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OCKET
AMJET
EADER

CHEMICAL SYSTEMS DIVISION

CSD's Modern Versatile Ramjet Facility Located at Coyote,

California

CORNELL UNIVERSrTY LIBRARY

04 032 820

THE POCKET RAMJET READER

Cover Design - Reproduction of the Figure from the German Patent
Issued to Albert Fono in 1928 for a Ramjet Engine

CHEMICAL SYSTEMS DIVISION

^\i/<^.

UNITED
TECHNOLOGIES.

Foreword

High technology discipUnes almost always bristle with the specialized jargon peculiar
to their field. Although an understanding of the jargon can be obtained from textbooks
and treatises, in many cases this may be both difficult and time consuming. It is a purpose
of this booklet to provide in one source a basic, simplified explanation of the terms,
elements, and operating parameters of ramjet technology. Armed with the basic
information contained herein, the reader should be able to participate knowledgeably in
discussions, presentations, and other business activities that involve ramjet propulsion
systems. He should also be able to use this information as a basis for extending his
knowledge of this complex and challenging field with more advanced technical data.

Portions of the material contained herein have been obtained from the following
sources:

Twenty-Five Years of Ramjet Development, William H. Avery, Jet Propulsion, Vol.

25, No. 11, November 1955, pp 604-614

Aircraft and Missile Propulsion, Vols. I and II, M. J. Zucrow, John Wiley & Son, Inc.,

1958

Aircraft Propulsion, P. J. McMahon, Harper & Row Publishers, Inc., 1971

These references provide excellent treatments of ramjet technology and are
recommended to those readers who desire a more detailed and comprehensive
understanding of the subject matter.

CONTENTS

1. HISTORICAL PERSPECTIVE 1

2. GENERAL CHARACTERISTICS OF RAMJETS 4

3. AIR INDUCTION SYSTEMS 19

4. FUEL MANAGEMENT SYSTEMS 29

5. LIQUID-FUELED RAMBURNERS 32

6. SOLID-FUELED RAMBURNERS 39

7. BOOSTERS 43

8. RAMJET TESTING 46

9. APPLICATIONS FOR INTEGRAL ROCKET RAMJETS 51

10. CSD, A LEADER IN RAMJET PROPULSION 54

Artist's Rendition of Air
Intercept Missile (Top),
Advanced Anti-radiation
Missile (Center), and
Advanced Long Range
Air-to-air Missile.

Historical Perspective

Origin of the Ramjet

The concept of the ramjet engine is
attributed to a Frenchman, Rene Lorin,
who first described such a device in 1913.
Since he did not envision flight at
supersonic speeds, his analysis of ramjet
propulsion was based on propelling
bodies at subsonic speeds, and he
concluded that the ramjet engine would
have a low thermal efficiency. A British
patent issued in 1926 discloses the
application of two ramjet-like devices for
propelling artillery shells, but there is no
evidence that the devices were ever built.

The first patent (German) that disclosed
the use of ramjet engine as a propulsion
system for supersonic flight was issued to
Albert Fono in 1928. Again, there is no
evidence that his engine was ever built.
As a resuh of work begun around 1933,
a French patent was issued to Rene Leduc
in 1935 on the design of a ramjet-
propelled airplane. By 1935 Leduc had
tested the thrust of a small unit at speeds
up to 679 mph, and the results were so
encouraging that the Air Ministry
authorized design and construction of a
research airplane to be propelled by a
ramjet engine. As conceived by Leduc,

Ramjet-propelled Airplane as Conceived by Leduc

atmospheric air entered an annular scoop
surrounding the cabin, the air
temperature was raised by fuel burners
located near the midsection of the
airplane, and hot gases were discharged
at the rear with a large jet velocity. Active
work on the airplane was begun in 1938,
and during 1939 several engine
components were tested over the Mach
number range 1.65 to 2.35. The
development program was interrupted by
World War II, but the experimental
airplane was completed and subsonic
flight tests were conducted in 1949.

During World War II a great deal of
effort was expended on ramjet engine
development in Germany, Great Britain,
and the United States. The Germans
studied the application of the ramjet
engine to fighter aircraft, conducted tests
of subsonic ramjet engines, and
investigated the possibilities of soHd fuels
and suspensions of metals (such as
aluminum in fuel oil). The Germans
made extensive studies on supersonic
spike-type diffusers; much of this
research was in connection with applying
ramjets to the propulsion of artillery
shells.

British work on the ramjet engine
(originally termed the Athodyd, for
aerothermodynamic duct) began during
World War II but was largely confined to
theory. Some experimental work was
performed on small engines primarily for
missile propulsion.

In the United States the potential of the
ramjet engine was first pointed out in
1 94 1 , but not until 1 944 was serious effort
expended on ramjet engines for
supersonic propulsion. The first work of

consequence began at the Applied
Physics Laboratory of Johns Hopkins
University under sponsorship of the
Navy's Bureau of Ordnance. The first
flight tests in 1945, using a 6-inch
diameter engine burning heptane, are
probably the first experimental
demonstration of the acceleration of a
ramjet engine in supersonic flight. These
early flight experiments were conducted
at low altitudes with the most
rudimentary fuel controls to maintain
maximum thrust output. Consequently,
many of the problems associated with
flight at high altitude and with flight
maneuvers were not encountered. The
encouraging results of these low-altitude
tests interested the military services in
supporting development programs for
several ramjet engines to propel
supersonic guided missiles. Thus, the
Navajo long range missile, the Bomarc
interceptor missile, and the Talos missile
were designed with ramjet propulsion
systems.

The Ramjet Recession

Despite these early systems, some of
which progressed into full operational
use, interest in ramjets abated, mostly
following cancellation of the Navajo
strategic missile program in favor of the
Atlas ballistic missile. One of the major
considerations in choosing Atlas as the
nation's first intercontinental missile was
that it followed a pure ballistic trajectory
after the conclusion of powered flight
some six minutes after launch. The
Navajo, on the other hand, employed a
large liquid rocket booster (almost

identical with the engine employed for
Atlas) to reach supersonic speed and
thereafter had to operate its ramjet
engines for several hours to achieve its
intended range. For the intercontinental
missile, therefore, a ballistic trajectory
was preferable. There were also serious
technical questions about rehability of
the airborne power supply operated by a
hot gas turbine, on which success of the
mission depended critically.

For nearly two decades interest in
ramjets was limited essentially to
research; no new ramjet propulsion
systems were specified for operational
vehicles. Meanwhile, both solid and
liquid rockets continued as primary
power sources for launch vehicles and
tactical missiles, attaining an advanced
state of development in the process.

There were two main reasons for the
lack of interest in ramjets. First, the
earlier, more modest requirements for
tactical missiles could be met by solid
propulsion systems. Second, even though
the airbreathing ramjet offered much
higher performance than a rocket with its
self-contained source of oxygen, the
ramjet had to be boosted to supersonic
velocity before it could operate. This
requirement made necessary a separate
rocket for the boost phase, incurring the
increased complexity of two separate
propulsion systems.

The Ramjet Resurgence

The recent revival of ramjet propulsion

stems from changes both in technology
and in tactical missile requirements. The
significant technology change was
emergence of the int egral rocket ra mjet,
which combines the rocket boost and
ramjet sustain functions in one efficient
propulsion system.

Tactical missile requirements must be
responsive to the international threat
environment, which has changed because
of the ability to detect launch platforms at
longer range. Because of this improved
ability, missiles must be launched much
further from their targets. They must be
capable of longer range and higher speeds
at all altitudes and in many cases be under
power all the way to the target. To obtain
these capabilities in a missile of minimum
volume requires the high performance of
a low cost airbreathing propulsion
system. The IRR fulfills these
requirements effectively. Over the next
five to 10 years, therefore, IRRs are
expected to form one of the major
propulsion systems in the nation's
arsenal.

Designing and testing IRRs requires a
wider range of skills than rockets, and the
test facilities need to be much more
elaborate. Nevertheless, IRRs offer such
advantageous operating characteristics
that they are now being specified as
propulsion systems for several next-
generation vehicles. It seems clear that
the day of the ramjet has finally arrived,
and that the IRR will become a
prominent member of the family of jet
propulsion systems during the 1980's.

General Characteristics of Ramjets

The ramjet engine is one of the
youngest of the family of jet propulsion
devices that includes the rocket and the
turbojet. Airbreathing ramjets give much
higher fuel efficiency than rockets since
ramjets use inlet air as a source of oxygen.
Self-sufficient rockets, on the other
hand, must carry their own oxidizer and
bear the consequent weight penalty.
Accordingly, although rockets must be
chosen for propulsion outside the earth's
atmosphere, ramjets generally
outperform rockets if there is a ready
supply of air.

Because the ramjet depends only on its
forward motion at supersonic speeds to
effectively compress intake air, the engine
itself employs no moving parts. It is
therefore capable of a simplicity,
lightness of construction, and high flight
speed not possible in other air-breathing
engines. These features, plus the high
thermal efficiency it can achieve, make
the ramjet a particularly attractive choice
for propelling vehicles at supersonic
speeds.

One other significant difference
between rockets and ramjets is thrust at
zero speed. Rockets can deliver thrust at
any speed, even standing still, whereas a

ramjet requires an auxiliary boost system
to accelerate it to its supersonic operating
regime so that its forward motion can
compress the inlet air. To operate at
practical efficiency a ramjet must be
moving at about Mach 1 .5 or greater so
that the margin of thrust over drag will be
satisfactory.

MACH NUMBER

The Mach number is the ratio of the speed
of a body with respect to a surrounding fluid
(such as air) to the speed of sound in the
fluid. An aircraft travelling at Mach 2 is
moving twice as fast as the local speed of
sound. The Mach number may also be the
ratio of the speed of the fluid to the speed of
sound. A stream of air exiting from a ramjet
diffuser may be moving at Mach 0.2 or
about 150 miles per hour.

Ramjet Components

The basic ramjet engine consists of an
air inlet or diffuser, a combustor, and an
exhaust nozzle. The diffuser admits air to
the engine, reduces the air velocity, and
develops ram pressure. The combustor
adds heat and mass to the compressed air
by burning a fuel. The nozzle converts

Basic Components of Ramjet Engine

some of the thermal energy of the hot
combustion products to kinetic energy to
produce thrust.

Although the operating principle of a
ramjet engine appears simple, both the
equations that define ramjet parameters
and the process by which a ramjet is
designed are much more complex than
those for a soUd propellant rocket.
Everything that happens inside a rocket is
isolated from its external surroundings.
Except for secondary effects of
acceleration, flight maneuvers, or
aerodynamic heating, internal processes
of the rocket are independent of its
environment. In fact, the only things
affected by the atmosphere are vehicle
drag and thrust level (which depends on
the pressure at the nozzle exit plane). The
thrust level can therefore be computed by
equations that essentially depend only on
the internal parameters of the rocket
motor.

By comparison, in a ramjet engine the
thrust level is subject to the dynamic
interaction of several factors, including
the pressure developed in the diffuser,
angle of attack, vehicle velocity, and
ambient pressure or altitude. This
additional complication of operating
parameters means that understanding
how a ramjet works requires a modest
exposure to certain principles of
aerodynamics. Some of these principles

will be touched on in the pages that
follow.

Operating Principle

A discussion of a ramjet engine can be
simplified by assuming that the ramjet is
stationary, and that air approaches the
engine at a velocity equal to the vehicle
speed. As air enters the inlet, adiabatic
compression causes an increase in
temperature and a decrease in velocity.

ADIABATIC COMPRESSION

The reduction of volume of a substance
without heat flow, in or out.

The air is further heated by combustion
of the fuel which also increases the mass
flow, typically between 5 and 10%. The
high-temperature compressed gases then
are expanded in the nozzle and
accelerated to high velocity. The thrust
developed by the engine is the net rate of
change of momentum of the gases passing
through the engine and is equal to the
mass flow rate of the air plus burned fuel
times the jet velocity minus the flow rate
of air times the air velocity. The effective
net thrust on the vehicle will be somewhat
less than the engine thrust because of skin
friction drag on the air flowing around
the ramjet vehicle.

