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Full text of "Interface of materials and structures on airframes. Part 3, Design problems in aircraft structures including proceedings of Monterey Symposium / [by] Ulrich Haupt."

LIBRARY 

TECHNICAL REPORT SECTION 
NAVAL POSTGRADUATE SCHOOk 
MONTEREY. CALIFORNIA 93940 



NPS-57Hp71111A 



United States 
Naval Postgraduate School 




INTERFACE OF MATERIALS AND 


STRUCTURES 


ON AIRFRAMES 




PART 3 




DESIGN PROBLEMS IN AIRCRAFT 


STRUCTURES 


INCLUDING 




PROCEEDINGS OF MONTEREY ! 


3YMP0SIUM 


Ulrich Haupt 




October 1971 





Approved for public release; distribution unlimited 



FEDDOCS 

D 208.14/2:NPS-57HP71111A 



V 



INTERFACE OF MATERIAL AND STRUCTURES 
ON AIRFRAMES 

PART 3 

DESIGN PROBLEMS IN AIRCRAFT STRUCTURES 
INCLUDING 
PROCEEDINGS OF MONTEREY SYMPOSIUM 



Ulrich Haupt 
October 1971 



NAVAL POSTGRADUATE SCHOOL 
Monterey, California 

Rear Admiral A. S. Goodfellow, Jr., USN Milton U. Clauser 

Superintendent Provost 



ABSTRACT: 

The proceedings of the Monterey Symposium on Design Problems in 
Aircraft Structures provide a basic survey of design problems 
from the engineer's viewpoint. Further analysis of the present 
situation draws attention to some essential aspects which are 
not yet generally recognized. This leads to the conclusion that 
recent design problems cannot be solved on a technological level 
alone. An organizational effort is needed to disseminate available 
information. Beyond this, the complexity of interactions must 
be understood more thoroughly and this requires an educational 
effort on a broad basis. A practical and systematic approach 
toward the solution of these problems is developed. 

The present report covers the final phase of a project under the 
title Interface of Materials and Structures on Airframes. This 
project is supported by: Naval Air Systems Command 

Work Request No. WR- 2-6059 

dated l6 July 1971 



TABLE OF CONTENTS 



Foreword 

I. Proceedings of Symposium on Design Problems 

in Aircraft Structures 1-1 

List of Participants 1-1 

M. U. Clauser -- Welcoming Remarks 1-3 

U. Haupt -- Introductory Remarks 1-5 

W. H. Sparrow -- Translating High-Strength 
Steel Characteristics into Effective 
Structural Designs 1-10 

J. C. Ekvall -- Fatigue, Relating Past 

Experience to Design 1-22 

W. C. Dietz - Fracture Mechanics Considerations 

in Design and Manufacture 1-39 

W. S. Hyler -- Some Comments on Fail-Safe 

Design 1-67 

W. T. Shuler -- Impact of Procurement Practices 

on C-5 Structural Design 1-79 

F. F. W. Krohn -- Structural Design Considerations 

for a Space Shuttle 1-80 

J. E. Fischler -- Structural Design Considerations 

for Advanced Aircraft 1-107 

J. W. Ellis -- Some Practical Aspects of Risk 

Evaluation 1-135 

D. A. Shinn -- Advanced Metallic Structures I-1U7 

F. L. Cundari -- Summarizing Remarks 1-151 

II. Some Unsolved Questions and Various Comments II-l 

1. Regarding an acceptably low risk II-2 

2. Regarding development of metallic 

materials 11-10 



II 



3. Regarding present trends in fatigue 

criteria 11-17 

k. Regarding present trends in fracture 

mechanics and fail-safe design 11-22 

5. Regarding evaluation and application 

of structural metals 11-27 

6. Regarding more general design problems 11-31 

7. What other questions are considered 

pertinent to design problems? 11-36 



III. Some Basic Considerations and Conclusions III-l 

1. Synopsis of Present Situation III-l 

1.1 Introduction III-l 

1.2 Survey of Other Findings Concerned 

with Present Situation III-2 

1.3 Purpose of Present Project III-U 
l.U Method of Approach III- 5 

2. Outline of Design Problems III -6 

2.1 General Remarks III-6 

2.2 Basic Concepts III-6 

2.3 Design Problems in the Realm of 

Technology III-8 

2 .k Design Problems Beyond the Realm 

of Pure Technology III-ll 

2.5 Policy Problems Affecting Design 111-13 

2.6 Designer's Viewpoint III-16 

3. Unexplored Aspects of Design Problems III-18 

3.1 General Remarks III-18 

3.2 Engineering Curricula III-18 



III 



3.3 Continuing Engineering Education 111-20 

3,U Engineering Professionalism 111-20 

3.5 Practical Considerations about 

Educational Aspects 111-21 

3.6 Practical Considerations about 

Information Systems III -2 3 

3.7 Practical Considerations about Overall 
Perspective III-2U 

h. Conclusions 111-27 

References 111-29 

Distribution List 111-30 



IV 



FOREWORD 



This report forms the final part of the project Interface of 
Materials and Structures on Airframes and includes the Proceedings 
of the Monterey Symposium on Design Problems in Aircraft Sturctures. 

The project Interface of Materials and Structures on Airframes 
has been sponsored by the Naval Air Systems Command under the 
cognizance of the Structures Administrator and has been conducted 
at the Naval Postgraduate School, Monterey, California. Two 
earlier reports were published as Part 1: Basic Design Considerations 
and as Part 2: Outline of Decision Process in Structural Design. 
They were concerned with basic aspects of the design process as 
they influence problems of interaction between materials and 
structures . 

In order to coordinate these considerations with recent 
experience in industry and to recognize those design problems 
which most companies have in common, a symposium on Design Problems 
in Aircraft Structures was held at the Naval Postgraduate School 
in Monterey, California, on July 15 and l6, 1971. About twenty 
invited participants represented aerospace industry, government 
agencies, and research institutes. Prepared talks on particularly 
significant aspects of design problems were given by engineers 
thoroughly familiar with the present state of the art and 
discussion sessions followed. 

The proceedings of the symposium are given in Section I and 
related questions with various comments are listed in Section II to 
stimulate general thoughts regarding design problems. 

Section III contains some basic considerations and conclusions, 
attempting to bring together the viewpoints of design engineers, 
engineering management, and government agencies. It begins with 
a consideration of technological problems and recognizes that they 
are being approached systematically and competently on an 
engineering level. This leads, however, to the realization that 
such work on a purely technological level will not suffice to 
solve our design problems. Additional organizational and 
educational efforts will be required. Some of them have been 
recommended already by committees concerned with these problems. 
Some more aspects are added and an integrated approach is 
suggested on the final pages of this report (pages III-21et seq.). 

There is no consensus of opinions in this field. Conclusions 
must be based on the subjective interpretation of facts. They 
always have to be submitted to much discussion and careful 
examination. The considerations of Section III incorporate ideas 
developed by widely scattered people. Any shortcomings and 



V 



controversial aspects of the discussion are the obvious responsibility 
of the coordinator of this project. The conclusions must not be 
construed as necessarily representing the attitude of the Navy 
Department . 

The participants of the symposium and many individuals 
throughout the aerospace industry made this report possible by 
extending a spirit of full cooperation and giving generously of 
their sparse time and hard-gained experience to discuss problems 
which were not always easily defined. 



VT 



SECTION I 

PROCEEDINGS OF SYMPOSIUM ON 
DESIGN PROBLEMS IN AIRCRAFT STRUCTURES 



SECTION I 

PROCEEDINGS OF SYMPOSIUM ON 
DESIGN PROBLEMS IN AIRCRAFT STRUCTURES 



The symposium on Design Problems in Aircraft Structures was 
held at the Naval Postgraduate School in Monterey, California, on 
July 15 and l6, 1971. 

List of Participants 

Cundari, Frank L., CDR, U. S. Navy, Structures Administrator, 

Code Air 320B, Naval Air Systems Command, Washington, D. C. 20360 

Dickenson, Kenneth H. , Chief of Stress, Research and Engineering 
Division, The Boeing Company, P.O. Box 3955, Seattle, 
Washington 98124 

Dietz, W. C, Director, Airframe and Structures Technology, General 
Dynamics Fort Worth Division, P.O. Box 7^8, Fort Worth, 
Texas 761OI 

Ekvall, John C, Senior Design Specialist, Advanced Design and 
Laboratories, Lockheed- California Company, Burbank, 
California 91503 

Ellis, Joe W. , Supervisor, Advanced Design, North- American Rockwell 
Corp., International Airport, Los Angeles, Calif. 90009 

Fischler, Jerome E., Branch Chief, Advanced Design, Douglas 

Aircraft Co., 3855 Lakewood Blvd, Long Beach, Calif. 908OI 

Haupt, Ulrich, Associate Professor, Department of Aeronautics, Naval 
Postgraduate School, Monterey, Calif. 939^0 

Heldenfels, Richard R., Assistant Director for Structures, NASA 
Langley Research Center, Hampton, Va. 23365 

Hyler, Walter S., Senior Advisor, Mechanical and Systems Engineering, 
Battelle, 505 King Avenue, Columbus, Ohio 1+3201 

Krohn, F. F. W. , Manager, Structures and Thermo- Protective Systems, 
General Dynamics Convair Aerospace Division, P.O. Box 1128, 
San Diego, Calif. 92112 

Langen, William A. , Director, Aero- Structures Dept, Naval Air 
Development Center, Johnsville, Warminster, Pa. 1897^ 



1-1 



McQuillen, Edward J., Head, Applied Mechanics Research, Naval 
Air Development Center, Johnsville, Warminster, Pa. 1897^ 

Rosenfeld, Murray S. , Staff Scientist, Aero- Structures Dept., 

Naval Air Development Center, Johnsville, Warminster, Pa. 1897*+ 

Ryan, Edwin, Chief, Structures Branch Air 5-302, Naval Air Systems 
Command, Washington, D. C. 20360 

Shinn, Donald A., Chief, Aeronautical Systems Support Branch, Air 
Force Materials Laboratory, Wright- Patterson Air Force Base, 
Ohio 45^33 

Shuler, William T., Chief, Structural Engineering, Lockheed-Georgia 
Company, Marietta, Georgia 3006l 

Sparrow, W. H. , Chief of Structural Technologies, Vought Aeronautics 
Company, P.O. Box 5907, Dallas, Texas 75222 

Esary, James D , Associate Professor, Department of Operations 
Analysis, Naval Postgraduate School, Monterey, Ca . 939^+0 

Leonesio, Robert B , Assistant Professor, Department of Material 
Science and Chemistry, Naval Postgraduate School, Monterey, 
Calif. 939^0 



1-2 



WELCOMING REMARKS 



M. U. Clauser 

Provost 

Naval Postgraduate School 



It is with some feeling of nostalgia that I come "back into 
this closer contact with men from the aircraft industry. After 
having spent quite a number of years building airplanes , I am 
glad to tell you how very pleased I am to welcome you here at 
the Postgraduate School and how appropriate I think it is to have 
a conference of this sort at this time. I am sorry that Admiral 
Goodfellow is in Washington and is not able to also extend his 
most hearty welcome to you. We are also very thankful to the 
Naval Air Systems Command for the sponsorship of this symposium 
and we welcome Commander Cundari in the efforts to organize and 
help sponsor this work. 

I would like to take a few moments , because I have this 
sort of background and continuing interest, to speculate a little 
bit about the future but I would like to get a running start 
from the past. The 19^0's were certainly the great days for the 
aircraft industry as I remember. There was the dictate of 
President Roosevelt that we were going to build 50,000 airplanes 
and it really happened during World War II. Then came the very 
interesting transition to jet airplanes and even rocket airplanes 
and I can remember the time I spent on the design of the Skyrocket 
airplane which held the world's altitude record and several speed 
records and how we had to work our way through some new thinking. 
Then, in the fifties, we turned to rockets with all the trials 
and tribulations and in the sixties we had the ability to really 
have ballistic missiles and space programs, with the emphasis 
on complex systems and ultra-reliability. 

Now, as we come to the seventies, it is very interesting to 
see this return to emphasis on the aircraft but there are some 
trends which, it seems to me, pose a little bit of a dilemma. 
It is this problem of complexity. Costs have gone up as an 
exponential function. With our concern about the cold war and 
about our competition in the scientific war, as you might call 
it, and our space competition with Russia, somehow we always 
felt that a great nation could afford the increase in price. 
We seem to still want to make things even more complex but we 
must recognize that in the last few years the nation is no longer 
willing to support an exponential rise in costs . 

If you add together the R&D costs on such things as DoD, 
AEC, and NASA, you get a nearly constant growth rate of about 
20$ per year from World War II to at least the middle sixties. 
But at the same time, the population was growing only 1 or 2%, 
and the gross national product only about 5$- This trend cannot 
continue indefinitely. 

1-3 



Further, we probably can no longer accept a reduction in 
the numbers of airplanes we buy in order to be able to obtain 
the greater complexity. The military services clearly cannot 
afford to procure just one very costly airplane . I think the 
new situation throws a double burden on the designer. He must 
now make the equipment more effective and at the same time 
cheaper . 

There will be a great responsibility on the designer and 
many people will be watching as the realities of the situation 
do unfold. It is a real challenge to get people together who 
will think about how we are going to handle the complexity of 
design and at the same time can make the airplane continue to 
be more effective and not just keep going up in price . 

Here at the Postgraduate School, as in a number of universities, 
we certainly realize the needs and problems which our graduates 
will face. We recognize the obligation to give them the intimate 
knowledge to cope with the decision-making and operating problems 
of the future. So it is a pleasure to find a way to interact 
with the aircraft industry more intimately and it is a pleasure 
to have you here. Anything we can do to make your stay here more 
profitable, more enjoyable, we stand ready to do it. I thank 
you very much for the opportunity to welcome you this morning. 



I-U 



INTRODUCTORY REMARKS: 
THE NATURE OF DESIGN PROBLEMS 



Ulrich Haupt 
Naval Postgraduate School 



Design problems in aircraft structures have existed as long 
as there has been aircraft design. They usually found their 
solution by following a general pattern along the lines of 
engineering judgment and experience, experiments and analysis. 
In case of special difficulties the specialist had to exert 
special efforts, usually resulting in more experiments and more 
refined analytical methods . Adding up all these efforts over 
more than six decades, we see the results in form of jet trans- 
portation and space flight and there is good justification for 
taking pride in substantial accomplishments. 

Now, however, we have reached a stage where recent experience 
forces us to reappraise our situation. We feel that we are close 
to the boundaries of our present concepts and that the complexity 
of our problems is somewhat frightening. In the past, we could 
approach a problem by isolating it and applying all the available 
expert knowledge and experience. This is still the case for many 
of those present and future problems which are the bread and 
butter of the engineering profession. Yet the really challenging 
problems which we have to face become increasingly difficult and 
usually we cannot isolate them anymore. They are actually of a 
different nature. 

Think of our situation in fatigue. We began by isolating 
the problem of crack initiation. This took us to the problem of 
rate of crack propagation, then to the problem of fail-safe 
design, then to the problem of residual strength and fracture 
mechanics. In each one of these fields we have to deal with a 
subset of problems -- including material selection, structural 
concepts, stress level and stress concentration, manufacturing 
methods, load spectrum, and service life. Fatigue is a problem 
of detail design but it frequently has its roots in early 
decisions far upstream in the design process and may result in 
catastrophic consequences due to a special combination of 
circumstances way downstream in the service life. All of them 
are interwoven. 

In addition to such problems of detail design, we have 
another kind of problems which are typical of advanced design. 
Examples are : 



1-5 



How do we define the criteria when we want to apply 
newly developed materials to present-day aircraft -- 
considering weight, cost, risk, life cycle, uncertainties 
of load spectrum, etc? 

How do we optimize materials and structural configuration 
for thermal and chemical environments as complex as we 
have to face them for supersonic cruise or space shuttle? 

These challenging problems in detail design and in advanced 
design have in common that they incorporate a large number of 
parameters . Most of them are interdependent and many of them are 
not clearly defined at all. Beyond these aspects of interdependence 
and quantity, a new qualitative aspect is introduced. Up to now, 
the main emphasis in aircraft has been on airworthiness. For 
our new type of design problems, however, we have to combine 
airworthiness with design optimization under very complex 
conditions. This takes us into a new field. For both detail 
design as well as advanced design we have to develop an outlook 
which is much wider than the conventional knowledge, experience, 
and expertise of the specialist. This question of basic outlook 
is closely connected with design problems. 

-* -* * 

It may be helpful if we look at developments which have 
become visible in engineering education. Throughout the 1950 's 
increasing emphasis was given to science curricula until we 
realized in the later 1960's that engineering aspects had been 
seriously shortchanged. We came to understand the basic 
difference between science and engineering. The scientific 
approach starts from a given problem and proceeds by analytical 
and experimental methods -- simplifying and clarifying a problem 
to its skeleton in order to establish and understand basic 
principles. On the other hand, the engineering approach envisions 
a goal and proceeds by defining the problem, creating alternative 
possibilities, analyzing them by scientific methods and making a 
decision about the optimum solution — taking into account all 
the inherent complexities and practical consequences in order 
to find a practical answer. 

So we finally began to realize that engineering is more than 
applied science. The engineer must have a different attitude and 
a different viewpoint than the scientist. Slowly, much too 
slowly, there is a growing emphasis on creative engineering, 
value judgment, and interdisciplinary responsibilities. It 
will take some time until this process becomes clear and the 
results will penetrate into industry, but a new trend is visible. 



1-6 



There seems to be an analogy between the situations in 
education and in industry. Science and engineering in education 
have their counterparts in the professional activities of 
analysis and design in industry. In education we are realizing 
that engineering incorporates science as a most essential 
ingredient but has to reach into wider aspects. Correspondingly 
in industry design may include analysis as a most essential 
ingredient but again it has to reach into wider aspects. However, 
I should not belabor this point because some of my friends hold 
different opinions on it. 

There is full agreement that our analysts have been at the 
forefront of developments during the last few decades . In the 
field of structures they have, among many other things, developed 
a beautiful system of finite elements for basic airworthiness 
calculations and they are working on mathematical aspects for 
structural optimization. Yet our actual difficulties are not 
covered by these somewhat abstract considerations of airworthiness 
and optimization. Our difficulties are of a very practical 
nature with an endless number of real-life complexities. Beyond 
analytical aspects we have to 

recognize any possible mode of failure; 

evaluate uncertainties and risks; 

establish a value system for any optimization procedure. 

Here we have design problems in the fullest and widest sense 
of the word. For their solution we require an analytical mind 
coupled with an imaginative spirit. This is expressed in the 
designer's intuition as it has been applied in the past when 
the designer could handle a few parameters on his sliderule . 
Yet for our present design problems with an ever-increasing 
number of parameters we have to develop a new methodology. 

Such a new methodology is a challenge to do in a systematic 
way what we have been doing intuitively. Two different aspects 
can be considered: 

firstly, a technological effort to reduce the risk of 
structural failure; 

secondly, a methodical effort to establish a process of 
decision making under complex conditions. 

Regarding the first aspect, namely the technological effort, 
a decisive step is taken in the new Air Force program on Advanced 
Metallic Structures. This program is directed toward improved 
technologies in materials, structures, and manufacturing and 



1-7 



emphasizes the importance of a systematic transfer of information. 
It is an eminently practical, hardware -oriented program. As such 
it does not sponsor any ideas which are still somewhat vague in 
themselves. 

Regarding the second aspect, the methodical effort toward 
the decision making process, this is unfortunately still in a 
category of vagueness. Perhaps it depends mostly on becoming 
aware of the problem and on communicating about it. There will 
not be any quick and ready solution but it seems that fundamental 
elements could crystallize sooner if a common concern would be 
developed. A great effort is exerted by specialists regarding 
mathematical aspects of optimization but before we can ever hope 
to use any mathematical refinements, we have to clarify some of 
the most basic aspects of the decision-making process. 

Decision making is an essential part of design. Many 
variations can be found in other fields and it should be easy to 
point out how primitive the state of the art is . From the 
highest level of deciding about national policies and war or 
peace to the very personal level of choosing a marriage partner, 
important decisions are being made in a somewhat haphazard way. 
Any self-respecting gambler would estimate his chances more 
thoroughly. 

There is nothing mysterious about decision making and 
optimization. The basic requirements are an adequate problem 
statement and objectivity -- and this is where we usually fall 
short. Clearly defined values as well as uncertainties have to 
be included. The values will generally be of a variety of 
dimensions -- weight, cost, time, reputation, etc. Utility 
theory or similar methods of decision making provide a basic 
tool to express these mult i -dimensional values on a single 
scale. The corresponding detail manipulations can be left to 
operations analysts but every designer will have to acquaint 
himself with the underlying concepts . 

Notions like uncertainties and value systems still have a 
long way to go before they become household words in aircraft 
design. Yet it is the designer who has to establish a merit 
function, basing it on his experience and making it clearly 
visible so that any parameters can easily be submitted to 
scrutiny and modification. This field is still wide open and 
largely unexplored. It concerns both detail design and advanced 
design and represents design problems of a very different nature 
from those to which we have been accustomed. 

In the subsequent presentations and discussions we want to 
stay at first on familiar grounds and consider some specific 
problems of technology which are of much concern for new materials, 



1-8 



including fatigue, fracture mechanics and fail-safe. Afterwards 
we will see whether we can proceed toward some basic aspects of 
methodology. This should take us to some ill-defined problems 
which, nevertheless, may be of fundamental importance. 

There are many viewpoints which sometimes look quite 
different but frequently can be reconciled when some good will 
exists and a real effort is made. Good will can be created as 
soon as we become aware how important a problem is and how much 
we need a solution. How successful our efforts will be — that 
is often beyond our jurisdiction. All we can do is to make a 
sincere attempt -- and that is the purpose of this symposium. 



1-9 



TRANSLATING HIGH-STRENGTH STEEL CHARACTERISTICS 
INTO EFFECTIVE STRUCTURAL DESIGNS 



W. H. Sparrow 

LTV Aerospace Corp. 

Vought Aeronautics Division 



ABSTRACT: Development of test methods to characterize high- 
strength steels is discussed up to the present fracture 
mechanics methods. These test methods are related 
to service experience by specification requirements, 
structural criteria, and design practices. Examples 
are shown for parts which failed unexpectedly- -pointing 
out the need for materials and structures engineers 
to improve their ability to communicate with design 
engineers . 



I would like to review briefly some of the attempts of the 
materials engineers to characterize high-strength steels for the 
design engineer. For the purpose of this discussion, high-strength 
steels will be defined as those having an ultimate strength of 
200,000 psi or above, used in the range of -65° to ^00°F. These 
steels have yield-to-ultimate ratios from .80 to .98 with minimum 
elongations of approximately 5$> in 2 inches. 

My basic thesis is that we have not provided the designer 
with many quantitative numbers to work with for load-carrying 
capacity designs. We give him tensile strength and modulus and 
that is about all he can put into a formula. When you talk about 
things like ductility, impact strength, corrosion resistance, 
these are just comparative values. There is no ready way to put 
them into a numerical calculation and use them. 

I would like to go back, since I have about as much gray 
hair as most of you here, and think of the 30 's when we looked 
at the mechanical properties, including impact and ductility, 
of a new steel. If they were similar to one which had already 
a lot of good service life, you assumed that you could endorse 
it for design. Then, when we got into the war, much acceleration 
went on, and we used a big variety of steels with a lot of odd 
combinations of chemistry because of alloy shortages. We worked 
pretty much from the basis of tensile strength and hardenability, 
and we used all kinds of things which probably would make us 
shudder today. We also worked with a lot of specialized tests, 
including notched tensile fatigue tests. But again, we had a 
hard time relating those numbers to a specific design and 
idealized specimens bore very little resemblance to actual parts. 



1-10 



Now the latest thing is the fracture mechanics approach and 
here we are getting a little closer to the numerical idea. But 
the first thing for which we have to characterize a steel is K^ c 
and many of you know the war which is going on about the type of 
specimen to use and the wide variety of values for a given heat 
of steel. The designer is pretty much at the mercy of an empirical 
system when he is making a part out of a slightly different 
cross-section or form. He cannot make a clear-cut decision 
without a lot of specialized help and specialized work in the 
laboratory, and even with all that, we still have failures. 

In the early 1950' s, we got enamored with the idea that we 
could take some of our more conventional steels, specifically 
good old SAE ^3^0, and move them up to the 260,000 psi heat-treat 
level. A classical paper by Melcon of Lockheed-California* was 
published in 1953 and many of the aircraft companies, including 
mine, did work of their own and decided that this was a promising 
approach to get the weight down. We proceeded on the F-8 Crusader 
airplane, which was designed in the 1953 period and flew in the 
early part of 1955 and which had a considerable amount of steel 
parts in the 260,000 heat-treat level, particularly in the landing 
and arresting gear systems. 

We had some problems and I would like to review several 
typical examples to show you how the designer and the materials 
and structures engineers can get themselves into a trap. I am 
sure all of you have dozens more in your own experience. Some 
of the problems occurred fairly early in the programs; others 
took time to develop in the environment in the fleet. 

Fig. A-l shows the main landing gear inner cylinder shell, 
made of 260,000 heat-treat steel, with a clevis at the upper 
end. The interesting thing in this design was that the designers 
found the space in this upper cylinder could be used as a pneumatic 
bottle for an emergency actuation of flaps and gears in case of 
a hydraulic failure. This was not part of the landing gear system 
at all. Every time the engine was running, the cylinder was 
pressurized to 2000 psi and when the engine was shut down, it 
bled off slowly by leakage. Explosive failure occurred not during 
a landing but while the airplane was being pivoted on the carrier 
deck around one wheel which resulted in maximum stress at the 
clevis . 

We did not have an electron microscope at the time but it 
was obviously a very brittle failure. All the standard tests 
had been performed but we never had a failure nor any problem at 



*Melcon, M.A. , Ultra High Strength Steel for Aircraft Structures, 
Product Engineering, October 1953- 



1-11 



all with this part in the plant at any time. No matter what we 
did, we could not explain at that time exactly what was going 
on. Being 260,000 heat treat, we did not use electrolytic cad 
plate on it. We knew about hydrogen embrittlement, so we tried 
to keep away from that area. At that time, vacuum cadmium 
plating was not developed and so we just had an organic finish 
on these parts. 

Subsequently, when we got the electron microscope several 
years afterward, we found out that some minor scratches in the 
paint had permitted the part to corrode in a very small area, 
about 10 or 15 thousandths in diameter. The little bit of 
corrosion resulted in enough hydrogen to hydrogen-embrittle 
the part in that local area and also to give it some stress 
corrosion. The designer had veen quite ingenious but the 
combination of maximum stress in the clevis, hoop tension due 
to internal pressure, H- embrittlement, and minor stress corrosion 
resulted in failure. The part is now vacuum cadmium plated and 
painted and is no longer pressurized as a pneumatic bottle. 

Fig. A-2 shows the horizontal stabilizer shaft which is 
heat-treated to 260,000. Through a series of offset tapered 
pins, a little over an inch in diameter, it is attached to the 
actuating horn of Fig. A-3. We found a few of these shafts 
cracked and after some checking --it was also in the pre -electron 
microscope era -- we decided that a washer at the tapered pin 
might be getting misaligned. So we designed a specially shaped 
washer to fit the inside contour of the shaft and to avoid local 
high stresses. We also added some more paint in local areas 
and did not seem to have any more problems. 

Then, about six or eight months later, we found some cracks 
again. We could not put our finger on the definite thing but 
we made a few more changes and still have had occasionally some 
cracked parts. Examination of these later parts by electron 
microscope revealed hydrogen embrittlement and we have tried 
to trace the source of it and to keep control over the processing. 
The part is vacuum cadmium plated and does not get any electro- 
lytic cadmium. So here we have a design which, from the 
designer's point of view, looked like an efficient and good way 
to get high strength and close tolerance and where impact and 
ductility values did not indicate any reason why this would 
not work. 

Fig. A-3 shows the stabilizer horn which is also a 260,000 
heat-treat part. For this part it turned out that the 260 
high heat treat was specified because the structures engineers 
were told by value engineering that this would cost hardly any 
more. Like most structures people, they were looking to a growth 
version of the airplane and left an extra margin where it can 
be done without overpenalizing design and weight. So this 



1-12 



is one part where we moved the heat treat back to 200. With 
U3U0 steel you have to stay out of the 220-2U0 range for 
metallurgical reasons, just as a precautionary move. 

Then, recently, we had a failure again. The reason was 
typical for the way things can happen. We had a new vendor 
who came out a little high in hardness . He pulled the test 
part which went with it, got the ductility, and felt there was 
no problem because he was giving a little more than required. 
However, when we dropped the heat-treat to 200,000, we decided 
that we could electrolytic ally cad plate it because, generally 
speaking, there is not any hydrogen embrittlement problem with 
200,000 heat treat. Now with 246,000 heat treatment and electro- 
lytic cad plating, we had hydrogen embrittlement coming out the 
ears. Again the designer did his job but there was not enough 
communication about the processing precautions with the 
manufacturing people. 

Fig. A-k is a swinging arm for the arresting gear at the 
260,000 heat treat level. It gets a certain amount of impact 
and we had failures as shown. This was a good old traditional 
case where we had a big discussion among the structures, design 
and material engineers about how much interference fit these 
pressed-in bushings could have without getting above the thresh- 
hold stresses for stress corrosion cracking. However, in cad 
plating there was some tolerance which could get involved 
occasionally. You will say in pressing them in you should have 
shaved the excess cad plating off. But we had a manufacturing 
operation which deep-freezed the bushing and dropped it in. 
So they shrunk it down to under size and did not have to get the 
cad plate shaved off but when it came back to room temperature, 
we were above the stress allowables by just enough to get us 
into trouble. 

This part is now 200,000 heat treat. There seems to be 
some magic about this number. Things like hydrogen embrittlement 
and stress corrosion are not very critical at this level. For 
any bushings which we have now in high-strength parts, both 
aluminum and steel, we try to prevent water entrance. We provide 
a section around the bushing for a sealant groove to keep the 
water out of the interface between the bearing and the fitting 
even though it is press fit. We forbid deep freezing bushings 
because of the tolerance and condensed water problem. 

Fig. A-5 is a 260 H.T. bellcrank on a droop mechanism. It 
was designed for stiffness and did not need the 260,000 strength 
but it failed in a fatigue sense. This was pretty much an 
interior part when it was on the deck and the droop was retracted. 
However, there was full flowing air going around this part 
during the extended position and it got enough corrosion to 



1-13 



provide a pit. Again, the heat treat was dropped down to 200,000 
and then the corrosion effects did not appear to be as critical 
in fatigue. We also went to more protection, did a little 
aluminum spraying in the beginning, and then went finally to 
electrolytic cad. 

Fig. A-6 seems to be hard to explain. This is the main 
fin beam in both the F-8 and A-7 airplanes and it is about 
^0 x 35". It is forged from a U3U0 billet and in the center 
where normally the worst material would be, the part is split 
in two directions. There are holes drilled into it, and it 
has only primer for protection. The part is 260,000 heat treat, 
has 3 million flight hours on it, and has never failed. By 
the way, the note "crack" in the figure only refers to a specimen 
which they tried to straighten during a heat treat operation. 

