CAP). They will be in constant radio communication and using their radar 
discretely to identify targets. 

Military Avionics Systems Ian Moir and Allan G. Seabridge 
© 2006 John Wiley & Sons, Ltd. ISBN: 0-470-01632-9 





]/ UUi'SiiWlIf/ 

Figure 6.1 Typical battlespace scenario. 

4. Defence suppression may be using radar for terrain following or for seeking targets. 

5. Maritime operations in the form of rotary- or fixed- wing units will be conducting open 
ocean or sonar searches to locate and identify surface units or submarines. This will 
involve the use of radar and passive or active sonar, with intelligence sent to headquarters 
(HQ) by data link. Communication with other units will be controlled by use of high 
frequency (HF), very high frequency (VHF), ultrahigh frequency (UHF), shortwave 
marine band or data link. A radar altimeter enables the aircraft autopilot to maintain an 
accurate height over the sea surface regardless of changing atmospheric conditions. 

6. Offensive operations will be using radar for detecting targets, and launched radar guided 
missiles will also be emitting. 

7. Naval forces will be conducting their own operations in close cooperation with their own 
naval and marine forces. This will include the use of surveillance radar, self-defence radar 
and communications. Their communications include marine band shortwave for com- 
munication with merchant vessels or fisheries vessels, as well as very low frequency 
(VLF) for communicating with submarines. 

8. Land forces will be similarly employed with their own units and deploying a wide range 
of radar and communications system. 

As if this situation is not complex enough, modern warfare attracts the attention of the media 
with their attendant TV and sound satellite links and mobile telephone traffic. 

The radio-frequency spectrum covered by the emitters used by these forces is broad, as 
illustrated in Figure 6.2. No single item of equipment can cover this range for transmission or 
reception. Hence, most communications and radar systems are designed for use in specific 
bands. These bands are usually designated by international convention, as detailed in 
Chapter 7. 

The main role of electronic warfare is to search these radio-frequency bands in order to 
gather information that can be used by intelligence analysts or by front-line operators. The 









Navigation Aid 

HFl- VHF Spectrum I 

1 I 

J. i 

t & wirin 

g resonant frequencies 

HF Commun 


I G 







| Altimeter 

fJme I Weath ^ r 



HF frond 

UL» P * ' 


HF Commun cations 



UHF 1 Band 


Ground Based 

10kHz 100kHz 1MHz 

Figure 6.2 Radio-frequency spectrum. 





information gained may be put to immediate effect to gain a tactical advantage on the 
battlefield; it may be used to picture the strategic scenario in peace time, in transition to war, 
or during a conflict. It may also be used to devise countermeasures to avoid a direct threat or 
to deny communications to an enemy. It must also be observed that such tactics are deployed 
by all sides in a conflict - in other words, the listeners are themselves being listened to. 

The drive for intelligence is derived from a continuous need to be one step ahead of any 
potential adversary at all times - in peacetime, in transition to war, during actual conflict and 
in post-war peacekeeping operations. A typical cycle of intelligence is shown in Figure 6.3. 

Changing needs 



dissemination to 
trusted allies 

Surveillance platforms 
Co-operative exchange 



Figure 6.3 Intelligence cycle. 



The cycle of intelligence begins with a requirement to gather information on a particular 
scenario. This may be tactical - the observation of a conflict - or it may be strategic - 
observation of a potential adversary's build-up of forces, their disposition and strength, and 
to identify new assets in the enemy inventory. 

This requirement leads to a set of orders to collect information. This may be by means that 
include land, sea, air and space platforms, and is backed up by background information and 
espionage. Nations will also exchange information, although usually selectively. Raw 
information is analysed to identify new information or changes from previous intelligence. 
It is collated with other sources and with historical data. It is validated for accuracy and 
reliability by comparison with other intelligence and by other sources. It has now become 
'intelligence' and is disseminated by secure means to trusted users. 

Tactical users will make use of the intelligence to modify their battle plans and tactics. 
The intelligence may result in changes to the original requirement, and political situations 
may result in changing needs. Thus, new direction will be provided to the collectors of 

As well as obtaining intelligence, military forces use electronic warfare actively to evade 
detection and to pursue aggressive attacks on enemy radar-guided weapons. Figure 6.4 
illustrates some aspects of electronic warfare broken down into major subdivisions which 
will be described below. 

In addition to all other forces in the electronic warfare (EW) scenario, the Air Forces play 
their own role. Figure 6.5 shows the high-flying EW aircraft gathering and analysing signals, 
and the low-flying tactical EW aircraft accompanying strike forces to counteract enemy 
defences. The high-flying surveillance platform is equipped with a range of sensors and 
receivers to cover the broad range of emitting systems on the ground and in the air. A vast 
amount of data is collected and analysed in real time to provide information of use to forces 
on the ground, and to provide a basis for intelligence to be used in the longer term. The low- 


Signals Intelligence 











Electronic Counter 


Electronic Counter 

Counter Measures 



J amming 


Anti ESM 



Figure 6.4 Electronic warfare elements. 



Figure 6.5 EW airborne roles. 

flying aircraft is often equipped with a more selective range of sensors to identify and attack 
specific targets. 

The aircraft type and the sensors and mission systems selected for these aircraft are 
determined by the requirement to perform strategic or tactical electronic warfare, and to 
obtain the appropriate intelligence. This requirement can be derived from analysis of a top- 
level national requirement such as 'defence of the realm'. This can be progressively 
decomposed or broken down into subsets of requirements which lead to the definition of 
a particular role (Price, 2005). 

6.2 Signals Intelligence (SIGINT) 

Intelligence is collected from a number of different sources to form a strategic picture. These 
sources include: 

• ELINT or electronic intelligence; 

• COMINT or communications intelligence. 

Confirmation of electronic warfare intelligence is usually performed by comparison with 
local information collection and photographic evidence, including: 

• HUMINT or human intelligence; 

• IMINT or image intelligence; 

• PHOTINT or photographic intelligence. 

The first two items on this list are often gathered by high-flying EW aircraft on long duration 
patrols, usually flying a patrol on the friendly side of a border and beyond missile 



Other Aircraft 

Missile Guidance Radar, 

Ground based 



Missile site Radar 

Forward Command 

Post or mobile 

missile site 

Figure 6.6 Users of radar systems. 

engagement range. The aircraft is often a converted commercial type providing accommoda- 
tion for a flight crew and a mission crew of operators able to detect, locate and identify 
sources of radio-frequency emissions at very long ranges. Their task is a combination of 
routine gathering and identification together with an ability to spot new or unusual emitters 
or patterns of use. This they perform with their own knowledge and experience and by using 
databases of known emitter characteristics. This is reinforced by the ability to communicate 
with external agencies to obtain further information. 

The intelligence obtained from analysis of this electronic information is complemented by 
human intelligence in the field and by photographic intelligence which is used to confirm the 
existence and precise locations and types of target. 

6.2.1 Electronic Intelligence (ELINT) 

Figure 6.6 shows some examples of radar emitters, or systems operating in the radar bands 
that are likely to be of interest to ELINT collection aircraft. These include: 

• Ground-based surveillance radars scanning borders looking for airborne or land-based 
intruders and forming a defensive security screen; 

• Missile site or anti-aircraft artillery (AAA) radars scanning for threats and preparing to 
lock on and to track targets for directing defensive weapons; 

• Forward command post radars providing advanced and localised warning of intrusion in 
order to direct local defences; 

• Land and naval forces operating their own radar systems for detection and target tracking; 

• Other fixed-wing aircraft and helicopters operating with their own characteristic radar 

The ELINT system must provide a wide area coverage, preferably as near to spherical 
as possible with few 'shadows' as may be caused by wing tips, fin or fuselage masking. 
Figure 6.7 illustrates the functions to be performed by an ELINT system. The antennas are 



Spherical coverage 
Sensitive detection 
Direction of arrival 

Transmit Intelligence 
Receive Intelligence 



Aircraft position 
Route determination 




Data base 

Figure 6.7 Functional overview of an ELINT system. 

located on the aircraft to provide suitable coverage of the scenario to be monitored and detect 
an arriving signal and its direction of arrival (Do A). The signal is analysed to identify the 
source and its DoA, and to scan intelligence received from other sources to try to confirm the 
signal source. This is fused with the aircraft navigational data so that a picture can be 
provided showing the source relative to the ELINT aircraft. The crew will interpret the 
information and provide the information to other operators. An example system block 
diagram is shown in Figure 6.8. 




Navigation System 

Figure 6.8 Typical ELINT block diagram. 


This system receives signals via a number of antennas situated on the aircraft to provide 
maximum spherical coverage. These antennas are connected to preamplifiers or amplifiers by 
appropriate cables. This may include the routing of low-loss coaxial cables from wing-tip 
antennas to fuselage-installed receivers and amplifiers. The signals are processed by the 
mission computer to add labels or colour for ease of identification. 

Operators are able to ask for further signal analysis to extract key signal characteristics, 
and may also ask for comparisons to be made with similar signals held in a database. With 
this analysis it is possible to identify the type of transmitter, which may enable identification 
of the type of installation or vehicle that made the original transmission. 

An aircraft with a large number of operators can process many signals and is able to build 
up a picture of emitters over a wide area. Each operator will deal with signals from a 
particular band, logging each signal on receipt. The operator's workstation is equipped with 
a roll-ball and keyboard, or a touch screen, which allows the operator to annotate the signals, 
call up analysis or database checks and to store signals. The tactical commander is able to 
retrieve the received and processed signals and build up a composite picture. 

The operators and the pilots work as a team to capture the best possible picture of the 
signal environment. Data link communication allows ground stations or operational 
commanders to join the team and to use other sources of intelligence to direct a specific 
search. The identified emitter remains in the real-time display, tagged as friendly or hostile 
with its characteristics. 

The system can be used to identify radar signals from many sources, including: 

• Fixed ground or airfield radar; 

• Mobile missile battery radar; 

• Ship radar; 

• Aircraft radar; 

• Missile radar; 

• Submarine radar. 

Skilled analysis and comparison with the intelligence database entries enables users to 
identify threat types by radar type, vehicle class and sometimes individual vehicle, especially 
ships where the number of high-value assets is small. 

There will be highly classified threats that will need specific antennas and analysis 
techniques for identification. This situation arises as national security agencies develop new 
transmitters using different bands or different countermeasure techniques to avoid detection. 
This will always be a continuous activity in electronic warfare. 

The database of historical intelligence, the flight plan and tactics and the collected 
intelligence would be of value to an enemy if the aircraft were to be forced to land or if it 
were destroyed. The data storage devices must not be captured intact, and for this reason are 
usually fitted with an explosive charge to ensure complete destruction. 

6.2.2 Communications Intelligence (COMINT) 

Figure 6.9 depicts some examples of users of communications. These users employ bands 
that are mandated for peacetime use such as VHF and UHF for air traffic control or shipping 
lanes, as well as satellite communications and data link for long-range encrypted data 
communications. All forces use a variety of frequency bands. Also shown in this figure is the 



Other Aircraft 

Other Forces 

Command HQ Distress 


Figure 6.9 Users of communications. 

Command Post 

unauthorised listener - the electronic warfare listeners of all participants in a conflict, as well 
as those agencies not directly involved but who want to gather more intelligence. It should be 
noted that this activity also takes place in peacetime and may include listening to friends as 
well as enemies. 

Communications intelligence (COMINT) is gathered by scanning the normal commu- 
nications frequency bands and locking on to detected transmissions. In peacetime it may be 
possible to receive in clear speech, but this is extremely unlikely in times of tension or during 
conflict. However, a great deal of intelligence can be obtained from the following 
characteristics of communications activity: 

• The location of individual transmitters; 

• The locations of groups of transmitters and the numbers in the groups; 

• The frequency of the transmission carrier; 

Provide protection 
Counter threat 

Spherical coverage 

Sensitive detection ih Detect 

Direction of arrival ^^BThreat 

Transmit Intelligence 

Receive Intelligence 


Aircraft position 
Route determination 

Figure 6.10 Functional view of a COMINT system. 





Signal Analyser 

Mission Data 




Data Link 

Navigation System 

Figure 6.11 Typical COMINT block diagram. 

• The style of the operator; 

• The relative intensity of messages; 

• Intervals between message groups; 

• Periods of silence; 

• Periods of activity, especially sudden or unusual activity; 

• Overall pattern of communications during various states of force readiness. 

For these reasons the information-gathering aircraft attempts to obtain a position fix on a 
transmitter and records the activity for later analysis. Depending upon the communications 
frequency, it is not always possible to obtain an accurate fix on a particular platform within a 
task group. However, observation over a period of time allows an overall picture to be built 
up regarding a potential foe's force structure and intended electronic order of battle (EoB). 
The antennas and receivers are optimised for reception over a broad band with high-gain, 
high-noise rejection. An example COMINT system block diagram is shown in Figure 6.11. 
Over a period of several years a potential foe's electronic communications are collected, 
analysed and catalogued in order that both normal (peacetime) and abnormal (high-readiness 
states) may be recognised and understood. This enables responding forces to respond in kind 
by elevating readiness states and, if necessary, imposing restrictions of critical emissions. 
The overall effect is akin to an electronic form of 'cat and mouse', with neither side hoping 
inadvertently to disclose their readiness state or possible future intentions. Large-scale 
peacetime training exercises provide major intelligence-gathering opportunities. 

6.3 Electronic Support Measures 

Information on immediate threats is gathered by an electronic support Measures (ESM) 
system. This consists of a collection of sensitive antennas designed to detect signals in 
different frequency bands. The antennas are often grouped in a wing-tip pod. This allows a 
wide angle of view without obscuration by the fuselage, and also enables a fix on the signal 



Figure 6.12 Wing-tip ESM pod installation - Nimrod MR2. 

source to obtain an accurate direction of arrival (Do A) of the signal. Figure 6.12 shows the 
ESM pods mounted on the Nimrod MR 2 wing tips. 

An effective ESM system rapidly identifies the signal band and location and determines 
the signal characteristics depending upon a number of discriminators. A signal analyser then 
examines the signal characteristics to identify the type of transmitter and the level of threat 
posed. Even the most cursory of analysis can establish whether the emitter is associated with 
surveillance, target tracking or target engagement. This analysis can compare the signal with 
known emitter characteristics obtained from an intelligence database or threat library and 
known signal types confirmed and new emissions identified and categorised. Every signal 
identification is logged with date, time and intercept coordinates, along with the known or 
suspected platform type, and the results are stored. A typical block diagram is shown in 
Figure 6.13. 




Signal Analyser 

Mission Data 





Data Link 

Navigation System 

Figure 6.13 Example ESM block diagram. 



Time of arrival 

Pulse width 



Example pulse shape 

Figure 6.14 Important ESM parameters. 

of arrival 

Scan rate 
Scan pattern 

Signals received by the electronic support measures system may in some cases be analysed 
instantaneously to produce an identity for the transmitter of each signal received. Figure 6.14 
shows some parameters of a signal that are essential for the understanding of the type of 
transmitter producing the signal. The nature of the pulse shape is used to determine the 
particular type of transmitter. The scan rate and the pattern of the scan also provide invaluable 
information about the mode of the transmitter. It is possible to detect the antennas changing 
from scanning mode to lock-on to tracking and hence determine the threat that the transmitting 
station poses. 

As well as providing threat information, ESM is used by maritime and battlefield 
surveillance aircraft as a passive or listening sensor which adds important information to 
other sensors. It is especially useful when tracking submarines, where the use of the aircraft 
radar would be a source of intelligence to the submarine commander. 

The salient signal characteristics or discriminators identified during the ESM collection 
and identification process include the following: 

1. Signal frequency. Owing to the RF atmospheric propagation and transmission character- 
istics, the operating frequency is the first indicator of radar type as all RF emitters have to 
compete in the same physical world. 

2. Blip/scan ratio. Examination of the blip/scan ratio will give preliminary indications of 
scan rate, sector scan width and possibly radar/emitter beamwidth. 

3. Scan rate. The higher the scan rate, then generally the more likely is the threat of 

4. Scan pattern. Search, track, track- while- scan (TWS) and ground-mapping (GM) modes 
will exhibit particular characteristics. 

5. Signal modulation. Pulse, pulse compression, pulsed Doppler (PD), a continuous wave 
(CW) and other more sophisticated forms of modulation are indicative of the emitter 
mode(s) of operation and likely threat level. 

6. Pulse repetition frequency (PRF). High PRF associated with a tracking mode signifies an 
imminent engagement. 

The combination of analysis of all these modes of operation and when they are employed 
either singly or in combination is vital to establishing the likely capabilities and intentions 


of a threat platform, especially when used in combination with other intelligence 

ESM may be employed at a strategic intelligence-gathering level using an AWACS or 
MPA aircraft to build the overall intelligence picture and electronic order of battle (EOB). 
Alternatively, such information may be gathered and utilised at a tactical level using radar 
warning receivers (RWR), whereby information is gathered and used at the strike platform 
level to enable strike aircraft to avoid the most heavily defended enemy complexes during 
the mission. 

6.4 Electronic Countermeasures and Counter-countermeasures 

Electronic countermeasures (ECM) and electronic counter-countermeasures (ECCM) take 
the form of interfering or deceiving the enemy's radio and radar systems in order to negate 
their use or, worst of all, compromise their performance. On occasions it is difficult to 
distinguish between 'chicken' and 'egg' as so many issues are considered during the design 
phase and then hastily need to be re-evaluated once a real conflict begins. 

Therefore, the authors have chosen to consider these issues together rather than separately, 
as it is indeed a rapidly evolving process. The deployment of EW and successes and failures 
are invariably and rapidly recognised as the conflict develops and as both sides are inclined 
to receive untimely and unexpected unpleasant surprises. In some cases this is due to an 
inexact appreciation of the capabilities of the foe or where ELINT has not been able to 
provide the complete picture. Also, given the frailties of humankind, there is also a tendency 
to 'lose the recipe' and to relearn the hard way the lessons derived from a previous conflict. 

These countermeasures, or 'jamming' as they are often loosely called, may be divided into 
two categories: 

• Noise jamming; 

• Deception jamming; 

6.4. 1 Noise Jamming 

Active noise jamming is often performed by identifying an enemy detection system and 
broadcasting white noise at high power levels. For communication systems, noise jamming 
could employ the broadcast of music or other audio features designed to deny the use of the 
particular service. This effectively swamps the input circuitry of detection systems and 
prevents it from operating. 

The effectiveness of a jamming system depends on a number of aspects of the system, for 

Transmitter power output; 

Transmission line losses between the transmitter and the radiating antenna; 

Antenna gain in the direction of the receiver to be jammed; 

• Transmitter bandwidth. 

The amount of energy delivered into a target transmitter depends on similar aspects of the 
target such as: 



• Incoming jamming power; 

• Receiver bandwidth; 

• Antenna gain; 

• Radar cross-sectional area of the aircraft being masked. 

In order to be effective, the jamming transmitter must be able to emit sufficient power in the 
bandwidth of the target receiver to mask its intended signal or to simulate a deceptive signal 

Most jamming transmitters are designed to operate over the range of frequencies expected, 
and, as has been shown above, this is extremely wide given the range of communications 
devices, search radar and guidance radar types to be found on the modern battlefield. In the 
case of radar, the bandwidth covered is often in the range 2-18 GHz. The jammer is most 
effective if it can be designed to target a specific frequency range or type of threat, in which 
case the power output is concentrated into a narrow spectrum. Given that a jammer must 
operate against a wide range of emitters, its power must be spread over an increased 
spectrum (Figure 6.15). 

Figure 6.15 portrays a frequency spectrum in which four targets exist; the jammer has the 
task of nullifying each of the four targets. Three different techniques are shown: 



Target 1 Target 2 Target 3 Target 4 

Target 1 

Swept Spot | 

s~\ ^\ 

A L 

Target 2 

Target 3 

Target 4 

Target 1 

Multiple Spot 

Target 2 

Target 4 

Figure 6.15 Rudimentary jamming techniques. 



Radar beam sweep Jamming signal 

Received at radar 




Figure 6.16 Inverse noise gain jamming. 

1. Barrage jamming. In this example, jamming power is spread across the entire spectrum 
encompassing the targets. This results in a very low jamming power density (W/MHz) to 
the point that none of the targets is adversely affected. 

2. Swept spot jamming. Swept spot jamming concentrates sufficient power in a narrow 
bandwidth to negate each target. The jammer switches to each of the targets in turn but is 
only present for a low-duty cycle. This may suffice if the target receiver saturation and 
automatic gain control capabilities are modest but will not suffice for higher-performance 

3. Multiple-Spot Jamming. Multiple- spot jamming divides the energy between the targets, 
effectively jamming them in parallel rather than sequentially. This requires a more 
sophisticated jamming transmitter. 

The foregoing explanation is in itself very superficial. In reality the radar is unlikely to be 
transmitting continuously on a fixed frequency; modern radars have a considerable degree of 
frequency agility and can often change frequency and even signal modulation on a pulse-by- 
pulse basis. This makes noise jamming more difficult to achieve effectively. 

These techniques are rudimentary and are not particularly effective except against the 
most primitive equipment and more sophisticated techniques may be employed. 

Figure 6.16 illustrates the principle of inverse noise gain jamming. The target signal is 
analysed and a pattern of noise is generated on time that complements the original incoming 
signal. This results in a return signal received at the target radar that is a continuous noise 
pattern, thereby masking the return from the aircraft skin. With high power this can be used 
to swamp the return, thus denying the enemy range information. With even higher powers it 
is possible to enter the sidelobes of the threat radar to deny angle information. Bur nth rough 

Burnthrough range is the range at which the strength of the radar echo becomes greater than 
that of the jamming noise. The radar return is proportional to 1/R 4 since it must travel to the 
target and return to the host radar. The jamming signal only travels in one direction, and 
is thus proportional to l/R 2 . The more closely an aircraft approaches the victim radar source, 
the more likely is the radar signal to break through the jamming noise (see Figure 6.17 which 
illustrates the principle). 

In Figure 6.17 a plot is shown comparing the received power (dB) against range in nautical 
miles, and the effect of 1/R 2 and 1/R 4 for jammer and radar respectively can be clearly seen. 




Range R 


\ Burn -Through ~8to12dB 


Jamming oc 1 / 

^ 2 

' *~ ■ — Target Return oc 1 


Range (nm) 

Figure 6.17 Effect of burnthrough. 

However, at some point close to the radar, the target return signal will exceed the jamming 
signal by a suitable margin and the radar will prevail. The threshold associated with 
burnthrough is generally assumed to be of the order of 8-12 dB. At ranges greater than 
this the jammer has the advantage. 

This balance depends upon a multitude of factors including the relative performances of 
jammer and radar transmitter and receivers, the antenna gain and sidelobe characteristics, 
the aspect of the engagement, etc. A radar antenna with low gain or poor sidelobe per- 
formance will be vulnerable to clutter, as already described in Chapter 4, and noise jamming 
is in effect a man-made form of clutter. Conversely, the higher the performance of a radar 
and the better the ability to discriminate against clutter, the more robust it will be in a 
jamming scenario. 

Another significant disadvantage of noise jamming as a countermeasure extends to the 
jamming platform itself. By virtue of transmitting relatively high power, the jammer itself 
becomes a beacon whereupon the foe can use the jamming emissions as a source of 
guidance. Hence, many modern systems have a home-on-jam (HO J) mode to enable the 
jammer itself to be attacked while radiating. 

6.4.2 Deception Jamming 

Deception jamming employs more sophisticated techniques to negate the performance of the 
radar. If subtly employed, the radar and radar operator may not realise that countermeasures 
are being used. Some typical techniques used to break the radar-tracking loops previously 
described in Chapter 4 are: 

1. False target generation. If the modulating characteristics of the target radar are known, it 
is possible to transmit pulses that will appear as multiple targets in the victim radar. 


Hence, by using the jamming transmitter with diligence and transmitting replica pulses 
after a time delay, these false, multiple, spurious targets will appear in subsequent radar 
range sweeps. An intelligent radar operator should realise that his radar is being deceived 
but may have a problem in deciding which of the multiple returns is the correct one. 

2. Range gate stealing. This is a variation on the technique described above where one false 
pulse is generated that appears in the victim's radar at the same range as the jammer. It is 
then possible to capture the range gate with the artificial pulse; in particular, if the false 
pulse appears to be stronger than the original in the victim receiver, it is possible to 'steal' 
the range gate by progressively altering the false range. If desired, the range gate may be 
left on a false value or moved off to coincide with clutter, whereupon the target lock will 
be lost. 

3. Angle track breaking. Similarly, there are ways of breaking the angle track mechanism, 
especially if the tracking mechanism of the victim radar is well understood. For example, 
in a conscan radar, angle track may be broken if the jamming signal is modulated at a 
frequency that approaches that of the conscan modulation frequency of the subject radar. 
This presupposes that the angle tracking method and conscan rate are known, which may 
not be the case in a wartime situation. Other simple ways of angle deception include 
terrain bounce, cross-eye, cross-polarisation and double cross. 

4. Velocity gate stealing. This is similar to range gate stealing except that the incident signal 
is re-radiated back to the victim radar, initially without modification. Progressively the re- 
radiated signal is amplified and masks the original Doppler component upon which the 
velocity gate is centred. The deceiving radar may then steal the velocity gate in a similar 
manner to the range gate stealer described above. 

Modern radars are inherently resistant - although not impervious - to jamming owing to a 
range of features inherent in the design. These characteristics are as follows: 

• Low antenna sidelobes; 

• Wide dynamic range with fast-acting automatic gain control (AGC); 

• Constant false alarm rate (CFAR) reduction; 

• Sidelobe blanking. 

When these features are employed together with a range of other technology advances that 
evolved throughout the late 1980s and early 1990s, including greatly increased RF 
bandwidth, sensor fusion and the application of artificial intelligence techniques, then 
significant advances may be achieved. These developments have all contributed towards 
greatly enhanced radar performance. These techniques are outside the scope of this book and 
in many cases are classified. 

6.4.3 Deployment of the Jamming Platform 

The airborne jamming assets may deployed in two possible ways: 

1. Self- screening platforms with their own on-board EW suite. The complexity and intensity 
of the modern battlefield is such that most platforms carry their own protection suite, also 
sometimes referred to in US parlance as aircraft survivability equipment (ASE). 


2. Escort or stand-off jammers with a specialised EW role. The escort jamming role has been 
provided in the past by aircraft such as the F-4 Wild Weasel and EA-6 Prowler. Recently, 
the F-16C/J has taken on this role for the US Air Force and the F-18E/F is under 
development to replace the Prowler in the near future with the EF-18G. Such aircraft may 
also perform a stand-off jamming role, although this may also be performed by platforms 
with lower performance. 

In reality, a modern conflict depends much upon the blending and merging of both asset 
types, depending upon the nature of the engagement. Jamming assets also offer comple- 
mentary assistance to stealth platforms where they are deployed as low observability is easier 
to maintain in an aggressive EW environment, which has proved to be the case in recent 
Kosovo and Iraq engagements. 

Escort jammers that accompany the main force are often referred to in US Air Force 
parlance as Wild Weasel squadrons and comprise strike aircraft types modified to perform a 
dedicated EW support and suppression role. The task is to precede or accompany the strike 
force, selectively jamming and confusing enemy defence radar and communications. These 
aircraft may also be armed with anti-radiation missiles (ARMs) that use the threat radar 
beam to guide themselves to the radar. 

Aircraft operating in support of an attack force may also station themselves in a stand- 
off position outside the range of ground defences while maintaining a patrol so that an 
appropriate noise jamming signal can be used to confuse defences. Care must be taken 
that the jamming supplements and does not diminish the effectiveness of the attacking 

6AA Low Probability of Intercept (LPI) Radar 

All these countermeasures depend upon the detection of the victim radars ' emissions and 
upon having some prior knowledge of the frequency of operation and modulation techniques 
employed. The most obvious counter of all is to avoid detection as far as possible by utilising 
low probability of intercept (LPI) techniques. LPI techniques must be designed into the radar 
at the outset and involve a number of trade-offs where increasingly sophisticated design (and 
cost) is balanced against a lower probability of interception. Some of the design considera- 
tions include the following: 

1 . A reduction in peak power and an increase in the period of integration will result in the 
same overall detection capability for reduced peak power. 

2. An increase in receiver bandwidth using spread spectrum techniques and a reduction in 
peak power, effectively spreading the modulation data across a wider band, will make the 
task more difficult for the jammer. 

3. The radar has a much higher gain than that of a radar warning receiver (RWR) antenna 
and, while potentially disadvantageous during transmit, it has significant advantages 
during receive. Balancing peak power against antenna gain can yield benefits, and the aim 
is usually to increase antenna gain while reducing peak power. For effective LPI radars a 
design aim is to achieve a main beam gain of +55 dB above the first sidelobes. Other 
considerations include a high-duty cycle reducing peak power, low receiver losses and a 
low receiver noise factor. 


6.5 Defensive Aids 


An aircraft operating in a hostile military environment needs to be equipped with measures 
for self-defence. The crew will have been briefed on the threats on their outward and return 
transits, as well as enemy defences in the area to be attacked or where an engagement is to 
take place. This will be based on intelligence and will be in accordance with the most up-to- 
date intelligence compilation. 

However, during the mission, the pilot must be warned of real tactical threats to the 
mission, and must have the means to minimise their effectiveness. The most common threats 
to low-flying aircraft are: 

• Small Arms fire; 

• Radar-guided anti-aircraft artillery (AAA or triple- A). 

• Shoulder-launched surface-to-air missiles (SAM); 

• SAM from ground sites, vehicles or ships. 

Appropriate countermeasures include a means of detecting the threat and luring the threat 
away from the aircraft or causing the missile to detonate prematurely or far enough away 
from the aircraft so that no damage is sustained. This combination of sensor and counter- 
measure is often referred to as a defensive aids subsystem or DASS, and often abbreviated to 
Def-Aids. Figure 6.18 shows an aircraft equipped with a set of threat detectors and 
countermeasure subsystems. 

There is little that can be done by a defensive aids system to have a significant impact on 
small arms fire and AAA, although counters may be devised for AAA gun tracking systems. 
High-speed evasive manoeuvres on the low-level run in to the target and may be firing a gun 
at ground sites may be a palliative, but the risk of a hit remains. Most aircraft are designed to 
minimise the catastrophic effects of missile or shell fragments by physical separation of 
critical equipment and wiring to reduce the probability of common mode damage effects. For 
those weapons employing active sensing there are mechanisms for reducing their effectiveness. 

To counter weapons or systems utilising some form of electronic system or guidance, a 
defensive aids subsystem may include any or all of the following subsystems, depending 
upon the role and the intensity of the threat: 

Typhoon DASS pods 
Sensors & Towed 

Figure 6.18 Example of an aircraft equipped with a DASS. 








eS 6 

Up to 

■ Sensors ■ 

Figure 6.19 Functional layout of the radar warning receiver. 

• Radar warning receiver; 

• Missile warning receiver; 

• Laser warning receiver; 

• Countermeasure dispenser (CMD) - chaff or flares; 

• Towed decoy. 

On a military aircraft these systems will have the capability of interfacing with the aircraft/ 
mission avionics system by using MIL-STD-1553B data buses or other cost-effective 
commercial data bus equivalents. 

6.5.1 Radar Warning Receiver 

A typical radar warning receiver (RWR) is depicted in Figure 6.19. Sensors are located 
strategically around the peripheries of the aircraft - typically four sensors placed at the wing 
tips or sometimes at the top of the fin. The objective, as far as is possible, is to provide full- 
hemisphere horizontal coverage around the aircraft in order that the crew may detect and be 
alerted to potential RF threats. Each of these sensors may provide up to 90° conical coverage, 
although in some cases the angular reach may be less than this. A typical antenna used in this 
application would be a spiral antenna with an angular coverage of 75° but with a gain of 
-10 dB. 

This figure should be compared with the 55 dB gain that would be the design point for a 
LPI radar - a difference of 45 dB or a factor of 32,000. This illustrates in part the 
disadvantage that the RWR faces while operating against a modern state-of-the-art radar. 
Other considerations such as the use of sophisticated spread spectrum modulation, radiated 
power management and advanced signal processing indicate why it is conceivably possible 
for a sophisticated AESA radar such as the AN/APG-80 as used on the F-22 to operate 
almost invisibly to some medium-capability RWR equipment. 

The quadrant-located spiral antennas detect and to some extent direction find (DF) any 
emissions within their respective area of coverage. Demodulated signals are analysed by the 


signal processor and categorised against a known threat library according to the following 

• Frequency of operation; 

• Modulation type; 

• Signal strength; 

• Direction of arrival. 

In some cases an audio tone may be derived to provide the pilot or observer with audio cues - 
typically a tone equivalent to the PRF of the incoming radiation. 

In early systems the processed outputs were displayed upon a plan position indicator (PPI) 
in a manner that depicted the angle of arrival according to a clock format with 12 o'clock 
dead ahead. In the late 1960s/early 1970s, when these systems were operationally deployed 
for the first time during the Vietnam War, this information would be presented on standard 
CRT green phosphor displays. Relative signal strength was shown by the length of the line 
from the centre of the clock, while the coding of the line into solid, dashed or dotted 
portrayal was indicative of the modulation type or possibly the band of operation. Early 
systems such as the air radio installation (ARI) 18228, as employed on the UK F-4K/M 
Phantom, used a hard-wired implementation to code specific threats and were therefore 
cumbersome to reprogramme. 

The advent of digital processors now means that the threat library is coded in software 
allowing for rapid updates using a suitable software loading device. On a modern system, 
particularly since the advent of AMLCD colour displays, display symbology is much more 
likely to utilise stylised colour-coded symbology which is much more easily recognised by 
the pilot or radar operator, especially in a stressful combat environment. 

Typical frequency coverage of a RWR system extends from 2 to 18 GHz and embraces a 
wide range of electronic threats across the RF spectrum. Modern RWR equipment offers a 
much more dynamic response to specific threats than was possible with the first-generation 

6.5.2 Missile Warning Receiver 

A missile warning receiver operates on a similar principle, except the missile warning 
systems (MWS) operate by detecting infrared (IR) or ultraviolet (UV) emissions during and 
following a missile launch. A typical system is portrayed in Figure 6.20. 

Although the conical coverage of an IR/UV sensor may be as much as 110°, these systems 
often provide the option of expansion to six rather than four hemispherical sensors - see the 
F-35 example in Chapter 9. Apart from the sensors, in an overall sense the system works in a 
very similar fashion to the RWR above. Threat analysis is undertaken within a central signal 
processor/computing unit, and the results are output to a suitable tactical display. In modern 
systems this will be a colour tactical display. 

An example of a missile warning system is shown in Figure 6.21. 

6.5.3 Laser Warning Systems 

A laser warning system again uses similar principles, except that the sensors are operating in 
the laser band. The example shown covers the 0.5-1.8 urn band and addresses the threat 





q 9 p -i 


I vv| 

~| Controller 


o o 

^— Sensors — ► (Optional) 

Figure 6.20 Missile warning receiver layout. 


posed by the following lasers: doubled NdYAG, ruby, GaAs, NdYAG, Raman shifted 
NdYAG and exbium glass lasers. Angle of arrival (AoA) is claimed to be within 15° rms 
and sensor angles are 110°. (Figure 6.22). 

6.5.4 Countermeasure Dispensers 

The defensive aids will be equipped with a technique generator that interprets the threat and 
defines a suitable defensive response using the following: 

• Chaff; 

• Flares; 

• Towed decoy. 

Figure 6.21 Missile warning system (SAAB Avitron). 



Figure 6.22 Typical laser warning system (SAAB Avitron). Chaff and Flares 

A chaff and flare dispenser is fitted to many aircraft so that appropriate mixes of chaff and 
flares can be selected and deployed to confuse missile seekers or defence radars. This can be 
done by providing alternative decoy targets for seekers, or by disguising the aircraft by 
changing its radar return so that operators cannot set up an aiming solution. Chaff and flares 
are usually deployed as 'last ditch' countermeasures against an incoming missile. Their 
effective protection zone is to the rear of the aircraft, and they offer no protection against 
missile engagements in the forward hemisphere. 

Chaff consists of reflectively coated strips of plastic or metal foil. The strips are designed 
to a half- wavelength (A/2) of a typical homing radar. The chaff can be dispensed in patterns 
or blooms to disguise the dispensing aircraft with a view to confusing a radar operator. When 
released into the turbulent airflow, chaff disperses rapidly (blooms) and for a brief period of 
time generates a very large, static radar image between the target aircraft and the threat radar, 
which is probably in tracking mode or CW illumination mode, providing radar guidance for 
an in-flight missile. If the timing is right, radar or missile lock may be broken. 

Flares are a countermeasure against IR homing missiles. When deployed, a flare burns 
with an IR wavelength similar to that of the target aircraft IR signature. It works by initially 
appearing in the missile seeker head coincident with the target aircraft, but is left behind as 
the target aircraft performs evasive manoeuvres. Its thermal image is designed to be longer 
than the target aircraft and it then becomes the preferred IR target for the missile. 

Timing of deployment is critical. Too soon and the divergence of target aircraft and flare 
will be detected and the flare ignored, too late and the missile will detonate on the flare and 
fragments may hit the target anyway. 

Flare deployment can be used in a 'saturation' mode during periods of extremely high 
risk, where the target aircraft is in very close proximity to a missile launcher and has no time 
to manoeuvre if missile launch is detected, for example, transport aircraft carrying out low- 
level air drops or landing on captured airfield in hostile territory where MANPAD or Stinger 
IR missiles may be launched within a few hundred feet of the aircraft. In these circumstances 
the crew may choose to pre-empt target launch detection by the tactical deployment of 
multiple flares in the high-risk zone. 



Figure 6.23 Example of flare dispensing. 

Figure 6.23 shows an example of a C-130 Hercules deploying multiple flares. Towed Decoy 

The towed decoy is essentially a heat source that is towed behind the target aircraft on a 
cable. Located in a wing-tip pod, a wing-mounted pylon pod or deployed from inside the 
aircraft, the towed decoy is released and extended on a cable which restrains the device and 
provides a source of electrical power. The purpose is to cause infrared seekers in missiles to 
home on to the decoy rather than the jet-pipes of the towing aircraft. Any explosion should 
be sufficiently distant so as not to cause damage from the explosion or from missile 
fragments. The decoy can be rewound or it may be jettisoned by cutting the cable if there is a 
failure of the rewind mechanism. An example is shown in Figure 6.24. 

Towed Decoys 

Fibre optic aerial decoy 

Incoming missile homes on 
the de coy 

Figure 6.24 Example of a towed radar decoy. 




AN/APG-79 AESA Radar 
AN-APG 67 RWR Suite 

AN/ALQ-214 EW 
Countermeasures Suite 


AN/ALQ-214 EW 




Figure 6.25 Simplified overview of F/A-18E/F countermeasures suite. 

6.5.5 Integrated Defensive Aids Systems 

In order to convey the complexity and extent of self- screening EW systems on modern 
combat aircraft, the example of the F/A-18E/F Super Hornet will be briefly analysed. This 
aircraft has the following countermeasures suite fit: 

• AN/APG-79 AESA radar; 

• AN/ALR-67 radar warning receiver; 

• AN/ALQ-214 integrated defensive electronic countermeasures (IDECM); 

• AN/ALE-47 countermeasures dispenser; 

• ALE-50/55 towed decoy. 

6.5.5. 1 AN/APG-79 AESA Radar 

The Raytheon APG-79 active electronically scanned array radar is an 1100-element radar 
that has all the advantages and flexibility inherent in this type of radar. In particular, 
flexibility of mode of operation, high scan rates, sophisticated modulation and signal 
processing and LPI features give the aircraft significant operating advantages in a hostile 
EW environment. AN/ALR-67 Radar Warning Receiver 

The RWR suite is an integrated suite comprising the following components: 

• Countermeasures computer; 

• Countermeasures receiver; 

• Low-band integrated antenna; 



ALR>67(V)3 System Components 
Figure 6.26 AN/ALR-67 RWR components (Raytheon). 

• 6 x integrated antenna detectors (two low band and four high band); 

• 4 x quadrant receivers. 

The countermeasure receiver receives inputs from the two low-band antennas and from the 
four high-band antennas via the respective quadrant receivers. The quadrant receivers 
provide preconditioning to reduce transmission losses between antenna and receiver. The 
receiver digitises and categorises the received signals and is able to handle a dense pulse 
environment while at the same time handling faint signals from distant threats. 

The countermeasure computer incorporates a 32 bit machine with the application software 
encoded in Ada. The software structure enables complete reprogramming of the master 
threat file without any software changes. (Figure 6.26). The system weighs less than 100 lb - 
well under the normal weight of a system of this kind. AN/ALQ-214 Integrated Defensive Electronic Countermeasures (IDECM) 

The IDECM system is a radio-frequency countermeasures (RFCM) suite comprising the 
following units: 

• Receiver; 

• Processor; 

• Signal conditioning amplifier; 

• Modulator/techniques generator; 

• Two optional plug-in transmitters may also be used. 

The weight of the system, including the rack, is ~168 lb. If the ALE-50 towed decoy option 
is included, a further 541b is added. The fibre-optic towed decoy actually transmits the 
jamming signal according to the top-level architecture shown in Figure 6.27 and the units 
shown in Figure 6.28. 

After the interception of the incoming victim radar signal, the appropriate counter- 
measures are applied and the RF is converted to light energy for transmission down the fibre 
cable to the decoy. The light energy is converted to RF energy and amplified by the travelling 
wave tube (TWT) transmitter. The resulting jamming signal is transmitted to the target radar. 





t Receiver/ ^^^ 
Processor "^^T 

1 I F 

ent LJ P 

Aircraft -ALQ-21 4 

; On-Board !~ , 

| Transmitters •;« 

! (Optional) ! ! 

• • 






ALE-50 Decoy 

Figure 6.27 AN/ALQ-214 concept of operations. AN/ALE-47 Countermeasure Dispenser 

The AN/ALE-47 countermeasure dispenser - a successor to the ALE-39 - is able to dispense 
up to 60 expendables comprising chaff, flares or radar decoys. This is all achieved under 
computer control, enabling the pilot to achieve the optimum mix of expendables and 
deployment sequence for a given threat scenario. 
The system has four main modes of operation: 

1. Automatic. The countermeasure system evaluates the threat data from the on-board EW 
sensors and merges them with stored threat data to determine the optimum dispensed 
stores mix. The system automatically dispenses this countermeasure load. 

Figure 6.28 AN/ALQ-214 units (BAE SYSTEMS). 


2. Semi-automatic. The countermeasure system determines the optimum stores mix as for 
the automatic mode, but the crew activate deployment. 

3. Manual. The crew manually select and initiate one of a number of preselected 

4. Bypass. In the event of a system failure the crew can reconfigure the system in flight. 


Lynch Jr, D. (2004) Introduction to RF STEALTH, SciTech Publishing inc. 

Price, A. (2005) Instruments of Darkness: The History of Electronic Warfare, Greenhill Books. 

Stimson, G.W. (1998) Introduction to Airborne Radar, 2nd edn, Scitech Publishing inc. 

7 Communications 
and Identification 

7.1 Definition of CNI 

All military aircraft need certain computing sensing and computing resources apart from the 
mission sensors and weapons to enable them to complete their mission. These are: 

1. Communications. The ability to be able to communicate by either voice or data link 
means with cooperative forces, be it wingmen in the same flight of aircraft, airborne 
command centre or troops on the ground. 

2. Navigation. The military platform needs to be able to navigate with sufficient accuracy 
to a target, rendezvous point, waypoint, or initial point as dictated by the mission 

3. Identification. The rules of engagement for a given theatre of operation will necessitate 
the classification and identification of a target before permission to engage is given. 

The American military refer to this collection of resources as communications, navigation, 
identification (CNI). 

Some of the CNI sensors such as air data, radar altimeters and inertial systems are 
autonomous to the platform, in other words the platform needs no other input from third- 
party sources. Others such as communications, radio navigation beacons and global 
navigation satellite systems (GNSSs) require the participation of other organisations to 
respond or the provision of a network of aids or a constellation of satellites to provide 
the navigational framework. Military platforms use a combination of all these sensors with 
the additional rider that, for certain covert stages of a mission, no emissions are made by the 
platform as radio silence - more correctly known as emission control (EMCON) - 
procedures are enforced. 

Military Avionics Systems Ian Moir and Allan G. Seabridge 
© 2006 John Wiley & Sons, Ltd. ISBN: 0-470-01632-9 







100 MHz 



100kHz - 
Figure 7.1 CM RF spectrum. 


(~ 4Ghz) 


(Uplink - 1626.5 to 1660.5MHz 

Downlink - 1 530.0 to 1 559.0MHz) 


(1575.42 (L1) & 1227.6 (L2)MHz) 


(1060 Mhz) 

(960 to 1215MHz) 

UHF Comms 

(225 to 400Mhz) 

VHF Comms 

(118 to 156 Mhz) 

Marker Beacons 

(75 Mhz) 

HF Comms 

(3 to 30Mhz) 


(200 to 1600kHz) 

(90 to 110kHz) 

7.1.1 RF Spectrum 

The RF spectrum associated with the CNI functions is shown in Figure 7.1. The 
CNI spectrum covers a range of different equipment spanning almost five decades from 
100 kHz to 4 GHz and comprising a range of functions as described below. Communica- 
tions and identification are addressed within this chapter while navigation is discussed in 
Chapter 8. For ease of reference, the equipment is listed in ascending order of operational 

1. Communications: 

• High-frequency (HF) communications; 

• Very high-frequency (VHF) communications; 

• Ultrahigh-frequency (UHF) communications; 

• Satellite communications (SATCOM); 

Data links. 


2. Identification: 

• Air traffic control (ATC) mode S; 

• Traffic collision and avoidance system (TCAS); 

• Identification friend or foe (IFF). 

With one or two exceptions, this equipment is all freely available for use by the civil 
community as well as by the military platform. All operational frequencies are published on 
aeronautical charts to ensure safe and successful integration and interoperability of all traffic 
within the wider airspace. There are a few exceptions, namely: 

1 . Civil traffic does not usually use the UHF communications band. Military users may also 
use UHF SATCOM which is not widely available. 

2. Civil traffic would not ordinarily be equipped with TACAN. 

3. Certain GPS codes offering more accurate navigation capabilities may be denied to the 
civil user community. 

4. IFF is compatible with ATC modes S but not available to civil users. 

7. 1.2 Communications Control Systems 

The control of the aircraft suite of communications systems, including internal communica- 
tions, has become an increasingly complex task. This task has expanded as aircraft speeds 
and traffic density have increased and the breadth of communication types have expanded. 
The communications control function is increasingly being absorbed into the flight manage- 
ment function as the management of communication type, frequency selection and intended 
aircraft flight path become more interwoven. Now the flight management system can 
automatically select and tune the communications and navigation aids required for a 
particular flight leg, reducing crew workload and allowing the crew to concentrate more 
on managing the on-board systems. 

7.2 RF Propagation 

The number of antennas required on-board an aircraft to handle all the sensors, commu- 
nications and navigation aids is considerable. The CNI aspects of RF systems integration on 
a fighter aircraft have already been described in Chapter 2. 

Civil aircraft adopted for military applications also have a comprehensive CNI antenna 
complement. This is compounded by the fact that many of the key pieces of equipment may 
be replicated in duplicate or triplicate form. This is especially true of VHF, HF, VOR and 
DME equipments. Figure 7.2 shows typical antenna locations on a Boeing 777 aircraft; this 
is indicative of the installation on most civil aircraft operating today, particularly those 
operating transoceanic routes. Owing to their operating characteristics and transmission 
properties, many of these antennas have their own installation criteria. SATCOM antennas 
communicating with satellites will have the antennas mounted on the top of the aircraft so as 
to have the best coverage of the sky. ILS antennas associated with the approach and landing 













Figure 7.2 Typical aircraft CNI antenna (Boeing 777 example). 

phase will be located on the forward, lower side of the fuselage. Others may require 
continuous coverage while the aircraft is manoeuvring and may have antennas located on 
both upper and lower parts of the aircraft; multiple installations are commonplace. In 
addition to these antennas, military aircraft will have additional communications fitted 
commensurate with their military role. 

In aviation, communications between the aircraft and the ground (air traffic/local 
approach/ground handling) have historically been by means of voice communication. 
More recently, data link communications have been introduced owing to their higher data 
rates and in some cases superior operating characteristics. As will be seen, data links are 
becoming widely used in the HF, VHF and UHF bands for basic communications but also to 
provide some of the advanced reporting features required by FANS. In the military 
community, data links have a particular significance in relation to target reporting and the 
sharing of tactical and targeting information, as will be described in the section on network- 
centric operations. The most common methods of signal modulation are: 

1. Amplitude modulation (AM). This type of modulation concentrates the information being 
carried by the transmission in relatively narrow sidebands. AM communications are 
susceptible to noise and jamming. 

2. Frequency modulation (FM). FM modulation is more sophisticated and spreads the 
transmission across a wider frequency spectrum than AM, thereby reducing the vulner- 
ability of the signal to interference and jamming. This technique is generically known as 
spread spectrum modulation and can be used in a number of ways using differing 
modulation techniques. The spread spectrum is a useful adjunct to low probability of 
intercept (LPI) systems where the intention is to make the task of an adversary detecting 
the signals more difficult. 




or Ship 

Figure 7.3 HF communications signal propagation. 

There are many extremely sophisticated methods of signal modulation. In the military 
environment these are used to maximise the performance of the signal under adverse 
operating conditions while minimising the probability of intercept. 

7.2. 1 High Frequency 

High frequency (HF) covers the communications band between 3 and 30 MHz and is a very 
common communications means for land, sea and air. The utilised band is HF S SB/AM over 
the frequency range 2.000-29.999 MHz using a 1 kHz (0.001 MHz) channel spacing. The 
primary advantage of HF communications is that this system offers communication beyond 
the line of sight. This method does, however, suffer from idiosyncrasies with regard to the 
means of signal propagation. 

Figure 7.3 shows that there are two main means of propagation, known as the sky wave 
and the ground wave. 

The sky wave method of propagation relies upon single- or multiple-path bounces between 
the earth and the ionosphere until the signal reaches its intended location. The behaviour of 
the ionosphere is itself greatly affected by radiation falling upon the earth, notably solar 
radiation. Times of high sunspot activity are known adversely to affect the ability of the 
ionosphere as a reflector. It may also be affected by the time of day and other atmospheric 
conditions. The sky wave as a means of propagation may therefore be severely degraded by a 
variety of conditions, occasionally to the point of being unusable. 

The ground wave method of propagation relies upon the ability of the wave to follow 
the curvature of the earth until it reaches its intended destination. As for the sky wave, the 
ground wave may on occasions be adversely affected by atmospheric conditions. Therefore, 
on occasion, HF voice communications may be corrupted and prove unreliable, although HF 
data links are more resistant to these propagation upsets, as described below. 

HF communications are one of the main methods of communicating over long ranges 
between air and ground during oceanic and wilderness crossings when there is no line of 



sight between the aircraft and ground communications stations. For reasons of availability, 
most long-range civil aircraft are equipped with two HF sets with an increasing tendency 
also to use HF data link if polar operations are contemplated. 