The overall process can be more clearly
understood by examining the thrust
equation. Let m^ be the mass rate of flow
of inlet air and V^ be its velocity. Then
riiaVa is the momentum rate of the inlet
air stream. The exhaust gas is coming out
of the nozzle at a velocity Vg, sometimes

(m^ + mf)V^

m = mass rate of flow of inlet air
a

111. = mass rate of flow of fuel
V„ = velocity of inlet air

V = velocity of exhaust stream (sometimes designated V- or c)
e J

referred to as the jet velocity, Vj. The
momentum rate of the exhaust gas is (iha
+ iTif) -Vg where ihf is the rate of mass
addition in the engine due to burning the
fuel. The thrust of the engine is,
therefore, simply the net rate of change in
momentum at a steady state condition
and is given by:

F = (ifia + mf) Vg - rfiaVa
This equation for change of
momentum in a ramjet bears further
examination, since each term affects the
complex interactions associated with
ramjet operation. First, the exhaust
velocity Vg is identical with the symbol c

commonly used for exhaust velocity of a
rocket engine. In both ramjets and
rockets, c = c*Cp, where c* is the
characteristic exhaust velocity and Cp is
the nozzle thrust coefficient. The value of
c* obtained from theoretical calculations
is a measure of the energy available from
the propellant and depends on combustor
pressure, mixing efficiency, and residence
time. The value obtained under test
conditions thus becomes a measure of
combustion efficiency. Cp depends on
pressure in the combustion chamber
(hence on velocity, altitude, and inlet
efficiency in ramjets) and on nozzle

^* ^ M t t|Lt I t t t I I! t t t

Distribution of Internal Pressure in a Ramjet Engine

RESOLUTION OF FORCES IN A
RAMJET

The force causing the increase in
momentum of the inlet air acts in the same
direction as the air stream and appears as a
pressure drop in the ramjet. Now, the
pressure at any point in the duct is
perpendicular to the surface. Since the
product of pressure and area is force, for
each unit area of surface there is an applied
pressure and a corresponding force. Each of
these force vectors can be resolved into
radial and longitudinal components. The
radial components cancel, but the
longitudinal components are algebraically
additive. The resultant longitudinal force is
directed forward and equals the thrust of the
ramjet, which is numerically equal but
opposite in direction to the force causing the
change in momentum of the air stream.

Unit Area

Cancelling
radial forces

Resolution of Forces for Pressure
Applied to Unit Area of a Duct

configuration.

The term V^ is flight speed, sometimes
expressed in terms of Mach number.
Although the equation shows V^ causing
a decrement in thrust, this negative term
is offset by the term containing m^, irif.

and Vg, all of which are direct functions
of V^. In fact, if V^ is zero, the thrust is
zero. As V^ increases, the term (rh^ +
rhf)Vg increases more rapidly than rhaVa
so thrust increases steadily and usually
becomes maximum in the range of Mach
3 to Mach 5. Thereafter the negative
term begins to dominate, so thrust falls
off.

For these reasons, ramjet thrust
calculations are considerably more
complicated than those for rockets. In
making such calculations it is convenient
to convert the familiar conservation
equations of mass, momentum, and
energy to forms involving Mach number
and to combine them into expressions

STREAM THRUST

A fluid flowing through a conduit is
subject to three forces: the pressure acting
over the bounding end surfaces and the force
exerted by the inner surface of the conduit.
The resultant of these forces is equal to the
rate of change in momentum of the fluid.
For calculations involving the thrust due to
a moving fluid it is convenient to regard the
sum of the pressure-area force and the rate
of change of momentum as a single term,
called the stream thrust. Foriexample, at a
point X the stream thrust may be defined
as:

Fg= •^x-'^o-PoCAj^-Ao)

Fg = gross thrust of the ramjet

^x- stream thrust for the internal flow

at station x
^o ~ stream thrust for the internal flow

at station
Po = ambient pressure
A^ = cross-sectional area at station X
Aq = cross-sectional area at station

PROPERTIES OF THE ATMOSPHERE

The atmosphere that a ramjet engine
encounters over its range of operating
altitudes is quite different from what we
experience near sea level. Air density atop
the tallest mountains (about 35,000 feet) is
only 31 percent of the density at sea level. At
100,000 feet it is less than 2 percent. The
pressure exerted by the atmosphere also
decreases with altitude. From a value of
over 2100 pounds per square foot (14.7
pounds per square inch) at sea level it is
reduced to only 23 pounds per square foot at
100,000 feet.

The temperature of the atmosphere
behaves quite strangely. Starting from a
value of 58 F at sea level it decreases steadily
till at an altitude of 36,000 feet it has fallen to
-68 F. From 36,000 to 65,000 feet the
temperature is constant. From 65,000 to
100,000 feet (which approaches the upper
operating limit for ramjets) the temperature
rises slightly. Above 100,000 feet it rises to
170 F at 180,000 feet, than falls to -28 F at
260,000 feet, and rises again to 188 F at
380,000 feet.

1.00

2000

o
o

1500

T3
C
3
O

1000

500-

50 75 100

Altitude, Thousands of Feet

50 75 100

Altitude, Thousands of Feet

100-1

a
o

1200-1

« 1100-

I 1000-

(4-1

o 900

— I —
25

— 1 —
50

— 1 —

100

Altitude, Thousands of Feet

50-

■100

■50-

50 75 100

Altitude, Thousands of Feet

CHOKED FLOW

When flow in a duct or passage is such
that the flow upstream of a certain critical
section cannot be increased by a reduction of
downstream pressure, the flow is said to be
choked.

employing the stream thrust as a
parameter. The stream thrust is a
particularly useful quantity in ramjet
calculations because the difference in
stream thrust between two stations is
equal to the thrust exerted in an axial
direction on the duct walls between the
two planes. Moreover, when the local
Mach number is unity, as at a throat or
choking section of the duct, the stream
thrust becomes a direct measure of the

exit stream thrust of a ramjet equipped
with a non-expanding exit nozzle. Thus,
stream thrust for unit mass flow and
nozzle area depends only on
thermodynamic characteristics of the
exhaust gas. It is therefore a useful
measure of the combustor performance
(analogous to the characteristic velocity,
c*, employed for rocket engines).

Operational Characteristics of
Propulsion Systems

Selecting a vehicle propulsion system
involves consideration of many aspects of
its performance and use. The
performance characteristics of primary
importance are (1) thrust per unit frontal
area, (2) thrust per pound of engine

25-1

o
o

3
O"
(Z3

•a

•a
a
S
g

20-

15 —

10—

Booster rocket

Reciprocating engine
• Sustainer rocket

Ramjet

Turbojet with afterburner

Turbojet

— \ 1 \ r

1.5 2.0 2.5 3.0 3.5

Mach Number

4.0

Thrust per Unit Frontal Area for Propulsion Systems

12 3 4 5
Flight Mach Number

Thrust to Weight Ratio
for Propulsion Systems

weight, (3) fuel consumption rate per
pound of thrust, and (4) speed and
altitude boundaries for efficient
operation. To these performance
characteristics must be added such
considerations as cost, flexibility in
installation, and reliability.

Since the lift-to-drag ratio of
supersonic vehicles is one quarter or less
that achievable in subsonic types, the
thrust required for a high speed vehicle to
carry a given payload becomes relatively
large. Engine drag thus becomes a
significant part of the overall drag, so
thrust per unit frontal area is a
characteristic of primary importance in

engine selection. Above Mach 2, ramjets
are superior to turbojets or sustainer
rockets, which generally have relatively
low thrust-to-weight ratios, in terms of
this parameter.

The same reasons that make thrust per
unit area important for supersonic
propulsion apply to the thrust delivered
per pound of engine weight. In this
respect ramjets are markedly superior to
turbojets.

Fuel specific impulse determines the
range of the vehicle and is accordingly the
principal discriminator for long range
missions. For short range applications it
may be of little significance since other
factors, such as available volume of the
vehicle, may be dominant. The fuel
specific impulse of ramjets is relatively
poor compared to other airbreathers
until speeds above Mach 1.5 are reached,
but above this point it is superior to that

SPECIFIC IMPULSE

The specific impulse is the number of
pounds of thrust delivered by one pound of
propellant burning in one second. Specific
impulse is given in seconds. In ramjets,
the propellant is simply the fuel.

Ia=Fg/Wa

If=Fg/Wf=Ia/f

Ig = air specific impulse

lf= fuel specific impulse

Fg = gross thrust (due to the internal flow)

Wg = weight rate of flow of air
Wf = weight rate of flow of fuel
f=Wf/Wa= fuel-air ratio

6000-

5000-

o
u

OT 4000-

6 3000-

a 2000-

1000—

• Reciprocating engine

-i 1 1 1 1 r

0.5 1.0 1.5 2.0 2.5 3.0

Mach Number
Fuel Specific Impulse for Propulsion Systems

-| 1

3.5 4.0

of all other chemical propulsion systems.

By comparison, the specific impulse of
rocket propellants at sea level is about
250 seconds. These propellants contain
more oxidizer by far than fuel
(typically 70 to 80% of total propellant
weight). It would therefore be expected
that the specific impulse of a ramjet
would be several times as great because
no oxidizer has to be carried on board. In
fact, the specific impulse of common fuels
in ramjets is from 1000 to 1500 seconds
over the normal range of flight Mach
numbers.

As the speed of ramjets is increased
over about Mach 4, the rapid increase in
air stagnation temperature causes design
difficulties due to structural heating.

STAGNATION TEMPERATURE AND
PRESSURE

When a gas is decelerated so that its final
speed is zero, its kinetic energy of motion is
converted partially to an increase in static
pressure and partially to heat. Its
temperature rises to a final value termed the
stagnation temperature, which is related to
the Mach number of the gas before
deceleration. The stagnation pressure is the
static pressure achieved under these
conditions.

T= tod +0.2m2)

T = stagnation temperature in degrees

Rankine (Fahrenheit plus 460)
tg = local air temperature in degrees Rankine
M = flight Mach number

3000-1

1 1 r

12 3'

Flight Mach Number

Stagnation Temperature as a Function
of Flight Mach Number

Higher heat transfer rates in denser air
cause the temperature Umit to be reached
at a somewhat lower speed at sea level. At
the "thermal boundary," about Mach 4,
materials problems for both airframe and
engine become severe. Unless active
coohng were provided, the vehicle would
have to operate at a red heat and would
have to be made from expensive high-
temperature metal alloys. Furthermore,
there is a loss of fuel effectiveness in
heating air that is already very hot.
Thermal efficiency of the engine
decreases because of dissociation of the
products of combustion into molecular
fragments. This process absorbs energy
and therefore limits the temperature rise

t3 120-

E^ioo-

•a so-
ts

CS
V3

3
O

H 40

-S 20H

2 0-

60-

1 r

2 3

Mach Number

Ramjet Operating Limits*

*Assumptions

Conventional (hydrocarbon) fuel
Uncontrolled aerodynamic heating
Subsonic combustion

that can be attained in the engine.

The altitude ceiling is reached when
pressure in the combustion chamber falls
too low for efficient combustion.
Moreover, in most vehicle designs a
somewhat lower ceiling would be
imposed by the need for air pressure to
provide hft and maneuverability for the
vehicle. The most serious drawback of
the pure ramjet is its inability to produce
thrust at zero speed and the
accompanying strong dependence of
thrust on flight velocity. Evaluation of
ramjet applications must, therefore,
always consider the combination of
ramjet and boost power plant and judge
the combination in terms of fuel economy
and engine weight.

Although subsonic ramjets are
feasible, their performance is low, so it is
in the supersonic flight regime that
ramjets display advantages over other
propulsion systems. Most of the
discussion in the following pages will
therefore be based on supersonic ramjets.

Air inlet and
supersonic diffuser

Subsonic i

■ diffuser ~*"|"*'
/

Combustor

Nozzle

Normal
shock

Oblique
shock

Station

Station Number

1
2
3
4
5
6

Location

Vehicle flow field immediately upstream of the air induction system

Capture station - beginning of internal flow system

Cowl lip

Diffuser exit - combustor entrance

Combustor exit

Nozzle throat

Nozzle exit

Liquid-fueled Ramjet Nomenclature

Production costs of ramjet engines
tend to be low in comparison with
turbojet or piston engines because of the
lack of rotating machinery. In addition,
the large advantage of ramjet engines in
thrust per pound of engine weight leads to
significant cost savings for systems with
the same thrust level.