Now you could say that the reason for no failure is that 
this part is lowly stressed or is in a mild environment. Well, 
this part does not have a sustained load like in a pressure 
cylinder or in a pressed-in bushing. It is designed for a 
gust condition which rarely occurs and the total spectrum life 
is easy. But since you have to stay out of the 220-2^0 range 
for this type of steel, the next jump down to 200 would have 
been too big. The design also has no sharp corners and the 
holes are straight for standard close tolerance bolts. Besides, 
it is an interior part in warm, dry air and is protected from 
the engine heat by a stainless steel shield. Even with all 
these good things I am still surprised that we have a zero 
failure rate. 

This last example shows that you should not panic if a 
catastrophic failure occurs on some part at the same strength 
in the, same steel. We had some other parts in the airplane 
which were not loaded any higher but were in more exposed areas 
and failed. It is hard to tell the designer when he is really 
in trouble about corrosion unless the environment can be very 
accurately defined. 

I might say that we have had some cleanliness problems in 
steel but we tried to control them by specifying values for 
reduction of area in specified locations in each heat of steel. 
This seemed to give the fewest problems along the way in 
unpredicted failures. We felt this was our protection but, 
at the same time, how do you explain an empirically picked 
number to the designer? 

Well, I just wanted to show you some histories of how we 
have given the designer some problems and have complicated 
them right along. Designers and structures engineers need to 
understand the effect of stress raisers, residual stress due to 



I-lU 



processing and induced assembly stresses due to clevis type 
parts or other misalignments. Effects from stress corrosion 
and hydrogen embrittlement have to be considered in selecting 
strength level. When working with high-strength steels, 
process engineers, designers, tool planners, and manufacturing 
personnel need to be educated in the catastrophic effects of 
grinding, plating, cold-straightening, cleaning, weld repairing, 
etc., without proper post stress relief treatments. 

The designer also needs more education along the lines of 
NDT so that he knows what defect size can be detected. Even 
if we are great believers in magnaflux, you still can magnaflux 
a part and not find a crack which is there. I think, almost 
intuitively, our own people have backed up from welding ideas 
in high strength steels quite a bit. Talking about steels of 
300,000 psi, they are wary of sustained stresses and are trying 
to keep the stress level down at a fairly conservative level. 
That hurts from the weight and cost point of view pretty fast. 
We worry not only about things that happen in our own shop 
but we are also indoctrinated by the troubles everyone else 
has. Risks are involved in this area and when you want to move 
off to something new, you have to convince not just your engineering 
and manufacturing management but you are involved pretty strongly 
with the operating management even when they are not engineering 
oriented. 

Most failures in high- strength steels result from three 
major sources, namely: (l) improper design; (2) improper 
processing; (3) undetected flaws. Now we are giving the designer 
a new input from fracture mechanics, with new tools and new 
concepts and there is no table where he can look up the answers. 
Few failures are the result of fracture toughness per se but 
its effect on premature failure must be understood. 

I would like to conclude in saying that there is no 
substitute for communication. We have to get engineering, 
manufacturing, and quality control together but I think we 
also will have to develop a more definitive and quantitative 
way of characterizing high-strength steels than the present 
empirical data. 



1-15 



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1-21 



FATIGUE -- RELATING PAST EXPERIENCE 
TO DESIGN 



J. C. Ekvall 
Lockheed- -California 



Abstract: Three general approaches are used to design aircraft 

structures for a long service life: control of design 
details, establishment of design stress levels, and 
analysis combined with experiments. These approaches 
are discussed and illustrated by examples from previous 
experience. Examples include the use of the stress 
severity concept and how it relates to fatigue life, 
fatigue quality index, and design stress levels of 
similar structures -- emphasizing that determination 
of safe life is not yet an exact science and must be 
based on previous experience. 



The previous speaker gave us some examples of how to make 
use of experience. As you know, there is no good theoretical 
approach for predicting fatigue life of an aircraft component 
and so, in order to circumvent this problem, previous experience 
is used to form a baseline from which to work. 

Structures which have performed adequately over a long 
period of time provide assurance that a similar design for a 
similar application will perform just as well. However, problems 
arise when trying to exploit new materials and new technology 
where previous experience is lacking. In this case, experience 
must be supplemented by adequate test data, although it is not 
always easy to obtain test data applicable to what is going to 
happen ten or twenty years hence . In this presentation I would 
like to discuss how our previous experience is used to the best 
advantage in coming up with new designs and how it relates to 
the overall problem of design, detail design and advanced design. 

Figure B-l illustrates how service experience is utilized 
in the design of new aircraft and where the information is used 
in relation to each phase of the program. A satisfactory service 
life can be achieved by approaching the problem on three 
different levels : 

a. Control of aircraft design details -- based on service 
experience of design details from previous structures; 

b. Establishment of design stress levels — based on stress 
levels utilized in previous aircraft structures; 



1-22 



c. Theoretical and experimental considerations -- for new 
designs where previous experience is not available. 

For any new aircraft structure each of the three above approaches 
must be utilized to achieve a good design from a fatigue point 
of view. Therefore, I would like to illustrate the use of these 
three approaches with some examples. 

In the first approach, information gathered from experience 
has to be presented in a simple format and brought to the 
designer's attention so that he will look at the data and try 
to avoid these problems . Figure B-2 shows a common type of 
design detail in connection with access holes. Small holes in 
the vicinity of a bigger hole inevitably mean superposition of 
stress concentration. An acceptable solution can be achieved 
for many applications by eliminating the fastener hole in the 
vicinity of highest stress concentration due to the cut-out. 
For the preferred solution of using a clamp-on door, consideration 
must be given, of course, to the effect of fretting fatigue due 
to rubbing action between the clamp-on door and the structure. 

Figure B-3 shows another type of problem which is often 
encountered in service and which consists of cracks in the 
corners of door cut-outs. Attempts have been made to try to 
reinforce these areas to reduce the stress levels and prevent 
the crack; however, this has not always been successful. The 
best approach in this case is to make a fairly generous radius 
in the corners of these types of cut-outs, using the rule -of - 
thumb shown in Fig. B-3- With this type of design concept, 
fewer cracking problems have developed in service . 

Figure B-k illustrates another area which may easily be 
overlooked as not a structural problem. A forging of this type 
needs to be attached to some type of jig for machining to final 
dimensions . Tooling holes are drilled at various locations 
for the purpose of holding the forging to the jig during 
machining. Often these tooling holes are located where the 
bending stresses are essentially zero. However, the shear 
stresses in a beam are usually high where the bending stresses 
are low and fatigue cracks due to high shear stresses have 
occurred at these locations. Also, often the tooling holes do 
not have as good a surface finish as attachment holes in other 
parts of the structure. For example, tooling holes may have 
scratches or other imperfections which amplify the stress 
concentration effect of the hole. As shown in Fig. B-k, plugging 
tooling holes with interference fit fasteners can increase the 
fatigue life 1+- fold while reducing the fatigue quality level, 
K^est? "k° less than k.O. On our new designs we prefer to reduce 
the stresses by reinforcing locally around the hole. 



1-23 



Figure B-5 refers to a problem which is often overlooked. 
This is fatigue cracking as a result of deflections from one 
part of the structure induced on another part of the structure . 
In this example the nacelle fillet region had a doubler attached 
to the nacelle skin to prevent localized buckling. To prevent 
fatigue cracking at the end of the doubler it was necessary to 
extend the doubler and install additional rows of fasteners to 
stiffen up the area. 

These simple types of examples provide the designer with 
some basic information as a first line of defense against fatigue 
cracking problems. Next, let us consider the selection of 
design stress levels. There are two ways of specifying design 
stress levels, either by putting limitations on the design 
ultimate tensile stress of the material or on the stress due to 
lg loading conditions . 

Figure B-6 shows one technique of selecting preliminary 
design stress levels. Based on service experience on previous 
aircraft, it relates maximum allowable design stress level to 
the number of flights when fatigue cracking problems were 
encountered. Most designs fall within the cross hatched band, 
therefore the percentage of ultimate tensile strength which can 
be utilized for a new design depends on the fatigue life that 
is desired in future service. 

The results of spectrum fatigue tests of components can 
also be utilized to aid in the selection of design stress 
levels. The fatigue quality index, K, is determined from 
fatigue test results as illustrated in Fig. B-7. The analysis 
is based on the test spectrum which was applied to a given 
structural component up to the time when a fatigue crack was 
initiated, and a set of S-N curves for various Kj. values 
representative of the material in which the crack was started. 

With these data, a fatigue analysis is conducted for several 
Kj. values, determining S n/N for each case, and finally 
calculating by interpolation on the K^ value which corresponds 
to E n/N = 1. This K-t value is designated the fatigue quality 
index K, of the structural component. 

A number of components have been analyzed and compiled as 
shown in Fig. B-8. This Figure shows the distribution of K 
values obtained from h-Z test results. The fatigue quality index 
ranges from slightly below 3-0 to above 5.0, with a mean of 
3.65 or 3.7. So we can say that the fatigue quality of aircraft 
structures is something a little worse than an open hole which 
would have Kt = 3.0. For a new design, you might not want to 
use the average K value as a basis for structural design since 
you would have only a 50$ chance of passing a fatigue test. 
Therefore, we usually try to design for a fatigue quality level 



I-2U 



of about U.O. In some areas of the structure a poorer quality 
or higher K value can be tolerated because the stresses have 
to be low for reasons other than fatigue. This type of analysis 
is just a one -parameter (K) analysis. There are a lot of other 
variables that have an effect on fatigue life, but with k3 
results, you really do not have enough data to consider more 
variables in the preliminary stages of design. 

Figure B-9 shows how to use the K value to arrive at a 
preliminary design stress level. For a new application, you 
first develop spectrum loading conditions which the structure 
is likely to encounter during its service life . Then you 
perform a fatigue analysis with the same S-N data that was 
used for the analysis of fatigue test results. If various 
materials are being considered, you can conduct the analysis 
using S-N data for the various materials to obtain a relationship 
between the design stress or reference stress and the fatigue 
quality index as shown in Fig. B-9- For this particular example, 
if you picked a K value of k as being the type of quality you 
think can be achieved in the structure, then a ^3,000 PSI 
stress level would be permitted for 202U-T3 aluminum alloy. 
This analysis was conducted for a particular service life so 
that the values in Fig. B-9 are all for the same number of 
flights or the same number of flight hours anticipated in 
service . 

Assuming you have done the best job you can and are coming 
up with satisfactory design stress levels and design qualities, 
however, there are still new materials, new fasteners and new 
approaches that you want to apply to the structure for which 
you have no previous experience. In this case, you will have 
to conduct various types of tests to develop information and 
correlate the data with similar tests on structures in which 
you have confidence. So, let's look at some of the types of 
tests that are conducted and the results of some example cases . 

Figure B-10 shows some results of two -row lap joint tests 
with aluminum countersunk fasteners. The peculiar S-N curve 
is the result of three types of failure. In the low-cycle 
high-stress region, the failure occurs where you would normally 
expect it, i.e., originating at the center of a fastener hole 
where the highest stress concentration is located. In the 
transition section, failure occurred in the countersunk sheet 
away from the fastener holes, and in the high-cycle low-stress 
region, the failure took place in the plain sheet also away 
from the fastener ho3.es. In the high-cycle region (> 300,000 
cycles), the cracks initiated at the edge of the attachment 
holes in the interface between the two sheets. This is where 
the greatest rubbing action takes place and failures were caused 
by fretting fatigue. The data in Fig. B-10 illustrates dangers 
of extrapolating from previous experience, say less than 



1-25 



300,000 cycles, into a different region of life expectancy. For 
this example, you would get over a 50$ reduction in stress 
level over what you'd expect with a normal extrapolation from 
low cycle part of the S-N curve. 

Fatigue tests have been conducted on simple lap joints 
using lubricants, shims, etc. to try to eliminate the fretting 
fatigue type failure. Aluminum bronze between mating surfaces 
worked pretty well, teflon films did not work very well. How- 
ever, for each application, tests have to be conducted to find 
out each technique that will work because with fretting fatigue 
the same technique doesn't always work for different applications. 

Figure B-ll is another example of a fretting fatigue 
failure. This is the case of a countersunk hole in the skin of 
the fuselage and the example shown is from a pressurized panel 
test. Fatigue cracks initiated in the black area in the 
transition region from the countersunk to the cylindrical portion 
of the hole opening due to the flexure and fretting action 
when the fastener was rubbing against the sheet material in 
this general region. To avoid this type of problem, it is 
necessary to install more fasteners or larger fasteners to 
reduce the bearing and bending stress . 

Figure B-12 shows the results of a series of tests on 
three-row lap joints conducted to investigate what type of 
fastener might be best for a given application. None of these 
failures represent bolt failures . The normal theory for the 
design of mechanically fastened joints is that if you use 
interference fit fasteners, you are going to get an improvement 
and if you use high clampup, you are going to get an improvement. 
So, combining these effects, you should get a terrific improvement, 
Oddly enough, the results did not come out that way. We got a 
very small amount of difference between the various fasteners 
tested. Only the two upper fasteners gave a little better 
result; however, it still isn't a dramatic improvement. The 
main reason for this anomally, I think, is again that fretting 
fatigue failure is coming into the picture and it is limiting 
the degree of improvement you can get by changing the fastener 
and the interference. Therefore, the state of the art of the 
material fastener system was pushed to the point where fretting 
fatigue type failures were occurring and something else must be 
done to avoid this problem. 

Figure B-13 shows the results of some other tests, where a 
series of tests were conducted to evaluate various design 
alternatives in order to come up with a good design. This 
example is a typical fuselage joint encountered in an aircraft 
structure. A continuous stringer is attached to the skin and 
the continuous frame inside of the stringer is attached to the 



1-26 



skin by an angle which is cut out in the stringer region. The 
first concept was the simplest one without any attachment between 
stringer and frame. Crack initiation was at 10,000 cycles in 
the angle at the rivet attachment to the stringer. 

The next design used a slightly different stringer shape 
which allowed a rivet attachment between stringer and frame . 
This gave an almost 9-fold improvement in crack initiation time. 
The crack was in the stringer at the rivet attachment to the 
frame which moved the fatigue critical point away from the skin. 

A third design resulted in an additional 10-fold improvement 
in crack initiation time. This was due to an added clip on the 
side of the frame so that the load was more evenly distributed 
across the stringer. This shows how simple joints can be 
developed in the laboratory by running simple tests . This is 
probably the best way to go in many cases since the analysis 
is unreliable and rather difficult. It does indicate what 
good detail design can accomplish for you. 

Another approach for coming up with a good design is to 
use finite element analysis techniques. Figure B-lU shows 
four different designs for a wing-to-fuselage joint with tapered 
stringers. The first two designs have a combination shear and 
tension connection, the last two designs are double shear 
joints. Normally, a tension-shear type connection is not a 
very good design from a fatigue point of view, but you can make 
it work provided you put enough material in the right places. 
For this particular application, it was considered because of 
the simplification of final assembly. 

The finite element analysis takes into account the 
flexibility of the fastener systems, so the load transfer 
from skin to stringers is more or less properly accounted for. 
The resulting stress concentration or stress severity factors 
are shown at various places in Fig. B-lU. The finite element 
analysis model is very useful for looking at various methods 
of shaping the material and for reducing the stress concentration 
down to a minimum. It permits one to come up with more or less 
an optimum design for each concept instead of running many 
fatigue tests. The example shows how a good design can be 
evolved using computer analysis techniques in conjunction with 
a minimum amount of test data. 

The finite element type of analysis results in a stress 
severity factor and does not really predict the lifetime. So 
a correlation was developed between the stress severity factor 
(obtained from finite element analysis) and the fatigue quality 
index which is obtained by fatigue analysis of test failures. 
The correlation is shown in Fig. B-15 which indicates that a 
stress severity factor of about 3 corresponds to a fatigue quality 
index of about k . 



1-27 



After design details and stress levels have been established 
and basic testing has been conducted, the final proof comes by 
conducting a full-scale fatigue test on the complete aircraft 
structure. There you integrate the effects of adjoining structure 
which you cannot simulate adequately in the laboratory on 
components. Fatigue is not an exact science yet, but much can 
be learned from experience and applied to new types of design. 
The use of this experience in conjunction with a suitable fatigue 
test program will provide a reasonable assurance of adequate 
service life. 



1-28 



INFORMATION USED 



METHOD OF ACHIEVING 
REQUIRED SERVICE LIFE 



PHASE OF 
PROGRAM 





OLDER STRUCTURE 
GOOD FATIGUE 




DESIGN 
EQUAL 


NEW STRUCTURES 
OR SUPERIOR TO 






QUALITIES PROVEN 
IN SERVICE 


OLDER STRUCTURE 


, 








< 






ADOPT ALLOWABLE 
STRESS LEVELS FROM 
PAST EXPERIENCE 




PROVIDE SUFFICIENT 
MATERIAL TO KEEP STRESSES 
BELOW ALLOWABLES 






i 


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CORRELATION BETWEEN 
TESTS ON PREVIOUS 
AIRCRAFT STRUCTURES 
& SERVICE EXPERIENCE 




CONDUCT FATIGUE ANALYSIS 
AND TESTS TO DEMONSTRATE 
THAT SERVICE BEHAVIOR 
WILL BE SATISFACTORY 








'! 






SERVICE 












EXPERI 


ENCE 



PRELIMINARY 
DESIGN 



DESIGN 

DEVELOPMENT 

AND QUALIFICATION 



OPERATION 



BOW SERVICE EXPERIENCE IS UTILIZED 
IN THE DESIGN OF NEW AIRCRAFT 



Fig. B-l 




HOLE AT POINT 
Of HIGHEST STKHS 
CONCENTRATION 



AVOID 



NO HOLE IN REGION 
OF HIGHEST STRESS 
CONCENTRATION 



ACCEPTABLE 



PREFERRED 



AVOID PLACING CUTOUT ATTACHMENTS AT POINTS OF 
HIGH STRESS CONCENTRATION 



Fig. B-2 



1-29 



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1-30 



NACELLE 



WING-NACELLE 
FILLET 




A 



NACELLE 
DEFLECTIONS 




FILLET ATTACHMENT SCREWS 



NARROW FILLtT SKIN DOUBLER 
(2 ROWS OF SPCTWELDS) 



FILLET SKIN 



WIDE FILLET SKIN DOUBLER EXTENDED TO HERE 
(4 ROWS OF SPOTWELDS) 

SECTION A-A 



LOCAL BUCKLING OF THE WING-NACELLE FILLET, INDUCED BY DEFLECTIONS OF 
THE NACELLE, CAUSED FATIGUE CRACKING IN THE FILLET SKIN AT THE EDGE OF 
THE NARROW FILLET SKIN DOUBLER 12 ROWS OF SPOTWELDS). THE DESIGN WAS 
IMPROVED BY THE USE OF A WIDE FILLET SKIN DOUBLER 14 ROWS OF SPOTWELDS). 
NO CRACKS OCCURRED WHERE THE WIDE FILLET SKIN DOUBLER WAS INCORPORATED. 



EFFECT OF INDUCED DEFLECTIONS 



Fig. B-5 



1-31 



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CHEAP AND 
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DATA NORMALIZED FOR 
ALUMINLM ALLOYS - 



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10 10' 

SERVICE LIFE, FLIGHTS 



10 



Fig. B-6 



1-32 



STRESS 



STRESS 




REAL SCALE FATIGUE DATA 



N(REAL SCALE) 

STANDARD S-N OATA FOR MATERIAL- 
CONDITION, MEAN STRESS, ENVIRONMENT 

FATIGUE QUALITY INDEX "K" IS THE 
INTERPOLATED S-N CURVE = K 

WHICH MAKES S -77 = 1.00 

N 




N CURVES 



LOG In OR N LOG N 

NUMBER OF FLIGHTS OR CYCLES 



LABORATORY TEST DETERMINATION OF FATIGUE QUALITY INDEX 



Fig. B-7 



MBAN OF AIRCRAFT COMPONENTS 



BASED ON 42 TEST 
RESULTS 



NUMBER OF 
TESTS 



LESS THAN 1C£ 




3.0 4.0 

FATIGUE QUALITY INDEX, K 



LESS THAN 10$ 



5-0 



DISTRIBUTION CURVE FOR AIRCRAFT STRUCTURAL COMPONENT TEST VALUES OF K 



Fig. B-8 



1-33 



70 



6o 



50 



to 



30 



FATIGUE CRITICAL LOCATION OH TRANSPORT 
WING LOWER SURFACE 




FATIGUE QUALITY INDEX ~K 



RELATION BETWEEN FATIGUE QUALITY INDEX-K 
FOR A GIVEN FATIGUE LIFE 



Fig. B-9 



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FAILURE IN COUNTERSUNK SHEET 
THROUGH THE FASTENER HOLES 



FAILURE IN COUNTERSUNK SHEET 
AWAY FROM THE FASTENER HOLES 




FAILURE IN PLAIN SHEET 
AWAY FROM THE FASTENER 
•TOLES 



-*-»- NO FAILURE 



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J I I I I I I I 



10 5 10 6 

CYCLES TO FAILURE 

FAILURE MODES FOR LAP JOINT SPECIMENS 



10' 



Fie;. B-10 




FRETTING FATIGUE 



Fie. B-ll 



1-35 



Ti-1-8-5 TLB-960 
HL-51 Hi-Lok 

Ti-6*4, TLV-100 
(Taper-Lok) 

Broach Bolt 

Ti 6-4, HL-11V 

(Hi-Lok) 

TL-100 Taper-Lok 

Huck Crimp 

Steel Hi-Tigue 

HL-51 Hi-Lok 
Ti-Hi-Tigue 







80,120 (Standard Interference) 




52,680 


(Clearance Fit) 




1+4,510 


(Std. Interference) 




42.465 




41,370 


(Clearance Fit) 




38.330 


(Standard Interference) 




36.580 


(Ream and Deburred Hole) 






34.920 


(interference Fit) 






34,870 


(interference Fit) 






33,410 


(Deburred Holes - Interference Fit) 


1 1 1 


l 1 1 1 1 1 1 



50 

FLIGHTS TO FAILURE - 100 

FLIGHT-BY-FLIGHT SPECTRA FATIGUE TEST RESULTS OF 
TITANIUM AND STEEL FASTENERS 



100 



Fig. B-12 



1-36 



TYPE OF 
DETAIL 



SPECIMEN CONFIGURATION 
(CONSTANT AMPLITUDE LOADING AT R = 0) 



CYCLES TO 
CRACK 
TNTTTtTTnn 



CYCLES TO I 
FAILURE 



NO ATTACHMENT 
BETWEEN FRAME 
AND J-STRINGER 
• STRINGER 




T 



CRACK IN ANGLE 
AT RIVET 
ATTACHMENT 



10,000 



63,800 



t t 



3F 



i— STRINGER 

f ami-; 



f 

SIDE VIEW 



Z-ST RINGER 
RIVETED TO 
FRAME 



,— ■ STRINGER 





STRINGER AND 
CLIP RIVETED 
TO FRAME 



W 1 



+ * * t 






^Z 



U) 



CRACK THRU 
STRINGER AT 
RIVET ATTACH- 
MENT TO FRAME 



f SIDE VIEW 



(TOP VIEW SAME AS ABOVE) 

CRACK INITIATION IN 
.STRINGER AT RIVET ATTACH 
MENT 




FAILURE 
TH.TU STRINGER AT 
RIVET ATTACMENT 
TO FRAME 



119,000 



3,000 



DEVELOPMENT FATIGUE TESTS OF FUSELAGE DESIGN 1ETATL 



Fig. B-13 



1-37 



TYPE OF 
JOINT 



TOP VIEW OF 
DESIGNS 



ORIGINAL DESIGN 
LIGHT 
TENSION-SHEAR 

REDESIGNED 

HEAVY 
TENSION-SHEAR 



SPECIMEN CONFIGURATION 




mimit rwirj 




— N0 



JOT.1 BH rjDIM 



TEST LIFE 
FLIGHTS TO FAILURE* 



16,600 



STRINGER 74,000 
SKIN, 36,000 



REDESIGNED 
LIGHT SHEAR 




INITIAL 92,200 
FINAL 111,520 



fTRDCIB CWCI 



REDESIGNED 
HEAVY SHEAR 




INITIAL 99,000 



* FATIGUE TESTS CONDUCTED UNDER REPRESENTATIVE FLIGHT -BY-FLIGHT FATIGUE TESTING 



Fig. B-lk 



STRESS 
SEVERITY 
FACTOR, 
SSF 



K = 0.8 + 1.2 SOT 




RAJTOI OF VALUES 



BOTES: K VALUE CALCULATED FROM FATIGUE TEST 
RESULTS OF PANEL TESTS SO THAT 
r n/» = 1.0 



2 — 



3 U 5 6 

FATIGUE QUALITY INDEX - K 
CORRELATION BETWEEN STRESS SEVERITY FACTOR AHD FATIOUE 



Fig. B-15 
1-38 



FRACTURE MECHANICS CONSIDERATIONS 
IN DESIGN AND MANUFACTURE 



¥. C. Dietz 
General Dynamics - Fort Worth 



Abstract: Problems encountered on the F-lll in detail design and 

inspection procedures are discussed; subsequent programs 
toward establishing service life integrity and fracture 
mechanics data are described; and methods for deter- 
mination of inspection intervals and crack retardation 
are given. The designer's problem is summarized as 
shown in Fig. C-39, ^0 -- emphasizing the need for 
intimate association of the designer with manufacturing 
and inspection operations . 



As you are all well aware, there are many basic considerations 
involved in structural design. These have been briefly summarized 
and shown on Fig. C-l. I would like , however, to discuss two 
specific items concerned with service life, Detail Design and 
Manufacturing, as they have been responsible for most of the 
difficulty in service life structural certification of the F-lll, 
and in fact, I believe are the major contributors to service 
life problems on most aircraft. 

Before discussing these items in detail, I believe it is 
important to get a feel for the significance of the structural 
problems which have occurred in military aircraft service usage 
and to view them from the proper perspective. An examination 
of the statistics covering a period of about eight recent years 
indicates the incidents of inflight primary structural failures 
of military aircraft as shown on Fig. C-2. Only a fraction of 
the failures shown have resulted in loss of the aircraft but the 
magnitude of the problem is evident. It is also apparent that 
the majority of failures occurred in the wing structure. 

Fig. C-3 shows the loss statistics for USAF fighters due 
to all noncombat causes. An average yearly loss of 139 airplanes 
translates to a little over 11 airplanes destroyed per 100,000 
flight hours. More specifically the losses due to structural 
failures have been about one per year or an average rate of 
.103 per 100,000 flight hours, or just under 1% of the total 
USAF noncombat fighter losses. Fig. C-h presents the same 
data by aircraft type. 

I would now like to discuss a few of the problems which 
occurred in the structural history of the F-lll during structural 
certification and service usage. The most significant of these 
have been associated with the high heat treat steel parts and 



1-39 



have been the result of the detail design or the fabrication and 
inspection aspects. Most of the problems, as shown on Fig. C-5? 
were isolated in the laboratory during full-scale fatigue tests 
or during the course of inspection and proof test of the airplane . 
However, there has been one inflight structural failure. You 
will note that our experience is in part similar to the experience 
of LTV described in an earlier presentation in that three of the 
problems were due to stress corrosion cracking, none of these 
were service induced. 

There is a considerable amount of high heat treat steel 
used in flight critical parts on the F-lll airframe. The 
definition of critical in this case is a part or component 
essential to maintaining structural integrity of the primary 
flight structure. The total usage of steel represents approximately 
20$ of the structual weight of the airplane or about 5*000 pounds. 
As shown in Fig. C-6, the use of steel in critical parts is 
essentially concentrated in the heavily loaded structures, such 
as the wing carrythrough box, wing pivot fittings, tail support 
bulkhead and longerons . With few exceptions , the steel employed 
in the F-lll is D6AC, a derivative of an alloy which was initially 
developed by Ladish. It is similar in many of its characteristics 
to U3U0 except it does not exhibit a blue brittle heat treat 
range. Extensive use is made of welding on the three major wing 
fittings; however, this has not been a source of any service or 
test problems . 

The F-lll structural test program included a full-scale static 
test and a full-scale fatigue test article. Fig. C-7 shows the 
fatigue test article set-up for testing. The fatigue test 
originally was planned to be conducted, as shown on this chart, 
on a complete airplane including wing, fuselage and tail. As 
the program progressed it was revised to test individual major 
components to expedite the program. This was particularly 
feasible for the F-lll due to the nature of the wing attachment to 
the fuselage inherent in the variable sweep design. 

The first failure in the airplane fatigue test program 
occurred in the wing carrythrough structure approximately 2.\ 
years ago, as shown on Fig. C-8. Approximately 1800 cycles of 
kg load had been applied when failure occurred in this wing 
support structure. The origin of the failure was a taper-lok 
bolt which attaches an aluminum door to the rear spar structure. 
The crack at the point of unstable crack growth was approximately 
.2", and progressed across the sculptured lower plate of the box. 
In the bolt hole adjacent to the origin of the failure, it was 
discovered on post-mortem examination that a fatigue crack also 
existed. 

In view of this early failure, an extensive investigation 
was undertaken, and it soon became obvious that there were two 
problems. As shown in the magnified pictures of Fig. C-9, it 
was found that some rather rough drilled and reamed holes 



I-UO 



existed. The profilcorder measurements shown indicate the 
surface condition. Measured with a profilometer, the surface 
finish is RMS of 200 to 250. There was also concern about the 
stress levels in view of the evidence of fatigue in the adjacent 
hole to the failure origin, and a stress survey was conducted 
in the steel flange corner radius. Strain gages were installed, 
as shown on Fig. C-10, on the flange and slug area not only on 
a full-scale test specimen which was set up specifically for this 
purpose, but also on the static test airplane and on a flight 
test aircraft. With a nominal stress of approximately 80,000 
psi in the slug area, a factor of about 2.5 times this value 
existed on the top edge of the flange . 

The solution therefore involved both design and manufacture. 
Figure C-ll is a simplistic chart but it makes the point. With 
the high stress concentration as measured and the rough hole 
surface finish a low fatigue life is indicated. It was necessary 
to lower the stress level and to reduce the taper -lok bolt loads 
by design, and additionally, revise the manufacture technique 
to improve bolt hole finish and obtain better bolt fits to 
achieve a lower effective stress concentration. 

As quite a large number of parts had already been manufactured, 
retrofit considerations were of prime importance. A considerable 
amount of analytical and test effort was required to develop a 
simple retrofitable fix that could be applied without causing 
other problems within the constraints of minimum rework to the 
high heat treat steel structure. The final design is shown in 
Fig. C-12 and represents the design correction aspect of the problem. 
The gusset design, though not evident from the chart, required 
precise sculpturing to obtain the optimum stress distribution 
necessary to reduce the flange stresses and to achieve minimum 
bolt loads. The significant reduction in stress level achieved 
is apparent from the chart. 

At the same time and during the period of several months of 
testing of the design improvement change, process improvements 
were also being vigorously pursued. The resulting changes are 
shown on Fig. C-13? and involved the development of special 
multi-fluted carbide tapered reamers, and revised manufacturing 
procedures including such things as use of improved coolants 
for the reaming process. An extensive test program on effects 
on fatigue life of bolt interference fit was also conducted, and 
as a result the installation tolerances were considerably 
tightened. 