HF data link (HFDL) offers an improvement over HF voice communications owing to the 
bit encoding inherent in a data link message format which permits the use of error-correcting 
codes. Furthermore, the use of more advanced modulation and frequency management 
techniques allows the data link to perform in propagation conditions where HF voice would 
be unusable or incomprehensible. An HFDL service is provided by ARINC using a number 
of ground stations. These ground stations provide coverage out to ~2700 nm and on occasion 
provide coverage beyond that. Presently, HFDL ground stations are operating at the 
following locations (Figure 7.4): 

1. Santa Cruz, Bolivia. 

2. Reykjavik, Iceland. 

3. Shannon, Ireland. 
Auckland, New Zealand. 
Krasnoyarsk, Russia. 
Johannesburg, South Africa. 
Hat Yai, Thailand. 

8. Barrow, Alaska, USA. 

9. Molokai, Hawaii, USA. 

10. Riverhead, New York, USA. 

11. San Francisco, California, USA. 

12. Bahrain. 

13. Gran Canaria, Canary Islands. 



7.2.2 Very High Frequency 

Voice communication at very high frequency (VHF) is probably the most heavily used 
method of communication used by civil aircraft, although ultrahigh frequency (UHF) is 

Figure 7.4 HF data link ground stations. 


Ground Aircraft 

Transmitter/ Transmitter/ 

Receiver Receiver 

Figure 7.5 VHF signal propagation. 

generally preferred for military use. The VHF band for aeronautical applications operates in 
the frequency range 1 18.000-135.975 MHz with a channel spacing in past decades of 25 kHz 
(0.025 MHz). In recent years, to overcome frequency congestion, and taking advantage of 
digital radio technology, channel spacing has been reduced to 8.33 kHz (0.00833 MHz) 
which permits 3 times more radio channels in the available spectrum. Some parts of the 
world are already operating on the tighter channel spacing - this will be discussed in more 
detail later in the chapter in the section on global air transport management (GATM). 

The VHF band also experiences limitations in the method of propagation. Except in 
exceptional circumstances VHF signals will only propagate over line of sight. That is, the 
signal will only be detected by the receiver when it has line of sight or can 'see' the 
transmitter. VHF transmissions possess neither of the qualities of HF transmission and 
accordingly neither sky wave nor ground wave properties apply. This line-of- sight property 
is affected by the relative heights of the radio tower and aircraft. This characteristic applies 
to all radio transmissions greater than ~ 100 MHz, although the precise onset is determined 
by the meteorological conditions prevailing at the time (Figure 7.5). 

The formula that determines the line-of-sight range for VHF transmissions and above is as 

i? = 1.2V#t+1.2V#a 

where R is the range (nautical miles), H t is the height of the transmission tower (ft) and H a is 
the height of the aircraft (ft). 

Therefore, for an aircraft flying at 35 000 ft, transmissions will generally be received by a 
100 ft high radio tower if the aircraft is within a range of around 235 nautical miles. 

Additionally, VHF and higher-frequency transmissions may be masked by terrain, by a 
range of mountains, for example. These line-of-sight limitations also apply to equipment 
operating in higher-frequency bands and mean that VHF communications and other 
equipment operating in the VHF band or above - such as the navigation aids VOR and 
DME - may not be used except over large land masses, and then only when there is adequate 
transmitter coverage. Most long-range aircraft have three pieces of VHF equipment, with 
one usually being assigned to ARINC communications and reporting system ACARS 
transmissions, although not necessarily dedicated to that purpose. The requirements for 
certifying the function of airborne VHF equipment are given in Advisory Circular AC 20- 
67B (1986), while RTCA DO-186 (1984) specifies the necessary minimum operational 
performance standards (MOPS). There are advanced techniques that may be used in 
sophisticated military equipment that mitigate against these fundamental limitations. Such 
systems are said to possess an over-the-horizon (OTH) capability. 


A number of VHF data links (VHFDL) may be used, and these are discussed in more 
detail later in the chapter. ACARS is a specific variant of VHF communications operating on 
131.55 MHz that utilises a data link rather than voice transmission. As will be seen during 
the discussion on future air navigation systems, data link rather than voice transmission will 
increasingly be used for air-to-ground, air-to-ground and air-to-air communications as higher 
data rates may be used while at the same reducing flight crew workload. ACARS is dedicated 
to downlinking operational data to the airline operational control centre. The initial leg is by 
using VHF communications to an appropriate ground receiver, thereafter the data may be 
routed via land-lines or microwave links to the airline operations centre. At this point it will 
be allowed access to the internal airline storage and management systems: operational, flight 
crew, maintenance, etc. 

All aircraft and air traffic control centres maintain a listening watch on the international 
distress frequency: 121.5 MHz. In addition, military controllers maintain a listening watch 
on 243.0 MHz in the UHF band. This is because the UHF receiver could detect the 
harmonics of a civil VHF distress transmission and relay the appropriate details in an 
emergency (second harmonic of 121.5 MHz x 2 = 243.0MHz; these are the international 
distress frequencies for VHF and UHF bands respectively). 

7.2.3 Satellite Communications 

Satellite communications provide a more reliable method of communications using the 
International Maritime Satellite Organisation (INMARSAT) satellite constellation which 
was originally developed for maritime use. Now satellite communications, abbreviated to 
SATCOM, form a useful component of aerospace communications. In addition there are 
dedicated and secure military satellite systems not addressed in this book for obvious 
reasons. In this publication, SATCOM is described as it uses similar principles of operation 
and is also used in conjuction with the global air transport management (GATM) described 

The principles of operation of SATCOM are shown in Figure 7.6. The aircraft communicates 
via the INMARSAT constellation and remote ground earth station by means of C-band uplinks 
and downlinks to/from the ground stations and L-band links to/from the aircraft. In this way, 
communications are routed from the aircraft via the satellite to the ground station and on to the 
destination. Conversely, communications to the aircraft are routed in the reverse fashion. 
Therefore, provided the aircraft is within the area of coverage or footprint of a satellite, 
communication may be established. 

The airborne SATCOM terminal transmits on frequencies in the range 1626.5- 
1660.5 MHz and receives messages on frequencies in the range 1530.0-1559.0 MHz. 
Upon power-up, the radio-frequency unit (RFU) scans a stored set of frequencies and 
locates the transmission of the appropriate satellite. The aircraft logs on to the ground earth 
station network so that any ground stations are able to locate the aircraft. Once logged on to 
the system, communications between the aircraft and any user may begin. The satellite to 
ground C-band uplink and downlink are invisible to the aircraft, as is the remainder of the 
earth support network. 

The coverage offered by the INMARSAT constellation was a total of four satellites in 
2001. Further satellites are planned to be launched in the near future. The INMARSAT 
satellites are placed in earth geostationary orbit above the equator in the locations shown in 
Figure 7.7. 







Figure 7.6 SATCOM principles of operation. 

C-Band: 4-6 MHz 
L-Band: 1530 -1660 MHz 

54W 15.5W 

Figure 7.7 INMARSAT satellite coverage - 2001. 


Table 7.1 SATCOM configurations 

Configuration Capabilities 

Aero-H/H+ High gain. Aero-H offers a high-gain solution to provide a global capability and 

is used by long-range aircraft. Aero-H+ was an attempt to lower cost by 
using fewer satellite resources. Provides cockpit data, cockpit voice and 
passenger voice services 

Aero-I Intermediate gain. Aero-I offers similar services to Aero-H/H+ for medium- and 

short-range aircraft. Aero-I uses the spot beam service 

Aero-C Version that allows passengers to send and receive digital messages from a PC 

Aero-M Single-channel SATCOM capability for general aviation users 

1. Two satellites are positioned over the Atlantic: AOR-W at 54° west and AOR-E at 15.5° 

2. One satellite is positioned over the Indian Ocean: IOR at 64° east. 

3. One satellite is positioned over the Pacific Ocean: POR at 178° east. 

Blanket coverage is offered over the entire footprint of each of these satellites. In addition 
there is a spot beam mode that provides cover over most of the land mass residing under each 
satellite. This spot beam coverage is available to provide cover to lower-capability systems 
that do not require blanket oceanic coverage. 

The geostationary nature of the satellites does impose some limitations. Owing to low 
grazing angles, coverage begins to degrade beyond 80° north and 80° south and fades 
completely beyond about 82°. Therefore, no coverage exists in the extreme polar regions, a 
fact assuming more prominence as airlines seek to expand northern polar routes. A second 
limitation may be posed by the performance of the on-board aircraft system in terms of 
antenna installation, and this is discussed shortly. Nevertheless, SATCOM is proving to be a 
very useful addition to the airborne communications suite and promises to be an important 
component as future air navigation system (FANS) procedures are developed. 

A number of different systems are offered by SATCOM as described in Table 7.1. A 
SATCOM system typically comprises the following units: 

• Satellite data unit (SDU); 

• Radio-frequency unit (RFU); 

• Amplifiers, diplexers/splitters; 

• Low-gain antenna; 

• High-gain antenna. 

7.3 Transponders 

There are a number of different interrogators and transponders used on military aircraft. 
These are as follows: 

1. Distance measurement equipment (DME) is used as a navigation aid for both civil and 
military aircraft (see Chapter 8). 



2. Tactical air navigation (TACAN) is used as a navigation aid for military aircraft solely 
(see Chapter 8). 

3. ATC mode S is used both on military and civil aircraft, usually in association with the 
traffic collision avoidance system (TCAS). ATC mode S and TCAS are described below. 

4. Automatic dependent surveillance - address mode (ADS/A) is used to support FANS and 
GATM developments during oceanic crossings using HF communications. 

5. Automatic dependent surveillance - broadcast mode (ADS/B) is used to support FANS 
and GATM developments over land using VHF communications. 

6. Identification friend or foe (IFF) is used by the military specifically for identification of 
threat aircraft. IFF is compatible with ATC mode S and works on the same frequencies of 
1090 MHz (TX or interrogator) and 1090 MHz (RX or transponder) but carries additional 
identification codes specifically for military purposes. 

DME and TACAN are described under the Communications and Navaids section while ATC 
mode S is described under the GATM section of Chapter 8. 

7.3.1 Air Traffic Control (ATC) Transponder - Mode S 

As a means to aid the identification of individual aircraft and to facilitate the safe passage of 
aircraft through controlled airspace, the ATC transponder allows ground surveillance radars 
to interrogate aircraft and decode data which enables correlation of a radar track with a 
specific aircraft. The principle of transponder operation is shown in Figure 7.8. A ground- 
based primary surveillance radar (PSR) will transmit radar energy and will be able to detect 
an aircraft by means of the reflected radar energy - termed the aircraft return. This will 

1090 MHz 



1030 MHz 





Beacon Video 

I Radar I 
l Scope J 

Radar Video 

Figure 7.8 Principle of transponder operation. 


enable the aircraft return to be displayed on an ATC console at a range and bearing 
commensurate with the aircraft position. Coincident with the primary radar operation, a 
secondary surveillance radar (SSR) will transmit a series of interrogation pulses that are 
received by the on-board aircraft transponder. The transponder aircraft replies with a 
different series of pulses which give information relating to the aircraft, normally aircraft 
identifier and altitude. If the PSR and SSR are synchronised, usually by being co- 
boresighted, then both the presented radar returns and the aircraft transponder information 
may be presented together on the ATC console. Therefore, the controller will have aircraft 
identification (e.g. BA 123) and altitude presented alongside the aircraft radar return, thereby 
greatly improving the controller's situational awareness. 

The system is also known as identification friend or foe (IFF)/secondary surveillance radar 
(SSR), and this nomenclature is in common use in the military field. On-board the aircraft, 
the main elements are as listed below: 

1. ATC transponder controller unit for setting modes and response codes. 

2. A dedicated ATC transponder unit. 

3. An ATC antenna unit with an optional second antenna. It is usual to utilise both upper and 
lower mounted antennas to prevent blanking effects as the aircraft manoeuvres. 

The SSR interrogates the aircraft by means of a transmission on the dedicated frequency of 
1030 MHz which contains the interrogation pulse sequence. The aircraft transponder replies 
on a dedicated frequency of 1090 MHz with a response that contains the reply pulse 
sequence with additional information suitably encoded in the pulse stream. 

In its present form the ATC transponder allows aircraft identification - usually the airline 
call-sign - to be transmitted when using mode A. When Mode C is selected, the aircraft will 
respond with its identifier together with altitude information. 

More recently, an additional mode - mode S or mode select - has been introduced with 
the intention of expanding this capability. In ATC mode S the SSR uses more sophisticated 
monopulse techniques that enable the aircraft azimuth bearing to be determined more 
quickly. Upon determining the address and location of the aircraft, it is entered into a roll call 
file. This, together with details of all the other aircraft detected within the interrogator's 
sphere of operation, forms a complete tally of all the aircraft in the vicinity. Each mode S 
reply contains a discrete 24 bit address identifier. This unique address, together with the fact 
that the interrogator knows where to expect the aircraft from its roll call file, enables a large 
number of aircraft to operate in a busy air traffic control environment (see section 7.3.2 for 
details of the traffic collision avoidance system). 

ATC mode S has other features that enable it to provide the following: 

• Air-to-air as well as air-to-ground communication; 

• The ability of aircraft autonomously to determine the precise whereabouts of other 
aircraft in their vicinity. 

Mode S is an improved conventional secondary radar operating at the same frequencies 
(1030/1090 MHz). Its 'selectivity' is based on unambiguous identification of each aircraft 
by unique 24 bit addresses. This acts as its technical telecommunications address, but does 
not replace the mode A code. There are also plans for recovery of the A and C codes via 
mode S. 


Apart from this precise characterisation of the aircraft, mode S protects the data it 
transmits owing to the inclusion of several parity bits which means that up to 12 erroneous 
bits may be tolerated by the application of error detection and correction algorithms. For 
transmission, these parity bits are superimposed on those of the mode S address. 

Finally, mode scan may be used to exchange longer, more varied data streams, which can 
even be completely unplanned. To do this, mode S transmissions between the station and the 
transponder use highly sophisticated 56 or 1 12 bit formats called frames. They fall into three 
main categories: 56 bit surveillance formats, 112 bit communication formats with a 56 bit 
data field, which are in fact 'extended' surveillance formats (uplink COMM-As and 
downlink COMM-Bs), and 112 bit communication formats with an 80 bit data field (uplink 
COMM-Ds and downlink COMM-Ds). This feature will be of use in facilitating the 
transmission and interchange of flight plans dynamically revised in flight which is one of 
the longer-term aims of FANS. 

Mode S also has the capability of providing a range of data formats, from level 1 to level 4. 
These are categorised as follows: 

1. Level 1. This is defined as the minimum capability mode S transponder. It has the 
capability of reply to mode S interrogations but has no data link capability. All the 
messages provided by level 1 are short (56 bit) messages. 

2. Level 2. These transponders support all the features of the level 1 transponder with the 
addition of standard length data link word formats. This can entail the use of longer 
messages (112 bit). Some of the messages are used for TCAS air-to-air communication 
while others are utilised for air-to-ground and ground-to-air communication as part of the 
enhanced surveillance data access protocol system (DAPS) requirements. 

3. Level 3. The level 3 transponders embrace the same functionality as level 2 with the 
additional ability to receive extended length messages (ELM) which comprise 16 
segments of information, each containing a 112 bit message. 

4. Level 4. Level 4 has the full functionality of level 3 with the capability of transmitting 
ELM messages of up to 16 segments of 112 bit word messages. 

Originally it was envisaged that ATC mode S would be the primary contender to provide the 
CNS/ATM functionality by providing large block transfers of information. More recently it 
has been realised that VDL mode 4 might better serve this need, and levels 3 and 4 are no 
longer required. 

When used together with TCAS, ATC mode S provides an important feature for FANS, 
that of automatic dependent surveillance - A (ADS-A). This capability will assist the safe 
passage of aircraft when operating in a direct routing mode. 

7.3.2 Traffic Collision and Avoidance System 

The traffic collision and avoidance system (TCAS) was developed in prototype form during 
the 1960s and 1970s to provide a surveillance and collision avoidance system to help aircraft 
avoid collisions. It was certified by the FAA in the 1980s and has been in widespread use in 
the United States in its initial form. The TCAS is based upon a beacon interrogator and 
operates in a similar fashion to the ground-based SSR already described. The system 
comprises two elements: a surveillance system and a collision avoidance system. The TCAS 


detects the range bearing and altitude of aircraft in the near proximity for display to the 

The TCAS transmits a mode C interrogation search pattern for mode A and C transponder 
equipped aircraft and receives replies from all such equipped aircraft. In addition, the TCAS 
transmits one mode S interrogation for each mode S transponder equipped aircraft, receiving 
individual responses from each one. It will be recalled that mode A relates to range and 
bearing, while mode C relates to range, bearing and altitude and mode S to range, bearing 
and altitude with a unique mode S reply. The aircraft TCAS equipment comprises a radio 
transmitter and receiver, directional antennas, computer and flight deck display. Whenever 
another aircraft receives an interrogation it transmits a reply and the TCAS computer is able 
to determine the range depending upon the time taken to receive the reply. The directional 
antennas enable the bearing of the responding aircraft to be measured. The TCAS can track 
up to 30 aircraft but only display 25, the highest-priority targets being the ones that are 

The TCAS is unable to detect aircraft that are not carrying an appropriately operating 
transponder or that have unserviceable equipment. A transponder is mandated if an aircraft 
flies above 10 000 ft or within 30 miles of major airports; consequently, all commercial 
aircraft and the great majority of corporate and general aviation aircraft are fitted with the 

The TCAS exists in two forms: TCAS I and TCAS II. TCAS I indicates the range and 
bearing of aircraft within a selected range; usually 15-40 nm forward, 5-15 nm aft and 10- 
20 nm on each side. The system also warns of aircraft within ±8700 ft of the aircraft's own 

The collision avoidance system element predicts the time to, and separation at, the 
intruder's closest point of approach. These calculations are undertaken using range, closure 
rate, altitude and vertical speed. Should the TCAS ascertain that certain safety boundaries 
will be violated, it will issue a traffic advisory (TA) to alert the crew that closing traffic is in 
the vicinity via the display of certain coloured symbols. Upon receiving a TA, the flight crew 
must visually identify the intruding aircraft and may alter their altitude by up to 300 ft. A TA 
will normally be advised between 20 and 48 s before the point of closest approach with a 
simple audio warning in the flight crew's headsets: 'TRAFFIC, TRAFFIC TCAS I does not 
offer any deconfliction solutions but does provide the crew with vital data in order that they 
may determine the best course of action. 

TCAS II offers a more comprehensive capability with the provision of resolution 
advisories (RAs). TCAS II determines the relative motion of the two aircraft and determines 
an appropriate course of action. The system issues an RA via mode S, advising the pilots to 
execute the necessary manoeuvre to avoid the other aircraft. An RA will usually be issued 
when the point of closest approach is within 15 and 35 s, and the deconfliction symbology is 
displayed coincident with the appropriate warning. 

A total of ten audio warnings may be issued. Examples are: 




Finally, when the situation is resolved: 'CLEAR OF CONFLICT'. 



ATC Mode S 
Transponder 1 

I — y/\ Audio 
I *1_ kj System 







o o o 


Figure 7.9 TCAS architecture showing related equipment and displays. 

TCAS II clearly requires a high level of integration between the active equipment. 
Figure 7.9 shows the interrelationship between: 

• TCAS transmitter/receiver; 

• ATC mode S transponders; 

• VSI display showing vertical guidance for TAs and RAs; 

• Optional horizontal situational indicator for RAs that could be the navigation display; 

• Audio system and annunciators; 

• Antennas for ATC mode S and TCAS. 

This is indicative of the level of integration required between ATC mode S transponders, 
TCAS, displays and annunciators. It should be noted that there are a variety of display 
options and the system shown does not represent the only TCAS option. 

More recently, further changes have been introduced to TCAS II - known as TCAS II 
change 7. This introduces software changes and updated algorithms that alter some of the 
TCAS operating parameters. Specifically, change 7 includes the following features: 

• Elimination of nuisance warnings; 

• Improved RA performance in a multiaircraft environment; 

• Modification of vertical thresholds to align with reduced vertical separation minima 
(RVSM) - see the section on global air transport management (GATM) and the civil 
equivalent future air navigation system (FANS); 

• Modification of RA display symbology and aural annunciations. 

The change 7 modifications became mandatory in Europe for aircraft with 30 seats or more 
from 31 March 2001 and for aircraft with more than 19 seats from 1 January 2000. The rest 


of the world will be following a different but broadly similar timescale for implementation. 
Change 7 is not mandated in the United States but it is expected that most aircraft will be 
equipped to that standard in any case. Further information can be found on AC- 12955 A, 
RTCA DO-181 DO- 185 certification and performance requirements for TCAS II and mode S. 

7.3.3 Automatic Dependent Surveillance - Address Mode (ADS-A) 

ADS-A will be used to transmit the aircraft four-dimensional position and flight plan intent 
based upon GPS position during oceanic crossings. The communications media will be 
SATCOM or HF data link (HFDL). ADS-A requires the aircraft to be fitted with an FMS and 
CDU and with some means of displaying message alerts and annunciation. 

7.3.4 Automatic Dependent Surveillance - Broadcast Mode (ADS-B) 

ADS-B will be used to transmit four-dimensional position and flight plan intent based upon 
GPS position using line-of-sight VHF communications. Either mode S or digital VHF radio 
will be used to transmit the data. ADS-A requires a cockpit display of traffic information. 

7.3.5 Identification Friend or Foe (IFF) 

There are two ways in which IFF equipment may be used: 

• Providing 360° coverage in order to be able to respond to interrogation and receive 
transponder returns from friendly aircraft in any direction. In this respect the operation is 
very similar to the airborne operation of ATC mode S when used in association with the 

• Used in association with a primary radar sensor in order to be able specifically to identify 
targets appearing within the radar scan. This operates in the same way as a ground 
surveillance radar interrogating aircraft in the vicinity of an airfield. 

An example of the first type is the advanced IFF (AIFF) AN/APX-1 13(V) used on-board the 
F-16 aircraft but which is typical of equipment of this type (Figure 7.10). IFF equipment is 
sometimes referred to in a generic sense as IFF mark XII which relates the generic family of 

Figure 7.10 IFF set AN/APX-1 13(V). (BAE SYSTEMS) 


|FF Primary Radar 

Interrogator - 1 .090Ghz ~ 3Ghz for AWACS 

Transponder - 1 .030 Ghz Airborne intercept 

Figure 7.11 Co-boresighting of IFF interrogator with radar. 

present IFF equipment. Equipment such as the AN/APX-113(V) has the following 

• Multiple antenna configurations - electronic or mechanical scan; 

• MIL-STD-1553 interface to connect to the rest of the avionics system units; 

• Ability to provide encryption capability; 

• Provision of growth to accommodate future functional modes. 

The use of the interrogator co-located and co-boresighted with the main radar creates a 
problem as illustrated in Figure 7.11. The primary radar will be operating at a higher 
frequency than the 1090/1030 MHz that the interrogator uses. An airborne early warning 
radar will be operating at ~3 GHz, while an airborne intercept (AI) radar will be operating 
higher still at ~ 10 GHz. For a given radar antenna size the beam width is inversely 
proportional to frequency, so, the higher the frequency, the more narrow is the beamwidth. 
The IFF beamwidth will therefore encompass the main beam and several sidelobes of the 
radar beam and may therefore be receiving returns from targets that are not of fundamental 
interest to the radar. This effect will be more pronounced for AI as opposed to AWACS radar. 

7.4 Data Links 

The use of voice was the original means of using RF communications. However, the use of 
speech has severe limitations; it is slow in terms of conveying information and prone to 
misunderstanding, whereas high bandwidth data links can delivery more information, if 
necessary incorporating error correction or encryption. In the avionics sense, typical data 
link users are portrayed in Figure 7.12. 
Primary users include: 

• Strategic airborne sensor platforms such as E-4, E-6, E-8, Global Hawk and satellites; 

• Tactical airborne sensors and shooters - F-15, F-16, F-18, Harrier, Tornado, Eurofighter 
Typhoon and tactical UAVs/UCAVs among others; 

• Shipborne sensors; 

• Land forces. 






Figure 7.12 Typical data link users. 

Many of the data links are limited to line-of- sight operation owing to the transmission 
characteristics of the RF frequencies being employed. However, the use of communications 
satellites to perform a relay function permits transmission of data over the horizon (OTH), 
thereby enabling intra- and intertheatre communications. 

Typical data packages that may be delivered by data links include: 

• Present position reporting; 

• Surveillance; 

• Aircraft survival, EW and intelligence information; 

• Information management; 

• Mission management; 

• Status. 

The primary data links used for communications between airborne platforms and space and 
surface platforms are as follows: 

1. Link 16. This is the most commonly used avionics data link and is usually manifested in 
avionics systems as the joint tactical information distribution system (JTIDS). The JTIDS 
is also compatible with the US Navy data link satellite tactical data information link J 
(S-TADIL J). Link 16 operates in the UHF frequency band in the same frequency range as 
identification friend or foe (IFF), distance measurement equipment (DME) and TACAN, 
as described below. 

2. Link 11. Certain strategic aircraft assets such as E-4 and Nimrod MRA4 associated with 
joint maritime operations also have the capability of operating with link 11 - a data link 
commonly used by naval forces. 



51 Frequencies! 








■ 126 Channels 



■ 126 Channels 




- DME Interrogater Operation - 

-DME Transponder Operation- 



1151 1213 
Mhz Mhz 

Figure 7.13 JTIDS frequency band. 

7.4.1 JTIDS Operation 

The characteristics of the JTIDS frequency band and how this is shared with other equipment 
is shown in Figure 7.13. The JTIDS characteristics are as follows: 

1. Data are transmitted in the UHF band between 969 and 1206 MHz. 

2. Frequency-hopping techniques are employed to provide ECM jam-resistant properties. 

3. A total of 51 channels are provided at 3 MHz spacing. 

JTIDS transmissions are constrained to avoid interfering with the IFF frequencies at 
1030 MHz (TX) and 1090 MHz (RX), and JTIDS is not employed within ±20 MHz of 
these frequencies. The problem of mutual equipment interference is one that has to be 
frequently faced on highly integrated military avionic platforms. 

Integration with the host aircraft avionics system usually takes the form shown in 
Figure 7.14. The JTIDS terminal and associated antenna are shown on the right of the 
diagram. The equipment shown in this particular case is the URC-138 terminal, a typical 
example of which is shown in the inset. Such a terminal - compatible with tanker/transport, 
fighter and helicopter environments will have the following capabilities/characteristics: 

Data rate 

28.8-238 kbps 


40 lbs 


12.5 in deep 

7.5 in high 

10 in wide 

(equivalent to 8ATR) 


750 W 



Data Link 

Upper Antenna 

I Platfo 

, Mil-Std- 
I 1553B 
1 Data Buses 


r— •►( 



C X+- 

{ r*~ 


Aircraft Sensors 

JTIDS Terminal 

Radar ) 




Electro-Optics ) 


Mission System 




( W 

Lower Antenna 



ESM ) 

JTIDS Terminal (Rockwell Collins) 
Figure 7.14 JTIDS integration with the avionics system. 

The JTIDS terminal typically interfaces with the host aircraft mission system computer via 
MIL-STD-1553B data buses. The host mission systems computer is connected in turn to the 
aircraft platform sensors embracing radar, electrooptics, ESM/EW and other operational 
sensors depending upon the host platform sensor fit. 

Such a system will in most cases include secure voice capabilities and the ability to 
transmit encrypted data. Clearly, in a real battlefield scenario there is a need to share 
classified information between some but certainly not all of the participants. With the data 
that many participants will utilise in a communications systems used in a military 
environment, therefore, there needs to be the capability of separating secure/classified 
data from the data that are 'in the clear' or open to all participants. 


7.4.2 Other Data Links 

Apart from JTIDS which specifically operates in the UHF band, other transmission 
techniques may be used to communicate between military platforms. These are: 


• HF data links; 

• Local cooperative data links. 

SATCOM and HF data links or HFDL are already used extensively by the maritime and civil 
aviation communities. The same transmission capabilities are open to the military commu- 
nity, except that in many cases data protection/encryption may be required depending upon 
the sensitivity of the message content. Therefore, many military communications systems are 
designed to include an encryption/decryption device at the front end - between the processor 
and transmit/receive elements. By using suitable encryption 'keys', the necessary levels of 
encryption may be achieved depending upon the sensitivity of the message content. 

Aircraft such as the F-22 Raptor use a local cooperative data link to aid in the data sharing 
and coordination of a group of aircraft embarked upon a shared mission. As outlined in 
Chapter 2, on the F-22 it was the intention to utilise two phased array related cooperative 
data links. These are: 

• A common high-band data link (CHBDL) or in-flight data link (IFDL) operating at around 
10 GHz and utilising three antenna locations to pass data between adjacent aircraft; 

• A cooperative engagement capability (CEC) using similar antenna configurations. 

It is not clear whether either or both of these facilities have been included in the final aircraft 

7.5 Network-centric Operations 

Network-centric operations are becoming the latest 'force multiplier' element of modern 
airborne warfare in the same way as air-to-air refuelling and the availability of digital signal 
processing have been in the past. The use of high-bandwidth digital communications, 
together with sophisticated signal processing capabilities and the high bandwidth of internal 
platform interconnective buses and highways, have enabled the data interchange between a 
variety of sensor and weapons platforms to ascend to much higher levels. The command and 
control (C 2 ) structure, sensor and weapon delivery platforms have become integrated at a 
level that would previously have been unimaginable. This connectivity, allied with the 
capability of ultrahigh-re solution sensors, enables target and threat data to be shared at all 
levels of the force structure with unprecedented speed and fidelity. 

The nature of network-centric operations may be appreciated by reference to Figure 7.15. 
This figure illustrates three tiers of interconnected centres or nodes, each of which are 
interconnected at the three respective levels and which are also interconnected between 
levels or layers by specific network nodes. In the figure, interconnecting nodes between 
layers are shown in black while supporting nodes within a layer are portrayed in white. 

The network comprises three layers, which in descending order of importance/authority 
are as follows: 

















< 20 Users 






Joint Data 






(Link 16/11) 





< 500 Users 




Joint Planning 






~ 1000 Users 

Figure 7.15 

Nature of network 






1 . Weapons control layer. This embraces a limited number of participants - 20 or fewer - all 
operating at the strategic level, all interconnected within the layer and also connected to 
subordinate layers to implement force control. It is at this level that the rules of 
engagement (RoE) for a particular task execution will be decided and implemented. 
Airborne platforms in this category may include platforms such as AWACS or similar 
airborne assets. 

2. Force control layer. This layer exercises control over the force structure, implementing the 
RoE and engaging targets on the basis of sensor and tactical information exchange. This 
joint data network will be typified by link 11 and link 16 users exchanging data at the 
tactical level and deciding the priorities according to different target types depending 
upon geographical and time currency of intelligence and target data. This layer may 
include up to 500 users including strategic and tactical force assets. Aircraft platforms 
may include maritime reconnaissance and fighter aircraft. 

3. Force coordination layer. This layer embraces a joint planning network invoking force 
coordination at the local or theatre level and exercising force coordination to achieve 
maximum force effect or to avoid 'blue-on-blue' fratricide engagements. In the aviation 
context this may include fighter aircraft, attack helicopters, UAVs and forward air 
controllers (FAC). This network may extend across 1000 users or contributors. 

It is noteworthy that the nature of the information changes as it migrates from the lowest to 
the highest level within this hierarchy according to the simple tabulation below: 


Information timecales 

Information accuracy 

Weapons control 


High accuracy 


Force control 



Force coordination 



Low accuracy 




u ■ 

Control ■ 

Sensor i 

• bensors .........................I...... .......... 



& Control 

Information \ 

-i— -• 

-I- Control 

...g^.J I ! — 






Figure 7.16 Sensor/shooter information grid. 

The doctrines associated with network-centric operations have many proponents, especially 
in the United States. One of the most celebrated and vocal proponents is USN Vice Admiral 
Arthur K. Cebrowski together with J.J. Garstka (1998). 

Key elements to the operation of a network-centric operation relate to the information grid 
interrelationship between the 'sensors' and 'shooters' involving the command and control 
element (Figure 7.16). This depicts the information flow and command links that exist 
between the detection of a target on the left to the engagement of the target on the right. It 
embraces the overlapping nature of information and engagement grids that determine the 
process by which information is processed between sensor, command and control and 
shooter to ensure that the necessary information is provided to the command function in 
order that a target may be correctly assessed, command and control may be exercised and 
battle damage assessment may be accomplished. 

The high-bandwidth communications available for intelligence and target data interchange 
between these functional entities mean that radar video or electrooptic images may be 
exchanged in near real time. Therefore, the decision time to identify, categorise and authorise 
target engagement has reportedly decreased from hours (1991 Gulf War) to tens of minutes 
(Afghanistan War), with the aim of reducing this to a matter of minutes in future conflicts. 

In the avionics environment this information exchange is achieved by using data links 
using a series of transmission means as described elsewhere in this chapter. On-board the 
airborne platform the availability of high-bandwidth fibre-optic or fibre-channel commu- 
nications as described in Chapter 2, Technology and Architectures. 


Advisory circular AC 20-67B, Airborne VHF communications installations, 16 January 1986. 
Advisory circular AC 20- 131 A, Air worthiness approval of traffic alert and collision avoidance systems 
(TCAS II) and mode S transponders, 29 March 1993. 


Advisory circular 129-55 A, Air carrier operational approval and use of TCAS II, 27 August 1993. 
Cebrowski, A.K. and Garstka, J. J. (1998) Network-centric warfare: its origin and future. Proceedings of 

Naval Institute, January. 
RTCA DO- 181, Minimum operational performance standards for air traffic control radar beacon system/ 

mode select (ATCRBS/mode S) airborne equipment. 
RTCA DO- 185, Minimum operational performance standards for traffic alert and collision avoidance 

systems (TCAS) airborne equipment. 
RTCA DO- 186, Minimum operational performance standards (MOPS) for radio communications 

equipment operating with the radio frequency range 1 17.975 to 137.000 MHz, dated 20 January 1984. 

8 Navigation 

8.1 Navigation Principles 

8.1.1 Introduction 

Navigation has been an ever-present component of humankind's exploitation of the 
capability of flight. While the principles of navigation have not changed since the early 
days of sail, the increased speed of flight, particularly with the advent of the jet age, has 
placed an increased emphasis upon accurate navigation. The increasingly busy skies, 
together with rapid technology developments, have emphasised the need for higher-accuracy 
navigation and the means to accomplish it. Navigation is no longer a matter of merely getting 
from A to B safely, it is about doing this in a fuel-efficient manner, keeping to tight airline 
schedules, and avoiding other air traffic - commercial, general aviation, leisure and military. 
Navigation of military aircraft has to comply with the same regulations as civil traffic when 
operating in controlled airspace. Platforms adopted from civil aircraft will retain the civil 
navigation systems as described in the companion volume 'Civil Avionics Systems' (Moir 
and Seabridge, 2003), some of which are described here for ease of reference. More than 
likely, legacy military platforms will be fitted with a bespoke system meeting most but 
possibly not all the latest requirements specified for controlled airspace and may on occasion 
need to operate with certain limitations until the necessary upgrades are embodied. 

Outside controlled airspace in operational theatres the navigational accuracy will be 
determined by the accuracy provided by the platform mission, weapons system and possibly 
the weapons being carried. Operational mission navigation constraints include the optimisa- 
tion of routing to avoid hazardous surface-to-air missile (SAM) and anti-aircraft artillery 
(AAA). Routing may also take into account the need to maximise the aircraft stealth or low- 
observability characteristics. Therefore, while military aircraft need to adopt all the 
necessary features to enable them to operate safely alongside civil aircraft in today's 
crowded airspace, they also add a further layer of complexity to the mission management 

This section summarises some of the modern methods of navigation, leading to more 
detailed descriptions of how each technique operates. A later section in the chapter relates to 

Military Avionics Systems Ian Moir and Allan G. Seabridge 
© 2006 John Wiley & Sons, Ltd. ISBN: 0-470-01632-9 


future air navigation system (FANS) requirements, also known within military circles as 
global air transport management (GATM). 

The main methods of navigation as practised today may be summarised and simplified as 

• Classic dead-reckoning navigation using air data and magnetic, together with Doppler or 

• Radio navigation using navigation aids - ground-based radio-frequency beacons and 
airborne receiving and processing equipment; 

• Barometric inertial navigation using a combination of air data and inertial navigations 
(IN) or Doppler; 

• Satellite navigation using a global navigation satellite system (GNSS), more usually a 
global positioning system (GPS); 

• Multiple- sensor navigation using a combination of all the above. 

The more recent the pedigree of the aircraft platform, the more advanced the navigational 
capabilities are likely to be. However, it is common for many legacy platforms to be 
retrofitted with inertial and GNSS navigation, the accuracy of which far outshines the 
capability of the original system. 

8.1.2 Basic Navigation 

The basic navigation parameters are shown in Figure 8.1 and may be briefly summarised as 

1. An aircraft will be flying at a certain height or altitude relative to a barometric datum 
(barometric altitude) or terrain (radar altitude). 

2. The aircraft may be moving with velocity components in the aircraft X (V x ), Y ( V y ) and Z 
(V z ) axes. Its speed through the air may be characterised as indicated airspeed (IAS) or 
Mach number (M). Its speed relative to the ground is determined by true airspeed (TAS) 
in still air conditions. 

3. The aircraft will be flying on a certain heading; however, the prevailing wind speed and 
direction will modify this to the aircraft track. The aircraft track represents the aircraft 
path across the terrain and will lead to the destination or next waypoint of the aircraft. 
Wind speed and direction will modify the aircraft speed over the ground to ground speed. 

4. The aircraft heading will be defined by a bearing to magnetic (compass) north or to true 
north relating to earth-related geographic coordinates. 

5. The aircraft will be flying from its present position - defined by latitude and longitude to a 
waypoint also characterised by latitude and longitude. 

6. A series of flight legs - defined by way points - will determine the aircraft designated 
flight path from the departure airfield to the destination airfield. 

As has already briefly been described, there are sensors and navigation techniques that may 
be used solely or in combination to navigate the aircraft throughout the mission. 

The relationship of the different axis sets is shown in Figure 8.2. These may be 
characterised as follows: 



IAS or 


& Direction) 




Aircraft Present 


Latitude Longitude 




(or Variation) 



Drift Angle 

Figure 8.1 Basic navigation parameters. 

1. Earth datum set. As shown in Figure 8.2, the earth axis reference set comprises the 
orthogonal set E x , E y , E z , where: 

• E x represents true north; 

• E y represents east; 

• E z represents the local gravity vector. 

Figure 8.2 Earth-related coordinates. 



2. The orthogonal aircraft axis set where: 

• A x is the aircraft longitudinal axis (corresponding to the aircraft heading); 

• A y is the aircraft lateral axis; 

• A z is the aircraft vertical axis (corresponding to E z ). 

For navigation purposes, the accuracy with which the aircraft attitude may be determined is a 
key variable for Doppler navigation systems in which the Doppler velocity components need 
to be resolved into aircraft axes. Similarly, attitude is used for IN axis transformations. 

The navigation function therefore performs the task of manoeuvring the aircraft from a 
known starting point to the intended destination, using a variety of sensors and navigation 

The classic method of navigation which has been in used for many years is to use a 
combination of magnetic and inertial directional gyros used together with airspeed 
information derived from the air data sensors to navigate in accordance with the parameters 
shown in Figure 8.1. This is subject to errors in both the heading system and the effects of en- 
route winds which can cause along-track and across-track errors. In the 1930s it was 
recognised that the use of radio beacons and navigation aids could significantly reduce these 
errors by providing the flight crew with navigation assistance related to precise points on the 

8.2 Radio Navigation 

For many years the primary means of navigation over land, at least in continental Europe and 
the North American continent, was by means of radio navigation routes defined by VHF 
omniranging/distance measuring equipment (VOR/DME) beacons as shown in Figure 8.3. 
By arranging the location of these beacons at major navigation or crossing points, and in 
some cases airfields, it was possible to construct an entire airway network that could be used 
by the flight crew to define the aircraft flight from take-off to touchdown. Other radio 
frequency aids include distance measuring equipment (DME) and non-distance beacons 
(NDB). The operation of the radio navigation and approach aids is described elsewhere in 
this chapter. 

Waypoint 1 

Waypoint 2 

Waypoint 3 

NDB 2 

Figure 8.3 Radio navigation using VOR, DME and automatic direction finding (ADF). 


Figure 8.3 shows: 

1. Three VOR/DME beacon pairs: VOR 1/DME 1, VOR 2/DME 2 and VOR 3/DME 3 
which define waypoints 1 to 3. These beacons represent the intended waypoints 1, 2 and 3 
as the aircraft proceeds down the intended flight plan route - most likely an identified 
airway. When correctly tuned, the VOR/DME pairs succesively present the flight crew 
with bearing to and distance from the next waypoint. 

2. Off-route DME beacons, DME 4 and DME 5, may be used as additional means to locate 
the aircraft position by means of the DME fix obtained where the two DME 4 and DME 5 
range circles intersect. As will be seen, DME/DME fixes are a key attribute in the modern 
navigation system. 

3. Off-route NDB beacons may be used as an additional means to determine the aircraft 
position by obtaining a cross-fix from the intersection of the bearings from NBD 1 and 
NDB 2. These bearings are derived using the aircraft ADF system. 

4. In the military context: as well as these beacons TACAN and VORTAC beacons may be 
used specifically. TACAN has the particular advantage that it may be used in an offset 
mode where a navigational point may be specified in terms of a range and bearing offset 
from the TACAN beacon itself. TACAN also has certain features whereby it may be used 
to determine the range and bearing to other formations, eg a tanker aircraft, thereby 
facilitating airborne rendezvous operations. 

Thus, in addition to using navigation information from the 'paired' VOR/DME or TACAN 
beacons that define the main navigation route, position fix, cross-fix, range or bearing 
information may also be derived from DME or NDB beacons in the vicinity of the planned 
route by using automatic direction- finding techniques. As has already been described in 
Chapter 7, a major limitation of the radio beacon navigation technique results from line-of- 
sight propagation limitations at the frequencies at which both VOR and DME operate. As 
well as the line-of-sight and terrain-masking deficiencies, the reliability and accuracy of the 
radio beacons can also be severely affected by electrical storms. Over longer ranges, 
LORAN-C could be used if fitted. 

Owing to the line-of-sight limitations of these radio beacons, these navigation techniques 
were only usable overland where the beacon coverage was sufficiently comprehensive or for 
close off-shore routes where the beacons could be relied upon. 

8.2.1 Oceanic Crossings 

In 1969 the requirements were already specified for self-contained long-range commercial 
navigation by advisory circular AC 121-13. The appropriate document specified that self- 
contained navigation systems should be capable of maintaining a maximum error of ±20 nm 
across track and ±25 nm along track for 95% of the flights completed. Two systems were 
addressed in the specification: one using Doppler radar and the other using an inertial 
navigation system (INS). 

In June 1977, the North Atlantic (NAT) minimum navigation performance specifications 
(MNPS) were altered to reflect the improved navigation sensors - see advisory circular AC 
120-33. This defined the separation requirements for long-range navigation over the North 
Atlantic. The lateral separation was reduced from 120 to 60 nm while retaining the previous 
vertical separation of 2000 ft. Statistical limits were specified as to how long an aircraft was 



allowed to spend 30 nm off-track and between 50 and 70 nm off-track - the latter actually 
representing an overlap with an adjacent track. The standard deviation of lateral track errors 
was specified as 6.3 nm. 

The Doppler radar system was specified as being an acceptable navigation means applying 
within certain geographical boundaries. Eastern and western entry points or 'gateways' were 
specified as entry and departure points into and out of the North Atlantic area. These 
gateways were identified as a number of specific named NDB or VOR beacons on both sides 
of the ocean. The North Atlantic transit area was specified as being the oceanic area bounded 
by the eastern and western gateways and lying between the latitude of 35°N and 65°N. By 
the standards of the allowable navigation routes available to today's aviators, this represented 
a very restricted envelope. 

The aircraft equipment requirements were also carefully specified: 

• Dual Doppler and computer systems; 

• Dual polar path compasses; 

• ADF; 

• VOR; 

• One LORAN receiver capable of being operated from either pilot's station. 

8.3 Inertial Navigation Fundamentals 

The availability of inertial navigation systems (INS) to the military aviation community 
during the early 1960s added another dimension to the navigation equation. Now flight crew 
were able to navigate by autonomous means using an on-board INS with inertial sensors. By 
aligning the platform to earth-referenced coordinates and present position during initialisa- 
tion, it was now possible to fly for long distances without relying upon LORAN, VOR/DME 
or TACAN beacons. Waypoints could be specified in terms of latitude and longitude as 
arbitrary points on the globe, more suited to the aircraft's intended flight path rather than a 
specific geographic feature or point in a radio beacon network (Figure 8.4). The operation of 
inertial platforms is described in detail in section 8.8. 

The specifications in force at this time also offered an INS solution to North Atlantic 
crossings as well as the dual-Doppler solution previously described. The inertial solution 
required serviceable dual INS and associated computers to be able to undertake the crossing. 

Waypoint 1 

Waypoint 2 

Waypoint 3 

Figure 8.4 Fundamentals of inertial navigation. 


There were also limitations on the latitudes at which the ground alignment could be 
performed - 76° north or south - as attaining satisfactory alignment becomes progressively 
more difficult the nearer to the poles the INS becomes. 

Advisory circular AC 25-4 set forth requirements for operating an INS as a sole means of 
navigation for a significant portion of the flight. These requirements may be summarised as 

1. The ability to provide the following functions: 

• Valid ground alignment at all latitudes appropriate for the intended use of the INS; 

• The display of alignment status to the crew; 

• Provision of the present position of the aircraft in suitable coordinates: usually latitude 
from +90° (north) to -90° (south) and longitude from +180° (east) to -180° (west); 

• Provision of information on destinations or waypoints; 

• Provision of data to acquire and maintain the desired track and the ability to determine 
deviation from the desired track (across-track error); 

• Provision of information needed to determine the estimated time of arrival (ETA). 

2. The ability to comply with the following requirements: 

• ±20 nm across track and ±25 nm along track; 

• Maintainenance of this accuracy on a 95% probability basis at representative speeds 
and altitudes and over the desired latitude range; 

• The capacity to compare the INS position with visual fixes or by using LORAN, 
TACAN, VOR, DME or ground radar (air traffic control). 

3. The provision of a memory or in-flight alignment means. Alternatively, the provision of a 
separate electrical power source - usually a dedicated stand-alone battery - able to 
support the INS with full capability for at least 5 min in the event of an interruption of the 
normal power supply. 

8.4 Satellite Navigation 

The foregoing techniques were prevalent from the 1960s to the 1990s when satellite navigation 
became commonly available. The use of global navigation satellite systems (GNSS), to use the 
generic name, offers a cheap and accurate navigational means to anyone who possesses a 
suitable receiver. Although the former Soviet Union developed a system called GLONASS, it is 
the US global positioning system (GPS) that is the most widely used. The principles of satellite 
navigation using GPS will be described in detail later in the chapter. 

GPS receivers may be provided for the airborne equipment in a number of ways: 

1. Stand-alone GPS receivers, most likely to be used for GPS upgrades to an existing 
system. These are multichannel (typically, 12-channel) global navigation satellite system 
(GNSS) receivers - the B777 utilises this approach. 

2. GPS receivers integrated into a multifunction receiver unit called a multimode receiver 
(MMR). Here, the GPS receiver function is integrated into one LRU along with VOR and 
ILS receivers. 


8.4. 1 Differential GPS 

One way of overcoming the problems of selective availability is to employ a technique called 
differential GPS (DGPS). Differential techniques involve the transmission of a corrected 
message derived from users located on the ground. The correction information is sent to the 
user who can apply the corrections and reduce the satellite ranging error. The two main 
techniques are: 

1. Local-area DGPS. The corrections are derived locally at a ground reference site. As the 
position of the site is accurately known, the satellite inaccuracies can be determined and 
transmitted locally to the user, in this case by line-of-sight VHF data link. The local-area 
DGPS system under development in the United States is called the local-area augmenta- 
tion system (LAAS) and is described below. 

2. Wide-area DGPS. The wide-area correction technique involves networks of data collec- 
tion ground stations. Information is collected at several ground stations which are usually 
located more than 500 miles apart. The correction information derived by each station is 
transmitted to a central location where the satellite corrections are determined. Correc- 
tions are sent to the user by geostationary satellites or other appropriate means. The wide- 
area augmentation system (WAAS) being developed in the United States is outlined 

Note that differential techniques may be applied to any satellite system. For GPS the basic 
accuracy without selective availability is about ±100m as opposed to ±8m when the full 
system is available. The DGPS developments under way in the United States are intended to 
improve the accuracy available to civil users. 

8.4.2 Wide-area Augmentation System (WAAS) 

The operation of WAAS, shown in Figure 8.5, is described as follows: 

1. WAAS is a safety-critical system that augments basic GPS and will be deployed in the 
contiguous United States, Hawaii, Alaska and parts of Canada. 

2. WAAS has multiple wide-area reference stations which are precisely surveyed and 
monitor the outputs from the GPS constellation. 

3. These reference stations are linked to wide-area master stations where corrections are 
calculated and the system integrity assessed. Correction messages are uplinked to 
geostationary earth orbit (GEO) satellites that transmit the corrected data on the 
communications LI band to aircraft flying within the WAAS area of coverage. Effec- 
tively, the GEO satellites act as surrogate GPS satellites. 

4. WAAS improves the GPS accuracy to around ±7 m which is a considerable improvement 
over the 'raw' signal. This level of accuracy is sufficient for Cat I approach guidance. 

Some problems were experienced in initial system tests during 2000. Commissioning of the 
WAAS requires extensive testing to ensure integrity levels, accuracy, etc., and the system 
was finally commissioned during 2004. 






• Wide-area Reference Station (WRS) 
^f Wide-area Master Station (WMS) 

Figure 8.5 Wide-area augmentation system. 

8.4.3 Local-area Augmentation System (LAAS) 

The operation of LAAS is shown in Figure 8.6 and described below: 

1. LAAS is intended to complement WAAS but at a local level. 

2. LAAS works on similar principles except that local reference stations transmit correction 
data direct to user aircraft on VHF. As such, the LAAS coverage is restricted by VHF 
line-of-sight and terrain-masking limitations. 


GPS Receiver 
VHF Transmitter 
Monitor Service 

Figure 8.6 Local-area augmentation system. 



3. LAAS improves the GPS accuracy to about dzlm, close to the higher GPS level of 
accuracy. This level of accuracy is sufficient to permit Cat II and Cat III approaches which 
are described more fully in Chapter 9. 

Implementation is as before, with final deployment in 2006, although these timescales are 
apt to slip, as has the implementation of WAAS. According to present plans it is expected 
that LAAS will be deployed at up to 143 airfields throughout the United States. It was the 
anticipation of LAAS implementation that caused the United States to modify its stance 
upon the implementation of MLS as an approach aid successor to ILS, the space-based GPS 
system being seen as more flexible than ground-based MLS. 

8.5 Integrated Navigation 

Integrated navigation, as the name suggests, employs all the features and systems described 
so far. An integrated navigation solution using a multisensor approach blends the perfor- 
mance of all the navigation techniques already described together with GPS to form a totally 
integrated system. In this case the benefits of the GPS and IN derived data are blended to 
provide more accurate data fusion, in the same way as barometric and IN data are fused 
(Figure 8.7). 