Efficient design of supersonic vehicles
requires close coordination of the
interface between power plant and
airframe. Designing airframe and power
plant as a unit accordingly places a

premium on engine flexibility in redesign
to accommodate desired changes in
dimensions or performance. Because of
the simplicity of the ramjet, small changes
in scale or performance may usually be
accepted without extensive redesign,
retooling, or test programs.

No power plant, however attractive
from the standpoint of performance or
operating characteristics, can succeed in
commercial or military applications
unless it is reliable. The need for
reliability in complex power plants or

Air inlet and

supersonic diffuser

I / I Subsonic i
I ' 'h diffuser -*^^

Normal
shock

Oblique
shock

Station

I. /.I s

Combustor

•H"* — Nozzle

\\<

Air injector

Mixing device
Solid fuel grain
Centerbody

Station Number

2
3
4
5
6

Location

Vehicle flow field immediately upstream of the air induction system

Capture station - beginning of internal flow system

Cowl Up

Diffuser exit - combustor entrance

Combustor exit

Nozzle throat

Nozzle exit

SoUd-fueled Ramjet Nomenclature

those requiring close tolerances tends to
result in extensive quality control, hence
greater cost and decreased production
rates. The simplicity of the ramjet engine,
with its complete absence of moving parts
exposed to hot gases, makes it extremely
attractive in this regard.

Liquid-fueled Ramjets

The characteristics that distinguish the
liquid-fueled ramjet (LFRJ) are the fuel
delivery system, with which fuel is

introduced, and the combustor, which
includes a flameholder, the combustion
zone where heat is released, and a nozzle
through which the burned gases are
ejected rearward at high velocity. The
LFRJ requires a separate fuel storage
system that can supply fuel to the delivery
system. There must also be a fuel control
system to adjust fuel rate to air rate
(which varies with vehicle altitude and
flight speed) and control flight speed of
the vehicle as desired. Some form of
auxiliary power supply must be provided

.^^

Ranvjet fuel

•3" Podded combustor
-J-''

_ Integral rocket ramjet combustor

Typical Engine Configurations

to furnish power to drive the control
system.

Solid-fueled Ramjets

Solid-fueled and liquid-fueled ramjets
are related in the same way as solid-
propellant and liquid-propellant rockets.
The main characteristic that distinguishes
the solid-fueled ramjet (SFRJ) is the
absence of fuel tankage, delivery, and
control systems, since the fuel is entirely
contained in the combustor at the
beginning of the duty cycle. In addition,
the combustor is usually simpler because
there is no liquid phase fuel to be
atomized and mixed with air in the
combustor. Instead, there is an air
injector to increase the turbulence of the
air as it enters the combustor so as to
improve flameholding. There may be a
mixer to ensure that fuel-rich and air-rich
gases are thoroughly mixed to improve
combustion efficiency, which is always a
key consideration in any combustion
process involving a gas and a solid.

Integral Rocket Ramjets

Early ramjet systems employed a
separate detachable booster to achieve
ramjet takeover speed. However, this
scheme was not always well suited to
launcher installations or to other
operational requirements. For example,
it meant dropping a fairly heavy piece of
hardware earthward, so launches were
limited to uninhabited areas.

Dependence on a tandem booster
ceased with the conception of the integral
rocket ramjet (IRR), successfully reduced
to practice by CSD, which employs a
dual-purpose combustor that first serves
as a rocket combustion chamber for
booster propellant cast into it. The
propellant burns and accelerates the
vehicle to a high speed. Then inlet air is
allowed to enter the combustor where it
encounters either a liquid or a solid fuel.
The fuel then burns in the combustor in
the normal manner of a ramjet.

Because the boost rocket operates at
1000 to 2000 pounds per square inch and

Dual purpose combustion chamber

Ramjet fuel

Nozzle clamp

Booster nozzle

Inlet port cover

Booster Operation

Dual purpose combustion chamber
Ramjet fuel

Nozzle
clamp ■

Booster nozzle

Inlet port cover

Transition

Dual purpose combustion chamber
Ramjet fuel

Ramjet Operation
Operating Sequence of Integral Rocket Ranyet

the sustain ramjet operates generally at
less than 100 pounds per square inch, two
nozzles are normally required.
Moreover, since the boost nozzle has a
smaller throat diameter than the ramjet,
the boost nozzle must be expelled before
ramjet operation begins. This scheme of
operation is quite workable; however,
some other techniques that achieve the
same overall effect of boost-sustain
operation are available (and will be
described later).

The simplicity of the IRR makes, it
aerodynamically "cleaner," more
reliable, and lighter than a ramjet with a
separate booster. In some form the IRR
will doubtless be one of the leading
propulsion systems of the 1980s.

Ducted Rocket

Strictly speaking, the ducted rocket is
not a ramjet. However, it is an
airbreathing close cousin and its
operational characteristics are so similar
to a ramjet that the two systems can be
considered together for all practical
purposes. The configuration of the
ducted rocket can be considered similar
to an LFR J whose fuel tank is replaced by
a fuel-rich sohd propellant grain. The
amount of oxidizer in the grain is just
sufficient to sustain combustion in the
absence of air. The fuel-rich gas
generated by the grain mixes with inlet air
in the combustor, or aft mixer, and is
exhausted through the nozzle. The major
problem is to mix air and exhaust gas
streams thoroughly so as to obtain high
combustion efficiency. The advantage is
that the ducted rocket can attain higher

THRUST MARGIN

The thrust margin is the ratio of the
difference between thrust and vehicle drag to
the vehicle drag. The term therefore
indicates the fraction of thrust (as a function
of drag) available to accelerate the vehicle in
level flight. If chmb is involved, the thrust
margin must also include a term for weight.

thrust margins at low supersonic speeds
than the integral rocket ramjet. The
ducted rocket's performance depends
greatly on air inlet angle and velocity,
Mach number of the gas from the solid
fuel gas generator, impingement angle,
and air/ propellant ratio. The significant
difference between fuel grains for the
SFRJ and the ducted rocket is that the
ramjet grain does not sustain combustion
without air, since it normally contains
Httle or no oxidizer. The ducted rocket
grain supports combustion (because of a
higher oxidizer content), so many of the
ramjet-oriented problems relating to
flameholding and recirculation are not as
important.

In principle, because the ducted rocket
contains part of its oxidizer, it does not
have a performance potential as high as a
pure ramjet. This disadvantage is offset
by increased operational flexibility. The
ducted rocket therefore represents one of
the simplest forms of ramjet-type engines
in that there is a reduced dependence on
flight parameters. In most applications
the ducted rocket is used with an integral
rocket booster, so the combustor
functions initially as a chamber for a solid
propellant rocket motor. It is
advantageous if the ducted rocket

Booster nozzle

-—-> .■ ■ ■ ■ ■ ■ ■■■■ ■ ■ ■■ •■ ' . ■■

c.'v.^V'Vr-'i "''ii'ii^

Port cover

Boost

Booster Propellant

Inlet

\r- Fuel grain (gas generator)
I ,^ Sustain nozzle

Sustain

Ducted Rocket Configuration

combustor can be made to operate
efficiently without mixing aids or
flameholding devices, thus eliminating
the problems of trying to fit a solid
booster grain in and around the various
aids and devices. Moreover, the axial

momentum of the effluent from the fuel
generator can be preserved and
combustor pressure losses can be
minimized while achieving complete
combustion and mixing and therefore
high combustion efficiency.

Air Induction Systems

The diffuser transforms the kinetic
energy of the air entering the engine into a
pressure rise, called the ram pressure.
The magnitude of the ram pressure is a
function of flight speed and the design
characteristics of the supersonic and
subsonic sections of the diffuser.

When the ramjet is operating, air from
the atmosphere enters the engine. After
the velocity of the air has been reduced
and its static pressure increased by the
supersonic diffuser, the air enters the
subsonic diffuser and is compressed still
further. It then flows into the combustor
where it is heated to 3000 to 4000 F by
continuous combustion of fuel. The hot
gaseous products of combustion are then
expanded in the exhaust nozzle section
and are ejected from the engine with a
velocity exceeding that of the entering air.

In the usual ramjet, the air
approaching the engine at supersonic
speed must be slowed to a subsonic value
low enough that it will not blow out the
flame in the combustor. A Hnear
supersonic flow can be reduced to a
subsonic flow only if it passes through a
normal shock wave. It is characteristic of
a shock wave that the subsonic flow
leaving the shock is at a higher static

pressure than the supersonic flow
entering. In all cases, shock waves are
accompanied by a decrease in available
energy: the stronger the shock, the greater
the decrease. However, the flow at
supersonic velocity can be reduced to
subsonic velocity by causing the
supersonic flow to pass first through one
or more oblique shocks and finally
through a weak normal shock. Under
these conditions the loss in available
energy is smaller, and the flow leaves the
weak normal shock at a velocity slightly
less than Mach 1.

The subsonic diffuser (through which
the stream must pass next on its way to
the combustor) further reduces the Mach
number of the flow to about 0.2 to 0.4 at
the entrance to the combustor. Because
of this deceleration there is an additional
rise in the static pressure above the rise
resulting from the shock waves.

Ideally, a diffuser configuration should
be chosen that will compress the
supersonic approach stream with
continuous reduction of the air speed to a
final value appropriate to the through
duct. The diffuser should do this by
converting the kinetic energy of the
stream to pressure energy with no energy

PRESSURE DISTURBANCES — SHOCK WAVES

A disturbance originating from a source is
propagated in all directions at the speed of
sound through the fluid surrounding the
source. If the source is moving at subsonic
speed, it is in effect trying to catch up with the
sound waves (disturbances) that its motion
produces. At subsonic speed, however, the
acoustic speed is always larger than the speed of
the moving body or source. The body therefore
always moves into a fluid that has already
undergone changes because of the motion of the
body. That is, the fluid ahead of the body may
be said to become aware of the presence of the
body because the latter propagates disturbance
signals ahead of itself. Thus, when a body
moves at subsonic speeds, the disturbances it
creates are said to clear away from it.

The situation is quite different when the body
moves at supersonic speed. The wave front of
the disturbance created by the body lags behind
the point on the body that created the
disturbance, so the disturbance wave front
cannot overtake the moving body.
Consequently the moving body is always
outside and ahead of the disturbance wave front
it produced. The different disturbance wave
fronts are enveloped by a conical surface, called
a Mach cone, the shape of which is related to the
speed of the body. When a body moves with
supersonic speed, all of the disturbances in the
flow are confined to the Mach cone. In the
regions outside the cone, the fluid medium is
unaffected by the moving body. The conical
separating surface therefore forms a wave front
called a Mach wave, which is a weak
compression shock.

In the wave system formed at the nose of a

vehicle moving at supersonic speed the wave
front is very steep. In traversing the wave there
is a large pressure rise, called a shock wave. The
shock phenomenon is a more or less instant
compression of the gas, so it is not a reversible
process. Energy for compressing the gas
flowing through the shock wave is derived from
the kinetic energy that the gas possessed before
the shock. Because the process is irreversible,
the kinetic energy of the gas leaving the shock is
less than that corresponding to reversible
compression between the same pressure hmits.
The reduction in kinetic energy appears in
heating the gas to a temperature above that for
the reversible compression process.
Accordingly, there is a decrease in the available
energy of the gas.

There are several different kinds of shock
waves, each with particular characteristics. In
some cases the shock wave is stationary with
respect to the body upon which it is formed.
This shows that the speed of propagation of the
shock wave is equal to the speed of the body,
otherwise the stationary relationship could not
be maintained. When the shock wave is formed
so that it is perpendicular to the direction of the
flow, it is termed a normal shock (here
"normal" is used to mean "perpendicular"
rather than "usual").