These detail changes in design and manufacture were 
subsequently proven by subjecting a full-scale test specimen to 
the equivalent of 2U,000 flight hours or 6 service lives with- 
out failure . 

As noted previously, the fatigue test program was revised 
to test major components, and Fig. C-lU shows the wing in the 
fatigue test fixture. The F-lll contractual requirements are 

I-Ul 



that the airplane be designed and tested to four lives. At 
just over three lives in the initial test of the wing a failure 
occurred at the point shown in Fig. C-15. Figure C-l6 is a 
photograph of a production part with the failure line indicated. 
The failure initiated in the lower plate of the fitting at one 
of the fuel flow holes machined in the integral ribs . 

Figure C-17 shows the fractured part at the point of failure 
origin. The test had been run in repeated spectrum loaded blocks 
so the block markers were quite clear and permitted the determination 
of the crack growth rate. The critical crack depth for this 
part is about jy" which results in a surface crack length of 
approximately £■". We were well aware that these fuel flow holes 
did represent a stress concentration and this area was used as a 
fatigue analysis control point. Moreover, a number of element 
tests of the part had previously been run satisfactorily to 
specification requirements. It was determined on examination of 
these original test specimens and the full-scale wing test that 
the surface finish of the test specimens was better than the 
fatigue test full-scale wing. 

When this part was initially designed, we were not as adept 
at using finite element analysis as we have since become and 
subsequently, a fine grid finite element analysis was made of 
this area as depicted in Fig. C-l8. The average stress field of 
103,000 psi, used as a reference, corresponds to the maximum 
stress expected in the fatigue test spectrum. The close 
correlation of the calculated values to measured values, which 
were taken from strain measurement of a full-scale part is 
evident. The origination of the fatigue crack was at the left 
lower corner of the fuel flow hole where the stress levels 
indicate approximately 190 ksi. 

Again, the problem existed of having a large number of parts 
manufactured and installed on airplanes which were flying. While 
this particular failure represented no immediate service problem, 
in view of the relatively long demonstrated life, improvement 
was required. Figure C-19 shows the application of an epoxy 
boron reinforcement on this steel plate. It was fortunate that 
the outside surface of the fitting was relatively flat without 
any ribs or stiffeners thus providing a good surface for bonding 
of the composite. Through development test and analyses, a basic 
design and orientation of boron laminates was selected which was 
most compatible with the modulus of elasticity of the steel, and 
provided acceptable bond shear loads. The objective of this 
reinforcement was fatigue enhancement, not static load since 
static testing had been successfully completed, so the design 
was tailored to specifically accomplish the objective of reducing 
the stress levels at the fatigue critical point. Figure C-20 is 
a photograph of this somewhat unique application of filamentary 
composites as a reinforcement. It is quite easily applied in 
production in an autoclave, and can be installed on complete 



I-U2 



wings with use of portable tools employing electric heaters and 
pressure bags. The reinforcement weighs 6 pounds, and by- 
analysis essentially doubles the fatigue life of the part. This 
design has been tested on a complete full scale wing through 
five lives without evidence of any fatigue distress. 

Figure C-21 illustrates the effectiveness of the reinforcement. 
The values are measured stress levels from a full scale test wing 
and show the significant reductions and uniformity of stress 
level achieved. The finish of the rib intersections with the 
plate were also improved. 

I would like to touch on one other problem which resulted 
in a flight failure. Figure C-22 indicates a manufacturing 
defect in a wing pivot fitting which escaped detection during 
manufacture. The small light zone at the base of the black area 
is the only region showing evidence of fatigue. Failure occurred 
during a kg maneuver after a relatively short time in service. 
While we had been concerned in the past with the problem of 
fracture of these high strength materials , as a result of the 
aforementioned test experience, a rigorous fracture mechanics 
test and analytical approach had not been applied. As a result 
of this flight failure an extensive fracture mechanics program 
was initiated and has progressed a long way toward understanding 
and quantifying crack propagation and brittle fracture phenomenon. 
Some of these actions taken on the F-lll will undoubtably benefit 
and be reflected in future programs . 

Figure C-23 shows the elements which essentially constitute 
the service life integrity program for the F-lll as modified to 
include fracture mechanics considerations. Involved in quantifying 
the Fracture Mechanics inspection interval calculations is the 
prediction of crack growth rates and critical crack sizes as 
indicated. This required a considerable test and analysis program 
as an adequate test data base and analytical procedures to 
accomplish these analyses did not exist. The basic elements 
of this program are shown in Figure C-2i+. Essentially, the 
procedure starts with the service usage spectrum used for fatigue 
analysis with the additional consideration of the temperature 
and chemical environment. As it is not practical to perform 
an analysis for every part of the airplane, it was necessary to 
isolate the critical parts which, if they fail, would cause 
catastrophic failures. In some instances the criticality of 
parts could be determined by simple examination. However, in a 
large number of cases extensive failure analyses were required. 
Flaw growth models were also required that were representative 
of the structure and correlated with spectrum test. It became 
apparent that there was a considerable difference between 
spectrum loading and constant amplitude cyclic loadings on 
crack propagation in steel and this type testing was included in 
the specimen test program. Another unique procedure adopted for 
the F-lll was to proof test airplanes on a production basis and 



I-U3 



this was made an integral part of the fracture control program. 
Also, a statistical risk assessment analysis was performed which 
was useful in establishing the relative criticality of important 
structural part. 

Figure C-25 illustrates the basic proof test concept. The 
premise is that the test will demonstrate that a crack greater 
than a critical depth will not exist at a given stress level 
and temperature. Further, in actual operation, the airplane 
will in all probability experience some lesser stress level and 
at a higher temperature than the proof test values . Advantage 
can be taken of both the lower stress and the fact that steel 
parts have a considerably reduced fracture toughness level at a 
low temperature to provide a margin as illustrated which can be 
converted to flight hours for a crack, if present, to propagate 
to critical size. In essence, the procedure is simple however 
there are a considerable number of problems involved in the 
practical execution. Figure C-26 is a superimposed photograph 
of an F-lll airplane subjected to positive and negative loads 
in the cold test chamber. As previously mentioned, the assumption 
is that any cracks which may exist in the structure are below 
their critical crack size for the conditions tested or catastrophic 
failure would result. 

The basic fracture mechanics data program conducted for the 
F-lll involved not only Convair but also a number of other test 
laboratories participated, including those at Boeing, Battelle, 
Aerospace Corp., NASA, AFML and AFFDL. A large number of tests 
were required to isolate the variables shown in Fig. C-27. Close 
to a thousand specimens of D6AC 220-2U0 heat treat steel were 
run in this basic data program. Figures C-28 to C-30 illustrate 
some of the types of specimens and typical test machines which 
were used to obtain this data. 

An unforeseen difficulty was encountered in establishing the 
variation of K]_ c with temperature which is required to permit 
prediction of critical crack size. It was found that D6AC did 
not have a single Ki c variation with temperature as illustrated 
by Fig. C-31. This departure from the expected situation was 
determined to be a function of the cooling rate during the heat 
treat quench from a material temperature of 600°F down. Values 
of K]_ c at room temperature were found to vary from Uo to 95 
depending on this quench rate. It was therefore necessary to 
determine the heat treat histories of all the parts installed 
in the completed aircraft, and to take into account the fracture 
toughness value that would produce the shortest or most conservative 
inspection interval. It should be noted that the other material 
properties: impact strength, elongation, yield strength and 
ultimate strength, were within specifications values and did 
exhibit this variation. Also, in spite of the variability of 
fracture toughness no difference is discernible in crack 
propagation rates . 



1-1* 



The data plotted on Fig. C-32 is typical of the data derived 
from the test program and illustrates the relative influence of 
environment on crack growth for constant amplitude WOL type 
specimens. The Ak range for the data presented is of the order 
of 10 to 60. Similar data from the complete test program have 
been assembled and will be published shortly by the Air Force 
Material Lab. 

As previously noted, it became apparent during the course of 
the program and from evaluations of data available that the 
constant amplitude test specimen program data was not consistent 
with results from spectrum tests. A considerable retardation 
effect was evident from these spectrum loaded tests depending 
upon the order of the applied loading and environment. To resolve 
this difference an extensive spectrum/environment effects program 
was initiated as summarized in Fig. C-33- A baseline program 
was run with certain conditions held constant as noted and variations 
evaluated against this baseline. There was a total of 109 
specimens in this program and again a number of laboratories 
participated in addition to General Dynamics. Figures C-3^- and 
C-35 show some of the typical test set-ups used. 

To establish an analytical correlation, a mathematical model 
was developed by Dr. Wheeler of Convair, as shown on Fig. C-36. 
This model accounts for the crack retardation which takes place 
when a large initial load is followed by subsequent smaller loads. 
The derivation of this equation has been published and is 
available in the engineering literature. The significance of 
the retardation effect is shown in Fig. C-37- The particular 
test illustrated was run to a 5g maximum load and a mission 
developed spectrum in a JP-^ fuel environment. The spectrum 
loads were applied in 58 random load levels. An analysis based 
on constant amplitude test data would predict that the crack 
would propagate to a .20" depth after U00 hours. The specimen 
as tested took 3j200 hours to progress to this depth. A prediction 
based on using the retardation approach previously discussed is 
shown for comparison. 

I would now like to discuss one of the specific aspects of 
the design problem today. In the past, fatigue analyses were 
performed and tests conducted on the assumption that the service 
life was a function of the time required to incubate and initiate 
a crack and to have it progress to failure. The current concept 
interjecting Fracture Mechanics starts with the assumption that 
there is a pre-existing crack. Figure C-38 is an artist's 
rendition but it illustrates the obvious fact that the left-hand 
curves will produce an appreciably shorter life than the right-hand 
curve. It is also generally recognized, as noted on the chart, 
that there are a large number of factors involved in the development 
and growth of flaws. The net result is that the design task, if 
it is to take into account Fracture Mechanics considerations 
directed toward specific requirements, will be considerably more 
complex. 

I-U5 



In conclusion, I have summarized on Figures C-39 a nd C-Uo 
the Designers problem as it can be viewed today in regard to 
achieving adequate and safe service life. To cope with these 
problems, new methods and procedures need to be developed and 
integrated into the design process. 



I-U6 



• STATIC STRENGTH 

• FLUTTER AEROELAST ICITY 

• SERVICE LIFE 

• APPLIED LOADS (Magnitude Spectrum) 

• INTERNAL LOAD DISTRIBUTION 

• DETAIL DESIGN (Concept, Analytical Methods. Materials) 

• MANUFACTURING (Fabrication & Inspection) 

• SERVICE USAGE (Operational. Maintenance! 



Fig. C-l 



M/UTAPV AIR&PAFr /A/-F£/GSYr 

rA/LUP£ or f>/?/mat?v <srp{/er(/f£ 





WING FAILURES 




F-4 


18) 


B-57 


(2) 


F-5 


(1) 


C-5 


(1) 


F-8 


(61 


C-124 


(1) 


F-84 


(2) 


C-133 


(11 


F-8b 


(1) 


T-33 


(1) 


F-100 


(3) 


T-37 


(1) 


F102 


(3) 


A-I 


(?) 


F-104 


Hi 


A-4 


(2) 


B-26 


(11 


A-6 


ID 


B-52 


(3) 


F 111 


(11 




| TOTAL 


I42)| 





FUSELAGE /TAIL 


F-101 ID 


F-105 (3) 


B-52 (6) 


C-47 ID 


C-130 (I) 


C-133 <ll 


KC-135 (1) 


T-33 (1) 




1-37 ill 




TOTAL - (16) 



• ABOVE FAILURES OCCURRED DURING 196? THROUGH SEPTEMBER 1970 TIME PERIOD 
EXCEPT FOR B-52 AND B-57 FAILURES. 

• B-52 AND B-57 FAILURES OCCURRED DURING 1959 THROUGH 1969 TIME PERIOD 

• FAILURES INVOLVE HOIH USAF AND USN AIRCRAFT AND IN SOME CASFS 
RESULTED IN LOSS OF AIRPLANE. 



IS-/-FB ■ 7/ 



Fig. C-2 



I-U7 



ogmf /ww-d&w&ir /zy&yxsf ewz&mee 



1962 1969 INCLUSIVE' 



• i SAI FIGHTERS DESTROYED ALL NON COMBAT CAUSES 

• Avcrag si per Year 

• Average Rate per 1 000 Flighl Hours 



> - 



• USAr FIGHTERS DESTROYED IAIIHRL OF PRIMARY STRUCTURE 

• Average Losses per Year., . 

• Average Rate per 100 000 Fligl I Ho n 

• Portion ol Total Fighter Losse; Ca i I by Failure 
of Primary Structure . . 



. ! 25 



• FAILURE OF PRIMARY STRUCTURE IS CAUSE OF ABOUT 1 % OF ALL USAF NON -COf/.BAT 
FIGHTER LOSSES 



NORTON AFB DOC. AS-XI6 INCLUDES F-4, F-5, F-86, F-100, 
F-IOI, F-102, F-104, F-105, F-106, 4 F-IIIA I D i 



Fig. C-3 



1962 19h f . INCLUSIVE' 



AIRCBAFT 


LOSSES 
ALL CAUSES 


LOSSES- FAILURE 
OF PRIMARY 
STRUCTURE 


TOT PLIGHT LOSS RATE LOSS RATE FAILURE OF 
HO Urc 1 ALL CAUSES PRIMARY STRUCTURE 
i(P?r W.000 Fit Hrs) (Per 100,000 Fit Hrs ) 


F-4 
F-5 
F-8h 
F-100 
1 101 
F-102 
F-104 
F-105 
F-106 
F-111AE D F 


167 
9 

44 
328 

9Q 
132 

82 
193 

47 

10 





1 

3 
1 
1 
! 

2 

1 


2 165 094 
45.787 

231.713 
2 543.027 
1.122 795 
1 572,305 

332,016 
1 088 619 

538 278 
85,073"-- 


1 71 

18 99 
12 90 
8 82 
8 40 
24 70 
17 73 
8 73 




432 
118 
089 
064 
.301 
184 



Total 


1 111 


10 


9.689,622 


11 47 


103 



N r P Doc. AS-XI6 

'Not Con parable Due fo Piograr: 
•As of 2 July 1971 



Fig. C-k 



I-U8 



Fs//<5mAz?z/&id. //&7&*r 



• MOSl SIGNIFICANT PROBLEMS HAVE BEEN ASSOCIATED WITH HIGH HEAT 
TREAT STEH 

• DETAII DESIGN 

• FABRICATION INSPECTION 

• SOURCE OF PROBLEM IDENTIFICATION 



• FATIGUE TEST (4) 

• SERVICE USAG1 111 

• INSPECTION'PROOF II SI 13 S C.C 



facture 



Fig. C-5 



OeTMMt &7£& &?#r& 



STABILIZER HORN-RI& 



rupper torque 
tube 




496 KJACELuE FORMER 
/VIN& CARRY TROUGH BOA (WCTB.) 



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I-U9 



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/f-4 #m4GF-a/ffier 7z%z/g&c 




Fig. C-8 



1-50 



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o ■ C 






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UPPSf 

fvm- 

Beum ! i. . •'. 

em 

SLUG 



fw n o 

A-5 (Sutictest) a 

AIRPLANE '2» Q 



P .. a* '» i " s/f . « 



180 ! 
IW ; 

140 

100 
80 



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UPPER FLANGE STRESSES 






!2o * | BELCW BOLT STRESSES _] 

100 * *v^ 



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to- ¥■ 



* 5 



100, 



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**« .•«« .».« 



1-51 



&r/&/£ PES/4A/ 'G4C7DX& 




FATIGUE LIFE 



f*t* .'V K CI* 



Fig. C-ll 



Bi&!r0FM0/?fAfM&t3?£r -£60, 



(R\ 



t 3 'STAT/C FESTAKr •,- , ;;r- 

A-3 &W/0 TEST ART - M0C [ '■' \ - 5U56ET 




Fig. C-12 



1-52 



• MREE-STAGE REAMING SEQUENCE 

• IMPROVED COOLANT FOR FINISH REAMING 

• POWER FEED IAPER REAMING 

• 18-1 IBIDE FINISHING REAMER : ' 15 RMS 

• I REASEO PROTRUSION OF TAPER LOK BO: 



Fig. C-13 



A-4 mm ue 7Esnm» 




Fig. C-li+ 



1-53 



//// WAG /te&FAti&y- &M5R &//&MOE 



. 




HUM 6H& 



Fig. C-15 




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Fig. C-l6 



1-5^ 



Oman FtowwoAffav j-4 Fm new Me 



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!N£HE£ 



OS 



s^M^wr^p^mm 




t 4 b ft ID IZ 14 lb l& Z0 Z.2 24 Zb Z8 3D J2. 

TEAT &LC76K 



Fig. C-17 



FUEL FLOW HOLE NO. 4 



AV6 FIELD 

sroEse 

THIS APEA 
IS 103 



ALL STRESSESMARE 
FORMAX.FAFI&UE LOAD 
*RSURES FROM A -4 
STRAIN SURVEV 




Fie. C-18 



1-55 



F/NAL ^&£M&77W&&2&W&M0l&e 




Fig. C-19 




OOUBLER INSTALLED 
ON FV» 22 WPF 



Fig. C-20 



1-56 



W/Na PlMT F/TTIN6 - (&NTER SPWAREfl) 

STRESS COMPARISON BEFORE * AFTER BORON REINFORCEMENT 







{740)? /I4- 
927 




BAS/£ 



/09 -/2± ?* 2 /S/ <fj'9_ 70t-__ 62 & 

\jl± \jl0_ \rff? \jJ6_ \ 96 8 \S6 3 



W/rH 80R0A' 

mil STRES5ES SHOWN ARI MAXIMUM 5TCESSE I FATIGUE SPECTRUM 

{exmtwi7&>flmA&-:i* spxcrtM/r7D&m>x. 757. ort/w) 



W-S74-- K90f-2A 



Fig. C-21 



i£frm&'0wrf/mv& jazh/aw 




steel 28z nominal 
Bonpep Aluminum honeycoms 



Anomaly in steel fracture 






Fig. C-22 



1-57 



66*V/e€ L/F6 /A7£G/?/ry 



SER ICE OPERATIONS 



MISSIOI 

| , < | 



LOADS SPECTRA 



FLIGHT LOADS 
GROUND LOADS 
ENVIRONMENT 



■- 1 I" I 



if 



FATIGUE 
TEST 

PROGRAMS 



i riGUE anal ,';is 



I I'FKAME 
LANDING GEARS 



INSPECTION INTEP A 
CALCULATIONS 



CRACK GROWTH 
CRITICAL CRACK 
SIZE 



• CRACK GROWT-H 
DATA 

• FRACT. TOUGHNESS 
DATA 



OPEI DS "EC PROG 



STRUCT. Ft T TEST 



static ri • 



SERVICE-LIFE 
ANALYSIS 



FLEET 
INSPECTION 
PROC-RAM 



PARAMETRIC 
FATIGUE 
ANALYSIS 



I 



USING 
COMMAND 
PLANNING 






Fig. C-23 



BASIC ELEMENTS OF F-lll FRACTURE CONTROL PROGRAV. 



PROOF TEST 

AND NDE 

'Initial Flaw Size' 



SERVICE USAGE 

SPECTRUM 
iLoads-Ternp. - 
Environnenti 



CRITICAL PART 

SURVEY 
(Critical Arias and 

Type Flaws' 



BASIC FRACTURE 
MECHANICS 



<Ki„. K 



'Nc- Msec 



da Iti 



N~^ 



Mil 

INSPECTION 

INTERVAL 

CALCULATIONS 



RISK 
ASSESSMENT 



FLAW GROWTH 
' IDEL 

DEVELOPMENT 



X 



CORRELATION OF 
FLAW GROWTH 
MODEL WITH 

SPECTRUM TESTS 



FLEET 
INSPECTION 
PROGRAM 



Fig. C-2U 



1-58 



PROOF TEST CONCEPT 



CRACK DEPTH, 
a, IN. 



• RESIDUAL STRENGTH 




• CRACK GROWTH 




„de/r cr&e <zw 






1 








1 


_4 — U 1 




""■"" \-rm-) 



STRESS, <r, KSI 



FLIGHT TIME, HRS. 



rtf-f68->G/?63 



Fig. C-25 



-<&V r 0QOCF7£&7'( : 2.4* Z356s) 




Fie. C-26 



1-59 



BASIC I R A i. T tj RE MECHANICS DATA 
D6ac 220/240 KS I 



* VARIABLES INVESTIGATED 


TYPE DATA 


K|c | K| SC( Scda/dl | da/dN 


M|| 


SPECIMENS: ICT, ST DOB > 
MATERIAL: (Plate Forging, Prod. Parts) 
HEAT TREAT PROCESS (Ti ' itures, 
Quench Rate, Quenchant, Circulation) 

THICKNESS . 10 25 to 75 in ) 
DIRECTIONALITY: ( RT, RW, WR TR. WT. TW i 
TEMPERATURE: 1 -65° F to 300° F ) 


v 

V 
V 

V 

V 


v 
V 
V 

V 
« v 


V 

V 

V 
V 


V 

\ 


ENVIRONMENTS: (Dry Air, Lab Air, jp-4, 
Distilled Water. 3-1/2* NaCI, Prussian 
Blue, Relative Humidity) 




V 


V 




FREQUENCY 11,6, 50, 130, 600 cpm) 
LOAD RATIO: (0.1. 0.3, 0.51 






V 




DEPTH TO LENGTH: ( a/2c . 09 to 56 ) 

DEPTH TO THICKNESS: <a/t=.29to 99) 








V 
/ 


*NO. SPECIMENS IGD/TOTAL): (722/930) 


462/556 


98/124 j 129/217 


33/33 



Fig. C-27 



rWCXC /WC7W6 MFCWW/6S S&ECMEMS 




'S-iCSV 



Fig. c-28 



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Fig. C-29 



rwc4t ok/aw 7iEsrser-c/p 




• COM! " 

• I TESTS 

• N1AUY CONTROUED 

• ir'll T'SilNG 



Fig. C-30 



1-61 



Dbac,Z20/240K3l 







40 00 120 

TEMPERATURE T 



ux> 



if J* 7/ 



Fig. C-31 



g&p } wz*-tfew&zwr# wjm%& fM/zm^® 



ii 



2x10 




i I.VitNS 



• NSTAN1 I' ■ ■ 

•■ = 0:3, 

i 



• THIN -Data • in. 

sp imens 

• THIC" - Data on 7 
I 

. CLATIVEHOSTI ITY OF ENVIRONMENTS 






Fig. C-32 



1-62 



SPECTRUM/ENVIRONMENTAL EFFECTS PROGRAM 
D6ac, 220/2*0 KSI 



APPROACH 



ESTABLISH BASELINE DATA 

• CONSTANTS: K|C Flaw, Spectrum, Material Form, Thickness, Unit Stress 

• 4 ENVIRONMENTS: Dry Air Rel. Hum. JIM Fuel, Dist. Water 

• 2 or More Crack Growth Curves/Environment 



• EVALUATE EFFECTS OF: 

• SUSTAINED LOADS 

• COMPRESSION LOADS 

• FRACTURE TOUGHNESS 

• MATERIAL FORM 

• SPECIMEN THICKNESS 



• STRESS LEVEL (Unit Stress) 

• SPECTRUM SHAPE 

• LOAD SEQUENCE 

• SURFACE FLAW SHAPE 

• FLAW TYPE 



NUMBER Of TESTS 



• GD SPECIMENS/FLAWS 

• TOTAL SPECIMENS 



48/77 
109 






Fig. C-33 



s&Eer/tm/ resrawr/zot /mstzw/gvtht/qy 




■ 



Fig. 0-3^ 



1-63 



S&ECm/M '/Xr/5V£ 7EST &ET-UP 



i\ — 




I 



■&;:■' 



Fig. C-35 



*##• z$y &6VS 



ANAl ' riCAl APPROACH 



• a + Ry ?a p 
Cp-1 



• a ♦ R v - a D 

/ Ry ^ 

C P = U p ^T 



da 
dN 



■f(AK) 




a ■ a Q + /c p f(AK)dN 



Fig. C-36 



I-6U 



3U&*C£/ : ij<HV£0\SP£C/M£N* t C&ACK G#OMH *& fl?££>/CreO 



TEST PLAN: 



• 5.0 G MAC SPECTRUM 

• 58 LOAD LEVELS 

• RANDOM SEQUENCE OF 
LOAD LEVELS 




5 6 7 8 9 10 11 

NO. OF 200 HOUR TEST BLOCKS 



16 17 18 

Fit 27* m 0739 
2CXCI970 



Fig. C-37 



/TMW&mawfivrgMfiE 




I INCUBATION 

II MICRO MACRO CRACK DEVELOPMENT 
III CRACK PROPAGATION 



FACTORS 



A - INTRINSIC MATERIAL PROPERTIES 

B DESIGN STRESS LEVELS LOAD SPECTRUM 

C • DETAIL DESIGN 



A B-C-D E 

ABC 

A-B-C-E 



D - MANUFACTURING PROCESSFS 
E ENVIRONMENT 

an tr* >.*w* 



Fig. C-38 



I-65 



• ACHIEVEMENT OF ADEOUATE SERVICE LIFE AT ACCEPTABLY LOW RISK IS DEPENDENT 
ON A LARGE NUMBER OF FACTORS, (some of which are beyond the D>rert Control ot 
the Designer) MANY OF THE FACTORS WHICH MUST BE CONSIDERED ARE IN TURN 
SUBJECT TO A LARGE NUMBER OF VARIABLES. STATISTICAL ASSESSMENTS OF 
SERVICE LIFE AND RISK CAN BE USED TO ASSIST IN THE DECISION MAKING 
PROCESS; HOWEVER. WHAT CONSTITUTES ACCEPTABLY LOW RISK HAS NOT BEEN 
ESTABLISHED. 

• TO ACHIEVE DESIRED PERFORMANCE, INCREASING USE IS BEING MADE OF MORE 
EXOTIC MATERIALS AND MANUFACTURING PROCESSES 

• HIGH STRENGTH METALLIC ALLOYS 

• FILAMENTARY COMPOSITES 

•HIGH FATIGUE RESISTANCE FASTENER SYSTEMS 
•WELDING, ADHESIVE AND DIFFUSION BONDING 

• SURFACE FINISHING 

IN ORDER TO ASSURE THAT THE EXPECTED BENEFITS IN SERVICE LIFE ARE ACHIEVED 
THE DESIGNER MUST HAVE A MORE INTIMATE ASSOCIATION WITH THE MANUFACTURING 
AND INSPECTION OPERATIONS AND PROVIDE MORE SPECIFIC INFORMATION AND 
DIRECTIONS 

Fig. C-39 



• THE CURRENT TREND TOWARD A DESIGN CRITERIA BASED ON THE ASSUMPTION OF 
PRE-EXISTING FLAWS AND A DEFINITION THAT FATIGUE LIFE IS BASED ON THE 
TIME A FLAW BECOMES OF DETECTABLE SIZE REQUIRES NEW KNOWLEDGE ON THE 
PART OF THE DESIGNER OF THE VARIABLES AFFECTING SERVICE LIFE INCLUDING 
WHAT THE PRODUCTION MANUFACTURING PROCESSES WILL PRODUCE AND IN- 
SPECTION CAPABILITIES WILL FIND 

• TO AVOID THE PROBLEM OF CATASTROPHIC FAILURE IN MONOLITHIC STRUCTURE 
THERE HAS BEEN A TREND TOWARD CONSIDERATION OF FAIL SAFE DESIGNS 
THIS DESIGN CONCEPT HOWEVER PRESENTS A NUMBER OF PROBLEMS 

• PERFORMANCE DEGRADATION (Weiyht) 

• COST - INITIAL AND MAINTENANCE 
(Additional Piece Parts and Joints* 

• VERIFICATION BY TEST OF FAILURE MODES 

A FAIL SAFE STRUCTURE IS NOT FAIL SAFE UNLESS THE FIRST FAILURE CAN BE 
READILY DETECTED AND CORRECTED 



Fig. C-J+O 



1-66 



SOME COMMENTS ON FAIL-SAFE DESIGN 



Walter S. Hyler 
Batte lie -Columbus 



Abstract: Tensile design properties are based on strength 

distribution curves with mean stress and standard 
deviation. No similar curves are available for 
fracture resistance. Data are presented to show the 
large scatter of K lc , depending on slight processing 
variations as well as on location within the same 
sheet. A coefficient of variation indicates whether 
processing is well under control. Improved material 
characterization will be necessary for fail-safe 
design with thick, moderately thick, and thin 
structures --again requiring close interrelation 
between design and materials and processing. 



The talk by Mr. Dietz this morning was very impressive to 
me, probably not only because it was done so well, but also 
because it touched upon many concerns that I have relative to 
fail-safe or damage tolerant design. One thing which came out 
clearly was his statement about the need for positive interaction 
between the designer and the materials and process engineers, 
since in damage -tolerant design of high-performance systems, we 
are concerned with fracture and fatigue-crack propagation for 
which analytical methods still are being developed and for which 
optimum materials processing is necessary. This is where I 
wanted to start my discussion and what I wanted to amplify on 
because I believe that it really is a fairly critical point. 

In the old days — and I say this with a smile because I really 
don't know about structural design in the old days --it seemed 
we were interested in ultimate strength, in limit load, and in 
developing positive margins, and then we would go on to the 
next design problem. With our high-performance airplanes, we 
have more complexities. We may still be computing ultimate 
load, limit load, and positive margins, but with damage tolerance 
we are also concerned with establishment of minimum detectable 
flaw sizes, crack propagation by fatigue, establishment of 
inspection intervals and useful life , and the strength of flawed 
structure . 

For handling ultimate and yield load design calculations, 
we have learned to employ and have become comfortable with the 
use of design allowables based on statistically established minimum 
design properties. It is obvious that we have not nearly approached 
that point yet in terms of fatigue-crack propagation and fracture; 
but as we evolve design concepts, such as fail-safe or damage 
tolerance, then I think that we will have to give some attention 

1-67 



to the establishment of minimum design allowables for these 
behaviors . 

Both fatigue -crack growth and fracture are sensitive to 
geometric factors in component geometry, to composition and 
structure of the material, and to processing details in the 
fabrication operations. The measures of these behaviors, such as 
K Q , Kj c , and da/dN versus A K have grown out of the development 
of fracture mechanics as a stress -analysis tool and can be used 
in design analysis. In the discussion today, I make reference 
to only one of these measures, Kj ; however, the philosophical 
implications also apply to problems associated with plane-stress 
fracture and fatigue-crack propagation. 

While Kj c has historically been considered a constant for a 
material under a plane-strain stress state, this is an over- 
simplification when one considers the reality of materials in 
structural components. If one simply considers a material in a 
product form, such as plate, processed by procedures reasonably 
well under control, we know from countless tensile tests that 
there is a variation in tensile properties of 15 to Uo ksi, 
depending upon the material and its strength level. Similarly, 
one can expect the same material to exhibit a variation in Kj c 
of some appreciable, but generally unknown, magnitude. Now 
consider the same thick plate material fabricated into a useful 
but complex piece of structure with distinct changes in thickness 
and plan-form and finally heat treated to the desired strength 
level. If the material is quench-rate sensitive or sensitive 
to some other processing procedure, one can expect that the 
variability in Kj c from location to location may be appreciably 
greater than in the prior case. A design philosophy such as 
damage tolerance that is based in part on fracture toughness 
considerations is not on too solid a footing unless there is a 
realistic appraisal of this variation. 