Such an integrated system is a precursor to the introduction of the advanced navigation 
capabilities that will comprise the future air navigation system (FANS). FANS is designed to 
make more efficient use of the existing airspace such that future air traffic increases may be 
accommodated. Some elements of FANS have already been implemented, others will take 
several years to attain maturity. A key prerequisite to achieving a multisensor system is the 
installation of a high-grade flight management system (FMS) to perform the integration of all 
the necessary functions and provide a suitable interface with the flight crew. 




Figure 8.7 Integrated GPS and inertial navigation. 


8.5. 1 Sensor Usage - Phases of Flight 

When assessing which navigation sensors to use for various phases of flight, the navigation 
accuracy, equipment availability and reliability and operational constraints all need to be 
taken into account. The advent of GPS with its worldwide coverage at high levels of 
accuracy has given a tremendous impetus to the navigation capabilities of modern aircraft. 
However, system integrity concerns have meant that the certification authorities have stopped 
short of relying solely on GPS. 

Bearing in mind physical and radio propagation factors, and the relative traffic densities 
for various phases of flight, a number of requirements are specified for the use of GPS. 
Advisory circular AC 90-94 specifies the considerations that apply for the use of GPS as a 
sole or supplementary method of navigation. 

These considerations apply for the following phases of flight: 

1. Oceanic en route. Operation over long oceanic routes means that the aircraft will be 
denied the availability of most of the line-of-sight radio navigation aids such as NDB, 
VOR, TACAN, etc. LORAN-C may be available in some circumstances. The aircraft will 
need to depend upon an approved primary long-range method of navigation. For most 
modern transport aircraft that means equipping the aircraft with a dual- or triple-channel 
INS or ADIRS. Supplementary means such as GPS may be used to update the primary 
method of navigation. Aircraft using GPS under instrument flight rules (IFR) must be 
equipped with another approved long-range navigation system: GPS is not certified as a 
primary and sole means of navigation. Certain categories of GPS equipment may be used 
as one of the approved long-range navigation means where two systems are required. The 
availability of a functioning receiver autonomous integrity monitor (RAIM) capability is 
also important owing to the impact that this has upon GPS integrity. Providing RAIM 
is available, the flight crew need not actively monitor the alternative long-range 
navigation system. 

2. Domestic en route. Once overland, most of the conventional navigation aids may be 
available, unless the aircraft is transiting a wilderness area such as Siberia. For the most 
part, NDB, VOR, TACAN and LORAN-C will be operational and available to supplement 
GPS. These ground-based systems do not have to be used to monitor GPS unless RAIM 
failure occurs. Within the United States, Alaska, Hawaii and surrounding coastal waters, 
IFR operation may be met with independent NDB, VOR, TACAN or LORAN-C 
equipment. This may not necessarily be the case outside the US NAS. 

3. Terminal. GPS IFR operations for the terminal phases of flight should be conducted as for 
normal RNAV operations using the standard procedures: 

• Standard instrument departures (SIDs); 

• Standard terminal arrival routes (STARs); 

• Standard instrument approach procedure (SIAP). 

The normal ground-based equipment appropriate to the phase of flight must be available, 
out, as before, it does not need to be used to monitor GPS unless RAIM fails. 

4. Approach. In the United States an approach overlay programme has been introduced by 
the FAA to facilitate the introduction of instrument approaches using GPS. The key 
features of the GPS overlay programme are described below. 


8.5.2 GPS Overlay Programme 

The GPS overlay programme allows pilots to use GPS equipment to fly existing VOR, VOR/ 
DME, NDB, NDB/DME, TACAN and RNAV non-precision instrument approach proce- 
dures. This facility only applies in US airspace and was introduced in February 1994. The 
approach aid appropriate to the type of approach being flown must be available for use, but 
need not be monitored provided RAIM is available. In April 1994, 'phase III' approaches 
introduced the first GPS specific approaches with GPS specifically included in the title. For 
these approaches the traditional avionics need not be available - either ground-based or 
airborne equipment - provided RAIM is available. For aircraft fitted with GPS without a 
RAIM capability these navigation aids must be available. 

8.5.3 Categories of GPS Receiver 

The different types of GPS are mandated in technical standing order (TSO) C-129a. This 
categorises the different GPS receivers by three major classes: 

1. Class A. This equipment incorporates a GPS receiver and the navigation capability to 
support it. 

2. Class B. This consists of GPS equipment providing data to an integrated navigation 
system such as a flight management system or an integrated multisensor navigation 

3. Class C. This includes equipment comprising GPS sensors which provide data to an 
autopilot or flight director in order to reduced flight technical errors. 

This classification therefore categorises the GPS equipment types according to function. 
TSO C-129a also specifies which class of equipment may be used for the typical flight 
phases described above. It also specifies whether the RAIM function is to be provided by the 
GPS or the integrated system. 

8.6 Flight Management System 

It is clear from the foregoing description of the aircraft navigation functions that navigation 
is a complex task and becoming more so all the while. FMS functionality has increased 
rapidly over the last decade, and many more enhancements are in prospect as the future 
features required by FANS are added. A typical FMS will embrace dual computers and dual 
Multifunction control and display units (MCDUs) as shown in Figure 8.8. In military bomber 
and transport aircraft the system implementation is likely to be in the form portrayed in this 
figure. For military fighter aircraft the functions will be similar but embedded in the avionics 
system navigation computers and mission computers as appropriate. 

Figure 8.8 is key to depicting the integration of the navigation functions described above. 
System sensor inputs, usually in dual-redundant form for reasons of availability and integrity, 
are shown on the left. These are: 

• Dual INS/IRS; 

• Dual navigation sensors: VOR/DME, DME/DME, etc; 



Air Data 

Clock & 
Fuel Sensors 





Autopilot & 

Figure 8.8 Typical flight management system (FMS). 

• Dual GNSS sensors - usually GPS; 

• Dual air data sensors; 

• Dual inputs from on-board sensors relating to fuel on-board and time. 

These inputs are used by the FMS to perform the necessary navigation calculations and 
provide information to the flight crew via a range of display units: 

• Electronic flight instrument system (EFIS); 

• Communications control system; 

• Interface with the autopilot/flight director system to provide the flight crew with flight 
direction or automatic flight control in a number of predefined modes. 

The FMS-crew interface is shown in Figure 8.9. The key interface with the flight crew is via 
the following displays: 

1. Captain's and first officer's navigation displays (ND), part of the electronic flight 
instrument system (EFIS). The navigation displays may show information in a variety 
of different ways. 

2. Control and display units 1 and 2, part of the FMS. The CDUs both display information 
and act as a means for the flight crew manually to enter data. 

The FMS computers perform all the necessary computations and show the appropriate 
navigation parameters on the appropriate display. The navigation displays show the 
navigation and steering information necessary to fly the intended route. These are colour 
displays and can operate in a number of different formats, depending upon the phase of 



Compass Rose 
Heading Lubber 
DME Distance 
VOR bearing 
ADF Bearing 
Intended Flight Path 

NAV Mode 
ILS Frequency 
VHF Frequency 
ILS Frequency 
Time to Waypoint 











Figure 8.9 FMS control and display interface. 

8.6.1 FMSCDU 

The FMS CDU is the key flight crew interface with the navigation system, allowing the 
flight crew to enter data as well as having vital navigation information displayed. A typical 
FMS CDU is shown in Figure 8.10. The CDU has a small screen on which alpha-numeric 
information is displayed, in contrast to the pictorial information displayed on the EFIS 
navigation displays. This screen is a cathode ray tube (CRT) monochrome display in 
early systems; later systems use colour active matrix liquid crystal display (AMLCD) 
(see Chapter 11). The tactile keyboard has alpha-numeric keys in order to allow manual 
entry of navigation data, perhaps inserting final alterations to the flight plan, as well as 
various function keys by which specific navigation modes may be selected. The line keys 
at the sides of the display are soft keys that allow the flight crew to enter a menu-driven 
system of subdisplays to access more detailed information. On many aircraft the CDU is 
used to portray maintenance status and to execute test procedures using the soft keys and the 
menu-driven feature. Finally, there are various annunciator lights and a lighting control 

Examples of the data displayed on the CDU are indicated in Figure 8.11. The CDU 
displays the following parameters using a menu-driven approach: 

1. An ETA waypoint window. This shows the estimated time of arrival (ETA) at the 
waypoint, in this case waypoint 15. 

2. Early/late timing information. This represents the earliest and latest times the aircraft can 
reach the waypoint given its performance characteristics. 

3. Information on the runway - an ILS approach to runway 27. 

4. Wind information for the approach - wind bearing 290. 



Flight Phase 



Alpha Keys 


{jo h h q~i o © b 




Figure 8.10 Typical FMS control and display unit. 





5. Information on the navigation aids being used: VOR, DME and ILS/LOC. 

6. An ANP/RNP window. This compares the actual navigation performance (ANP) of the 
system against the required navigation performance (RNP) for the flight phase and 
navigation guidance being flown. In this case the ANP is 0.15 nm against a RNP of 0.3 nm 
and the system is operating well within limits. 

Runway 27 

ETA Window 
Waypoint 15 

Figure 8.11 Typical FMS CDU display data. 



Nav Data Base: 

- Airways 

- Airports 

- Runways 

- Routes, SIDS, 
STARS etc 

- Procedures 

- Flight Plans 

- Winds 

Performance Model 



Display Data 


Figure 8.12 Top-level FMS functions. 

8.6.2 FMS Functions 

Nav Sensors: 

- INS 

- GPS 


- ILS 

- ADF 

Fuel State 

Air Data 

The functions of the FMS at a top level are shown in Figure 8.12. This diagram gives an 
overview of the functions performed by the FMS computers. These may be summarised as 

1. Navigation computations and display data. All the necessary navigation computations are 
undertaken to derive the navigation or guidance information according to the phase of 
flight and the sensors utilised. This information is displayed on the EFIS navigation 
display or the FMS CDU. Flight director and steering commands are sent to the autopilot 
for the flight director with the pilot in the loop or for the engagement automatic flight 
control modes. 

2. Navigation sensors. INS, GPS, VOR, ILS, ADF, TACAN and other navigation aids 
provide dual-sensor information to be used for various navigation modes. 

3. Air data. The ADCs or ADIRS provides the FMS with high-grade corrected air data 
parameters and attitude information for use in the navigation computations. 

4. Fuel state. The fuel quantity measurement system and the engine-mounted fuel flow- 
meters provide information on the aircraft fuel quantity and engine fuel flow. The 
calculation of fuel use and total fuel consumption is used to derive aircraft and engine 
performance during the flight. When used together with a full aircraft performance model, 
optimum flight guidance may be derived which minimises the fuel consumed. 

5. Sensor fusion and Kalman filter. The sensor information is fused and validated against 
other sources to determine the validity and degree of fidelity of the data. By using a 
sophisticated Kalman filter, the computer is able to determine the accuracy and integrity 
of the navigation sensor and navigation computations and determine the actual navigation 
performance (ANP) of the system in real time. 


6. Communications management. The system passes information to the communication 
control system regarding the communication and navigation aid channel selections that 
have been initiated by the FMS in accordance with the requirements of the flight plan. 

7. Navigation database. The navigation base contains a wide range of data that are relevant 
to the flight legs and routes the aircraft may expect to use. This database will include the 
normal flight plan information for standard routes that the aircraft will fly together with 
normal diversions. It will be regularly updated and maintained. A comprehensive list of 
these items includes: 

• Airways; 

• Airports - approach and departure information, airport and runway lighting, 
obstructions, limitations, airport layout, gates, etc; 

• Runways including approach data, approach aids, category of approach (Cat I or Cat 
II/III) and decision altitudes; 

• Routes, clearance altitudes, SIDS, STARS and other defined navigation data; 

• Procedures including notification of short-term airspace restrictions or special 

• Flight plans with standard diversions. 

• Wind data - forecast winds and actual winds derived throughout flight. 

8. Aircraft performance model. The inclusion of a full performance model adds to the 
systems ability to compute four-dimensional (x, y, z, time) flight profiles and at the same 
time make optimum use of the aircraft energy to optimise fuel use. 

The FMS provides the essential integration of all of these functions to ensure that the overall 
function of controlling the navigation of the aircraft is attained. As may be imagined, this 
does not merely include steering information to direct the aircraft from waypoint to 
waypoint. The FMS also controls the tuning of all the appropriate aircraft receivers to 
navigation beacons and communications frequencies via the communications control units 
and many other functions besides. The flight plan that resides within the FMS memory will 
be programmed for the entire route profile, for all eventualities, including emergencies. More 
advanced capabilities include three-dimensional navigation and the ability to adjust the 
aircraft speed to reach a waypoint within a very small time window (typically ±6 s). The 
various levels of performance and sophistication are summarised in Table 8.1. Military 
aircraft such as the Boeing multirole maritime aircraft (MM A) will be fitted with an FMS 
developed to provide all these features for a civil operator. The FMS will incorporate 
additional specific modes of operation to facilitate performance of the mission, e.g. flying 
mission attack profiles at low level. 

The FMS capabilities will be examined in a little more detail. 

8.6.3 LNAV 

Lateral navigation or LNAV relates to the ability of the aircraft to navigate in two 
dimensions, in other words, the lateral plane. LNAV was the first navigation feature to be 
implemented and involved navigating aircraft to their intended destination without any other 
considerations. LNAV comprises two major implementations: 



Table 8.1 Summary of FMS capabilities 





Four-dimensional navigation 

Full performance based navigation 

Future air navigation system (FANS) or 
global air transport management (GATM) 

The ability to navigate laterally in two dimensions 

The ability to navigate laterally in two dimensions 
plus the ability to navigate in the vertical plane. 
When combined with LNAV, this provides 
three-dimensional navigation 

The ability to navigate in three-dimensions plus 
the addition of time constraints for the 
satisfaction of time of arrival at a waypoint 

The capability of four-dimensional navigation 
together with the addition of an aircraft specific 
performance model. By using cost indexing 
techniques, full account may be made of the 
aircraft performance in real time during flight, 
allowing optimum use of fuel and aircraft energy 
to achieve the necessary flight path 

The combination of the full performance model 
together with all the advantages that FANS or 
GATM will confer, eventually enabling the 
concept of 'free flight' 

• Airway navigation; 

• Area navigation or RNAV. 

8.6.4 Airway Navigation 

Airway navigation is defined by a predetermined set of airways which are based primarily on 
VOR stations, although some use NDB stations. In the United States these airways are 
further categorised depending upon the height of the airway: 

1. Airways based on VOR from 120 ft above the surface to 18 000 ft above mean sea level 
(MSL) carry the V prefix and are called victor airways. 

2. Airways using VOR from 18 000 ft MSL to 45 000 ft MSL are referred to as jet routes. 

Each VOR used in the route system is called either a terminal VOR or a low- or high-altitude 
en-route VOR. Terminal VORs are used in the terminal area to support approach and 
departure procedures and are usable up to ~25nm; they are not to be used for en-route 
navigation. Low-altitude en-route VORs have service volumes out to a range of 40 nm and 
are used up to 18 000 ft on victor airways. High-altitude VORs support navigation on jet 
routes and their service volume may extend to a range of ~200 nm from the ground station. 
Clearly, the range of the VOR beacons is limited by line-of-sight propagation considerations, 
as has already been described. 

Airway width is determined by the navigation system performance and depends upon 
error in the ground station equipment, airborne receiver and display system. 


8.6.5 Area Navigation 

Many aircraft possess an area navigation (RNAV) capability. The on-board navigation 
together with the FMS can navigate along a flight path containing a series of waypoints that 
are not defined by the airways. Navigation in these situations is not confined to VOR beacons 
but may use a combination of VOR, DME, LORAN-C, GPS and/or INS. Random routes 
have the advantage that they may be more direct than the airway system and also that they 
tend geographically to disperse the aircraft away from the airway route structure. Advisory 
circular AC 90-45A defines the regulations that apply to aircraft flying two-dimensional 
RNAV in IFR conditions in the US national airspace system. Advisory circular AC 90-45 A is 
the original guidance on the use of RNAV within the United States, while advisory circular 
AC 20-130A is a more recent publication on navigation or FMS systems using multiple 
sensors - including GPS - and is probably more relevant to the sophisticated FMS in use 

8.6.6 VNAV 

Following on from the LNAV and RNAV capabilities, vertical navigation (VNAV) proce- 
dures were developed to provide three-dimensional guidance. Present VNAV systems use 
barometric altitude, as it will be recalled that the GPS satellite geometry does not generally 
provide accurate information in the vertical direction. Whereas DGPS systems such as 
WAAS will address and overcome this issue, these systems will not be available for some 
time. Advisory circular AC 90-97 provides guidance for the use of VNAV guidance in 
association with RNAV instrument approaches with a VNAV decision altitude (DA). One 
disadvantage of using barometric means to provide the VNAV guidance function is the non- 
standard nature of the atmosphere. Therefore, VNAV approaches embrace a temperature 
limit below which the use of VNAV decision height is not permitted. If the temperature on a 
particular day falls below this limit, then the flight crew must instead respect the published 
LNAV minimum decision altitude (MDA). 

8.6. 7 Four-Dimensional Navigation 

The combination of LNAV and VNAV provides a three-dimensional navigation capability. 
However, in a busy air traffic management situation the element of time is equally important. 
A typical modern FMS will have the capability to calculate the ETA to a specific waypoint 
and ensure that the aircraft passes through that point in space within ±6 s of the desired time. 
Furthermore, calculations can be made in response to an air traffic control enquiry as to when 
the aircraft can reach an upcoming waypoint. By using information regarding the aircraft 
performance envelope, the FMS can perform calculations that determine the earliest and the 
latest possible time within which the aircraft can reach the waypoint. The ability to 
determine this time window can be of great use in helping the air traffic controller to 
maintain steady traffic flow during periods of high air traffic density. 

8.6.8 Full Performance Based Navigation 

If the FMS contains a full performance model provided by the aircraft manufacturer, then 
even more detailed calculations may be performed. By using the aircraft velocity and other 


dynamic parameters, it is possible to compute the performance of the aircraft over very small 
time increments. By using this technique, and provided that the sensor data are sufficiently 
accurate, the future dynamic behaviour of the aircraft may be accurately predicted. Using 
this feature, and knowing the four-dimensional trajectory and gate speeds that are detailed in 
the flight plan, the aircraft can calculate the optimum trajectory to meet all these require- 
ments while conserving energy and momentum and ensuring minimum fuel burn. When this 
capability is combined with the increasing flexibility that FANS will provide, further 
economies will be possible. Today, most FMS systems are being developed with these 
emerging requirements in mind such that future implementation will depend upon system 
software changes and upgrades rather than aircraft equipment or architecture modifications. 

8.6.9 FMS Procedures 

Although the foregoing explanations have concentrated on performance enhancements, the 
assistance that the FMS provides the flight crew in terms of procedural displays cannot be 
forgotten. Typical examples include: 

• Standard instrument departure (SID); 

• En-route procedures; 

• Standard terminal arrival requirements (STAR); 

• ILS approach. 

Examples of these procedures are given in the companion volume (Moir and Seabridge, 

8.6.10 Traffic Collision and Avoidance System (TCAS) 

The TCAS combines the use of the ATC mode S transponder with additional computing and 
displays to provide warning of the proximity of other aircraft within the air traffic control 
system. The operation of ATC mode S and the TCAS is described in detail in Chapter 7. 

8.6.11 GPWS and EGPWS 

While the TCAS is designed to prevent air-to-air collisions, the ground proximity warning 
system (GPWS) is intended to prevent unintentional flight into the ground. Controlled flight 
into terrain (CFIT) is the cause of many accidents. The term describes conditions where the 
crew are in control of the aircraft, but, owing to a misplaced sense of situational awareness, 
they are unaware that they are about to crash into the terrain. The GPWS takes data from 
various sources and generates a series of audio warnings when a hazardous situation is 

The terrain awareness and warning system (TAWS) embraces the overall concept of 
providing the flight crew with prediction of a potential controlled flight into terrain. The new 
term is a generic one since the ground proximity warning system (GPWS) and enhanced 
GPWS became associated mainly with the Allied Signal (now Honeywell) implementation. 
The latest manifestation is designed to provide the crew with an improved situational 
awareness compared with previous systems. The FAA is presently in the process of 


specifying that turbine-equipped aircraft with six seats or more will be required to be 
equipped with a TAWS by 2003. Advisory circular AC 25-23 addresses the airworthiness 
requirements associated with the TAWS. 

The GPWS/TAWS uses radar altimeter information together with other information 
relating to the aircraft flight path. Warnings are generated when the following scenarios 
are unfolding: 

• Flight below the specified descent angle during an instrument approach; 

• Excessive bank angle at low altitude; 

• Excessive descent rate; 

• Insufficient terrain clearance; 

• Inadvertent descent after take-off; 

• Excessive closure rate to terrain - the aircraft is descending too quickly or approaching 
higher terrain. 

Inputs are taken from a variety of aircraft sensors and compared with a number of algorithms 
that define the safe envelope within which the aircraft is flying. When key aircraft dynamic 
parameters deviate from the values defined by the appropriate guidance algorithms, 
appropriate warnings are generated. 

The installation of GPWS equipment for all airliners flying in US airspace was mandated 
by the FAA in 1974, since when the number of CFIT accidents has dramatically decreased. 

More recently, enhanced versions have become available. The EGPWS offers a much 
greater situational awareness to the flight crew as more quantitative information is provided, 
together with earlier warning of the situation arising. It uses a worldwide terrain database 
which is compared with the present position and altitude of the aircraft. Within the terrain 
database the earth's surface is divided into a grid matrix with a specific altitude assigned to 
each square within the grid representing the terrain at that point. 

The aircraft intended flight path and manoeuvre envelope for the prevailing flight 
conditions are compared with the terrain matrix and the result is graded according to the 
proximity of the terrain, as shown in Figure 8. 13: 


High Red 
High Yellow 

Medium Yellow 

Medium Green 
Light Green 

Clear Screen 

Figure 8.13 Principle of operation of the EGPWS (TAWS). 


Terrain responses are graded as follows: 

• No display for terrain more than 2000 ft below the aircraft; 

• Light-green dot pattern for terrain between 1000 and 2000 ft below the aircraft; 

• Medium-green dot pattern for terrain between 500 and 1000 ft below the aircraft; 

• Medium-yellow dot pattern for terrain between 1000 ft above and 500 ft below the 

• Heavy-yellow display for terrain between 1000 and 2000 ft above the aircraft; 

• Heavy-red display for terrain more than 2000 ft above the aircraft. 

This type of portrayal using coloured imagery is very similar to that for the weather radar and 
is usually shown on the navigation display. It is far more informative than the audio 
warnings, given by earlier versions of GPWS. The EGPWS also gives audio warnings, but 
much earlier than those given by the GPWS. The earlier warnings, together with the 
quantitative colour display, give the flight crew a much better overall situational awareness in 
respect of terrain and more time to react positively to their predicament than did previous 

8.7 Navigation Aids 

As aviation began to expand in the 1930s, the first radio navigation systems were developed. 
Initially, these were installed at the new growing US airports, and it is interesting to note that 
the last of these early systems was decommissioned as recently as 1979. 

One of the most prominent was the 'radio range' system developed in Italy by Bellini and 
Tosi, which was conceived as early as 1907. The operation of the Bellini-Tosi system relied 
upon the transmission of morse characters A (dot-dash) and N (dash-dot) in four evenly 
spaced orthogonal directions. When flying the correct course, the A and N characters 
combined to produce a humming noise which the pilot could detect in his earphones. 
Deviation from the desired course would result in either the A or N characters becoming 
most dominant, signifying the need for corrective action by turning left or right as 

Following WWII, the International Civil Aviation Organisation (ICAO) produced inter- 
national standards that led to the definition of the very high-frequency omnirange (VOR) 
system which is in widespread use today and is described below. 

The use of radio navigation aids is important to military aircraft as much of their operation 
involves sharing the airspace with civil users. Therefore, military aircraft, especially those 
adopted from a civil aircraft platform, will utilise a similar suite on navigation aids. Typical 
aids are: 

1. Navigation aids: 

• Automatic direction finding (ADF); 

• VHF omnirange (VOR); 

• Distance-measuring equipment (DME); 

• Tactical air navigation (TACAN); 


• Long range navigation (LOR AN). 


2. Landing aids: 

• Instrument landing system (ILS); 

• Microwave landing system (MLS). 

8. 7. 1 Automatic Direction Finding 

Automatic direction finding (ADF) involves the use of a loop direction finding technique to 
establish the bearing to a radiating source. This might be to a VHF beacon or a non-distance 
beacon (NDB) operating in the 200-1600 kHz band. Non-directional beacons, in particular, 
are the most prolific and widely spread beacons in use today. The aircraft ADF system 
comprises integral sense and loop antennas which establish the bearing of the NBD station to 
which the ADF receiver is tuned. The bearing is shown on the radio magnetic indicator 
(RMI) in the analogue cockpit of a 'classic' aircraft or more likely on the electronic flight 
instrument system (EFIS), as appropriate. ADF is used by surveillance aircraft such as MPA 
on an air sea rescue mission to home on to a personal locator beacon used by downed airmen 
or installed in life rafts. 

8.7.2 Very High-frequency Omnirange (VOR) 

The VOR system was accepted as standard by the United States in 1946 and later adopted by 
the International Civil Aviation Organisation (ICAO) as an international standard. The 
system provides a widely used set of radio beacons operating in the VHF frequency band 
over the range 108-1 17.95 MHz with a 100 kHz spacing. Each beacon emits a morse code 
modulated tone which may be provided to the flight crew for the purposes of beacon 

The ground station radiates a cardioid pattern which rotates at 30 r/min generating a 30 Hz 
modulation at the aircraft receiver. The ground station also radiates an omnidirectional signal 
which is frequency modulated with a 30 Hz reference tone. The phase difference between the 
two tones varies directly with the bearing of the aircraft. At the high frequencies at which 
VHF operates there are no sky wave effects and the system performance is relatively 
consistent. VOR has the disadvantage that it can be severely disrupted by adverse weather - 
particularly by electrical storms - and as such it cannot be used as a primary means of 
navigation for a civil aircraft. 

Overland in the North American continent and Europe, VOR beacons are widely situated 
to provide an overall coverage of beacons. Usually these are arranged to coincide with major 
airway waypoints and intersections in conjunction with DME stations - see below - such 
that the aircraft may navigate for the entire flight using the extensive route/beacon structure. 
By virtue of the transmissions within the VHF band, these beacons are subject to the line-of- 
sight and terrain-masking limitations of VHF communications. Advisory circular AC 00- 
31A lays out a method for complying with the airworthiness rules for VOR/DME/TACAN. 

8.7.3 Distance-measuring Equipment (DME) 

Distance-measuring equipment (DME) is a method of pulse ranging used in the 960- 
1215 MHz band to determine the distance of the aircraft from a designated ground station. 



DME Equipment 



I 125.5 | 

( > 




Reply Pulses 

Interrogation Pulses 


Ground DME Station 


Figure 8.14 DME principle of operation. 

The aircraft equipment interrogates a ground-based beacon and, upon the receipt of 
retransmitted pulses, unique to the on-board equipment, is able to determine the range to 
the DME beacon (Figure 8.14). DME beacons are able to service requests from a large 
number of aircraft simultaneously but are generally understood to have the capacity to 
handle ~200 aircraft at once. Specified DME accuracy is ±3% or ±0.5 nm, whichever is the 
greater (advisory circular AC 00-31 A). 

DME and TACAN beacons are paired with ILS/VOR beacons throughout the airway route 
structure in accordance with the table set out in Appendix 3 of advisory circular AC 00-31 A. 
This is organised such that aircraft can navigate the airways by having a combination of 
VOR bearing and DME distance to the next beacon in the airway route structure. A more 
recent development - scanning DME - allows the airborne equipment rapidly to scan a 
number of DME beacons, thereby achieving greater accuracy by taking the best estimate of a 
number of distance readings. This combination of VOR/DME navigation aids has served the 
aviation community well in the United States and Europe for many years, but it does depend 
upon establishing and maintaining a beacon structure across the land mass or continent being 
covered. New developments in third- world countries are more likely to skip this approach in 
favour of a global positioning system (GPS), as described later in the chapter. 

8.7 A TACAN 

Tactical air navigation (TACAN) is military omnibearing and distance-measuring equipment 
with similar techniques for distance measurement as DME. The bearing information is 
accomplished by amplitude modulation achieved within the beacon which imposes 15 and 
135 Hz modulated patterns and transmits this data together with 15 and 135 Hz reference 



pulses. The airborne equipment is therefore able to measure distance using DME interroga- 
tion techniques while using the modulated data to establish bearing. 

TACAN beacons operate in the frequency band 960-1215 MHz as opposed to the 108- 
118 MHz used by DME. This means that the beacons are smaller, making them suitable for 
shipborne and mobile tactical use. Some airborne equipment have the ability to offset to a 
point remote from the beacon which facilitates recovery to an airfield when the TACAN 
beacon is not co-located. TACAN is reportedly accurate to within ±1% in azimuth and 
±0.1 nm in range, so it offers accuracy improvements over VOR/DME. 

TACAN also has the ability to allow aircraft to home on to another aircraft, a feature that 
is used in air-to-air refuelling to enable aircraft to home on to the donor tanker. 

8.7.5 VORTAC 

As most military aircraft are equipped with TACAN, some countries provide VORTAC 
beacons which combine VOR and TACAN beacons. This allows interoperability of military 
and civil air traffic. Military operators use the TACAN beacon while civil operators use the 
VOR bearing and TACAN (DME) distance-measuring facilities. This is especially helpful 
for large military aircraft, such as transport or surveillance aircraft, since they are able to use 
civil air lanes and operational procedures during training or on transit between theatres of 

8.7.6 Hyberbolic Navigation Systems - LORAN-C 

Hyperbolic navigation systems - of which long range navigation (LORAN) is the most 
noteworthy example - operate upon hyperbolic lines of position rather than circles or radial 
lines. Figure 8.15 illustrates the principle of operation of a hyperbolic system in a very 
elementary manner. This shows hyperbolic solid lines which represent points that are 
equidistant from the two stations. These points will have the same time difference between 
the arrival of signals from the blue-master and blue-slave stations (the term secondary station 

Figure 8.15 Principle of operation of a hyperbolic navigation system. 



is probably a better and more accurate description). This in itself will not yield position, but, 
if a second pair of stations is used - angled approximately 45° to the first - shown as dashed 
lines, then position can be obtained. The relative positioning of the lines in this dual-chain 
example shows that three outcomes are possible: 

1. At point A the lines cross at almost 90°, and this represents the most accurate fix. 

2. At point B the lines cross at a much more acute angle and the result is a larger error 

3. At point C there are two possible solutions and an ambiguity exists that can only be 
resolved by using a further station. 

LORAN-C is the hyperbolic navigation system in use today and was conceived in principle 
around the beginning of WWII. Worldwide coverage existed in 1996 and new facilities were 
being planned in the late 1990s. LORAN operates in the frequency band 90-110kHZ as a 
pulsed system which enables the ground wave to be separated from the sky wave, the ground 
wave being preferred. A LORAN chain will comprise at least three stations, one being 
nominated as the master. The time difference of arrival between the master and slaves allows 
position to be determined. Each of the stations in a chain transmits unique identifiers which 
allow the chain to be identified. A typical example of a LORAN-C chain is shown in 
Figure 8.16 which shows the north-eastern US chain. 

Within the defined area of coverage of the chain, LORAN-C will provide a user with a 
predictable absolute accuracy of 0.25 nm. A typical chain will have over 1000 nm operating 
range coverage. LORAN-C is also capable of relaying GPS positional error within the 
transmissions. LORAN-C is expected to remain in commission until at least 2008. Advisory 
circular AC 20-121 A provides information to assist with the certification of LORAN-C 
navigation systems for use within the United States and Alaska. 




M - Seneca 
W - Caribou 
X - Nantucket 
Y - Carolina Beach 
Z - Dana 

Figure 8.16 Typical LORAN-C chain - north-eastern United States. 



8.7.7 Instrument Landing System 

The instrument landing system (ILS) is an approach and landing aid that has been in 
widespread use since the 1960s and 1970s. The main elements of the ILS include: 

1. A localiser antenna centred on the runway to provide lateral guidance. A total of 40 
operating channels are available within the 108-1 12 MHz band. The localiser provides 
left and right lobe signals which are modulated by different frequencies (90 and 
150 Hz) such that one signal or the other will dominate when the aircraft is off the 
runway centre-line. The beams are arranged such that the 90 Hz modulated signal will 
predominate when the aircraft is to the left, while the 150 Hz signal will be strongest to 
the right. The difference in signal is used to drive a cross-pointer deviation needle such 
that the pilot is instructed to 'fly right' when the 90 Hz signal is strongest and 'fly left' 
when the 150 Hz signal dominates. When the aircraft is on the centre-line, the cross- 
pointer deviation needle is positioned in the central position. This deviation signal is 
proportional to azimuth out to ±5° of the centre-line. 

2. A glideslope antenna located beside the runway threshold to provide lateral guidance. 
Forty operating channels are available within the frequency band 329-335 MHz. As for 
the localiser, two beams are located such that the null position is aligned with the desired 
glideslope, usually set at a nominal 3°. In the case of the glideslope, the 150 Hz modulated 
signal predominates below the glideslope and the 90 Hz signal is stronger above. When 
the signals are balanced, the aircraft is correctly positioned on the glideslope and the 
glideslope deviation needle is positioned in a central position. As for the localiser needle, 
pilots are provided with 'fly up' or fly down' guidance to help them to acquire and 
maintain the glideslope (see Figure 8.17 for the general arrangement of the ILS). 
Figure 8.18 illustrates how guidance information is portrayed for the pilot according 
to the aircraft position relative to the desired approach path. On older aircraft this would 

Glide Slope 




[90 Hz] 

ILS Characteristics: 


Localiser transmitting between 108- 

Glide Slope transmitting between 

Audible Morse code tone for 

'Hard' Pairing of ILS Localiser, Glide 
Slope and associated DME to ease 
flight crew workload 

Figure 8.17 ILS glideslope and localiser. 



Right of 

i N Centre 

v Line 

Glide Slope 








Figure 8.18 ILS guidance display. 

be shown on the compass display, while on modern aircraft with digital cockpits this 
information is displayed on the primary flight display (PFD). The ILS localiser, glideslope 
and DME channels are paired such that only the localiser channel needs to be tuned for all 
three channels to be correctly aligned. 

3. Marker beacons are located at various points down the approach path to give the pilot 
information as to what stage on the approach has been reached. These are the outer, 
middle and inner markers. Location of the marker beacons are: 

• Outer marker approximately 4-7 nm from the runway threshold; 

• Middle marker ~3000 ft from touchdown; 

• Inner marker ~ 1000 ft from touchdown. 

The high approach speeds of most modern aircraft render the inner marker almost 
superfluous and it is seldom used. 

4. The marker beacons are all fan beams radiating on 75 MHz and provide different morse 
code modulation tones which can be heard through the pilot's headset. The layout of the 
marker beacons with respect to the runway is as shown in Figure 8.19. The beam pattern is 
±40° along track and ±85° across track. The overall audio effect of the marker beacons is 
to convey an increasing sense of urgency to the pilot as the aircraft nears the runway 

A significant disadvantage of the ILS system is its susceptibility to beam distortion and 
multipath effects. This distortion can be caused by local terrain effects, large man-made 



+/- 85 



- Runway 


ILS Markers: 

All markers transmit on 75MHz 

Outer marker: 

400Hz- 2 tones/sec 

Middle marker: 

1300Hz- dash-dot /sec 

Inner marker: 

3000Hz- six dots/sec 



-3000 feet- 

— 4 to 7 nautical miles- 

Figure 8.19 ILS approach markers. 

structures or even taxiing aircraft which can cause unacceptable beam distortion, with the 
glideslope being the most sensitive. At times on busy airfields and during periods of limited 
visibility, this may preclude the movement of aircraft in sensitive areas, which in turn can 
lead to a reduction in airfield capacity. More recently, interference by high-power local FM 
radio stations has presented an additional problem, although this has been overcome by 
including improved discrimination circuits in the aircraft ILS receiver. 

8.7.8 Microwave Landing System (MLS) 

The microwave landing system (MLS) is an approach aid that was conceived to redress some 
of the shortcomings of the ILS. The specification of a time-reference scanning beam MLS 
was developed through the late 1970s/early 1980s, and a transition to the MLS was envisaged 
to begin in 1998. However, with the emergence of satellite systems such as the GPS there 
was also a realisation that both the ILS and MLS could be rendered obsolete when such 
systems reach maturity. In the event, the US civil community is embarking upon higher- 
accuracy developments of the basic GPS system: the wide-area augmentation system 
(WAAS) and local-area augmentation system (LAAS) have already been outlined. In 
Europe, the United Kingdom, the Netherlands and Denmark have embarked upon a modest 
programme of MLS installations at major airports. 

The MLS operates in the frequency band 5031.0-5190.7 MHz and offers some 200 
channels of operation. It has a wider field of view than the ILS, covering ±40° in azimuth 
and up to 20° in elevation, with 15° useful range coverage. Coverage is out to 20 nm for a 
normal approach and up to 7 nm for back azimuth/go-around. The co-location of a DME 
beacon permits three-dimensional positioning with regard to the runway, and the combina- 
tion of higher data rates means that curved-arc approaches may be made, as opposed to the 
straightforward linear approach offered by the ILS. This offers advantages when operating 



20 degrees -v 20,000 Feet 

15 degrees -v^ ^<^^^^ 

/ 20 nm 


3 degrees -/ 

" -40 Approach ELEVATION 
degrees Azimuth 

/ - 1 A . 

.'\ +40 


7 nm / ^Af 

\ degrees 


-'"' /X 

^ / X 


V / 20 nm 


Figure 8.20 Microwave landing system coverage. 

into airfields with confined approach geometry and tactical approaches favoured by the 
military. For safe operation during go-around, precision DME (P-DME) is required for a 
precise back azimuth signal. 

A groundbased MLS installation comprises azimuth and elevation ground stations, each of 
which transmits angle and data functions which are frequency shift key (FSK) modulated 
and which are scanned within the volume of coverage already described. The MLS scanning 
function is characterised by narrow beam widths of around 1-2° scanning at high slew rates. 
Scanning rates are extremely high at 20 000 deg/s which provides data rates that are around 
10 times greater than is necessary to control the aircraft. These high data rates are very useful 
in being able to reject spurious and unwanted effects due to multiple reflections, etc. 

Typical coverage in azimuth and elevation for an MLS installation is shown in Figure 8.20. 

8.8 Inertial Navigation 

8.8. 1 Principles of Operation 

The availability of inertial navigation systems (INS) to the military aviation community 
during the early 1960s added another dimension to the navigation equation. Now flight crew 
were able to navigate by autonomous means using an on-board INS with inertial sensors. By 
aligning the platform to earth-referenced coordinates and present position during initialisa- 
tion, it was now possible to fly for long distances without relying upon TACAN or VOR/ 
DME beacons overland or hyperbolic navigation systems elsewhere. Waypoints could be 
specified in terms of latitude and longitude as arbitrary points on the globe, more suited to 
the aircraft's intended flight path rather than a specific geographic feature or point in a radio 
beacon network (Figure 8.13). This offered an enormous increase in operational capability as 
mission requirements could be defined and implemented using an indigenous on-board 




Attitude Rates: 









Axis Set 














1 ' 





Figure 8.21 Principles of inertial navigation. 

sensor with no obvious means of external detection except when the need arose to update the 

The principles of inertial navigation depend upon the arrangement of inertial sensors such 
as gyroscopes and accelerometers in a predetermined orthogonal axis set. The gyroscopes 
may be used to define attitude or body position and rates. 

The output from the accelerometer sensor set is integrated to provide velocities, and then 
integrated again to provide distance travelled (Figure 8.21). First in the military field and 
then in the commercial market place, inertial navigation systems (INS) became a preferred 
method for achieving long-range navigation such that by the 1960s the technology was well 

The specifications in force at this time also offered an INS solution to North Atlantic 
crossings as well as the dual-Doppler solution previously described. The inertial solution 
required serviceable dual INS and associated computers to be able to undertake the crossing. 
There were also limitations on the latitudes at which the ground alignment could be 
performed - 76° north or south - as attaining satisfactory alignment becomes progressively 
more difficult the nearer to the poles the INS becomes. 

For civil operators, accuracy requirements set forth requirements for operating an INS as a 
sole means of navigation for a significant portion of the flight. These requirements were 
described earlier in the chapter. 

8.8.2 Stand-alone Inertial Navigation System 

For reasons of both availability and accuracy, systems were developed with dual and triple 
INS installations. A typical triple INS installation of the type used by modern wide-body 










Figure 8.22 Typical triple INS system. 



INS #3 

INS #2 

I r __I__. 
I | GPS 

| L 










commercial transport aircraft is presented in Figure 8.22, showing three INS units integrated 
with the other major systems units. This type of system would be representative of an INS 
installation of a large aircraft before the availability of satellite sensors in the 1990s. By this 
time the gimballed IN platform would have been replaced by a more reliable strapdown 
system similar to the Litton LTN-92 system. 

This integrated system comprised the following units: 

1. Dual sensors: 

• VOR for bearing information; 

• DME for range information; 

• Air data computer (ADC) for air data; 

• Provision for a dual GPS interface. 

2. Controls and displays: 

• Control and display unit (CDU); 

• Electronic flight instrument system/flight director (EFIS/FD); 

• Mode selector unit. 

3. Other major systems receiving INS data for stabilisation or computation: 

• Weather or mission radar; 

• Flight management system (FMS); 

• Autopilot. 

The weight of this system, which comprised three LTN-92 platforms with back-up battery 
power supplies, two CDUs and two mode selector units, was in the region of 234 lb. 




■► Time 

Figure 8.23 Sensor fusion of air data and inertial sensors. 

By integrating the air data information with inertially derived flight information, the best 
features of barometric and inertial systems can be combined. Figure 8.23 illustrates the 
principle of sensor fusion where the short-term accuracy of inertial sensors is blended or 
fused with the long-term accuracy of air data or barometric sensors. 

Means of taking external fixes were evolved so that longer-term inaccuracies could be 
corrected by updating the INS position during long flights. Some fighter aircraft systems 
such as Tornado also added a Doppler radar such that Doppler-derived data could be 
included in the navigation process. The availability of on-board digital computers enabled 
statistical Kalman filtering techniques to be used to calculate the best estimated position 
using all the sensors available. 

The fundamental problem with the INS is the long-term and progressive accrual of 
navigation error as the flight proceeds, and, irrespective of the quality or type of the 
gyroscopes used, this fundamental problem remains. 

8.8.3 Air Data and Inertial Reference Systems (ADIRS) 

The system illustrated in Figure 8.22 utilises stand-alone ADCs; however, the use of ADCs in 
many new large aircraft systems was superceded by the introduction of air data modules 
(ADMs) in the late 1980s. The new integrated air data and inertial reference system (ADIRS) 
developed in the early 1980s combined the computation for air data and inertial parameters 
in one multichannel unit. As large civil plaftforms are increasingly adapted for use in the 
transport, air refuelling, anti-surface warfare (ASuW) or surveillance roles, this 'commer- 
cial' implementation is finding use in military applications. 

Taking the B777 as an example, the primary unit is an air data and inertial reference unit 
(ADIRU) which provides the main source of air data and inertial information. This unit is 
supported by an attitude and heading reference system (AHARS) which on the B777 is 



Rate Gyro 
Package (6) 

Package (6) 

Sensor Processing 



L C R 

Data Buses 

/ Skewed 

Sensor Set 

Figure 8.24 B777 ADIRU. 

called the secondary attitude air data reference unit (SAARU). This provides secondary 
attitude and air data information should the primary source, the ADIRU, become totally 

The B777 ADIRU is shown in Figure 8.24. There are six laser rate gyros (LRGs) and six 
accelerometers included in the unit. It can be seen that both sets of sensors are arranged in a 
hexad skew-redundant set in relation to an orthogonal axis set. This means that, by resolving 
the output of each of the six sensors in the direction of the axis set, each sensor is able to 
measure an element of the relevant inertial parameter - body rate or acceleration - in each 
axis. This provides a redundant multichannel sensor set with the prospect of achieving higher 
levels of accuracy by scaling and combining sensor outputs. Additionally, the output of 
erroneous sensors may be detected and 'voted out' by the remaining good sensors. This 
multiple-sensor arrangement greatly increases the availability of the ADIRU as the 
performance of the unit will degrade gracefully following the failure of one or more sensors. 
The ADIRU may still be used with an acceptable level of degradation until a replacement 
unit is available or the aircraft returns to base and only has to be replaced following the 
second failure of a like sensor (e.g. second rate or accelerometer sensor). By contrast, the 
failure of a sensor in an earlier three-sensor, orthogonally oriented set would lead to a sudden 
loss of the INS. 

Coupled with the dual-hexagonal sensor arrangement, there are four independent lanes 
of computing within the ADIRU, each of which computes a wide range of navigation 



North velocity 

East velocity 

Ground speed 



Wind speed 

Wind direction 

True heading 

Magnetic heading 

True track angle 

Magnetic track angle 

Drift angle 

Flight path angle 

Inertial altitude 

Computed airspeed 

Mach number 

Altitude rate 


Total air temperature (TAT) 

Static air temperature 

True airspeed 

Static pressure (corrected) 

Impact pressure 

Corrected computed airspeed 

Corrected Mach number 

Corrected total pressure 

Corrected static pressure 

CG longitudinal acceleration 

CG lateral acceleration 

CG normal acceleration 

Flight path acceleration 

Vertical speed 

Roll attitude 

Pitch attitude 

Track angle rate 

Corrected angle of attack (AoA) 

Roll attitude rate 

Pitch attitude rate 

Heading rate 

Body yaw rate 

Body pitch rate 

Body roll rate 

Body longitudinal acceleration 

Body lateral acceleration 

Body normal acceleration 

Finally, the ADIRU interfaces with the remainder of the aircraft systems by means of triple 
flight control ARINC 629 digital data buses: left, centre and right. The unit is provided with 
electrical power from a number of independent sources. 

Further information on typical ADIRS implementations is given in the companion volume 
(Moir and Seabridge, 2003). 

8.8.4 Inertial Platform Implementations 

There are two methods by which the IN function may be achieved. These are as follows: 

1. Gyrostabilised platform. In the gyrostabilised platform the sensing elements - gyroscopes 
and accelerometers - are placed on a platform that is itself stabilised to maintain a fixed 
position in space. This requires fine servomotors and mechanisms to maintain this 
stabilisation, and consequently gyrostabilised platforms are expensive to manufacture 
and tend to be unreliable. All of the earlier platforms were implemented in this fashion 
and many are still in service today. 

2. Strap down or analytical platform. The advent of digital computation and its application to 
avionics applications enabled the introduction of the strapdown or analytical platform. In 
this implementation the sensors are strapped directly on to the body of the vehicle and the 
necessary axis transformations to convert from the vehicle to space axes are performed 
numerically using digital computers. Originally this technique was used for military 
applications; today, virtually all IN platforms work in this way. The benefits of strapdown 






Axis Set 



Axis Set 



Axis Set 




Axis Set 

x(N), y(N), z(N) 







True North 




Axis Set 



Axis Set 



Axis Set 




Axis Set 

x(N), y(N), z(N) 

Vehicle Body 


Axis Set 

x(B), y(B), z(B) 

Strapdown or 'Analytical 1 
Figure 8.25 Inertial platform implementation. 

platforms are that they are easier and cheaper to manufacture and are more reliable as 
they contain none of the servomotors and mechanisms that are a feature of the 
gyrostabilised platform. Consequently, they are more reliable by a factor of around 3. 

Both the gyrostabilised platform and the strapdown platform need to undergo a series of 
axis transformations before they may be used for navigation on the surface of the earth 
(Figure 8.25). These axis transformations are as follows: 

• Space axes to the Greenwich meridian; 

• Greenwich meridian axes to true north; 

• True north axes to great circle; 

• Great circle axes to vehicle body (strapdown only). 

These transformations are fully described in the following pages. 

The initial space axes are shown in Figure 8.26. This initial reference shows a generic axis 
set: X® , Y® , Z® as a set of axes determined in space by the platform. For ease of reference 
it may be assumed that the inertial axis Z® coincides with the earth axes Z^> at the outset. 

8.8.5 Space Axes to the Greenwich Meridian 

Given that Z® and Z^> already coincide, the completion of the earth axis transformation is 
achieved by rotating the X® axis to coincide with the Greenwich meridian by rotating by Q t 
to X^ E \ The bold axes now represent the earth reference set X^ E \ Y^ E \ Z^ (Figure 8.27). 

8.8.6 Earth Axes to Geographic Axes 

As it is unlikely that the navigation task will begin precisely aligned with the earth axes, 
account needs to be taken of where the platform is on the surface of the earth when powered 



Site of Alignment - 
in inertial space 

Site of Alignment 
- on eartl i 

Polar Axis of Earth 
North Pole 

-►y u 

(J) is the latitude 
A- is the longitude 


is the wander azimuth 

Figure 8.26 Initial space reference axis set X®, Y^\ Z®. 

up. This process is also known as platform alignment and will be described separately 
shortly. The alignment process rotates the X axes from X^ to X^ by the angle of longitude. 
The process also aligns Y^ such that it points north by rotating by the angle of latitude <P. 
The outcome is that Z^ is in line with the local earth vertical (Figure 8.28). 

,0 Z (E) 

Polar Axis of Earth 

Site of Alignment - 
in inertial space 

Site of Alignment 
- on eartl 

(J) is the latitude 
A is the longitude 


is the wander azimuth 

Figure 8.27 Space axes to the Greenwich meridian X^ E \ Y^ E \ Z^ E \ 



Site of Alignment - 
in inertial space 

Site of Alignment 
- on eartl 

z ( V E) 
Polar Axis of Earth 


North Pole 

(J) is the latitude 
A is the longitude 


is the wander azimuth 

Figure 8.28 Greenwich meridian axes to geographic axes X^ G \ Y^ G \ Z^ G \ 

8.8.7 Geographic to Great Circle (Navigation) 

By aligning the platform at the wander angle a, the platform may be used to navigate a great 
circle route which represents the shortest possible path between two points on the surface of 
the globe (Figure 8.29). 

Site of Alignment - 
in inertial space 

Site of Alignr tent 
- on earth 

z<V E) 
Polar Axis of Earth 

Cj) is the latitude 

A is the longitude 

^ is the wander azimuth 

Figure 8.29 True north axes to great circle route (navigation) X^ N \ Y^ N \ Z^ N \ 



8.8.8 Great Circle/Navigation Axes to Body Axes (Strapdown) 

In the case of the analytical/strapdown platform, the navigation axes X^ , Y ^ , Z^ have to 
be realigned to the vehicle body axes X^ B \ Y^ B \ Z^ respectively. This is achieved by 
executing the following rotations in turn: 

• Rotating in the yaw axes by W\ 

• Rotating in the pitch axis by 0; 

• Rotating in the roll axis by (P; 

Refer to Figure 8.30. 

8.8.9 Platform Alignment 

The process of platform alignment follows the processes defined in Figure 8.31. The process 
is split into the following phases: 

• Entry of the present position latitude and longitude of the platform into the INS. 

• Platform levelling - both course and fine phases. 

• Gyrocompass alignment. 