In many situations involving shocks, the
direction of a supersonic flow is changed
sufficiently that the gas is compressed in such a
way that a shock front is formed that is inclined
with respect to the initial flow direction. Such
shocks are termed oblique or angle shocks.
Where an oblique shock is formed, the fluid
stream is deflected toward the shock.

losses. There are several types of diffusers
that can fulfill this function satisfactorily.
The particular type that is "best" depends
on Mach number (that is, air speed of the
ramjet). Some diffusers with favorable
internal flow may cause unacceptably

high external drag. The optimum
performance of many diffusers is shown
only at or near a single design point and
worsens rapidly with changes in angle of
attack or Mach number. Practically,
the diffuser must usually be selected to

Shock wave

Gas Properties in a Ramjet Engine

perform well over a range of Mach
numbers and angles of attack.

Since the steady state performance of
supersonic diffusers is well understood,
they may be designed for accurately
predictable air flow reception and

pressure recovery under steady-state
conditions. Time-dependent phenomena
are less well understood. For example,
spike diffusers operating at an off-design
condition sometimes display an
oscillatory phenomenon in which part of

ANGLE OF ATTACK

The angle of attack is the angle between a
reference line fixed with respect to an
airframe (usually the longitudinal axis) and
the direction of movement of the body. The
angle of attack affects ramjet performance
because the incident air stream is no longer
parallel to the diffuser centerline, so the inlet
shock train is shifted. This shift must be
taken into account by the inlet design.

the air compressed supersonically spills
outside the inlet. This phenomenon,
called "buzz," involves rapid forward and
backward movement of the shock pattern
at the diffuser inlet. This movement is
accompanied at high Mach numbers by
pressure oscillations that may be of
destructive intensity.

Problems also exist in designing and
testing unsymmetrical configurations
such as scoops or off-axis inlets.
Moreover, when the diffuser is
sufficiently close to the combustor,
asymmetries in the subsonic flow of the
inlet can affect combustor operation.

Inlet Designs

One of the most difficult problems in
designing ramjets is in connection with
air inlet systems, especially for flight at
very high speeds. With any type of system
there is a problem of regulating the inlet
flow of air as flight speed is varied. One
solution is to vary exhaust nozzle area
and also inlet area by mechanical means.
However, the more usual scheme is to
design the air intake so that the shock can
travel back and forth to accommodate
changes in air flow.

The design of air induction systems for
supersonic airbreathing missiles is
influenced by both external and internal
factors. External factors include

compatibility with the launcher,
maximum allowable missile length and
weight, restrictions on ground clearance
of the aircraft, and placement of aircraft
structures such as aerodynamic surfaces
and landing gear. Internal factors to be
considered include packaging of the
missile guidance and control system,
warhead, and propulsion system
components. Even with all of these
constraints the inlet must deliver an
adequate supply of air to the engine over
a wide range of flight operating
conditions, and must do so with a
compact, well-integrated design that
offers both low cost and low drag.

To strike a reasonable balance among
all of these conflicting requirements
usually means designing an inlet for each
new application. As a result, several
types of inlets have been developed to
meet various combinations of
requirements.

There are four basic types of inlets:
normal shock, internal contraction, two-
dimensional, and three-dimensional. A
normal shock inlet is essentially a circular
duct, slightly smaller in diameter at the
leading edge. Since there is no supersonic
diffuser, the transition of the inlet air
from supersonic to subsonic flow occurs
across the normal shock which resides at
the inlet plane. The performance of this
type of inlet is rather poor compared to
inlets where external compression serves
to weaken the normal shock at or near the
cowl lip.

Subsonic diffuser
Supersonic Constant-area

Combustor

^-dif fuser — |— throat — U- > -4*

^/ .^

Oblique shocks r- Normal shock

Oblique shock ^Normal shock

Internal Contraction Inlet
with Constant-area Throat

Internal Contraction Inlet
with Bell-shaped Entrance

Two-dimensional Inlet

An internal contraction inlet is
essentially an inverted rocket nozzle. A
rocket nozzle accelerates subsonic
exhaust gas to supersonic speed. The
original high pressure of the gas is
simultaneously reduced to atmospheric
pressure, and the initial high temperature
is reduced by about half. An internal
contraction inlet decelerates supersonic
inlet air to subsonic speed.
Simultaneously it raises the temperature
of the gas and compresses it from its
original atmospheric pressure to some
higher value.

For this type of inlet to perform

Centerbody (support
not shown)
Cowl Up

Axisymmetric Inlet

efficiently, the strong normal shock in the
throat should be weakened by oblique
shocks upstream. This inlet theoretically
gives the highest performance of any
design at some single set of operating
conditions. However, the normal shock
tends to move in or out of the throat when
the engine is operated off the design
point. Accordingly, the throat is
purposely elongated to help stabilize the
normal shock. This inlet therefore does
not operate as easily over a wide range of
conditions as the types more generally
employed. Despite its lack of flexibihty,
the internal contraction inlet is

particularly good for such applications as
artillery-type ordnance where constant
velocity is maintained over a rather flat
trajectory. The reason for its
attractiveness for such appHcations is not
only its high performance but also its low
cost.

A two-dimensional inlet has a more or
less rectangular cross section. The
wedge-shaped supersonic diffuser
consists of one or more ramps which turn

1.0-1

Subcritical operation -

Unstart -

Critical operation-
Supercritical operation — >
(The relative weight flow can
be greater than unity because
of precompression from a wing,
for example. )

Relative Weight Flow 1.0

CRITICAL, SUPERCRITICAL, AND SUBCRITICAL OPERATION

There are three distinct conditions under
which a ramjet engine diffuser can operate,
depending on the heat released in the
combustor. When the heat released is just
enough that the back pressure at the exit section
of the subsonic diffuser causes the normal
shock to be positioned at the inlet thoats, the
operation is said to be critical; this is the design
condition.

Supercritical operation occurs when the heat
released in the combustor is below the design
condition. The back pressure at the outlet
section of the diffusion system becomes too
small to maintain the normal shock at the inlet.
The excess pressure (or energy) associated with
the internal flow must therefore be dissipated
inside the diffusion system by a strong shock
wave forming in the diverging portion of the
diffuser. In other words the normal shock
moves into the inlet.

The opposite condition occurs in subcritical
operation. If the heat release in the combustor is
increased, the static pressure at the exit of the
subsonic diffuser is greater than can be achieved
under the design condition. The normal shock
wave moves upstream, is expelled from the
diffuser, and continues to move toward the
vertex of the supersonic diffuser. Behind the
normal shock wave the flow is subsonic. Since
the shock wave is detached from the inlet the
incoming air spills over the cowl of the diffuser.

increasing vehicle drag and possibly leading to
instability (buzz).

These three operating conditions can be
related conveniently by means of a plot of
pressure recovery versus relative weight flow of
air. Pressure recovery is an efficiency factor, the
ratio of the actual pressure immediately
downstream of the diffuser to the theoretical
stagnation pressure. Relative weight flow is the
ratio of actual to theoretical weight flow. When
one of these parameters is plotted against the
other for actual ramjet operating conditions, a
curve of characteristic shape is produced.
Above the critical point (that is, in supercritical
operation) the diffuser operates with poorer
efficiency because the normal shock has moved
into the inlet. However, ramjet operation is
stable.

This range of stable operation is called the
supercritical margin. As the relative weight flow
decreases below that corresponding to critical
operation (for example, because of an increase
in drag that reduces Mach number), the normal
shock moves out of the inlet and the inlet begins
to operate subcritically. If the relative weight
flow drops low enough, the engine will unstart.
At the end of the transition from booster
operation and the beginning of ramjet
operation, the inlet must be operating within
the supercritical margin for ramjet takeover to
occur and stabilize.

the air flow, introducing oblique shocks
which decelerate the flow until at or near
the cowl lip (depending on the balance
between design and operating conditions)
a normal shock is finally formed.

A three-dimensional inlet is either
circular or elliptical in cross section. If it
is circular, it is called axisymmetric and
its diffusers have circular cross sections at
any point along their length. The
supersonic diffuser is essentially conical.
A variation of the three-dimensional inlet
is a half-axisymmetric inlet in which the
diffusers are rounded rather than wedge-
shaped.

Behavior of Inlet Shock Train

The shock train in the inlet adjusts
itself according to flight speed (inlet

Cowl
Normal shock
Bow shock

Vehicle body

Centerbody

Cowl

Front Side

Half-axisymmetric Inlet

Mach number) and pressure
requirements of the engine. If the
pressure supplied by the diffuser for a
particular operating condition equals the
design pressure, the inlet is said to be in
critical operation. If the pressure
supplied is greater than the required
pressure, the shock train moves into the
inlet and the operation is termed
supercritical. If the pressure supplied is
less than the required pressure
(subcritical operation), the shock train

Single Ramp

Leading Ramp

Double Ramp Triple Ramp

Inlet Behavior at Design Mach Number

Third ramp shock
Second ramp shock
Bow shock

Spillage

air flow — ~{ _

/ r- Sonic line
'Z^--'"''"'.,^ Terminal normal

shock

Subcritical Operation

At Design Mach Number St
Critical Operation H

Supercritical Operation

Behavior of Triple Ramp Inlet at Design Mach Number-

Above Design Mach Numb

moves out of the inlet and causes the
airflow to spill over, sometimes leading to
instability ("buzz").

In some inlets employing internal
compression a very troublesome
phenomenon can occur. Under certain
flight conditions when combustor
pressure is raised too high, the normal
shock can move so far upstream that
supersonic flow inside the inlet is lost.
When this happens the inlet loses its
ability to compress the incoming air
efficiently, causing a rapid loss of ramjet
thrust and flight speed. This phenomenon
is called an inlet unstart.

Sometimes it is possible to install the
inlet under a wing or in the nose cone in

the body flow field. The wing can then
serve as aprecompression device, making
a rather efficient arrangement for the
inlet. With precompression, both the
relative flow rate and the pressure
recovery can be greater than 100% of the
performance of the isolated inlet.

Inlet Location

Inlets are usually located to give peak
performance at the high angles of attack
(5° to 10°) typical of cruise operational
high altitude. Locating the inlets depends
largely on the airframe, particularly the
placement and arrangement of wings and
tail surfaces. There may be one, two, or

o^-.o

Four Axisymmetric Inlets

Two Axisymmetric Inlets Under Wing
and One Axisymmetric Inlet Under Body

Two Axisymmetric Inlets

One Axisymmetric Inlet

□

One Two-dimensional Inlet

Two Axisymmetric Inlets Under Wing
and One Two-dimensional Inlet Under Body

oQo

Two Axisymmetric Inlets Cheek-mounted

<>-<>

Two Two-dimensional Inlets

Half-axisymmetric Inlet Chin-mounted
Under Precompression Wing

four inlets that are axisymmetric, half-
axisymmetric, or two-dimensional.
Different types of inlets may be used in

multi-inlet configurations, especially in
SFRJs where one or more separate inlets
often supply bypass air to an aft or

secondary mixer section of the
combustor. Single inlets may be located
on the bottom of the vehicle in the nose
cone or body flow field (chin mount);
dual inlets may be located on the sides
(180° side mount or 90° cheek mount);
four inlets may be in a cruciform
arrangement. Generally inlets are located
in line with tail or dorsal fins to minimize
drag.

In LFRJs, after air leaves the diffuser it
sometimes passes through the fuel tank
by means of a transport duct of constant
area. The air then flows into the turn and
dump region of the combustor (where the
fuel injectors and flameholders are
located). If a flow-straightening device
(an aerogrid) is required, it is located at
the exit plane of the diffuser. Finally the
air enters the forward section of the
combustor.

In SFRJs, after air leaves the diffuser it
passes through the turn and dump region
of the inlet to the air injector system,
which may incorporate flow
straightening devices that improve

©©

©o©

Aerogrid for Square Diffuser Exit

flameholding. The air is then dumped
into the forward end of the combustor.
Early ramjet vehicles such as Navajo,
Bomarc, orTalos (all of which required a
separate detachable rocket booster)
placed inlet, combustor, and nozzle in a
separate pod or nacelle attached to the
main airframe, which contained the fuel
tank. More recent design practice is to
integrate the entire IRR into the vehicle
body.