Figure D-l provides some rationale for the above statement. 
In this figure are shown for three generic materials (with their 
representative F^ u levels), the critical surface crack length 
associated with various Kj c levels. If one decides, based on 
limited testing of simplified shapes that "reasonable" Kj c 
levels should be 60, 60, and 25 ksi inch (for steel, titanium, 
and aluminum, respectively), and if the real material in a complex 
component shows variations of -30 percent or more (which is not 
unlikely), the possibility exists that a flaw will grow to 
instability long before it can be detected with high probability . 
This is a consequence that may accompany inadequate determination 
of minimum Kj c values. For example, for the steel alloy in 
Figure D-l, the carck length is shown to be 0.l6 inch. However, 
if the distribution of Kj c in real parts can vary by -30 percent, 
application of the limit load stress might result in failure 
with a surface crack of only 0.07 inch; hardly a detectable 
crack with high probability. 



1-68 



For static properties, as stated before, the aerospace 
industry has "become accustomed to the use of statistically based 
design allowables, such as the A and B values in MIL-HDBK-5. 
I would like to review briefly some features of A and B values, 
since some of the ideas are instructive in considering fracture 
and fatigue-crack propagation. Figure D-2 shows the analysis 
of tensile yield strength for 7075-T6 sheet. 

In this figure are shown two bell-shaped distribution curves 
based on an analysis of test data. The solid curve was obtained 
from test samples sectioned from one sheet. The dashed curve 
came from a large number of quality control tests on many lots 
of the alloy. It is seen that in either case the properties 
fall in a broad range, and for this collection of data the 
single sheet data covered over half the range obtained from the 
mult Hot test data. It is from this type of analysis that the 
aluminum industry developed the concept of A values as a minimum 
design mechanical property. Thus, the A value is a minimum 
strength above which it is expected that 99 percent of all 
future production will lie with a confidence of 95 percent. I 
have shown on the figure what I terms an A' value that represents 
a maximum strength value, below which 99 percent of all future 
production will lie with a confidence of 95 percent. 

For this alloy, the "process capability" is such that 
98 percent of all production is expected to lie between A and A' . 
The term "process capability" implies that for a given production 
process schedule, there will be a recognizable distribution in 
strength properties, definable by a reasonable population of 
test data. The range of properties reflects acceptable limits 
on composition and structure and on all the controls utilized 
in each processing step. 

When we consider fracture as exemplified by Kj , it may be 
that for small quantities of a given material, heat treated 
carefully and consistently, that Kj c will be a reasonably constant 
value. If one can convince one's self that such values represent 
the material in a complex structure, it (Kj c ) provides a good 
engineering tool to employ in damage tolerant design. However, 
many modern aerospace structural materials are sensitive to 
various states of processing so that unless one knows the "process 
capability" of a material with regard to fracture, one may be 
led to unsafe design conclusions. The next several figures 
and accompanying discussion amplify this point. 

Figure D-3 shows the distribution curve obtained from 
quality control data on current production Ti-6A1-^V mill- 
annealed plate. The statistical analysis resulted in the 
A, A', X, B and B' values indicated. One concludes from the 
analysis that the process capability of this alloy is such that 
98 percent of expected future values will lie between 118 ksi 
and lU6 ksi. 



I-69 



Figure D-k shows the relationship between F, and Kj c for 
the same titanium alloy in about the same thickness range. In 
this figure, each horizontal line, vertically tic marked at 
the ends, marks the minimum and maximum Kj c value observed for 
a given plate with the appropriate TYS. Two or more tests 
were conducted for each plate. The excessively large scatter 
is disturbing since the material is essentially that which 
can be purchased according to the confines of present military 
specifications and, as such, corresponds reasonably well in 
tensile properties to the material shown in Figure D-3. 
Consequently, the ordinate scale on Figure D-1+ is tic marked 
to indicate A', B 1 X, B, and A values from Figure D-3. 

From data, as shown in Figure D-U, one must establish a 
representative Kj c value for use in damage -tolerant design 
considerations. There are several possibilities for this 
selection: 

(l) Assume a typical value. A natural selection would 
be a Kj c value associated with t he me an strength, 
X; thus, a K-r of about 65 ksi ■ inch. 



5 



(2) Establish a trend line, such as the downward sloping 
solid line in Figure D-U, and select a minimum Kj 
value associated_with the A' value; thus, a Kj c of 
about U2 ksi inch.* 

(3) By statistical analysis, using regression techniques 
and introducing probability and confidence (not 
necessarily 99? 95) , establish a minimum curve and 
select a minimum Kj c value associated with the A' 

valuer. Such an approach would yield a Kj c of 35 

ksi ' inch, or less . 

(h) The fourth alternative is to lump all of the data 
and compute by statistical techniques the X and s 
associated with the data and a minimum value based 
on some agreed-upon probability and confidence. A 
K-r value about the same as in Alternate 3 would 
be anticipated. 

Neither Alternates 3 nor h provide a design value for Kj 
that would excite the designer (based on this data collection; 
since this results at a limit load stress of F^/l.5 in the 
need to confidently find surface cracks less than about 0.25 
inch according to Figure D-l. Alternates 1 and 2 appear less 



* The A' value is chosen since the TYS distribution indicates 
that this is an expected value for this alloy. 



1-70 



desirable, since they ignore the distinct probability that 
substantially lower values of Kj c will actually be present in 
the material (particularly for Alternate l). There is some 
belief that for fracture critical components the assumed 
Alternate 1 value can be screened with quality-control tests. 
This rationale appears to be potentially erroneous if the 
distribution of Kj c values in a single plate shows a range of 
expected values nearly as broad as that from a large sampling 
of plates -- along the lines suggested for TYS in Figure D-2 . 

It should be stated that other information than shown in 
Figures D-3 and D-k (relative to composition and microstructure 
and their influence on K-r for this alloy) suggests that the 
extensive scatter may not be real for this alloy, providing 
some further examination of process control is done in order 
to sort out what contributes to the variability and how process 
controls can be modified to lessen variability. This is the 
essential point that Mr. Dietz was addressing in promoting 
much closer interaction between design people and materials 
and processing people. 

Another material of considerable interest these days is 
D6AC, a modification of 1+3^-0. Figure D-5 shows the distribution 
curve for yield strength for this alloy based on a large sampling 
of producer and user quality control data. It was a normal 
distribution with X at 2lU ksi. The process capability was 
such that 98 percent of future production would be expected 
to fall within the range 200 ksi to 228 ksi with a confidence 
of 95 percent. 

Figure D-6 begins to delineate the real-life situation 
for this alloy with regard to fracture, or Kj c . This is a very 
busy figure that requires some preliminary orientation. The 
rectangular border represents the boundary of l6 different 
plates of D6AC (2 feet by 3 feet) all processed similarly by 
General Dynamics /Fort Worth and identically heat treated 
according to a procedure that I have identified as heat treatment 
A. Each plate had been laid out to provide a variety of specimens 
for tensile tests, fracture toughness, etc. On Figure D-6, 
each of the sixteen plates is identified by a distinct symbol. 
The symbols on the drawing represent the approximate location 
in each plate from which a fracture specimen was taken (usually 
more than one fracture test was conducted and Figure D-6 contains 
all of the data). The numerical values adjacent to each symbol 
represent the observed Kj c values. 

Within the circle at the lower right-hand corner of the 
figure are data for six different plates. It is seen that Kj c 
for the pl ates a t this confined location ranges from about 59 
to 86 ksi yinch. Within the circle at the lower left-hand corner 
also are data for six plates, where some plates in this location 
duplicate those in the right circle. For this second circle, 



1-71 



the range in Kj c is from about 55 to 91 ksi '/inch. Where plates 
are represented in both circles, a plate with Kj c on the high 
side of the range in the right circle may not be similarly 
ordered in the left circle. In fact, if one were to draw a 
number of circles at various other locations on the plate 
plan-form and examine the details of the data within each circle, 
comparing the data with that in other circles, one would conclude 
that with this particular heat treatment a fairly large variation 
in Kj c can be expected within a plate and among the many plates . 
Detailed perusal of the figure shows this range to be from 
slightly less than 50 ksi ' inch to slightly more than 

90 ksi inch. 

Figure D-7 shows the same fracture -toughness data plotted 
as a function of TYS, with the A', X, and A values from Figure D-5 
tic marked on the ordinate. Of the l6 plates, 11 had accompanying 
tensile test data, so that the horizontal lines represent the 
plate tensile yield strength; the tic marks at the ends are the 
maximum and minimum Kj c values in the plates from two or more 
tests . The lowest horizontal line on the figure is for the 
five other plates, where TYS is an assumed, but reasonable, 
value. It is evident from this display that the heat-treat 
process capability with regard to TYS is quite satisfactory. 
However, one concludes that although the fracture toughness 
information also is consistent, the inordinate scatter in results 
may prejudice the utilization of the material since the capability 
of heat-treatment A would yield a design allowable for Kj c of 
approximately 37 ksi inch (for 99? 95 on probability and 
confidence). 

Figures D-8 and D-9 illustrate a happier story. Figure 
D-8 shows Kj c data from eight plates of D6AC that were identically 
heat treated, but in a different way than the previous 16 plates. 
This heat treatment, I have identified as heat -treatment B. The 
plotting scheme in Figure D-8 is the same as for Figure D-6. 
Examination of the somewhat fewer data in Figure D-8 shows 
(l) an encouraging uniformity in fracture toughness within a 
plate and among the plates, and (2) a substantial increase in 
fracture toughness. The range in Kj c values is from about 

91 to 102 ksi /inch. 

Figure D-9 shows the Kj c values plotted against TYS as in 
Figure 7; however, there were tensile data for only two of the 
plates. The middle line at TYS of about 220 ksi is for the 
remaining six plates. Since so few tensile data were available, 
it is difficult to say any more than that there appears to be 
a greater spread in TYS associated with heat-treatment B than 
for A. More important, however, while the fracture test results 
for D6AC with heat -treatment A suggest a possibly severe sensitivy 
to nonuniform quench rates along the plan-form and from plate 
to plate, heat -treatment B suggests that its process capability 
is compatible with achieving high and uniform fracture toughness . 



1-72 



Not only that, but a designer familiar with the use of statistical 
design allowables for static strength would have a much more 
comfortable regard for an "A" value for fracture toughness of 
about 83 ksi ^ r inch to which manipulation of these data lead 
one. 

This type of materials and processes research, which has 
been conducted at General Dynamics/Fort Worth, interfacing with 
the stress and structures people is the way that confidence will 
be built into the use of the damage tolerant design concept. 

Figure D-10 is illustrated to recap some of the statistical 
measures that were employed in preparing several of the other 
figures . Information shown for the three materials and two 
strength attributes are coefficient of variation, C; average 
strength, X; and standard deviation, s. As one examines the 
right-hand column (s), it is immediately evident that the 
standard deviation for fracture toughness of heat -treatment A 
for D6AC is excessively high, certainly suggesting that processing 
is yielding an intolerable variability in results. On the other 
hand, heat-treatment B exhibits an s value even less than that 
for TYS, when processing has been examined and developed for 
the strength attribute of importance. 

In the column for coefficient of variation, which is a 
measure of dispersion, s, normalized by the average value of 
strength, several features stand out. First, the C value for 
Ti-6A1-Uv alloy for TYS is in excess of 50 percent higher than 
the comparable value for the other two materials . This suggests , 
based on this and other information, that process control for 
this alloy probably can be improved. This process control, in 
turn, may significantly alter the fracture toughness display of 
Figure D-U. Recent work at North American Rockwell relative to 
the B-l is beginning to bear this out. Finally, the almost 
order of magnitude difference between C values for TYS and 
fracture toughness (A heat treatment) for D6AC reinforces the 
idea that process development and control for fracture toughness 
was needed. 

This discussion has been an attempt to quantify, by examples, 
my thoughts and those expressed independently by Mr. Dietz and 
others at this meeting that the development of the damage tolerant 
design concept presents a challenge in the aerospace industry 
that encompasses design engineers as well as materials and 
processing personnel. The challenge is not only in developing 
and understanding the design concepts utilizing fracture mechanics, 
Equally important is to recognize that the mechanical behaviors 
associated with damage tolerance, fracture toughness, and 
fatigue-crack propagation are structure and process sensitive. 
This sensitivity will require perhaps an order of magnitude 
increase in materials and process research, development, and 
characterization on aerospace systems contracts in order to assure 
that the material in future structural components will have high 
and uniform fracture toughness. 

1-73 



FIGURE D-1 

CRITICAL SURFACE-CRACK LENGTH VALUES FOR SEVERAL GENERIC 
MATERIALS WITH ASSUMED K| C VALUES 



Steel, 


Ftu 


= 280 Ksi 


Ti, 


Ftu 


= 130 Ksi 


Al. 


Ftu 


= 75 Ksi 


K|C 




2C, inch 


K IC 




2C, inch 


K IC 




2C, inch 


20 




0.017 


20 




0.078 


15 




0.14 


40 




0.07 


40 




0.31 


20 




0.25 


60 




0.16 


60 




0.70 


25 




0.39 


80 




0.28 


80 




1.24 


30 




0.56 


100 


r>n ce 


0.45 

«mirirn ilar cnH 


100 

arp f lauu a 


nr\ S 


1.70 

= F*../1 R and I 


35 




0.77 
7c. 



90 1- 



c 

2 

</) 

> 

c 



Multi-lot 

n > 4500 
X = 80 ksi 
s = 2.I3 ksi 




J 



70 l - 



Single piece 
n = I78 
X = 78 ksi 
s = 1. 25 ksi 



ALCLAD 7075 -T6 Sheet 

FIGURE D-2. DISTRIBUTION OF TENSILE YIELD 
STRENGTH VALUES FOR ALCLAD 
7075-T6 SHEET 



1-lh 



I60i- 



Ti-6AI-4V plate 
Mill annealed 
t>I.O inch 
n = 2l9 
X = I32 ksi 
s = 5.6 ksi 




100 



FIGURE D-3. DISTRIBUTION OF TENSILE YIELD 
STRENGTH VALUES FOR TI-6AI-4V 
PLATE 



Ti-6AI-4V plate 
Mill annealed 




40 50 60 70 80 90 

K jc , Plane Strain Fracture Toughness 



100 



FIGURE D-4. RELATIONSHIP BETWEEN TENSILE YIELD STRENGTH 
AND K Ic FOR Ti-6AI-4V ANNEALED PLATE 



1-75 



240 



[ 



D6AC plate 
(Nominally 220-240, ksi) 
n=3684 
X=2I4 ksi 
s = 5.73lksi 




FIGURE D-5. DISTRIBUTION OF TENSILE YIELD 
STRENGTH VALUES FOR D6AC 



i 


k 






M ft fir* rt v ^s t t 


r^ 




81 g ApprOX. O TT 






76 2 ® _ .62.3 


60.8 578 




55.5 540 A ®70.5^ 
O 634 VA73.3 










061.8 <gj 057.5 






58 5 59 3 
89.0 81.5 ° O 619 *54.0 


64.2 
V 
60.4 
9 




V 888 V 56 7 70 667 ® V 

§®48.7 A (^76.8 A 528 


N ?93 60~.3 






72.7 


x 88.9 734 


71.6 


O T T 


▼ w 


a. 




< 5 . 8 






59.5 






^TT^x ™° 5 i 9 

/^ 54.6^\ a * 






59 °/^~~-\ 




, / ® , 56.0 ® 7 64 


/\. 7 ^.6 608 


92 
i 


• 3 /894 771 /1-* 66 "L 52 4 

| U 894 •^•66 6 # r) V T7 9 6 

V 714 ° / 835^*^*784 


60.0 / ^ ^ \ 
® 66.0®, ^.6 




\ 91.0 75 3 


86.IU / 


\ 


' 


V 



FRACTURE TOUGHNESS DISTRIBUTION IN SIXTEEN PLATES OF D6AC 

FIGURE D-6 



1-76 



V 


£OU 


— A* 










*N 








_c 










o> 


220 




• 1 — i — 


H i 


— ■ 


c_ 


i 1 1 

, , i 1 . 


k_ 


S i ' 






f 1 






C/) 












"O 




— X 








cu 




J 






>- 


210 


— 






CD 










en 










c 










.<!> 










h- 


200 


— A 

1 I 


1 1 


1 1 



40 



50 60 70 80 90 

Fracture Toughness of D6AC Plate 



100 



FIGURE D-7. RELATIONSHIP BETWEEN TENSILE YIELD 
STRENGTH AND FRACTURE TOUGHNESS 
OF D6AC PLATE, HEAT TREATMENT A 



Approx. 3 tt 



100.5 

O 



96.5 
O 



91.0 
O 



— 95.1 



91.8 



91.2 



x 

o 

a. 
a. 
< 



91.4 




99.2 



98.3 
97 2 



92.5 

• 


92.8 



94.7 

• 


101.9 
8 






942 


94.2 

* 90.6 
8 



O «j • 

96.3 95.4 93 2 


96.9 



FRACTURE TOUGHNESS IN EIGHT PLATES OF D6AC 

FIGURE D-8 



1-77 



A' 



"\ 



> *- 



y 



230 

sl 

a> 
c 
<u 

l_ 

oo 220 
<v 

« 210 

y> 

c 
d) 

h- 

200|— A 

40 50 60 70 80 90 100 

Fracture Toughness of D6AC Plate 

FIGURE D-9. RELATIONSHIP BETWEEN TENSILE YIELD STRENGTH 
AND K Ic FOR D6AC PLATE, HEAT TREATMENT B 



ALCAD 7075-T6 

Tensile Yield Strength - 0.027 

Ti6AI-4V 

Tensile Yield Strength - 0.042 

D6AC Steel (220-240 Ksi) 

Tensile Yield Strength - 0.027 
Fracture Toughness, A— 0.17 
Fracture Toughness, B— 0.042 



80 



132 



2.13 



5.6 



214 


5.73 


64.5 


11.0 


94.8 


4.0 



C = Coefficient of Variation = -=- 

X 

FIGURE D-10. STATISTICAL MEASURES FOR THE THREE ALLOYS 



1-78 



IMPACT OF PROCUREMENT PRACTICES 
ON C-5 STRUCTURAL DESIGN 



W. T. Shuler 
Lockheed-Georgia 



Abstract: Contractual factors which affected the structural 

design of the C-5 are analyzed and evaluated, including 
guaranties on performance, low weight empty, and crack- 
free fatigue life as well as fixed price contract and 
program for concurrent development and production. No 
revolutionary change in procurement practices is recom- 
mended, but areas of possible improvement are indicated 
and discussed. 



Only the abstract is available. 



1-79 



STRUCTURAL DESIGN CONSIDERATIONS FOR 
A SPACE SHUTTLE 



F. F. W. Krohn 
General Dynamics - San Diego 



Abstract: Details of structural arrangement and thermal protective 
system for a space shuttle booster are shown and 
corresponding design conditions are given including 
thermal expansion, creep, shock impingement, acoustic 
environment. Design considerations for high-temperature 
materials, and structural features for components 
incorporating a safe -life or a fail-safe approach are 
discussed -- indicating a new order of complexity 
compared to aircraft design. 



I would like to present to you design considerations for a 
space shuttle booster. The design requirements on space shuttle 
vehicles are a unique combination of the design requirements for 
aircraft and rocket launch vehicles. This is something brand new 
and the combination of these requirements will impose on us quite 
a number of new problems in materials and design. 

For those who might not have followed the space shuttle 
project, there have been four teams engaged in design studies. 
McDonnell-Douglas has a design effort on the booster as well as 
on the orbiter in Phase B. North-American Rockwell and General 
Dynamics Convair have a split in tasks, with us at General 
Dynamics doing the booster and North-American doing the orbiter. 
Then there is a team of Grumman and Boeing and the fourth is one 
company again -- Lockheed. 

The objective of the space shuttle system is to put payload 
cheaply in orbit with re-usable vehicles. The different versions 
which have been studied accomplish this with marked difference 
in cost. Some of the systems throw away tanks which weigh up 
to 60,000#. Payload cost in orbit varies from $75 for fully 
re-usable to $200 plus for partially re-usable systems. 

I will address myself to a system which mounts the orbiter 
parallel to the axis of the booster -- a polar orbit design 
mission with a payload of Uo,000# and an orbiter with two engines 
requiring a staging velocity of 10,800 ft/sec. This results in 
the most difficult booster and this vehicle system was studied 
in depth. However, we had a re-direction as we found out rather 
lately, through side studies, that it would be much more advantageous 
to have three engines on the orbiter. This brings the staging 
velocity down to less than 8000 ft/sec, even for the polar missions. 
As stated I will, however, address myself to the particular system 

with 10,800 ft/sec staging velocity. 

1-80 



The approximate lift-off weight on the pad, with the orb iter 
mounted, is slightly in excess of 5 million pounds for the 
combination -- a rather staggering weight. As far as the booster 
is concerned, our lift-off weight is ^,l88,OOC#. The landing 
weight is 638,000#. The rest is propellants . 

The vehicle is really a huge tanker with 12 rocket engines 
mounted at its base. To give you some idea about the dimensions, 
the booster has a wing span of lkk feet and a total length of 
256 feet. For cruise back it has JP propellants of approximately 
lUU,000#. It has an ascent propellant weight of 3,382,000#. 
Of the total vehicle weight, essentially more than half is 
oxidizer. 

Now let us get a little bit into the description of this 
vehicle system. It takes off vertically. The orbiter is 
connected with a reversible link system which is sized in such a 
fashion that the orbiter can be ejected even with the orbiter 
engines dead. This is accomplished by the thrust on the booster. 
In normal operation the orbiter has its thrust run up to 
approximately 50$ at time of separation. There is still a very 
large force in the horizontal direction on the linkage system. 

After the orbiter has been separated, the vehicle will 
assume an angle of attack of approximately 60° and spend its 
kinetic energy very rapidly. The whole heating cycle is exceedingly 
short and the heat-up is very steep. Here, you will see, we have 
very severe implications for structural design. Everything on 
the outside will be hot and any inside structure will just lag in 
temperature. The vehicle, during the time of re-entry, is 
controlled by a set of reaction controls because the tail is 
fully ineffective and also the wings are not very effective at 
this time. 

We have 12 jet engines in this vehicle. They are deployed 
and run up one after the other following re-entry. The last 
engine will have run up approximately two minutes after the first 
has been deployed. The vehicle then will cruise home or it will 
cruise to the next airfield. The vehicle has the capability just 
like some of the orbiters to ferry back to the launch pad. 

In structures arrangement there is a very large difference 
between the design requirements for 10,800 ft/sec staging velocity 
and a vehicle staging with less than 8000 ft /sec. A 10,800 ft /sec 
vehicle requires a substantial thermal protection system (TPS). 
An 8000 ft/sec vehicle can be built around what we call the heat 
sink principle. Here temperatures are controlled by allocation 
of sufficient mass in the surface. Use of heavy aluminum gauges 
will reduce temperatures to 300°F over most of the vehicle 
surface . 



1-81 



Now let us look at the structural arrangement shown in Fig. F-l. 
The vehicle has a thermally protected backbone structure consisting 
of thrust structure, LH2 tank, intertank, LO2 tank and forward tank 
extension. All major vehicle loads are introduced into this thermally 
protected core. 

The wing is link-attached with vertical and drag links. Once 
we fuel the LH2 tank, the tank wall will go to approximately -100°F 
and shrink very substantially while everything around it will stay 
at room temperature. So, all supports have to be statically 
determinate . 

The orb iter attachments are backed up by two large bulkheads 
which are essentially external in the region of the LH2 tank and 
one very deep forward bulkhead which is inside the LO2 tank. This 
forward support takes the brunt of the loading. 

Figure F-2 shows the heat shield which shrouds the body. It 
is composed of large semi structural elements. We did not follow 
the usual route of having TPS panels of 2 or h feet square . Instead 
we have large panel shells with slip joints and a horseshoe element 
directly attached to the wing because the temperatures of the 
upper wing surface and this shroud are approximately the same. 
One must make sure to get compatible expansion. The best way to 
enforce this is to attach the heat shield to the wing directly as 
a root web. All other shell elements are supported in the center 
and expand fore and aft with respect to the substructure. 

The great advantage of this arrangement is that it reduces 
the potential leak area in seals and slip joints to less than 
1000 ft length compared to some miles in a conventional panel 
system. This makes it possible to control the containment of 
purge gas. There is a purge envelope over the LH2 tank separated 
by a bulkhead from the purge cavity over the L02 tank. Both 
cavities are filled on the launch pad with dry nitrogen. 

In up-f light, we bleed off the purge gas to avoid large 
internal pressures on our heat shield. Just very shortly before 
staging, the bleed-off vents are closed and kept closed during 
the time of maximum heating. We will have some loss of purge gas 
during this time without any doubt. We have a self-imposed spec 
of permissible leakage area (not to exceed 350 square inches) and 
this permits us to go through this environment with a quite high 
degree of assurance that we will not get hot plasma into the 
purge cavities. After we have suffered through the re-entry, the 
tank pressure will be reduced and the vents will be opened to 
take air on board for pressure equalization. So much for the 
general description and operational aspects of the vehicle 
effecting the structure . 



1-82 



The design criteria shown in Fig. F-3 indicate for the launch 
mode an ultimate factor of l.k x limit and for the aircraft mode 
an ultimate factor of 1.5 x limit. Buried in this arbitary 
limit-to-ultimate factor are 80 million dollars . That is the cost 
to conduct a full-scale separate static test. 

Now, what are we really after? We are after a vehicle which 
is capable to take limit loads at the end of its useful structural 
life. We have proposed to the contracting agency to leave these 
factors in for design but to drop the whole static test program 
except for development testing. Instead, we would go through four 
lifetimes of operational flights and then subject the structure to 
1.15 times limit load. We feel that this is a much more useful 
type of testing. Arbitrary design criteria which we have set up 
with the ratio of yield-to-ultimate factor on one outdated 
aluminum alloy in mind most certainly are suspect in our time 
and age. 

The tanks, which compose a very large portion of the structure, 
are not really designed by these external load factors. They are 
designed by fracture mechanics or, to put it another way, by proof - 
testing requirements . The tests are conducted with a proof 
pressure which precludes the presence of a crack which could grow 
to critical size with actual operational pressures and projected 
crack growth during 150 flights. This imposes a very severe burden 
on the materials community because now we really need to know 
crack growth rates and cannot be content anymore with K-j_ c alone. 
All our gages are thin enough to be neither plane strain nor 
plane stress, but something in between. And here we have a very 
substantial lack of data, even on the 2219 material which we have 
selected and which is excellent for the application. 

One of the new criteria coming in is creep. This is a wide 
open field. We have limited our creep to .2% plastic deformation. 
We do not know if this is right because we don't know what really 
happens with a panel that has experienced this much creep. One 
could have panel flutter. One also has steadily changing conditions 
as far as overall dynamic behavior of the structure in flutter is 
concerned. We have designed around many of these problems, just 
to be on the safe side. 

Figure F-U shows peak limit load intensities and the load 
envelope of the vehicle with 25 different loading conditions. 
As you can see, not much is designed by the aircraft mode of 
operation, mostly localized structure only. Figure F-5 shows that 
the tanks are designed for safe life and only a relatively small 
portion of the structure is designed for fail-safe criteria. 

We made a valiant attempt to get fail-safe tanks and looked 
into crack stoppers as they are presently employed on commercial 
jet fuselages. We found that they would have to be at a prohibitively 



1-83 



close spacing of approximately hjt" . We were quite uncomfortable, 
as aircraft designers always are, with all this safe life 
situation. 

Then, by fracture mechanics analysis, we found to our delight 
that the tanks will leak long before we get to critical crack 
size. We really could fly with a subcritical, leaking crack on 
quite a number of missions before it would become critical. We 
have to make very sure that we install sensors in the purge 
cavities to find out if we have somewhere a slow leakage . This 
should warn us and we then could take a closer look at the tank 
wall. 

Figure F-6 shows the structural backbone of the vehicle with 
thrust structure, LH2 tank, intertank, LO2 tank, and forward tank 
extension. Orbiter support points as well as landing gear 
attachments are shown. A heat shield is mounted approximately six 
feet away from the rocket nozzle exit plane. Before we go into 
these components and their design logic, let us have a look at 
the crew compartment. 

Figure F-7 shows the structure of the crew compartment which, 
however, is not designed to structures requirements at all for a 
very unique reason. Severe noise levels caused by the captured 
shocks between booster and orbiter noses require this component 
to be designed for noise suppressions, e.g. with heavy skins. 

The crew compartment and the electronic compartment behind it 
are suspended from the outer heat shield shell which will expand 
substantially during heat-up in re-entry. The shell is conventional 
aluminum alloy construction. We have fused silica glass windows 
and there is a local area around them which is designed to 
the rmo- structural requirements. The side enclosures can be swung 
open and the crew can eject out to both sides. This is required 
since the orbiter is mounted on top. Seats can be swiveled for 
vertical and horizontal flight. 

We have stringers, very few of them, on the inside and the 
frames are on the outside. The whole enclosure is enveloped with 
low-density fiberglas insulation and 1" dynaflex over it to maintain 
the temperature environment in the compartment. 

Figure F-8 shows the forward tank extension with nose gear 
support and a large JP tank located here for stability during 
re-entry. The forward segment of the heat shield is supported 
here with two links and guide rollers. When we re-enter, the heat 
shield which is bolted to the sub-structure at the rear end of 
the lox tank, expands at this point approximately 6" forward. 
Tank extension construction material in the forward portion is 
202U. In the aft portion it is 2219, principally for compatibility 
with the wall of the lox tank which will see temperatures of 
-320°F. 



I-8U 



Figiire F-9 shows the lox tank which is made of 2219 aluminum 
and designed principally for fracture mechanics considerations. 
The nose gear loads do not tax this structure very much at all 
and there is a minimum amount of stiffening. We have integral T 
stiffeners longitudinally and truss-type frames. The reason for 
not going to a web-type frame is, of course, that the tank shrinks 
and we had to tailor the stiffness of the frame very carefully 
in order to prevent large pull forces on the tank stiffeners in 
chill down. The rather massive bulkhead at the orbit er forward 
attachment was mentioned before. 

All fittings are designed of two elements. Why not three for 
fail-safe? The reason is that we always would have covered one 
of the elements and therefore could not have inspected one of the 
elements visually. 

The lateral load on the forward orb iter attachment is 1.25 
million pounds limit and also the vertical load is very substatial 
but, nevertheless, all this does not design the structure here 
because stiffness requirements are critical. We had to meet a 
natural frequency with the orbiter mounted on top of approximately 
.8 cps. We made .9 cps after many trials and many computer runs. 

No large thermal stresses are encountered on a lox tank 
because the whole structure system will cool down simultaneously. 
We are definitely working in the mixed mode as far as fracture 
mechanics is concerned. Tank walls are approximately l/8" 
in the upper portion. There are also quite a number of baffles 
mounted in the tank. 