During these processes the sensors are used to sense misalignment between the platform and 
the desired platform datum. The platform is inched towards these datums over a period of 
several minutes and, when the datums are reached, the platform is considered to be aligned 
and in a suitable state to be used for navigation. The platform attitudes for alignment 
correspond to the navigation axes X^ G \ Y^ G \ Z^ described in Figure 8.31 above. 

x (N) , y (N) , z (N) are navigation (great 
circle) referenced axes 

V (B) (B) _(B) 

referenced axes 

z ( ' are vehicle body 

7 — o 

* x (N) 





is yaw 

is pitch 


is roll 

*{> Yaw 

Figure 8.30 Great circle/navigation to body axes (strapdown) X^ B \ Y^ B \ Z^ B \ 











JWilitary aircraft (fast alignment) ~ 3 to 8 minutes 
Civil aircraft ~ 10 minutes 

Figure 8.31 Process of platform alignment. 

8. 8. 9. 1 Platform Levelling 

In the levelling phase the platform sensors and servomotors drive the platform to ensure that 
the Z direction corresponds to the local vertical at the point on the earth's surface where the 
platform is located. Now one of the three platform axes is correctly aligned (Figure 8.32). 

8. 8. 9. 2 Gyrocompass Alignment 

Gyroscopic compassing uses a similar technique and commences once the fine alignment 
process is under way. In this case the Y axis is driven to align with true north. In an 
orthogonal axis set, when the Z axis corresponds to the local earth vertical and the Y axis 



X represents the Easterly Axis 
Y represents the Northerly Axis 
Z represents the local Vertical Axis 
Figure 8.32 Aligned platform axes. 







Gyro Dynamic 
range- 10 5 

Accelerometer Strapdown 

Dynamic Range Technology: 

~10 6 

-20-30 Hz 

RLG Technology - 
Dynamic Range 
~10 7 




~ 2000 Hz 

Combined IN 

& Air Data 


• Military 
Civil | 




— >- Military 

IN Technology 


Cheaper, Lower 


More integrated 





Figure 8.33 Historical development of inertial platforms. 

corresponds to true north, the X axis corresponds to east and the platform is aligned and 
ready to perform navigation tasks. 

The accuracy of navigation depends to some degree upon the accuracy of the alignment 
process, so, in general, but also within reason, the longer the alignment, the better is the 
accuracy. Accurate alignments are difficult above ±70° north or south as greater inaccuracies 
are experienced. 

8.8. 10 Historical Perspective - Use of Inertial Platforms 

The historical development of IN platforms is shown in Figure 8.33. Gimballed technology 
came into prominence in the 1960s, followed by strapdown in the early 1980s. In both cases 
the military avionics community was the first to exploit the technology. In the late 1980s and 
early 1990s the civil community developed the integratedADIRS concept which rapidly 
became adopted as the primary means of navigation. Meanwhile, the US military in 
particular were developing the satellite-based global positioning system (GPS). More 
recently, IN/GPS coupling has been adopted which enables the fusion of IN and GPS 
sensors in a similar fashion to baro-IN fusion. Both loosely coupled and tightly coupled 
implementations are commonly used. 

Key attributes of the stabilised and strapdown platforms are as follows: 

Gimballed platform 

Strapdown technology 

Gyro dynamic range ~105 
Accelerometer dynamic range ~106 
Calculations undertaken ~20-30Hz 
Laser ring gyro (LRG) dynamic range 
Digital computing technology 
Calculations undertaken ~2000Hz 



Table 8.2 Typical INS performance and physical characteristics 

Parameter Value 

Navigation accuracy 0.8nm/h 

Velocity accuracy 2.5 ft/s rms 

Pitch/roll accuracy 0.05° rms 

Azimuth accuracy 0.05° rms 

Alignment 3-8 min 

Volume 4-8 ATR/MCU 

Weight 20-30 lb 

Power 30-150 W 

Acceleration capability 30 g 

Angular rate capability 400 deg/s 

MTBF (fighter environment) 3500 h 

MTBF (civil environment) 10 000 h 

The future trend for IN/GPS products is to use cheaper, lower-performance IN sensors, 
smaller packages and with increasing integration. 

The performance and physical characteristics of a typical LRG strapdown performance of 
1996 vintage are summarised in Table 8.2. 

8.9 Global Navigation Satellite Systems 

8.9.1 Introduction to GNSS 

Global navigation techniques came into being from the 1960s through to the 1990s when 
satellites became commonly available. The use of global navigation satellite systems 
(GNSS), to use the generic name, offers a cheap and accurate navigational means to anyone 
who possesses a suitable receiver. Although the former Soviet Union developed a system 
called GLONASS, it is the US ground positioning system (GPS) that is the most widely used. 
The European Community (EC) is developing a similar system called Gallileo which should 
enter service in the 2008-2010 timescale. A comparison of the three systems is given in 
Table 8.3. 

GPS is a US satellite-based radio navigational, positioning and time transfer system 
operated by the Department of Defense (DoD) specifically for military users. The system 
provides highly accurate position and velocity information and precise time on a continuous 
global basis to an unlimited number of properly equipped users. The system is unaffected by 
weather and provides a worldwide common grid reference system based on the earth- fixed 
coordinate system. For its earth model, GPS uses the world geodetic system of 1984 (WGS- 
84) datum. 

The Department of Defense declared initial operational capability (IOC) of the US GPS on 
8 December 1993. The Federal Aviation Administration (FAA) has granted approval for US 
civil operators to use properly certified GPS equipment as a primary means of navigation in 
oceanic and certain remote areas. GPS equipment may also be used as a supplementary 
means of instrument flight rules (IFR) navigation for domestic en-route, terminal operations 
and certain instrument approaches. 


Table 8.3 Comparison of global navigation satellite systems 



Soviet Union launched 1982 

24 satellites (only 10 in orbit in 2000) 
Three planes 
Inclination 64.8° 
Height 19130 km 


United States - early 1990s 


Europe - 2008-2010 

24 satellites (29) 
Six planes 
Inclination 55° 
Height 20 180 km 

30 satellites (27 + 3) 
Three planes 
Inclination 55° 
Height 23 616 km 

8.9.2 Principles of Operation 

The principles of satellite navigation using GPS are illustrated in Figure 8.34. GPS 
comprises three major components as characterised in the figure: 

1. The control segment embraces the infrastructure of ground control stations, monitor 
stations and ground-based satellite dishes that exercise control over the system. 

/ Space 
/ Segment 

L1 [1575.42MHz] 

&L2 [1227.6MHz] 



L1 [1575.42MHz] 
&L2 [1227.6MHz] 

Ground Antennae 

Figure 8.34 Principles of GPS satellite navigation. 


2. The space segment includes the satellite constellation, presently around 25 satellites, that 
forms the basis of the network. 

3. The user segment includes all the users: ships, trucks, automobiles, aircraft and hand-held 
sets. In fact, anyone in possession of a GPS receiver is part of the user segment. 

The baseline satellite constellation downlinks data in two bands: LI on 1575.42 MHz and L2 
on 1227.60MHz. A GPS modernisation programme recently announced will provide a 
second civil signal in the L2 band for satellites launching in 2003 onwards. In addition, a 
third civil signal, L5, will be provided on 1176.45 MHz on satellites to be launched in 2005 
and beyond. Finally, extra signals for military users (Lm) will be included in the LI and L2 
bands for satellites launched in 2005 and beyond. 

GPS operation is based on the concept of ranging and triangulation from a group or 
constellation of satellites in space which act as precise reference points. A GPS receiver 
measures distance from a satellite using the travel time of a radio signal. Each satellite 
transmits a specific code, called course/acquisition (CA), which contains information on the 
position of the satellite, the GPS system time and the health and accuracy of the transmitted 
data. Knowing the speed at which the signal travelled (approximately 186000miles/s) and 
the exact broadcast time, the distance travelled by the signal can be computed from the 
arrival time. 

The GPS constellation of 24 satellites is designed so that a minimum of five are always 
observable by a user anywhere on earth. The receiver uses data from a minimum of four 
satellites above the mask angle (the lowest angle above the horizon at which it can use a 

GPS receivers match the CA code of each satellite with an identical copy of the code 
contained in the receiver database. By shifting its copy of the satellite code in a matching 
process, and by comparing this shift with its internal clock, the receiver can calculate how 
long it took the signal to travel from the satellite to the receiver. The value derived from this 
method of computing distance is called a pseudorange because it is not a direct measurement 
of distance but a measurement derived from time. Pseudorange is subject to several error 
sources; for example, ionospheric and tropospheric delays and multipath. In addition to 
knowing the distance to a satellite, a receiver needs to know the exact position of the satellite 
in space; this is known as its ephemeris. Each satellite transmits information about its exact 
orbital location. The GPS receiver uses this information to establish precisely the position of 
the satellite. Using the calculated pseudorange and position information supplied by the 
satellite, the GPS receiver mathematically determines its position by triangulation. The GPS 
receiver needs at least four satellites to yield a three-dimensional position (latitude, longitude 
and altitude) and time solution. The GPS receiver computes navigational values such as 
distance and bearing to a waypoint, ground speed, etc., by using the known latitude/longitude 
of the aircraft and referencing these to a database built into the receiver. 

8.9.3 Integrity Features 

The GPS receiver verifies the integrity (usability) of the signals received from the GPS 
constellation through a process called receiver autonomous integrity monitoring (RAIM) to 
determine if a satellite is providing corrupted information. At least one satellite, in addition 
to those required for navigation, must be in view for the receiver to perform the RAIM 
function. Therefore, performance of the RAIM function needs a minimum of five satellites in 


view, or four satellites and a barometric altimeter (baro-aiding) to detect an integrity 
anomaly. For receivers capable of doing so, RAIM needs six satellites in view (or five 
satellites with baro-aiding) to isolate the corrupt satellite signal and remove it from the 
navigation solution. 

RAIM messages vary somewhat between receivers; however, generally there are two 
types. One type indicates that there are insufficient satellites available to provide RAIM 
integrity monitoring. Another type indicates that the RAIM integrity monitor has detected a 
potential error that exceeds the limit for the current phase of flight. Without the RAIM 
capability, the pilot has no assurance of the accuracy of the GPS position. Areas exist where 
RAIM warnings apply and which can be predicted - especially at higher latitudes - and this 
represents one of the major shortcomings of GPS and the reason it cannot be used as a sole 
means of navigation. 

8.9.4 GPS Satellite Geometry 

The geometry of the GPS satellites favours accurate lateral fixes. However, because a 
number of the visible satellites may be low in the sky, determination of vertical position is 
less accurate. Baro-aiding is a method of augmenting the GPS integrity solution by using a 
non-satellite input source to refine the vertical (height) position estimate. GPS-derived 
altitude should not be relied upon to determine aircraft altitude since the vertical error can be 
quite large. To ensure that baro-aiding is available, the current altimeter setting must be 
entered into the receiver as described in the operating manual. 

GPS offers two levels of service: the standard positioning service (SPS) and the precise 
positioning service (PPS). The SPS provides, to all users, horizontal positioning accuracy of 
100 m or less with a probability of 95% 300 m with a probability of 99.99%. The PPS is more 
accurate than the SPS; however, this is intended to have a selective availability function 
limiting access to authorized US and allied military, federal government and civil users who 
can satisfy specific US requirements. At the moment, the selective availability feature is 
disabled, making the PPS capability available to all users pending the availability of differential 
GPS (DGPS) solutions to improve the SPS accuracy. This step has been taken pending the 
development of differential or augmented GPS systems which will provide high accuracy to 
civil users while preserving the accuracy and security that military users demand. 

The basic accuracy without selective availability is about zblOOm as opposed to ±lm 
when the full system is available. Developments are under way in the United States to 
improve the accuracy available to civil users. These are: 

• The wide-area augmentation system (WAAS) to improve accuracy en route; 

• The local-area augmentation system (LAAS) to improve terminal guidance. 

8.10 Global Air Transport Management (GATM) 

The rapidly increasing commercial air traffic density is leading to a pressing need to improve 
the air transport management (ATM) system by all available means and move on from the 
techniques and technologies that have served the industry for the last 40 years. This 
evolution will embrace the use of new technologies mixed with existing capabilities to 
offer improved air traffic management. The aims and objectives of ATM and a full 


description of the future air navigation system (FANS) may be found in Chapter 12 of the 
sister publication 'Civil Avionics Systems' (Moir and Seabridge, 2003). GATM is the 
military version of FANS and has to be compatible in all respects to enable the interoper- 
ability of civil and military aircraft within controlled airspace. This section provides a brief 
overview of some of the key features. 

To this end, the air traffic control authorities, airline industry, regulatory authorities and 
airframe and equipment manufacturers are working to create the future air navigation system 
(FANS) to develop the necessary equipment and procedures. In order to be able to use 
controlled airspace on equal terms with commercial users, military platforms will need to 
embody GATM objectives. 

The areas where improvements may be made relate to communications, navigation and 
surveillance, commonly referred to as CNS. The key attributes of these improvements may 
be briefly summarised as follows: 

• Communication. The use of data links to increase data flow and permit the delivery of 
complex air traffic control clearances. 

• Navigation. The use of GPS in conjunction with other navigational means to improve 
accuracy and allow closer spacing of aircraft. 

• Surveillance. The use of data links to signal aircraft position and intent to the ground and 
other users. 

These headings form a useful framework to examine the GATM improvements already made 
and those planned for the future. 

8. 10. 1 Communications 

The main elements of improvement in communications are: 

• Air-to-ground VHF data link for domestic communications; 

• Air-to-ground SATCOM communications for oceanic communications; 

• High-Frequency data link (HFDL); 

• 8.33 kHz VHF voice communications. 

8. 10. 1. 1 Air-to-Ground VHF Data Link 

The emergence of data links as means of communications versus conventional voice 
communications has developed recently in the commercial community; they have long 
been used for military purposes. 

Voice links have been used in the past for communications between the air traffic control 
system or ATM and the airline operational centre (AOC). The use of data links, controlled 
and monitored by the FMS or other suitable method on-board the aircraft, facilitates 
improved communication with the AOC and ATM systems. These data links may be 
implemented using one or more of the following: 

• VHF communications; 

• Mode S transponder; 

• Satellite links. 


Data link communications are being designed to provide more efficient communications for 
ATC and flight information services (FIS). Although these systems essentially replace voice 
communication, there will be a provision for voice back-up in the medium term. 

Flight plan data, including aircraft position and intent in the form of future waypoints, 
arrival times, selected procedures, aircraft trajectory, destination airport and alternatives, will 
all be transferred to the ground systems for air traffic management. The data sent to the 
ground ATM system will aid the process of predicting a positional vector for each aircraft at 
a specific time. This information will aid the task of the ground controllers for validation or 
reclearance of the flight plan of an aircraft. Furthermore, use of the required time of arrival 
(RTA) feature will enable the air traffic controllers to reschedule aircraft profiles in order that 
conflicts do not arise. 

For ATC flight service and surveillance, VHF data link (VDL) communications will be 
increasingly used for domestic communications. VHF communications are line-of- sight 
limited, as has already been explained. A number of options exist: 

1. VDL mode 1. Compatible with existing ACARS transmitting at 2.4 kbps. This mode 
suffers from the disadvantage that it is character oriented. 

2. VDL mode 2. Data only transmitted at 31.5 kpbs. As well as having a higher bandwidth, 
this protocol is bit rather than character oriented, making it 50-70% more efficient than 
the ACARS protocol. VDL mode 2 is able to support controller to pilot data link 
communications (CPDLC). 

3. VDL mode 3. Simultaneous data and analogue voice communications using time division 
multiple-access (TDMA) techniques. 

4. VDL mode 4. Used with the 1090 MHz signal of ATC mode S. 

It is expected that the introduction of data link technology will benefit all users owing to a 
more efficient and less ambiguous nature of the messages passed. Significant improvements 
in dispatch delays and fuel savings are expected as these technologies reach maturity. 

8. 10. 1.2 Air-to-Ground SATCOM Communications 

SATCOM is a well proven data link that, as has already explained, is limited at very high 
latitudes in excess of about 82°. The SATCOM system is supported by the INMARSAT 
constellation already described earlier in the chapter. HF Data Link 

Modern technology enables HF data link transmissions to be more robust than HF voice and 
therefore less susceptible to the effects of the sunspot cycle. HF data link provides primary 
coverage out to 2700 nm and secondary coverage beyond that should propagation conditions 
be favourable. There is extensive cover by ground stations located in the northern hemi- 
sphere such that HF data link is a viable alternative to SATCOM for north polar transitions. 
Refer to the communications and Navaids description earlier. 

8. 10. 1.4 8.33 kHz VHF Voice Communications 

Conventional VHF voice channels are spaced at intervals of 25 kHz throughout the spectrum. 
A denser communications environment has resulted in the introduction of digital radios that 


permit spacing at 8.33 kHz, allowing three channels to be fitted in the spectrum where only 
one could be used previously. With effect from 7 October 1999, these radios have already 
been mandated in Europe for operation above 20 000 ft and will follow in the United States 
within a number of years; one of the difficulties in predicting the timescale is the vast number 
of radios that have to be replaced/retrofitted. Protected I LS 

Within Europe some ILS installations suffer interference from high-power FM local radio 
stations. Modifications have been mandated that introduce receiver changes to protect the 
ILS systems from this interference. 

8.10.2 Navigation 

A number of navigational improvements are envisaged: 

1. Introduction of required navigation performance (RNP) and actual navigation perfor- 
mance (ANP) criteria. This defines absolute navigational performance requirements for 
various flight conditions and compares this with the actual performance the aircraft 
system is capable of providing. 

2. Reduced vertical separation minima (RVSM). 

3. Differential GPS (DGPS) enhancements: 

• WAAS - described earlier; 

• LAAS - described earlier. 

4. Protected ILS. 

5. Introduction of the microwave landing system (MLS) in Europe. 

8. 10.2. 1 Area Navigation (RNAV) 

Area navigation (RNAV) systems allow the aircraft to operate within any desired course 
within the coverage of station-referenced signals (VOR, DME) or within the limits of a self- 
contained system capability (IRS, GPS) or a combination of these. RNAV systems have a 
horizontal two-dimensional capability using one or more of the on-board navigational 
sensors to determine a flight path determined by navigation aids or waypoints referenced to 
latitude and longitude. In addition, the RNAV system provides guidance cues or tracking of 
the flight path. Many modern RNAV systems include a three-dimensional capability to define 
a vertical flight path based upon altimetry, and some include a full aircraft and engine 
performance model. 

The performance of pre-RNAV systems has historically been defined according to the 
following criteria: 

• Along-track error; 

• Across-track error; 

• Flight technical error (FTE). 

The total navigation error is the root sum square (RSS) of these elements for a given 
navigation means or phase of flight. 


The availability of the navigation capability is defined at 99.999%, and the integrity 
requirement for misleading navigation information is set at 99.9999%. 

8. 10.2.2 RNP RNA V and Actual Navigation Performance 

The actual navigation performance (ANP) of the aircraft navigation system is represented by 
a circle that defines the accuracy of the aircraft navigation system for 95% of the time. The 
value of the ANP is derived by taking the value of all the navigation sensors and statistically 
weighing them against the other sensors. After a period of time a degree of confidence is 
established in which are the most accurate sensors and therefore the ANP value is 
established. The 95% probability circle is that which is compared with RNP to decide 
whether the navigation system performance is good enough for the route segment being 
flown. The ANP and RNP values are displayed on the FMS CDU such that the flight crew 
can readily check on the navigation system status. Should the ANP exceed the RNP value 
for a given route sector for any reason - for example owing to a critical navigation sensor 
failing - the crew are alerted to the fact that the system is not maintaining the accuracy 
necessary. This will result in the aircraft reverting to some lower-capability navigational 
means. In an approach guidance mode it may necessitate the crew executing a go-around and 
reinitiating the approach using a less accurate guidance means. 

8. 10.2.3 Required Navigation Performance (RNP) 

The RNP defines the lateral track limits within which the ANP circle should be constrained 
for various phases of flight. The general requirements are as follows: 

1. For oceanic crossings the RNP is ±12 nm, also referred to as RNP- 12. 

2. For en-route navigation the RNP is ±2nm (RNP-2). 

3. For terminal operations the RNP is ±1 nm (RNP-1). 

4. For approach operations the RNP is ±0.3 (RNP-0.3). 

Other specific RNP requirements may apply in certain geographical areas, e.g. RNP-4 and 
RNP-10 (Figure 8.35). 

It is clear that this represents a more definitive way of specifying aircraft navigational 
performance, versus the type of leg being flown, than has previously been the case. Other 
more specific criteria exist: RNP-5 (also known as BRNAV or area navigation) has already 
been introduced in parts of the European airspace with the prospect that RNP-1 (also known 
as PRNAV or precision navigation) will be introduced in a few years. There are precision 
approaches in being - notably those in Juneau, Alaska - where RNP-0.15 is required for new 
precision approaches developed for mountainous terrain. 

8. 10.2.4 RNA V Standards within Europe 

Two RNAV standards are being developed in Europe: 

1. Basic RNAV (BRNAV). BRNAV was introduced in 1988 and is equivalent to RNP-5 for 
RNAV operations. Navigation may be accomplished by using the following means: 









Within 12 nm 


Within 2 nm 

Within 1 nm 

Within 0.3 nm 


Actual Navigation 

Performance (ANP) 

95% of time 

Figure 8.35 ANP versus RNP requirements. 


• VOR/DME with a 62 nm VOR range limit; 

• INS with radio updating or limited to 2h since the last on-ground position update; 

• LORAN-C with limitation. 

• GPS with limitation. 

Until 2005, primary sources of navigation will be DME/DME, VOR/DME and GPS. 
Advisory circular AC 90-96 on the approval of US operators and aircraft to operate under 
instrument flight rules (IFR) in European airspace designated for basic area navigation 
(BRNAV), 20 March 1998, approves the operation of US aircraft in European airspace 
under the application of existing advisory circulars. 

2. Precision RNAV (PRNAV). PRNAV is intended to be introduced at some time in the 
future but not before 2005. PRNAV will invoke the use of navigation under RNP-1 
accuracy requirements or better. RVSM 

One of the other ways of increasing traffic density is the introduction of the reduced 
separation vertical minima (RVSM) criteria. For many years aircraft have operated with a 
2000 ft vertical separation at flight levels between FL290 and FL410. As traffic density has 
increased, this has proved to be a disadvantage for the busiest sections of airspace. 
Examination of the basic accuracy of altimetry indicated that there were no inherent 
technical reasons why this separation should not be reduced. Accordingly, RVSM was 
introduced to increase the available number of flight levels in this band and effectively 
permit greater traffic density. The principle is to introduce additional usable flight levels such 
that the flight level separation is 1000 ft throughout the band, as shown in Figure 8.36. 

Originally a trial was mounted in 1997 to test the viability of the concept on specific flight 
levels - FL 340 and FL360 as shown in the figure. RVSM is now implemented throughout 
most of Europe from FL290 to FL410, introducing six new flight levels compared with 


Seperation Levels Trial Phase Full 

wef1960 commencing 27/3/97 Implementation 


FL390 ■ 





FL330 ■ 





Figure 8.36 RVSM - insertion of new flight levels. 

before. All the specified flight levels on the North Atlantic were implemented in 2001. Other 
regions in the globe will have RVSM selectively implemented to increase air traffic density 
according to Figure 8.36 and Table 8.4. 

8. 10.2.6 RVSM Implementation 

At the time of writing, the plans for the worldwide implementation of RVSM are as shown in 
Table 8.4, and many have been implemented to plan. The Federal Aviation Authority (FAA) 
RVSM website lists the most recent schedule and level of implementation. 

RVSM operation requires the aircraft to possess two independent means of measuring 
altitude and an autopilot with an accurate height hold capability. The operators of RVSM- 
equipped aircraft are not taken on trust: independent height monitoring stations survey 
aircraft passing overhead, measuring actual height compared with flight plan details, thereby 
assuring the performance of each aircraft and operator. RVSM implementation therefore 
embraces a watchdog function that ensures that all users are conforming to the RVSM 
accuracy and performance provisions. 

8. 10.2.7 Differential GPS Enhancements 

DGPS enhancements are being developed for en-route and precision landings in the United 
States. The GPS enhancements - WAAS and LAAS implementations - in the United States 
have already been described earlier in the chapter and their introduction should lead to the 
following accuracies being achieved as a matter of course: 

1. WAAS is anticipated to yield an accuracy of ~7m which will be sufficient for Cat I 

2. LAAS is expected to provide enhanced accuracies of ~1 m which will be sufficient for 
precision approaches Cat II and Cat III. 



Table 8.4 RVSM implementation schedule - worldwide 

RVSM status - Americas and Europe 

North Atlantic 

West Atlantic route system (WATRS) 

Europe tactical (UK, Ireland, Germany, Austria) 


South Atlantic 

Canada North domestic 

Canada South domestic 

Domestic US - phase l a 

Domestic US - phase 2 

Caribbean/South America 

March 1997 
October 1998 
24 January 2002 
1 November 2001 
24 January 2002 
April 2001 
24 January 2002 
24 January 2002 
April 2002 

FL3 10-390 
FL3 10-390 

Coordinate with US domestic 
1 December 2004 FL350-390 

Late 2005-2006 FL 290-390 


RVSM group established 

RVSM status - Asia/Pacific 



Western Pacific/South China Sea 

Middle East 
Asia-Europe/South of Himalayas 

February 2000 


Tactical use 


November 2001 


21 February 2002 



November 2003 


November 2003 


a DRVSM plan to be finalised not later than January 2002 on the basis of ATC simulation results and user inputs. 

The introduction of DGPS technology is also envisaged for Europe and the Far East. In 
Europe there are two programmes in the planning stage that will enhance satellite navigation. 
The European Space Agency (ESA), the European Commission (EC) and the European 
organisation for the Safety of Air Navigation (Eurocontrol) are working together on the 
development of a global positioning and navigation satellite system (GNSS) plan. The GNSS 
programme is being carried out in two phases: 

1. GNSS-1. This involves the development of the European geostationary navigation overlay 
system (EGNOS) which will augment the US GPS and Russian GLONASS systems. 

2. GNSS -2. This involves the development of a second-generation satellite navigation 
system including the deployment of Europe's own satellite system - Galileo. At the 
time of writing the EU nations were wrangling about budget increases needed to fund the 
programme, so delay appears to be inevitable. 

8.10.3 Surveillance 

Surveillance enhancements include the following: 


• ATC mode S; 


• Automatic dependent surveillance A (ADS- A); 

• Automatic dependent surveillance B (ADS-B). 

The operation of TCAS and ATC mode S has already been described in Chapter 7, but their 
use in a FANS/GATM context will be briefly examined here. 

When operating together with a mode S transponder and a stand-alone display or EFIS 
presentation, TCAS is able to monitor other aircraft in the vicinity by means of airborne 
interrogation and assessment of collision risk. TCAS II provides vertical avoidance 
manoeuvre advice by the use of RAs. TCAS II will soon be made mandatory for civil 
airliners - aircraft with a weight exceeding 15 000 kg or 30 or more seats - operating in 
Europe. This will be extended to aircraft exceeding 5700 kg or more than 10 seats, probably 
by 2005. 


Moir, I. and Seabridge, A. (2003) Civil Avionics Systems. Professional Engineering Publicating/ 

American Institute of Aeronautics and Astronautics. 
Advisory circular AC 121-13, Self-contained navigation systems (long range), 14 October 1969. 
Advisory circular AC 120-33, Operational approval for airborne long-range navigation systems for flight 

within the North Atlantic minimum navigation performance specifications airspace, 24 June 1977. 
Advisory circular AC 25-4, Inertial navigation systems (INS), 18 February 1966. 
Advisory circular AC 90-94, Guidelines for using global positioning system equipment for IFR en-route 

and terminal operations and for non-precision approaches in the US national airspace system, 

14 December 1994. 
Technical standing order (TSO) C-129a, Airborne supplementary navigation equipment using global 

positioning system (GPS), 20 February 1996. 
Advisory circular AC 90-45A, Approval of area navigation systems for use in the US national airspace 

system, 21 February 1975. 
Advisory circular AC 20-130A, Airworthiness approval of navigation or flight management systems 

integrating multiple sensors, 14 June 1995. 
Advisory circular AC 90-97, Use of barometric vertical navigation (VNAV) for instrument approach 

operations using decision altitude, 19 October 2000. 
Advisory circular AC 25-23, Airworthiness criteria for the installation approval of a terrain awareness 

and warning system (TAWS) for Part 25 airplanes, 22 May 2000. 
Advisory circular AC 00-31 A, National aviation standard for the very high frequency omnidirec- 
tional radio range (VOR)/distance measuring euipment (DME)/tactical air navigation (TACAN) 

systems, 20 September 1982. 
Advisory circular AC 20- 121 A, Airworthiness approval of LORAN-C navigation systems for use in the 

US national airspace systems (NAS) and Alaska, 24 August 1988. 
Advisory circular AC 00-31 A, National aviation standard for the very high frequency omnidirec- 
tional radio range(VOR)/distance measuring equipment (DME)/tactical air navigation systems, 

20 September 1982. 
Advisory circular AC 90-96, Approval of US operators and aircraft to operate under instrument flight 

rules (IFR) in European airspace, March 1998. 
Federal Aviation Authority (FA A) RVSM website: 1 .htm 

9 Weapons Carriage 
and Guidance 

9.1 Introduction 

Thus far the technologies associated with the overall integration of the avionics system and 
all the associated sensor technology have been described. In this chapter, exemplar platforms 
will be described along with the weapons they carry. By this means the concept of the total 
integration of avionics system, sensors and weapons will be outlined so that the reader can 
gain and understanding of the entire weapons system. To set a historical context and to 
illustrate the breadth of typical weapons systems, the following systems are described: 

• F-16 mid-life Update and subsequent developments; 

• AH-64 C/D Longbow Apache; 

• Eurofighter Typhoon; 

• F-22 Raptor; 

• Nimrod MR4; 

• F-35, formerly the joint strike fighter (JSF). 

Figure 9.1 shows a comparative timescale for these developments. There are some 
interesting observations that may be made: 

1 . The relative ease of modifying an existing platform, especially where there are a large 
number of aircraft manufactured - the F-16 (>4000) and AH-64 (>1000) are good 
examples. The huge investment in developing and purchasing aircraft in these numbers 
means that there is an inducement to modifying existing platforms instead of buying new 
airframes. There is a limit to this, as eventually the aircraft will run out of fatigue life and 
will have to be replaced. 

Military Avionics Systems Ian Moir and Allan G. Seabridge 
© 2006 John Wiley & Sons, Ltd. ISBN: 0-470-01632-9 





, 1 . 

i j F-16A/B&C/D 




■ + \ F-16MLU~[ l 


-► AH-64C/D 

► F-16E/F 



F/A-22 nnnnnnnnnnnno 

Nimrod MR4 nnnnnnnnncz^) 


Figure 9.1 Comparative development programmes. 

2. The Typhoon and Raptor have both suffered from political factors. Both platforms are 
very capable aircraft that were conceived as air superiority fighters during the latter stages 
of the Cold War and have lacked total political commitment since. The result has been in 
both cases a very long development phase of ~15 years leading to entry into service 
(EIS), and both aircraft are still evolving to adapt to deployment in air-to-ground roles. 

As will be seen, all of these aircraft are able to carry combinations of weapons options best 
suited for air-to-air and air-to-ground roles in which they are deployed. In many cases, 
common weapons are capable of being carried and released from more than one platform 
type, and this offers economies of scale in each weapon manufacture, and interoperability 
between different nations and air forces - a point recognised in NATO some time ago. The 
key has lain with the standardisation of the weapons/aircraft mechanical and electrical 
interfaces. The latter has taken on increasing significance as 'smart weapons' have evolved 
that require information exchange with the platform and sensors. 

The standard stores interface - MIL-STD-1760 - will also be described, as will the 
capabilities of common air-to-air and air-to-ground missiles that are used on the platforms 
listed above. 

9.2 F-16 Fighting Falcon 

The F-16 is a multirole fighter of which more than 4000 have been produced. The F-16 was 
selected by the US Air Force as a result of the lightweight fighter (LWF) contest in which the 
General Dynamics (now Lockheed Martin) YF-16 underwent a fly-off competition against 
the Northrop YF-17 and was successfully declared the winner in 1976. The F-16 was smaller 
than its contemporary, having a single engine versus two engines for the YF-17. The YF-16 
was a more innovative design than the YF-17, which in many respects was derived from the 
Northrop F-5. One particularly innovative feature was that the aircraft was designed to be 


negatively aerodynamically stable, that is, with the aircraft centre of gravity (CG) behind or 
close to the aerodynamic centre of pressure (CP). This unstable design necessitated the 
provision of a quadruplex fly-by-wire (FBW) computer system which carried out the 
necessary flight control calculations and signalled the demands to the flight control actuators 
by electrical means rather than using the conventional lever and push rod configuration. This 
resulted in a safe, highly efficient design that provided a fast control response while also 
being programmed to provide the pilot with pleasant handling characteristics. 

The F-16 is a classic example of a multirole aircraft serving in a variety of roles with a 
large number of air forces. The airframe has been grown to provide increased all-up weight, 
fuel and weapons loads; engine thrust has increased to accommodate the additional airframe 
mass without any loss in performance. Also interesting within the context of this book is the 
evolution of the avionics system to provide ever greater weapons systems capability within 
the aircraft. The F-16 MLU avionics systems architecture is depicted in a simplified form in 
Figure 9.2. 

9.2.1 F-16 Evolution 

The F-16 evolution over more than 20 years can be summarised as follows. The original F-16 
was designed as a lightweight air-to-air day fighter. Air-to-ground responsibilities trans- 
formed the first production F-16s into multirole fighters. The empty weight of the block 10 
F-16A is 15 600 lb. The empty weight of the block 50 is 19 200 lb. The A in F-16A refers to a 
block 10 to 20 single-seat aircraft. The B in F-16B refers to the two-seat version, hence the 
abbreviation A/B, C/D, etc. The block number is an important term in tracing the evolution 
of the F-16 and is particularly useful in tracking the evolving avionics systems configurations 
as the aircraft has developed. 

The F-16A, a single-seat model, first flew in December 1976. The first operational F-16A 
was delivered in January 1979 to the 388th Tactical Fighter Wing at Hill Air Force Base, 
Utah. The F-16B, a two-seat model, has tandem cockpits that are about the same size as the 
single canopy in the A model. Its bubble canopy extends to cover the second cockpit. The 
various F-16 models may be summarised as follows: 

1. Block 1 and block 5 F-16s were manufactured through 1981 for the US Air Force and for 
four European air forces. Most block 1 and block 5 aircraft were upgraded to a block 10 
standard in a programme called Pacer Loft in 1982. 

2. Block 10 aircraft (312 total) were built through 1980. The differences between these early 
F-16 versions are relatively minor. Block 15 aircraft represent the most numerous version 
of the more than 4000 F-16s manufactured. The transition from block 10 to block 15 
resulted in two hard points added to the chin of the inlet. The larger horizontal tails, which 
grew in area by about 30% are the most noticeable difference between block 15 and 
previous F-16 versions. 

3. F-16C/D aircraft, which are the single- and two-seat counterparts to the F-16A/B, 
incorporate the latest cockpit control and display technology. All F-16s delivered since 
November 1981 have built-in structural and wiring provisions and systems architecture 
that permits expansion of the multirole flexibility to perform precision strike, night attack 
and beyond- visual-range (BVR) interception missions. All active units and many Air 
National Guard and Air Force Reserve units have converted to the F-16C/D, which is 
deployed in a number of block variants. 


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4. Block 25 added the ability to carry AMRAAM to the F-16 as well as night/precision 
ground-attack capabilities and an improved radar, the Westinghouse (now Northrop- 
Grumman) AN/APG-68, with increased range, better resolution and more operating 

5. Block 30/32. This configuration adopted two new engines - block 30 designates a General 
Electric F110-GE-100 engine, and block 32 designates a Pratt & Whitney F100-PW-220 
engine. Block 30/32 can carry the AGM-45 Shrike and the AGM-88A HARM, and, like 
the block 25, it can carry the AGM-65 Maverick. 

6. Block 40/42. F-16CG/DG gained capabilities for navigation and precision attack in all 
weather conditions and at night with the LANTIRN pods and more extensive air-to- 
ground loads, including the GBU-10, GBU-12 and GBU-24 Paveway laser-guided bombs 
and the GBU-15. Block 40/42 production began in 1988 and ran through to 1995. 
Currently, the block 40s are being upgraded with several block 50 systems: the ALR-56M 
threat warning system, the ALE-47 advanced chaff/flare dispenser, an improved perfor- 
mance battery and Falcon UP structural upgrade. 

7. Block 50/52. Equipped with a Northrop Grumman APG-68(V)7 radar and a General 
Electric F 110-GE-129 increased performance engine, the aircraft are also capable of 
using the Lockheed Martin low-altitude navigation and targeting infrared for night 
(LANTIRN) system (see Chapter 5). Technology enhancements include colour multi- 
functional displays and a programmable display generator, a new modular mission 
computer (MMC), a digital terrain system, a new colour video camera and colour 
triple-deck video recorder to record the pilot's head-up display view and an upgraded 
data transfer unit. In May 2000, the US Air Force certificated block 50/52 variant F-16s to 
carry the following: 

CBU- 103/104/105 wind-corrected munitions dispenser; 
AGM-154 joint stand-off weapon (JSOW); 
GBU-31/32 joint direct attack munition (JDAM); 

• Theatre airborne reconnaissance system. 

Beginning in mid-2000, Lockheed-Martin began to deliver block 50/52 variants equipped 
with an on-board oxygen generation system (OBOGS) designed to replace the original 
liquid oxygen (LOX) system. 

8. Block 50D/52D. Wild Weasel F-16CJ (CJ means block 50) comes in C-model (one-seat) 
and D-model (two- seat) versions. It is best recognized for its ability to carry the AGM-88 
HARM and the AN/ASQ-213 HARM targeting system (HTS) in the suppression of 
enemy air defences (SEAD) mission relinquished by the F-4. This system allows HARM 
to be employed in the 'range-known' mode, providing longer-range engagements against 
specific targets. This specialised version of the F-16, which can also carry the ALQ-119 
electronic jamming pod for self-protection, became the sole provider for Air Force SEAD 
missions when the F-4G Wild Weasel was retired from the Air Force inventory. Although 
F-18s and EA-6Bs are HARM capable, the F-16 provides the ability to use the HARM in 
its most effective mode. 

9. Block 60 or F-16E/F. In May 1998 the UAE announced selection of the block 60 F-16 to 
be delivered between 2002 and 2004. The upgrade package consists of a range of modern 
systems including conformal fuel tanks for greater range, new cockpit displays, an 
internal infrared (IR) sensor suite, a new mission computer and other advanced features 



including a new active electronically scanned array radar (APG-80) as described in 
Chapter 4 (see below under F-16 E/F). 

9.2.2 F-16 Mid-life Update 

Towards the end of the 1980s the US Air Force began to study an avionics upgrade for the F- 
16A/B and also attracted interest from four European NATO air forces that also operated the 
F-16. These air forces were those of the Netherlands, Belgium, Norway and Denmark. In the 
event, the end of the Cold War led the US Air Force to abandon their plans to adopt the mid- 
life update (MLU) as it was called, and their F-16A/Bs were retired instead. In all, a total of 
343 NATO aircraft were modified as follows: 

Air force 

Number modified 




The MLU programme was undertaken by an industrial partnership formed from partners 
from the participating nations. The key technical improvements included: 

1. Introduction of a new modular mission computer (MMC) to provide processing for 
weapons control, stores management and head-up display (HUD). Helmet-mounted 
display (HMD) features were also added. The MMC is a derivative of the common 
integrated processor (CIP) developed for the F-22 (see F-22 in this chapter) (Figure 9.3). 
This is a modular avionics rack populated by line replaceable modules (LRMs) as 
described in Chapter 2 under the JIAWG architecture. There are a total of 30 LRM slots 
provided in the MMC rack, of which 21 are populated; the remaining nine are available 
for growth. The use of LRMs allows the adoption of two-level maintenance, thereby 
eliminating intermediate-level (I-level) test equipment. The MMC replaces three LRUs 

Figure 9.3 MLU modular mission computer. 


from the previous architecture, reducing volume, weight and power consumption. The 
MMC has four functional areas: 

• Data processing set (DPS) associated with weapons MUX bus control; 

• Avionics display set (ADS) interfacing with the HUD; 

• Avionics I/O (AIOS) interfacing with the avionics units; 

• Power set (PS) controlling the power supplies and conditioning. 

2. Improvements to the AN/APG-66 radar bring it to the AN/APG-66(V)2A standard. 
Improved operational modes include ten-target track- while-scan (TWS) capability, 64 : 1 
Doppler beam sharpening (DBS) and enhanced air-to-air and ground-mapping modes. 
Target detection and tracking ranges have been improved by ~25%. Rationalisation of 
the digital signal processor (DSP) and radar processor (RP); these have been replaced by 
a single unit with accompanying improvements in weight, volume and power dissipa- 
tion. Introduction of a radar MUX bus to integrate the radar units has provided further 

3. Advanced identification friend or foe (AIFF) AN/APX-113(V) incorporating both IFF 
interrogator and transponder functions. 

4. Wide-angle HUD. 

5. Multifunction displays (MFDs). 

6. Data entry/cockpit interface set integrating communication and navigation functions. 

7. Side stick controller and throttle grip to give hands-on throttle and stick (HOTAS) 

8. Improved data modem (IDM) to provide target data sharing capability using V/UHF and 
secure communications. 

9. Digital terrain system using the terrain profile matching (TERPROM) principle. 

10. Global positioning system (GPS). 

1 1 . Navigation/targeting pod provision - LANTIRN. 

Many of these items were subsequently added to the US Air Force block 50 F-16s, but under 
a series of budget allocations and modification programmes. 

The top-level architecture for the F-16 MLU is shown in Figure 9.2. The MMC performs 
the primary bus controller function for a total of five MIL-STD-1553B (MUX) buses: MUX 
A, MUX B, MUX C, MUX D and MUX W. The figure shows three of these MUX buses: 

1. MUX A is the avionics bus that interconnects all the main avionics functions: air data, 
IFF, radar altimeter, inertial navigation unit (INU), up-front controls, data entry, flight 
control, IDM and TERPROM. 

2. MUX D is the displays bus interfacing with the two MFDs, wide-angle HUD and HMD 
among others. The main avionics units - AN/APG-66 (V2) radar, MMC and stores 
management system (SMS) - straddle both MUX A and MUX D buses. 

3. MUX W is the weapons bus controlled by the stores management system (SMS) and 
interfaces to the weapons stations by means of standard MIL-STD-1 760/1 553B inter- 

There are a total of nine stores stations (STA) provided, as shown in Table 9.1. In addition, 
the two new chin-mounted stations - 5A and 5B - are also shown. As well as carrying 
























































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weapons and fuel tanks on designated stations, the system is also capable of carrying the 
following pods for specific missions: 

• LANTIRN pods - navigation and targeting; 

• Reconnaissance pod; 

• AN/ALQ-131 ECM pod. 

The MLU/block 50/52 modifications have given an enormous mission capability far 
exceeding that of the original F-16 A/B configuration. All new F-16s being delivered are 
delivered to block 50/52 standard or block 60 - see below. 

9.2.3 F-16 E/F (F-16 block 60) 

The latest avionics upgrades to the F-16 were originally planned as block 60 as a 
continuation of the block numbering system. Block 60 - now officially termed the 
F-16E/F - has been ordered by the United Arab Emirates (UAE) and represents a totally 
new avionics system. Compared with the original F-16A/B, the E/F version has increased the 
maximum take-off weight by 50% (to 22 700 kg/50 000 lb) and increased engine thrust by 
35% (to 32 500 lb/145 kN). In the F-16E/F, approximately 70% of the aircraft structure is 
new; additional conformal fuel tanks may be fitted to improve range and a new dorsal 
avionics bay has been introduced down the spine of the aircraft (Figure 9.4). 

The avionics system for the F-16E/F offers a huge advance in capability over previous 
systems. The key attributes are: 

1. Introduction of an active electronically scanned array (AESA) radar - AN/APG-80 - with 
a marked increase in radar performance and capabilities as described in Chapter 4. In 
certain circumstances, the increase in radar range offered by the AESA radar represents a 
55% improvement over the block 50/52 radar (Chapter 4, Figure 4.29). The radar is 
cooled by a dedicated liquid cooling system with the radar ECS located in the dorsal spine 
equipment bay and the coolant lines run forward to the radar. The radar also uses COTS 
processor and fibre-channel (FC) technology to handle the high-bandwidth processing and 
data handling tasks. 

2. The avionics systems integration is performed by a dual-redundant FC that is capable of 
transferring data at 1 Gbit/s. This provides much greater bandwidth than the 1553B MUX 
buses, allowing more data to be passed between aircraft systems. Avionics computing 
resources are based upon Power PC COTS technology, and the 1.3 million lines of 
software code have been written in C++. Existing/legacy software has been rewritten 
from the original Jovial or Ada languages into C++. 

3. Introduction of an integrated FLIR and targeting system (IFTS) developed by Northrop 
Grumman and designated as the AN/AAQ-32. This system offers the navigation and 
targeting functions offered by the LANTIRN system but with separate navigation and 
targeting pods. The system comprises: 

• A passive navigation FLIR located on the upper fuselage just ahead of the cockpit; 

• A targeting FLIR minipod which looks similar to the XR Sniper pod being introduced 
on to Air National Guard (ANG) block 30s and US Air Force block 50s having a 
similar chisel nose; however, it is more likely to be based upon the Northrop Grumman 



Figure 9.4 F-16E/F (Lockheed Martin). 

and Rafael technology embodied in the Litening family of EO pods which is outlined 
in Chapter 5. (Table 5.3). 

4. Improved Falcon Edge electronic warfare (EW) suite. 

5. Improvements to the displays - introduction of three (5x7 in) colour MFDs and use of a 
25° x 25° HUD. 

6. Improvements to the flight control system to provide safety features such as automatic 
recovery from a deep stall and automatic ground collision avoidance. The flight control 
software has been rewritten in the commercial language C++. 

The F-16 can carry a wide variety of weapons to satisfy a number of different roles (see 
Table 9.1 which displays typical weapons loads for four different missions). Table 9.1 is a 
simple portrayal of the weapons carried by the many air forces who operate the F-16. 
Many operators will have their own configurations of indigenously produced missiles 
and ECM pods. The wide range of weapons carried is a testimony to the flexibility of the 


9.3 AH-64 C/D Longbow Apache 

The Army AH-64 C/D Longbow Apache is a formidable attack helicopter. The original 
Hughes AH-64 A/B Apache became a McDonnell Douglas Helicopter Company (MDHC) 
product before they in turn were taken over by Boeing. 

9.3. 1 Baseline System 

The original AH-64A/B Apache was fitted with a state-of-the-art electrooptics suite that 
comprised the following capabilities: 

• Day TV for use by day and low-level light conditions providing monochrome (black and 
white) imagery. 

• FLIR- viewed imagery, real world and magnified during day, night and adverse weather. 

• Direct vision optics (DVO) viewing real world in colour and with magnified images 
during day and low-light conditions. 

These features are integrated in the integrated AN/AAQ-1 1 target acquisition and designator 
sight (TADS) and AN/ASQ-170 pilot's night-vision sight (PNVS) into the combined TADS/ 
PNVS described in Chapter 5. 

The US Army procured a total of 824 of the A/B variants of Apache, and several overseas 
countries also purchased this baseline variant of the aircraft. By the late 1980s, studies were 
under way to upgrade the aircraft, and this led to the AH-64C/D Longbow variant. This 
programme entails the remanufacture of the original 64A/B version and converts it to the 
later variant. The US Army intend to procure a total of 227 helicopters to the 64D variant 
fitted with the fire control radar (FCR) and 531 to the 64C variant which does not have the 
FCR but includes all the other avionics modifications. 

The United Kingdom has also procured 67 of the Longbow Apache as the WAH-64, and 
these have entered service with the British Army. Other foreign nations that have purchased 
or have shown an interest in the Longbow Apache are: the Netherlands - 30; Singapore - 8, 
with an option for 20 more; Japanese Defence Force - 12, with a requirement for up to 60; 
Israel - 12 converted AH-64As plus nine new-build; Egypt - 36 rebuilt AH-64 As; and 
Kuwait - 16. 

9.3.2 Longbow Apache 

The key elements of the AH-64C/D upgraded avionics systems are: 

1. Mast-mounted fire control radar (FCR) with the designation AN/APG-78, more com- 
monly known as the Longbow fire control radar. The FCR is a multimode millimetric 
wave (MMW) radar and provides the pilot with the capability of detecting, classifying 
and prioritising stationary and moving targets on the ground and in the air. The radar has 
four modes of operation: 

• Air-targeting mode (ATM): this detects, classifies and prioritises fixed and rotary wing 



Electrical Power 
Management System 


Aircraft Survivabilty: I 


IR Jammer 

RF Jammer 

Radar Warning 
Laser Warning 


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- Data Modem 

- UHF(AM) 

- VHF (FM & AM) 

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[MUX 4] 

MIL-STD-1553B MUX 3 

Figure 9.5 Longbour Apache avionics architecture. 

• Ground- targeting mode (GTM): this detects, classifies and prioritises ground and air 

• Terrain-profiling mode (TPM): this provides obstacle detection and adverse weather 
pilotage aids to the Longbow. 

• Built-in test (BIT) mode: this monitors radar performance in flight and isolates 
electronic failures before and during maintenance. 

2. Introduction of a new sensor, the radio-frequency interferometer (RFI) - AN/APR-48A. 
The RFI is capable of monitoring a broad spectrum and identifying threat radars and also 
when threat emitters are tracking the helicopter. The system has a stored library of threats. 

3. Improved navigation using a dual-channel enhanced global positioning system (GPS)/ 
inertial navigation system (INS) or EGI. 

4. Improved Doppler velocity rate sensor. 

5. The UK WAH-64 has a slightly different fit in terms of radios and IFF. There will be a 
new EW equipment fit called the helicopter integrated defensive aids system (HIDAS) 

• AN/AAR-57 common missile warning system (CWMS); 

• AN/APR-48 RFI (as for AH-64C/D); 

• Sky Guardian 2000 radar warning receiver (RWR); 



• Type 1223 laser warning system; 

• IR jamming system: either the AN/ALQ-144 as already fitted to the AH-64 or the AN/ 
AAQ-24 Nemesis directed IR countermeasures (DIRCM) system fitted to C-130 

The new avionics is hosted in extended forward avionic bays (EFABs) each side of the 
cockpit. The EFABs are conditioned by means of two closed-cycle environmental control 
systems. The Apache is flown by a two-man crew: a pilot and a copilot gunner (CPG). The 
new mission system introduced by Longbow allows the crew member to use the FCR to seek 
air targets while the CPG uses the TADS/PNVS system to identify ground targets. 

The Longbow Apache avionics system is shown in Figure 9.6. It can be seen that this 
federated architecture using 1553B MUX buses bears some similarity to the F-16 MLU 
architecture. There are four MUX buses: 

1. MUX bus 1 integrates the electrical system and the communications suite with two 
systems processors. The communications suite includes a communication interface unit, a 
data modem, UHF (AM), VHF (AM and FM), IFF and secure communications. 

2. MUX bus 2 is similar to an avionics bus in function, interfacing the following functions 
with the systems processors: 

• Navigation; 

• Flight control; 

• Display processors, which in turn drive the display suite; 

• Weapons processors; 

• Data management; 

• Aircraft survivability equipment (ASE) including: 

- radar warning: AN/APR-39A (V) 

- laser warning: AN/AVR-2A 

Mast Mounted Assembly 

Radio Frequency 

RF Missile 

RF Seeking 

Figure 9.6 Fire control radar, RFI and RF-guided Hellfire. (Lockhead Martin) 


-radar jammer: AN/ALQ-136(V) 
- IR jammer: AN/ALQ-144(V). 