Forward control surface

Strongback-

Warhead— /
Dual axisymmetric inlets

Tube-in-hole
air injector -

Example of Missile Powered 1

Fuel Management Systems

Fuel Delivery

It is convenient to subdivide liquid fuel
management systems into two parts: fuel
delivery and fuel control. The fuel
delivery system supplies fuel from the
storage tank by pumping or
pressurization. A pump-fed system may
be driven by any prime mover, but a gas
turbine is often employed since the
turbine is usually lighter than, say, an
electric drive motor and batteries.
Besides, batteries deteriorate in storage
and must therefore be replaced
periodically.

Gas for pressurized systems may be
obtained by bleeding ram air from the
inlet, or it may come from a high pressure
storage vessel through suitable reducing
valves. High pressure gas may also be
produced when needed by combustion of
liquid or solid fuel with air. An even
better way, since it is independent of an
air supply, is to generate the gas from a
self-contained solid-propellant grain of
suitable composition. When gas pressure
is required, the sohd grain is ignited and
furnishes sufficient gas for the duty cycle
of the ramjet. A system of this type is
particularly advantageous for

Boost igniter
-Aft port cover Ejectable boost nozzle

Ramjet nozzle -

d-fueled Integral Rocket Ramjet

applications where volume is limited and
low cost is a dominant consideration.
For example, certain chemical
compositions developed by CSD can
produce a larger volume of gaseous
nitrogen from a given volume of solid
than is obtainable from an equal volume
of liquid nitrogen.

Ramjet-powered vehicles are subject to
large acceleration forces during boost
(say, zero to 1200 miles per hour in four
seconds) and during flight maneuvers at
high speed. It is therefore important to
ensure that liquid fuel can be delivered
from the tanks under any flight
conditions. This requirement often
means devising some sort of positive
expulsion system so that the storage
volume is continuously made smaller as
the liquid is fed to the ramburner. In this
way the storage volume is always just
large enough to contain the remaining
liquid so only liquid can enter the fuel line
leading to the combustor. Accordingly,
storage tanks are frequently equipped
with collapsing or expanding bladders of
soft metal or elastomeric materials. When
gas pressure is applied to an expanding
bladder system the bladder expands
against the fluid, forcing it out of the
tank. Conversely, the bladder may
surround the fluid so that when it
collapses (from exterior pressure), the
fluid it contains is forced out. The
challenge in designing a positive
expulsion system is to ensure that the
bladder expands or collapses predictably
and without tearing, and that it empties
the tank almost completely. CSD has
developed and refined the technology
employed in making positive expulsion

systems that meet these requirements.
Fuel Control

Fuel control of a Hquid-fueled IRR is a
task whose complexity depends on the
range of operating requirements to be
met. Even for test stand conditions the
requirements are sufficiently complex
that preset controls are seldom usable.
Pneumatic (open loop) controls can be
used for relatively simple situations, but
more complex missions usually dictate a
more complex hydromechanical or
electronic (closed loop) device. The fuel
control system must match fuel flow with
air flow so as to maintain fuel-to-air ratio
within limiting values on both lean
(blowout hmit) and rich mixtures.
Operation of the fuel control system is
therefore closely interrelated with
conditions in both inlet and combustor.
In addition, the system must maintain an
appropriate initial flow of fuel during the
transition from rocket to ramjet
operation, control inlet pressure margin,
and limit flight Mach number.

Sometimes for special applications it is
possible to simplify or eliminate some of
these requirements. A simple feedback
system can maintain thrust, speed, or
combustor pressure constant for test
stand operation or for sustained level
flight. Or, for example, fuel flow might
be preprogrammed during certain
portions of the duty cycle of the
propulsion system. However, for a flight
profile that must respond to a wide range
of maneuvering requirements some sort
of adaptive control is necessary which
senses pressure in the air induction

system and receives commands and flight
trajectory data from the vehicle guidance
and control system.

A second level of complexity is reached
with an adaptive control for flight at
variable altitude or range to a fixed
terminal target. Although altitude and
speed change, there is no significant
problem with large changes in angle of
attack that affect inlet operation. Finally,
the most complex adaptive control is
needed to intercept a fast-moving, evasive
target because angle of attack may now
become a limiting parameter.

Operation of an LFRJ from sea level to
altitudes greater than 60,000 feet will lead
to air flow rates that vary by 10- to 15-
fold. In addition, there will be required a
twofold to fourfold variation in fuel flow
at a given Mach number and altitude
between a lean value for low-drag cruise
and a rich value for acceleration and
maneuver conditions. The fuel meter
must therefore be designed to control fuel
flow within, say, 5% over a 50-fold
variation in total flow. This stringent
requirement leads to a need for precise
construction and close tolerances.

Measurements of input parameters for
speed control may be obtained in a

number of ways. For example, fuel rate
may be set proportional to ram pressure,
which is approximately proportional to
the air rate near the design Mach number.
Mach number may be determined from
the ratio of ram pressure to static
pressure. Adjusting ram pressure by
correcting fuel flow causes the vehicle to
achieve the desired Mach number. If
velocity control is desired, a
measurement of the total temperature
may be used as a basis for converting
Mach number to velocity.

Mechanical, hydraulic, or electronic
systems may be employed to achieve the
desired accuracy and quickness of
response of the control system. The
system must be designed to be insensitive
to forces arising from rocket boost or
flight maneuvers. It must take into
account the possible effects of vehicle
angle of attack on both sensing elements
and control requirements. In addition,
any possibility of unstable operation
resulting from interaction between the
combustor or the diffuser and the
metering system must be prevented.
Hamilton Standard Division has
successfully adapted aircraft fuel control
technology to ramjets.

Liquid-fueled Ramburners

Another component requiring a high
degree of skill in engineering design and
analysis is the ramburner. Quite complex
factors enter into securing high rates of
heat release from the fuel with efficient
combustion at all required altitudes and
flight speeds. As an example, to minimize
frontal area and drag, the velocity of the
incoming air must be as high as possible.
On the other hand, the air velocity at the
entrance to the combustor must not
exceed about 250 to 300 feet per second
because of the complex interaction of
aerodynamic factors and chemical
kinetics.

The ramburner for an IRR includes the
combustor, port covers, igniter, and
thermal protection system. Engines for
early ramjet vehicles such as Talos or
Bomarc were installed in an external
nacelle; in more recent vehicles the
combustor is integrated with the vehicle
body.

A so-called dump combustor is
characterized by a rearward facing step or
sudden expansion of the duct that creates
a recirculation zone. Dump combustors
may be concentric with the vehicle or
side-mounted, depending on the overall
vehicle arrangement.

Port covers seal the air inlets to the
combustor during booster operation.
Then they must be removed by one of
several methods so that ram air can be
admitted during the transition from
booster to ramjet operation. The igniter
is the device that starts combustion of the
ramjet. The thermal protection system
insulates the combustor and nozzle
structure from the high temperature of
the combustion gases.

The combustors for LFRJs and SFRJs
are quite different because of their
significantly different combustion
processes and because the SFRJ
combustor contains all of the fuel. The
high velocity flow into the combustion
zone of an LFRJ is accompanied by a
high turbulence level as well as wakes,
eddies, and recirculation zones generated
by upstream spars, air intakes, and the
like. These phenomena create problems
of extreme complexity in devising
straightforward design methods for fuel
injection, vaporization, and controlled
distribution. The problems are
somewhat like those encountered in
designing injectors for liquid propellant
rocket engines, and must to a large degree
be solved by educated cut-and-try

VI
Nozzle

Flameholding
. Jet breakup, evaporation, mixing (spreading)
Fuel injection (penetration)
- Inlet flow

h L

Midstream fuel
and pilot injector

Y///(//////y^/C////^/////////////y ////////////////,

Air

inlet

'-v ' v/' ' ' VV/'y '

Wall fuel and
pilot injector

Recirculating

zone
Base fuel and
pilot injector

• >/////^^^yT/ 7 ?///)^//////?///J ( ^^/?/???/??^/,

l^TTx

Dump burner

Directed flow zone

Fuel Injector and Flame Stabilization Processes in Ramjet Combustors

procedures. Most of these methods
involve using an arbitrary number of fuel
injection points (nozzles) and varying the
geometrical arrangement and nozzle
vaporization characteristics semi-
empirically until a satisfactory fuel
distribution pattern has been found.
More quantitative methods can be
applied in predicting the rate of fuel
spreading from sources of various shapes
in a turbulent stream and in predicting

the degree of vaporization that will occur
in a known time for a particular type of
nozzle.

Flameholders

Early ramjet combustors encountered
a major problem in maintaining a stable
flame in the high velocity flow through
the combustor (some 250 to 300 feet per
second). The problem was solved by

Baffle
(gutter-type)

Air Flow

Dump angle A'-, — ^^

Can

Flame Stabilizing Devices

introducing baffle flameholders in the
duct. Behind these flameholders,

recirculation zones were formed where
the flow was sufficiently slow that a stable

flame could be estabhshed. The same
result was achieved by passing the inlet
air to the combustor through holes
or slots in a perforated plate or cone
mounted in the duct. Combustors with
the first type of flameholders have come
to be called baffle-type combustors. The
latter type is known as a "can"
combustor.

It was found quite early that a
relationship exists between flameholder
dimensions and flame blow-off velocity,
and that blow-off velocity depends on the
fuel-air ratio of the stream impinging on
the baffle.

Heat Release Rate

With fuel distribution and stability
limits defined in terms of combustor
geometry and air flow, one further major

Centerbody
support

Subsonic
diffuser

Fuel jets /- Baffle- type flameholder

Spike-type supersonic diffuser Pilot flame -i Exhaust nozzle ■

General Arrangement of Ramjet with Baffle-type Flameholder

Centerbody
support

Subsonic
diffuser

Fuel jets

Spike-type supersonic diffuser Can-type burner

Exhaust nozzle -

General Arrangement of Ramjet with Can-type Burner

design criterion is needed to complete the
requirements for combustor design. For a
given cross-sectional area, heat release
rate determines the length of combustor
required to achieve a desired combustion
efficiency.

It turns out that the heat release rate for
a baffle combustor depends on the rate
(or angle) at which the flame spreads into
the unburned material from the
stabilizing baffles. The required
combustor length is somewhat greater
than the distance needed for the flame
from one baffle to spread to the wall or to
contact the flame spreading from another
baffle. In a can combustor the unburned
material enters the combustion zone in
the form of jets that are gradually
consumed. Here the "jet length"
determines the required combustor
length.

In effect, both baffles and cans may be
thought of as devices to introduce sources
of ignition from burned material into the
entering combustible mixture. The rate
of heat release thus depends on mixing
rate and rate of combustion once an
ignition source has been provided.

Ramburner performance and

Flush wall
fuel injectors

Flame fronts

Flameholding Lr^r^i, V

-Torch

Recirculation zone
(flameholder)

centerbody with . .
fuel injectors 'Sniter

Axisymmetric Multistep (Dual Dump)
Combustor with Centerbody

Midstream

fuel

injectors

Flush wall
fuel injectors

/--^

-^-^z^-

Flameholding ^^,>??£:r:::i^

centerbody 1

with fuel injectors

Torch
igniter

Flame front

Axisymmetric Combustor with Centerbody

combustion instability have been shown
to depend on inlet air temperature. At
low altitude flight conditions, which
generally occur at low Mach number,
performance is decreased because of low
inlet temperature and high air flow rates.
This combination of conditions, when

Midstream and flush wall fuel injectors

■ Recirculation zone
Torch (flameholder)
igniter

Flame front

Axisymmetric Combustor with Center Dump Inlet

coupled with fuel that is injected under a
low differential pressure, produces large
fuel droplets with poor penetration,
increases the difficulty of transporting the
droplets into recirculation zones through
gas turbulence, makes vaporization
poorer, and causes incomplete mixing of
fuel with air and nonuniform
combustion.

Combustion Instability and Efficiency

Since the early days of ramjet engine
development, combustion instability has
been a problem of major concern.
Unstable, periodic fluctuation of
combustion chamber pressure that has
been encountered in ramburners arises
from several causes having to do with
combustion mechanism, aerodynamic
conditions, real or apparent shifts in fuel-
to-air ratio or heat release, and acoustic
resonance. From a physical standpoint
the probable source of instability is the
dynamic behavior of the recirculation
zone. Both skill and ingenuity are needed
to explain and correct combustion
instability when it appears.