Figure F-10 shows the intertank section. It connects the 
LO2 tank with the LH2 tank. It is crammed full with equipment. 
Again we are dealing with an aluminum structure, local reinforcements 
in titanium, and here, fortunately, we can go inward and get our 
stiffness by depth. So, no advanced high stiffness materials are 
employed . 

The canard and the orbiter drag link are attached in this 
region. The all-movable canard is streamlined during re-entry 
since reaction controls are used. This effectively eliminates 
temperature differences between upper and lower surfaces. Regarding 
the critical orbiter connections we are becoming quite cautious 
and are toying with the idea to subject any of the critical 
orbiter connecting elements to proof loads for qualification. 

Figure F-ll shows the IH2 tank, made of 2219 aluminum, 
approximately 120 feet long and 33 feet in diameter. On the 
outside of the tank we have the frames, including the orbiter 
aft attachment bulkheads . 



1-85 



These two bulkheads experience their main load at the moment 
of separation but, interestingly enough, the exceedingly turbulent 
flow around these vehicles forced us into the ultimate of stiffness 
design on the frames because otherwise we could probably have 
fatigued this area half-way through the first flight at the point 
where the stiffeners join the frames. Whereas we quite succeeded 
in meeting the overall natural frequency requirements for the 
combined booster and orbiter with aluminum frames in this location 
here, we now had to beef up the caps with beryllium for local 
stiffness . 

Of course, we could take other materials here. It is beryllium 
because its price has come down substantially and we laminated it. 
Even if we should have a crack in one of the laminations we still 
do not lose our stiffness because a very localized area would be 
effected. Of course, we could have taken another route by lowering 
the height of the stiffeners joining the bulkhead here but this 
would have resulted in a substantial weight penalty. 

Figure F-12 shows the landing gear support structure. Essentially 
it is a wheel well, structurally fully enclosed. Vertical and drag 
links introducing the loads into bulkheads and into longerons 
which are integral with the tank wall provide an attachment system 
which permits the decrease in dimensions which takes place when 
we fuel the tank. 

But principally we selected this system because of the need 
for a meaningful drop test arrangement without involving too much 
of the vehicle. The arrangement permits drop testing the landing 
gear separately and with a concrete barrel on top of it because 
it is supported in a determinate fashion and loads into the 
structure can be determined accurately. The cost of testing 
requires very careful consideration and is quite important for 
the structural design of future large vehicles. 

Figure F-13 shows the wing-body-TPS arrangement. A drag 
brace taking the drag loads into the longeron of the tank walls 
is the only fore and aft restraint of the wing and the horseshoe 
heat shield attaches directly to the root rib here. Wing and 
heat shield expand from this attachment point. All the dimensions 
are steadily changing and there are slip joints to arrive at an 
expansion-compatible structures system. 

A metallic seal at the fore and aft end of the TTS horseshoe 
section is provided by a bellows. We have corrugated skins all 
over this vehicle because aerodynamic smoothness is of little 
concern. As a matter of fact, if one would start with a smooth 
surface on a high-temperature vehicle subjected to flash heating, 
after the first flight you would not have it any more . This 
necessary approach makes aircraft aerodynami cists shudder. 



1-86 



Figure F-lU shows the cryogenic insulation system. Again, it 
is something rather unique. All the cryogenic insulation systems 
of the past have been one-shot systems but here we need long-time 
reliability because maintenance in a tank, which is 10 stories high 
and has 33 feet diameter / could be prohibitive. 

An external foam insulation system would not work because its 
specific heat is such that it would melt due to the high outside 
temperatures , in spite of the short heating times . So we need the 
heat sink of the structure to more or less protect the foam. We 
have a gas layer system without an inner liner, reconciling ourselves 
from the beginning with the fact that even with a plastic liner 
hydrogen would penetrate sooner or later during a 10-year life span. 

The unique polyphenyene oxide foam has a capillary system into 
which the hydrogen penetrates. A gas layer is formed which is 
kept in equilibrium by capillary forces at the surface. The system 
is structurally beautiful because it does not impose large shear 
forces on the adhesive joints. This is very important because it 
is a tough proposition to hold on a cryogenic insulation system 
at -U21°F and we have more than a third of an acre of it. Besides, 
all the small columns of the system are so nicely stabilized 
against each other that a heavy man can walk on it. 

Figure F-15 shows the thrust structure. This is a fail-safe 
system with multiple load paths, partly with trusses, partly with 
shear webs providing a benign mode of failure which a truss does 
not give us. The basic structure is titanium, reinforced by 
boron -aluminum. Their coefficients of thermal expansion are 
close, 5-0 x 10"" and 3-1 x 10"°, respectively. Graphite epoxy 
with a coefficient of 1 x 10 _D is not compatible with a metallic 
structure and we would have to make the whole article of graphite 
epoxy. 

The problem is also that boron epoxy and graphite epoxy are 
deficient in high-temperature applications . If graphite epoxy is 
heated to 350°F in the presence of moisture its strength will 
deteriorate. We have tested boron aluminum rather successfully 
up to 700°F. There is an additional consideration. You might 
have high temperature only for a short time but you need a certain 
heat sink capability to prevent the temperature from going up 
too fast. Lately we have done exceedingly well with boron and 
graphite polyimides which would have a substantial cost advantage 
but we were not quite sure of them at the time when we conducted 
this study. 

Let me point out the importance of weight savings . For every 
pound of advanced composite we save one pound of weight in 
stiffness critical areas. We attempt to put one pound of weight 
in orbit for approximately $75, say $100. Let us look at the 
orbiter first. Without any change in sizing, with 100 flights 
at $100 a pound we generate a value of $10,000. Assuming an 
average 60$ payload factor, this makes it $6000 per pound or, to 
show definitely a profit, say $5000 per pound. 

1-87 



For this particular booster we need a 5 pounds saving to put 
one more pound of payload into the orbiter. This means for the 
booster one pound of weight saved is approximately $1000. It 
shows that advanced materials at a cost of about $350 per pound 
don't have to frighten our community as long as we apply them in 
such an economical fashion. The installed cost of aluminum is 
anywhere between $35 to 150 per pound, depending on the quantity 
of production. As we are progressing into this new world of 
advanced materials, we will probably employ them in more complex 
shapes. Right now, the easiest way to employ them economically 
is just as reinforcements in localized areas. I am afraid we 
will always have a mix of many materials. The world is not 
built out of one material . 

Figure F-l6 shows the base heat shield which is very lightly 
loaded. Here we are really going all-out with exotic materials. 
There are beryllium truss systems, providing a quadruple load path 
and in a conservative fail-safe approach. The performance of these 
trusses on tests was just fantastic and columns with an L/p 
approaching 300 are somewhat of an eerie sight. We have them 
ball- jointed to prevent lateral loading which shows what you can 
do if you watch the peculiarities of a material. 

The outer region of the bulkhead consists of corrugated 
rene kl panels while the center panels consist of coated corrugated 
columbium. The reason is that the temperatures at the periphery 
are reduced due to air circulation, particularly on the launch 
pad. These are smaller panels, different from the heat shield 
arrangement on the rest of the vehicle. It is a slip- jointed 
system. 

Figure F-17 shows the wing arrangement, housing the engines 
with cutouts over a large portion of the structure. The acoustic 
environment of the engines with an output of around 172 decibels 
requires high stiffness surface panel construction in this area 
which merits every type of scrutiny we could put on it. The 
typical wing structure appears to be a throwback to old days 
with heavy concentrated load members of titanium and semi- 
structural covers. We have a fail-safe system of ribs and spars 
and corrugated surfaces with the corragation running in the fore 
and aft direction. 

The reason for this arrangement becomes apparent when we look 
at the wing structure in Fig. F-l8. The skin of Inconel 7l8 goes 
up to 1350°F and can expand and flex between the attachments at the 
node points while the heavy titanium spar cap has a maximum 
temperature of 300°F . We have to get used to thinking in terms 
of compatible heat-up because a panel could quickly deform to an 
unacceptable level and move into a critical panel flutter mode. 



1-88 



Figure F-19 shows the vertical stabilizer which consists of 
conventional, integrally stiffened type of titanium construction. 
The thermal environment is rather mild because the vertical surface 
is shaded during re-entry. Only the leading edge is heat-sinked 
to cope with plume impingement from the orbiter . 

Figure F-20 shows the canard structure which again is rather 
conventional. The internal structure and substructure skin are 
titanium but skin gages and outer surface material are determined 
by thermal environment considerations. This vehicle flies by 
wire so there are no linkages to the crew department. 

Figure F-21 summarizes the design criteria for the booster 
TPS. However, the creep factor has been lowered in the meantime 
to 2 life cycles with concurrence of NASA because the penalty 
would have been very severe and all structures in the TPS are, 
after all, but secondary types of structure. With a cumulative 
creep of .2% and a factor of 2, our plastic deformation in actual 
flight experience will be .1%. Panel flutter might become 
critical after such plastic deformation. Flight monitoring and 
replacement, if required, should be considered since actual 
conditions cannot be simulated in laboratory type tests . 

Figure F-22 shows the pressures to which the outer TPS is 
subjected. They are quite substantial on the lower surface but 
rather mild on the upper surface. 

The heating environment is shown in Figure F-23 and you can 
see that we are only approximately 100 seconds above 1000 F, 
that's all. So you heat up and you cool down and everything on 
the inside will lag substantially. Consequently, the influence 
of the local heat sink on or near the outer surface will assume a 
remarkable influence on the temperatures which we actually 
experience . Figure F-24 shows a plot of TPS temperature versus 
thickness for corrugated and smooth TPS. With increasing gage 
thickness the temperature goes down so you really can tailor 
your temperatures for a vehicle like this . 

Now let us look to what we have tailored these temperatures. 
Figure F-25 shows the booster body temperatures for the TPS . We 
can use rene kl throughout and heat sink it down to the criticality 
of this material. The required maximum gage we have determined 
is .058. On the wing, the situation is a bit different and rather 
serious due to nose shock impingement.. Temperatures and various 
materials which have been used on the aerodynamic surfaces are 
shown in Fig. F-26. 

Figure F-27 shows what this flash heating will do to various 
types of TPS panels. This is very interesting. The total stress 
consists of thermal stresses due to temperature gradient and 
bending stresses due to air loads. For the first configuration 



I-89 



of .016 rene Ul the total stress is 1+3,700 psi compared with an 
allowable of UlOO psi. So we applied heat sinking and went to 
.055 rene Ul which gave us a total stress of 37,700 psi compared 
with an allowable of 13,700 psi. We still could not make the 
grade. 

We found out that in flash-heating one cannot employ the type 
of structure which we have used since the beginning of metal 
aircraft design. So we went to the third configuration shown in 
Fig. F-27. Because of the reduced moment of inertia it forced 
us to a much closer frame spacing on the lower surface but with 
essentially everything working at the same temperature we had a 
total stress of 11,800 psi compared with an allowable of 
13,700 psi. 

Figure F-28 shows the frames supporting the body TPS which 
are about 9" deep. We first had a titanium frame with a corrugated 
web but after we had found out the very high significance of 
heatsinking, we went to a quite conventional aluminum construction 
with additional beryllium caps to maintain the temperature low 
enough . 

Regarding the application of beryllium I would like to point 
out that the principal aluminum beam is designed to take limit load 
plus. However, we are stiffness critical. The weight of the 
beryllium caps on these frames is in the order of 6000 pounds but 
the total weight saving is far in excess of that. As far as 
structural reliability is concerned, even a full crack across the 
beryllium cap would have very little influence on the stiffness 
of the beam. 

Figure F-29 shows a panel TPS which we had to add on the 
wing as we went from 10,200 to 10,800 ft/sec staging velocity. We 
have flex clips for connection to the very substantial spar cap 
which is really a heat sink arrangement employed to keep strength 
allowables high with low temperatures. It is a system of floating 
panels on the outside, ^0" long. The material in the shock 
impingement area is coated columbium and the rest of the wing 
lower surface is HS-188. 

Fortunately, with the reduction in staging velocity we can 
eliminate this TPS and even the body TPS, going to heat -sink tank 
walls. The weight penalty for going from a staging velocity of 
8000 ft/sec to 11,000 ft/sec is of the order of 50,000 pounds 
for this booster, in addition to the dramatic effect on design 
criteria. 



1-90 



Figure F-30 shows the materials we use on this vehicle and 
the limitations we impose on them. With all the lessons we have 
paid for we are quite reluctant to jump in with new materials 
except maybe by first getting our feet wet with a semi-structural 
application. All this refers to the B-9U vehicle; the lower 
staging velocity vehicles are not as demanding. As a matter of 
fact, all these exotic materials fall more or less by the wayside 
with the lower staging velocities . 

What I wanted to present to you was one of the most difficult 
vehicles that we treated and I hope that I have inspired your 
thinking along these lines a little bit. It is a completely new 
ballgame . 



1-91 



BOOSTER STRUCTURAL ARRANGEMENT 




WING/FUSELAGE 
ATTACH LINKS 



CREW 
COMPARTMENT 



TPS 

FORWARD ATTACH LINK 
N.L.G. JPTANK 
SUPPORT 
STRUCTURE 



Fig. F-l 



B-9U HEAT SHI ELD SHELL 



O 



FIXED SHELL SECTION (TYP) 
(UNIDIRECTIONAL GROWTH 
CAPABILITY)- 



LONGITUDINAL 

GROWTH 

DIRECTION 



ASSEMBLY SPLICE 



NOSE SECTION 



N.L.G. DOOR' 




'ASSEMBLY SPLICE (TYP) 
'SUPPORT LINKS (TYP) 
"CANARD FAIRING 



ROLLER RECEPTICLE 
FITTING (TYP) 



FIXED AFT 
UPPER SECTION 



FIXED LOWER 
T.E. SECTION 



Fig. F-2 



1-92 



DESIGN CRITERIA 
DESIGN FACTORS 



© 





FACTOR OF SAFETY 


PROOF 
FACTOR 




COMPONENT 


YIELD 


ULTIMATE 


APPLIED ON 


RArN PROPELLANT TANKS 

PERSONNEL COMPARTMENTS 

WINDOWS, DOORS, HATCHES 

AIRFRAME STRUCTURE 

PRESSURE VESSELS 

PRESSURIZED LINES + FITTINGS 

ALL COMPONENTS 
(ABORT CONDITIONS) 

ALL COMPONENTS 
(THERMAL STRESSES) 


■ 


1.10 

1.10 

1.00 

1.10 
1.50 
1.00 

1.10 
1.10 

1.10 

1.00 
1.00 


1.40 
1.40 

1.50 
2.00 

3.00 
1.40 
1.50 

2.00 
2.50 

1.40 

1.00 
1.25 


F.M. 

1.50 

2.00 
TBD* 
TBD* . 

1.50 
1.50 


MAX. OPERATING PRESSURE 
+ DYNAMIC HEAD. 
LOADS ( + LIMIT PRESSURE) 
PROOF PRESSURES 

LOADS (+ LIMIT PRESSURE) 

MAX. OPERATING PRESSURE ONLY 

PROOF PRESSURE 

MAX. OPERATING PRESSURE ONLY 

BOOST & ENTRY LOADS 

AIRCRAFT MODE LOADS 

MAX. OPERATING PRESSURE • 
MAX. OPERATING PRESSURE 

ABORT LOADS (+ LIMIT PRESSURE) 

THERMAL FORCES (+ FLT. LOADS) 
THERMAL FORCES (ALONE) 



F.M. = FRACTURE MECHANICS 
ASSUMED SERVICE LIFE - 100 MISSIONS 

♦TO BE DETERMINED ON AN INDIVIDUAL COMPONENT BASIS DEPENDING ON DESIGN APPROACH AND COMPONENT 
CRITICALITY. 61CV6360 



Fig. F-3 



BOOSTER B-9U 

PEAK LIMIT LOAD INTENSITIES 



o 



1 1 HR GROUND HEADWINDS 

2 1 HR GROUND TAILWINDS 

3 1 HR GROUND SIDEWINDS 

4 LIFTOFF + GROUND HEADWINDS 

5 LIFTOFF + GROUND TAILWINDS 

6 LIFTOFF + GROUND SIDEWINDS 

7 MAX ALPHA-Q HEADWINDS 

8 MAX ALPHA-Q TAILWINDS 

9 MAX BETA-Q (2400) 

10 3G MAX THRUST 

11 BOOSTER BURNOUT 

12 BOOSTER RECOVERY 

13 BOOSTER SUBSONIC GUST 

14 BOOSTER 2 POINT LANDING 

15 BOOSTER 3 POINT BRAKED ROLL 

16 BOOSTER 2 G TAXI 

17 1 DAY GROUND HEADWINDS 

18 1 DAY GROUND TAILWINDS 

19 1 DAY GROUND SIDEWINDS 

20 TWO WEEK GROUND HEADWINDS 

21 TWO WEEK GROUND TAILWINDS 

22 TWO WEEK GROUND SIDEWINDS 

23 BOOSTER 2.5 G POSITIVE MANEUVER 

24 BOOSTER -1. G NEGATIVE MANEUVER 

25 BOOSTER MAX OPERATING PRESSURE 




O -4000 - - o 



I- 




S* 



■' cor- 

1 

\r > COND. 9 \ l 

^ifr_^ \-y\ ) 

A COND. 3^ "s? 

U, COND. 8— -y \Jr[± 



/ 

COND. 10 



COND. 6 /Jf 
X CONO. 11 



Fig. F-U 



6ICV6364 



1-93 



BOOSTER STRUCTURE DESIGN APPROACH 




7 



COMPONENT 


INHERENT FEATURES 


DESIGN 


PROOF TEST 






APPROACH 


REQUIREMENT 


LO„ h LH TANKS 
2 2 


NOT PRACTICAL TO DESIGN FAIL-SAFE 




LO o= 1 23 
LH 2 , o= 1 13 


FOR PRESSURE. STRINGERS ACT AS 
CRACK ARRESTORS FOR FLIGHT LOADS 


SAFE-LIFE 


WING 


HIGH FAIL-SAFE CAPABILITY 


FAIL-SAFE 


... 


CANARD BOX 


SOME FAIL-SAFE CAPABILITY 


SAFE-LIFE 


... 


CANARD PIVOT 


IMPRACTICAL TO DESIGN FAIL-SAFE 


SAFE-LIFE 


... 


TPS 


IMPRACTICAL TO DESIGN FAIL-SAFE 


SAFE- LIFE 


... 


VERTICAL TAIL 


HIGH FAIL-SAFE CAPABILITY 


FAIL-SAFE 


... 


CREW CABIN WALL 


FAIL-SAFE WITH TEAR STOPPERS ONLY 


FAIL-SAFE 


a - 1.5 


ORBITER SUPPORT 


FAIL-SAFE WITH MULTI-ELEMENT CAPS 


FAIL-SAFE 





FRAMES 


AND WEBS 







Fig. F-5 



BODY - STRUCTURAL COMPONENTS 



© 




Fig. F-6 
I-9U 



61CV6226 



CREW COMPARTMENT 



O 



FWD ATTACH 
PTS 




INTER-COMPARTMENT 
BULKHEAD 



CREW ACCESS HATCH 

Fig. F-7 



FORWARD SKIRT STRUCTURE 



TPS GUIDE 
ROLLER 

MACHINED FRAMES 
Al ALLOY 



FORWARD JP TANK 
AND SUPPORTS -=d 



N.L.G. SUPPORT 
STRUCTURE TITANIUM 
TUBES 




LANGE 
FOR ATTACHMENT 
TO L0 2 TANK 



BUILT-UP SHELL 
STRUCTURE 
Al ALLOY 



Fig. F-8 



1-95 



LIQUID OXYGEN TANK 



© 




^—INTEGRAL 
STIFFENER 



BOLT FLANGE 
DETAIL-TYPICALN 
BOTH ENDS 



ANTI-VORTEX 

BAFFLE 

INSTALLATION 



RECIRCULATION FLOW 
DIVIDER 



RADIAL 

SUPPORT BEAMS 
ORBITER FORWARD 
ATTACH BULKHEAD 



ANTI -SLOSH BAFFLES (8) 

MATERIAL: 2219 A 
ALLOY 





SUMP 



, L0 2 DUCT 
^/' /FLANGE (4) 

\ - 

DOME 
STIFFENERS 



SUMP AND ANTI-VORTEX 
BAFFLE DETAIL 



LAUNCH TOWER 
ATTACH PROVISION 



TEGRALLY 
STIFFENED SKINS- 
WELDED ASSEMBLY 




TYPICAL STABILIZING 
FRAME DETAIL 



Fig. F-9 



INTERTANK ADAPTER STRUCTURE 



ACCESS DOOR 



CANARD SUPPORT 
BEARING - OUTBOARD 
(BOTH SIDES) 



INTERMEDIATE 
FRAMES (2) 



CANARD 

ACTUATOR ACCESS 
(BOTH SIDES) 

LOoTANK ATTACH 
FLANGE 

ORBITER SUPPORT FRAME 




ORBITER DRAG LINK 
ATTACH FITTINGS (2) 
TITANIUM 

ORBITER SUPPORT FRAME 
^CANARD SUPPORT 
FRAMES (2) 

CANARD SUPPORT 
BEARING - INBOARD 
(BOTH SIDES) 



LO, SUPPLY LINE 
EXIT (BOTH SIDES) 



STIFFENED INTERTANK 
SKIN 



LO2 SUPPLY LINE 
EXIT (BOTH SIDES) 



Fie. F-10 



I-96 



LIQUID HYDROGEN TANK 



© 



MILLED INTEGRALLY 
STIFFENED SKINS - 
WELDED ASSEMBLY 
ORBITER LINK ATTACH FITTINGS 
ORBITER AFT ATTACH BULKHEADS 

FRAME STABILIZING 
LINKS (TYPICAL) 

BOLT FLANGE ,-£: 

BOTH SIDES^/fe^C^ 



STABILIZING FRAMES 
(TYPICAL) 




WING DRAG LINK 



TPS SUPPORT 
FRAMES 



DETAIL OF WING - FUSELAGE ATTACHMENT 



Fig. F-ll 



MAIN LANDING GEAR SUPPORT STRUCTURE 



WHEEL WELL TO 
LH 2 TANK 
SIDE LOAD LINK^ \DRAG LINK 



CARRY-THROUGH 
OR "SADDLE" BEAMS 
(2) TITANIUM 
FORWARD LINKS 
VERTICAL & SIDE LOADS 



BULKHEAD 
STATION 3020 

M.L.G. DRAG STRUT TO 

WHEEL WELL FITTING 





STRUCTURAL 
WHEEL WELL 
ENCLOSURE 
Al ALLOY 



BULKHEAD 
STATION 3295 

/-" 

VERTICAL LOAD 

LINK 
DRAG LONGERON 
ON LH 2 TANK 



M.L.G. MAIN 
TRUNNION FITTING 
(2 PER SIDE) 



Fig. F-12 



1-97 



WING / BODY / TPS ARRANGEMENT 



O 




WING-TO-BODY 
SIDE LOAD LINK 



WING-TO-BODY 
FILLET FAIRING 



WING-TO-BODY 
VERTICAL ATTACH LINKS 



WING DRAG LINK 



WING-TO-BODY 
SIDE LOAD LINK 



DRAG LOAD 
TANK LONGERON 



Fig. F-13 



CRYOGENIC INSULATION LH 2 TANK 



ADHESIVE 

CREST 7343/1% SI LANE 

OR EPON 934 




PPO (POLYPHENYENE OXIDE) FOAM 
30" X 18" PANELS 



PANEL JOIN 
ADHESIVE BONDED 



THRUST STRUCTURE 



BORON/ALUMINUM COMPOSITE 
THRUST POSTS (12) 



TITANIUM SKIN 



BORON/ALUMINUM 
COMPOSITE STRINGERS 



THRUST SECTION 
LH 2 TANK ATTACH ,*■' 
FLANGE -^ 



TITANIUM 
BULKHEADS 



WING ATTACH 
FITTING -TITANIUM 




VERTICAL STABILIZER 
ATTACH FITTINGS - TITANIUM 

HOLDDOWN 
FITTING (4) 
TITANIUM 



TITANIUM - BORON/AI 
COMPOSITE REINFORCED 
THRUST BEAMS (4) 



> BORON/ALUMINUM 

COMPOSITE TRUSSES (TYP) 



Fig. P-15 



BASE HEAT SHIELD 



O 



RENE' 41 BASE BULKHEAD 



RENE' 41 SUPPORT CHANNELS 

TITANIUM CORRUGATED SKIN 




COATED COLUMBIUM CORRUGATED PANEL 
(CENTER FOUR PANELS ONLY) 



RENE' 41 CORRUGATED PANELS 



RENE' 41 REINFORCING RING 



BERYLIUM TRUSS TUBES (TYP) 

Fig. F-l6 



1-99 



WING ARRANGEMENT AND SUPPORT POINTS 



O 



LEADING EDGE 

COATED 

COLUMBIUM 



SPAR 
TRUSS 
TITANIUM 



JP FUEL TANK 



SKINS TITANIUM CORRUGATED 

ACPSTHRUSTERS 

'SIDE LOAD LINKS 




BODY SUPPORT ATTACH 
FITTINGS TITANIUM 

SPAR 

TITANIUM CORRUGATED 



E LEV ON 



TITANIUM 



DETACHABLE 



Fig. F-17 



WING STRUCTURE 

MULTISPAR CORRUGATED SKIN CONCEPT 



© 



MULTISPAR PRIMARY 
BENDING STRUCTURE 



TITANIUM WEB 




300° F 
MAX. 



ACCOMMODATES SKIN 
EXPANSION RELATIVE 
TO SPAR 



1-100 



VERTICAL STABILIZER 




CORRUGATED 
SPAR WEB 



SPAR-TO-THRUST 
STRUCTURE ATTACH 
FITTINGS 



Fig. F-19 



CANARD STRUCTURE 



Q 




ACTUATOR TRUSS 
SUPPORT SYSTEM 



VENT HOLES 



ACTUATORS 




CARBON-CARBON 



• SUBSTRUCTURE SKIN & INTERNAL STRUCTURE - TITANIUM 

• UPPER SURFACE TPS OUTER SKIN - COATED COLUMBIUM 

• LOWER SURFACE TPS OUTER SKIN - RENE 1 41 




LEADING EDGE 
TRUSS RIBS 



Fig. F-20 



1-101 



DESIGN CRITERIA, BODY TPS 



LIFE 100 FLIGHTS, NO REFURBISHMENT 

AIRLOAD/TEMPERATURE HISTORY 

ACOUSTICAL ENVIRONMENT 

VIBRATION 

PANEL FLUTTER 
FACTOR OF SAFETY = 1.4 
FATIGUE FACTOR OF 4 ON CYCLES 
CREEP FACTOR OF 4 ON LIFE 
TOTAL CUMULATIVE CREEP - 0.2% 
LEAKAGE 350 IN. 2 EQUIVALENT 



Fig. F-21 



PANEL PRESSURES, TPS SHELL 



STATION 1000 1458 2040 2510 

0.163 0.342 |o.506 




LIMIT PRESSURES (PSI) 




LOAD 
CONDITION 


STATION 
1485 


STATION 
2040 


STATION 
2510 


STATION 
3800 


MAXIMUM aq 
ASCENT 

P„ = 6 - 5 P'i 
P lntern a l = 6 -° P sl 


-1.30 
-1.20 


-1.00 
+ A-0.90 

-.70 


-0.90 

rv 

+ — V-0.4^ 

w 

-0.40 


-1.50 
-0.90 


MAXIMUM g 
ENTRY 

p^ = ps! 

P. i = P s ' 
'internal 




+2.20 


+0.03 
+ 1.95 



+2.00 




a- 

+ 1.80 


MAXIMUM aq 

ENTRY 

P„ = P si 

P- . i = psi 
internal 


+0.08 
+2.10 


+0 05 
+ 1.50 


+0.03 
+ 1.65 


+0.02 

+ — 1+0. 10 
+ 1. 10 



M DENOTES CRUSHING PRESSURE 



Fig. F-22 



1-102 



BOOSTER TEMPERATURE PROFILE 



TEMPERATURE (F) 




100 200 300 

TIME (SECONDS) 



400 500 



BODY LOWER SURFACE, 6= DEGREES 



Fig. F-23 



r 



TPS TEMPERATURE VS. THICKNESS 



O 



PEAK TPS 
TEMPERATURE T(F) 



2100 
2000 
1900 
1800 
1700 
1600 
1500 
1400 
1300 
1200 
1100 
1000 




■ I I I 1 



0.05 0.10 

TPS THICKNESS, t (INCHES) 



0.15 



Fig. F-2U 



1-103 



B-9U BOOSTER BODY TEMPERATURES 



© 




1,700 
STA. 1400 



Fig. F-25 



THERMAL PROTECTION SYSTEM (AERO SURFACES) 



O 



TEMPERATURE (F) 




510 A 



A INNER SKIN TEMPERATURE 



GZZZZZZZ3 TITANIUM 
' 1 INCONEL 718 



fSS\SVASS3 RENE 41 
llllllllllllllll HS 188 



3 COATED COLUMSIUM 
CARBON/CARBON 



•TEMPERATURE INCLUDES 20OF FOR 
EFFECT OF SEMI-FLAT SURFACE 



Fig. F-26 



I-lOU 



TPS PANEL CANDIDATES 




LOCATION 
ON TPS 



LOWE! 
SURFACE 



SIDE 



SIDE 

& 

TOP 



PANEL CONFIGURATION 



MATERIA! 



1.25 JJ 



1.251 



-6.00- 



i.oo! y 



imJZjT 



p— 4.60 — >j 



RENE' 41 



RENE' 41 



RENE' 41 



INCONEl 
718 



TITANIUM 
6AL-4V 



SHEET 

THKNESS 

(F) 



.016 



.055 _ 



.055 



.035 



.030 



TEMP 



1.750 



1,600' 



1,600* 



1,180' 



695' 



MAX. 
THERMAL 
GRADIENT 
STRESS 

IPSU 



39,200 



35,700 



44,500 



6,000 



«aj INCLUDES EFFECTS OF HEAT SINKING AND 
SURFACE GEOMETRY 



TOTAL 
STRESS 
(PSD 



43.700 



37,700 



11,800 



59,000 



17,500 



ALLOW 
STRESS 
(PSI) 



4,100 



13,700 



13,700 



64,500 



18,500 



Fig. F-27 



INTERNAL STRUCTURE - BODY TPS 



O 




MECHANICAL 

SKIN 

FASTENERS 



INTEGRALLY 
MACHINED STIFFENER 

BERYLLIUM CAP 

■SEMI -SMOOTH 
RENE' SKIN 



SECTION A-A 



A !► 
SUB-FRAME 



Fie. F-28 



1-105 



r 



WING LOWER SURFACE TPS 



O 



SPAR CAP 



CORRUGATED 
INNER SKIN 

INSULATION 



OUTER SKIN 
(SEMI -FLAT) 



STAND-OFF 




NBD 



Fig. F-29 



BOOSTER TPS MATERIAL TEMPERATURE LIMITS 



o 



MATERIAL 



TEMP. 
(°F) 



TITANIUM 



CARBON/ 
CARBON 



800/ 
650 



INCONEL7I8 1,350 

RENE' 41 1,650 

HAYNES 188 1,900 

TD NICr 2,200 

COLUMBIUM 

752 2,500 



3,000 



RATIONALE 



STRESS CORROSION CRACKING LIMITS USE AT TEMP. ABOVE 650°F TO VERY SHORT 
TIME & LOW STRESS. HIGHER LIMIT OF 800°F MAY BE CONSIDERED FOR BOOSTER 
COMPONENTS WHERE TIME & STRESS CONDITIONS PERMIT 

l,350 o F IS AGE TEMP. OPERATION ABOVE THIS TEMP. SEVERELY DEGRADES 
MATERIAL 

ABOVE 1,600°F, INTEGRANULAR OXIDATION OCCURS. TIME ABOVE THIS ON 
BOOSTER IS SHORT ENOUGH TO BE ACCEPTABLE 

CREEP LIMITED. TIME AT TEMP. AFFECTS TOTAL CREEP. BOOSTER ACCEPTABLE 
BECAUSE TIME IS APPROXIMATELY 1 MINUTE PER MISSION 

LIMITED BY CREEP & HIGH TEMP. STRAIN CAPABILITY. TIME AT TEMP & LOADS 
OF BOOSTER PERMIT HIGHER TEMP. APPLICATION 

2,500°F IS UPPER LIMIT FOR COATING REUSABILITY 
BASED ON PLASMA-ARC TESTS 



Fig. F-30 



1-106 



STRUCTURAL DESIGN CONSIDERATIONS FOR 
ADVANCED AIRCRAFT 



J. E. Fischler 
Douglas -Long Beach 



Only the figures and an abstract are available. 