3. MUX bus 3 is a weapons bus integrating the weapons processors, the pylon interface unit 
(PIU), Hellfire missiles and the AN/APG-78 fire control radar. 

4. MUX bus 4 is a bus dedicated to integrating the fire control radar units. 

See Figure 9.6 which depicts: 

• The AN/APG-78 radar mast assembly; 

• The AN/APR-48A; 

• RF-guided AGM-114 Hellfire missile. 

9.3.3 Modernisation of TADS/PNVS 

Towards the end of the 1990s a programme was initiated to modify and upgrade the TADS/ 
PNVS. The M-TADS and M-PNVS programmes (also known as Arrowhead) upgrade 
the system by replacing six of the eight existing system line replaceable units (LRUS), as 
well as introducing improved technology with an improved performance, Arrowhead draws 
upon in-service experience to make the system truly two-level maintenance capable. The 
LRUs are also being designed so that in many cases LRMs may be replaced instead of the 
entire LRU. 

The external appearance of the M-TADS/PNVS is similar to the original TADS/PNVS 
apart from the PNVS shroud which now houses two sensors: a new FLIR and an in- 
built image intensifier (effectively an integrated 18 mm NVG tube). The pilot will be able 
to use either FLIR or an image intensifier, depending upon which is best for the conditions 
that prevail. Ultimately there are plans completely to fuse the output of both sensors 
(Figure 9.7). 

The new FLIR will have an improved 4 x 480 detector, as opposed to the original 
1 x 180, and therefore image quality will improve. This performance improvement has a 
particular bearing in avoiding wires at low level and in adverse weather. Automatic 
boresighting (harmonisation) is incorporated, with the process taking about 2min after 
being initiated by the CPG. A number of other improvements including a new multitarget 

Pilotage System f— ^ Pilotage 

M fltt Sw Advanced 
Pilotage IS I I Technology 

Image- WW / l\ FLiR 


ffWk ♦ Laser 
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Pilotage j^^ \ | » Designator 

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Technology %&* ^ Tracker 

Targeting System ' Cay TV 

Figure 9.7 M-TADS/PNVS - Arrowhead. (Lockhead Martin) 


tracker system (capable of tracking up to six targets) and extended-range algorithms will 
enable target tracking at ranges around twice that of the existing system. 

The M-TADS/PNVS is delivered as a field retrofit kit that can be installed in about 8 h. 
Deliveries commenced in 2004. 

9.3.4 Weapons 

The Apache is capable of carrying the following weapons: 

1. Up to 16 AGM-114 Hellfire missiles. The latest AGM-114 K has a dual-seeker capability: 
pulsed radar or semi-active laser guidance and a range in excess of 8 km (5 miles). 

2. Hydra 70 rocket pod carrying up to 1970 mm folding-fin aerial rocket (FFAR) unguided 

3. M 320 30 mm 'chain gun' located under the chin of the aircraft and capable of carrying up 
to 1200 rounds of ammunition. The chain gun can be slaved to the pilot/CPG's helmet 

The Hellfire missiles and pylon interface units (PIUs) are interfaced with the weapons MUX 
bus as shown in Figure 9.5. Four missiles can be accommodated on each M229 missile 
launcher (Figure 9.8). 

Although there have been test firings of other missiles, including the AIM-9 Sidewinder, 
AIM-28 air-to-air stinger (both heat-seeking missiles) and the British laser-guided Star- 
streak, none of these has been deployed for operational service, (see Figure 9.9 for typical 
weapons loads of 16 x Hellfire missiles and a mixed Hellfire missiles (eight) plus rockets 

9.4 Eurofighter Typhoon 

The Eurofighter Typhoon had its origin in an operational need in the 1980s to provide an air 
superiority fighter for the European NATO air forces to counter the threat of the Soviet Union 
and Warsaw Pact countries. After the decision of France not to participate, the project became 

Figure 9.8 Apache M229 missile launcher. (Lockhead Martin) 



Figure 9.9 Typical Longbow Apache weapons. 

a four-nation joint project with the participating partners being Britain, Germany, Italy and 
Spain. The project had to accommodate national as well as common equipment fits and both 
single-seat and two-seat versions, the latter for training rather than operational purposes. 

The Typhoon configuration had previously been very successfully demonstrated using a 
single- aircraft flight demonstrator called the experimental aircraft programme (EAP) which 
first flew in 1986. This aircraft demonstrated cardinal-point technologies including colour 
multifunction displays, an integrated utilities management system (UMS), the first of its type 
incidentally to fly anywhere in the world, and a digital fly-by-wire system to control the 
highly unstable aircraft. At the time it was the first aircraft flying in Europe with extensive 
use of MIL-STD-1553 buses. This aircraft flight demonstrator was funded by the UK 
Ministry of Defence (MOD), together with UK Industry, and with some help from German 
and Italian Industry. It flew for around 2 years, gathering vital data about the aircraft 
dynamics and the interaction of the new digital systems, and proved to be a highly successful 
venture gaining valuable experience that would be used during the design of the Typhoon 
(Figure 9.10). The aircraft is now at Loughborough University in the United Kingdom. 

The aircraft is now in service with the air forces of the four participating nations. The 
avionics architecture is shown in Figure 9.11. The full capabilities of the system may be best 
explained by describing the following elements: 

Figure 9.10 Eurofighter Typhoon (Eurofighter GmbH). 

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• Sensors and navigation; 

• Displays and controls; 

• Flight control; 

• Utilities control; 

• Systems integration; 

• Survival/countermeasures; 

• Weapons. 

9.4.1 Sensors and Navigation 

The Typhoon sensors include the following: 

1. Captor radar. This is an X-band (8-12 GHz) radar multimode pulse Doppler radar. A 
track- while- scan (TWS) mode can track, identify and prioritise up to 20 targets 
simultaneously. Air-to-ground modes include a ground moving target indication 
(GMTI), spot mapping and surface ranging. A synthetic aperture radar (SAR) mode 
has the capability of high resolution for specific mapping purposes. Sophisticated 
frequency analysis techniques provide a non-cooperative target recognition capability 
where the signal returned from a target aircraft may be analysed and its signature 
recognised as being from a particular aircraft type. At some stage, Typhoon may be 
retrofitted with a European AESA radar with technology developed jointly from the 
United Kingdom, Germany and France AMSAR programme. A small demonstration 
array has been tested, and a full-scale array of 1000 or more elements is being flown on a 
test bed aircraft. 

2. Infrared search and track (IRST). This is a second-generation IRST system called 
PIRATE and was described in Chapter 5. It provides passive IR detection in the 
MWIR (3-5 urn) and LWIR (8-1 1 urn) bands. 

3. IFF interrogator and transponder. An IFF interrogator and mode S transponder compatible 
with the NATO IFF Mk 12 standard. 

4. FLIR targeting pod. The aircraft will have the ability to carry a contemporary FLIR 
targeting pod, as yet this capability is not operational. 

5. Dual INS/GPS. A laser-rate gyro-based INS together with GPS provides better naviga- 
tional accuracy within several metres. A terrain avoidance warning system (TAWS) based 
upon TERPROM working with the INS/GPS and covert radio altimeter allows passive 
low-level navigation and terrain avoidance. 

6. Air data. Triplex air data sources provide high integrity data to the FBW system. 

9.4.2 Displays and Controls 

The displays and controls include the following: 

1. HOTAS capability providing hands-on throttle and stick control of sensors, weapon 
control and communications and cursor control. A total of 24 selector buttons are 
provided (12 on each control). 

2. Direct voice input (DVI) with 200 commands and a response time of 200 ms. A 95% 
recognition capability is claimed. 


3. Wide-angle HUD with a 35° x 25° FOV. 

4. Three multifunction head-down displays (MHDDs) using colour AMLCD technology. 
Any of the displays - usually the centre display - can show a moving map using digital 
terrain data to portray the position of the aircraft. If necessary, the target and threat 
scenario may be overlaid, providing the pilot with complete tactical awareness. 

5. Helmet-mounted sighting system (HMSS) with an HMD providing a binocular system 
with up to 40° FOV. 

9.4.3 Flight Control 

The FBW is a full-authority active control technology (ACT) digital system to provide 
carefree handling of the aircraft using all-moving foreplanes mounted near the nose, wing 
trailing edge elevons, leading edge slats, rudder and airbrake. The system has quadruplex 
digital flight control computers, each containing eight Motorola 68020 processors and 
specially designed ASICS to achieve the necessary levels of safety. The flight control 
computers, sensors and flight control actuators are connected using a MIL-STD-1553B data 
bus and dedicated links where necessary. The flight control bus interfaces to the avionics bus 
via a dedicated interface. 

9.4.4 Utilities Control 

Control of the aircraft utilities systems such as fuel, environmental control, brakes and 
landing gear, secondary power systems, and OBOGS are by means of dedicated controllers 
connected to a utilities MIL-STD-1553B bus. Also connected to this bus are the full- 
authority digital engine controllers (FADECs) for the Eurojet 2000 engines and a main- 
tenance data panel. This philosophy in part follows the rationale of an integrated utilities 
management demonstrated on the EAP described above. 

9.4.5 Systems Integration 

The aircraft uses a combination of 20 Mbit/s fibre-optic STANAG 3010 and standard 1 Mbit/s 
MIL-STD-1553 buses to integrate the various avionics subsystems. The STANAG 3910 bus 
combines high data rate 20 Mbit/s fibre-optic transfers by using wire-based 1553 control 
protocol as described in Chapter 2. To see how these high-speed buses integrate the Typhoon 
avionics system, refer to Figure 9.11 which offers a very simplified portrayal; in fact there 
are a total of two STANAG 3910 and six MIL-STD-1553B in total to integrate all the aircraft 
avionics subsystem. The aircraft-level data buses may be simply described as follows: 

1. STANAG 3910 avionics buses. The avionics and attack buses interface with the sensors 
and displays. There are dedicated interfaces to the defensive aids subsystem (DASS) 
and flight control system. Two display processors are connected to both the avionics 
and utilities bus. The stores management system interfaces with the dedicated weapons 

2. MIL-STD-1553B flight control bus. The flight control system has a dedicated data bus 
interconnecting sensors, flight control computers and actuator assemblies. There is a 
dedicated interface connecting the flight control and utilities buses. 


3. MIL-STD-1553B utilities bus. A dedicated bus interconnects the utility control system 
(UCS) computers, FADECs and maintenance data panel which facilitates servicing the 

4. MIL-STD-1553B weapons bus. The dedicated 1553/MIL-STD-1760 weapons bus inter- 
faces with the 13 weapons stations as described below. 

9.4.6 Survival I Countermeasures 

Aircraft survival and countermeasures are provided by an integrated defensive aids suite 
(DASS) which integrates the following equipment: 

1. Wide-band receiver (100 MHz to 10 GHz) providing 360° radar warning receiver (RWR) 
coverage in azimuth and an active jammer using antennas located on the wing-tip pods 
and the fuselage. 

2. A pulse Doppler missile approach warning (MAW) system is fitted which uses antennas 
located at the wing roots and near the fin. This system warns of the approach of passive as 
well as actively guided missiles. Improvements are expected to enhance this system using 
either IR or UV detectors. 

3. Laser warning receiver (Royal Air Force only). 

4. Towed radar decoy (Royal Air Force only). This is a derivative of a system already 
deployed by the RAF on Tornado and other aircraft. 

5. Chaff and flare dispenser. 

9.4.7 Weapons 

The Typhoon is able to carry a wide range of weapons and stores to satisfy the operational 
needs of the four participating nations and export customers. The Typhoon has a total of 13 
weapons stations, four under each wing and five under the fuselage. The full complement of 
weapons that may be carried is shown in Tables 9.2 and 9.3. Figure 9.12 illustrates several of 
these weapon fit options. 

Figure 9.12 Typhoon weapon carriage options. (BAE SYSTEMS) 






































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F/A-22 RAPTOR 357 

9.5 F/A-22 Raptor 

9.5. 1 Introduction 

The F/A-22 was the outcome of a US Air Force DemVal fly-off between two competing 
designs for the advanced tactical fighter (ATF): the Lockheed YF-22A and the Northrop YF- 
23 A. The aircraft was selected as the winner in December 1990 and progressed to the 
engineering manufacturing and development (EMD) phase. After a protracted development 
phase the aircraft is now entering service with the US Air Force with an original planned 
purchase of 648 aircraft, subsequently progressively reduced to 232 aircraft owing to rising 
costs. The aircraft designator was recently changed to F/A-22 from F-22 to reflect an 
increased emphasis upon the attack role. 

The JIAWG architecture for the F/A-22 has already been described in Chapter 2, and 
many of the development and technology obsolescence issues have already been outlined. 
The F/A-22 is the only survivor of the triad of programmes for which the JIAWG 
architecture was intended: the US Navy A- 12 [formerly the advanced tactical aircraft 
(ATA)] was cancelled in 1991, and the US Army RAH-66 Comanche (formerly the LHX) 
followed in 2004. Chapter 2 also outlined the integrated RF aperture sharing architecture 
adopted in an initial form by the F/A-22 and to be carried forward to the Lockheed Martin F-35. 

The common integrated processor (CIP) integrated rack/LRM technology did find another 
application in the modular mission computer (MMC) on the F-16E/F, albeit with a 30- 
module rack rather than the 2 x 66 module racks of the CIPs. The CIP is reported to contain 
33 signal processors and 43 data processors of seven different types that are interconnected 
to provide a fault-tolerant network. The CIPs are liquid cooled using a poly-a-olefin (POA) 
coolant and a sophisticated environmental control system (ECS) comprising both forward 
and aft cooling loops. The forward loop comprises the POA liquid coolant loop while the aft 
loop is more conventional using an air cycle machine (ACM) and air/fuel heat exchangers to 
dump rejected heat into the fuel. This system enables the modules to be maintained in a 
friendly environment of ~68°F or 20°C, allowing LRM mean time between failure (MTBF) 
figures of ~25 000 h per module to be attained. 

Given that the F/A-22 architecture and the associated technology have already been 
thoroughly described in Chapter 2, this portrayal will be confined to the major subsystems 
that comprise the avionics system (Figure 9.13): 

• The APG-77 AESA radar; 

• Electronic warfare (EW) and electronic support measures (ESM); 

• Communications, navigation and identification (CNI); 

• Displays; 

• Vehicle management system (VMS); 

• Weapons. 

9.5.2 AESA Radar 

The APG-77 radar is arguably one of the most advanced in service. The concept and benefits 
of the AESA are described in Chapter 4. The APG-77 has a 1500-element array with a 
claimed detection range of 125 nm against aim 2 radar cross-section target (Figure 4.29). 
Perhaps more important is the flexibility that an active electronically scanned array confers 



















Common Integrated 
Processor #1 


Common Integrated 
Processor #2 

High Speed Data Bu 

Figure 9.13 F/A-22 - avionics system top-level architecture. 






where the beam or beams may be switched virtually instantaneously across the entire scan 
pattern. Added to this is the fact that the radar can effectively operate simultaneously in 
several modes, providing the utmost in operational flexibility. 

The use of frequency agility to switch the operating frequency of the radar makes it 
difficult for a potential foe to categorise and classify the radar emissions. The use of 
sophisticated pulse burst modulation techniques enables information to be extracted that 
would not be possible using conventional techniques. The availability of a complex array of 
signal processors enables all the necessary target characteristics to be analysed and the 
characteristics of the radar to be modified to confuse or nullify the opponent's radars 
(Fulghum, 2000). The F/A-22 and the APG-77 radar are depicted in Figure 9.14. 

9.5.3 Electronic Warfare and Electronic Support Measures 

The F/A-22 combines its high levels of stealth with a sophisticated array of EW and ESM 
sensors. The aircraft is able to jam enemy radars using on-board jammers operating in the 2- 
18 GHz frequency band, but this would not be the preferred mode of operation for a stealthy 
aircraft since those transmissions may be detected. The trump card is that the aircraft can 
operate undetected - or at least with a very low probability of detection - while gathering 
vital data about the enemy radars and the electronic order of battle (EOB). The aircraft 
carries a range of integral EW/ESM sensors and apertures conformally, that is, within the 
envelope of the aircraft and flush with the surface (Figures 2.32 and 2.34 and Table 2.1). The 
aircraft passive EW/ESM antennas include: 

• Six radar warning receiver (RWR) spiral antennas: four located port and starboard, 
forward and aft, one each top and bottom operating over the frequency range 2-18 GHz; 



Figure 9.14 F/A-22 as Air Force photo and APG-77 radar. (Northerp German) 

• Twelve situational awareness (SA) spiral antennas: six port and six starboard operating 
over the frequency range 2-18 GHz; 

• Six SA spiral antennas: three port and three starboard operating over the frequency range 
0.5-2 GHz. 

In certain modes of operation, RWR and SA inputs may be combined to provide azimuth 
(AZ) and elevation (El) direction finding on the emitting source. This facility is called 
precision direction finding and allows accurate angular measurements of emitters to be 
correlated with radar returns or stored in a threat library. 
The antennas associated with active ECM are: 

1 . six active ECM log periodic antennas transmitting in the 2-6 GHz frequency band. These 
antennas are located on the waterline and so are effective in countering threats both above 
and below the aircraft. 

2. six active ECM log periodic antennas transmitting in the 6-18 GHz frequency band. 
These are located on the waterline. 

3. two ECM spiral antennas located top and bottom and receiving signals over the frequency 
range 2-1 8 GHz. 


While precise details of the performance and capabilities of the EW/ESM system have not 
been revealed for obvious reasons, the sophistication of the antennas suite is clearly 
indicative of a highly significant capability. 

9.5.4 CNI 

With one or two exceptions, the CNI system is relatively conventional, comprising the 
following equipment and functions: 

• VHF communications; 

• Secure voice communications; 

• UHF communications; 

• Inter/intraflight data link (IFDL) - a cooperative data link that allows all the F/A-22s in a 
flight to share data automatically about the status of the aircraft and the targets being 
engaged; the antennas are narrow-beam steerable arrays; 

• Dual Litton LRG-100 inertial reference set (IRS)/GPS; 


• JTIDS/Link 11; 

• ILS: glideslope, localiser and marker receivers; 

• Microwave landing system (MLS) - growth/space provision; 


• IFF interrogator - this incorporates an electronically scanned linear array; 

• IFF transponder. 

All CNI antennas are conformal to preserve the aircraft stealth characteristics; 

9.5.5 Displays/Cockpit 

The display suite comprises a HUD and colour AMLCD head-down displays and an 
integrated control panel as follows (Figure 9.15): 

1. The wide-angle HUD has a 30° (H) x 25° (V) FOV. 

2. The integrated control panel (ICP) provides the primary means of inputting data into the 
avionics system. 

3. Two 3 in x 4 in up-front displays (UFDs) portray CNI settings, warnings and cautions and 
standby flight instrumentation and fuel indications. 

4. The primary multifunction display (PMFD) is the primary pilot's display providing 
navigational data such as waypoints and flight route as well as SA information about the 
threat scenario. 

5. Three secondary MFDs (SMFDs) are used to display tactical information as well as 
aircraft management data: checklists, system status, engine thrust, stores management, 

6. HOTAS controls using the side-stick controller and both throttles. Up to 60 functions may 
be controlled via HOTAS. 

7. NVG compatible lighting. 



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FOV: 30°x20° 

Integrated Control 
Panel (ICP) 

Up-Front Display 

Primary Multi- Function 
Display (PMFD) [8"x8 M ] 

Secondary Multi- Function 
Display x 3 [ 6.25" x 6.25" ] 


Figure 9.15 F/A-22 cockpit schematic. 

9.5.6 Vehicle Management System 

The vehicle management system (VMS) integrates the flight control, engine control and 
aircraft utilities control into one major subsystem. The integrated vehicle subsystem 
controller (IVSC) integrates all the utilities subsystems: 

• Environmental control system; 

• Fire protection; 

• Auxiliary power generation system (APGS); 

• Landing gear; 

• Fuel system; 

• Electrical system; 

• Hydraulics; 

• Arresting system. 

9.5.7 Weapons 

The weapons on the F/A-22 may be carried in three ways (Table 9.4): 

• Inside the centre weapons bay; 

• Inside the left and right side bays; 

• External carriage of fuel tanks and missiles for ferry purposes. 

9.6 Nimrod MRA4 

The Nimrod MRA4 is a total redesign of the original MR1/2 aircraft with totally new 
systems and remanufactured wings and empennage. The MRA4 first flew on 27 August 2004 









































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and is due to enter service in 2008. The Nimrod MRA4 has a federated architecture as shown 
in Figure 9.16. The key MIL-STD-1553B data buses are: 

• Air vehicle bus integrating NAV/FMS, FDDS/displays and utilities control functions; 

• Communications bus integrating the communications systems; 

• Sensor bus interfacing the maritime search radar, ESM, magnetic anomaly detector 
(MAD), acoustics and electrooptic sensors; 

• Mission bus integrating mission control, station control and the DASS; 

• Defensive aids subsystem (DASS) interfacing to the EW/ESM suite. 

Besides this main framework there are many other buses including Ethernet lOBaseT and 
fast switched ethernet (FDX) 100BaseT COTS buses as most of the mission suite is COTS 
based. Most of the data buses interface directly with the two main computers. 
The main functional areas are: 

• Navigation and displays; 

• Utilities control; 

• Communications; 

• Mission system and sensors; 

• DASS. 

9.6. 1 Navigation and Displays 

The navigation and displays are based upon civil flight deck architectures using ARINC 429 
data buses for intercommunication within the subsystems as well as interfacing to the dual- 
redundant MIL-STD-1553B air vehicle bus. 

The flight management system (FMS) is a dual-computer, dual multifunction control and 
display unit (MCDU) system based upon that flying in commercial Boeing 737s, albeit with 
additional modes of operation. The civilian FMS is described in detail in the companion 
volume (Moir and Seabridge, 2003), Chapter 8 - Navigation. The dual-FMS computers 
and captain's and first officer's navigation displays and MPCDs are connected as shown in 
Figure 9.17. 

The flight deck display system (FDDS) is likewise based upon an Airbus flight deck 
display implementation. Two display computers with in-built redundancy supply six colour 
multifunction displays that are usually configured as follows (see Moir and Seabridge 
(2003), Chapter 7 - Displays): 

• Captain: primary flight display (PFD) and navigation display (ND); 

• First officer: PFD and ND; 

• Central displays: associated with alerts and warnings, system synoptic displays and status 

This is typical of civil aircraft platforms that are adapted for military use and allows a cost- 
effective way of adapting proven civil systems and integrating with the specific needs of the 
military aircraft mission system. 

The air data, attitude and inertial/GPS navigation data are provided by units connected to 
the air vehicle MIL-STD-1553B bus. 



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Heading Lubber 
DME Distance 
VOR bearing 
ADF Bearing 
Intended Flight Path 

NAV Mode 
ILS Frequency 
VHF Frequency 
ILS Frequency 
Time to Waypoint 





oo© 1 











Figure 9.17 Nimrod MR4 FMS based upon the Boeing 737 system. 

9.6.2 Utilities Control 

Control of the utilities systems is accomplished by the following: 

1. An engine monitor unit connecting to the four BR 710 full- authority digital engine control 
(FADEC) units - one for each engine. The baseline FADEC was designed with ARINC 
429 data bus interfaces, and this unit connects the engine controllers to the aircraft systems. 

2. Dedicated controllers for the following subsystems: 

• Environmental control system (ECS); 

• Auxiliary power unit (APU); 

• Braking and steering; 

• Electrical system. 

3. Four multisystem utilities controllers based upon the integration principles proven on the 
experimental aircraft programme (EAP) - the technology demonstrator forerunner to 

9. 6.3 Communications 

The communications system has its own dedicated 1553B bus connecting the following: 

• Several V/UHF radio sets; 

• HF communications; 


• Link 11 and link 16 data links; 



• Secure communications; 

• Internal communications. 

With a mission crew complement of up to 11 operators, it is possible, indeed likely, that 
several will be communicating at any one time using a combination of these communications 
assets either in secure or non-secure modes. 

9.6.4 Mission and Sensor Systems 

The main computers connect to the sensor and mission buses. The sensor bus connects the 
following equipment: 

1 . Searchwater 2000 MR radar which provides the primary mission sensor. The radar has the 
following modes: 

• Anti-submarine warfare (ASW) and anti- surface warfare (ASuW) capability in open 
seas or littoral waters; 

• Synthetic aperture radar (SAR) with swath capability; 

• Inverse SAR (ISAR) mode; 

• Pulsed Doppler mode for air-to-air operation; 

• IFF interrogator. 

2. Magnetic anomaly detector. 

3. An extensive acoustics/sonics suite. 

4. Electrooptics. The aircraft is fitted with a retractable Northrop Grumman Nighthunter II 
EO turret with a combination of sensors: 

• FLIR using MWIR and LWIR detectors; 

• Laser range finder; 

• Colour TV using CCD devices. 

This system embodies Litening pod and F-35 electrooptic surveillance and detection system 
(EOSDS) technology to provide leading-edge sensing capabilities. 
The mission bus connects the following: 

1. Electronics support measures (ESM) suite. 

2. Interface with the DASS bus. 

3. Connection of the seven mission workstations and the station management system which 
in turn controls the seven station control units. 

9.6.5 DASS 

The DASS bus connects the following 

• Radar warning receiver; 

• Missile warner; 

• Towed decoy array; 

• Chaff/flare dispenser; 



• Future additions are likely to include: laser warning, directional IR countermeasures 
(DIRCM) and jammers. 

9.6.6 Weapons and Stores 

The weapons and stores load includes the following located in the bomb bay and various 
underwing stations: 

• Spearfish; 

• Air-to-surface/anti-ship missiles (Harpoon); 

• Mines; 

• Search and rescue stores; 

• Flares and smoke markers. 

9.7 F-35 Joint Strike Fighter 

The evolution of the JIAWG (F/A-22) and JAST (F-35) architectures is described in 
Chapter 2. Two key features that distinguish these architectures are: 

• The centralisation of the avionics computing function into multiprocessor signal and data 
processing resources in integrated avionics racks with extensive use of high-bandwidth 
fibre-optic buses for interconnection; 

• The rationalisation of RF systems into a common integrated sensor system (ISS) utilising 
shared apertures and frequency conversion modules. 

The F-35 avionics architecture is depicted in Figure 9.18 and comprises the following major 
subsystems interconnected by fibre channel (FC): 






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Integrated Common 
Processor #1 

Integrated Common 
Processor #2 

■ Fiber Channel - 



IEEE 1394Firewire— 

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Figure 9.18 F-35 - avionics system top-level architecture. 


• Central integrated computing resource comprising two integrated common processors 

• AESA radar; 

• Integrated EW/CNI and EO systems; 

• Display suite; 

• Vehicle management system (VMS); 

• Weapons. 

9.7.1 Integrated Common Processors 

The F-35 common integrated processors (ICPs) comprise the backbone computing resource 
for the aircraft. This function is packaged into two racks with 23 and 8 slots respectively and 
consolidates all the signal and data processing tasks formerly managed by a range of 
dedicated processors. The ICPs use COTS components to enable future upgrades to be 
readily incorporated. The present baseline uses Motorola G4 Power PC microprocessors. At 
initial operational capability the ICP will be capable of performing calculations at the 
following rates: 

• Data processing: 40.8 billion operations/s (40.8 GOPS); 

• Signal processing: 75.6 billion operations/s (75.6 GOPS); 

• Image processing: 225.6 billion multiply /accumulate operations/s. 

The design incorporates 22 modules of different generic types: 

• Four general-purpose (GP) processing modules; 

• Two general purpose I/O (GPIO) modules; 

• Five special I/O modules; 

• Two image processor modules; 

• Two switch modules (these are 32 port FC switches that interconnect the FC elements); 

• Five power supply modules. 

Future growth allows for an additional power supply and eight more digital modules. The 
ICP uses Green Hills Software Integrity commercial RTOS for data processing and Mercury 
Systems' commercial multicomputing operating system (MCOS) for signal processing. The 
CNI and display computers also use the Integrity RTOS to provide an upgrade path to allow 
for future developments. 

9.7.2 AESA Radar 

The AN/APG-81 AESA radar has all the multimode operation and benefits of the active 
radar as described in Chapter 4. The 1200 array AESA radar is said to have a detection range 
of 95 nm against aim 2 cruise missile target. 

9.7.3 Integrated EW/CNI and EO Systems 

The EW subsystem comprises two subsystems: 

• The ESM subsystem from Northrop Grumman; 

• The radar warning and electronic countermeasures system from BAE SYSTEMS. 


The integrated CNI subsystem integrates the following elements: 

• Aircraft communications; 

• Data links; 

• Navigation; 

• Radar altimeter; 

• Identification and interrogation. 

These systems use an integrated antenna suite comprising one S-band (2-4 GHz), two UHF 
(300-1000 MHz), two radar altimeters (-4 GHz) and three L-bands (1-2 GHz) per aircraft. 
The EO systems include two major subsystems as already described in Chapter 5: 

• Electrooptic targeting system (EOTS) providing an EO targeting system (Figure 5.35); 

• Distributed aperture system (DAS) providing 360° coverage (Figure 5.36). 

9.7.4 Displays Suite 

The F-35 display suite extends the F/A-22 display philosophy to a single panoramic 
8 in x 20 in viewing area. Two 8 in x 10 in screens provide the multifunction display system 
(MFDS) projection displays, each with a 1280 x 1024 pixel display resolution. The system is 
effectively split into two such that, if one half fails, the system can continue operating with 
the remaining 'good' half. 

A helmet-mounted display system (HMDS) replaces the conventional HUD, resulting in 
significant cost and weight savings. The HMDS displays flight-critical, threat and safety 
information on the pilot's visor. The system is also able to display imagery derived from the 
distributed aperture system (DAS) or via a helmet-mounted day/night camera. 

Pilot's commands are initiated using the HOTAS system. 

9.7.5 Vehicle Management System 

The vehicle management system (VMS) comprises three vehicle management computers 
which perform calculations for the aircraft flight systems: flight control, fuel systems, 
electrical system and hydraulics, among many others. The vehicle management computers 
are interfaced via IEEE 1394b data buses and continuously compare computations across all 
three computers to ensure integrity. In the event of a disagreement between the units, 
arbitration techniques permit the 'voting out' of a defective processor or sensor input. 

The system also has 10 remote interface units (RIUs) that collect data for interfacing to the 
VMS computers. These are distributed at suitable points throughout the airframe to act as 
collection or distribution agencies for all the VMS I/O signals. 

9.7.6 Weapons 

The F-35 will be cleared for the following range of weapons and stores: 

1. Internal carriage: 

• Joint direct attack munition (JDAM); 

• CBU-105 wind-corrected munitions dispenser (WCMD); 

• Joint stand-off weapon (JSOW); 



• Paveway II guided bombs; 


2. External carriage: 

• Joint air-to- surface stand-off missile (JASSM); 

• AIM-9X Sidewinder; 

• Storm Shadow cruise missile; 

• External fuel tanks. 

9.7.7 Gun 

The gun has yet to be selected - the US Air Force (F-35A) will have an integral gun whereas 
the marine (F-35B) and carrier (F-35C) versions will have an external gun pod. 

9.8 MIL-STD-1760 Standard Stores Interface 

The military standard, MIL-STD-1760, was developed to ensure a standard interface 
between weapons or stores and the carriage aircraft electrical and avionics systems. This 
greatly facilitates the carriage of a particular weapon type across a wide range of platforms, 
reducing development effort and maximising operational flexibility. MIL-STD-1760D was 
released on 1 August 2003. 

The aim of MIL-STD-1760 is to provide a common interface between aircraft and stores 

• The electrical and fibre-optic interfaces at aircraft stores stations, the interface on mission 
stores the interface on carriage stores, and the characteristics of umbilical cables; 

• Interrelationships between aircraft and stores interfaces; 

• Interrelationships between stores interfaces at different stores stations on the aircraft. 

The electrical interfaces covered by the standard are: 

• Aircraft station interface (ASI); 

• Carriage store interface (CSI); 

• Carriage store station interface (CSSI); 

• Mission store interface (MSI). 

Refer to Figure 9.19 which is equivalent to Figure 1 in MIL-STD-1760D. 

There are many possible combinations given the large number of weapons stations 
on modern fighter aircraft and the wide variety weapon types. Figure 9.20 portrays two 

1. Figure 9.20a shows a very simple stores interface. 

2. Figure 9.20b depicts a complex arrangement where two intelligent stores are commu- 
nicating via the aircraft communications network. This might be typical of a EO pod 
handing off target coordinates to a smart weapon elsewhere on the aircraft. 





' ' n I 3 jY^X MISSION 

j/ff/U store 




Figure 9.19 Aircraft store configuration examples (MIL-STD-1760D). 

The MIL-STD-1760 signal set is listed in Table 9.5. These are further classified for particular 

• Class I - basic interface set; 

• Class I A - basic interface set plus the auxiliary power lines; 





Figure 9.20 Examples of stores interfaces (MIL-STD-1760D) 


Table 9.5 MIL-STD-1760 signal set. 

Aircraft Store 

Signal lines 

< < < < High bandwidth 1 > > > > 

< < < < High bandwidth 2 > > > > 

< < < < High bandwidth 3 > > > > 

< < < < High bandwidth 4 > > > > 
<<<< MUX A >>>> 
«« MUXB >»> 

< < < < Low bandwidth > > > > 
<<<< Fibre optic 1 >>>> 
<<<< Fibre optic 2 >>>> 

Discrete lines 

Release consent >>>> 

Interlock >>>> 

<<<< Interlock return 

Address BIT 4 »» 

Address BIT 3 »» 

Address BIT 2 >>>> 

Address BIT 1 »» 

Address BIT »» 

Address BIT parity >>>> 

<<<< Address return 

<<<< Structure ground 

Power lines 

28 V DC power 1 »» 

<<<< Power 1 return 

28 V DC power 2 »» 

<<<< Power 2 return 

115 VAC phase A »» 

115 VAC phase B »» 

115 VAC phase C »» 

<<<< 115 VAC neutral 

270 V DC power »>> 

<<« 270 V DC return 

Auxiliary signals 

Auxiliary signals provide 
an additional set of 28 V DC 
power, 115 VAC power, 
270 V DC power and interlock 
and structure ground signals 

Note: <<<< denotes a signal transiting left; >>>> denotes a signal transiting right. 
A combination of the two indicates a bidirectional signal. 


• Class II - basic interface set excluding high-bandwidth 2 and 4 signals and fibre optics 1 
and 2; 

• Class II A - as for class II plus the auxiliary power lines. 

This classification scheme therefore allows the stores interface to be standardised but takes 
account of the fact that a smart weapon employing fibre-optic communication may not be 
used at every station. The auxiliary power lines allow for the fact that certain stores such as 
EO pods may have much higher electrical power requirements than simple stores. 

High-bandwidth signals may be one of two types: type A from 20 Hz to 20 MHz and type 
B from 20Mhz to 1.6 GHz. Low bandwidth signals are those between 300 Hz and 3.4kHz. 

The specific provisions of the D standard which has been recently introduced are: 

• New provisions for 270 V DC power; 

• Additional data time tagging criteria; 

• Characterisation of the GPS RF signals. 

9.9 Air-to-Air Missiles 

Some of the common air-to-air missiles are briefly described below. These are: 

• AIM-9 Sidewinder; 

• AIM- 120 advanced medium-range air-to-air missile (AMRAAM); 

• AIM- 132 advanced short-range air-to-air missile (ASRAAM). 

These missiles are in service today and are also intended to arm aircraft in development such 
as the F/A-22 and F-35. 

9.9. 1 AIM-9 Sidewinder 

The AIM-9 Sidewinder is a short-range IR guided missile fitted on the aircraft of many air 
forces today. The missile has a long pedigree, being developed during the 1950s with the first 
successful firing in September 1953. The first missiles to be launched in anger were fired by 
the Chinese Nationalist F-86s in 1958 against Mig-17s of Communist China. The missile has 
since undergone many upgrades, with the latest version in operational service being the 
AIM-9M which has the capability of all-aspect engagements. In the early days the electronic 
technology utilised was the vacuum tube whereas it is solid-state today. Kopp (1994) gives an 
excellent insight into the development and capabilities of the Sidewinder. 

Under normal launch conditions the missile generates an audio tone in the pilot's 
earphones when the IR seeker senses a target. The pilot then releases the missile which 
tracks the IR energy being emitted by the target and homes on to the target using a blast 
fragmentation warhead to kill the target. Another method is to use the Sidewinder expanded 
acquisition mode (SEAM) where the missile seeker head is slaved to the aircraft radar. As 
the radar tracks the target the missile seeker head is slaved to the radar boresight, being 
pointed in the direction of the target. 

Finally, for advanced systems using a helmet-mounted display (HMD) it is possible to use 
slaving cues from the pilot's helmet-mounted sight where the missile takes its guidance from 
the direction the pilot is looking. This provides the pilot with a 'first look-first shot' capability. 



Figure 9.21 AIM-9 Sidewinder. 

The AIM-9 complements the longer-range AMRAAM to give a total capability to the 
aircraft self-defence capability (Figure 9.21 and Table 9.6). 

9.9.2 AIM- 120 AMRAAM 

The AIM- 120 advanced medium-range air-to-air missile (AMRAAM) was developed as the 
follow-on missile to the semi-active AIM-7 Sparrow and UK Skyflash, both of which had 
semi-active radar guidance. Semi-active guidance suffers from the disadvantage that the 
launch aircraft has to illuminate the target aircraft throughout the engagement as the missile 
semi-active guidance head is only able to track an illuminated target. This reduces the ability 
of the launch aircraft to engage multiple targets and also makes the launch aircraft vulnerable 
during the flight time of the missile. The AMRAAM overcomes this problem by having an 
active radar seeker guidance capability. As well as the guidance improvements, the 
AMRAAM is smaller, faster and lighter than its predecessor and has improved capabilities 
against low-altitude targets. 

The development of AMRAAM grew out of a joint agreement between the United States 
and several NATO countries that also included the ASRAAM with a view to establishing a 
common technology baseline and joint production. In the event, the joint agreement lapsed 
and the United States continued with development of the missile alone but as a joint US Air 
Force/US Navy programme. The missile exists in a number of versions: 

Table 9.6 ASRAAM characteristics 








Short-range air-to-air missile 

IR guidance single sensor for early versions, 

12x128 FPA for advanced versions 
Supersonic Mach 3+ 
10-18 miles depending upon altitude 
85.5 kg (1901b) 
2.87 m (9 ft 6 in) 
13 cm (5 in) 
Blast annular fragmentation 



Figure 9.22 AIM- 132 ASRAAM. (BAE SYSTEMS © 2005) 

1. AIM-120A, the original production version with deliveries commencing in 1988. This 
variant requires hardware modification to reprogramme the missile. 

2. The AIM-120B and AIM-120C versions presently in production. These variants feature 
smaller control surfaces to facilitate internal carriage in the F/A-22. They are also 
software programmable which enhances operational flexibility. 

3. A preprogrammed product improvement (P 3 I) version. This features improved software 
reprogrammability, advanced countermeasures and improvements in the propulsion system. 

9.9.3 AIM-132 ASRAAM 

The advanced short-range air-to-air missile is a European development to replace the AIM-9 
Sidewinder. It was initiated in the 1980s by Germany and the United Kingdom but the two 
nations were unable to agree upon the details of the joint venture; Germany left the project 
in 1995 and initiated its own version of the improved Sidewinder - IRIS-T. The United 
Kingdom continued with the development of ASRAAM and began to equip its aircraft with 
the missile in 1998. The Australian Air Force purchased the missile in 1998 for use on the 
F/A-18 (Figure 9.22 and Table 9.7). 

9.10 Air-to-Ground Ordnance 

The earliest air-to-ground missiles were used during the Vietnam War, examples such as 
BullPup and Helldog were typical examples. As the war developed it was realised that the 
release envelopes for close air support weapons were very hazardous for the pilot when used 

Table 9.7 ASRAAM characteristics 









Short-range air-to-air missile 
IR guidance plus strapdown IN 
Supersonic Mach 3+ 
300 m to 15 km 
100 kg (2201b) 
2.73 m (8 ft 11 in) 
16.8 cm (6.6 in) 
Blast fragmentation 


against concentrated ground fire; small arms fire could down an aircraft just as easily as anti- 
aircraft artillery (AAA) or surface-to-air missiles (SAMs). 

The effectiveness of air power used against strategic targets was boosted with the advent 
of laser-guided bombs (LGBs) used against bridges and bunkers and anti-radar missiles 
(ARMs) to engage air defence and missile fire control radars. Later generations of this 
technique, augmented by inertial navigation (IN) and GPS guidance, have further improved 
accuracy and effectiveness. 

The air-to-ground missiles and ordnance described in the section include: 

• Wind-corrected munition dispenser (WCMD); 

• Joint direct attack munition (JDAM); 

• AGM-88 high-speed anti-radiation missile (HARM); 

• Air-launched anti-radar missile (ALARM); 

• Storm Shadow/SCALP EN. 

9. 10. 1 Wind-corrected Munition Dispenser 

The wind-corrected munition dispenser (WCMD) programme was developed to provide im- 
proved guidance for submunition dispensers. These weapons were developed for deployment 
at low level against heavily defended targets and are intended to produce an area denial 
effect on targets such as airfields. During Desert Storm, low-level tactics were not the 
preferred weapon of choice, preference being given to precision weapons launched from 
medium level. Nevertheless, such weapons can have merit in use against certain types of 
target and the WCMD is intended to upgrade existing dispensers and provide greater 
accuracy by countering the effects of wind. 

The WCMD includes a tail kit to add to existing ordnance to provide greater accuracy. The 
guidance is purely by inertial means with no GPS guidance added. 

9. 10.2 Joint Direct Attack Munition 

The joint direct attack munition (JDAM) is a tail kit to be added to existing 'dumb' bombs to 
allow their use in a precision mode. The JDAM kit consists of a guidance package that fits on 
to the tail section of the existing bomb and strakes fitted along the side of the bomb to 
enhance aerodynamic effects. 

The guidance used is a tightly coupled INS/GPS system that can operate in both GPS- 
aided INS and INS-only modes of operation. The specified accuracies of circular error 
probabilities (CEPs) are 13 m and 30 m respectively, although it is understood that accuracies 
better than this are regularly achieved. JDAM kits are being fitted to a range of bombs 
including 250, 500, 1000 and 20001b existing ordnance, some with a hard target penetration 

As well as improving accuracy, the aerodynamic characteristics of a JDAM-fitted bomb 
allow a lateral footprint to be accommodated. This means that a bomb may be deployed 
within reason against targets that do not lie immediately on the track of the launch aircraft. 
This offers significant operational flexibility so that an aircraft can fly over a dense target 
environment, engaging targets lying to the left and right of the aircraft track and allowing 
multiple targets to be engaged in a single pass. The combination of this operational 


capability combined with the accuracy of the JDAM guidance package can deliver 
devastating effects, as has recently been demonstrated. 

9.10.3 AGM-88HARM 

The AGM-88 HARM missile was developed as a successor to early ARMs such as the AGM- 
45 Shrike that was used during the Vietnam War. The missile was fitted to the US Air Force 
F-4G Wild Weasel aircraft before they were withdrawn from service in the early 1990s. 
Certain versions of the F-16 now discharge this role, and Navy F/A-18s are also capable of 
deploying the missile. 

The launch aircraft provides the missile with data relating to the radar of the target to 
be engaged. The missile seeker head acquires the target and, after launch, uses the RF 
energy to home in on the radar, destroying it with a high-explosive (HE) direct 
fragmentation warhead. 

The target radar may counter an ARM by using emission control (EMCON) procedures, 
namely by switching off the radar when missile launch is detected. In certain circumstances 
these may achieve the desired operational effect if the launch aircraft is screening a 
multiaircraft raid. In other situations it may just result in a lost missile. 

9.10.4 ALARM 

The air-launched anti-radiation missile (ALARM) is a UK missile developed to fulfil the 
same mission as HARM. ALARM entered service in the early 1990s and was successfully 
used during Desert Storm. The missile has the capability of ascending to 40000 ft and 
loitering if the enemy radar is switched off. This is accomplished by deploying a parachute 
which allows the missile to descend slowly but still in an active mode. If the target radar 
recommences radiating, the missile will detect the target, release the parachute and fall to the 
target using gravity. The missile also has a memory so that, if the target fails to reradiate, the 
missile will attack the last known position. 

9.10.5 Storm Shadow/SCALP EN 

Storm Shadow is the outcome of a joint UK/French development of the Apache air-breathing 
missile; the two governments have evolved a common technical solution, albeit the precise 
implementations are slightly different. Storm Shadow is a stealthy cruise missile of 
^1300 kg (28601b), capable of delivering a conventional warhead at ranges of over 
250 km (156 nm) against range of high- value targets such as C3 facilities, airfields, ports, 
ammunition/storage depots and bridges. 

The guidance and warhead attributes of Storm Shadow are: 

• Guidance based upon a terrain profile matching (TERPROM) system with integrated GPS; 

• Terminal guidance using an IR sensor and an autonomous target recognition feature; 

• A highly lethal bomb Royal Ordnance augmented charge (BROACH) warhead. 

Storm Shadow will be deployed on the Tornado GR4/4A, Harrier GR7 and Eurofighter/ 



www, eurofighter- star steak, net 

Flight International F-16E/F cut-away drawing. 

Flight International AH-64C/D Longbow Apache cut-away drawing. 

Flight International Typhoon cut-away drawing. 

Flight International F/A-22 cut-away drawing. 


Fulghum, D. A. (2000) New F-22 radar unveils future. Aviation Week and Space Technology, 1 February. 
Kopp, C. (1994) The Sidewinder story - the evolution of the AIM-9 missile. Australian Aviation, April. 
MIL-STD-1760D (2003) Interface standard for aircraft/store electrical interfaces, 1 August. 
Moir, I. and Seabridge, A.G. (2003) Civil Avionics Systems. Professional Engineering Publishing/ AIAA. 

1 Vehicle Management 

10.1 Introduction 

The utility systems are a collection of fluid, air, mechanical and electrical systems associated 
with the provision of sources of power or energy to perform the general or utility functions of 
the basic air vehicle. This control is usually obtained by the performance of some functional 
activity resulting in the appropriate control of that energy to impart flow or motion. These 
systems are also known as: 

• Flight systems; 

• General systems; 

• Aircraft systems; 

• Power and mechanical systems; 

• Vehicle systems. 

Figure 10.1 shows the aircraft considered as a set of systems to illustrate the position of 
utility systems relative to the airframe, the avionics and the mission systems. These systems 
can and do operate autonomously, but there are important system interactions that need to be 
considered in the design of each system. Examples are: 

1. Fuel is often used to maintain the aircraft centre of gravity within certain limits. This is 
especially relevant to an unstable fighter aircraft where eg has an impact on manoeuvr- 
ability and is dependent on fuel status and external weapons carriage. 

2. Fuel is also used as a cooling medium, for example in fuel-cooled oil coolers, and its 
temperature needs to be carefully monitored. 

3. Undercarriage oleo switches are used as an indication of weight on wheels, and these 
signals are used to inhibit some functions. 

Military Avionics Systems Ian Moir and Allan G. Seabridge 
© 2006 John Wiley & Sons, Ltd. ISBN: 0-470-01632-9 



An Aircraft 



Utility Systems 


Avionic Systems 

The major structural 

aspects of the 






Structural Integrity 

The systems that enable 

the aircraft to continue 

to fly safely throughout 

the mission 

The systems that enable 

the aircraft to fulfil its 

operational role: 


* Process 

* Control 



\^> Flight deck displays 

Js \ Flight deck warnings 

- rv ^ > Prognostics & health 


Accident data recording 

Predominantly Hard wired 
interconnected and 
mechanical systems 

Predominantly Data Bus 
interconnected information 
management systems 

Figure 10.1 The aircraft as a set of systems. 

The utility systems of an aircraft have a particular set of characteristics that make them 
challenging to interface and control, for example: 

1. They are predominantly mechanical or electrical. 

2. Their operation usually involves a transfer of energy. 

3. They have a diverse range of input and output characteristics. 

4. In earlier aircraft they were controlled by relay logic, hydromechanical devices or 
individual controllers and indicators/warnings. 

5. In current aircraft there is a need to connect them to data bus interconnected avionic 
systems, especially to gain access to the human-machine interface with controls and 
electronic displays. 

6. In high-performance, highly integrated aircraft systems there is a need to improve 
performance and increase knowledge of system operation and failure modes. 

In the modern aircraft the utility systems must coexist with the avionic systems. There is an 
interdependency - the avionic systems need electrical power and cooling air, both of which 
are generated by the set of utility systems, and the utility systems need to be connected to the 
human-machine interface in the cockpit, which is largely generated by the avionic systems. 
In addition there are data generated within each group of systems that must be exchanged. 
Examples of such data include: 

• Air data - altitude, airspeed, Mach number, attitude; 

• Weight on wheels; 

Stores configuration; 



• Navigation steering commands; 

• Fuel centre of gravity. 

The vehicle systems exist to provide the airframe with the capability to perform as an air 
vehicle and to provide suitable conditions for the carriage of the crew and the avionic or 
mission systems. This they do by providing controlled sources of energy to move the vehicle, 
and to control its direction, providing electrical energy and cooling, as well as providing for 
the health and survival of the crew. There is a need for interconnection or integration with 
the mission avionics systems to improve the effectiveness of the air vehicle. In a modern 
aircraft the vehicle management system has the ability to facilitate this interconnection by 
its connections to the aircraft data bus systems, and by the fact that it stores information 
about the performance of the vehicle systems in a format that is compatible with the 
mission system computing architecture. Thus, the predominantly mechanical and electrical 
utility systems can be interconnected with the predominantly data- and information-driven 
avionic systems. This chapter will provide a brief description of the utility systems and their 
key interfaces with avionics. 

10.2 Historical Development of Control of Utility Systems 

The characteristics of utility systems resulted in a gradual development from predominantly 
mechanical and hydromechanical systems towards a system where electrical and electronic 
control of functions became feasible and attractive. Although this progress removed the 
dependence on mechanical items such as cams, levers, control rods or wires and pulleys, it 
did result in a large number of individual solutions. The consequence of this was a large 
number of separate control units of differing technologies and shapes, each with its own 
installation issues and setting and repair tasks. 

A determined effort to reduce this diversity was attempted in the mid-1980s which 
resulted in microprocessor-based architectures for control. Figure 10.2 (Moir and Seabridge, 
1986) shows the reductions in the number of control units achieved as a result of an 
experimental aircraft project in the United Kingdom. 

Conventional control system 

Utilities system management 


- 20 to 25 Dedicated LRU's 

- 6 Power switching relay units 

- Extensive overhead of wiring, 
connectors, mounting trays 

-Dedicated instruments/display panels 


- Four dedicated LRU's 

- Dedicated 1 553B data bus 

- Distributed intelligence 

- Integrated power switching devices 
_ Interfaces with modern digital 

cockpit and multi-function displays 

Figure 10.2 Historical view of utility systems control equipment. 