Aside from considerations of
instability, the main concern in
ramburner design is to achieve high
performance. For IRRs, fuel distribution
in the combustor (especially in the
recirculation zones) and flameholder
design have the greatest impact on
performance. Obtaining high
combustion efficiency at low overall fuel-
to-air ratio requires high gas temperature
within the recirculation zone. With both
coaxial and side entry dump combustors,
injecting fuel flush with the wall in the

dump region will achieve this goal and at
the same time improve flame stability
limits. The only bad feature of this
arrangement is that efficiency drops off as
overall fuel-to-air ratio increases toward
the stoichiometric value (that is, neither
fuel-rich nor fuel-lean). Conversely,
midstream injection gives high efficiency
under stoichiometric conditions but
flame stability limits are compromised.
Therefore, some combination of injection
sites often gives performance superior to
either of these arrangements alone.

Port Covers

Port cover operation is a key element in
the performance of an IRR. The cover
(or covers) must withstand booster
chamber pressures of 1,500 to 2,000
pounds per square inch without being
ejected through the inlet. During the
brief transition period after booster

STOICHIOMETRIC RATIO

The stoichiometric ratio is a proportion of
chemical substances which is exactly correct
for a specific chemical reaction, with no
excess of any reactant. It is necessary to
specify the reaction since the stoichiometric
ratio is different if different products are
formed. For example, methane burns with
oxygen to form water and carbon dioxide or
carbon monoxide. More oxygen is needed to
form carbon dioxide from a given quantity
of methane than to form carbon monoxide.

COMBUSTOR EQUIVALENCE RATIO

The ratio of actual to stoichiometric fuel-
to-air ratio (often designated (p).

operation and before ramjet takeover, the
cover must be ejected reproducibly,
reliably, and without damage to the
combustor or nozzle thermal protection
system. It would be desirable if the port
cover were self-ejecting under ram air
pressure. Then, as the booster chamber
pressure decayed to zero, the cover would
be ejected automatically without the need
for any other mechanical device.

Both monolithic and segmented port
covers have been employed successfully.
However, several newer concepts are of
considerable interest because they offer
the possibility of reducing the size of
ejected pieces or eliminating them
entirely. One approach is to mechanize
the port cover, so that it can be opened
without being ejected. For example, the
cover could be hinged, louvered, or
sliding. In each case an actuating device
would be needed. These approaches
would be particularly good for small
dump openings. Moreover, the port
structural elements would create
recirculation zones that would act as
flameholders.

Another approach is to make the port
cover from high strength chemically
treated or heat treated frangible glass.
When the glass is broken (by a hard
metallic pin that penetrates the hard
outer surface or by a small detonating
device) it breaks into granules shaped like
rock salt and about the size of the original
glass thickness. This method seems best
suited for combustors with center dump.

Still another idea is to employ a
consumable cover consisting of a support
grid plate covered with a layer of solid
propellant reinforced with metal

screening. After the propellant burns
away, ram air flow consumes or ejects the
screen while the grid remains in place.
The grid could be arranged to act as a
flameholder.

Igniter

The igniter for a liquid-fueled IRR is
usually a one-shot pyrotechnic type
which is actuated briefly during ramjet
takeover. In early ramjets continuous
ignition was necessary to attain high
efficiency or to prevent blowout at lean
fuel-to-air ratios.

Thermal Protection System

The thermal protection system
maintains the combustor and the nozzle
below their maximum allowable
temperatures. For moderate heat loads
(Mach number less than 2.5) and
relatively short durations, air cooling,
film cooling, and radiation cooling
techniques are adequate, although most
of them add some complexity, cost, or
weight. The preferred technique is
ablative cooling, whereby the surface
layers of the protective material are
charred and vaporized. In this way they
absorb the heat of the combustion gases
and keep it from the structure being
protected. Suitable ablative materials
include inorganic oxides such as silica,
magnesia, or asbestos in phenolic, epoxy,
or silicone elastomers.

These materials char at accurately
known rates that depend on temperature
and velocity of the hot gas, so they are
easily applied to a particular duty cycle by

adjusting their thickness. The main
considerations in using these materials
are to ensure that the booster grain is
securely bonded to the ablator and to
retain the charred ablator in the
combustor throughout the ramjet duty
cycle. These requirements are not always
easy to meet, since the best ablators tend
not to bond to propellant very well, and
after they have become charred they tend
not to remain attached to the combustor
wall. The usual techniques for dealing
with these problems are to coat the
ablator with some material that bonds
well to both the ablator and the booster
propellant or to retain the ablator in the
combustor by means of some mechanical
device.

The maximum temperature of a
component protected by an ablative
material during long cruise missions is a
function of the thermal conductivity of
the fully charred ablative, the char
thickness, and the effect of external
aerodynamic heating resulting from
cruise conditions. After all vaporizable
material of the ablator has been driven

off, decomposition in depth of the ablator
is complete. Only the char remains, and
the thermal conductivity of the char is the
only property that affects the
temperature gradient through it.
Predicting the thermal conductivity of
char accurately at high temperature is
therefore of utmost importance.

Nozzle

The nozzle is often considered to be
part of the combustor, possibly because it
is often convenient to manufacture the
two components as part of the same
assembly. Most ramjet nozzles have a
fixed throat area. It is possible to devise
two-position nozzles so as to optimize
ramjet performance for two different
thrust levels. It is even possible to design
continuously variable nozzles that can
optimize performance at all flight
conditions. In practice, however, the
complexity and added weight of such
nozzles as well as the impact on the rest of
the propulsion system rarely justify these
approaches.

Solid-fueled Ramburners

Ramburners for SFRJs include some
components not present in their liquid-
fueled counterparts: solid fuel grain, air
injector, mixer, and bypass. (They also do
away with fuel control and delivery
systems.) In discussing these components
it is important to recognize that there are
two basic combustor configurations for
SFRJs. In the nonbypass arrangement all
of the air from the inlets passes through
the port (the open passage that runs the
length of the fuel grain). In the bypass
configuration a significant portion of the
inlet air is admitted downstream of the
fuel.

Nonbypass Configuration

The nonbypass configuration is lower
cost and uses simpler metal parts.

primarily because of the simpler interface
between combustor and inlets. But there
are two disadvantages. First, to stabilize
combustion there must be a rear-facing
step of suitable height to obtain
recirculation zones and the gas velocity
(or Mach number) within the fuel grain
port must be limited. These requirements
(which in effect set the flameholding
conditions) are defined in terms of a
critical step height (or port-to-injector
area ratio) and a maximum allowable
port Mach number. In practical system
designs these requirements often impose
an unacceptable limit on the amount of
fuel that can be carried. Moreover, at
higher ramjet thrust levels the combustor
nozzle throat area must be a larger
fraction of the combustor area (which is
usually fixed by the vehicle design), so the

Distributed air injector
Port cover

Solid fuel grain f- Thermal protection system

Nozzleless booster propellant

Air induction A
system—^

Boost igniter -

Ranyet nozzle"

Nonbypass-type, Solid-fueled Integral Rocket Ramjet

fuel port area must be further enlarged to
keep the port Mach number below its
allowable limit. Enlarging the port area
further reduces the amount of fuel that
can be contained within the combustor.
The second disadvantage of the
nonbypass configuration is its tendency
to low combustion efficiency, mainly
because of incomplete mixing of air and
fuel gas within the combustion chamber.
Efficiency can be improved by
incorporating mixing devices in the
combustion chamber or in a mixing
section downstream of the fuel grain.

Bypass Configuration

The alternate approach to combustors
is to bypass a portion of the captured air
and inject it downstream of the fuel grain
directly into the mixing chamber. This
method greatly improves combustion
efficiency and has another effect that is
even more beneficial. Since not all of the
air passes through the fuel grain port, the
port area can be reduced without causing
the air flow to exceed the allowable Mach
number. The reduced port area increases
the fuel load. Because of the dual effect

on combustion efficiency and fuel load,
overall performance of the bypass engine
can be considerably greater than the
nonbypass.

It must be emphasized that for certain
applications the thrust requirements are
not high enough for limiting fuel load to
be the deciding factor on overall vehicle
performance. The primary consideration
in selecting a configuration is often
combustion efficiency rather than fuel
load. In one application for a low thrust
engine, for example, preliminary
calculations showed that a bypass engine
offered 27% greater range than a
nonbypass engine. However, 20% of this
difference was attributable to lower
combustion efficiency and only 7% to
Hmiting fuel load. For this application,
therefore, the nonbypass engine could
compete quite effectively if its
combustion efficiency could be
improved.

Fuel Grain

Fuel grains for solid-fueled IRRs
almost always have a perforation that
runs the length of the grain to permit inlet

Port cover .
Tube-in-hole injector-

Air induction
system

Solid propellant sliver
for ramjet ignition
Solid fuel Mixing

grain •, device -, /^ <.

■Thermal protection system
Ranyet nozzle-
SoUd boost grain

Bypass air duct-

Boost igniter

Port cover Ejectable

boost nozzle

Bypass-type, Solid-fueled, Integral Rocket Ran^jet

Case

Thermal protection system
Fuel
Boost propellant

Boost/ Fuel Grains with Stress Relief

air to flow through it. Although an end-
burning grain could be employed (as in a
ducted rocket, which was described
earlier), practical considerations
essentially dictate a fuel grain with a
round hole or one with spokes of some
sort extending into the port.

In the IRR, the boost propellant grain
is in the same combustion chamber as the
sustain fuel. Like a conventional soHd
propellant, the boost propellant contains
its own oxygen for combustion.
However, the fuel grain for the sustain
portion of IRR operation requires
oxygen from the atmosphere. Although
for some applications it may be
appropriate to have a small percentage of
an oxidizing material mixed with the fuel,
the amount is insufficient to sustain
combustion in the absence of
atmospheric oxygen.

Flameholding

The basis for all modern developments
of SFRJs has been the ability to stabilize

flames by means of a rearward facing step
located at the forward end of the fuel
grain. The energy released by
combustion occurring in the recirculation
zone downstream of the step sets up
conditions that propagate the flame
down the remainder of the fuel grain.

There are two parameters that affect
flameholding. The first, called the
injector area ratio, is the ratio of the fuel
port area to the area of the air injector.
An equivalent parameter of more
physical significance is the ratio of the
height of the rearward facing step to the
fuel port diameter. Either of these terms
can be used so long as it is remembered
that the blockage area must be adjacent
to the fuel grain if combustion is to be

Rearward Facing Step with
Controlled Recirculation Zone

sustained in the recirculation zone. The
second parameter that affects
flameholding is the velocity (Mach
number) in the fuel grain port. It has been
found that the injector area ratio must be
greater than some critical value,
otherwise flameholding cannot be
achieved at any Mach number in the fuel
grain port. If this ratio is greater than the
critical value, combustion can be
stabilized over a range of port Mach
number up to some limiting value.
However, if the port number exceeds the
limiting value, blowoff will occur in any
case.

Two techniques have been developed
to improve the capability of the step
flameholder, which causes turbulent and
distorted air flow. The first is to inject air
directly into the recirculation zone to
intensify the heat release there, since the
combustible mixture in the recirculation
zone is fuel-rich. The second, which is a
little simpler in practice, is to use a so-
called tube-in-hole injector. Here the
annular sleeve in the combustor inlet
smooths inlet flow and proportions the
air for entrainment into the flameholder
recirculation zone.

Combustion Efficiency

Left to itself, the gas near the fuel grain
surface is fuel-rich and that near the
center of the port is air-rich. Unless this
difference is equalized, combustion
efficiency will be poor. In the nonbypass
engine configuration, mixing is promoted
by a vaned mixer located immediately
downstream of the fuel grain. The
enhanced mixing dramatically improves

combustion efficiency. The bypass
engine gives excellent combustion
efficiency more easily than the nonbypass
configuration because the bypassed air is
admitted at right angles to the fuel-rich
combustion gas stream. This
arrangement in combination with a
simple mixer plate with orifices gives
quite efficient mixing (as well as natural
secondary flame stabilization) under all
conditions, yielding high combustion
efficiency.