Abstract: Structural Engineers must continue to investigate ways 
to save structural weight without losing structural 
integrity. Two computing programs were used to 
determine a basis for comparing advanced materials and 
types of construction with existing aluminum structure. 
One computing program determines the rate of crack 
growth from the material characteristics (Kq = critical 
stress intensity factor, R » stress range, AK = stress 
intensity factor, B and n = material characteristics 
and type of construction, and R c t = stress ratio of an 
unstiffened panel to one with stiffeners and/or crack 
stoppers). A second program solved for the probability 
of structural failure using the mean 1 g stress and 
coefficient of variation of mean load (dependent on the 
type of mission), mean strength and coefficient of 
variation of mean strength, and residual strength which 
decreased as a function of the crack length and panel 
stiffening at any time after first flight. Ways were 
shown to calculate the probability of initiation of a 
crack and its percentage of total fatigue damage. For 
ease of comparison, and reduction of computer time, 
all the materials compared were based on an undetected 
crack of 0.2 inches initially. The theoretical results 
were compared to actual tests on a 96 inch wide stiffened 
panel. Modifications to the theoretical coefficients 
were recommended due to practical considerations of the 
slowdown in crack growth when the crack tip reaches an 
attachment, when a stiff ener fails, when the fracture 
toughness, Kq and AK values vary with crack length, when 
the value of R c ^ varies with stiffener area and crack 
length, and when the stress spectra has occasional 
high stress cycles that modify the plastic zone ahead 
of the crack tip and actually slows down the rate of 
crack growth for the average cyclic stresses. 



1-107 



The materials compared were aluminum (202U-T3, 7075-T6), 
titanium and graphite epoxy. The great weight advantage 
of graphite -epoxy at as low as 10,000 psi 1 g stress and 
the low probability of failure allowed many more flights 
and longer inspection intervals (thus reducing Maintenance 
Costs). 

Desirable future analysis methods, material characteristic 
and type of construction research, and other structural 
research plans were illustrated as areas for Structural 
Development to reduce the structural weight and improve 
the Structural Reliability (Lower the Probability of 
Structural Failures ) . 

One of the most desirable results of the study are the 
charts that have been prepared that represent over two 
hundred data points . These charts give the Structural 
Analyst the opportunity to compare materials at different 
stress levels, mission load variations, mean strength 
variations, number of flights, structural efficiency, 
residual strength, and variation of material parameters 
all versus the Probability of Failure. These overall 
views of various materials and types of construction 
can be very useful for choosing a material and also 
determining what future research is essential for 
saving structural weight and improving structural 
reliability. 



1-108 



DIRECT OPERATING COST = CENTS/SEAT-MILE 
RETURN ON INVESTMENT = PROFIT PER YEAR/COST 



DOC 



FLIGHT OPER COSTS + DEPRECIATION + FLIGHT INSURANCE + MAINTENANCE + FUEL COSTS 
RANGE x PAYLOAD 



RETURN ON INVESTMENT 



PROFIT PER TRIP x NUMBER TRIPS PER YEAR 
ORIGINAL COST PER AIRCRAFT 



RANGE = 2.3 -^ (L/D) MAX LOG 10 — - 



WEIGHT AT INITIAL CRUISE 



WEIGHT AT FINAL CRUISE 



Fig. G-l 




POTENTIAL PROFIT - 
THOUSANDS 
OF DOLLARS 




11 


"INTERCONTINENTAL" VERSION 
"INTERNATIONAL" FARE STRUCTURE 








10 






PASSCNGOl LOAD 


FACRM ~tMl 


- i 




/ 












_^JSP8 












(rf^^T^,. iw"'"'j£^sstl 










M^^^i 










■^. ''^.'jl*^ 










" "ZSSfl 




















"^ ^^^* F. ■■'''^M 










■ 


BB^---" 




ACTUAL "WEAK 






mrm®'.*-'^ 


_1 


orar une — - 


■ ■ ■ — 








20 


00 29 


00 


an 


0* HM 



- 



■■ 



RANGE -NJIL 



''s$m& 




■■4, 



Fig. G-2 



1-109 



DISTRIBUTIONS OF COST AND WEIGHT 



COST 






' "■ """ ■ '. '. - ' ■ " ■ '. ■ ' / . ■ '. ■"■'•' . ■' . ■ r , / '. r / . 



■'.■'.■'.■'.'.■'.■'.'.■.. "° '■ ■'.■'.-'/.'.■ ■'.■'.■'.■'.'. 



v 'i'i v v v v •/ '/ v v yy i v y i 



#19% 






■.-.-.■ 

V... 




:>::x>: 



J 



6% 



WEIGHT 




l:;:;;;; 



48% 




18% 



STRUCTURE 



_J 




PROPULSION 



FUEL AND 
ELECT 



PAYLOAD OR 
OTHER COSTS 



Fig. G-3 



Wo 



RANGE CHART 



i .* 


^0 Wq 




— = 1 35— — = 156 
Wl w x 


10, 


. W! Wx 




"**w ■•-—= 0.74— ►— = 0.64 
W \N 


0.8 


^^•k. 






0.6 


^v**0 /s 


04 


1 1 1 1 1 N 



2 3 4 

RANGE (1000 N Ml) 



Fie. G-k 



1-110 



PAYLOAD vs GROSS WEIGHT 



0.40 



0.30 



W 



PAYLOAD 



W 



0.20 



TO 



0.10 - 



_ PASSENGER 



PAYLOAD INCREASE 40 ft 

PERCENT FOR 30 PERCENT 
STRUCTURAL WEIGHT RATIO 
REDUCTION (FOR SAME RANGE)- 




--- $—- 



100 200 300 400 500 600 
TAKEOFF WEIGHT (1000 LB) 



Fig. G-5 






NEW VEHICLE COST 
WITH CHANGE IN STRUCTURAL COST 



ASSUME 30 PERCENT STR WGT DECREASE = 60 PERCENT STRUCTURAL COST INCREASE (COMPOSITE 
STRUCTURE) 

30 PERCENT STR WGT - 40 PERCENT PAYLOAD GAIN 

15.6 PERCENT COST INCREASE 



P 6 * ° 68 -O.J1 = 

L 13 ° J 



ASSUMED VEHICLE COST = 25 MILLION 
1.156 x 25M = 28.9 MILLION = (NEW COST) 



Fig. G-6 



1-111 



RETURN ON INVESTMENT 



$2000/FLIGHT AT 60 PERCENT LOAD FACTOR (SST) 

$2000 x 4 FLTS/DAY x 300 FLTS/YR 

RETURN ON INVESTMENT = 

(45% BREAKEVEN L.F.) 25,000,000 

R.O.I. = 9.6 PERCENT PER YEAR 

NEW AIRCRAFT IS 30 PERCENT LESS STR. WGT. 

25% W 

NEW STR. WGT. FRACTION = = 19.2% W 

1.3 

LET ALL STR. WGT. DIFF. BE ADDED TO PAYLOAD 

OLD PAYLOAD FOR BREAKEVEN = 0.45 x 18% W = 8.1% W 

8.1 + (25% W - 19.2% W) 

RATIO INCREASE IN REVENUE = = 1.72 

8.1 

STRUCTURAL COST HAS GONE UP 60% = 15.6% FOR TOTAL AIRCRAFT 

NEW BREAKEVEN = X 

1.72 

(0.60 L.F. - X) = (0.60 - 0.45) X = 0.377 

1.156 

NEW PROFIT/FLT. AT 60 PERCENT L.F. = S2980/FLT (USING 60% - 37.7%) 

2980 x 4 FLT/DAY x 300 

NEW RETURN ON INVESTMENT = 

(AT 37% BREAKEVEN L.F.) 28,900,000 

NEW R.O.I. = 12.4 PERCENT/YEAR 

PERCENT IMPROVEMENT = 12.4/9.6 - 1 = 30 PERCENT 

Fig. G-7 



MAINTENANCE GAINS 



(1) MAINTENANCE = TWICE ORIGINAL COST OR 25 PERCENT OF D.O.C. 

(2) ASSUME 20 PERCENT OF MAINTENANCE COST FOR ACCESS, INSPECTION, AND REPAIRING STRUCTURE 
AT SCHEDULED INSPECTIONS. 

(3) ASSUME GOAL IS TO DOUBLE INSPECTION INTERVAL (i.e.. 9,200 FLT TO 18,400 FLT) 

20% 

(4) PROGRAM COST REDUCTION = 2 x 2 x 25M = 5 MILLION 

20% 

DOC. COST REDUCTION a x 25% = 2.5% 

2 

Fig. G-8 



1-112 



GAIN IN DIRECT OPERATING COST 

STRUCTURE WEIGHT SAVING AND DOUBLE INSPECTION INTERVAL 

OLD CONFIGURATION 



A U U.U. * — 

NEW CONFIGURATION 










OLD CONFIGURATION 


NEW CONFIGURATION 




/ PRICE = 25M\ 
^MAINT = 50MJ 


( 


PRICE = 28.9M\ 
MAINT = 45M ) 


DEPRECIATION/TRIP 

(12 YR) AND 14,400 TRIPS 


$1730 




$2010 


MAINTENANCE/TRIP 


$3480 




$3120 


FUEL COST/TRIP 
AT $0.01493/N Ml 


$4300 




$4300 



FUEL = 0.48 WGT 
WGT = 600,000 LB 

RANGE x PAYLOAD 



9510 



AD.O.C. = 1 - 9430 = 1-1.41 
1.4 

AD.O.C. = -0.41 



1.0 



$9510 



Fig. G-9 



1.4 
(40 PERCENT BETTER FOR 
30 PERCENT STR WGT 
SAVING) 



$9430 



CRACK LENGTH AS A PERCENTAGE OF FATIGUE LIFE 




N = FATIGUE LIFE IN CYCLES 
N=30,000 



N=45,000 

N=70,000 

N=110,000 

N=180,000 
N=280,00O 

N=550,000 
N=4,000,000 



20 40 60 80 100 

PERCENTAGE OF FATIGUE LIFE CONSUMED, n/N (PERCENT) 

Fig. G-10 



PR71GEN-20b37 



1-113 



A SAFETY GOAL (P T, 10 ,0 ) C SUPERSONIC SPECTRUM E SUPERSONIC SPECTRUM 



B SUBSONIC SPECTRUM 

t 3000 hrs. R 10 000 hrs 



PROBABILITY OF 
STRUCTURAL FAILURE 



K 1500 hrs, R 4000 hrs t, = 3000 hrs, R - 4000 hrs 

D SUBSONIC SPECTRUM F T ( . - 50,000 hrs 



I 3000 hrs R 4000 hrs 



G T c = 30,000 hrs 




10 



20 • 30 40 

FLIGHT HOURS (1000) 



50 



FIGURE 11. PROBABILITY OF STRUCTURAL FAILURE OF FATIGUE 
CRITICAL WING COMPONENT VERSUS FLIGHT HOURS 

Fig. G-ll 



RELIABILITY 



60 



PROBABILITY 
DENSITY 




X OR Y 



RELIABILITY 

pR (X > Y) = PROBABILITY THAT THE STRENGTH X IS GREATER THAN 
THE APPLIED STRESS Y 



R = f G(X) f(X) dX 

2 oo 

x 
G(X) = J g(Y) dY 



IF Y' IS NOT A RANDOM VARIABLE BUT ALWAYS OCCURS (1g STRESS), THEN 
R = pR (X ^ Y) = SHADED AREA UNDER f(X) 



Fig. G-12 



PR71-GEN-2C 



I-llU 



PROBABILITY 

DENSITY 

FUNCTION 

OF LOAD 

OR STRENGTH 







1 


TME = T HOURS TIME = 


I t\ I 










2 / 


I f 










' U 


I 
x l 1 


z I 






O 


1 






< 




H 


1- 1 






O 


a |5 


§1 


< 1 
1- 1 














1/ o 




O 


UJ I 1- 

1 1: 


11 1 / 
1 / 


z I 
* 1 


Q:/ o 

Fl < 

0/ 

47 


I < 

I -J 


UJ 

O 


H PER 
TREN 


1 
< \ 

a 1 
1 




\ | 


O 
4 
O 


= 99T 
DUAL S 


» / 

</> \ / 
OC I / 

2 1/ 


t- I 

< 1 

I I 

h- 1 

\ 
z \ 

UJ \ 

<£ \ 
1- \ 

V) I 

z \ 

< \ 

UJ \ 






1.5 LIMIT 


D / UJ 
-1 / DC 

</ 

/ 


y 

/ \ 
/ \ 






V /' ' 


\ 


S \ 


LARGER OVERLAP AREA 
IN RESIDUAL STRENGTH 


WITH DECREASE 


\ / 






\ 




^^^- EXPECT ED FAILURE PROBABILITY 


-J^ 


^^ 


C_/ FUNCTION OF AREA OF OVERLAP 



LOAD OR STRENGTH ? 



Fig. G-13 



PR71-GEN-20448 



PROBABILITY OF STRUCTURAL FAILURE 

DEPENDS ON: 



^L 



_ 



MEAN STRENGTH IN PSI 



MEAN LOADING IN PSI (ONE g STRESS) 
STANDARD DEVIATION OF LOADING 



MEAN LOADING (ONE g) 

STANDARD DEVIATION OF STRENGTH 
MEAN STRENGTH 

RESIDUAL STRENGTH 



(MEASURE OF SPECTRA RANGE) 



yira 
W TAN 

W 



XT 



c 

w 
a 



FRACTURE TOUGHNESS = VARIES WITH MATERIAL, DIRECTION OF LOADING, CRACK STOPPERS = 
(TYPE OF CONSTRUCTION); SIZE OF CRACK TO PANEL WIDTH, THICKNESS. 

RATIO OF STRESS IN THE REGION OF THE CRACK TIP IN AN UNSTIFFENED PANEL TO THAT IN A 
STIFFENED PANEL 

FINITE PANEL WIDTH CORRECTION FACTOR 

PANEL WIDTH 

HALF CRACK LENGTH 



Fig. G-lU 



1-115 



PROBABILITY OF STRUCTURAL FAILURE 

ALSO DEPENDS ON: 



da 

dN 

da 

cIN 



RATE OF CRACK GROWTH PER CYCLE 
BAK n 



(1-R) K r 



SK 



AK 
B 



A CT 



-R) C v/i 



W TAN 



RANGE OF STRESS INTENSITY FACTOR 



VARIES WITH MATERIAL. TYPE OF CONSTRUCTION, LOCATION OF CRACK TO ATTACHMENTS, 
TYPE OF ATTACHMENTS 

"MIN MINIMUM STRESS 

"MAX maximum stress 

FRACTURE TOUGHNESS 



Fig. G-15 



PROBABILITY OF FAILURE AFTER A CRACK OCCURS 



P (FAILURE) = 



WHERE: 



^s 



erfc 



- 1 





1.5 L c 



ju l 1 - 2.575 



u s 



1/2 



L s = LIMIT STRENGTH 



FACTOR OF SAFETY 



1.5 BETWEEN THE DESIGN LIMIT LOAD 
AND THE 99TH PERCENTILE VALUE OF 
ULTIMATE STRENGTH 



PR71-GEN-20S32 



Fig. G-16 



1-116 



PROBABILITY OF STRUCTURAL FAILURE vs 
STRENGTH, LOAD AND NUMBER OF FLIGHTS 

ALUMINUM WING 



PROBABILITY 

OF 

STRUCTURAL 

FAILURE 



r" 



.j 




2024 T3 

1G - 13.500 PSI 

N FLTS " 33751 
K c » 85.000 
FAILURE AT 
N - 36.251 

1G - 10.000 PSI 
N FLTS = 158,850 

7075 T6 



1G - 


10,000 PSI 


N FLTS 


= 50.100 


1G = 


10,000 PSI 


Vts 


■ 51,350 


1G - 


10.000 PSI 


N FLTS 


= 62.600 



1G 



13,500 PSI 
31.350 



ONE 

RUN " 



ONE 
RUN 



FLTS 
1G - 13.500 PSI 
N FLTS = 50.000 
o L /y L ■ 0.174 
K c VARIES 

CIRCUMFERENTIAL 
CRACK 



2.0 3.0 4.0 5.0 

ft s /H L = MEAN STRENGTH/MEAN LOAD (1G) 



Fig. G-17 



PR71 GEN-20455 



PROBABILITY OF STRUCTURAL FAILURE vs 
STRENGTH, LOAD AND NUMBER OF FLIGHTS 

TITANIUM WING 



1.0 

-4 



10 



10 



-8 



10 12 



PROBABILITY 

OF io 

STRUCTURAL 
FAILURE 10 

10 



10 



10 



16 

20 
-24 

-28 
32 





^«^i 


^ 








1G STRESS = 28.637 PSI 
Vs ' ° 07 


^.i "V 










^ 


\.' ^"N* 3 


10" 10 FAIL/HR 








- 


>v 2 ! ^O 






























^V<0 i %fc 


A' N F 


LTS = 12600 

» 3.06 
• 1250 

LTS = 1385 ° 
■ 17.27 








fo 


A 2 N p 

L , 

IN 
__A 3 N F 






^\ 1 \ 




■ 1250 
Ts " 11.350 








Y 




_iN 


= 1174 
= 1250 



1.0 



2.0 30 4.0 5.0 6.0 7.0 

Ui s /fi l = MEAN STRENGTH /MEAN LOAD (1G) 



PR7I-GEN-20449 



Fig. G-18 



1-117 



PROBABILITY OF STRUCTURAL FAILURE vs 
STRENGTH, LOAD AND NUMBER OF FLIGHTS 



GRAPHITE/EPOXY 



PROBABILITY 

OF 
STRUCTURAL 



FAILURE 10 




3.0 4.0 5.0 6.0 7.0 8.0 

Ms/^l = MEAN STRENGTH/MEAN LOAD (1G) 



Fig. G-19 



PR71GEN-2045; 




SUBSONIC JET TRANSPORT (DOMESTIC ROUTiy 

M >- MACH = 2 SPECTRUM 
„ - MACH - 3 SPECT RUM . 

401 ^~ MATERIAL-TITANIUM 



MEAN STRESS, lg 
(1000 PSD 





ifftE 20 MEAN STRESS VERSUS FATIGUE LIFE 
FOR SUPERSONIC TRANSPORT 



Fie. G-20 



1-118 



RADIAL LOAD DUE TO CABIN PRESSURE 



1MO 



1000 



800 



RADIAL 
LOADING m 

(LB/IN.) 



400 



200 



DC^ DC7 DC* DC-9 
DOUGLAS AIRCRAFT — 



OC10 



Fig. G-21 



FUSELAGE FAIL-SAFE CRITERIA 



• ALL FLIGHT AND PRESSURE 
STRUCTURE FAIL-SAFE 

• TITANIUM CRACK STOPPERS IDC 8 TYPE 



• 2024 T3 USED FOR FUSELAGE SHELL AND 
LOWER WING 

• RECOMMENDED INSPECTIONS BASED ON 
ANALYSIS AND TESTS 

•LANDING GEAR, PYLON AND FLAP WILL 
BREAK AWAY WITHOUT TANK RUPTURE 



DOUGLAS CRITERIA 

(100% LIMIT) 




FAR. CRITERIA 

(80% LIMIT) 



PRM-OC 10-IIIS6O 



Fig. G-22 



1-11° 



AXIAL LOAD 

(INERTIA PLUS PRESSURE) 



HOOP TENSION 




PR«»-aE»44294 



Fig. G-23 



COMPARISON OF MATERIALS FOR DAMAGE TOLERANCE 

7075 T73 VERSUS 2024 T3 












I „. CROSS 

MATERIAL C RESIDUAL 

PSJ IN STRENGTH < 

PSI 

1 

707S-T73 | 100942 29600 

2024T3 -168631 ' 40850 




TRANSVERSE TWO BAY CRACK WITH 
BROKEN CENTRAL LONGERON 



T 



GROSS 

^ RESIDUAL 

PSKTM J STRENGTH, o, 
|M| 



4 __L 

7075T73 65787 

I | 

2024 T3 88090 



20000 

2*800 



-> •» 




L -V 



■t 



*+ 



li I ,'. 



LONGITUDINAL TWO BAY CRACK WITH 
BROKEN CENTRAL CRACK STOPPfR 



Fig. G-2U 



1-120 



EFFECT OF THICKNESS ON STRESS 
INTENSITY FACTOR 



STRESS 

INTENSITY 

FACTOR 




THICKNESS 



Fig. G-25 



PR71-GEN-20535 



RESIDUAL STRENGTH OF UNREINFORCED PANEL 



FAILURE STRESS VERSUS FINAL CRACK LENGTH 



FAILURE STRESS VERSUS INITIAL 
CRACK LENGTH 



STRESS 




CRACK LENGTH 



i , it; 



Fig. G-26 



1-121 



RESIDUAL STRENGTH CHARACTERISTICS 
OF STIFFENED PANEL 

STRINGER CRITICAL 



Ult 



STRESS, S 



F.S. 



FAILURE DUE TO 
STIFFENER FAILURE 
AFTER CRACK ARREST 




2a=2s 
CRACK LENGTH. 2a 

Fig. G-27 



PR71-GEN 20&40 



RESIDUAL STRENGTH CHARACTERISTICS 
OF STIFFENED PANEL 



SKIN CRITICAL 



STRESS , S 



FAILURE OF 
STIFFENER 
HEAVY STRINGER) 




URE DUE TO SHEET 
URE AFTER CRACK 
EST 



• REGION OF STABLE 

CRACK GROWTH 



2a=2s 
CRACK LENGTH, 2a 

Fig. G-28 



PB71-GEN-JOSM 



1-122 



ct 



FUSELAGE PANELS 

(FAIL SAFTY) 



- 










- 






/ / 


/ 


1 


1 


! 


i 





10 20 30 40 
TOTAL CRACK LENGTH 



EXPLOSIVE FAILURE 
Op - *c* Ret 




SLOW PROPAGATION 



TOTAL CRACK LENGTH 



PR69-GEN 24288 



Fig. G-29 



FATIGUE-CRACK PROPAGATION IN A PANEL 
REINFORCED WITH ALUMINUM -ALLOY STRINGERS 
SPACED AT 6 INCHES 



3 10 ~ 

> 



CO c 
T3 T3 



a: 



< 

< 
a. 
o 
en 
Ql 

o 



10" 3 - 



3 10 -e 



50 

- 1 - 



a (mm) 
100 



/!= 0.58 



UNSTIFFENED PANEL- 
(EXPERIMENTALI 




200 




CIRCLES REPRESENT 
ACTUAL DATA POINTS 
FOR STIFFENED PANEL 



STRINGER 
CENTERLINE 



2 3 4 5 6 7 

HALF-CRACK LENGTH, a ( INCHES ) 



-l 



10 



io- 2 3 



-10" 



J 10- 



■O T3 



PR71-GEN-20b3l> 



Fig. G-30 



1-123 



GROSS RESIDUAL STRENGTH CURVES FOR 
CURVED PANEL NO. 15 



GROSS 

RESIDUAL 

STRESS G R 

(KSI) 

(HOOP 

TENSION) 



1 BAY TESTS 1 AND 2 




2 BAY TESTS 3 AND 4 



— O— AXIAL STRESS = 
—CD — AXIAL STRESS 
= 28,495 PSI 



65.114 



\ K c - = 80,573 PSI / IN. 
\ 



16 24 16 

TOTAL CRACK LENGTH (IN.) 

Fig. G-31 




PR69-GEN-2*287 



CRACK 
LENGTH, 
INCHES 
FROMC 



COMPARISON 
OF THE FATIGUE CRACK GROWTH 

RESULTS FOR TEST 2 OF PANEL 2 AND TEST 1 OF PANEL 4 




CYCLES (N x lo- 2 



PR69-OEN-24299 



Fig. G-32 



I-12U 



EFFECT OF OVERLOAD IN 
CONSTANT AMPLITUDE TEST 



CRACK 
LENGTH 

( mm) 



50 

30 
20 

10 




O MOMENT OF APPLICATION 
OF PEAK LOAD I 



20 



50 100 200 

NUMBER OF KILOCYCLES 

Fig. G-33 



500 



PH I I -GEN-20538 



COMPARISON OF CRACK GROWTH RATE 
PARAMETERS WITH TEST RESULTS 



CRACK 

LENGTH 

(INCHES) 



16 

14 

12 

10 

8 

6 

4 

2 



OBEST MATCH K c = 180,000; B 
(UP TO STRINGER FAILURE] 

D K c = 100,000; B, = 0.05 x 10 
(AFTER STRINGER FAILURE R 



CRACK REACHED 
RH IN ATTACH 



CRACK AT EDGE OF RIVETS- 
0.5 x 10" 




T ^-CRACK IN 



2ND 
STRINGER 



OBEST MATCH: K c - 180.000; 

B, = 0.05 x 10~ u 

O K c = 100.000; B, = 0.05 x 10 13 
(AFTER CRACK IN 2ND STRINGER) 



PANEL TEST RESULTS 

A 3 BEST MATCH; K c = 100.000: B = 3.0 x 10"' 3 ; R CT » FUNCTION Lj/W, 
<! K c ■ 85,000; B = 3.5 x 10 ,3 ; R CT = CONSTANT 
Ct 2 K„ ■ 120,000; B - 3.0 x 10~' 3 ; R^ T = CONSTANT 



10 15 20 25 30 

CYCLES (1000) 

Fig. G-3 1 ^ 



35 



40 45 



50 



PR71GEN-?0447 



1-125 



DELAY OF CRACK PROPAGATION AS A FUNCTION 
OF PEAK LOAD MAGNITUDE 



180 
160 



120 



KILOCYCLES TO GROW 
CRACK INDICATED 100 
INCREMENT OF 
LENGTH AFTER 
APPLICATION OF S D1: 80 



BO 



40 



20 



PEAK LOADS APPLIED AT CRACK LENGTH 
2a = 5mm AND 2a ■ 10mm 




n - n = NUMBER OF CYCLES AFTER PEAK LOAD 
TO EXTEND THE CRACK FROM 
2a » 5mm TO 2a = 6mm 


// 


- • n„-n„» NUMBER OF CYCLES AFTER PEAK LOAD 
TO EXTEND THE CRACK FROM 
2a ■ 10mm TO 2a = 12mm 




4> 

10mm TO 12mm v y^r 


S m = 7 KG/mm 2 
S = 3 KG/mm 2 




SHEET THICKNESS 1mm 
PANEL WIDTH 70 mm 

i 


— — — •" ™~ ™~ 5mm TO 6mm — - 1 

i I I i i 



10 



14 15 16 

<^ AK IKG/mm* , 



17 



18 



19 



20 



R71-GEN-20531 



Fig. G-35 




HALF CRACK LENGTH • IK 



CRACK 
LENGTH. 

INCHES S 

FROM $ 



■£• 


O TtST DAT* / 


• 






7 / / " 


71 ALCLAO 707S-T7 


)UM / < 








//C-- 




CRACK ARRESTED / 








^— 




at mm HOUS— ' < 
< 


s 

:s 










1 ° 








: 


V 


PROJECTED CRACK 


f o° 






GROWTH 


CURVE — v 

_ — -^~ 


>' 

















CYCLES IN " 10" : 



4K. KSIViN 





- R 

: ak 


Ml 

Kll • N 




















inn 


































"■AX 


19 4 KSI 




10 














































( 




* 






1 1 



an. 
ksvIn. 




HALF CRACK LENGTH I IN 



PR69-GEN-24333 



Fig. G-36 



1-126 



CONCLUSIONS 



(1) PROBABILITY OF STRUCTURAL FAILURE GOOD TOOL FOR SELECTING MATERIALS AND TYPES OF 
CONSTRUCTION. 

(2) ONE g STRESS MOST IMPORTANT 

(3) SPECTRA VERY IMPORTANT 

(4) STRENGTH VARIATION NEEDS MORE TEST DATA 

(5) FRACTURE TOUGHNESS AND RATE OF GROWTH NEEDS MORE ANALYSIS AND TESTING USING FULL 
STRUCTURAL COMPONENTS. 

(6) THE INSPECTION INTERVAL AND STRUCTURAL WEIGHT SAVING INFLUENCES THE: 

(a) RETURN ON INVESTMENT 

(b) DOC. 

(c) MAINTENANCE COSTS 

(7) GRAPHITE/EPOXY WILL BE A GOOD CHOICE FOR WEIGHT SAVING AND STRUCTURAL INTEGRITY 
(LIFE) IF 1g STRESS LEVEL IS APPROXIMATELY 10,000 PSI. 



Fig. G-37 



PR71-GEN-20444 



AVERAGE AIRFRAME COSTS 
PER POUND OF STRUCTURE 



1970 DOLLARS 



The cost may be described by: 





Cost = 


(M + L) x [increments] 








Where 

M = 


c r/i . i r . Purchased Weiqht 

Raw Material Cost x -= — ,., f — 

Flyaway Weight 






L 


All Direct Labor Includ 


ing Overheads 








[increments] = (1 + R) x 


(1+E) x (1+G) x (1+P) 




R = 


increment for RDT&E 








E - 


increment for ECP's 








G = 


increment for G&A 








P 


increment for Profit 














FACTORED 




RAW MATERIAL 


FLYAWAY WT 




TOTAL COST PER LB 


MATERIAL 


COST S/LB 


PURCHASED WT 


LABOR COST 


200 UNITS 500 UNITS 


Aluminum 


0.66 


0.359 


105 


148 107 


Steel 


0.18 to 1.00 


049 


113 140 


196 166 


Titanium 


11.15 


0.308 


227 


366 309 


GFRP 


5.31 


0.862 


165 


237 171 


Beryllium 


360.00 


0.308* 


656 


1750 1480 



"Assumed Same as Titanium 



Fig. G-38 



1-127 



SPECIFIC STRENGTH vs SPECIFIC MODULUS 



METALS vs COMPOSITES 



40 



30 



SPECIFIC 
STRENGTH, 

Fill/' 



(10 5 IN.) 20 



10 



-ALUMINUM 

-TITANIUM 
-STEEL 




10 



II = MULTILAYER 



n 



ORIENTATION 

EXPECTED 1980 
PRESENT 



^ 




MORGANITE COMPOSITE ill) (TYPE II RAE FIBER) 
BORON COMPOSITE (II) 



7\ >— CELANESE GRAPHITE 
* COMPOSITE (II) 



I 






1 



BERYLLIUM- 



20 30 40 50 

SPECIFIC MODULUS, E/P (10 7 IN.) 