Number of 

Number of 
Power Drives 

Engine & Associated Systems 



Fuel Gauging & Fuel Management 



Hydraulic Systems including Gear & 



ECS/Cabin Temperature 



Secondary Power Systems 



Miscellaneous Systems - LOX 
contents; Electrical system 
monitoring; Probe heating etc 






Figure 10.3 Early indication of utility system signal numbers. 

Figure 10.3 (Moir and Seabridge, 1986) gives an indication of the number of individual 
signals connected to control units in a conventional aircraft in the 1980s. This situation also 
needed to be resolved because each signal generally required more than one wire and the 
resulting proliferation of wires led to large and heavy cable harnesses with the attendant 
installation issues and the need for large multipin connectors. It was estimated that a 
considerable mass and maintenance penalty could be reduced by the use of a utility systems 
specific data bus. 

The resulting solution was based on a small number of processing units widely spaced 
in the airframe. Interfacing was local to those units so that lengthy wiring runs were reduced 
to the distance between utility system components and the processing units. Any signals 
required by the cockpit or by other systems were transferred by a standard serial data bus. 

What emerged from this activity was a generic architecture that allowed utility systems to 
become integrated with the major avionics systems and the various types of data bus. The 
structure illustrated in Figure 10.4 shows the key interfaces that need to be respected in the 




Vehicle Systems Components 


| Hardware Interfacing | 


Functional Processing | 

Data Bus Interfacing] 

Avionics Data Bus 

Vehicle Systems Data Bus 

Avionic systems 
including cockpit 

Other aircraft systems 

Figure 10.4 Principal utility system interfaces. 


modern aircraft. The agreement of standards and conventions for designing these interfaces 
early in the product life cycle enables the designers of the numerous systems of the aircraft to 
work independently, secure in the knowledge that their systems will work in harmony. 
Thus, the main objectives of the processing units became those of: 

• Interfacing with hardware components and converting analogue and discrete signals into 
data format, as well as converting digital commands into a power command; 

• Performing whatever control function was required; 

• Receiving signals from, and passing signals in an appropriate fashion to, a data bus. 

10.3 Summary of Utility Systems 

The utility systems provide the basic functions that enable the airframe to fly and allow the 
safe carriage of crew, passengers and stores. Not all of these systems will be on every aircraft 
type, and the varying range of ages of aircraft in service means that the technology will be 
different. The utility systems may be subdivided into a set of subsystems. 

10.3. 1 Mechanical Systems 

Those systems associated with control of the flight path of the aircraft and with landing, 
steering and braking: 

• Primary flight controls; 

• Secondary flight controls; 

• Landing gear; 

• Wheels, brakes and tyres; 

• Arrestor hook/brake parachute; 

• Actuation mechanisms. 

10.3.2 Crew Systems 

Systems associated with the comfort of the crew in normal flight, the safe escape of crew and 
passengers under emergency conditions and the continued well-being in the presence of 
chemical and biological threats: 

• Crew/passenger escape; 

• Aircrew clothing; 

• Life support; 

• Oxygen/on-board oxygen generation system; 

• Canopy jettison/fragmentation. 

10.3.3 Power Systems 

Those systems associated with providing a source of energy in the form of thrust, rotational 
or air off- take, electrical or hydraulic energy: 

• Propulsion; 

• Secondary power; 


• Emergency power; 

• Electrical power generation and distribution; 

• Hydraulic power generation and distribution. 

10.3.4 Fuel System 

The system that stores fuel and controls the transfer of fuel from the storage tanks to the 
main propulsion units, providing information to the crew on the status of fuel quantity and 
location within the aircraft tanks. 

• Fuel system architecture; 

• Fuel gauging; 

• Fuel feed/venting; 

• Fuel management. 

10.3.5 Air Systems 

Systems associated with the provision of clean air at appropriate temperature and humidity 
to provide a safe and comfortable environment for crew and passengers, to enable systems 
equipment to operate throughout a wide range of ambient temperature conditions and to 
provide air for operation of actuators, demisting or de-icing systems: 

• Cooling system; 

• Cabin cooling air distribution; 

• Equipment cooling air distribution; 

• De/anti-icing; 

• Canopy demist. 

10.3.6 Electrical Power Distribution Systems 

These systems are associated with the distribution and protection of electric power 
throughout the aircraft: 

• Primary power distribution and protection; 

• Secondary power distribution and protection; 

• Power switching; 

• Load shedding and restoration; 

• Internal lighting; 

• External lighting; 

• Probe heating; 

• Ice detection and protection; 

• Windscreen de-icing. 

10.3.7 Vehicle Management System (VMS) 

An integrated computing system to perform data acquisition, functional control and energy 
control for the utility systems: 


• System architecture; 

• System interfaces; 

• Functional requirements. 

In some more recent systems the VMS also integrates engine control with flight control to 
form an integrated flight and propulsion control (IFPC) system. In others the VMS will 
comprise the overall integration of flight control, propulsion control and utilities control 
within the same entity. 

This system is essentially the integrating mechanism for utility or vehicle systems - the 
control and monitoring functions of the systems are brought together in a single computing 
architecture. This is very close to being 'avionics' in nature - it is an open architecture 
system with appropriate input/output and data bus interfacing. Although implemented as a 
separate system today, it is most likely that future aircraft will incorporate this function into a 
single avionic and mission computing structure. 

10.3.8 Prognostics and Health Management 

A system that acquires data from the aircraft systems in order to diagnose deteriorating 
performance of systems and to provide a prediction of failure mode and time in order to 
provide a preventive maintenance approach with significant improvement in systems 
availability and operating costs: 

• Data acquisition; 

• Data processing; 

• Data recording. 

10.4 Control of Utility Systems 

Many of the subsystems described above require some functions to be performed in order to 
ensure that the subsystems operate satisfactorily. Although many of the systems are 
predominantly mechanical in nature, the functions are increasingly being performed by 
software in a computing system. There are many advantages to this, since modifications to 
functions can more easily be achieved by a software change rather than by redesigning and 
manufacturing mechanical components. 

Figure 10.5 shows a general arrangement of a utility management processing system that 
allows the following key aspects of interfacing and control to be performed: 

1 . It provides a means of interfacing a number of different types of vehicle system input 
component to the processing system characterised by a diversity of type, range, source 
impedance and slewing rate. These interfaces include: 

• relay or switch discrete 28 V or V; 

• fuel gauge probe AC capacitance or fibre optic; 

• gearbox speed pulse probe; 

• actuator position potentiometer or variable differential transformer; 

• temperature thermistor or platinum resistance; 

• demand position sensor, e.g. throttle lever. 





Avionics MIL-STD-1553B Data Bus 





(26lbs/140 watts each) 



m "^ ■ 


■ r^i I 
^^ -<? J| 

** 1 



X " 


(44lbs/230 watts each) 



Data Bus 

System Total: 

740 watts 

Figure 10.5 World's first integrated utility management system. 

2. Provide a means of interfacing a number of different types of effector (a device that reacts 
to a low-energy input signal to perform a high-energy transfer task, e.g. an actuator or 
pump) to the processing system characterised by diversity of type, range, load impedance/ 
resistance and reactance. These interfaces include: 

valve 28 V or V discrete power; 

DC motor DC power drive 28 V DC or 270 V DC; 

actuator drive low- voltage analogue; 

Torque motor low-current servo drive; 

Fuel pump high-current AC motor drive; 

Filament lamp Lamp Load (high in-rush current). 



3. It provides a means of connecting to the aircraft data bus structure. These buses include: 

MIL-STD-1553 (Def Stand 00-18, Stanag 3838); 

ARINC 429; 

ARINC 629; 

IEEE 1394 (firewire); 

Fibre-optic buses. 

4. It provides a means of performing subsystem control functions in software. 

5. It provides a means of performing control functions in hardware such as application- 
specific integrated circuits (ASICs), programmable logic devices or hardwired means. 

A typical subsystem may have a combination of software and hardware control imple- 
mentations depending upon the nature of the system operation and the level of integrity 

Figure 10.5 depicts the key features of the first integrated system of its type to fly in the 
world: the British Aerospace (BAe) - now BAE SYSTEMS - experimental aircraft 
programme (EAP). This was a single- aircraft flight demonstrator that proved key technol- 
ogies for Eurofighter Typhoon and that first flew in August 1986. 

Each of the processing units shown in this system provided a facility for interfacing with 
the components of the utility systems. For economic reasons it would be ideal if these units 
were identical in construction, with only the software load providing each unit with its own 
characteristics. There are many practical reasons why this rarely happens. Figure 10.6 shows 
a generic form of such a processing unit. 

As the application of utility management has advanced through a number of military and 
commercial aircraft types, so technology has advanced in areas of semiconductors, memory 
and data transmission systems. The nomenclature has now changed to vehicle management 
systems to reflect broader applications than solely aircraft. 

The vehicle management system can be designed to observe the same architectural 
principles as the avionics and mission systems, for example data bus type, data bus protocols, 
processor and memory type, software language, redundancy requirements and physical 

Data Bus Terminal Bus Controller 

~^ A 

Data Bus 



v> Processor/Memory 

Power Supply Unit 

Figure 10.6 Generic arrangement of a utility management systems processor. 



Vehicle Management 

Remote Input/Output 
Units (RIO) 

High Speed Data Bus 
Figure 10.7 Modern VMS architecture. 

enclosure. Although contemporary aircraft often implement vehicle systems control in a 
separate processing structure to avionics, the two are likely to merge in the future to meet 
demands for more efficient on-board processing and reduced equipment mass. 

The modern vehicle systems management architecture has developed to take account of 
technology improvements, as well as a radical rethink of the installation aspects of the 
system. It makes considerable sense to install high-throughput, high-dissipation processors 
and high-capacity memory in a comfortable, cooled environment to achieve the highest 
availability possible, while changing the method of interfacing with the system components. 

Therefore, units associated with input and output interfacing functions, commonly known 
as remote input/output (RIO) units, can be designed as simple, rugged items packaged so that 
they can be installed in remote areas of the airframe that may be harsh in terms of 
temperature and vibration. Local connections to components will be short, and high-speed 
data links connect the RIOs to the main data bus or processing system. A typical architecture 
with a combination of VMS computers and RIOs is shown in Figure 10.7. 

A characteristic of the majority of utility systems is that they do not generally lend 
themselves to multiple-redundant architectures. Obvious exceptions are flight control 
(generally triple or quadruple redundancy) and propulsion control (generally duo-duplex). 
Otherwise the utility systems are a collection of singular control problems with single 
components to control. As a result, a number of strategies have emerged that are still practised 
in the software-dominant systems currently available for controlling the systems and for dealing 
with failures. A key determining factor in the choice of method is the integrity requirement or 
safety requirement for the system. These strategies can be summarised as follows: 

1. Some systems such as flight control and propulsion control have their own control 
systems, but may make use of utility management to gain access to data in the utility 
systems or to access the cockpit displays and other avionic systems. 



2. Some systems such as fuel systems are wholly controlled and monitored by utility 
management with redundancy or with graceful degradation achieved by distributing 
control among several utility management processors. 

3. Some systems are mainly controlled by utility management but may have alternative 
hardwired means of obtaining manual override or reversionary control, albeit with some 
degradation in performance. 

4. Some systems have no connection to utility management at all. Usually these are systems 
that are concerned with the safety of the aircraft or the crew and may be required to 
operate when the majority of other on-board systems have failed, for example the ejection 

10.5 Subsystem Descriptions 

The utility systems will be described in more detail below. More detailed descriptions 
can be found in Moir and Seabridge (2001 and 2002), Jukes (2004) and Pallet! (1992). 
Figure 10.8 shows a generic block diagram which can be used to represent each of 
the following subsystems. This shows a variety of inputs that determine the status of the 

• Demand from the pilot operating a device in the cockpit - the human-machine interface; 

• Sensor inputs from components in the system; 

• Other system inputs which may be relevant data or commands; 

• Energy to provide power for the processor; 

• Feedback from the system to close the loop on any control processes. 

Also shown is a variety of outputs: 

• Waste products in the form of heat, noise or energy which must be dissipated; 

• Signals and data to the cockpit to be presented on the displays or warning system; 


Waste products 

Flight deck 

Figure 10.8 Generic utility system block diagram. 


• Signals to other systems to demand a response or to provide information; 

• Commands for energy to enable an event to take place. 

10.5. 1 Mechanical Systems Primary Flight Controls 

In its most basic form, a flight control system is a set of rate demand control loops that allows 
the pilot to manoeuvre and control the aircraft attitude in pitch, control and yaw axes. In the 
modern aircraft this function is provided by a system architecture with appropriate safety and 
integrity implemented by bus interconnections, computing and software, control law design, 
actuation systems and pilot's inceptors. The flight control system is always considered to be 
flight safety critical, particularly in unstable air vehicles, since loss of the system may result 
in loss of the aircraft and severe injury or death of the crew or the overflown population. 
Multiple, similar redundancy or multiple redundancy with dissimilar back-up is used as a 
design solution. In most modern solutions the flight control function is performed in a 
dedicated set of computers, and not integrated with vehicle systems management. However, 
improved component reliability and the emergence of robust methods of software design 
coupled with high-speed data buses may lead to a change in this position. 

Systems such as flight controls are flight safety critical since a catastrophic failure can lead 
to loss of the aircraft. For this reason the system must be designed so that no single failure, 
and very often no multiple failures, will lead to loss of control. Common mode failures must 
be avoided, and hence the power sources and supporting sensors must be carefully selected. 
For these reasons, flight control systems have tended to be self-contained, with minimal 
interaction with other systems. Hence, in most modern aircraft the flight control system 
computing, sensors and actuators are virtually independent of all other systems other than 
providing sources of attitude and air data to other systems and relying on hydraulics and 
electrical power for sources of energy. 

In modern aircraft two types of flight control may be used. All those aircraft derived from 
civil transport aircraft and some fighter aircraft platforms exhibit a stable form of flight 
control where the aircraft centre of pressure is located aft of the aircraft centre of gravity 
(CG) position. This means that the aircraft is naturally stable and will resist natural 
perturbations during flight. This natural stability is reflected in the aircraft handling 
characteristics and therefore in the level of integrity required of the system. The disadvantage 
of this configuration is that the tailplane trim forces act in the same direction as the aircraft 
weight vector and the design is aerodynamically inefficient. 

In the second category the aircraft is naturally unstable, usually in the pitch axis; either for 
reasons of high manoeuvrability requirements or because of aerodynamic/stealth drivers 
(F-117A and B-2). A diagrammatic representation of the forces acting upon an unstable 
aircraft such as the EAP is shown in Figure 10.9. 

It can be seen that the CG is aft of the centre of pressure and that the aircraft is unstable. 
The saving grace is that the trim or manoeuvre forces act in the same direction as the lift 
vector, making the aircraft more responsive. The control surface configuration of the EAP is 
depicted in Figure 10.10. This diagram also gives some idea of the primary and secondary 
flight control surfaces, and a very similar configuration was adopted on Eurofighter Typhoon 
owing to the successful handling characteristics demonstrated by this aircraft. The fact that 
the aircraft is highly unstable results in a quadruplex flight control system architecture for 



Trim force 

Trim force 

Figure 10.9 Longitudinal control forces on an unstable platform. 

sensors, computing and actuators since the total loss of flight control would result in loss of 
the aircraft. Similar quadruplex flight control architectures are common on many modern 
high-performance aircraft. 

The outline digital quadruplex architecture to control the EAP airframe is shown in 
Figure 10.11. The key elements are: 

• Flight control computer (FCC); 

• Aircraft motion sensing unit (AMSU); 

• Pilot input sensors; 

• Airsteam direction sensing vanes; 

• Air data; 

• Actuator drive unit. 

Modern vehicle management systems with high-speed data buses and high-reliability 
systems may change this philosophy and lead to more functional integration. In some cases 

Example of flight 
control surfaces - 
EAP (British 

(J) Foreplane for pitch control 
and stabilisation and 
performance optimisation 

(D Intake scheduled 
for performance 
(D Leading edge droop 

scheduled for performance 
and stability 

(?) Primary controls 
(s) Secondary controls 

(p) Rudder for yaw 
trim, control 
and stabilisation 

;G placed well 
aft in the airframe 

'(?) Inboard and outboard 
flaperons for pitch 
control and stabilisation, 
roll trim and control 

Figure 10.10 EAP control surface configuration. 



Aircraft Motion 
Sensing Unit 








Figure 10.11 EAP quadruplex flight control system architecture. 

this type of functional partitioning represents a sound systems approach. However, where 
distributed systems are required to meet high-integrity requirements, it must be remembered 
that the interconnecting data buses need to meet the same requirements. This may constrain 
the application of many standard data buses, particularly those of a COTS origin; in those 
cases the COTS data bus chips may have to be redesigned and revalidated if the high- 
integrity requirement is to be supported. Some subsystem examples are described below. Secondary Flight Controls 

Secondary flight controls include the provision of secondary control surface mechanisms 
(flaps, slats, air brakes, speed brakes, pilot interfaces). Some of these systems are relatively 
simple logical controls with some airspeed or Mach number limitations. There is an 
increasing tendency to include secondary controls in with the primary flight control system, 
as shown in the EAP example above. The main utility interface is to collect information for 
the cockpit displays. Landing Gear 

Landing gear systems include the provision of undercarriage configuration and loads, 
sequencing, doors, locks, indications and warnings, and pilot interface. Position-indicating 
devices such as microswitches or proximity detectors are used to indicate the position of 
doors and the gear legs and wheels so that the correct sequence is followed. The utility 
system task is to perform logical checks on the switching sequence and to ensure that 
extensive built-in testing is performed to verify correct operation of the doors and gear 
sequence. In addition it is essential to understand when the undercarriage oleos are 



system hardware 

Undercarriage and 
Brake system 


Undercarriage monitoring 
Weight-on-wheel signalling 
Normal brakes control 
Emergency brakes control 

Figure 10.12 Landing gear system example. 

compressed - weight on wheels is used to signal to a number of systems that it is safe to 
operate. Examples of this is the use of weight on wheels to inhibit the arming and release of a 
weapon when the aircraft is on the ground, and the actuation of thrust reverse only when the 
aircraft is firmly on the ground. A typical system is shown in Figure 10.12. Wheels, Brakes and Tyres 

Wheels, brakes and tyres provide the mechanism for meeting the loads and performance 
required to match the aircraft role for landing, take-off and ground handling under normal 
and emergency conditions - for nose wheel steering, brakes/anti-skid, park brake, tyres, 
runway loads, etc. Braking and anti-skid are safety-critical, high-speed, closed-loop systems 
and may be duplicated to provide fail-safe operation. The utility control system reacts to 
demands from the pilot's pedals to activate the braking system, and differential speed sensors 
in the wheel detect the onset of a skid and adjust the brake demand accordingly. Arrestor Hook/Brake Parachute 

These subsystems are to provide suitable methods of arresting the aircraft to supplement the 
braking system. Aircraft operating from aircraft carriers will routinely be fitted with an 
arrestor hook. Many fast jets also carry a hook to enable them to engage with arrestor gear on 
an airfield should the need arise - in the event of failed brakes, reverse thrust or brake 
parachute, for example. These are simple logical functions which are usually separate from 
utility system control. 


10.5. 1.6 Actuation Mechanisms 

Throughout there will be a need to provide suitable actuator types and mechanisms for 
operating surfaces. Specifying these mechanisms requires a knowledge of surface travel, 
mechanical advantage, flight loads, failure cases, etc. There may be many doors, flaps or 
vents that need to be operated in response to manual or automatic demands. Hydraulic or air 
power may be used for this purpose. These functions may be simple logic but may have some 
dependencies on air speed, weight on wheels or other system modes that can be provided by 
utility systems control. 

10.5.2 Crew Systems 

10.5.2. 1 Crew Escape 

Crew escape systems provide safe aircrew escape under certain conditions of flight, 
including ejection seat performance, ejection clearances, escape doors, parachute require- 
ments, emergency location devices and survival equipment. Crew escape is a last resort and 
may be needed when all other functions have been lost. For this reason it is independent of 
all other systems. Increasingly sophisticated functions such as autoejection and automatic 
seat trajectory adjustments will be performed by integral controllers. An example of a crew 
escape system is shown in Figure 10.13. 

1 0. 5. 2. 2 Aircrew Clothing 

Military aircrew are provided with crew clothing for anti-g, immersion, restraint, comfort 
and survival in the arduous conditions of a military jet environment. Clothing varies 
according to the type of aircraft - fighter and fast-jet aircrew need to wear clothing to 
protect them against high-g conditions and the ejection case; larger aircraft with multiple 

Figure 10.13 Modern military ejection set - F-35 JSF Seat. (Martin Baker) 


crew members tend to operate in a 'shirtsleeve' or flying overall environment. There are 
some simple functions to be performed such as inflation of garments to anticipate the rapid 
onset of g, but these functions can be independent of utilities control. 

There is an important avionic interface with crew systems at the oxygen mask and the 
helmet. The supply of oxygen under pressure to the mask and the action of pilot breathing in 
close proximity to the communications microphone can distort speech and contribute to 
noise in the pilot's headset. 

The interface at the helmet is also significant, since the additional mass of sensors and 
displays can have an impact on the specification of the escape system and may affect pilot 
health and safety as a result of load on the neck under high-g conditions. Life Support 

The crew are provided with protection against biological and chemical threats, and with 
emergency air/oxygen. This protection may be provided by means of the air-breathing 
system in the case of fast jets, or by portable respirators or personal biological/chemical 
packs. This is usually an independent function. Oxygen/OBOGS 

The aircraft is equipped with systems to provide gaseous oxygen or on-board generated 
oxygen/enriched air and the mechanism for providing oxygen to the crew. Most fast jets 
provide oxygen derived by liquid oxygen (or LOX) or in the form of oxygen-enriched air 
provided by OBOGS. Aircraft with a large cabin are provided with filtered air by the aircraft 
environmental control system. Gaseous oxygen is carried for emergencies, such as depres- 
surisation, where oxygen masks are deployed to individual crew members. Although the 
majority of these functions are independent of utilities, there is a need for control of the 
sequencing of the catalyst beds and monitoring of oxygen concentration in the on-board 
oxygen generation system which is becoming increasingly common to provide a measure of 
independence from supplies of liquid or gaseous oxygen. 

A typical OBOG system is shown in Figure 10.14 in this case for a two-seat aircraft, 
although the architecture is the same for a single- seat aircraft, but with only one regulator. 
The following description is from a Honeywell Aerospace Yeovil paper (Yeoell and 
Kneebone, 2003). 

Engine bleed air enters the preconditioning system element where the temperature is 
reduced, ideally to less than 70°C, and water is removed as far as possible. In addition it is 
normal at this stage to use a combined particulate and coalescing filter to remove potential 
contaminants, including freewater, that may still be contained in the inlet air. 

The OBOGS contains a pressure-reducing valve to reduce the inlet air pressure of the air 
supply to that required by the OBOG generator, typically 35 psig. 

The next system element is the oxygen generator or, more correctly, the OBOGS 
concentrator which uses multiple zeolite beds to produce the oxygen-rich product gas. 

The switching between the zeolite beds is achieved using solenoid-actuated pneumatic 
diaphragm valves controlled by the system monitor/controller which may be located in 
the vehicle management system. These valves are 'wear free' and allow the concentrator to 
be a 'fit and forget' system that requires no scheduled maintenance and exhibits high 













The system monitor/controller is a solid-state electronic device that monitors the PP0 2 
level of the OBOGS concentrator product gas and adjusts the cycling of the beds to produce 
the desired level of oxygen concentration for cockpit altitudes below 15 000 ft. This process 
is known as concentration control and means that no air mix, or dilution, of the product gas is 
required at the regulator, hence preventing the ingress of any smoke or fumes from the cabin 
into the pilot's breathing gas supply. 

The breathing gas then passes to the pilot's breathing regulator, in this case a panel- 
mounted unit is shown, but ejection seat and pilot mounted devices can also be used. 

1 0. 5. 2. 5 Canopy Jettison 

Emergency escape of the crew is enhanced by the provision of a canopy jettison or canopy 
fracture mechanisms to match the requirements for escape. Fast-jet canopies may be 
jettisoned by explosive or rocket motor to remove them rapidly from the aircraft prior to 
ejection. Alternatively, the transparency material may be fractured by an explosive device 
known as miniature detonating cord (MDC) bonded to the canopy material - the pilot then 
ejects through the fractured material. As this system is associated with crew escape, it is 
independent of all other systems. 

10.5.3 Power Systems 

10.5.3. 1 Propulsion System 

There is an increasing tendency for the propulsion unit or power plant to be developed by 
engine manufacturers as a self-contained item with integral starting, full- authority digital 
engine control units and power supplies. Propulsion control has also remained independent 
of the majority of airframe systems. This separation has been strengthened by a solution 
in which the engine manufacturer has traditionally been responsible for the design and 
integration of the total propulsion system. The propulsion system has evolved from early 
implementations where all the electronic control units were mounted on the airframe in a 
conditioned environment. This solution needed large cable harnesses to connect the units to 
the engine across the airframe/engine interface - a considerable installation and maintenance 
penalty. Improved electronic component reliability and build standards that are able to 
tolerate harsh temperature and vibration environments have resulted in engine-mounted 
control systems. This has led to almost total independence of the full- authority digital engine 
control system with self-contained cooling, vibration isolation and power supplies, the 
only connections to the airframe being for the pilot's demand (throttle box) and cockpit 
displays. Secondary Power 

Provision of power to start engines and to maintain electrical and hydraulic power during 
ground operations. This is performed by an auxiliary power unit (APU) which is a small gas 
turbine started by battery and used to drive a generator and hydraulic pump, as well as to 
provide air to starter turbines on the engine(s). In some cases auxiliary power for 
maintenance and engine start is provided by ground power units. The APU controller is 
usually self-contained with demands from the engine start sequence provided by utility 







tommy \\w \ uit 

Figure 10.15 A secondary power/emergency power system. 




Hydraulic Power 


Electrical Power 




management. In some instances the APU may be started and used in flight to provide an 
assisted start to an engine that has flamed out. Emergency Power 

Provision of power during engine failure conditions to ensure safe recovery of the aircraft or 
safe crew escape - includes the emergency power unit (EPU), ram air turbine (RAT), fuel 
cells, etc. There are instances where the aircraft main engine(s) may cease to provide thrust 
or motive power. In this case an emergency power unit is used to allow the aircraft to be 
flown to a safe flight condition so that the engine can be relit. A ram air turbine (RAT) is the 
usual method used, in which a turbine is dropped into the airstream and used to power 
electrical or hydraulic power-generating devices. This system will be commanded to operate 
by utility management when loss of rotation of both engines is detected. A secondary power 
system together with emergency power provision is shown in Figure 10.15. 

This figure shows how an airframe-mounted accessory gearbox (AMAD) derives shaft 
power from the engine and drives the 3000 psi (F-15), 4000 psi (Tornado) or 5000 psi (part of 
F/A-18E/F and V-22) hydraulic system by means of an engine-driven pump (EDP). In the 
event that both engines are lost, emergency power is provided by an emergency power unit or 
by a ram air turbine (RAT) until such time that the engines may be relit. If a single engine is 
lost it is possible in some cases for the remaining engine to cross-drive both AMAD gear 
boxes and maintain full electrical and hydraulic services. Electrical power - typically 
115 VAC, three-phase or 270 V DC - is provided by generators which supply the power to 
the electrical power distribution system. Emergency electrical power may also be derived 
from the emergency power unit or the RAT. The AMAD is also used to start the engine using 
a start motor which uses air or electrical power as appropriate. The system is subject to 


extensive monitoring, with temperatures, speed, pressures, status, etc., being signalled to the 
various controlling LRUs. Alternatively, these control and monitoring functions may be 
embedded in an integrated VMS. 

10.5.4 Electrical Power Generation and Distribution 

The electrical power generation and distribution system provides AC and DC power at 
avionic equipment terminals with voltage and frequency characteristics defined by power 
system standards. The power system is generally designed to eliminate single-point failures 
and to maintain electrical supplies under defined failure conditions. The distribution system 
includes power circuit protection devices to isolate equipment faults and to protect the 
aircraft wiring. 

Under certain primary power source conditions, e.g. loss of engine or generator, it may be 
necessary to remove loads selectively from the bus bars, a technique known as load 
shedding. The distribution system design enables groups of electrical loads to be discon- 
nected to allow a progressive degradation in mission performance, while retaining critical 
avionic functions such as displays and controls, navigation and communications. 

Electrical power generation and distribution include the provision of electrical power 
derived from main engines and secondary or emergency power sources to meet the predicted 
electrical loads throughout the mission, and the provision of a suitable distribution and 
protection system. The system design also includes the provision of bonding and earthing, 
electrical hazard protection and compatibility with external power sources. Generator control 
circuits, which contain a means of sensing over- and undervoltage as well as under- and 
overfrequency, may be self-contained or may be incorporated into utility management. This 
control is safety critical because malfunction may lead to disconnection of the generators and 
subsequent loss of all electrical power. A key feature of utility management is the detection 
and announcement of system failures and the determination of corrective action which may 
include automatic load shedding. 

Most military aircraft power systems are conventional and use integrated drive generators 
(IDGs) to produce two or more channels of 115 VAC, three-phase, 400 Hz electrical power. 
Two slightly unusual examples of military aircraft power are described below: those for 
electrical power systems of the F/A-18E/F and the F-22. 

10.5.4. 1 F/A- 18E/F Super Hornet 

The F/A-18E/F uses a variable- speed constant-frequency (VSCF) cycloconverter power 
switching and commutation technique to provide 115 VAC and 28 VDC power to several 
aircraft buses. This type of system is widely used in the United States, being the primary 
means of electrical power on the F/A-18C/D, F-117A, U-2 and V-22, and has proved to be 
very reliable in service. This system is shown in Figure 10.16. F/A-22 Raptor 

The F/A-22 uses a 270 V DC electrical system that has been favoured by the US Air Force 
for the last 10 years or so. This system will also be used on the F-35 joint strike aircraft 
(JSF). Primary electrical power is generated at 270 V DC, but, as many legacy subsystems 
and components require 115 VAC and 28 VDC, converters are added to convert power to 







Main AC 





- 1260W 

■ 1260W 






~ 340W ~ 340W 



ivivaa ^— «w 

Main AC 







Control DC 

Control DC 





Figure 10.16 VSCF cycloconverter as used on the F-18E/F. 

cater for these loads. The electrical power distribution is provided by eight power 
distribution centres (PDCs): four distributing 115 VAC power and four distributing 270/ 
28 VDC power, as shown in Figure 10.17. 

Other systems associated with the utilities system generation of power are: 

1 . Hydraulic power generation and distribution - provision of hydraulic power derived from 
the main engines and secondary or emergency power sources to meet the predicted 
hydraulic loads throughout the mission, and provision of suitable system components to 
maintain power, provision of pipes, couplings and compatibility with external power 
sources. A key feature of utility management is the detection and announcement of 
system failures and the determination of corrective actions (see Figure 10.20). 

2. Fire detection and suppression - provision of methods for detecting fire or overheat, 
warning the crew and providing a means of suppressing/extinguishing the source of heat 
such as a hot gas leak or naked flame. Fire protection systems are usually independent of 
all other systems. 

10.5.5 Hydraulic Power Generation and Distribution 

The hydraulic system provides power mainly for flight control surface movement and 
landing gear operation. There are hydraulic circuits that are used to open doors and hatches, 
or to deploy sensors by providing motive power. Examples are camera bay doors, radar 
antenna rotation motors, electrooptic turret deployment and retraction and bomb-bay door 
operation. Hydraulic power is generated by engine-driven pumps at flowrates and pressure to 
meet the specification of the flight control system demands. The distribution system - pipes, 



270 VDC 

Main Aircraft 


270VDC Power 
115VAC Power (3 Phase) 
28VDC Power 

PDC - Power Distribution Centre 

270 VDC 

Main Aircraft 


No 1 270VDC BUS 

No 2 270VDC BUS 



| (No1 115VACBUS) 

115VACPDCs(2) j 



(No 1 270VDC BUS) 


"I r- 

I I 

( No 1 28VDC Bus ) | I ( No 2 28VDC Bus ) 

I I 


| (NO 2 115V AC BUS) 

I I 

115VACPDCs(2) J 

(No 2 270 VDC BUS) 

270/28VDC PDCs (2) ■ 270/28VDC PDCs (2) 

Figure 10.17 F-22 power generation and distribution system. 

reservoirs, accumulators and valves - ensures that the flight-critical services are isolated 
from non-flight-critical services to allow isolation in emergency cases. 

10.5.6 Fuel Systems 

The primary purpose of the fuel system is to provide a source of energy for the propulsion 
system. It does, however, have two subsidiary functions: one is to act as a form of movable 
ballast to maintain the aircraft eg within specified limits, and the other is to act as a heat sink 
for cooling system loads - heat is dumped from gearbox and avionic loads into a fuel-cooled 
heat exchanger. Key functional areas include: 

1 . Fuel system architecture - provision of a suitable shape and location of the internal and 
external fuel tanks to accommodate the fuel mass required to meet mission requirements, 
and provision of a gauging and transfer system to meet the tank layout. Requirements for 
the fuel system form the basis of the functional requirement and the determination of the 
distribution of interfaces and functions to utility management, implemented in software. 
An example system is shown in Figure 10.18 in the normal engine feed mode. Such a 
system has several different modes of operation including: 

• Engine feed from the forward and rear fuel groups to the left and right engine 





Figure 10.18 Typical fighter aircraft fuel system - simplified. 

• Fuel transfer from wing, underwing or conformal tanks in a preordained sequence; 

• Refuel/defuel; 

• In-flight refuelling; 

Fuel dump. 

2. Fuel gauging and level sensing provides a means of determining the quantity and 
location of fuel in the aircraft. Modelling tank shapes allows designers to define the 
location of gauging, level and density sensors to obtain accurate measurement of usable 
fuel under all flight attitude and manoeuvre conditions. Interface definition with the 
vehicle management system and pilot requirements for control and indications lead to a 
functional requirement and a human-machine interface definition. The number of fuel 
probes and the tank shapes dictate the accuracy of the fuel quantity measurement 
system. The variation in fuel properties and fuel density have to be carefully measured 
and taken into account. Depending upon the location and type of fuel, the density can 
vary by almost as much as 10% between the tropics versus the Arctic region. Fuel 
properties vary greatly at different locations in the world which can impact civil 
operations but have a huge impact upon military aircraft operating in remote locations 
in third- world countries. 

3. Fuel feed/venting - provision of tank interconnections, couplings, bonding and earthing, 
refuel and defuel, and provision of suitable venting for jettison, expansion, etc. 

4. Fuel management - provision of an appropriate mechanism for transferring fuel from 
tanks to engine to meet normal fuel demand, eg management and emergency/leak/battle 
damage conditions. Detection of component failures, leak detection and automatic 
reconfiguration and provision of system status to the cockpit displays. In unstable aircraft 
it may be necessary to ensure that the aircraft centre of gravity is maintained within strict 


limits so that the pilot is able to make manoeuvre demands to meet his operational 
demands without causing damage to the aircraft - the flight control system will ensure 
this. To do so safely, the flight control system requires information about the position of 
aircraft components. There are consumables on the aircraft that make a significant 
contribution to the eg, the main items being fuel and weapons and stores. As fuel is used, 
it is transferred from tank to tank and, although the range of variation in fuel eg is limited 
by the vehicle management system, its variation is sufficient to affect safe aircraft carefree 
handling. For this reason, in some aircraft types, the location of fuel is declared to the 
flight control system together with the initial store load and the use of stores during a 
mission. The flight control system is able to modify its control laws and limitations to 
protect the aircraft from damage. 

The simplified military aircraft fuel system as shown in Figure 10.18 comprises a number 
of fuel tanks: 

1. The fuselage tanks are usually separated into forward and rear fuel groups. 

2. Wing tanks are on the left and right. Total internal: 3986 1/1053 US gallon. 

3. Underwing tanks, to extend operational range, are on the left and right. Many aircraft can 
also carry another external tank on the fuselage centre-line. Total external: 56801/1500 
US gallons. 

4. Conformal tanks, for ferry purposes or to extend operational range, are on the left and 
right. Total conformal: 1700 1/450 + US gallons. 

The figures given relate to the F-16E/F block 60 configuration. 

The components that define the system configuration and fluid logic are transfer and refuel 
valves operating under the control of the fuel management function. Even on this simplified 
diagram a total of 20 transfer/refuel valves are shown. Fuel transfer is usually accomplished 
with the help of pressurised air in the case of the external tanks. Electrically driven fuel 
pumps assist the transfer of fuel from the wing tanks to the forward and rear fuel groups, 
while booster pumps assist the flow of fuel from these groups to the engine. On the relatively 
simple EAP demonstrator, which had no external tanks fitted and no in-flight refuelling 
capability, there were a total of 206 signals and 36 power drives associated with the fuel 
management and quantity gauging system. 

10.5. 7 Air Systems 

A considerable number of utilities functions are powered by bleed air from the main engines 
or the auxiliary power unit (APU). The functions powered from bled air taken from the 
engine compressors include the following: 

1. Cooling system - provision of an air refrigeration and decontamination system to 
condition air taken at high temperature and pressure from the engine connection. 

2. Cockpit/cabin cooling air distribution - provision of air at an appropriate temperature and 
humidity into the crew compartment to maintain crew comfort with minimum cabin 
noise. Controls are provided to enable adjustment of flow and temperature. 

3. Equipment cooling air distribution - provision of air at an appropriate temperature and 
humidity into the equipment bays/compartments to maintain continued equipment 
operation. The environmental control system (ECS) provides clean air at appropriate 



Heat Ex 

| ...1...>J L.-t.. 

?( \ © Second? 

^^ L^...t... HeatE 

■ I Avionics I 

— MS> 

WWfPPI I Mixing 
I ' -' I • Valve 




Shut-Off Valve 


Pressure reducing 
Shut-Off Valve 





Figure 10.19 Simplified air/environmental control system. 

temperature and humidity levels to ventilate equipment bays and to cool avionic 
equipment mounted in the bays. The supply of cooling air may be provided as ventilation 
by circulation in the bay, by washing air up the side walls of equipment or by providing 
air to a plenum chamber in the equipment mounting tray. In certain cases, air cooling is 
not sufficient and a liquid cooled (cryogenic) mechanism may be provided for such high- 
dissipation devices as radar transmitters or electrooptic sensor packages. 

4. De/anti-icing - provision of suitable methods for de-icing or anti-icing of the leading 
edges of appropriate surfaces - wings, empennage, intakes by hot air internal bleed or 
rubber boot inflation. 

5. Canopy demist - provision of hot air to demist internal canopy surfaces. 

A typical simplified layout for a military fighter aircraft is shown in Figure 10.19. This 
diagram depicts a bleed air supply taken from the engine compressor and fed through the 
ECS systems; the system effectively comprises two independent channels, each extracting 
air from the respective engine. The main attributes are: 

• Pressure-reducing shut-off valves (PRSOVs) to moderate the pressure of the engine bleed 
air to a constant pressure irrespective of throttle setting; 

• Primary and secondary heat exchangers to reject heat overboard; 

• Shut-off valves (SOVs) to provide an independent shut-off capability in addition to 

• Closed-loop, variable-mixing and turbine bypass valves to control the flow of warm and 
cold airstreams before entering the mixing plenum chamber; 

• Distribution system to allocate air of the necessary temperature to the cockpit and 
avionics bays; 


• A number or temperature and pressure monitoring points (as well as the very basic 
monitoring set shown in Figure 10.19, there will be a significant number of sensors to 
monitor system performance and status and to provide display data and warnings for 
display to the pilot when required). 

The foregoing system example is a greatly simplified portrayal. As well as this system there 
will be a number of other dedicated cooling subsystems - some of which will be using liquid 
cooling for high-power devices - that will also be interfacing with and rejecting heat into the 
system. On a relatively simple aircraft like the EAP demonstrator, which had a minimal 
mission fit, the number of signals associated with the ECS was ~59 with 14 power drives. 

10.5.8 Electrical Utilisation Systems 

The miscellaneous electrical utilisation systems listed below require simple logical control 
functions and may be independently controlled. However, in some cases there are depen- 
dencies such as weight on wheel interlocks, or automation such as automatic dimming of 
lighting. Typical electrical utilisation systems include: 

1 . Internal lighting - designed to provide a balanced lighting solution across the cockpit in a 
variety of lighting conditions with the ability to dim panels, instruments and equipment 
lighting. The system provides all power sources for all types of flight deck, cockpit and 
cabin lighting including integral lighting, floods, wander lamps, anti-flash and main- 
tenance bay lighting. 

2. External lighting - provision of lighting solutions for external lights, e.g. navigation, anti- 
collision beacons, high-intensity strobes, formation lights, refuel probe lights and land/ 
taxi lamps. 

3. Probe heating - provision of heating for external air data, temperature and icing probes. 

4. Ice detection and protection - provision of methods for detecting icing conditions and ice 
accretion and icing protection/de-icing systems. These systems may be used to provide a 
warning of icing conditions, or to initiate de-icing methods automatically. 

5. Windscreen de-icing - provision of anti/de-icing methods for canopy, windscreens or 
direct vision (DV) windows. 

10.5.9 Prognostics and Health Management 

The majority of aircraft in service today include a mechanism for capturing and recording 
failures that are detected during a flight. Most systems contain some form of in-built testing 
to monitor their continued satisfactory operation. Built-in test (BIT) is initiated automati- 
cally at system power-up (PBIT), or may be initiated by the ground crew or even by the 
aircrew (IB IT), or may take place continuously throughout system operation (CBIT). All of 
these forms of test will produce a result when a failure is detected, requiring some form of 
corrective action at the next convenient opportunity. 

This discrete form of recording failures has disadvantages: 

1 . The next landing site may not be the best place to perform maintenance, especially if the 
aircraft is on deployment or in transit through a site where spares and skilled mechanics 
may not be available. 



VMS data / Prognostics System 

VMS bus 

Avionics bus 

Avionics date 

Mission data 
Mission System bus 

Data acquisition 







Data from direct 

Figure 10.20 Prognostics system block diagram. 

2. It is better to be warned of an impending failure so that maintenance can be planned. 

3. With multiple-redundant systems it may be possible, even desirable, to despatch with a 
known failure, particularly in times of conflict. 

For these reasons, prognostics systems are developing that enable decisions to be made on 
the basis of known deteriorating performance and a statistical prediction of time remaining 
before action is required. All systems will be monitored for detected faults resulting from 
BIT, and system parameters will be measured to determine changes in performance (e.g. 
times, rates of changes, pressures, stress, etc.) and to measure structural limitations and 

Connections are made to systems to measure such signals as flow, pressure, rates of 
change, numbers of operations and duration of operation. This enables decisions to be made 
by comparing measured values with datum values. An informed condition report can be 
generated by the system which will inform the ground crew when a maintenance action 
should be performed. Connection of the prognostics system to data link allows information 
to be transmitted to the destination airfield so that spares and skilled ground crew are 
available to repair the aircraft. An example system block diagram is shown in Figure 10.20: 

• System architecture - provision of a computing architecture to meet the requirements for 
gathering of information, processing of information and interfaces with the customer's 
support environment; 

• System interfaces - provision of interface structures or remote interfacing units to match the 
vehicle system components requirements: I/O type, impedance, slewing rates, load, etc.; 

• Software - provision of functional requirements, algorithms and software. 

10.6 Design Considerations 

10.6.1 General 

Given the safety-critical and safety-involved nature of the utility systems, great care must be 
taken in the design of the utilities control system. This starts with developing a clear 



The set of vehicle 

requirements (R) 


R system 1 


C R system 2 J |> 
f R system 3 J |> 

( R system n j j> 


Requirement for VMS R VMS 

Equipment specification 
Operating system 
Functional partitioning 
Software requirement 
Interface control 
Hazard analysis 

• R system 

Project constraints e.g. Language, Standards, Data 
Bus type, design process 

Figure 10.21 Vehicle systems requirements. 

understanding of the requirements of the individual utility systems. This is not a trivial task. 
The total utility vehicle systems requirement is the sum of a large number of individual 
requirements of differing degrees of complexity. Figure 10.21 shows how the requirements 
flow from the individual systems, are then consolidated into a single vehicle system 
requirement and than used to influence the architecture and design principles. Functions 
are then isolated and allocated to individual processing units on the basis of integrity, 
separation, processor loading, etc. This leads on to development, testing and qualification of 
the total system. 

10.6.2 Processor and Memory 

An analysis of the functional requirements of the utility systems and the partitioning of those 
functions into individual processors provides an indication of the throughput requirement for 
the processor and the memory capacity. The designer must ensure that there is sufficient 
growth potential in the selection to account for the following eventualities: 

• Growth in the requirement during the early stages of detailed definition and design; 

• Changes resulting from errors in the design and encountered during testing; 

• Changes resulting from flight testing; 

• Growth potential to allow for system capability enhancement during use. 

The processor and memory configuration is shown in Figure 10.22. 

10.6.3 Interfacing 

The early implementation of utility systems management integrated the input and output 
interfacing with the processor and memory in a small number of processing units of a 
standard case size. The reason for this was to minimise the number of processing units and 
the number of suppliers in order to achieve a level of standardisation or commonality in the 
architecture. There is now a trend with modern vehicle management systems to segregate 




Operating System 

Built In Test 



Built In test 

Bus Control 

Processor throughput and memory sized to allow 100% growth in service 

Figure 10.22 Processor and memory configuration. 

the interfacing circuitry into remote input/output (RIO) units that can be distributed to the 
extremities of the aircraft. This reduces the majority of interface wiring to the locality of 
the RIO. The information is encoded in the RIO into a data format that is compatible with the 
processing unit. 

The interfacing circuits must take full account of the electrical characteristics of the 
component. The RIO should ideally perform a 'degree of goodness' check, confirming that 
the input signal is within reasonable bounds and ensuring that it is a valid value; in other 
words, that the data the sensor is providing lie within a window of upper and lower values 
that gives confidence that they are valid. 

10.6.4 Software 

The choice of language will probably be driven by the need to have a single common 
language for a project, and may be driven by a customer need to have a language common 
with other assets in their inventory. 

Whatever the language, there is a need for a robust design process that is consistent with 
the level of safety defined for the individual utility systems. This is usually safety critical or 
safety involved. 

10.6.5 Obsolescence 

The long life of military products - there are many examples of aircraft still in active service 
40 or more years after their initial design - means that component obsolescence is a certainty 
rather than a risk. An obsolescence plan is required to detail all components that are likely to 
become difficult to replace after a period of time. There is sufficient experience available to 
determine those components most likely to become obsolete as a result of technological 
advances, and there are existing mitigation plans that can be used as a model. However, 
obsolescence can also be driven by health, safety and environmental legislation, where 
materials once considered safe to use are no longer acceptable because they are dangerous to 
health in use, or their disposal poses a health or environmental hazard. This situation is less 
predictable and requires a continuous monitoring of legislation. 



Controls, Sof tkey <& 
voice inputs 


Display, voice and 
audio tone outputs 

Displays & controls 

Mission Systems 

Navigation System 

Vehicle Management System 

Fuel System 


Flight Controls 

Signals and 

Figure 10.23 Human-Machine Interface. 


Prognostics & Health 

Other systems 

V Commands 


The generally lower performance - in terms of throughput, clock rate, memory and 
executed instructions - of many of the processing devices used within VMS applications 
may offer alleviation to the obsolescence problem rather than protection from it. The 
problem of COTS is a major problem in the high-performance avionics systems, as already 
discussed in Chapter 2. 

10.6.6 Human-Machine Interface 

The human-machine interface in the modern aircraft cockpit is largely based on multi- 
function displays combined with multifunction or 'soft' keys, as described in Chapter 11. 
There will be some independent, hardwired controls, but these will be minimised to those 
essential instinctive controls. The main interface between the individual utility systems and 
the pilot will be provided by the vehicle management system. The VMS will interpret 
messages from the cockpit and use them in the control functions to place appropriate 
demands on a system to modify a control function or directly to command an action. Data 
from the systems are used in control functions and transferred to the cockpit displays or to 
other avionic systems via the appropriate data buses. In this way the pilot can make demands 
and observe the behaviour of the utility systems, making maximum use of the carefully 
designed cockpit layout. 

Two items of aircraft equipment straddle the boundary between utility systems and 
avionics. These items are the throttle lever handles and the stick top, associated with 
propulsion and flight controls respectively. Both items have been furnished with switches and 
controls to perform a number of avionic functions, over and above their basic use as a 
comfortable grip for the throttle and flight control demand levers. This allows the pilot to 
perform a number of control functions without removing hands from the two demand levers. 
This has become know as hands-on throttle and stick, or HOTAS. The positioning and 
actuation loads of the switches installed in the handles has to be designed with great care to 
meet ergonomic requires for instinctive finger movements. Figure 10.24 shows an example 
HOTAS arrangement. 





Throttle Grips 

Chaff/Flare Dispenser 

Radar Elevation 


Speed Brake 




Control Stick Grip 


Raid Mode 


Air to Ground 



Gun/Missiles trigger 

Wheel Steer 

Auto- Pilot/Nose Wheel 
Steer Disengage/g-li miter 

Recce Event Mark 

Sensor Conirol 

Air to Air 
Weapon Select 

Figure 10.24 HOTAS concept in the F/A-18 Hornet. 


Jukes, M. (2003) Aircraft Display Systems, Professional Engineering Publishing. 

Moir, I. and Seabridge, A.G. (1986) Utility systems management. Royal Aeronautical Society Journal - 

Aerospace, 13(7), September. 
Moir, I. and Seabridge, A.G. (2001) Aircraft Systems, Professional Engineering Publishing. 
Moir, I. and Seabridge, A.G. (2002) Civil Avionics, Professional Engineering Publishing. 
Pallett, E.H.J. (1992) Aircraft Instruments and Integrated Systems, Longmans Group. 
Yeoell, L. and Kneebone, R. (2003) On-board oxygen generation systems (OBOGS) For in-service 

military aircraft - the benefits and challenges of retrofitting. Aero India 2003 - International Seminar 

on Aerospace Technologies: Developments and Strategies - Flight Testing and Man Machine 

Interface, 8 February 2003. 

Further Reading 

Bryson Jr, R.E. (1994) Control of Spacecraft and Aircraft, Princeton University Press. 
Conway, H.G. (1957) Landing Gear Design, Chapman and Hall. 
Currey, N.S. (1984) Landing Gear Design Handbook, Lockheed Martin. 
Hunt, T. and Vaughan, N. (1996) Hydraulic Handbook, 9th edn, Elsevier. 
Lloyd, E. and Tye, W. (1982) Systematic Safety, Taylor Young. 
Pallett, E. H. J. (1987) Aircraft Electrical Systems, Longmans Group. 

Pratt, R. (2000) Flight Control Systems: Practical Issues in Design and Implementation, IEE Publishing. 
Principles of Avionics Data Buses (1995) Avionics Communications Inc. 

Raymond, E.T. and Chenoweth, C.C. (1993) Aircraft Flight Control Actuation System Design, Society of 
Automotive Engineers. 

1 1 Displays 

11.1 Introduction 

The history of aircraft displays can broadly be divided into three technology eras, the 
mechanical era, the electro-mechanical (EM) era and the electro-optical (EO) era. 

Although the design boundaries are clear, the time boundaries are vague. The catalyst for 
the electro-optical era was the dramatic increase in performance and capability of digital 
electronics in the late 1960s that led to an impetus to change the means to display information 
on the flight deck. Cathode Ray Tube (CRT) technology was the first multifunction display 
medium, latterly superseded by Active Matrix Liquid Crystal Display (AMLCD) technology. 
Both provide a more flexible means for the display of information than had hitherto been 
possible. Multifunction displays (MFDs) can show many formats on the same display surface 
and portray the same piece of information in a variety of different ways. 