The bypass engine also offers the
potential for throttUng a solid-fueled
ramjet by varying the fraction of the air
that passes through the fuel grain port.
The fuel flow rate changes with the fuel
regression rate ("burning rate"), which
can in turn be varied by changing the air
mass flow rate through the fuel grain
port. Thus, a simple damper in the
bypass duct can change the air bypass
ratio and therefore the fuel flow rate.

Air

Distributed Air Admission

L_^

Pilot flame

Pilot-stabilized Combustion

Boosters

The booster portion of an integral
rocket ramjet must (1) accelerate and
launch the vehicle to ramjet takeover
velocity, (2) provide high volumetric
loading (fill the available space with as
much propellant as possible), (3)
accommodate center of gravity
requirements dictated by the guidance
and control system, (4) survive the severe
air launch environment, including
vibration and exposure to temperatures
ranging from -65 F to 165 F, (5) and end
its boost phase with a reasonably sharp
and reproducible pressure decay for
reliable ramjet takeover.

Booster Grain

Booster propellants that can meet these
requirements are quite similar to ones
used in solid propellant rockets. For a
liquid-fueled IRR, the booster grain is
cast in the combustor. When the booster
grain has burned, the inlet port covers are
opened and transition to ramjet
operation occurs. For a solid-fueled
IRR, the booster grain may be cast as an
inner layer over the sustainer fuel grain,
in the combustor volume downstream of
the mixer (where there is no fuel grain), or

in both places.

A rather effective refinement is to
remove a "sliver" of the fuel grain from a
recess at the forward end of the
combustor and fill the vacated space with
booster propellant. When the bulk of the
booster propellant has been consumed,
the sliver of booster propellant burns
long enough to ignite the air-fuel mixture
at the onset of ramjet takeover. One to
two seconds of additional burning
usually can be provided after the rest of
the propellant has been consumed, and
this interval is adequate for reliable
ignition. One reason this method works
is that during transition the chamber
pressure falls rapidly (since it is not being
maintained by the main portion of the
booster grain, which has been
consumed). When the pressure
decreases, the burning rate of the sliver is
diminished also. The burning surface of
the sliver is sized so that the pressure it
can supply by itself is an insignificant
percentage of the expected sustainer
pressure.

Locating the sliver in the recirculation
zone is ideal for other reasons. First, the
air flow is entrained in the sliver flame
and swept over the fuel surface, which is

heated very effectively. Second, in the
recirculation zone the regression rate of
the sustainer grain tends to be lower than
for the remainder of the fuel. By
replacing this part of the fuel grain with
booster propellant the difference in
regression rate can be partially
compensated for. Third, the larger
effective volume of the recirculation zone
can lower the recirculation velocities and
improve flameholder stability.

Nozzle

The nozzle for the booster has a smaller
throat diameter than the nozzle for the
sustainer. The reason for the difference is
that the booster operates at a chamber
pressure of 1000 to 2000 pounds per
square inch whereas the sustainer
operates at less than 100 pounds per
square inch. One way to meet these
differing requirements is to eject the
booster nozzle at the end of booster
operation. A disadvantage is that large
fragments ejected from the engine might
damage the vehicle or its launching
platform.

To eliminate this disadvantage and to

lower propulsion system cost, some
alternative nozzle concepts can be
considered. The idea of using a
submerged nozzle leads to some rather
interesting variations of the basic IRR.
As defined for airbreathing propulsion
system applications, a submerged nozzle
is located on the main missile axis but
forward of the aft plane of the missile.
Being on the main missile axis, the
submerged nozzle must dump its
supersonic exhaust products through the
aft portion of the combustor chamber
and the ramjet nozzle. The submerged
nozzle makes possible some propulsion
systems that are lower cost or may have
advantages in certain situations over the
standard IRR.

Tandem Rocket Ramjet

The tandem rocket ramjet (not to be
confused with a ramjet having a separate,
detachable booster) offers cost
advantages for some applications
because some of the complexities of the
transition sequence are eliminated. In
this arrangement the booster grain is
forward of the sustain combustor. The

Booster-

Tandem Rocket Ramjet with Submerged Nozzle

^ —

Ms-

Combustor

exhaust gases from the booster travel
through a blast tube and exit through the
booster nozzle, which is located inside the
ramjet nozzle. At booster exhaustion,
ramjet takeover occurs as liquid fuel is
injected into the aft-located combustor.
The ramjet exhaust gases escape through
the annulus formed by the ramjet and
booster nozzles. While this arrangement
avoids the problems associated with port
covers, it presents the difficult tasJc of
developing the blast tube.

Nozzleless Booster

A concept advanced by CSD that may
lead to lower cost propulsion systems is
the nozzleless booster. The reason is that
for virtually all solid rockets the nozzle
turns out to be the most costly single
component. The expense is caused partly
by the need for expensive materials that
can withstand high temperature and high
velocity at the nozzle throat. Other
materials, usually ablative, are employed
for exit cones. In addition, the nozzle's
design and construction are complex.
The temperature-resistant materials
must'be artfully reinforced with structural
members that hold the nozzle assembly
together under high internal pressure.

At first thought it might seem that

removing the nozzle would cause so great
a loss in performance that this approach
would be impractical. There is a loss of
specific impulse from the normal 245 to
250 seconds to around 200 seconds, or
some 20%. However, the space formerly
occupied by the nozzle can be filled with
additional propellant, reducing the
performance decrement. Moreover, the
cost of mixing a slightly larger batch of
propellant is merely the cost of the
ingredients. The propellant ingredients
at less than a dollar per pound are far
cheaper than a nozzle at perhaps $ 1 00 per
pound. Overall, a nozzleless booster costs
20% to 40% less than one with a nozzle.
A significant advantage of the
nozzleless booster in the integral rocket
ramjet is that it completely avoids the
complexity associated with ejecting the
booster nozzle. The ramjet nozzle,
however, is retained and contributes a
small improvement to booster
performance. To derive the greatest
benefit from a nozzleless booster it is
necessary for the propellant properties to
be tailored to the appUcation. Generally
such a propellant requires as low a
pressure exponent as possible and a
higher burning rate, density, and specific
impulse than conventional rocket
propellants.

Ramjet Testing

When a flight propulsion system is
designed to operate close to its limits of
strength it must be tested extensively
before committing it to a flight vehicle.
Ground testing of such a system can never
be a perfect substitute for flight testing
but it is the next best thing because,
carefully done, it can uncover most of the
potentially severe problems of the system.
Moreover, it is generally possible to start
ground testing before the flight vehicle is
available, and results of ground tests can
often influence vehicle design.

In testing a ramjet at sea level it is
important to reproduce as closely as
possible the conditions thai the engine
will encounter in flight. Three main test
methods are used: freejet, semi-freejet,
and connected pipe.

Freejet Testing

Freejet testing involves producing a
stream of air at the pressure, temperature,
and Mach number corresponding to the
flight condition and of sufficient size to
encompass the whole engine. Only in this
way is it possible to reproduce the flow
conditions around the engine and
determine its overall performance.

It is important to ensure that no
reflected shocks from the facility or the
model disturb the flowfields approaching
the air induction system. When the inlet
is near the nose cone of the vehicle, the
disturbances from reflected shocks are
negligible. However, aft-mounted inlets
must be shielded from reflected shocks by
a jet stretcher. Often it is difficult to
simulate the flow field at high angle of
attack; in addition, angle-of-attack
simulation itself may be hmited by the
relative sizes of model and facility.

Freejet testing is generally expensive
because of the size and complexity of the
facility, the large mass flow rates of
conditioned air that are needed, and the
specialized test equipment that is
necessary. However, essentially all
pertinent flow fields can be accurately
simulated. This feature is important to
propulsion system performance,
particularly when the air induction
system is closely adjusted to the local
flowfield.

Connected Pipe Testing

Connected pipe tests are adequate for
trying the combustion system, since the

combustion process is not affected by
flow outside the engine. In these tests, air
at conditions corresponding to those
leaving the intake diffuser is fed directly
to the combustion system.

Semi-freejet Testing

nozzle must be expelled as part of the
transition to ramjet operation. The test
facility must provide some sort of door
through which the ejected nozzle can pass
and which can then be closed to maintain
altitude conditions for the ramjet phase
of operation.

Semi-freejet testing of ramjet engines is
a relatively new technique that reduces
cost compared to classical freejet testing.
This technique is particularly useful for
testing IRRs since it can demonstrate real
time transition from rocket booster firing
as well as ramjet performance at specific
altitude and Mach number conditions or
over full flight trajectory using simulated
external aerodynamic heating.

The basic approach in semi-freejet
testing is to subject the air induction
system alone to the freejet environment
rather than the entire integrated vehicle.
Its main deficiency is that the air
induction system (and, consequently, the
propulsion system) is not being
influenced by local flowfield effects. For
configurations with inlets located far
back on the vehicle body the impact is
small, since flow conditions at the inlets
are nearly the same as freestream at zero
degrees angle of attack. On the other
hand, inlets located forward on the
vehicle body are in a region where the
flow is less uniform. The reason for the
nonuniformity is the greater influence of
forebody shock and its accompanying
flowfield. As a result, angle of attack is
limited to lower values in semi-freejet
than freejet testing.

Testing IRRs adds one further
complication since the rocket booster

Aerodynamic Heating

At the high Mach numbers typical of
ramjet operation, aerodynamic heating
can be a significant factor. Valid freejet
and semi-freejet tests must therefore be
further complicated by some means of
imposing aerodynamic heat loads to the

Tunnel diffuser nozzle

Wind tunnel wall

Jet stretcher

Inlet

Vehicle

Inlet

Reflected shocks'

Use of Jet Stretcher to Shield Inlet
from Reflected Shock Waves

Steam boiler -

Control center

Workshop

Steam accumulat

-Water conditioner
s^ai"^

Cooling
water

Air supply

Heater fuel
Ramjet fuel

Facility for Testing Integral Rocket Ramjets

Transfer flex tubes ^Air heater
i^ \ Forward
L flexure

Aft flexure
Ramjet engine
Ejector-

- Thrust butts

Main air supply

Direct-connect Ramjet Facility

Hot air diverter valve -j

- Forward
Transfer flex tube ^ I flexure

9? \ /^Air

heater

Ramjet engine

-Freejet nozzle

r Vacuum

shroud

^ Thrust butts V Main air supply

Semi-freejet Ramjet Facility

Diffuser -

Shroud air mass flow
Shroud heater

Primary air

n [7

Air heater

Primary air mass flow

Servovah^e-

Duty cycle inputs

Control
panel/computer
status and
interface Feedback signals

'tUUW,

)^niinnn),nt,i,,i

Liquid oxygen

"♦

Computer

Setpoint commands Oxygen

command

Shroud

heater Feedback Heater
and air signals command
commands , '^ .

•To aeroheat
shroud

Shroud heat
controller

Servovalve

-O

-a —

Shroud air
controller

^kkkkkk AAA

CZ3

Oxygen
controller

Servovalve

Control and
monitor panel

-D—

Z7*"

Heater fuel
control

Servovalve

Servovalve
Shroud air

F = Flow rate

P = Pressure

T = Temperature

Fuel-

Control System for Air Flow and Heaters

ramjet which match the flight Mach engine. The temperature and mass flow

number experienced by the inlet. In rate of shroud air must be continuously

practice a separate stream of heated air is determined and controlled to match the

supplied to a shroud surrounding the operating conditions at the inlet.

Applications for Integral Rocket Ramjets

Generally speaking, IRRs are superior
in performance to boost/ sustain rockets
when the ratio of sustain to boost impulse
is favorable. As a rule of thumb, IRRs
are better if less than about 70% of the
total impulse (thrust times duration) of
the system is required for boost. The
characteristics of IRRs make them
particularly suitable for (1) extended
range tactical weapons, (2) powered
intercept at longer ranges (for engaging
maneuvering targets), (3) replacing
existing volume-limited rockets with a
propulsion system having much higher
performance, and (4) extended range
ordnance against high-speed ground
targets.

Ramjets are favored over other power
plants in applications in which (1)
appreciable range is required at speeds
greater than Mach 2, (2) engine weight or
drag is a significant fraction of the weight
or drag of the vehicle, (3) engine cost is of
major importance, and (4) the available
volume is insufficient for a rocket motor
because of its significantly lower
performance. IRRs make auxiliary boost
unnecessary and therefore remove a
primary disadvantage of the pure ramjet.
It is worthwhile to discuss briefly a few of

the applications in which IRRs are a pre-
eminent choice.