Fig. G-39 



60 



70 



PR69-GEN-22360B 



REDUCTION IN VEHICLE STRUCTURAL WEIGHT 
BY USE OF HIGH MODULUS COMPOSITES 



20 

PERCENT 40 

WEIGHT 

REDUCTION 

60 

80 
100 




1970 



1980 

YEAR 

Fig. G-UO 



1990 



2000 



PR69-ADCS-01967 



1-128 



REDUCTION IN TAKEOFF WEIGHT 

DUE TO 1985 TECHNOLOGICAL IMPROVEMENTS 



100 



RELATIVE 
MAXIMUM 
TAKEOFF 
WEIGHT 
(PERCENT) 



50 



TRANSPORT 



25% 
STR 
WT 



10% 
NON 
STR 
WT 



35% 
ENG 
WT 



20% 
SFC 



37 % 
REDUCTION 



PR69-ADCS-1739C 



Fig. G-Ul 



REDUCTION IN DIRECT OPERATING COST 

DUE TO 1985 TECHNOLOGICAL IMPROVEMENTS 



100 



22% 
REDUCTION 



RELATIVE 

DIRECT 
OPERATING 

COST 
(PERCENT) 



50 



1971 
TRANSPORT 



25% 
STR 
WT 



10% 
NON 
STR 
WT 



10% 



35% 
ENG 
WT 



20% 
SFC 



Fig. G-i+2 



I in •- )\\ II '. I .HIM 



I-129 



TRENDS IN MATERIALS USAGE 

PERCENT 
(FOR PROPULSION) 

COMPOSITES AI/MG T.TAN.UM STEELS SUPERALLOYS 



J47 (1950) 
J79 (1955) 
J93 (1960) 
GE4 (1965) 
TF39 (1965) 



NEXT GENERATION 5-10 



0% 







2, 



22% 
3 
1 
1 
1 



0% 



DEV 
ALTERNATE 



\ ALTERNATE _ 



loEV 

7 

loEV 

12 

I DEV 

32 



OEV 



DEV 




70% 

•1 

85 
24 
15 
18 
15 



_DEV_ 
95 



92 



65 



Fig. G-i+3 



SPECIFIC FUEL CONSUMPTION vs YEAR 



1.0 



0.8 



THRUST 

SPECIFIC 0.6 
FUEL 
CONSUMPTION 
LB/HR/LB 

0.4 



0.2 



% 

DEV 

10 

""foEV 

6,8 

"joEV 

72 

47 
50 



PB7 1OC10 11229 









MACH 


8 CRUISE 
















^SST 
















— - n __ 




THEORETK 


:al limit 










THERMAL EFFICIENCY = 76% 








PROPULSIV 


'E EFFICIEN 


CY = 100% 









1930 1940 1950 1960 1970 1980 1990 2000 

YEAR 



fr* K)0-l*4*7t 



Fig. G-kk 



1-130 



M x (L/D) 



RANGE PARAMETER IMPROVEMENT 




1.5 2.0 

MACH NUMBER 



PR69-ATC-1390 A 



Fig. G-^5 







HT REDUCTION OF STRUCTURAL SYSTEM IT 








GOAL 45% 
REDUCTION 33 TO 60% 



Mi;- 



nuawtiin 



Fig. G-U6 



1-131 



1950 



DEVELOPMENT MILESTONES 



STRUCTURAL AND DYNAMIC ANALYSIS 



1970 

-i-rrV 



1990 

-i- 



I 
I 

1 INTEGRATION OF DIRECT SYNTHESIS METHODS. 

DIGITIZED DRAWINGS, AND NUMERICALLY 
CONTROLLED MANUFACTURING 

1 ADVANCED USE OF DIRECT SYNTHESIS METHODS 

'■ FULLY INTEGRATED ANALYSIS OF LOADS, STRENGTH, AND LIFE 

DYNAMIC AEROELASTIC LOADS VIA INFLUENCE COEFFICIENTS 

■ INITIAL USE OF DIRECT SYNTHESIS METHODS 



ADVANCED COMPUTER ANALYSIS OF STRENGTH 



■PSD ANALYSIS OF GUST RESPONSE 



•INITIAL USE OF ELECTRONIC COMPUTERS (STRENGTH AND FLUTTER ANALYSIS) 

Fig. G-U7 



PR69-GEN 22381 



DEVELOPMENT TRENDS 

STRUCTURAL AND DYNAMIC ANALYSIS 

EMPHASIS 

• AUTOMATION OF PROCEDURES 

• INTEGRATION OF TECHNOLOGIES 

BENEFITS 

• ELEVATED VIEWPOINT FOR DESIGN ENGINEER 

• MORE EFFICIENT USE OF ENGINEERS 

• IMPROVED COMMUNICATION BETWEEN DISCIPLINES 

• INCREASED RELIABILITY, SERVICEABILITY, 

AND EFFICIENCY OF STRUCTURAL SYSTEM 

Fig. G-U8 



1-132 



PLANS FOR FUTURE DEVELOPMENT 



MATERIALS DATA 



EXTERNAL LOADS 




- t ~ ' — ■— i t r r i • ■ ' 

COMPUTER-DRAWN SECTIONS 



TAPE CONTROLLED MACHINE 



PR71-GEN-20459 



Fig. G-^9 



STRUCTURAL DESIGN BY DIRECT SYNTHESIS 

(FUTURE) 



DESIGN 
CRITERIA 



MATERIAL 
ALLOWABLES 



STRUCTURAL 
ENVELOPE 



BASIC 
AERO DATA 




COMPUTERIZED 
DIRECT SYNTHESIS 
(OR AUTOMATED 
REDESIGN) 



OPTIMUM 
STRUCTURE 



Fig. G-50 



1'Nt,', ..I N • ■ IH ■ 



1-133 



PRELIMINARY STRUCTURAL DESIGN 



PAST 

• SEVERAL 
CONFIGURATIONS 

• SEVERAL 
MATERIALS 




> 'PM$l€ ► 




PRESENT 

• MANY 
CONFIGURATIONS 

• SEVERIAL 
MATERIALS 



P 1 



Mm 



*f\' 



3LO-J 




3 & DETAIL 

V DEVELOPMENT 



FUTURE 

• ALL 
CONFIGURATIONS 

• ALL 
MATERIALS 




DETAIL DEVELOPMENT OF 
OPTIMUM CONFIGURATION 

J I I I ' L 



20 



40 60 

PERCENT OF TIME 



80 



100 



PR68-GEN-2D27! 



Fig. G-51 



I-13 1 * 



SOME PRACTICAL ASPECTS OF RISK EVALUATION 



J. W. Ellis 
Worth American Rockwell- -Los Angeles 



Abstract: Risk evaluation in aircraft design is still far from 
being a useful tool. An approach is shown by 
identifying uncertainty factors and impact factors 
for determining a risk exposure index. Continuing 
development is bound to take place in this field- - 
pointing toward an important tool in the design 
decision process. 



I almost feel that I should apologize for introducing such 
an elusive subject as risk evaluation in a meeting encompassing 
so many discussions on good solid, data-substantiated, hardware- 
oriented problems. Why then, is this subject included here? Why 
not leave this tenuous field to a management symposium? It is 
true that, historically, risk has been primarily the province 
of the manager, but the simple truth of the matter is -- the 
manager needs help! 

I am neither a program manager nor risk analyst, but in 
my activity in preliminary design, whenever decision time comes 
around, I feel a strong need for some methodology in the risk 
area. To be of value in the detail decision procedure, this 
methodology must, of necessity, produce quantitative information 
which can be entered into the decision process much the same as 
weight or cost. The goal of applying rational risk analysis 
at the detail design decision level is not yet attainable, but 
recongition of risk as a valid parameter in the decision process 
has been made, and the first steps toward rational analysis 
methods are being taken. 

The Decision Process 

Let us examine what is involved in the structural design 
decision process. I believe everything that goes through the 
mind of the decision-maker will fall into one of the four 
categories shown in Figure H-l. The first one, the cost-weight 
combination, of course, is a pretty large piece of our structural 
world. It is considered here as a combination because I think 
we have learned to deal with cost and weight as a trade. They 
are not treated independently as they once were . Through value 
engineering techniques which enable us to equate weight changes 
with airframe growth, and subsequently with cost, we are able 
to reconcile the weight and cost dichotomy. The second item, 
reliability, is rapidly being quantified through developments 
in fatigue analysis, fracture mechanics, fail-safe design 
techniques and statistical approaches to structural failure 



1-135 



prediction. Thus this factor is approaching the point where it 
can be included in a quantitative way in the decision process. 
Schedules, of course, have always occupied a commanding spot in 
the decision process. We not only want to get our airplanes out 
on time, but once we get locked into a master schedule, we know 
we have to contribute a lot of blood, sweat and tears, and 
frequently money, to change. Design decisions dependent upon 
schedule factors are usually made on a go/no-go basis and thus 
require but little methodology to rationalize once scheduling 
is established. 

The fourth decision factor is risk. It is a different type 
of consideration from the other three and is certainly the most 
elusive. Risk can be considered a quality of the cost/weight, 
reliability, and schedule factors in that it pertains to the 
likelihood of achieving goals in these areas, hence the bracket 
on the chart indicating that risk encompasses the other three. 
Certainly there can be no doubt that in spite of its elusiveness, 
it is essential to consider exposure to loss when we are about 
to make a design decision. Such consideration requires methodology. 
both to assess the risk, and to express it in quantitative terms 
compatible with the other factors under consideration. 

Risk Evaluation 

What is risk evaluation? It has been said that it is a 
sophisticated way of turning chicken and I think there is 
something to that; however, the subtle art of turning chicken 
at just the right time is one of the key requirements of 
remaining solvent in your job, your company, or your program. 
The manager who always takes minimum risk, goes down in flames 
when technology or his competition catches up with him, and the 
manager who disregards risk, follows the same path, but his 
flames have historically proven to be not only bigger but 
substantially hotter. 

Let us examine the basic elements of risk evaluation as 
shown in Figure H-2 . All risks involve both uncertainty and 
impact. There can be no risk without both of these items 
present. If you have a great uncertainty, but no impact as 
consequence of the uncertainty, then there is no risk; or 
conversely, if you have a great impact but you know with 
certainty how things are going to come out, there is no risk 
there either. So we have to look at both elements of this pair 
before we can say we have looked at risk. 

Much of the imput data on the nature of the uncertainty 
will be subjective. Generally it must be learned from opinions, 
and but little will be clearly defined. However, this data must 
be quantified or we cannot work with it from an analytical 
standpoint. Risk impact may have a number of forms, but is 
usually resolved in terms of dollars, schedule impact, or system 
performance impact. 

1-136 



After we have evaluated the uncertainty and the impact, I 
think a professional judgment will always be involved in arriving 
at the final decision, at least in the foreseeable future. 
I don't believe any of the risk people claim that we will have 
an automatic system that circumvents or eliminates the judgment 
factor. The analytical risk work is properly used to place 
this judgment on a sound factual basis. The decision may be 
based on one of several criteria. It may be a decision to 
minimize risk as is shown in the figure, but it may more often 
consist of a balancing of risk versus advantages. 

Stumbling Blocks 

There are a number of stumbling blocks which will always 
render the job of risk analysis difficult. Several of these 
are shown in Figure H-3« Uncertainty assessment most frequently 
is a subjective matter especially when we are dealing with a new 
subject where we have only limited historical data. It is 
subjective because we must gather our working information by 
asking people to express opinions on risk. We often will have 
no other way to obtain information. The uncertainty judgment 
required may be a complex one. Other uncertainties may be 
involved, and it is difficult for the human mind to correlate 
several subjective judgments. 

In the area of impact assessment, there are often multiple 
alternatives rather than a single- valued solution. Impacts may 
also be subject to a domino effect where one event will trigger 
off a number of others requiring the ability to analyze the 
system well enough to know the full extent of the dislocations 
created. 

Possibly the most formidable stumbling block is that of 
arriving at a decision after the evaluating information is 
obtained. Frequently, the evaluation may yield probabilities 
of impacts in cost, schedule, and system performance simultaneously; 
a mixture of apples, oranges, and grapes. A common denominator 
is required to reduce the problem to manageable form. 

Qualitative Risk Evaluation 

Several approaches to risk evaluation exist. Figure E-k 
presents a very simple one. I call it the qualitative approach. 
It is based on an off-the-top-of-the-head probability evaluation, 
quantified slightly by stating whether there is a high, medium 
or low probability of failure . Some recognition of the impact 
is taken by estimating, generally, what kind of trouble a 
failure would bring. The final decision then is made by 
invoking the powers of a "Big Daddy" in the organization. His 
decision may be arbitrary, but it is very often unassailable. 



1-137 



Before we sneer too much at this type of approach, I think we 
should recognize that it has two things going for it; It is 
head and shoulders above no risk analysis at all, and it is the 
most commonly used method of all. 

Risk Index Approach 

A significant advancement over the approach just described 
is in use in some segments of the industry today as a comparative 
method of risk analysis. This is known as the risk index method, 
and is diagramed in Figure H-5. Here we introduce a systematic 
quantification of the risk factor. First we identify the risk 
areas and then break them down into uncertainty factors and 
impact factors . Next we identify the uncertainty factors 
involved, observe their relationships, and segregate them. 
This simplification makes it easier to think about the problem 
as we gather subjective data from our personnel. We don't try 
to treat more than one factor, or one coupling of factors at a 
time. Rather than thinking about the whole problem, we look at 
one piece at a time and assign probability indices to each piece. 
These are usually expressed as percentage figures because we 
are used to thinking in percentages. 

Similarly, we look at the impacts, which may be in terms 
of cost, schedule, or performance and assign an index to them. 
Frequently this index can be reduced to the form of dollars . 
Finally a risk exposure index is derived by multiplying the 
uncertainty, or probability of failure index, by the impact 
index. Because of the somewhat arbitrary nature of the index 
factors, this method is most commonly used to compare the risk 
exposures of alternate solutions to a design problem. When used 
in this manner, this technique can be remarkably effective at 
the detail design level. Although useful as a means of comparison, 
this method is more of a straightening-out-of-thinking than it 
is real methodology. 

Figures H-6 and H-7 show an example for a real design 
problem. In this instance, consideration is being given to a 
design change to composite construction for an aircraft component, 
at a good saving in weight. Risks are involved, both in the 
basic material properties to be used in design, and in the 
validity of composite component analysis methods. In the first 
case, any error would be detected early in the program through 
element tests , while errors in analysis would become apparent 
only late in the program in component testing. Assigning 
subjective estimates of probability of failure and applying the 
risk index analysis results in a risk exposure index of $55jOOO 
as shown in Figure H-7. 



1-138 



The index may be interpreted in several ways . It may be 
compared with similar analyses of alternate approaches to 
achieving the same goals , or it may be treated in a probabilistic 
manner by declaring that any means of reducing the risk to zero, 
such as early tests, etc., which can be accomplished for 
$55,000 or less, is a good action to take. Another evaluation 
may be made by comparing the risk exposure index to the dollar 
value of the weight saved as determined through value engineering 
analysis. 

This is about the present limit to the application of risk 
analysis at the detail design decision level. The method has 
several basic deficiencies which limit its value. It does not 
consider a range of probabilities of failure versus various 
degrees of impact, only discrete points on the curve. It also 
cannot deal with the interaction of a number of probabilities 
and consequences. In addition there is no provision for converting 
the direct consequences of a failure to meet a goal into a 
relevant impact upon the system. Finally, there is no attempt to 
provide a valid comparison between the risk exposure and the 
advantages to be gained. It appears that a more sophisticated 
methodology is needed. 

Rational Probability Risk Analysis 

There has been some work done in the associated field of 
contract management and incentive fee allocation, employing a 
rational probability approach which may answer some of the 
objections stated above. In Figure H-8, I have diagrammed a 
hypothetical approach to risk evaluation using some of the 
elements of rational probability analysis and, in addition, 
answering some of the other objections stated above. Here we 
take the proposed concept and identify the risk areas as before. 
Then we set up a probability network which relates the various • 
facets of uncertainty involved in the problem. We relate them 
and assign values through a procedure similar to the one shown 
for the risk index analysis, except that, instead of discrete 
values for probabilities and consequences we describe them through 
probability curves relating probability of occurence with 
consequences over a range of values. Through the probability 
network analysis we arrive at an output which is a resultant 
probability versus consequences, again in the form of curve 
data rather than discrete points. However, we are probably not 
looking for the consequences as such, but rather for the response 
of the system to these consequences. So we do a sensitivity 
analysis to see what significance weight, delay, or any other 
result really has to the system. From this we derive a relevant 
impact, usually in the form of a cost, schedule, or performance 
factors. We would like to reduce this to a common denominator 
if possible, usually in terms of dollars. 



1-139 



We can next take the probability- impact data, put them into 
the form of a probability density function, and multiply probability 
times the impact as we did before for the risk index analysis. 
This will give us a peaked curve from which our maximum risk 
exposure is determined. 

The next step is to compare this risk exposure with the 
concept advantage. In order to reduce the advantages to a 
comparable base, they are defined in terms of system advantages 
through sensitivity analysis, and then reduced to a dollar basis, 
much as the impacts were treated. When all of this is accomplished, 
we are then able to compare the risk and advantage on a dollars 
versus dollars basis . 

The preceding example of an advanced risk analysis procedure 
was highly hypothetical, presented primarily to illustrate some 
of the problems which must be overcome in future development. 

Problem Areas 

One of the basic problem areas in risk analysis mentioned 
previously is that of subjective data collection. Figure H-9 
illustrates the form of the probability data which is sought 
from experts in the technical areas. On the left side is shown a 
cumulative distribution curve. It is fairly easy for people 
to think in terms of what is least and most, or best and worst. 
This type of information is obtained first and forms the first 
points on the curves. The subjects must then be led to estimate 
at some other level, estimating 80$ or 20$ probability for 
instance; there are methods of interrogation which are being 
developed in the behavioral sciences, involving such things as 
choice of gambles and bets, to try to coax this kind of information 
out of people . 

In addition to giving us these probability estimates, 
technical groups also have to tell us what the penalties for 
failure are. In this case, if we give them the cost, schedule, 
and performance variations, they can tell us what this means to 
the contract in dollars or to the reduction in system cost 
effectiveness. This data is generally derived analytically and 
is not subjective in nature. 

The output of the network probability analysis can also be 
represented by a cumulative distribution curve, similar to the 
inputs in form. These distribution curves, as shown in Figure 
H-9s can tell us a number of things in themselves. For instance, 
the left-hand figure tells us that, considering all facts, all 
impacts, we have a 75$ chance of meeting a certain performance 
level. Now, by applying the sensitivity conversion parameters 
for the system we can produce the right-hand figure. This tells 
us, for instance, that we have a 90$ probability that the loss 
will not be greater than a certain amount. 



I-lU0 



Some other non-engineering problem areas are outlined in 
Figure H-10. The determination of overall total impact is frequently 
difficult because of factors not amenable to engineering analysis. 
The company reputation might suffer damage beyond the actual 
accountable dollars and cents if a project fails or is delayed 
by the failure of a risk item. Delays may cause damaging 
interference with other projects and failures or schedule 
slippage could conceivably jeopardize other contracts. 

Most of our contracts have some kind of fee or penalty 
feature which can be included in the risk analysis. However, 
if we are striving to reach a wise and prudent decision through 
our risk evaluation, we may find that the contractural 
stipulations bias the decision unduly. Then there is the 
matter of self interest. Here again, the solution should be 
worked out on a mutual -basis with the customer. Unfortunately, 
very often the advantages you are trying to achieve accrue 
mainly to the customer and the penalties are all yours. 

Management acceptance can be x nearly insurmountable obstacle 
in some organizations. Among managers there seem to be two 
camps . There are those who agree that they most certainly need 
help in making decision involving risk. They recognize the value 
of quantitative help from the engineers who are most familiar 
with the system. Those in the other camp oppose any infringement 
upon their traditional ground by technical personnel. The 
solution undoubtedly will lie in development of viable risk 
evaluation methodology and the education of management in its 
use. 

In conclusion, I believe it is safe to say that in spite of 
the visible problems, we will assuredly see continued development 
in the methodology of risk evaluation and that it will eventually 
alter the nature of the process by which we make our structural 
decisions. If properly applied, it should lead us around a lot 
of the blunders that we have perpetrated in the past in the 
application of new materials and constructions. 






















I-lUl 



DESIGN DECISION 



1. COST - WEIGHT 



2. RELIABILITY 



3. SCHEDULE 



4. RISK 



UNCERTAINTY 

•SUBJECTIVE 
•QUANTITATIVE. 



Fig. H-l 



RISK EVALUATION 



IMPACT 

•$ 

• SCHED 

•PERF 



PROFESSIONAL 
JUDGEMENT 



RISK REDUCTION 
DECISION 



TSP71-0484A 



Fig. H-2 



NA-71-69! 2 



I-1U2 



STUMBLING BLOCKS 



f' 



•UNCERTAINTY ASSESSMENT 

- SUBJECTIVE 

- LIMITED HISTORICAL DATA 

- COMPLEXITY 

• IMPACT ASSESSMENT - 

-MULTIPLE ALTERNATIVES 

- DOMINO EFFECT 

- SYSTEM SENSITIVITY 

• DECISION 

- APPLES, ORANGES, AND GRAPES 

Fig. H-3 



QUALITATIVE APPROACH 






TOP-OF-HEAD PROBABILITY EVALUATION - HIGH, LOW, MEDIUM 



•SOME RECOGNITION OF IMPACT 






•ARBITRARY DECISION BY 'BIG DADDY' 



MOST COMMONLY USED 

Fig. R-h 
1-11+3 



RISK INDEX APPROACH 




IDENTIFY 

UNCERTAINTY 

FACTORS 



IDENTIFY 

IMPACT 

FACTORS 



ASSIGN 

PROBABILITY 

INDEX 



ASSIGN 
IMPACT 
INDEX 




Fxg. H-? 



RISK INDEX ANALYSIS 



QUESTION: SHOULD THE VERTICAL STABILIZER BE CHANGED TO 
GRAPHITE COMPOSITE CONSTRUCTION. 

ADVANTAGE: GOOD WEIGHT SAVING AT REASONABLE FABRICATION 
COST 

RISK AREAS: 1. BASIC MATERIAL CHARACTERISTICS 



UNCERTAINTY: 



COMPRESSION AND SHEAR 
ALLOWABLES USED IN DESIGN 
MAY NOT BE VALID 



IMPACT: REDESIGN OF ASSEMBLY AFTER 
ELEMENT TESTS 

COMPONENT PERFORMANCE 

UNCERTAINTY: VALIDITY OF GENERAL 
STABILITY ANALYSIS 

IMPACT: REDESIGN, TOOLING SCRAPPAGE, 
AND SCHEDULE SLIPPAGE AFTER 
COMPONENT TESTS 



Fig. H-6 

1-lUU 



RISK INDEX ANALYSIS 



RISK AREA 



PROBABILITY 



(P 



IMPACT 



RISK EXPOSURE 
R = P f X I 



MATERIAL 

PROPERTIES 

INADEQUATE 



0.30 



REDESIGN $ 50.000 



$15,000 



COMPONENT 

PERFORMANCE 

INADEQUATE 



0.10 



REDESIGN 100,000 

TOOLING 225,000 
SCHED SLIP 

(O'TIME) 75,000 



$400, 000 



40,000 



TOTAL RISK 
EXPOSURE 



$55,000 



TSP71-S6S7 



Fig. H-7 



NA-7I-69I 7 



IDENTIFY 

RISK 
AREAS 



RATIONAL PROBABILITY APPROACH 



PROPOSED 
CONCEPT 



T 



PROBABILITY 
NETWORK 



ASSIGN 
VALUES 



OUTPUT: 



PROB VS 
CONSEQUENCES 



X 



PROBABILITY 
ANALYSIS 



SENSITIVITY 
ANALYSIS 



IMPACT: 



COST, SCHEDULE 
OR PERFORM. 



REDUCE 

IMPACTS TO 

COMMON BASE ($) 



RISK 

EXPOSURE 

R = P X M$) 



I MP ACT _ 
ANALYSIS 



CONCEPT 
ADVANTAGE 



SENSITIVITY 
ANALYSIS 



ADVANTAGE 




REDUCE TO 


TO 




COMMON 


SYSTEM 




BASE ($) 




TSP71 S6S8 



ADVANTAGE 
ANALYSIS" 

na- / l -691 8 



Fig. H-8 



I-1U5 



PROBABILITY ANALYSIS 




757o PROBABILITY 
GOAL WILL BE MET 



FINAL PRESENTATION 

- S/C EFFECTIVENESS 

- COST TO PROGRAM 

- SCHEDULE 



TASK 1-10 



AT LEAST (OR 
MORE THAN 
REQUIREMENT) 

CONTRACTUAL 
REQUIREMENTS 

AT MOST (OR 

LESS THAN 
REQUIREMENT) 



COST 

SCHEDULE 

PERFORMANCE 



TSP70-4930 




NA-71-691 



Fig. H-9 



PROBLEM AREAS 



•SUBJECTIVE NATURE OF DATA 

•DETERMINATION OF TOTAL IMPACT 
EXAMPLES: 

- COMPANY REPUTATION 

- INTERFERENCE WITH OTHER ACTIVITIES 

- RELATED CONTRACTS 

•CONFLICTING CONTRACTUAL FEATURES 
• SELF INTERESTS 

•MANAGEMENT ACCEPTANCE 

Fig. H-10 



I-1U6 



ADVANCED METALLIC STRUCTURES 



D. A. Shinn 
Air Force Materials Laboratory- 



Abstract: The Air Force program on Advanced Metallic Structures 
has the objective to increase structural efficiency, 
integrity and reliability by sponsoring a systematic 
approach to critical structural problems . The 
technical requirements of this program are discussed-- 
emphasizing the dual aspects of applying technology 
and disseminating the resulting information to the 
technical community. 



During the preceding presentations we have been listening 
to many aspects of design problems. I would now like to give 
you the outline of an approach toward their solution. Precisely 
how these problems can be solved is, of course, up to the air- 
craft industry. Our approach is to contract to the best offer to 
do the job. 

An Advanced Development Program has been initiated by the 
Air Force Systems Command in the field of Advanced Metallic 
Structures. Its objective is to demonstrate more reliable and 
improved structural design methods, materials, fabrication 
processes, and evaluation techniques. The underlying reasoning 
can be found in some general trends which have become clearly 
visible. 

One consideration is that developments in structural 
design have in the past been spearheaded by challenging tasks. 
Traditionally, technological challenges have been posed by the 
development of numerous new fighter and attack aircraft which 
provided a proving ground for the designer's ingenuity. In 
the early 1950 ' s, an average of three new projects for fighter 
or attack aircraft were started each year. During the last 
decade and a half, however, this average has been less than one 
in three years, i.e. a shrinkage by an order of magnitude, 
permitting much less opportunity for new and varied structural 
developments . 

Another consideration is that confidence in new structural 
designs can be established only by structural testing. Here 
the lessons are learned for future developments. The record 
shows that for more than 70 systems, tested over a long time 
period, more than half had two or more major components failing 
below design load. The percentage has not improved but actually 
worsened for recent systems , indicating perhaps more rigorous 



I-1U7 



test conditions or, more likely, a basic need to improve 
analytical methods . A very small flaw can have drastic effects 
and cause the loss of an aircraft. We need to improve structural 
efficiency and to avoid structural failures but the lack of 
opportunity to demonstrate new structures has hindered technological 
developments . 

We can see this in its proper context if we realize how 
the development of structures depends on basic research which 
feeds into exploratory development, then into advanced development, 
engineering development, and finally into manufacturing 
technology. These all interact with each other. They all 
provide increased confidence in design, manufacturing, and 
integration. Our current problem arises from the fact that 
nothing has been done in the advanced development area of 
structures. Every one has been preoccupied with their own 
specialties, with no forcing function such as an ADP to make 
them communicate and learn from each other. 

The preceding discussion is an elaboration of the objective 
in the Advanced Development Program. Its purpose, as stated 
before, is fairly straightforward. The approach is actually 
to design and fabricate structures with best available technology 
and to test them as a check on analytical methods. Feedback to 
improve both analysis and test will be emphasized. Another part 
of the approach is to remove the constraints which are normally 
associated with systems development (in production) where time 
schedule and budget considerations usually militate against a 
systematic test and development program. A number of critical 
design problems have been set forth for investigation during the 
entire lifetime of the Advanced Development Program. The first 
of them uses a bomber aircraft wing carry-through structure as 
a baseline . 

Now for the philosophy on this ADP. We had to set forth 
some guidelines as to what we want done and so we picked eight 
technical areas which have to be carried all the way through 
the whole program: 

Fracture mechanics will be emphasized at all times 
across the board; 

Structural materials shall be exploited and evaluated 
with special consideration for heat treatment and 
improved processes using titanium, aluminum, and steel 
as basic candidate materials but not excluding other 
metals or certain mixtures of metals and composites; 

Design criteria will be somewhat more severe than in 
the baseline structure; 



1-11+8 



Structural design concepts shall employ safe life and 
damage tolerance considerations applied to primary 
structure which is vital to vehicle integrity; 

Structural analysis methods shall incorporate at 
least finite element and numerical solution methods 
for analysis and optimization of structural config- 
urations, including fracture and damage -tolerant 
analysis; 

Manufacturing methods shall include any new basic 
and secondary methods which are believed to meet the 
objectives of this program; 

Non-destructive inspection shall be evaluated and 
exploited, insuring inspectability within predetermined 
confidence limits; 

Information transfer, last but not least, is particularly 
emphasized at all times for the rapid education of the 
technical community and not just the contractor. 

These eight technical areas have to be considered throughout 
the four phases of the program: preliminary design, detail 
design, manufacturing, and full-scale test and evaluation. The 
idea of the program is to be as diverse as possible in the approach 
to the problem. The requirements call for starting out in the 
preliminary design phase with six concepts which will be weeded 
down until finally one is selected for manufacturing. Originally 
we envisioned that this cycle would be done once and, if it 
were so indicated, it would be done over again as an iteration. 
It remains to be seen whether iteration can be carried out or 
whether the costs will be too high. 

In conclusion, I would like to describe why we feel that the 
ADP is the way to go rather than to continue such effort through 
normal systems development. If you have a problem in a system, 
you are limited in your materials selection, your criteria are 
fixed for you, you have to use proven design concepts and 
conventional manufacturing methods, inspectability is a secondary 
objective, and testing and reporting are kept to a minimum due to 
time and money considerations. 