This chapter provides examples of 'glass' electro-optical military fighter cockpits, or crew 
stations as they are now known, tracing the evolution from the electro-mechanical gyro 
gunsight and the radar 'scope' to current electro-optical head-down, head-up and Helmet- 
Mounted Displays for advanced tactical and global situational awareness in today's digital 

The Head-Up Display (HUD) has been applied predominantly to military fast jet fighter 
aircraft. The optical principles of the Head-Up Display will be described, along with its 
principles of operation for air-to-air and air-to-ground weapons aiming. 

The rotorcraft community were the first to apply Helmet-Mounted Display (HMD) 
technology. The operating principles and the role of the Helmet-Mounted Display will be 
described. The Head-Up Display is placed on the head of pilots, greatly enhancing their 
ability to acquire and designate targets and to release weapons off-boresight. Indeed, as 
currently planned for the joint strike fighter, it is entirely likely that the Helmet-Mounted 
Display will become the primary flight instrument in future fast jet fighters, and the head-up 
Display will no longer be fitted. 

This chapter will also discuss the principles of operation and device features of the 
shadow-mask CRT and the active matrix LCD head-down displays presenting primary flight, 

Military Avionics Systems Ian Moir and Allan G. Seabridge 
© 2006 John Wiley & Sons, Ltd. ISBN: 0-470-01632-9 


navigation, topographical and tactical map information, systems and sensor (radar and FLIR) 
images and weapons information to the crew. 

Emerging and potential future display technologies of rear projection reflective LCD, 
transmissive LCD (liquid crystal on silicon - LCoS) and digital micromirror devices (DMD) 
will be reviewed, as well as the means to apply these technologies to large-area megapixel 
head-down displays. 

Finally, the crew- station ambient lighting conditions in day, dusk/dawn and night will be 
described, together with the principal accepted industry visual performance metrics and 
optical test methods to achieve display viewability. The requirements and means to achieve 
compatibility with night-vision imaging system (NVIS) devices will be discussed. 

A more in-depth discussion of the topic areas in this chapter can be found in Jukes (2004). 

1 1 .2 Crew Station 

Undeniably, the first use of electro-optical (EO) devices in the cockpit was during World War 
II with the display of airborne intercept radar contacts on the Cathode Ray Tube. Soon the 
CRT was used in the gyro gunsight to produce a collimated (focused at infinity) aiming 
reticle. Very quickly the CRT gyro gunsight developed into a sophisticated projection device 
through which the operator could correlate the position or vector of the aircraft or weapon 
with the outside world. The Head-Up Display (HUD) had been invented. 

As CRT technology improved, it became possible to augment the role of the radar 'scope' 
also to provide flight information. The multifunction head-down display (HDD) had been 
invented. However, CRT brightness and contrast technology limitations meant that early 
CRT multifunction displays were monochrome (green), and information presentation was 
limited to character/symbolic images. In these early 'glass' cockpits the moving map display 
was produced by optical rear projection of a 35 mm filmstrip topographical map. However, it 
was not long before further advances in colour CRT technology made it possible to present a 
full colour topographical map image in the severe lighting conditions and environment of a 
military fighter cockpit. 

1 1.Z 1 Hawker Siddley (BAe) Harrier GR.Mkl and GR.Mk3 (RAF) 
and AV-8A (USMC) 

The first UK aircraft to enter service with a 'designed-in' Head-Up Display was the vertical 
take-off Hawker Siddley (BAe) Harrier. These aircraft equipped four operational squadrons, 
one in the United Kingdom and four in Germany. The US Marine Corps (USMC) took 
delivery of 110 aircraft, designated the AV-8A. 

The cockpit of the Harrier GR.Mk3 is shown in Figure 11.1. The flight instruments are 
conventional electro-mechanical counter pointer type. In the centre of the instrument panel is 
the moving map. This large and complex instrument projects the image of a filmstrip 
topographical map on to a rear projection screen. The film transport mechanism is driven by 
aircraft inertial reference coordinates and heading to indicate aircraft current position and 
course/track with reference to ground topography. 

Above the moving map is the Head-Up Display. The HUD was small by today's standards. 
It was a refractive design with a 4 in exit lens and provided a modest instantaneous field of 



Figure 11.1 Harrier GR.Mk3 cockpit (RAeS). 

view of about 16°. The HUD itself contained only the CRT and the collimation optics. The 
CRT high- voltage supplies and the CRT beam deflection electronics were remote from the 
HUD to minimize space and weight in the cockpit. 

The HUD operated in cursive (stroke) mode to provide a daylight-viewable symbolic 
image. Display modes included: 

• Navigation; 

• Approach and landing; 

• Precise local fix (IN update); 

• Air-to-air attack (guns, rockets and missiles); 

• Air-to-ground attack (freefall and retarded bombs, CCIP manual and CCRP automatic 

1 1.2.2 McDonnell Douglas F/A- 18 Hornet 

The F/A- 18 Hornet is generally considered the true beginning of the electro-optical (EO) era 
and is pivotal in cockpit display design. The cockpit of the night-attack F/A-18C is shown in 
Figure 11.2 and Plate 1. 



Figure 11.2 McDonnell Douglas F/A-18C cockpit. 

The HUD incorporates a raster mode for night-time use, presenting a collimated outside- 
world image from the forward looking infrared (FLIR) pod overlaid with conventional stroke 

The left and right MFDs present limited colour formats using time-sequential liquid crystal 
shutter technology. This technology superimposes the additional dimension of red and orange 
symbology overlays onto the high-resolution, high-brightness green raster weapon imagery. 

The projected map display utilises a multipurpose, high-brightness, high-resolution, full- 
colour stroke and raster shadow-mask multipurpose CRT display (MPCD) with a 5 x 5 in 
square format usable screen area. The map image is generated remotely from the display in a 
digital map computer using CDROM technology. 

The cockpit lighting is fully compatible with night-vision goggles and at night the pilot 
may aid his night vision with the use of 'cat's eye' goggles attached to the helmet. 

In the early 1990s, McDonnell Douglas commenced development of the F/A18E & F 
Super Hornet. The centre CRT-based MPCD is replaced with a larger area (6 x 6 in square 
format display) employing AMLCD technology to provide enhanced brightness/contrast and 

The F/A-18 cockpit truly broke new ground, but its introduction represented only the tip of 
a technological iceberg in terms of the challenge for the cockpit designer to show to the crew 
the massive amount of data now made available by digital processing without saturating 
them with data overload. The answer is to present only those data required for the current 
phase of the mission and to configure the display format accordingly. Initially, this 
reconfiguration was performed by the operator who decided what to display and when. 
Unfortunately, excessive operator involvement was found to be counterproductive in terms of 
reducing workload. Today, progressively more sophisticated decision aids predict the crew 
information requirements and configure the display formats accordingly (Garland et aL, 1994). 

As will be seen in subsequent cockpit designs, there is a continuous struggle to reduce the 
bulk of the display device itself while increasing the display surface area and flexibility of 



Figure 11.3 Eurofighter Typhoon cockpit (BAE SYSTEMS). 

the information content. The ultimate aim, possibly to be achieved in the joint strike fighter, 
is to provide the operator with one contiguous, controllable display surface. 

11.2.3 Eurofighter Typhoon 

The Eurofighter consortium was formed in June 1986 by the three countries that developed 
the Tornado, namely the United Kingdom, Germany and Italy, and was shortly joined by 
Spain. The development programme was launched in November 1988 and the first two 
prototype aircraft undertook their maiden flights on 27 March and 6 April 1994. 

The Eurofighter Typhoon cockpit is shown in Figure 11.3 and Plate 2. The main 
instrument panel comprises three colour multifunction head-down displays (MHDDs). In 
the prototype aircraft these displays used shadow-mask CRTs to provide daylight- viewable, 
full-colour, high-brightness, high-resolution images in both cursive (stroke) and hybrid 
(stroke + raster) modes. In production the CRTs have been superseded with high-resolution 
6.25 x 6.25 in square format Active Matrix Liquid Crystal Displays. The MHDDs in- 
corporate 18 multifunction keys around the bottom, left and right edges of the display. 
Each key contains a daylight- viewable LED matrix of two rows of four 7.5 characters plus 

The HUD uses holographic technology to achieve an ultra wide 30° x 25° field of view 
(FoV). The HUD provides stroke (cursive) operation for daytime use plus raster for night- 
time use with outside-world sensor video. The HUD incorporates a sophisticated up-front 
control panel with a 4 x 3 in daylight- viewable LED matrix display. The HUD is the primary 
flight instrument. 

An HMD is planned, configured into two variants. The daytime variant provides 
symbology for the targeting and release of off-boresight weapons. The night-time variant 
adds night-vision goggles (NVGs) to the helmet to provide the pilot with enhanced night 



vision. The NVG image is electrically mixed with the CRT symbology image to provide a 
comprehensive night-time capability. 

To either side of the HUD the left and right glareshield panels provide essential controls 
and warnings. The right-hand panel incorporates the standby attitude display employing 
AMLCD technology. The farthermost part of the right-hand glareshield flips open to reveal a 
set of standby get-u-home instruments in the unlikely event that there is a major power failure. 

The Eurofighter Typhoon provides direct voice input (DVI) command control for non- 
mission-critical functions such as communications equipment. The DVI speech recogniser 
has a vocabulary of about 100 words. The DVI system is trained by the individual user to 
function under all operational conditions including high-g manoeuvres and low-speed passes 
with significant wind buffet (Birch, 2001). 

11.2.4 Lockheed Martin F-22 Raptor 

In April 1991 the USAF announced that it had selected the Lockheed Martin F-22 Raptor to 
meet its advanced tactical fighter (ATF) requirement for a fighter combining low observa- 
bility, supersonic cruise, long range and a very high level of agility. 

The F-22 advanced avionics system provides the pilot with fused situational awareness in 
the battlefield environment. The advanced crew-station layout shown in Figure 11.4 includes 
a number of large-format full-colour AMLCD multifunction head-down displays, numerous 
general- and special -purpose knobs and switches, a Head-Up Display and a Helmet-Mounted 
Display. The displays contain integral video processing and graphics generation (Greeley 
and Schwartz). 

A central single 8 x 8in primary multifunction display is flanked by three 6 x 6 in 
secondary multifunction displays. Two 3 x 4 in up-front displays are arranged either side 

Figure 11.4 F-22 Raptor crew station. 


of the HUD. Below the HUD there is an integrated control panel housing some dedicated 
switches and alphanumeric text read-outs. Bezel option selection buttons (OSBs) surround 
each of the displays and are used for menu navigation, paging and function select. 

The prototype aircraft also included touch- sensitive screens, voice recognition and three- 
dimensional audio, but these technologies were deemed to be immature and were not 
included in the production design. 

The primary multifunction display (PMFD) is used to display the situation display (SD) or 
the attitude director indicator (ADI) display. The SD is a tactical format showing the entire 
track file icons as well as navigation data, with multiple levels of OSB menus for feature 
control. SD symbology is centred on ownship. The ADI is a back-up to the HUD. 

The three secondary multifunction displays (SMFDs) are used to display either of the 

• Attack display (AD); 

• Defence display (DD); 

• Expand display (EXD); 

• Situation display - secondary (SD-S); 

• Stores management display (SMD); 

• Fuel display, engine display; 

• Mission data edit (MDE) display family; 

• Back-up integrated control panel display; 

• Flight test display (FTD); 

• Surface position display (SPD); 

• Electronic checklist (ECL). 

The up-front displays (UFDs) are used to display the communication, navigation and 
identification (CNI) display or the standby flight group (SFG) display. Additionally, the 
integrated caution, advisory and warning (ICAW) data are shown in the centre column of the 
CNI display. The SFG display shows a simplified small version of the ADI and HUD attitude 
data with digital read-outs for altitude, speed and heading. 

The Head-Up Display (HUD) provides a 24° horizontal x 20° vertical binocular wide- 
field-of-view monochrome (green) image. The HUD is the primary flight display (PFD) and 
provides attitude, flight path, navigation and weapons deployment symbology in general 
conformance to MIL-STD-1787. 

The Helmet-Mounted Display is the US standard joint helmet-mounted cueing system 
(JHMCS) with a 20° instantaneous field of view. The HMD is monochrome (green) 
monocular (right eye only), and is used to show similar symbology to the HUD with 
some special symbology for depicting sensor volume limits. Like the HUD, extensive user- 
editable controls are provided to declutter symbology. 

11.2.5 Boeing (McDonnell Douglas/ Hughes) AH-64D Longbow Apache 

The AH-64D Longbow Apache best illustrates the attack helicopter electro-optical 'glass' 
crew station. The Apache is a tandem, two-seat helicopter with advanced crew protection 
systems, avionics and electrooptics plus a weapon control system that includes the nose- 
mounted target acquisition and designation sight/pilot's night- vision sensor (TADS/PNVS). 
The copilot/gunner sits in the front seat, with the pilot in the higher rear cockpit. Both use 



Figure 11.5 AH-64D Longbow Apache Cockpit (RAeS). 

sophisticated sensors and systems for the detection and attack of targets, including the 
integrated helmet and display sight system (IHADSS), which provides a monocular helmet- 
mounted designator sight. The copilot/gunner has the primary responsibility for firing both 
the gun and the missiles, but can be overridden by the pilot in the back seat. 

The AH-64D crew stations, shown in Figure 11.5, each have two 6 x 6 in full-colour, high- 
resolution, active matrix liquid crystal (AMLCD) displays replacing the monochrome CRT 
displays of earlier versions. 

11.3 Head-Up Display 

The Head-Up Display has been proven over many years to be a means of providing flight 
navigation, aircraft data and weapon release parameters. The earliest gunsights provided 
little more functionality than an aiming system for the guns, but, with the introduction of 
increasingly sophisticated computational capability, complex manual air-to-air and air-to- 
ground weapon release became possible with enhanced accuracy. Now the HUD forms part 
of an integrated weapon aiming and release system where automatic modes of release are 
available for the different stores and release conditions. In recent years it has been possible to 
introduce sensor video from forward looking infrared (FLIR) on to the HUD such that 
operation by night can be achieved with much the same capability as daytime operation 
(Quaranta, 2002). 

The HUD is installed in an area of 'prime real estate' in the instrument panel, and 
therefore the HUD tends to be 'designed in' to the aircraft to optimise operational capability 
and performance within cockpit geometry and the available space envelope. The configura- 
tion of a HUD typically comprises the pilot's display unit (PDU) and a remote display 



processor. The PDU is rigidly fixed to the airframe and accurately aligned to the aircraft axes 
by a process known as boresighting or harmonisation. 

11.3.1 HUD Principles 

The optical principles of the HUD are straightforward, although to integrate these principles 
into the restricted space confines of a fast jet fighter crew station results in some ingenious 
light-bending solutions. 

The Head-Up Display injects a virtual image of an object (the symbology and/or sensor 
video) into the pilot's line of sight with the key attribute of being collimated, or focused at 
infinity, so that to the pilot the image appears to be in the same focal plane and be fixed on 
features of the outside-world scene. This attribute allows the aircraft systems to place 
symbology in a manner that is often referred to as 'in contact analogue' with (or conformal 
to) the real world. Conformal symbology is used to identify the flight path vector, the horizon 
and sightlines to targets in the real world. In addition, primary flight data (speed, height, 
heading, etc.) and flight guidance cues allow the pilot to fly the aircraft without having to 
refocus his eyes to look inside the crew station at his head-down instruments during critical 
phases of the mission. 

There are two means to form virtual collimated images (shown in Figure 11.6): 

1. HUDs using the principle of refraction are often known as collimating. The lens system 
generates a collimated image of display symbology reflected to the eye via a flat 
combiner. The instantaneous field of view is limited by the size of the collimating lens 
and the pilot has to move around the eyebox to see the total field of view. 

2. HUDs using the principle of reflection are often known as pupil forming. The lens relay 
system generates an intermediate image of the display symbology which is then 
collimated by reflection from the curved combiner. The instantaneous field of view is 
larger, defined by the size of the curved combiner. 

The curved combiner in the pupil-forming HUD may use the optical principle of diffraction 
provided by a hologram rather than reflection to provide optical power, and so this arrange- 
ment is sometimes also known as a diffractive HUD (or variously DHUD or DOHUD - 
diffractive optics HUD). The spectrally selective feature of a hologram is sometimes also 
used to fabricate the plane combiner/mirror. 

Collimating Optics 
Figure 11.6 HUD optical arrangements. 

Pupil - Forming Relayed Optics 



(less than f) 


(focal length f) 

Figure 11.7 Optical refraction. 

11.3.2 Collimating (Refractive Optics) Head-Up Display 

The optical principle of refraction is shown in Figure 11.7. The image source (object) is 
placed just inside the focal length,/, of the collimating lens, shown in the figure as a simple 
convex lens. The observer sees a magnified virtual image of the object. This is the principle 
of the magnifying glass. If the object is placed in the focal plane of the lens, then the rays of 
light refracted by the lens from each point on the object will be parallel, i.e. collimated. 

To form a practical HUD, as shown in Figure 11.8, the image source must be placed out of 
the line of sight. The collimating lens, together with field-flattening lenses, focuses the CRT 
image at infinity and the image is introduced into the pilot's line of sight by a semi-reflective 
plane mirror, known as the combiner, so called because the pilot sees the HUD image 
combined with the normal forward view of the outside world. A second mirror or prism is 
usually introduced between the image source and the collimating lens to produce a compact 
design for convenience of installation and, more importantly, to ensure there is no 
obstruction to the pilot should the need arise to eject from the aircraft. 

In practice the simple convex lens comprises a number of elements to minimise 
aberrations together with a number of elements in the image plane to 'flatten' the field. 

The HUD CRT generally has a faceplate with an active area typically 40-50 mm in 

Combiner / 

Pilots Eye 


Figure 11.8 HUD using refractive optics. 





-> "Porthole" ' Combiner 

Focal Plane Fold Mirror 
Figure 11.9 Refractive HUD instantaneous field of view. 

The combiner is mostly transmissive (typically better than 90%) to the wide-band 
day light- visible spectrum so as not to degrade the pilot's view of the outside world in 
low-light conditions. The CRT phosphor is chosen to have a narrow waveband emission in 
the green part of the visual spectrum. The combiner is optically tuned to reflect the CRT 
spectral wavelength with high efficiency (Jukes, 2004 - Chapter 6). 

11.3.3 Field of View 

The total field of view (TFoV) of any HUD is the total angle subtended by the display 
symbology seen with head movement from any location. 

The instantaneous field of view (IFoV) of a refractive HUD is the display field of view 
seen from one head position using one eye and can be simply derived from the size of the 
collimator exit lens as seen by the observer, reflected in the combiner. A typical arrangement 
shown in Figure 1 1 .9 for an HUD with a 150 mm (6 in) exit lens viewed 450 mm (1 8 in) from 
the combiner yields an instantaneous field of view (IFoV) of about 15°. 

If the pilot's head moves forwards towards the combiner, the instantaneous field of view 
will increase. Similarly, if the pilot's head moves laterally, angles further off- axis will be 
seen. In practice, of course, the pilot sees two portholes from one head position, one with 
each eye. This describes the binocular field of view (BFoV). 

It is possible to increase the instantaneous field of view in the vertical axis by adding a 
second combiner parallel to and vertically above the first. This has the effect of allowing the 
observer to view the CRT image through a second porthole vertically above and super- 
imposed over the first. This arrangement is shown in Figure 11.10. 

11.3.4 Collimating (Refractive) HUD - Examples British Aerospace Harrier GR. Mk 1 and GR. Mk3 

The Head-Up Display of the Harrier GR.Mkl, shown in Figure 11.11, and contained only 
those optical and mechanical supporting structures described above together with the CRT 
and its electron beam deflection yoke and focus magnet. All other electronics was contained 
remotely. This architecture minimized the size and weight of the pilot's display unit (PDU) 
to facilitate its installation into the cockpit. 




"Porthole" Dual / 



/ Ejection Line 

Focal Plane / Fold / 

Figure 11.10 Dual combiner (including FOV). 

The PDU had a 4 in collimating lens and provided an IFoV of about 16° and a TFoV of 
25°. The single combiner was servo-ed to travel in the fore-aft direction to maintain the 
aiming vector within the pilot's normal field of view under all normal manoeuvres. 

1 McDonnell Douglas/British Aerospace Night-Attack Harrier II 
(GR-7 and AV-8B) 

The night- attack Harrier II entered service in 1989 and incorporates a high-performance 
wide-field-of-view refractive HUD capable of day/night operation. The HUD shown in 



Solar Cell 

• Cathode Ray Tube 

Figure 11.11 Jaguar/Harrier HUD installation (Smiths). 



Figure 11.12 AV8B/Harrier GR7 night-attack HUD (with FOV) (Smiths). 

Figure 11.12 incorporates the CRT, the deflection amplifier, the high- voltage power supplies 
and all the services to make the HUD a stand-alone line replaceable unit. 

The HUD employs a dual combiner. The collimating exit lens is truncated in the fore and 
aft axes to allow the optical axis of an effectively larger-diameter exit lens to be installed 
nearer the pilot without encroaching upon the ejection envelope. In the Harrier II, the 
binocular IFoV is 20° azimuth x 15° elevation. The TFoV is 22°. 

The HUD presents daytime images in cursive (stroke) mode and night-time images from 
the FLIR sensor as a raster video image with stroke symbology overlay during the raster field 
retrace period. 

11.3.5 Pupil-forming (Reflective/ Dif tractive) Head-up Displays 

It is highly desirable to offer a larger field of view than that obtainable from a collimating 
HUD to support more aggressive manoeuvring and to give pilots the sensation of flying 
under visual (VMC) conditions at night using a projected FLIR image, enabling them to use 
familiar visual cues to judge speed and terrain clearances and to navigate and identify 

Figure 11.13 illustrates the significant operational improvement that would be achieved if 
the HUD IFoV were to be increased from the 20° x 15° practical limit of a refractive 
collimating HUD to an IFoV of 30° x 20°. 

Using the optical principle of reflection (shown in Figure 11.14), it is possible to place the 
collimator closer to the pilot and therefore achieve a larger IFoV without infringing the 
ejection line. 

If an object is placed within the focal length of a concave mirror, the rays of light are 
reflected by the mirror towards the observer, who sees a magnified virtual image of the 



30 degrees 

6 degrees 

3 degrees 

Collimating HUD 

Pupil Forming HUD 

Figure 11.13 Operational improvement of increased IFoV (BAE SYSTEMS). 

object. This is the principle of the shaving mirror. By placing the object in the focal plane of 
the mirror, the rays of light from each point on the object will be reflected by the mirror to 
emerge in parallel, i.e. collimated. 

As with refraction, it is necessary to design an optical path that allows the CRT image to 
be introduced into the pilot's outside- world line of sight without interfering with it. This is 
more complex than with the refractive arrangement. Some of the possible optical config- 
urations are shown in Figure 11.15 (Fisher). On the right is a class of diffraction optics 
designated 'off-axis', which represents the most elegant solution but is the most optically 
complex. The curvature of the combiner provides the principal collimating function but is 
too great to allow the use of a planar doublet to sandwich the reflective coating. The 


less than f 

A ► 

Image Source 

Concave Mirror 
(focal length f) 


Figure 11.14 Optical reflection. 









Single element 
off-axis design 

Figure 11.15 Some possible reflective optic HUD configurations. 

necessary protection has to be provided by two pieces of curved glass. This reduces the 
thickness and weight of the element but is more complex to manufacture than a plane 
combiner, and care must be taken to ensure that the apparent distortion of the real world seen 
through it is minimal. The large off-axis angle causes significant optical aberrations, which 
must be corrected by introducing compensating aberrations in the reflective combiner itself 
and in a complex relay lens. 

By comparison with the off-axis systems, the other class, termed quasi-axial, have a much- 
reduced critical angle of incidence for reflection. Aberrations are minimal and do not require 
compensating aberrations in the reflective element itself: the limited corrections necessary 
can be implemented in a simple relay lens. The Z-HUD (so called because the optical path 
makes the letter Z) has the further advantage that the reflecting function is separated from the 
combining function. While this is obviously attractive, it probably places the collimating 
mirror over the ejection line in a real installation. 

It is theoretically possible to realise these optical systems using conventional refractive/ 
reflective optical elements, but the optical efficiencies would be poor and so in practice they 
have been fabricated using holographic diffractive techniques (Jukes, 2004 - Chapter 6). 

11.3.6 Pupil-forming (Reflective/ Diffractive) HUD - Examples 

1 1.3.6. 1 F16 LANTIRN HUD - Multibounce Quasi-axial Configuration 

The first application of the multibounce quasi-axial pupil-forming diffractive HUD config- 
uration was in the F-16 as part of the USAF low-altitude navigation and targeting infrared for 
night (LANTIRN) system (Hussey, 1981). 

The optical arrangement is shown in Figure 11.16. Obviously, all glass elements above 
the glareshield must be transparent to allow the pilot an uninterrupted view of the real world. 
They must also sometimes be transparent to CRT light and sometimes reflect it strongly 
to allow the HUD image to reach the pilot. This apparent paradox is resolved by the use 
of highly angularly selective holographic optical coatings that will only reflect a narrow 
band of green wavelengths emitted by the CRT phosphor when incident at a particular 
critical angle. The effect of this is to make the real world seem to the pilot to have a slightly 
pinkish tinge. 

The complete HUD is shown in Figure 11.17 and Plate 3, together with its installation in 
the F-16 LANTIRN crew station. 




Figure 11.16 F-16 optical configuration (BAE SYSTEMS). 

The HUD provides an instantaneous field of view (IFoV) of 30° azimuth x 18° elevation 
and operates in both raster and stroke modes to provide day/night operation. It is fully 
compatible with night- vision goggles. Eurofighter Typhoon HUD - Single-element Off-axis Configuration 

The much more aesthetically elegant solution is the single-element off-axis configuration 
shown in Figure 11.18. This configuration achieves a wide-field-of-view HUD with no 

Figure 11.17 F-16 LANTIRN HUD (BAE SYSTEMS). 




Relay Optics 
Figure 11.18 Eurofighter pupil-forming optical configuration (BAE SYSTEMS). 

incursion into the ejection envelope and with no upper mirror to obstruct the upper field of 
view. It comprises a single optical element between the pilot and the outside world; the semi- 
transmissive curved collimating mirror/combiner. 

The elegant simplicity of this configuration belies its optical complexity, which arises 
because, by its very nature, the intermediate image subtends a significant 'off-axis' angle to 
the collimating mirror. This means significant optical correction needs to be applied to 
correct for distortions. 

The collimator must emulate a complex aspheric surface in order to ensure all rays of light 
from the reflected image emerge in parallel (i.e. collimated). It is only possible to fabricate 
this element using holographic techniques in which the hologram itself is computer generated. 

The relay lens is complex also. It contains several aspheric elements to provide compen- 
sation for the image distortions produced by the off-axis collimating combiner. Finally, 
complex geometric distortions have to be applied to the CRT image. These are produced by 
correspondingly distorting the electron beam deflection current drive waveforms. 

Notwithstanding the above complexities, the clean lines and low forward obscuration 
make this optical configuration the configuration of choice for high-capability, high- 
performance and wide-field-of-view applications. It is now introduced on to production 
prestige fast jet fighters such as Eurofighter Typhoon, Rafale, F-15 and Gripen. 

The Eurofighter HUD is shown in Figure 11.19 and Plate 4. It employs advanced 
computer-generated holographic optics to provide a 30° x 25° total field of view (TFoV). 
The instantaneous field of view is identical to the total field. It provides stroke (cursive), 
raster and hybrid modes of operation with outstanding display luminance of 2700 ft.L 
(9200 cd/m 2 ) in stroke (daytime) mode and 1000 ft.L (3500 cd/m 2 ) in raster (night-time) 
mode, this latter being viewable in daytime under cloud and haze. The outside-world 
transmission is 80%. 

The HUD also provides a comprehensive up-front control panel with a large-area daylight- 
viewable LED matrix display and programmable keys. The HUD is a high-integrity design 
and is used as the primary flight display in Eurofighter. 

11.3.7 Head-Up Display Functional Description 

Unlike the Harrier HUD discussed earlier, most modern HUDs contain all the electronic 
services to support the HUD functions, providing a low-level signal interface to the aircraft 






Figure 11.19 Eurofighter Typhoon HUD (BAE SYSTEMS). 

computer systems. A typical HUD block diagram is provided in Figure 11.20. The HUD 
comprises the following functional elements: 

1 . Optics assembly. A set of optical elements comprising the final collimating lens for the 
collimating HUD or the relay lens for the pupil-forming HUD, image field flattener lenses 
and the fold mirror with coatings and filters to reduce sunlight reflections. 

2. Combiner. A pair of optically flat parallel glass plates for the collimating HUD or a 
curved collimator for the pupil-forming HUD, with semi-reflective coatings tuned to the 
peak spectral emission of the CRT phosphor. 

3. Cathode ray tube. A high-brightness, high-resolution CRT used to produce bright, 
precision fine-line stroke- written graphics symbology and raster sensor video imagery. 

4. X and Y deflection amplifiers. These are high-precision power amplifiers that source current 
into the CRT X and Y magnetic deflection yokes to cause the CRT electron beam to trace out 
the stroke (cursive) symbology and also the raster sensor video scan waveforms. 

5. Video amplifier. Controls the CRT beam current by adjusting the CRT cathode bias with 
respect to the grid electrode to turn the beam on or off in stroke mode and modulate the 
beam with the sensor video image. 

6. Ramp generator. Strips synchronization pulses from the sensor video and generates the 
raster scan waveforms to align/harmonise the sensor video with the outside world. 

7. High-voltage power supply. Provides the final anode potential (usually around 18kV), the 
Al (focus) potential (usually around 2-4 kV) and the grid potential (usually around 200 V). 




A/C Power 

Stroke Bright-Up 

Sensor Video 

Stroke X /Y 

Low Voltage PSU 

High Voltage PSU 



Cathode Ray Tube 

X/Y Deflection Amplifier 

Figure 11.20 HUD functional block diagram. 

8. Low-voltage power supply. Provides all the low- voltage rails for the electronic circuits, 
typically + 5 V, ± 15 V, ± 20 V, etc. 

9. Chassis. Usually a complex precision casting, the chassis provides environmental 
protection to the HUD functional elements and affords a means of boresighting the 
HUD to the aircraft structure. 

The physical realization of these components into a fully operational HUD tends to be 
aircraft specific to match the operational capability requirements within the available 
cockpit envelope. The HUD electronic assemblies are packaged to make the best possible 
use of the available space envelope. The subassemblies of a typical HUD are shown in 
Figure 11.21. 

11.3.8 Image Generation 

To achieve maximum display brightness and contrast, the HUD day image (symbology) is 
written in stroke mode. The information content is in the X and Y beam deflection signals. 
The line width of the graphic elements is the CRT electron beam width. For best accuracy 
and finest spot size, the CRT is usually electromagnetically deflected and electrostatically 
focused. To avoid flicker, the entire image is redrawn or refreshed at a frequency greater than 
50 Hz (20 ms). The absolute brightness range is typically four orders of magnitude 
(10000:1) to span the ambient illumination range from bright sunlight to night-time 

Day/night Head-Up Displays provide the capability to present raster video images of 
forward looking infrared (FLIR) sensors on the HUD, requiring the HUD image to be 



Figure 11.21 Typical HUD subassemblies (BAE SYSTEMS). 

accurately registered with the outside world. It is necessary therefore to be able to adjust the 
size, the position and sometimes the orientation of the raster scan waveforms to match 
precisely the field of view of the sensor. 

A day/night HUD provides hybrid operation in which stroke- written symbology overlays 
the raster sensor image. A typical HUD FLIR hybrid (raster video plus stroke overlay) is 
shown in Figure 11.22. 

11.3.9 HUD Symbology and Principles of Use 

HUD symbology has historically been designed specific to the aircraft type, although some 
standardization is now emerging. Principal areas of symbol generation include the following: 



[7 10 

MflV * 

T ' Srr \j 

& ir 

1 il 

Figure 11.22 HUD FLIR image plus stroke symbology overlay (BAE SYSTEMS). 



Primary flight data; 

Navigation symbology; 

Air-to-air weapon-aiming symbology; 

Air-to-air weapon-aiming symbology. Primary Flight Data 

Primary flight data displayed on the HUD usually encompass: 

• Flight-path vector/marker (also known as velocity vector); 

• Attitude (pitch and roll); 

• Speed; 

• Altitude; 

• Heading; 

• Vertical speed (rate of climb/dive); 

• Angle of attack. 

A typical example of HUD primary flight symbology is shown in Figure 11.23. 

The flight-path marker presents the flight-path vector in 'contact analogue' form. This 
symbol indicates the instantaneous velocity vector, or direction, in which the aircraft is flying 
(and will continue to fly if the current manoeuvre is maintained). 

Attitude (pitch and roll) is also indicated in 'contact analogue' form as a horizon line in 
register with the real horizon, if visible. Additional lines (the pitch ladder) parallel to the 
horizon line indicate positive and negative pitch angles, usually at 5° intervals. 











Figure 11.23 HUD primary flight symbology (BAE SYSTEMS). 



Speed and height are presented as numerical values and a linear tape scale. Heading is 
usually presented as a numerical read-out on a horizontal scale, emulating a compass card 
viewed edge-on. Other numeric read-outs may also be included if they are generally helpful 
to the pilot to enable him to prosecute his mission effectively. These include instantaneous g, 
barometric set, waypoint lat/long, time/distance to go to next waypoint, etc. 

1 Navigation Symbology 

Navigation symbology is added to the primary flight information to provide en-route 
navigation to/from the mission target area. This information typically includes: 

• Lateral and vertical guidance to maintain the preselected flight plan; 

• Next waypoint position (latitude and longitude); 

• Bearing and range to next waypoint; 

• Time to go to next waypoint. 

On approaching the next waypoint, a conformal fix cross indicates the computed sightline to 
the waypoint on the basis of the best estimate of present aircraft position. The fix cross 
should precisely overlay the visually sighted waypoint. Any error represents an error in 
estimation of present position and can be corrected manually by the pilot. Air-to-Surface Weapon Aiming 

On entering the target area, the pilot selects the weapon to be deployed, and the HUD 
symbology changes accordingly. Figure 11.24 shows typical symbology for the release of 
dumb cast-iron, free-fall bombs on to a preplanned target. 










Figure 11.24 Typical HUD air- to- surface bomb symbology (BAE SYSTEMS). 


The following description is illustrative only: 

1 . A target designator box indicates the sightline to the target. The target designator should 
overlay the visually sighted target. The pilot should manually correct any errors. 

2. A continuously computed impact point (CCIP) indicates the sightline to the point on the 
ground where the selected bomb type will fall if fired now. This point is computed from a 
knowledge of the aircraft manoeuvre, height above target, bomb ballistics, ejection 
velocities and an estimation of wind over the target. 

3. A bombfall line indicates the locus of impact points on the ground of all future bomb 
releases and extends forwards from the CCIP marker taking due account of wind. The 
pilot steers to this line. 

4. A safe-pass height marker is often associated with the bombfall line to provide an 
indication of the fragmentation zone of the bomb. The bomb should not be released if the 
CCIP marker is close to or infringes the safe pass height marker. 

For a successful release the pilot should: 

• Fly the aircraft to position the bombfall line over the target ahead of the CCIP marker; 

• Continue to fly towards the target, tracking the target with the bombfall line - the target 
(and target marker) will move down the bombfall line as the distance to the target 

• When the target marker reaches the CCIP marker, the bomb should be released, either 
manually or automatically; 

Specific adjustments are made to the air-to- surface symbology for different weapons classes, 
different manoeuvres and of course different ballistics. Air-to- Air Weapon Aiming 

Should pilots elect to fire guns or missiles, then the appropriate aiming symbology appears to 
enable them accurately to aim and fire their weapons. Typical symbology is shown in 
Figure 11.25. 

A track line indicates the computed trajectory of the gun or missile on the basis of the 
current aircraft manoeuvre and weapon ballistics. A ranging circle is positioned on the line at 
the range of the target aircraft. A target designator box indicates the radar sightline to the 
target. Other information derived by the air-to-air radar about the target is also displayed. 
This might include, for instance, target motion relative to own aircraft motion (closure rate), 
possibly augmented with target slant range and g. 

Infrared heat-seeking missiles can in general be commanded to seek targets at angles 
significantly 'off-boresight'. A missile target marker indicates the missile sightline. 
Typically, the missile will be slaved to the radar. Once the missile detects an IR source, it 
will lock to that source and track it independently of the radar. Using the HUD, the pilot can 
confirm that the missile is locked to the correct source before launch commit. 

A firing solution exists when the radar target designator and the ranging circle are 
coincident and stable. Usually, the ranging circle will have a number of other bugs that 
indicate the minimum and maximum firing range, together with the probability of success 












Figure 11.25 Typical HUD air-to-air symbology (BAE SYSTEMS). 

based on the manoeuvres executed by both aircraft and on the manoeuvre required by the 
missile for a successful intercept post-launch. 

11.4 Helmet-Mounted Displays 

The field of view of the HUD is a very limited field of view when compared with the total 
hemisphere of regard afforded to the pilot in the bubble canopy of a modern fast jet fighter 
aircraft, and limited when compared with the field of regard of a modern radar (typically 
120°) and the acquisition cone of modern air-to-air missiles (typically 90°), as graphically 
shown in Figure 11.26. 

Helmet-mounted displays overcome this limitation. By the use of miniature display 
technology producing a display for each eye, combined with accurate head tracking, it is 
theoretically possible to present a stereoscopic, full-colour image to the user in any direction. 

Although display technology and the processing speed of graphics generators have not yet 
matured to the point where the virtual world image is possible in an aircraft environment, it 
will undoubtedly materialize in the near future (Garland et ai, 1994). Putting the HUD onto 
the pilot's head, even with the same field of view, but now free to move wherever the pilot is 
looking, makes it possible to cue, acquire, designate, track and release weapons at targets 



Radar FoV 





Figure 11.26 Field-of-regard comparisons. 

significantly off-boresight without having to manoeuvre the aircraft. The increase in 
operational effectiveness is a significant 'force multiplier'. 

Figure 11.27 shows some of the operational possibilities, which include: 

• Off-boresight target cueing by directing the pilot/weapon aimer to look in the direction of 
potential targets or threats detected by on-board sensors and/or advised by data link from 
cooperating aircraft and ground stations; 

• Designation 'off-boresight' once the target has been recognised - tracking accuracy of the 
on-board systems and weapons can then be monitored to assure lock is maintained; 

Figure 11.27 Possible uses of a HMD to cue, designate and aim 'off-boresight' 



• Weapon release can be commanded once a launch solution is reached without having to 
manoeuvre the aircraft to be within the traditional success cone determined by the HUD 
field of view; 

• Handover is facilitated between crew members and between cooperating aircraft flying 
the same mission. 

The first practical HMD recorded was made by Autonetics in the United States during the 
early 1960s. By 1970 a system had been flown in a US Navy F-4, which proved the key 
points. That HMD had a sighting reticle and operated with an optical tracker, facilitating the 
pointing of radar and missiles. 

The joint helmet-mounted cueing system (JHMCS) installed on the F/A-18 aircraft and 
other USAF and US Navy fast jets provides a monocular display for use in daytime 

The application of a binocular, day and night capability has taken longer to mature on fast 
jets. The Eurofighter Typhoon will probably be the first. However, the next generation of fast 
jet fighters, typified by the JSF, is likely to replace the HUD with an HMD, and to have the 
HMD as the primary flight instrument. 

1 1.4. 1 HMD Physiological and Environmental Aspects 

The HMD is considerably more than an HUD on the head; the helmet already provides a 
number of facilities and life- supporting functions with which the display components must 
be integrated. These are illustrated in Figure 11.28, including: 

• Communications microphone; 

• Earphones (possibly with active noise reduction); 

• Oxygen mask; 

• Retractable sun visor; 

Protect Head & Eyes 
During an Ejection 

Reduce Gare 
from Sunlight 

Laser Eye 



Attenuate Cockpit Noise 

(Active Noise Reduction- ANR) 

Life Support - 
Oxygen Mask 

Figure 11.28 Normal helmet functions. 


• Possibly NBC protection; 

• Head protection during ejection. 

The provision of a Helmet-Mounted Display requires the addition of some extra components, 

• An image source (e.g. Cathode Ray Tube) or more than one source if a binocular image is 

• Relay optics to position the HMD image at the focal plane of the collimator; 

• Collimating optics to generate a virtual image focused at infinity; 

• A partially transparent/partially reflective combiner to introduce the collimated image 
into the pilot's line of sight. 

Adjustments need to be made to centre the optical image axis onto the pilot's eyeball. 
Typically, these include: 

• The vertical position of the eyeballs with respect to the pilot's head; 

• The distance between the eyeballs, the interpupilary distance; 

• The eye relief between the eyeball and the first optical element. 

Furthermore, the additional HMD components must not introduce features that may 
jeopardise the integrity of the helmet to protect the pilot from hazards such as: 

1. Birdstrike. It is potentially possible that a birdstrike may cause the canopy to shatter. 

2. Ejection. In the event that pilots have to eject from aircraft, the helmet must protect them 
from the acceleration forces of the ejection itself and from the windblast (in excess of 
600 knot). 

The range of human anthropomorphic geometry is vast. Usually it is necessary to have a 
range of helmet sizes. The helmet must be a firm fit to the pilot's head so that no movement 
can occur once the optics have been 'boresighted' to aircraft axes, but at the same time be 
comfortable to wear and keep the head cool. Often this is accomplished using a permanently 
deformable liner. 

The weight of the helmet must be minimised to reduce fatigue. A typical helmet before the 
addition of an HMD weighs less than 2.0 kg (4.5 lb). Additionally, the CG of the helmet should 
be in line with the pivotal point of the head on the spine to minimize out-of-balance forces. 

With all the HMD optical components added to the helmet, the helmet becomes a high- 
value item. For this reason, more recent designs allow the HMD optical elements to be 
removed from the helmet to reduce its intrinsic value. 

Achieving all of these requirements represents a significant challenge to the Helmet- 
Mounted Display designer. The requirements are often conflicting. 

11.4.2 Head Tracker 

To be effective as a sight, accurate knowledge is required of the head-pointing angle. The 
head can move vertically, laterally and fore/aft with respect to the body torso and can rotate 
in all three axes about the pivotal hinge point of the head on the spine. 



Figure 11.29 Optical head tracker (Kentron). 

11.4.3 Optical Head Tracker 

A number of schemes have been used. An early design used a number of sources positioned 
around the cockpit to generate swathes of scanning beams. Sensors mounted on the helmet 
detected these beams, providing signals to the head-tracking electronics to decode the 
interrelationship between the scanning beams and the time of detection and derive the head 
pointing angles. 

A more recent scheme, and that being used on the Eurofighter Typhoon, uses twin- 
tracking CCD cameras to sense clusters of LEDs on the surface of the helmet (Figure 11.29). 

11.4.4 Electromagnetic Head Tracker 

Both ac and dc electromagnetic fields have been used. In general, a source transmitter placed 
just behind the pilot generates an electromagnetic field in the cockpit. A three-axis 
orthogonal sensor mounted on the helmet detects the local electromagnetic field and provides 
signals for the head-tracker electronics to decode the amplitude and phase signals and derive 
the head pointing angles. 

The electromagnetic field is significantly distorted by the metal structures within the 
cockpit, and it is necessary to map the cockpit to characterize it fully. This mapping needs to 
take account of movable features within the cockpit, for instance, seat height position. 
Usually it is necessary to map only one aircraft of a type, although the mapping will have to 
be repeated if there are any subsequent equipment changes which might impact the magnetic 
properties of the cockpit. 

11.4.5 HMD Accuracy and Dynamic Performance 

HMD pointing accuracy comprises two elements, the first an optical accuracy similar to that 
of a HUD system plus a tracker error, which increases rapidly at extended angles from 








5 5 

Off-Boresight Angle 

Figure 11.30 Comparison of HUD and HMD accuracy (BAE SYSTEMS). 

boresight and at the margins of the head motion box. Typical HUD and HMD system 
accuracies are shown in Figure 11.30 (Bartlett). 

Since the helmet-mounted sight is not fixed to the aircraft structure, it is necessary to provide 
a means to 'boresight' the HMD system before each mission. Typically, the boresighting 
procedure comprises sighting a collimated target conveniently located in the cockpit. 

Using a helmet-mounted sight to launch a weapon with 'smart' guidance has been shown 
to be highly effective, especially since the target can be acquired and designated and the 
weapon launched 'off-boresight'. The HMD accuracy is adequate for this task. However, 
using a helmet-mounted sight to target a fixed gun or release dumb weapons is generally 
recognised as being inappropriate. 

The dynamic performance requires careful consideration since rapid movements of the 
pilot's head (30deg/s) are entirely possible. System bandwidth and delays must be 
commensurate with the operational role in order to avoid disorientation that might be 
caused by 'swimming' symbology. For this reason it is usual closely to couple the head- 
tracking electronics with HMD symbol generation, minimising any data latency problems. 

11.4.6 HMD Optical Configurations 

The optical configuration should be selected according to operational use. Display perfor- 
mance needs to be balanced with complexity. Bigger is not necessarily better if it weighs 


Monoscopic // Stereoscopic Monoscopic 

U (Overlap) 

Figure 11.31 Monocular or binocular configuration. 

The choices broadly are as follows: 

1 . Sight or display. A sight is used simply as a target designator, possibly with the addition of 
simple direction-finding cues (such as look-up/down, left/right) and event annunciators 
(such as weapons lock/fire). A display generally adds other symbology such as primary 
flight data, weapons data and imagery at night. 

2. Monocular or binocular. Monocular is suitable for daytime use when the pilot's attention 
is mostly on the outside world. However, a binocular configuration is preferred at night 
when vision is augmented by other aids such as FLIR or NVGs. 

3. If binocular, then one image source or two. Both configurations have been used. If dual 
sources are used then it is possible to extend the azimuth field of view having a central 
area in which the image is seen by both eyes with two monocular areas either side, as 
shown in Figure 11.31. 

4. Off -visor or periscope optical system. The periscope optical system is easier to design and 
manufacture and therefore was the first to be used. Off- visor systems are less obstructive, 
more technically elegant but demand more complex solutions if wide fields of view are 

5. Day, night or day/night operation. It is a self-evident requirement that a daytime display 
must be viewable in direct sunlight. At night the image may be augmented with sensor 
video and/or night- vision goggles. 

The optical configuration choices for a Helmet-Mounted Display are identical to those used 
in Head-Up Displays. Typical arrangements are shown in Figure 11.32. The lens diameters 
are considerably reduced when compared with those of a Head-Up Display since the 
effective exit lens is close to the pilot's eyeball and thus a 30-40° field-of-view porthole is 
obtainable with lens diameters of around 1 cm. However, the intermediate optics to relay the 
image from its source (typically a miniature CRT) to the focal plane of the collimator are 




Pupil Forming 


curved combiner 

Pupil Forming 

off-axis off-visor 

curved combiner 


Figure 11.32 Typical optical configurations. 

often very complex since they need to bend the light path around the head without being 
obstructive and at the same time maintain the CG close to the head natural CG. 

Figure 11.33 shows a generic HMD optical system. The EO component elements are: 

1. Image source. Usually a high brightness, high resolution miniature CRT (though early 
sights used a small LED matrix. Emerging technologies are AMLCD (transmissive and 
reflective, electro-luminescent (EL), organic LED (OLED) and low-power laser scanning 
directly into the eye-ball. 

2. Relay and fold mirrors. The relay is a group of mirrors, lenses and prisms, which translate 
the image from the source into the correct place to be projected into the pilots' line of 
sight by the combiner. 

3. Combiner. The combiner projects the image from the relay lens into the pilots' line of 
sight. It may be flat or have optical power. It is semi-transparent to the outside world, but 
reflective to the specific wavelength of light from the image source. 

Image Source / i i 

\ i i- 

Relay Optic 

Fold Mirrors 


Figure 11.33 Generic HMD optical system. 



Figure 11.34 IHADS helmet (Honeywell). 

11.4.7 Helmet-Mounted Displays - Examples 

1 1.4. 7. 1 Integrated Helmet and Display Sight System (IHADS) 

The first users of HMD technology in aerospace applications were the rotorcraft community. 
The HMD weight issues are less significant in a rotorcraft, which is not able to pull as much 
g as a fast jet fighter. Neither is ejection or windblast a concern. 

Figure 11.34 shows the Apache AH-64 attack helicopter Honeywell- supplied integrated 
helmet and display sight system (IHADS) helmet worn by the copilot/gunner who has 
primary responsibility for firing both guns and missiles. 

The IHADS provides a 40° x 30° field-of-view, monocular (single-eye), collimated 
image from a miniature Cathode Ray Tube (CRT). The images constructed on the CRT are 
derived from the nose-mounted target acquisition and designation sight/pilot night-vision 
sensor (TADS/PNVS). The TADS/PNVS and the gun are slaved to the copilot/gunner head 
line of sight. The system has proven itself to be operationally effective; however, there 
have been some reports of fatigue after extended use at night, possibly caused by the 
contention created in the brain of a monocular sensor image in one eye and the real world 
view in the other. Helmet-mounted Sight 

The simplest and earliest form of Helmet-Mounted Display used in fast jet applications was 
the helmet-mounted sight (HMS), used in daytime operations to acquire and designate 
targets off-boresight. The HMS installed and worn by RAF Jaguar pilots is shown in 



Virtual Image 

of Aiming 


Parabolic Dark 

Figure 11.35 Jaguar helmet-mounted sight (BAE SYSTEMS). 

Figure 11.35. The Jaguar HMS comprises a simple fixed aiming reticle formed from an array 
of bright LEDs. Simple optics collimate the image, which is projected off the visor into the 
pilot's line of sight. Joint Helmet-mounted Cueing System (JHMCS) 

The JHMCS programme is the culmination of extensive trial evaluations performed using 
helmets such as the Kaiser agile eye. 

The JHMCS, shown in Figure 11.36, is designed to provide first-shot, high off-boresight 
weapons engagement capabilities enabling the pilot to direct weapons against enemy aircraft 

Figure 11.36 Joint helmet-mounted cueing systems (JHMCS) (Kaiser Electronics). 



while performing high-g aircraft manoeuvres. The system can also be employed accurately 
to cue the pilot to ground targets. Targeting cues and aircraft parameters are displayed 
directly on the pilot's visor. The JHMCS is designed to have low weight, optimised CG and 
in-flight replaceable modules to enhance operational performance - including the ability to 
be reconfigured in-flight to meet night- vision requirements. 

The JHMCS Helmet-Mounted Display has the following features: 

• A monocular field of view of 20° ; 

• An 18 mm exit pupil with an eye relief of 50mm; 

• The display module is compatible with US Air Force HGU-55/PU.S and US Navy HGU- 
68/P helmets; 

• The weight is 41b (1.82 kg) with mask. MTBF is 1 000 h. Eurofighter Typhoon HMD 

The Helmet-Mounted Display in development for Eurofighter Typhoon is shown in 
Figure 11.37. Features are: 

1. Provision of primary aircrew protection. 

2. Attachment to inner helmet. 

3. Provision of a lightweight but stiff platform for the optical components: 

• CRTs/optics/mirror; 

• Night- vision cameras; 

• Blast/display and glare visors; 

• Head tracker diodes (infrared). 