Antiaircraft Guided Missile

High speed and long range capabilities
are required in antiaircraft guided
missiles designed to cope with high speed
bomber or fighter attacks. High thrust
per unit frontal area is important in
minimizing drag and reduces volume
requirements in missile-firing
installations. The ramjet's need for
rocket boost makes it possible to attain
cruising speed quickly and reduces the
range of minimum engagement and time
to target. Specific fuel consumption is not
of primary importance (except for
volume-limited applications), but values
as high as those typical of rockets cannot
be tolerated for ranges greater than about
30 miles. The simplicity and low cost of
the IRR as well as its outstanding
performance at high speeds make it a
prime choice for this appHcation.

Target Drones and Remotely Piloted
Vehicles

In target drones, low cost of the power

plant is of primary importance. The
comparatively simple construction of the
IRR offers significant advantages for this
application.

The greater efficiency of the ramjet at
high speeds gives it an advantage for
remotely piloted vehicles requiring long
range with cruise speeds above Mach 2. It
is clear that the need for increasingly
longer standoff distances will call for
higher and higher attack speeds. With
progress in automatic control and target
identification techniques the emergence
of the IRR as the favored propulsion
system for remotely piloted vehicles
seems assured.

not maintain fuel-to-air ratio at optimum
levels over a large range of altitudes. It is
most applicable, therefore, to missions
that do not demand a large change in
operating altitude.

The ducted rocket shows its best
relative performance in smaller class
vehicles where the operating envelope is
reasonably hmited. The high fuel density
of the ducted rocket makes it attractive in
volume-Hmited systems. However, since
its fuel is expended at essentially constant
rate it is penaHzed if operation is required
over a large range of altitudes.

Cruise Missiles

Tactical Missiles

The characteristics of various IRR
propulsion systems can be compared in
terms of the apphcability of several such
systems to a wide range of tactical
missiles. In a rather comprehensive study
it was found for weight- and volume-
limited missiles that the liquid-fueled
IRR can contain the maximum total
impulse. It therefore gives the highest
performance, especially for missiles that
must operate over a wide range of altitude
and Mach number. The ability to control
fuel flow to match performance
requirements is a distinct benefit. On the
other hand, where velocity alone is the
dominant factor, as for example in
intercepting a maneuvering target, the
performance advantage of the liquid-
fueled IRR is reduced.

Although the solid-fueled IRR offers
good performance it provides only a
partial form of fuel control and it does

Another useful comparison can be
made for cruise missiles. Here the ramjet
offers the advantages of low volume and
best performance at high altitude (70,000
to 90,000 feet) and high Mach number
(3.5 to. 4.5). Its disadvantages are that it
cannot match the long range capability of
the turbojet if subsonic cruise at low
altitude is adequate to meet mission
requirements. Moreover, to achieve long
range the ramjet must cruise at high
altitude and high speed, so it needs a
small radar cross-section if it is to
penetrate defensive screens defensively.

The SFRJ offers the desirable
operational and storage features of a
solid rocket (low cost, simplicity, and
high reliability, with minimum handhng
and support equipment). It offers greater
range than boost-sustain, dual chamber,
or pulsed solid rockets. However, the
solid fuel grain limits a multiple mission
envelope (many Mach numbers and
altitudes). Like the LFRJ, it must cruise

at high altitude and high speed to obtain
long range.

Podded ramjets have for many years
been in the military inventory. The
Navajo intercontinental missile employed
a podded ramjet, but the program was
cancelled during development. The
Navy's Talos (surface-to-air missile)
became operational in 1959 and is
scheduled to be phased out in the 1980s.
The Air Force's Bomarc (also a surface-
to-air missile) was developed in the mid-
1950s and is still operational, as is the
Army's Redhead Roadrunner, a
supersonic high or low altitude target
drone. The podded LFRJ is the simplest
and therefore the lowest cost ramjet
system (remembering that an auxiliary
booster must be supplied to make the
ramjet work). The principal
disadvantage is large total volume
compared to an IRR, with higher drag
during boost.

A number of cruise-type missiles have
relied on low-cost turbojet engines for
propulsion. In 1953 the Matador, a
surface-to-surface missile, became
operational. Within the ensuing 10 years
the Snark, Mace, Quail and Scad
(decoys), and Hound Dog (air-to-surface
missile) were deployed. The Harpoon and
the air-launched cruise missile are under
development. The advantage of the
turbojet engine for these applications
(aside from its estabhshed technology) is
that it gives the highest specific range per
pound of any propulsion system. Against
this considerable advantage must be
weighed a number of significant

disadvantages. Turbojet thrust is limited
compared to ramjets or rockets.
Turbojets cost more than competitive
systems and are essentially limited to
operation below Mach 2. Moreover,
turbojets have lower thrust per pound
and lower thrust per unit cross-sectional
area than ramjets or rockets.

Cost

In any propulsion apphcation, aside
from strictly technical considerations of
operating and performance characteris-
tics there inevitably arises the question of
cost. It is not uncommon for a particular
propulsion system to meet technical
requirements with flying colors only to be
rejected because of unacceptably high
costs. Therefore, some cost comparisons
with other propulsion systems are in
order. Comparisons with rocket engines
are not given since for most of the
applications of interest a rocket will not
meet the requirements unless its size,
weight, or complexity increase to
unacceptable values.

Turbojet engines are now employed for
drones and other pilotless vehicles that
could be powered by integral rocket
ramjets. Based on 1976 figures, a turbojet
engine costs $40 to $70 per pound of
thrust. Liquid-fueled IRRs for the same
missions would cost $8 to $15 per pound
of thrust. Solid-fueled IRRs would cost
$5 to $10 per pound of thrust. The
arguments in favor of the IRR are thus
fully justified when cost is made a
deciding parameter.

CSD, a Leader in Ramjet Propulsion

In 1972 the management of United
Technologies recognized the growing
need for ramjet propulsion systems in the
nation's arsenal and resolved to take

advantage of existing technological skills
held within the corporation, combine
these skills, and compete for ramjet
propulsion programs. In carrying out this

CSD's Sunnyvale Center

CSD's Semi-freejet Facility at Coyote Center

resolve UTC assigned its Chemical
Systems Division the responsibility of
leading the corporate effort with major
support from United Technologies
Research Center and the Hamilton
Standard Division.

To carry out its ramjet work, CSD
expanded its existing test facilities to
provide the capabihty for semi-freejet
testing of integral rocket ramjets. Unique
in the industry, this test facility is
designed specifically to test booster
operation, transition, and subsequent
ramjet operating modes in one sequence
under both sea-level and simulated high-

altitude conditions. The fully automated
facility, located at the division's Coyote
Center, can handle up to 1000 pounds of
solid propellant. Ramjet trajectories can
be simulated through computer control
of altitude, fuel and airflow rates, and air
total temperature, including the effects
on structures of aerodynamic heating.

The large scale facilities at Coyote
Center, which include capability for
hardware manufacturing, are
supported by research-scale installations
at CSD's Sunnyvale Center. In addition,
Hamilton Standard Division has
extensive facilities for developing.

testing, and manufacturing fuel control
systems. United Technologies Research
Center is generously provided with
facilities for testing air induction systems
and was first to prove an integral rocket
ramjet in flight.

The combination of highly qualified
staff from all three organizations.

comprehensive and specialized facilities
for developing, testing, and
manufacturing all components of integral
rocket ramjets, and CSD as a system-
oriented lead organization represents a
powerful team for meeting the nation's
growing need for ramjet propulsion
systems.

INDEX

ablative cooling, 37

adiabatic compression, 5

aerodyamic heating, 47

aerogrid, 28

air density in ramjet, 21

air injector, 15

air specific impulse, 10

air velocity in ramjet, 21, 32

altitude ceiling, 12

angle of attack, 22

antiaircraft guided missile, 51

Athodyd, 2

atmosphere, properties, 8

axisymmetric inlet, 23, 25

baffle flameholder, 34

blow-off velocity, 34

Bomarc, 2, 28, 34, 53

booster grain, 43

buzz, 22

bypass configuration, 40

can-type burner, 34

characteristic exhaust velocity, 6

cheek mount, 27

chin mount, 27

choked flow, 9

combustion efficiency, 36, 42

combustion instability, 36

combustor, 4, 32, 39

combustor equivalence ratio, 36

components, ramjet, 4

connected pipe testing, 46

controlled recirculation zone, 41

critical operation, 24, 25

critical step height, 39

cruise missile, 52

density ratio (atmosphere), 8

diffuser, 4, 19

distributed air admission, 42

double ramp diffuser, 25

ducted rocket, 17, 18

dump combustor, 32

exhaust velocity, 6

flame blowoff velocity, 34

flameholder, 14, 33, 41

flame stabilization, 34

Fono, Albert, 1

forces, resolution of, 7

freejet testing, 46

fuel control, 14, 30

fuel delivery, 29

fuel grain, 40

fuel port area, 39

fuel regression rate, 42

fuel specific impulse, 10, 11

gas generator, 29

gas properties (in ramjet), 21

gas turbine, 29

half-axisymmetric inlet, 25

Harpoon, 53

heat release rate, 34

Hound Dog, 53

igniter, 32, 37

injector area ratio, 41

inlet designs, 22

inlet locations, 26

inlet shock train, 25, 26

inlet unstart, 26

internal contraction inlet, 23

integral rocket ramjet, 3, 15

jet stretcher, 46, 47

leading ramp diffuser, 25

Leduc, Rene, 1

lift-to-drag ratio, 10

liquid-fueled ramjet, 13, 14

Lorin, Rene, 1

Mace, 53

Mach cone, 20

Mach number, 4

INDEX

Mach wave, 20

Matador, 53

maximum port Mach number, 39

mixer, 15, 40, 42

momentum rates, 5

Navajo, 2, 28, 55

nonbypass configuration, 39

normal shock, 19, 20

normal shock inlet, 22

nozzle, 4, 38, 44

nozzleless booster, 45

nozzle thrust coefficient, 6

oblique shock, 19, 20

operating principle, 5

operational characteristics, 9

pilot-stabilized combustion, 42

podded ramjet, 15, 53

port cover, 32, 36

port-to-injector area ratio, 39

positive expulsion, 30

precompression, 26

pressure (atmospheric), 8

pressure disturbances, 20

pressure in ramjet, 21

pressure recovery, 24

production costs, 13, 53

Quail, 53

ram air, 29

ramburner, 32, 39

ramjet recession, 2

ramjet resurgence, 3

ram pressure, 19

rear-facing step, 39, 41, 42

Redhead Roadrunner, 53

relative weight flow, 24

remotely piloted vehicle, 51

Scad, 53

semi-freejet testing, 42

shock waves, 20

single ramp diffuser, 25

sliver igniter, 43

Snark, 53

solid-fueled ramjet, 14, 15

specific impulse, 10

stagnation pressure, 1 1

stagnation temperature, 1 1

stoichiometric ratio, 36

stream thrust, 7

stress relief, 41

subcritical operation, 24, 25

submerged nozzle, 44

subsonic diffuser, 19

subsonic ramjets, 14

supercritical margin, 24

supercritical operation, 24, 25

supersonic diffuser, 2, 19

tactical missile, 52

Talos, 2, 28, 32, 53

tandem rocket ramjet, 44

target drone, 51

temperature (atmospheric), 8

temperature in ramjet, 21

thermal boundary, 12

thermal efficiency, 12

thermal protection system, 32, 37

three-dimensional inlet, 25

throttling, 42

thrust at zero speed, 4

thrust (equation), 6

thrust margin, 17

thrust per unit frontal area, 9

thrust-to-weight ratio, 10

total impulse, 51

triple ramp diffuser, 25

tube-in-hole injector, 42

two-dimensional inlet, 23, 24

unstart, 26

velocity of sound, 8

NOTES

Chemical Systems Division
United Technologies

1050 East Arques Avenue

P. O. Box 358

Sunnyvale, California 94088

, I II >' I,
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