Figure 11.37 Eurofighter Typhoon helmet-mounted display (BAE SYSTEMS). 


4. Removable night- vision cameras, autodetach during ejection. 

5. Mechanical function: 

• Protection features in line with survivability limits; 

• Life support up to the limits of human functionality; 

• Comfort and stability to support display requirements. 

6. Display features: 

• 40° x 30° binocular field of view; 


Corresponding 40° x 30° binocular night-vision camera; 
Display of sunlight visible symbology and/or imagery. 

The optics arrangement (dual binocular) uses a 1 in high-brightness, high-resolution, 
monochrome (P53 green) CRT as the image source. A complex relay lens with a brow 
mirror introduces the relayed image into the focal plane of two spherical diffractive mirrors 
which are deposited on the visor by holographic techniques. The field of view is 40° and the 
exit pupil is 15-20 mm. The optical arrangement introduces significant geometric distortion, 
which is corrected electronically. 

The Helmet is a two-part arrangement (see Figure 11.38). 

The inner helmet fits inside the display outer helmet and can be swapped with a respirator 
hood version to give NBC protection: 

1. It facilitates individual user fitting. 

2. It maximises comfort and stability: 

• The brow pad is moulded to exact fit; 

• Air circulation around the head is assisted. 

3. It embraces a wide anthropomorphic range: 

• Advanced suspension system; 

• Lightweight oxygen mask. 

Helmet Helmet 

Inner Outer 

Figure 11.38 Eurofighter Typhoon two-part helmet (BAE SYSTEMS). 



4. It provides optimum hose/cable routing. 

The display outer helmet attaches to the inner helmet. It provides: 

1. Primary aircrew protection. 

2. A lightweight but stiff platform for the optical components: 

• CRTs/optics/mirror; 

• NVE cameras; 

• Windblast/display visor (clear) and glare visors. 

3. Head tracker diodes. 

4. Removable night- vision cameras. Joint Strike Fighter HMD 

The joint strike fighter (JSF) will have a binocular day/night Helmet-Mounted Display and 
no Head-Up Display. The HMD will provide all the targeting information and sufficient 
primary flight information to meet the mission objectives. 

Figure 11.39 shows a prototype of the JSF HMD. The helmet provides a 30° x 50° total 
field of view from two overlapping 30° x 40° monocular images. The optics is a pupil- 
forming arrangement with the relayed image collimated and introduced into the pilot's 
sightline by a diffractive holographic semi-transparent curved combiner introduced into the 
visor. The exit pupil is 18 mm. The image source is a high-resolution (SXGA) transmissive 
LCD, illuminated with a bright LED backlight. 

Figure 11.39 JSF helmet mounted display (VSI). 




Head Tracker 

Helmet Control 
Panel Interfaces 


Helmet Shell 

HMD Electronics Unit 

HMD Cockpit 


Aircraft Wiring 


Quick Release 

HMD Tracker 



Figure 11.40 Typical HMD system components. 

1 1.4.8 Helmet-Mounted Display Functional Description 

A Helmet-Mounted Display is a heavily distributed system. For obvious reasons, only the 
minimum of components are carried on the head to minimise weight. 

A typical HMD system is shown in Figure 1 1 .40 and comprises the following line 
replaceable units (LRUs): 

1. Basic helmet assembly. The basic helmet provides a high level of pilot protection and 
comfort and is fitted with communications equipment, a form-fitting system and display 
module-mounting points. The basic helmet is form fitted to an individual pilot and 
becomes part of his personal equipment. 

2. Display module assembly. This often simply clips on to the basic helmet shell and 
contains the display components in a single-sized lightweight module. Dependent on 
the functionality, the components located within the display module are dual-image 
intensifiers, dual CRT displays, optical assemblies, a dual-visor system, a helmet tracker 
receiver, autobrilliance sensor, battery pack and umbilical cable. 

3. Helmet electronics unit. This is located in the avionics bay and contains the main HMD 
system electronics functions including system interfaces, processing, display drive and 
helmet tracking. 

4. Helmet tracker transmitter. This small unit is mounted either on the aircrew seat or the 
canopy, and is part of the helmet tracker function. 

5. Boresight reticle unit. This provides accurate optical reference for boresighting the helmet 
tracker system. 



Electron Beam 

X Deflection 

A3 Final Anode 

A1 (+1 KV) 
Grid (-50V) 


A2 (+500V) 


Electron Gun 

Y Deflection 

Magnetic Current 

Deflection Yoke 





Mounting Ring 

Figure 11.41 Typical HMD miniature CRT. 

6. Quick-release connector cable assembly. This comprises the cable and connectors running 
to the HMD quick-release connector. 

Most current Helmet-Mounted Displays use a miniature CRT as the image source 
(Figure 11.41). The CRT has a faceplate image of typically 20 mm (3/4 in) diameter and 
an overall length of around 100 mm (4 in). The CRT image is generated by an electron gun 
and focused electrostatically. The final anode potential is typically 13kV and the electron 
beam current is around 100 mA. Two orthogonal scan coils electromagnetically deflect the 
electron beam. To achieve high-brightness daytime viewability, the image is constructed in 
cursive (stroke) mode. If the HMD is also used for displaying night-time sensor imagery, 
then the combined sensor plus symbology image is constructed in raster or hybrid 
(stroke + raster) mode. A fibre-optic faceplate produces a flat image plane and enhances 
the image contrast. 

Next-generation HMDs may employ an LCD as the image source together with a bright 
LED backlight. This technology would avoid the need for high voltages. 

1 1.4.9 Binocular Day/Night HMD Architectures 

The Eurofighter Typhoon and next-generation HMDs are seeking to integrate NVGs with the 
HMD image generation function. There are two means to achieve this, optically and 
electronically (Jukes, 2004 - Chapter 8). 

11.4.10 HMD Symbology 

A typical symbology format is shown in Figure 11.42. In all current HMD installations the 
HUD is generally the primary flight instrument. Primary flight data on the HMD play a 
secondary role and are only used subliminally for quick orientation. Pilots defer to the HUD 
if they become seriously disoriented. In this example, attitude information is presented as a 
simple line indicating pitch and roll against a fixed circle. The extensions at the end of the 



Auxiliary Targets 

Figure 11.42 Typical HMD symbology format (Thales). 

intersecting line indicate normal/inverted flight. Speed, height and heading are presented as 
simple numerical read-outs. 

The primary aspects of the display format relate to targeting data to several priority 
targets. If the target is not within the HMD field of view, then a line with an arrowhead 
indicates the direction in which the pilot should look to acquire the target visually. Radar 
data such as range to target and target g level are displayed alongside the target marker. If 
missiles are selected and have acquired a target, then a marker also shows the line of sight to 
the missile target. Additional information indicates probability of launch success. 

11.4.11 HMD as a Primary Flight Reference 

In the future it is probable that HMD systems will replace the HUD as the primary flight 
reference. Doing so will free up the significant cockpit real estate occupied by the HUD as 
well as greatly increase the look-through display field of regard available to the pilot for 
weapon aiming and sensor display. HMDs provide a virtually unlimited field of regard for 
off-boresight targeting, weapons employment and display of sensor information. 

The seeming advantages of this increased area of display space must be approached with 
caution, however. There is a strong potential for targeting sensor information presented 
off-boresight to draw the pilot's attention away from the traditional on-boresight flight 
reference information for longer periods of time. During periods of limited visibility, the 
increased off-boresight time could result in reduced spatial awareness if some sort of flight 
information is not presented in the off-boresight display. While the presentation of primary 
flight information in HUDs is well understood and documented, methods for presenting off- 
boresight primary flight information are not as mature. 

Ongoing work is being undertaken by the USNR Naval Air Warfare Center Aircraft 
Division Patuxent River and Boeing to design and evaluate operational HMD symbology 
formats that fulfil the primary flight reference (PFR) requirements (Foote). 


11.5 Head-Down Displays 

The multifunction head-down display, variously called the MFD, the MHDD, the MPD 
multipurpose display (MPD) and the multipurpose colour display (MPCD), provides the 
flight-deck/crew-station/cockpit designer with a flexible display media on which to present 
data in a variety of ways according to the information needed by the crew for the current 
phase of the flight or mission. 

At the outset of the electro-optical era, display capability was limited primarily to that able 
to be provided by the monochrome (usually green) CRT. The significant breakthrough came 
with the application of the shadow-mask CRT to airborne environments. The shadow-mask 
CRT is perhaps the first example of commercial off-the-shelf (COTS) technology applied to 
airborne applications. It took significant developments (albeit evolutionary rather than 
revolutionary) to make the shadow-mask CRT both bright enough and sufficiently robust 
to be able to be used in the severe environment of airborne applications. 

However, CRT technology is bulky, heavy and power hungry and requires extremely high 
voltages to provide a high-brightness display, posing serious limitations to the crew-station 
designer. The advent of AMLCD technology offered the promise of eliminating all those 
problems. In the event, AMLCD technology has proved as difficult if not more difficult to 
apply to airborne applications. Although small (i.e. flat), low weight and low voltage, the 
performance of an AMLCD varies with viewing angle (particularly an issue for instrument 
reading cross-cockpit), and it is temperature sensitive and fragile. Furthermore, as was learnt 
very painfully, extreme attention to manufacturing process control is required to achieve 
consistent product quality. This generally only comes with high- volume production. Custom- 
size low-volume aerospace product is to be avoided if at all possible. 

However, these problems are now largely solved and the AMLCD is the technology of 
choice. Display sizes for new crew-station designs are transitioning from the custom square- 
format sizes historically used in aircraft to rectangular laptop PC and professional/industrial 
glass sizes. The focus now is to ruggedise COTS AMLCD devices, not design custom 
aerospace glass. 

1 1.5. 1 CRT Multifunction Head-Down Display 

The Cathode Ray Tube was the first fully flexible display device to be used for primary flight 
displays in the cockpit. 

1 1.5. 1. 1 Shadow-mask CRT 

The shadow-mask CRT shown in Figure 1 1.43 and Plate 5 comprises an evacuated glass bulb 
in which an electron gun emits electrons at high velocity to impact on a phosphor screen. The 
electron gun comprises three cathodes, today usually the in-line configuration, although early 
applications used the delta configuration. 

The shadow-mask CRT is a large heavy device, and not easy to package. Typically, a 
6.25 in x 6.25 in CRT is 35 cm (14 in) long and weighs 5 kg (111b), including magnetic 
components and shield. 

Electron emission is modulated by three grid electrodes, focused by a multistage electron 
lens and finally accelerated towards the faceplate through a shadow mask. A series of red, 



Figure 11.43 Shadow-mask CRT. 

green and blue phosphor dots are deposited on the CRT faceplate in a process that uses the 
shadow mask itself to maintain precise registration. The electron gun, shadow mask and 
phosphor dot geometry is arranged so that electrons emitted by each cathode only illuminate 
phosphor dots of its designated colour. The composite beam bundle is deflected in the X and 
Faxes by a magnetic yoke to scan the CRT display surface. Modulating the individual beams 
produces a full-colour image comprising three registered primary colour images. 

The phosphors deposited on the CRT faceplate emit light in the red, green and blue 
primary colours. The red and green are narrow-band, short-persistence (<10|is) phosphors; 
the blue is broadband, longer-persistence (1 ms) phosphor. 

The gamut in which colours can be produced by a shadow-mask CRT is best described in 
CIE u' v' colour space. By suitable combination of cathode drive voltages, any colour can be 
produced within the triangle described by the three primary colours: 

u' v' Dominant wavelength 




610 nm 




550 nm 




460 nm 

Figure 1 1 .44 and Plate 6 shows the colour gamut of a shadow-mask CRT. Also shown is the 
gamut available from an AMLCD for comparison. The CRT has a purer blue, although it can 
be argued that this is of little operational advantage. 





0.1 Yellow 0.2 0.3 u 0.4 

CLE. 1976U.C.S. 
— Diagram 

0.2 0.3 

Figure 11.44 CRT and AMLCD colour gamut. 

To display a day light- viewable, full-colour topographical map in raster mode in the high- 
ambient illumination conditions to be found in a fast jet fighter requires a considerably 
brighter shadow-mask CRT than used in domestic TV. A specialised custom CRT was 
developed by Tektronix (later Planar) with funding from industry and the US DoD (Jukes, 
2004 - Chapter 5). X/Y Deflection Amplifier 

To produce a high-brightness, daylight- viewable, high-quality text and symbolic display, the 
shadow-mask CRT is usually operated in stroke (cursive) mode, that is, the CRT beam is 
caused to trace out the desired symbology as a pen plot. The beam current is modulated 'on- 
off ' when raising/lowering the pen, and 'off when moving (slewing) to the next symbol. In 
this way the beam can be moved more slowly (typically 10 times slower) than when 
producing a raster scan and the image is correspondingly brighter. Format content is 
therefore limited by the deflection writing rate. 

When displaying a topographical map or sensor video, the CRT is operated in conven- 
tional raster mode, that is, the CRT beam traces the whole display surface area from top to 
bottom as a series of left-to-right horizontal lines, then retraces back to the top to start again. 

The video field retrace time can be used to write stroke symbology over the raster. The 
slower stroke writing rate usefully emphasises the stroke symbology. This mode of operation 
is called 'hybrid'. Typical hybrid X, Y and video waveforms are shown in Figure 11.45. 

A detailed description of the constituent functional components of a CRT multifunction 
display can be found in Jukes (2004 - Chapter 5). 

1 1.5. 1.3 Shadow-mask CRT Characteristics 

The shadow-mask CRT has a number of characteristic features that require careful attention 
if acceptable performance is to be obtained in aerospace applications: 



X Deflection 

Y Deflection 

j4^Mj4Mj4M4mJ^ nFLTTJTj^M^ 

Raster Video/ 
Stroke Bright-Up 

Raster Period 

Stroke during 
raster retrace 

Raster Period 

Figure 11.45 'Hybrid' (stroke + raster) operation. 

1. Contrast. Unilluminated by the electron beam, the phosphor dots are a whitish-grey 
colour. A contrast enhancement filter has to be fitted to the CRT to reduce the sunlight 
reflections that otherwise would occur in the adverse lighting conditions of the 

2. Colour purity. Impure colours can be produced by small residual magnetic fields. Changes 
in the earth's magnetic field are sufficient, hence the need to provide protection afforded 
by a mu-metal shield. 

3. Convergence. It is essential to maintain the constituent parts of the image in perfect 
registration; that is, the three electron beams must converge at all points on the CRT 
faceplate. Convergence is maintained by static and dynamic electromagnets placed close 
to the electron gun. CRT MFD: Principles of Operation 

The block diagram of a typical colour CRT technology based MFD is shown in Figure 1 1.46. 
The functional elements of the CRT MFD comprise: 

• Shadow-mask CRT; 

• High- and low- voltage power supply; 

• X and Y deflection amplifier; 

• Video amplifier; 

• Ramp generator; 

• Encoder; 

• Key panel; 

• Microcontroller. F/A- 18 and A V-8B Multipurpose Colour Display (MPCD) 

The MPCD shown in Figure 1 1 .47 uses a taut shadow-mask CRT to achieve a full-colour, 
high-brightness, high-resolution display in stroke, raster and hybrid modes. The cardinal 
point performance specification is set out in Table 11.1 



A/C Power 

Low Voltage PSU 

High Voltage PSU 





Stroke X /Y 





Shadow Mask CRT 

Convergence Deflection 

X/Y Deflection Amplifier 


Figure 11.46 Colour shadow-mask CRT MFD block diagram. 

Figure 11.47 F/A-18 and AV-8B multipurpose colour display (MPCD) (Smiths). 



Table 11.1 Cardinal point performance specification of the MPCD 

Usable screen area 

5 in x 5 in square 


Luminance (ft.L): 















Contrast ratio at lOOOOft.C: 















Stroke writing speed 


Line width 

0.018 in (0.46 mm 

l) to 50 % point 


65 line pairs/in 


6.7 in wide x 7.01 

in wide x 17.3 in deep 




115 V, three-phase, 400 Hz, 180VA 


165 W at maximum brightness 


3500 hours 

11.5.2 AMLCD Multifunction Head-Down Display 

Active matrix liquid crystal display (AMLCD) is now the accepted technology for all new 
applications. The principle advantages of the AMLCD (with backlight) are: 

• Significantly less depth than a CRT and easier to package (typically <60mm); 

• Significantly less weight than the CRT (typically <1 kg); 

• Significantly less power than a CRT (typically <20W); 

• No high voltages; 

• No magnetic components and no influence from external magnetic fields; 

• Perfect registration (no misconvergence); 

• Fixed pixel (spot) size at all display brightness levels. AMLCD Display Head Assembly 

In an AMLCD, liquid crystal (LC) material is introduced between two glass plates (known as 
the active plate and the passive plate). The spacing between the glass plates is critical and is 
around 4 urn (LC material specific). 

The following is a much- simplified explanation of the operation of the cell. Reference 
should be made to Figure 11.48. 

Liquid crystal material is a viscous organic fluid containing long polymer chains, which 
have the property of rotating the plane of polarisation of light according to their alignment 
axis. The alignment axis can be changed with electrical bias. 

The plane of polarisation of light alters as it passes through the cell. Placing two crossed 
polarisers either side of the cell (bonded to the glass plates) makes the cell into a light valve. 



Figure 11.48 AMLCD cell 


High Aperture 
BL Matrix 

High Aperture 
BL Matrix 

High Aperture 
BL Matrix 

High Aperture 
BL Matrix 

With the cell unbiased, light is plane polarised by the first polariser, its plane of 
polarisation is rotated by 90° as it passes through the LC material and it then passes through 
the second polariser. In the biased state, light is not rotated by the LC material and does not 
pass through the second polariser. Greyscale is obtained by applying an intermediate voltage, 
which partially rotates the molecules so the cell is partially transmissive. 

Accurately maintaining the cell gap is vital for correct operation. To achieve this, the LC 
material is filled with very accurately machined glass beads in suspension that become 
randomly distributed throughout the cell and are sandwiched between the glass plates under 
carefully controlled pressure during manufacture. 

Row and column address lines form a matrix with an amorphous silicon thin-film 
transistor (TFT) at each intersection. The processing steps are similar to any silicon foundry 
process except in this case on to large glass sheets. The row and column drive signals to the 
TFT control charge applied across the local cell defined by the TFT itself plus an associated 
pair of transparent electrodes placed opposite one another on the active and passive plate. 
This charge is the bias for the local light valve. 

Red, green and blue colour filters are deposited on the passive plate in exact register 
with the subpixels to form a colour group or pixel. A number of colour group arrangements 
have been popular, as shown in Figure 11.49, but now the accepted norm is the R:G:B stripe. 

A typical X- video graphics adaptor (XVGA) compatible panel is shown in Figure 11.50. 
This panel has been ruggedised for the airborne environment. It is 6.25 x 6.25 in square and 
has 768 x 768 R:G:B pixels; a total of 1.8 million pixels. 

The TFT row and column address lines are brought out to the edge of the glass plate. That 
is a total of around 1700 connections for the XVGA panel described. The lines are driven by 
row and column driver integrated circuits (ICs). Drive voltages are typically in the range -5 
to + 15V. 

The driver ICs are mounted on printed circuit boards which are folded back at 90° to the 
AMLCD glass panel. The ultrahigh-density interconnect between the drivers and the 










double green 



Figure 11.49 Colour group pixel structures. 

AMLCD glass panel is made with tape automated bonding (TAB). The TAB process uses 
copper traces in a flexible poly amide base layer. The driver ICs are mounted on the TAB. 
Connections are made between the TAB and the glass panel using a pressure-sensitive 
anisotropic adhesive. The adhesive contains small silver balls that touch each other under 
pressure, and make contact between circuit traces in the TAB and circuit traces on the glass. 
The balls do not touch in the axis perpendicular to the pressure axis and isolation is 
maintained between adjacent traces. Once the connections are correctly made, the adhesive 
is heated and permanently sets. Backlight 

The AMLCD is a light valve but with a transmission of less than 7%. To obtain a usable 
display requires the addition of a bright backlight (Figure 11.51). This is an important 
distinction between the CRT and the AMLCD. In the CRT the image and the light source 
are one and the same, namely the phosphor light emitter. In the AMLCD the image is 
separated from the light source; the image is made by the LCD; the backlight is the light 

The fluorescent lamp is the technology of choice. A number of fluorescent lamp backlight 
configurations have been used, hot or cold cathode, single or multiple lamps. The single, 

Figure 11.50 AMLCD panel with drivers (Korry Electronics). 



Figure 11.51 Single, cold-cathode, serpentine backlight (Korry Electronics). 

serpentine, cold-cathode lamp has become the configuration preferred by most designers 
(Jukes, 1997). An optical stack is placed between the fluorescent tube and the backlight to 
provide a uniform diffuse light source to back illuminate the AMLCD. 

The luminance of the backlight must be controlled to match the ambient illumination. For 
normal daytime operation the backlight is operated as a conventional fluorescent tube; that 
is, the arc is struck with a high-frequency alternating current (typically in the tens of kHz 
range). To dim the lamp, the hf pulses are gated to provide pulse width modulation of the 

A detailed description of the constituent functional components of an AMLCD multi- 
function display can be found in Jukes (2004 - Chapter 5). AMLCD Characteristics 

There are a number of characteristic features of the AMLCD, different to those of the CRT, 
that require careful attention if acceptable performance is to be obtained in aerospace 

1 . Viewing angle. The optical performance of an AMLCD is critically dependent on the cell 
gap. At increasing viewing angles from the normal to the display the optical cell gap 
effectively increases and the optical performance of the cell degrades. It no longer acts as 
a perfect light valve, and some light 'leaks' through the cell. This light leakage reduces 
the display contrast off-axis. For adequate visual performance a contrast ratio of 50:1 is 
desirable, and 30:1 is acceptable over the normal viewing cone. This represents a 
considerable challenge for AMLCD technology. 


2. Grey scale. Grey scale is obtained by partially switching the cell between its 'on' and 'off 
states. The grey scale transmission curve is not linear. It is viewing angle dependent and 
requires temperature compensation since it is sensitive to temperature. 

3. Black level uniformity. This characteristic has been perhaps the most problematic feature 
of AMLCD (second only to viewing angle). It is particularly a concern for symbolic 
images painted on a black background; it is not so significant for video images which 
typically have little black content. Many factors can result in a blotchy black background, 
which can be most distracting. This feature, more than any other, has received adverse 
criticism for the AMLCD. Fortunately, improvements in materials and in process control, 
driven by the volume laptop PC market, are overcoming this deficiency. 

4. Thermal management. Many of the parameters associated with the performance of the 
AMLCD are temperature sensitive and it is necessary actively to manage the thermal 
environment of the AMLCD. The performance of the fluorescent lamp is temperature 
dependent also. Active thermal management strategies are discussed in detail in Jukes 
(2004 - Chapter 5). 

1 AMLCD Sourcing 

Airborne AMLCD displays have fairly obviously sought to retain the pre-existent electro- 
mechanical and CRT instrument sizes of 3ATI, 5ATI, ARINC C (6.25 in square) and ARINC 
D (6.7 in square). These are all unique to the aerospace industry. However, the critical 
performance parameters of a high-quality AMLCD are heavily dependent on good process 
control. Nevertheless: 

1. Good process control only comes with volume manufacturing experience. 

2. Volume manufacture is not compatible with aerospace glass requirements. 

3. AMLCD high- volume manufacturers are all in the Far East. 

4. Volume glass sizes are aimed at the laptop PC market. 

By bundling the whole aerospace industry requirements together, it has been possible to 
convince a few Far East AMLCD glass foundries to process high- volume batches of bare 
glass cells in common aerospace sizes and hold them in a benign environment to be custom 
ruggedised for each specific aerospace application later. 
The custom ruggedisation process typically comprises: 

• Bonding of a front anti-reflective cover glass; 

• Bonding of a rear indium tin oxide (ITO) coated heater glass; 

• TAB attachment of row and column driver printed wiring boards; 

• Mounting of the end product in a rugged frame; 

• Compensation for performance over the intended operating temperature range. 

1 AMLCD MFD: Principles of Operation 

The block diagram of a typical colour AMLCD technology based MFD is shown in 
Figure 11.52. 

The functional elements of an integrated display unit (IDU) comprise: 



A/C Power 

Low Voltage PSU 

Backlight Driver 










row column 

AMLCD Interface 

Applications Processor 

Figure 11.52 Active matrix liquid crystal display MFD block diagram. 

• The active matrix liquid crystal (AMLCD) display head assembly (DHA); 

• Backlight and associated backlight driver; 

• AMLCD interface; 

• Video processor; 

• Graphics processor;* 

• Input/output (I/O) interface;* 

• Applications processor;* 

• Keypanel; 

• Chassis, power supply and interconnect. 

Note that a dumb AMLCD MPD without integral graphics processing omits the functional 
elements marked with an asterisk. Integrated Display Unit 

A fully integrated AMLCD multifunction display is shown in Figure 11.53 (with another 
example given in Plate 7). This unit is intended for dual-use operation in rotary- wing, civil and 
military applications. Although shown with a 6.25 x 6.25 in square format AMLCD, the unit 
is a modular construction and has been designed to accommodate other AMLCD glass sizes 



Figure 11.53 Integrated display unit (Smiths). 

by replacing the front display-head assembly module with 10.4 in and 15 in rectangular 
format COTS AMLCDs. The cardinal point performance specification is set out in Table 11.2. 

11.6 Emerging Display Technologies 

1 1.6. 1 Microdisplay Technologies 

Microminiature LCDs and DMDs (digital micromirrors) have recently revolutionised digital 
projector technology. LCD/DMD projectors produced for the consumer market can fit in a 
fraction of a cubic foot, cost a thousand dollars and can illuminate a large viewing screen in 
daylight. These technological breakthroughs in the commercial projector market are at the 

Table 11.2 Cardinal point performance specification of the AMLCD 

Usable screen area 


Luminance (maximum) 
Luminance (minimum) 
NVG compatibility 
Viewing angle 
Interfaces (video) 

Interfaces (digital) 

6.25 x 6.25 in 

10.4 in and 15 in options 

768 x 768 (XVGA) stripe pixels 

256 per colour, 16 million colours 

350ft.L white (maximum) 

0.05 ft.L (dimming range > 10000:1) 

NVIS class B 

±55° horizontal 

525 line 30:60 Hz 

625 line 25:50 Hz 

ARINC 429 and MIL-STD-1553B 

Integral anti-aliased COTS graphics processor 

Integral application processor 









Figure 11.54 Microdisplay technologies. 

forefront of candidates for the development of new, cost-effective and robust airborne 
displays (Tisdale and Billings, 2001). 

There are currently three microdisplay technologies employed in display projectors, and 
these are shown in Figure 11.54. [For a more detailed description of these technologies, see 
Jukes (2004 - Chapter 9).] 

1 1. 6.2 High-intensity Light Sources 

What makes the CRT so useful, and therefore so difficult to replace, is that it functions as 
both an image modulator and a light source. Although the LCD and DMD have proved to be 
much more reliable image modulators than the CRT, there is still a need to find a viable new 
light source. Typical projector applications employ high-intensity arc lamps, but there are 
many obstacles to overcome in adapting arc lamps to avionics applications. Emerging 
alternative technology light sources are in development, such as light-emitting diodes 
(LEDs) and high-frequency fluorescent lamps. 

11.6.3 Transmissive LCD 

Transmissive LCDs are the dominant display device for low-end conference-room projec- 
tors. The operating principle of the transmissive LCD microdisplay is very similar to the 
direct-view AMLCD except that the device is of course much smaller and designed for 
projection applications. A typical T-LCD is shown in Figure 11.55 and can be thought of as 
the electronic equivalent of a 35 mm transparency. It is illuminated by a very intense light 

11.6.4 Reflective LCD 

Reflective LCDs, also known as liquid crystal on silicon (LCoS) displays, are rapidly 
emerging as the technology of choice for medium-end applications. In a reflective display, 
the light source is positioned so that light reflects from the LCD on to the projection screen. 



Figure 11.55 Typical transmissive LCD device. 

These devices utilise a backplane of conventional crystalline silicon technology that is 
aluminised to increase the reflectivity. An LCD material is applied and sealed with a cover 

For a monochrome display (as in an HUD or an HMD) only a single device is required, 
illuminated by a monochromatic light source. For a full-colour display, typically three LCoS 
devices are employed, with beam- splitting optics to generate three R:G:B images that are 
recombined and projected to the screen. The spectrum of a bright, white light source is split 
into its three primary components and illuminates three reflective micro-LCDs. The three 
images are recombined and, through a projector lens, are focused on to a rear view screen. 
The white light source is typically a mercury arc lamp. Dimming has to be achieved by 
mechanical/optical shuttering in the light path. The lamp will dissipate about 60 W, so forced 
air cooling is required. The functional principles of a reflective LCD projector are shown in 
Figure 11.56. 

Reflective LCD (or LCoS) devices themselves are multisourced and are in widespread use 
in tabletop projectors. The imaging device, shown in Figure 11.57, is essentially semi- 
conductor wafer technology using LC principles to polarise reflected light through a simply 
attached liquid crystal cell placed over the device. 

1 1.6.5 Digital Micromirror Device 

The DMD is the primary competitor to the LCD for the image-modulating component of a 
projector, appearing more widely in high-end applications. 

Because of the improved reflection efficiency and high switching speeds of DMD 
compared with LCoS, it is possible to build a compact projector with a single DMD device, 
obtaining full-colour rendition using sequential colour processing. This arrangement elim- 





Po la riser 

Figure 11.56 Reflective LCD projector (Kaiser Electronics). 

inates the need for a beam splitter and the subsequent problems of alignment/reconvergence 
of the three colour images. The result is a significantly clearer image than LCoS technology, 
and in smaller volume. 

The development of the digital micromirror device (DMD), by Texas Instruments, has 
yielded a robust, reliable, high-performance device which can now be packaged with other 
components to produce a range of display solutions. This resultant digital light projector 
(DLP) brings together the DMD with a light source, optics, colour filters and a projection 



Cover Glass 


Liquid Crystal 

Spacer Layei 


Silicon Substrate 

Planarised Mirrors 
Figure 11.57 Reflective LCD (LCoS) device structure. 

CMOS Circuitry 



Mirror Address 

Support v 

r \ 

Post XfflOtL-^i 



^Yoke Address 

Via 2 Contact 

Figure 11.58 DMD array (Texas Instruments). 

lens to produce an image on a display surface (i.e. rear projected on to a screen). The 
arrangement is now widely used for PC and cinema projection systems (Hornbeck). 

The DMD is an array of aluminium micromirrors as shown in Figure 11.58. These are 
monolithically fabricated over an array of CMOS random-access memory cells, each of 
which corresponds to a micromirror. This allows each mirror to be individually addressed, 
causing it to tilt by approximately 10°, limited by a mechanical stop. 

A bright light source illuminates the DMD array. Depending on the state of each 
micromirror, light is either reflected towards the projector objective lens or towards a light 
absorber on the wall of the projector. 

Full-colour operation is obtained by interposing a colour wheel comprising red, green and 
blue colour filters between the light source and the DMD device, as shown in Figure 11.59. 

DLP Board _^^fltffl 

v.. .*^M 

^ Projection 

^ """" Lens A 

p. = 

J - ' 

\ /^S^^- Optics 
Color Filter ^^ 

^ s Light Source 

Figure 11.59 DMD projector (Texas Instruments). 




V ^ LAMP/ 



Polarization Recovery 



RGB Image Modulation 
Color Recombination 



Image Magnification 



Uniformity Enhancement 
Viewing Cone Management 
Constrast Enhancement 

Figure 11.60 Projector optical elements. 

Complex algorithms in the associated processor/DLP chip set provide signals to each DMD 
pixel to produce a full-colour image with at least 256 grey levels per colour by suitable mark- 
space modulation of each pixel in synchronism with the rotation of the colour wheel. 

11.6.6 Rear-projection 'Big Picture' Head-down Display 

Rear projection is a possible alternative technology to direct- view AMLCD, able to provide 
large-area head-down instrument displays. Rear projection has the advantage of being 
independent of the size vulnerability of using COTS direct- view displays. 

The technology is scalable and obsolescence proof. A single optical engine design 
combined with slightly modified folded projection optics can be used for a variety of 
display sizes from, say, 4 in square to 32 in diagonal. As the commercial market drives 
higher-resolution and more efficient display devices (LCoS or DMD), improved image 
generation components can be incorporated without major redesign. There are no single 
high-cost components in the product design. 

The basic optical elements of a projection system are illustrated in Figure 11.60 and 

1. Illumination assembly. Every projection system using a non-emissive image source must 
have a source of illumination that contains sufficient energy in the red, green and blue 
wavelengths. Conference-room projection systems most commonly employ high-pressure 
mercury lamps. Other lamp technologies are in use (e.g. xenon and metal halide) but do 
not offer efficiency equivalent to the mercury lamps. The lamp should have the 
characteristics of a point source: that is, the light should emit from as small an area as 
possible. Collimation and shaping are performed to create a beam of near-parallel rays 
that will illuminate an area just slightly larger than the display devices. 

2. Image engine assembly. This is the heart of any projection system. While numerous 
architectures are possible, the choice of display device technology is the single most 
important influence in architecture selection. Options identified earlier are: 

• Transmissive LCD; 

• Reflective LCD or LCoS; 

• Digital micromirror (DMD). 

3. Projection optics assembly. This magnifies the image and provides a focused image at the 
rear of the screen, free of chromatic aberration, distortion and other optical defects. 

4. Screen assembly. The screen assembly plays a significant role in presenting the pilot with 
a high-quality image that is viewable in the luminous environment of the cockpit. For 



Figure 11.61 Rear-projection head-down display (Kaiser Electronics and Brilliant Technology). 

high-performance avionic applications, when compared with commercial applications, 
the screen must provide: 

• Improved uniformity; 

• Shape and steer of the viewing cone; 

• Enhanced contrast. 

Figure 11.61 shows an open-box view of a typical rear-projection head-down display. The 
lower section of the box contains the image engine and lamp. The projection lens is at the 
rear of the box. 

Multiple projectors can be arranged to fabricate a 'big picture' display encompassing the 
whole of the instrument panel area. The potential redundancy of this architecture makes it 
attractive. Overlapping portions of the optical engines achieve the expanded width of the 

Rear projection is the chosen technology for the joint strike fighter (JSF) crew station, 
using a multiple-projector arrangement to achieve a 20 in x 8 in 'big picture' display as 
shown in Figure 11.62 and Plate 8. 

11.6.7 Solid-state Helmet-Mounted Display 

Current CRT technology HMDs impose large design burdens on the engineer. They are 
bulky and heavy and require high voltage and expensive quick disconnect interfaces. 

The next generation of HMDs is in development using solid-state technology that will 
provide greater performance and lower cost of ownership owing to high reliability. Most 
importantly, they reduce helmet size and head-borne weight. 



Figure 11.62 Multiple projector 'big picture' HDD (Kaiser Electronics). 

Figure 11.63 shows the HMD in development for the joint strike fighter (JSF). 

The image sources are two 1280 x 1024 transmissive AMLCDs. They achieve a contrast 
ratio of 100:1 with a response time of <9ms. T-LCDs typically have a transmission of 
approximately 10%. A small LED backlight has been developed that can produce greater 
than 26000ft.L light output which is sufficient to achieve good readability in a lOOOOft.C 
(108 000 lux) day environment. The LED requires around 0.8 mA and dissipates approxi- 

Figure 11.63 JSF helmet-mounted display (Kaiser Electronics). 



Figure 11.64 JSF HMD optical arrangement (Kaiser Electronics). 

mately 1.5 W per T-LCD, a significant achievement. The electrical power requirements in the 
region of the pilot's head have decreased from approximately 13.5 kV required by the CRTs 
to only 5 V (Casey, 2002). 

The use of T-LCDs allows the use of a smaller optical chain, shown in Figure 11.64, 
resulting in less complex optics, reduced weight and a more compact design, which has a 
minimum impact on the helmet profile. 

The remote display processor receives incoming video from the sensor suite, digitally 
processes and overlays the video signals and performs helmet tracking and input/output (I/O) 
control functions. A real-time image- warping engine performs dynamic scaling and warping 
of high-resolution video data. Dynamic image warping is required to compensate for the 
geometric distortion of an off-axis optical design. 

11.6.8 Organic Light-emitting Diodes (OLEDs) 

The operation of an LED is based upon the junction of p-type and n-type materials. When a 
voltage is applied, electrons flow into the p-type material, and holes flow into the n-type 
material. An electron-hole combination is unstable; they recombine and release energy in 
the form of light. This can be a very efficient process. 

Light-emitting diodes, based upon semiconductors such as gallium arsenide, have been 
around since the late 1950s. These crystalline LEDs are expensive, and it is difficult to 
integrate them into small high-resolution displays. However, there is a class of organic 
compounds that have many of the characteristics of semiconductors in which p-type and 


Figure 11.65 OLED microdisplay (Cambridge Display Technology). 

n-type organic materials can be introduced to make light-emitting diodes. [For a more 
detailed description, see Jukes (2004 - Chapter 9)]. 

While the early OLEDs did not have sufficient efficiency or life to be commercially 
attractive, significant progress is being made to improve these factors. 

The advantage of these polymer devices is the ability to spin on the layers and, in some 
cases, to pattern the films with photolithography. An alternative approach is to employ an 
active matrix using standard semiconductor techniques. This is eminently suited to making 
microdisplays, because a small silicon chip can be used as the substrate and the necessary 
driver circuits can be incorporated into the silicon chip along with the matrix structure 

A typical OLED microdisplay device is shown in Figure 11.65. 

OLED technology is now emerging to be of benefit for both direct- view and microdisplay 
applications. OLEDs offer higher efficiency and lower weight than liquid crystal displays 
since they do not require backlights or reflective light sources (Howard). 

1 1.6.9 Virtual Retinal Displays 

The virtual retinal display (VRD) offers the potential for a virtual cockpit. The VRD paints 
an image directly on to the wearer's retina using a modulated, low-power beam of laser 
lights. Operating at extremely low power, there is no danger to the eye. The wearer sees a 
large, full-motion image without the need for a screen. [For a more detailed description see 
Jukes (2004 - Chapter 9)]. 

To create the image, the VRD uses a photon source (or three for a full-colour display) to 
generate a coherent beam of light. The only required components are the photon sources, the 
scanner and the optical projection system. Scanning is accomplished with a small micro- 
machined electro-mechanical scanner (Figure 11.66). The projection optics are incorporated 
into the front reflecting surface of a pair of glasses (Collins, 2003). 

A prototype helmet-mounted VRD display (Figure 11.67) has been developed for military 
rotorcraft applications. 



Figure 11.66 Virtual retinal display components (Micro vision Inc.). 

Figure 11.67 Prototype VRD helmet for rotorcraft applications (Micro vision Inc.). 

11.7 Visibility Requirements 

11.7.1 Military Requirements 

The requirements for legibility and readability of military combat aircraft displays can be 
found in the following documents: 






Joint service specification guide - crew systems - cockpit/crew- 
station/cabin handbook 
Joint service specification guide - crew systems - aircraft lighting 

Electronically/optically generated airborne displays 
Lighting, aircraft, night- vision imaging system (NVIS) compatible 
Cockpit lighting standard (NATO restricted document) 




Figure 11.68 HDD high ambient - sun rear scenario. 

The real-world natural ambient illumination is both too variable and too complex for every 
condition to be considered in the design of cockpit displays. However, it is useful to identify 
a few worst-case scenarios. Head-down Display: High Ambient - Sun Rear 

This scenario (Figure 11.68) is the high ambient illumination condition that degrades the 
visibility of emissive displays by 'washing out' the presented information. The condition is 
experienced when: 

1. The aircraft is flying straight and level at 30000 ft above 8/8 cloud. 

2. The solar disc is low (30° elevation) and to the rear of the aircraft. 

3. The display is bathed in direct sunlight - around 100 000 lux. 

4. Display specular reflections of the general cockpit area predominate. 

5. The forward ambient scene is diffused - around 15 000 lux. 

6. The aircrew helmet tinted visor is down. Head-down Display: High Ambient - Sun Forward 

This scenario (Figure 11.69) arises when the principle effect of the ambient illumination is to 
produce 'solar glare' which degrades the pilot's perceptual capability. The condition is 
experienced when: 



Figure 11.69 HDD high ambient - sun forward scenario. 

1. The aircraft is flying straight and level at 30000 ft above 8/8 cloud. 

2. The solar disc to very low (15° elevation) and forward of the aircraft. 

3. The display is in shadow and illuminated by diffused skylight - around 8 000 lux. 

4. Display specular reflections of the pilot's flying suit predominate. 

5. The aircrew helmet visor is down and the forward scene is dominated by the solar disc - 
around 110 000 lux. 

1 1. 7. 1.3 Head-Up Display and Helmet-Mounted Display: High Ambient - 


Practical flight experience indicates that the worst-case ambient illumination condition for 
the HUD transmissive display visibility is when the display symbology is presented against a 
cloud face, with a clear high luminance solar disc producing additional glare just outside the 
display field of view. 

This scenario (shown in Figure 11.70) is as follows: 

1. The aircraft is flying straight and level at 30000 ft above 8/8 cloud. 

2. The solar disc is within 15° of the display central vision line. 

3. The display field of view is dominated by an illuminated cloud face - 

4. The aircrew helmet visor is down. 

around 25 000 cd/m z . 



Figure 11.70 HUD/HMD high ambient - sun forward scenario. 

The ability of the crew to detect information presented on the HUD or HMD will again depend 
on the visual difference or contrast between the foreground image and the background. In this 
case the foreground image is the HUD symbology reflected off the display combiner. The 
background is forward ambient scene illumination modified by the combiner optics. As before, 
there is an element of forward ambient veiling glare, which reduces the crew's perceptual 
capability through the HUD/HMD. All of the above modified by the visor, if down. 

1 1. 7. 1.4 Low Ambient - Dusk/Dawn Transition 

The final scenario is the low dusk/dawn transient illumination condition, when the sun is 
forward of the aircraft and close to the horizon. The clear air conditions result in the cockpit 
being in hard shadow while the pilot is subject to solar glare. Reflective displays are most 
affected, because the limited solar glare is sufficient to degrade the pilot's perception of the 
display low reflected luminance. This scenario is as follows: 

1. The aircraft is flying straight and level at 30000 ft in clear air. 

2. The solar disc is forward of the aircraft close to the horizon line - 
forward ambient. 

3. The sky hemisphere is clear and of low luminosity - around 200 lux. 

4. The cockpit lighting control 'Night' is selected and set to maximum. 

5. The aircrew helmet ND visor is up. 

around 1 500 lux Night 

In practice, night ambient illumination is not a single condition but the summation of 
multiple and complex radiation sets based on lunar reflected sunlight, various sky-glow 
derivatives and individual starlight spectra: 

1. Night illumination can be assumed to consist of a diffused illumination over the range 
1.0 -0.0001 lux. 

2. The aircrew may be flying night adapted without any vision aids. 


3. Alternatively the crew may be using night- vision goggles to aid perception of the external 

The ability of the crew to detect information presented on the cockpit head-down displays 
will depend on the visual difference or contrast between the foreground image and its 
background. In practice there are three display components, two resulting from the ambient 
illumination plus a component of veiling glare resulting at the eye, all modified by the visor 
transmission, if down. These are: 

• Display emission; 

• Display diffuse and specular reflection; 

• Veiling glare. 

For further description of these terms, see Jukes (2004 - Chapter 12). 

11.7.2 US DoD Definitions and Requirements 

The requirements identified in the above referenced US DoD documents embrace the 
legibility and readability requirements for electronic and electro-optical display contrast and 
luminance in a single combined environment of both diffuse light [producing 108 000 lux 
(lOOOOft.C) on the display face] and the specular reflection of a glare source [with a 
luminance of 6800 cd/m 2 (2000 ft.L)]. 

There is also a 'minimum luminance difference' requirement, which basically requires 
displays to have a high luminance in addition to achieving adequate contrast to combat 
veiling glare. 

The combined diffuse and specular environment is intended accurately to simulate the 
lighting conditions in a fighter crew station in direct sunshine. The 108 000 lux (lOOOOft.C) 
diffuse requirement represents sunlight of 130 000 lux (12000ft.C) to 160 000 lux 
(15 000ft.C) (outside sunshine ambient at high altitude) passing through an aircraft canopy 
(typically 80 - 90% transmission) and striking a display somewhat off-axis. The 6800 cd/m 2 
(2000 ft.L) glare source represents a reflection of the sun from interior parts of the crew 
station or from the pilot's flight suit or helmet. 

The visibility requirements, under the prescribed conditions, are shown in Table 11.3. 

11.7.3 European (Eurofighter Typhoon) Definitions and Requirements 

The requirements identified in the Eurofighter Typhoon document describe the legibility and 
readability requirements for electronic and electro-optical display contrast and luminance in 

Table 11.3 MIL-STD-85762 Table II 

Information type 

Luminance difference (min) 

Contrast (min) 

Numeric only 



100 ft.L (343 cd/m 2 ) 


Graphics plus alphanumeric s 


Video (high ambient) (six grey shades) 

160 ft.L (550 cd/m 2 ) 


Video (dark ambient) (eight grey shades) 




a comprehensive but complex manner using the concept of perceived just noticeable 
differences (PJNDs). The concept is based on the premise that the ability of the crew to 
detect information presented on a cockpit display will depend on the visual difference 
between the foreground image and its background. 

The PJND values for a particular display device can be computed from three sets of 

• The display measured performance characteristics; 

• The 'worst-case' ambient lighting conditions applicable to that display; 

• A set of perception equations that represent a 'standard pilot's eye'. 

For further description, see Jukes (2004 - Chapter 12). 

Although technically elegant, and a viable analysis of product performance during product 
formal qualification testing, it is impractical to perform this level of testing on a 100% basis 
on series-production articles. 

1 1. 7.4 Viewability Examples AMLCD Head-down Display 

The simplified worst-case scenario is shown in Figure 11.71. The cardinal point performance 
specification is set out in Table 11.4. 

1 0,000 fc 
(108,000 lux) 


4% transmission 

40:1 contrast ratio 

1% diffuse reflectance 

Figure 11.71 AMLCD HDD visual performance. 


Table 11.4 Cardinal point performance specification of the AMLCD head-down display 

Active area 6 in square 

Backlight luminance 10 000 ft.L white (34 000 cd/m 2 ) 

Transmission 4% 

Contrast ratio 40: 1 

Diffuse reflectance 1% 

Specular reflectance 1 % 

Ambient illumination 1 000 ft.C diffuse sunlight ( 1 08 000 lux) 

2 000 ft.L point source (6 800 cd/m 2 ) 

The display background is made up of a diffuse component of sunlight reflected off the 
AMLCD black matrix, plus a specular component which is reflected off the front surface of 
the AMLCD, plus the transmission of the 'off (black) pixels. The display foreground is the 
AMLCD transmission of the 'on' (red, green and blue) pixels: 

Background luminance = (10000 x 1%) + (2000 x 1%) + (10000 x 4%)/40 
= 130 ft.L (450 cd/m 2 ) 
Display luminance = (10000 x 4%) = 400 ft.L white (1370 cd/m 2 ) 
Contrast ratio = 1 + display foreground/background 
= 1 + (400/130) =4.1 : 1 

1 1. 7.4.2 Head-Up Display 

The simplified worst-case scenario is shown in Figure 11.72. The cardinal point performance 
specification is set out in Table 11.5. 

The display background is the illuminated cloud component attenuated by the combiner 
transmission. The display foreground is the CRT emission attenuated by the relay optics 
transmission and the HUD combiner reflectance at the wavelength of the CRT phosphor: 

Background luminance = (7200 x 90%) = 6480 ft.L 

Display luminance = (10000 x 70% x 25%) = 1750 ft.L (6000 cd/m 2 ) 
Contrast ratio = 1 + display foreground/background 
= 1 + (1750/6480) = 1.27: 1 Night-vision Imaging System Compatibility 

Night-vision goggles (NVGs), also called the night-vision imaging system (NVIS), are 
passive, helmet-mounted, binocular image intensification devices. The NVIS operates by 
converting photons of the outside night scene into electrons using a gallium arsenide 
photocathode (Figure 11.73). The photocathode releases one electron for every photon 
it receives, thereby converting the light energy to electrical energy. The electrons are 
multiplied by passing through a wafer-thin microchannel plate, which is coated to 
cause secondary electron emissions that are accelerated by the electric field and finally 
collide with a phosphor screen. The phosphor screen then converts the electrons back into 



90% transmission 
25% reflection 

6,480 fl_ 

Foreground Emission — 

1 ,750 fl_ 
(6,000 cd/rrf ) 

Figure 11.72 HUD visual performance. 

photons, displaying the image, which now is amplified around 2000 times. The image is 
shown in green. 

For aviation use, the goggles are mounted on the front portion of the helmet and hang 
down in front of the pilot's eyes. Figure 11.74 shows one type of NVIS in which the image- 
intensified image is projected into the pilots' line of sight through a prismatic combiner. The 
viewing eyepiece sits about 20 mm in front of the eyes, which enables the pilot to view and 
scan the outside world by looking at the image, and also enables the pilot to see all the 
cockpit instruments and displays by looking underneath the goggles. 

To achieve compatibility of the crew- station lighting with the NVIS, the crew- station 
lighting should have a spectral radiance with little or no overlap into the spectral response of 
the NVIS image intensifier tubes. Figure 11.75 illustrates the requirements of NVIS- 
compatible crew- station lighting. 

The rationale and criteria for achieving compatibility of crew-station instruments, displays 
and lighting with night-vision goggles is encapsulated in the US DoD document MIL-STD- 
3009 which is derived from the earlier MIL-L-85762A. 

Table 11.5 Cardinal point performance specification of the HUD 

Bare CRT stroke luminance 
Relay optics efficiency 
Combiner reflectance 
Combiner daylight transmission 
Ambient illumination 

10000ft.L(34 000cd/m 2 ) 


25% tuned to CRT phosphor wavelength 


7 300ft.L (25 000cd/m 2 ) illuminated cloud face 




Channel Plate 
Scene Image Multiplier 

Objective Lens 

High Voltage 
Power Source 

Viewing Screen 


Intensified Image 


Figure 11.73 Diagram of an image intensifies 

To achieve compatibility, both the NVIS and the displays are fitted with complementary 
filters. Three filter classes are generally accepted for the NVIS: 

1 . The class A filter maximises the NVIS sensitivity but only allows blue, green and yellow 
lights to be used in the crew station. Red cannot be used in the crew station because the 

Figure 11.74 Cat's eyes NVIS. 



Cockpit Lighting 
Spectral Emission 

NVIS Spectral 

550nm 625nm 


Figure 11.75 Conceptual diagram of the spectral distribution of NVIS -compatible lighting. 

class A NVIS is extremely sensitive to colours with radiance at wavelengths longer than 
600 nm. 

2. The class B filter was developed primarily to allow three-colour CRTs to be used with the 
NVIS. Note that the red acceptable for use with class B is not a 'deep red', as might be 
expected in a full-colour display, but more an orange red. 

3. The class C filter was introduced to make HUD symbology visible through the NVIS by 
incorporating a 'notch' or 'leak' in the green part of the spectrum. 

For further description of the NVIS and NVIS compatibility, see Jukes (2004 - Chapter 12). 


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