Fabricability, inspectability, reliability, all these -ilities 
represent the thingc about which the presentations by Dietz and 
Shuler were concerned. In the ADP, we hope to improve all these 
-ilities up to the next higher level. We will have many more 
choices for materials. We can use innovative criteria. We can 
use the best possible design concepts. Inspectability will be a 
prime objective. Fabrication can consist of variable types, 
whatever we happen to need. Testing and reporting will be optmized, 



I-IU9 



With this type of a systematic approach, we should be able to 
advance the state of the art in structural design and the 
utilization of structural materials. 



1-150 



SUMMARIZING REMARKS 



CDR F. L. Cundari 
Naval Air Systems Command 



It is not easy to summarize these two days of conference. 
I was particularly impressed by the spirit of free discussion of 
the many technical problems existing in Aircraft Design. Most 
importantly, I have learned that there is a considerable amount 
of work to do in this area. 

In this short time, two days, we did not resolve the 
fundamental design problems. We did not have the time to study 
in depth all of the basic issues. However, we did expose most 
of the problems . It is important that we capitalize on all the 
issues which We did discuss . In order to provide a reasonable 
avenue for the conclusive discussion of the salient points, I 
am recommending that a short follow-up questionnaire be mailed 
to each participant. Hopefully you will respond with at least 
the same clarity with which you have been able to express your 
points so far. Each participant would have to address in more 
depth the principal issues which have concerned us here. 

What are the basic issues? What are the fundamental 
problems? How can we correct them? 

The methods for basic static structural design are common 
place. The procedures have been utilized for years. However, 
we have learned that these techniques are not sufficient. 
Structural design procedures must be improved by incorporating 
fracture and fatigue analyses tools into preliminary design. 

A second major area highlighted at this session is the 
contractual procedures .of the United States Government. These 
rules or regulations cannot be changed overnight and are not a 
direct issue of our conference. We must, however, realize the 
impact of the contract regulations on the design technology. We 
are not going to change government contracting overnight nor are 
we going to change the industrial response to these procedures 
any faster. We will always find it difficult to conduct the 
requisite engineering design, and evaluation of aerospace vehicles 
when the government underestimates the task or top management 
underbids a job. 

Let me quickly summarize those points which have been discussed 
that have an important impact on the technical details of 
structural design. 



1-151 



1. Fatigue and Fracture 

There is a lack of knowledge concerning the basic physical 
fundamentals of fatigue and fracture . It appears that 
the aerospace industry is utilizing some engineering 
approaches to the application of these disciplines 
in design. Most likely these procedures are post 
design reviews and often introduce costly modifications 
or redesign rather than original criteria. So the 
fundamental physical knowledge must be obtained and then 
it must be converted into sound design principles. 

2. Stress Corrosion 

It is a difficult task to introduce the current basic 
knowledge in this field into a methodological design 
procedure. Presently it seems to be either a marvelous 
skill employed by aerospace designers which sometimes 
approaches a level with pure mystic overtones - black 
magic. Often this important technical detail of possible 
stress corrosion is not considered until a failure 
occurs . This aspect of structural/material failure 
must be considered when selecting materials and 
configurations in preliminary design as well as the 
whole iterative design procedure, 

3. Hydrogen Embrittlement & Fretting Fatigue 

Closely coupled with the first two problems are the 
aspects of hydrogen embrittlement and fretting fatigue. 
Both areas have not been fully characterized nor 
quantified to allow the development of design procedures. 

U. Structural Safety Factor 

Reduction of the Safety Factor has been proposed in a 
few discussions. It would be extremely difficult to 
lower the factor of safety currently used in the 
military. Only if a sufficient population of data 
concerning the component or part in question is 
available, will the service organizations be able to 
accept a lower safety factor. 

5. Material Characteristics 

A more comprehensive data bank of material charactersitics 
must be provided. This data must be in a form which is 
easily interpreted and used by the design engineers. This 
information must include variations with temperature, 
both equilibrium and high gradient states; the effects of 



1-152 



the environment including short and. long terra exposures; 
and of course, fracture and fatigue characteristics, as 
well as the fundamental modules and strength data. 

This aspect of material characteristics is even more 
important for the new composite materials, which may- 
have many different reinforcing fiber orientations, 
matrix materials, fiber types, and lamina thickness. 

6. Materials Processing Control 

The control of the material manufacturing process has 
become an extremely important aspect of air vehicle 
construction. Manufacturing techniques and quality 
control capabilities must become a part of the basic 
design philosophy. The heat treatment and aging 
processes must be fully characterized to reveal the 
resulting effect on the material structural properties 
and the effects on the fabrication quality and cost. 

7. Service/Operational Data 

There is an important need to obtain more sophisticated 
and accurate operational usage data. This needs to be 
coupled with "lifetime remaining" type testing of 
critical vehicle components. Sufficient engineering 
knowledge must be available to develop the cumulative 
damage analysis tools for aircraft and missile design. 

It may be possible that the new emphasis being given to 
vehicle prototypes will allow new, untried structural 
innovations to be adequately demonstrated. This 
approach could permit a better evaluation of alternative 
candidate design of some major components. 

8. Cost Benefit Analysis 

The idea of incorporating the elements of operations 
analysis or risk analysis into the initial vehicle 
candidate designs is a step in the right direction. 
However, the objectives of a national defense force 
are not usually oriented towards pure dollar loss risks 
or continual profit margin minimums . The goals are 
oriented towards weapons performance superiority on 
a competitive basis. Therefore, we should be developing 
performance benefit analytical tools that evaluate the 
potential military combat benefit available for the cost 
of the candidate designs. This is more of a cost- 
benefit rather than a loss-risk type of analysis. It 
appears as though we should attempt to apply some of 
the techniques that have been developed in other fields 
and incorporate them in the design phases of the basic 
airframe . 



1-153 



There are many other important considerations in the design 
of air vehicles. I have attempted to list only the major items 
that have "be en discussed at these sessions. 

The objective of this conference was: "to recognize those 
fundamental design problems which all aerospace companies have in 
common, to clarify these aspects, and to indicate priorities and 
guidelines for a methodological development". I feel confident 
that we have discussed the problem areas. However, I believe we 
have yet to determine all of the approaches for the development 
of a methodological design process. I hope that our discussions 
will he continued until this comprehensive airframe design process 
is established. 

I wish to extend, again, the thanks of myself and everyone 
associated with the Navy and the Postgraduate School to all 
participants. Your expert knowledge and your dedication for 
improving our engineering methods have been exemplary. I know 
that you will all he willing to participate in the development 
of this design methodology. 

I particularly want to extend my warmest appreciation to 
Professor Ulrich Haupt for organizing and conducting this 
conference. His extraordinary interest in airframe structural 
design has kept the direction of the meeting pointed towards 
the important goals of this conference and he has been a guiding 
influence to all of us . 

Thank you. 



I-15U 



SECTION II 



SOME UNSOLVED QUESTIONS AND VARIOUS COMMENTS 



SECTION II 
SOME UNSOLVED QUESTIONS AND VARIOUS COMMENTS 



Discussions during the symposium began along the lines of 
the designer's problems as indicated by W. C. Dietz and shown in 
Figures C-39 an d C-UO. Some of the comments were made during 
the symposium and some more thorough comments continued coming 
in afterwards . A few of them came from engineers in industry 
who did not participate in the symposium but were interested 
in the subjects . 

The questions which were posed are of the kind for which 
there is no ready-made answer. The comments -- no matter how 
contradictory they sometimes are -- may help to stimulate thinking 
about some of those problems which can be anticipated but are not 
yet clearly visible. After questions have been formulated and a 
discussion has been started, it becomes easier to clarify one's 
own thought. Questions and comments in this section should be 
considered from such a perspective as somewhat of an opening bid. 
No claim is made that they represent a complete and thorough 
probe . 



II-l 



la. Is a concerted effort made to establish a quantitative 
plateau of acceptability for risks as a guideline for the 
engineer? 



Very limited. Some in establishing gust criteria, some 
in repeated load and allowable stress areas . 



An effort is made but in qualitative terms rather than 
quantitative. I, among others, have always felt that 
each structural comnonent should be assigned a level of 
risk based upon the consequences of failure and/or the 
probability of experiencing the critical conditions. 
Thus , we would have a variable margin of safety through- 
out the airplane. This has never been done to my 
knowledge -- though it has been discussed. 



No. Existing specifications talk about fatigue life, 
fail-safe and safe-life, s 3 structural integrity but not 
risk . Risk is a bad word that implies failure. Airlines 
specify a fatigue life for design of primary structure 
sufficient to "guarantee" a "crack-free" structure. The 
contract specifies how many flight hours and/or number of 
landings are required. If the structure "cracks" before 
these values the manufacturer must repair it at his 
own expense. The FAR regulation requires a fail-safe 
strength of 80 percent limit load as the ultimate fail-safe 
strength. This implies that the residual strength can drop 
from a positive margin of safety of greater than 1.5 
limit load initially to a minimum of 80 percent limit load 
at the end cf a structural maintenance interval and is still 
acceptable. Of course the manufacturer tries to provide 
an inspection interval small enough to discover cracks 
before the strength drops to the 80 percent limit level. 
Also, the manufacturer tries to provide crack stoppers and 
a low enough 1 g stress level so that the cracks grow at a 
slow rate rather than a fast crack growth rate . All these 
precautions help to reduce the failure probability. 
However, any loss in residual strength increases the 
possibility of an occasional load cf greater than 80 
percent limit load to 1.5 limit lead exceeding the residual 
strength. No failure probability is specified although, 
as was shown in Mr. Fischler's slides, a loss in residual 
strength means an increase in the probability of catastropic 
failure . 



II-2 



Although some attempts have been made to specify the probability 
of failure, (i.e., Air Force "Rational Probability of Failure 
for the Fleet" and Lundberg's 10 -10 Structural failure rate 
per hour), a "quantitative plateau of acceptability" must 
be established which allows "visibility" without excessive 
structural weight increases. "Visibility" can be achieved 
by comparing the failure probability of successful past 
designs to advanced future designs and trying to include 
the important variables in such a way to yield a lower risk 
with new materials and types of construction. 



* 



No. It appears that more static and fatigue testing of 
elements, components and entire vehicles coupled with a 
better assessment of service life failures will be required 
before we can define "acceptably low risk" quantitatively 
for the designer. 



No — a risk must be related to factors over which the engineer 
has some feel for and control, such as structural weight, 
cost, life, etc. Methods of risk analysis haven't been 
developed to the point where this can be done in the time 
span that most technical decisions must be made. 

#■ * # 



lb. What distinction should be made between commercial and 
military fields? 



A very considerable distinction should be made if one can 
accept the idea objectively that a big difference exists 
between the consequences of failure on a 250 passenger 
airliner and the consequences on a one or two man military 
aircraft equipped with in-flight escape devices . This 
means that design for military aircraft can and should be 
based on probabilistic considerations of loads and allowables 
which accepts the possibility of x failures in y hours of 
operation. 

* * •* 

Commercial vehicles should have a lower failure probability 
goal than now exists. The air transport of civilians risk 
should be comparable and less than other means of transport 
(i.e., car, bus, train or boat). Also, with the increase 
in the use of aircraft, (growing at an average rate of 15 
percent per year), failures must be fewer to physically 
maintain the same quantity of failures per year. In 
addition, with the advent of jumbo jets that carry twice 



H-3 



the amount of passengers as the DCS's and 707 's, a failure 
which results in a total passenger loss, will be so large 
a disaster that it is unacceptable even if less than 50 
percent as frequent. A purely economic consideration is 
the large financial loss from the law suits that result 
and the high cost loss per vehicle. 

The high failure rate of recent military vehicles has 
resulted in the loss of competent military personnel that 
should not be measured, in dollar value . The required 
additional development costs and repairs to correct their 
deficiencies after they occur has been unacceptable. 
Although less publicity is given to military failures, the 
loss has reduced our military strength to an unacceptable 
level. 

Therefore, the failure risk of both commercial and military 
vehicles must be reduced substantially. 

Since commercial vehicles have a desired life of approximately 
20 years with a higher yearly utilization than military 
vehicles with a desired life of approximately only ten 
years, then commercial vehicles must have a lower failure 
probability to account for these factors. Also, if the 
loss of a human being, whether a civilian or military, is 
at the same level, then the risk exposure per passenger 
mile should be the same. Since military vehicles normally 
carry a much smaller human paylcad than commercial, over 
an average shorter range, (since fighters are short range), 
these factors should also be considered. Military vehicles 
therefore could be designed for a higher risk of failure 
structurally if the acceptable criteria is to match the 
failures structurally per commercial passenger mile. 

Military vehicles, in general, are designed for higher 
load factors than commercial vehicles (i.e. a fighter is 
designed for 7«33 limit load factor and a commercial 
transport for approximately 2.5g). The load factor occurrences 
per mile of a fighter however are higher than the commercial 
vehicle. A higher design load factor for a fighter results 
in a lower 1 g stress level. Therefore, the rate of crack 
growth per hour of flight for a fighter is usually lower 
than for a commercial transport. Therefore, the loss in 
residual strength for a fighter, in spite of the higher 
frequency of loads, is less than for a commercial transport 
per hour of flight. These factors also must be considered 
when comparing the expected failure rate. To overcome the 
disadvantage of such a high design load factor for a 
fighter, high strength steels or titanium are frequently 
used. A slight reduction in strength allowables, (from 
260,000 p.s.i. to 200,000 p.s.i.) for steel would increase 



II-U 



the fighter life substantially. In conclusion, design of 
the transport or the military vehicle is very dependent 
on the desired life and the environment expected. A 
proper design, considering fracture mechanics. could result 
in an efficient design for either aircraft with an equivalent 
probability of failure (same failure rate per passenger 
mile ) . 

-x- * -* 

The goal for both military and commercial aircraft should 
be to obtain a low level of risk. However, the trade-offs 
for military and commercial aircraft are usually different. 
For commercial aircraft risk must be weighed against cost, 
whereas for military aircraft risk is weighed against 
performance primarily. Therefore, more costly solutions 
to design problems become necessary for a military 
aircraft as compared to a commercial aircraft. 



. . . Military should be willing to accept risks to help 
assure early availability of high performance systems. 

-* * -* 



lc. What is the trade-off between increased probability of 
failure and increased performance for military aircraft? 



None. Higher probability of failure should not be accepted 
to increase performance except in aircraft deliberately 
designed for short life or to meet national emergencies. 



. . . Many failures are a function of the poor selection of 
materials for structures, controls, propulsion, and poor 
inspection procedures. All these areas' reliability must 
be improved plus the use of new materials that are more 
damage tolerant (graphite epoxy at low stress levels), 
before the risk of designing for higher stress levels to 
reduce the structural weight and increase the probability 
of structural failure to get increased performance can be 
considered. A systematic failure probability analysis 
which has visibility and is accepted as a standard for 
comparison to get the failure rate lower must first be 
accepted and tested by actual flight performance before 
considering the trade off of parameters . 



H-5 



Obviously, increasing the probability of failure should 
result in a lighter, less expensive aircraft, but as the 
probability of failure increases, the overall cost of owner- 
ship of x aircraft in service will increase. 

#■#•■* 

Quantification not really possible. The military risks 
unknowns of new environments and new materials. Best 
thing would be to provide for tolerance of difficulties 
regardless of source as opposed to acceptance of unknown 
risks . 



Id. Should different levels for probability of failure be 
specified for different types of failure -- partly for 
psychological reasons? 



Yes, but first define "failure consequences." The probabilities 
should be tied not so much to the type of failure as to the 
effects of the failure. 



Yes . Structural failures are unacceptable by the public . 
The Martin 202 and Comet were types of structural failures 
that the public would not accept. Therefore, structural 
failures should have a very low value so that their 
occurrence should occur once every 20 years, or based on 
the expected passenger miles projected for the 1980's, 
values of structural failure that result in the loss of 
only one jumbo jet aircraft from a structural failure in 
20 years . 

Failures occur most frequently from pilot error, power 
failure, system failure, gear failures, and structural 
failures . To improve the overall failure rate a study of 
these failures must be done to uncover why the failures 
occur. Setting different levels of failure is too premature 
First a better data collection and failure analysis must 
be made to discover the underlying reasons for failures 
that need correction. 



Yes, depending on mode of failure, e.g. partial failure 
in fail-safe structure or slow crack growth in safe -life 
structures detectable by routine inspection and repairable 
will permit higher stress levels and conceivably better 
systems performance. 



II-6 



The probability of failure can not be predicted with any 
degree of accuracy at the present time, so it is not 
feasible to distinguish between different types of failure 
at the present time. Certain parts which can cause loss 
of an aircraft should be considered differently than parts 
that can be lost without seriously endangering the 
completion of a flight or mission. 



-* ■* 



The consequences of a local buckling failure and a tensile 
failure can be vastly different with respect to safety 
in flight and therefore we might tolerate a higher 
probability of buckling. 



#■ ■* * 



Clever design should be able to provide a forgiving or 
tolerant enough structure to make this question not 
relevant . 



* * 



le. How can the climate in the public and in congress be changed 
from an "anti-technology fad" to an understanding of the 
engineering process and the risks connected with new developments? 



You answered the question. If it is a "fad", time alone 
will improve the climate . 

*■ ■* •* 

The current "ant i -technology" fad, particularly in Congress, 
is in part related to an inability to communicate in a 
meaningful way. Tangled up in the problem are the economic, 
environmental, and social problems that have led to a 
near breakdown in meaningful exchange. It sometimes 
appears that in major weapons systems developments the idea 
is never clearly brought out that there are technical 
problems in such developments that challenge the then current 
state of engineering development and understanding, and 
that their solutions cannot be anticipated at the time a 
contract price is negotiated. Both the developing 
organization and the purchaser (generally, a DoD agency) 
will have to be more candid in discussing (l) possible 
problem areas, (2) the stretching of the state of the art 
to solve the problems, and (3) the risk of success or 
failure. Congress will have to learn how to listen, how 
to question, and how to evaluate with a minimum of political 
haymaking . 

* ■* •* 



II-7 



By establishing a reputation for telling the truth. Better 
communications with the legislators at all levels . How 
about a monthly newsletter from each Contractor reporting 
major test, flight, or service failures. With an 
explanation of their cause, correction and technical 
implications -- candid to the point of admitting errors 
of commission and omission. In this some of the calculated 
risks in design could be explained. 

■* •* -* 

Better presentation of the facts and alternatives by news 
media and television reporting are needed. The "anti- 
technology" factions seem to get their points well 
publicized even when there is little technical basis for 
some of the points. The publicity during the SST debate 
is a case in point. The public and many Congressmen seem 
to have difficulty distinguishing the difference between 
someone's theory and results based on experimental evidence. 

#• ■* # 

By the military services, the National Science Foundation, 
the President's Scientific Advisory Board, the Ford Foundation, 
the AIA, the AIM, the SAE, ASNE, ASTM and other groups 
being contacted to have joint meetings to decide on a course 
of action. Probably a T.V. public relations concern would 
have to help, plus newspapers, plus the college boards, 
and large corporations that need the technology. 



I don't believe patience is enough. We have to use the 
media to play up the good side. AIA, ATA, AIAA, SAE, 
etc. all are missing the boat here. 

* * * 

Hostility and fads can only be neutralized by education, 

■* -* ■* 



If. How much does the new outlook toward fatigue, fracture 
mechanics, and service life affect our traditional attitude 
about limit and ultimate load factors? 



The detailed consideration of fatigue, fracture, and service 
life in the design of complex aircraft systems can lead 
to better and safer structures . It may reverse the trend 
toward continued utilization of ultra-high strength materials 
of dubious flaw tolerance, which is the direction that 
ultimate and limit load concepts leads . It also is bound 
to stimulate the detailed materials and processes research 
necessary to develop flaw tolerant materials. 

* * -* 



II-8 



These are new tools which indicate that a critical crack 
length must never occur, that the decrease in residual 
strength due to a crack increases the risk of a load 
exceeding this strength and therefore increases the risk 
of failure of the entire vehicle . To decrease the risk 
of failure the margin of safety of 1.5 should be studied 
further to include the time factor of rate of crack growth, 
loss in residual strength, and probability of failure. 
Notice that for titanium and graphite epoxy at ultimate 
strength to one "g" stress must have a factor of more than 
3.75 (1.5 x 2.5 g) for a transport critical for the limit 
load factor for maneuver of 2.5. 

-* * * 

This is a healthy attitude provided it doesn't go overboard. 
Structural engineers have always considered fatigue, 
fracture mechanics, etc., at least since the 1930's. But 
it takes a political hot potato like the F-lll and its 
problems to dramatize the problem. We will always have 
to be concerned about static loads -- but with more and 
more emphasis on probability of occurrence. 



I don't think that fatigue, fracture mechanics, and service 
life is changing our attitude about limit and ultimate 
load factors at the present time. The limit and ultimate 
load factors used provide a sort of baseline configuration 
from which to work. Structural weight must then be added 
to meet other requirements such as structural life, damage 
tolerance, etc. It will be some time before structures 
can be designed on the basis of failure probability rather 
than on the basis of an ultimate load factor of 1.50. 



Design for service life makes the ultimate load factor 
concept obsolete and of little value. The design must be 
based on a load spectrum that includes a limit load 
defined as the load the vehicle must sustain with some 
prescribed probability of success on the last mission 
in its total service life or in each inspection interval. 



I'd say anyone that thinks traditional factors should be 
attacked does not really know the vagrancies of our 
knowledge base. I believe in rationalization but not 
exploitation! 

-* * * 

It has not "sunk in" that improved methods of analysis, 
materials with yield to ultimate strength ratio different 
from aluminum alloys, better load prediction methods, 



II-9 



process control etc., could obsolete the old traditional 
ultimate (ignorance) factor. Substantial improvements 
in systems performance and safety can be envisioned if 
arbitrary load factors are replaced with a design criteria 
system taking critical load or load spectra and critical 
strength or failure mode into account. All that can be 
said in defense of tradition is that such system is not 
yet developed. They should however not block the work 
needed since tradition is as incompatible with engineering 
as it is with military science. 

* * * 



2. Regarding development of metallic materials . 

2a. How much can reliability be improved by close control of 
processing? 

Substantially. However, a check is needed at the source 
of the material, before it is fabricated, and after it is 
fabricated, to assure compliance with the specified 
process . 

#■ * * 

A quantified answer to this question is difficult to come 
by. However, close control of processing, if it is 
economically acceptable (which is part of the risk analysis 
problem), can be expected to improve reliability, and may 
result in higher and safer design stress levels. This is 
really the reason why most major aerospace companies 
maintain highly motivated and competent materials and 
processing personnel. 



Reliability can always be improved by closer control, but 
getting the producer and processer to improve control, 
particularly with no real definition on how close is 
close, is impossible. Therefore, the only practical way 
of improving reliability in metallic materials behavior is 
to develop more definitive KDl/NDT techniques along with 
fracture mechanics analysis to predict the influence on 
structural strength. 

A practical way of establishing meaningful producer/ 
processer controls would be to define a financial 
responsibility. Presently, if a producer /processer gives 
you "bad" material, his only responsibility is to replace 
the material, not the configured part that usually costs 
an order of magnitude greater than the raw material. If 



11-10 



the producer/processer can be made to bear some of this 
expense, control will improve dramatically. 

* * * 

Modest improvements could be expected in the order of 10$ 
less improperly processed parts. More destructive testing 
of actual production parts is far more effective than 
200 or 300$ visual and non-destructive tests or the 
examination of standard test coupons. Particularly, 
applies to welded, bonded, cast, and forged parts. Is 
expensive in practice but compared to the loss of a 
vehicle, extensive investigation, or proof tests relatively 
cheap . 

#• #- * 

Perhaps a great deal for structural applications which 
are not tolerant of flaws. Process controls have to be 
compatible with fundamental reliability requirements and 
generally are so restrictive as to be premissive of no 
real variation. 



S ub s t ant i ally . 



•* 



-*■ * 



2b. Has there been a wrong emphasis in the development of 
metallic materials on increasing strength -density ratios 
instead of improving fatigue and fracture toughness? 



It is not entirely a case of wrong emphasis. There has been 
a real trend in the development and use of high strength- 
density metallic materials, in accordance with established 
design practices, in order to solve, presumably, the 
mutual problems of performance and weight. While fatigue 
and principally fracture become more critical with this 
trend, it is only recently (5 years or less) that acceptable 
organized design procedures have evolved, particularly to 
handle the fracture and crack propagation problems. Because 
of the attention that has been paid in characterizing the 
phenomena and in developing rational design procedures, 
less attention has been given to the alloy development 
activity. Inclusion of a damage tolerant design requirement 
in the B-l is certainly going to stimulate this kind of 
research, in view of the gains already evident. 

The slowness in the improvement of materials for fatigue and 
fracture resistance probably is more closely tied to 
availability of funds than to lack of ideas. Alloy 



11-11 



development for structure sensitive properties, such as 
fatigue and fracture, is a time-consuming and costly- 
operation. In comparison with construction materials and 
automotive materials, base sales of aerospace materials 
are small enough at present to discourage large-scale 
research of this kind without massive interest of DoD 
and NASA agencies in the form of funding. 

-* ■* -* 

The fatigue and fracture mechanics tools are relatively- 
new and they certainly need to "be exploited to the fullest. 
However, new metallics should be judged considering many 
factors. The strength-density factor is only one ingredient 
and should also be obtained. Frequently the fatigue 
strength after 10° cycles is a percentage of the F-^ u value. 
Also, the K^ c value is often a percentage of the F-fc u value. 
The residual strength, (initially Ftu)> decreases with 
crack size, stress range, stress intensity factor (Kq), 
and type of construction. 

The slowness of improvement in fatigue, or the recognition 
of a fatigue problem, is caused by the slow development of 
the finite element approach to discover the state of stress 
causing a fatigue problem and/or a fracture mechanics 
problem. As soon as the stress field is known the fatigue 
problem becomes obvious. 



Probably, although I think it is pretty much recognized 
today that tensile strength is an inadequate description 
of the usefulness of a material. It does provide a 
"floor" for comparison. 

* * * 

Possibly --we are approaching minimum gage restraints 
in many applications. Therefore, higher strength, per se, 
may not always be needed. Most of our systems are fatigue 
and crack growth critical now. Since fatigue, crack growth 
and fracture toughness do not generally increase with 
increasing strength, more emphasis should be placed on 
these properties in developing new and improved materials. 



This emphasis was probably proper in the past but in the 
future we must seek a proper balance between strength, 
fracture toughness and crack growth resistance of materials. 
This balance depends on the particular application. For 
parts with long cyclic life, the crack growth resistance 
is the most important characteristic (and the least studied 
and understood). The high strength of a material may not be 
usable in such a case because the material has a low 
resistance to crack growth or low fracture toughness. 

# * # 



11-12 



If this is a true assertion, ray answer "would, be yes. 
However, designing to accommodate any particular deficiency 
is an alternate if attributes more than balance deficiencies 



* -* 
Yes. 



Yes, I think so, since most of structural problems are 
associated with repeated tension loadings and little 
advantage can be taken of the high strength properties . 
Higher fatigue properties would give us a greater advantage. 

* ■* -* 

Yes. Emphasis on strength /density ratios has been beneficial 
in developing high strength materials up to the point where 
toughness and fatigue transitional effects become dominant. 
The strength/density emphasis was also necessary in order 
to highlight the need for materials where fracture toughness 
and fatigue keep pace with strength. However, strength 
levels have presently reached the point where emphasis 
must be placed on fracture toughness /fatigue criteria in 
order that these materials remain practical for utilization 
with a degree of confidence . 

* •* * 



2c. How much could be accomplished with additional funding for 
the development of metals? What kind of incentive is required? 



To take an example in regard to what can be accomplished 
with additional funding for research in metals, I cite 
recent B-l experience with Ti-6A1-Uv alloy. This alloy, 
the workhorse of the titanium family, has a minimum fracture 
toughness variously estimated to be in the range of 
30 - Uo ksi-inch 1 / . With the imposition of a fracture 
criterion for the airplane, funds from the Air Force, and 
good work, North American Rockwell now is specifying the 
same alloy with a minimum fracture toughness of 70 or 75 
ksi-inchV^ . The improvement in crack tolerance, by 
doubling the fracture toughness, is vitally important to 
the airplane . The ingredients or incentive to accomplishing 
this were (l) the requirement of fracture-safe design, and 
(2) funds to accomplish the development work. 

I would expect there are still improvements to be gained in 
the upgrading of fatigue and fracture behavior in metallic 
systems of aluminum, steel, and titanium. There probably 
will be modest percentage improvements for the first two 
alloy systems. I prefer to think that titanium has yet a 
way to go in useful alloy development. 



11-13 



The diffusion bonding of metallic parts into a rood 
compression and fatigue resistant structure with light 
weight foils has been attempted but the development costs 
are high. If the incentive of a bonus for higher strength 
small components is made with all the development costs 
sustained by the government, then metallurgists and structural 
engineers would devote their effort in that direction. Heat 
testing a matrix which is diffusion bonded and later 
etched out to reduce the wall sizes also needs to be tried 
with small components. 

Adding small quantities of high strength materials to low 
strength materials (Beryllium, Boron, or graphite in 
aluminum or magnesium or glass or an epoxy matrix) . 

Making sandwich structure with high fatigue strength of the 
core with high strength edge attachments could be developed. 

Etching of covers and diffusion bonding to metallic cores 
to make up light weight sandwich structure has been partially 
successful. However, a development program to set up a 
production method using Vee cores or double Vee cores or 
light weight channel sections has not been successfully 
tripri . The etching can provide fail safe crack stoppers 
and end attachments . 



The sky is the limit. Physicists say we are realizing 
maybe 10$ of atomic load capabilities . Profits are our 
incentive -- government should give us the opportunity. 



Magnitudes are not the most important -- put the money where 
the capability exists and the incentive exists, i.e., where 
the development falls into the company's normal sphere of 
business . 

Theoretical considerations along with practical attainments, 
say that steels with 500 ksi, aluminum with 150 ksi and 
300 ksi titanium are totally feasible. 

■* ■* ■* 

Funding should be concentrated in the critical, now better 
understood areas of fracture strength, fatigue strength, 
creep strength, and in stress corrosion and materials 
process control. 

The probabilities to obtain significant improvements are 
s irprisingly good, judging by recent breakthroughs like the 
improvement of creep strength of titanium alloys at high 
temperatures , etc . 



II-lU 



I am a composites enthusiast but I see aluminum and steel 
continuing to "be the workhorse structural materials for the 
next twenty years even if no significant improvements in 
properties are made. As for titanium it will continue to 
be used where the operating environment requires it. I 
cannot assess the probability of improvements. 



2d. How much can we learn from the fact that corrosion resistance 
and crack tolerance of aluminum alloys were improved considerably 
in the last few years by a relatively small effort after a 
stagnant situation for more than a decade? 



I am not sure of the phrase "relatively small effort". 
All the DoD agencies and NASA, as well as the aluminum 
companies, have spent considerable sums of money in 
developing alloys and tempers of high-corrosion resistance 
and crack tolerance. However, in answer to the questions, 
it was in part a response by a knowledgeable industry to 
a known set of technical deficiencies. It represented also 
an approach by the industry to protect its markets in the 
face of newly emerging structural materials, such as 
titanium alloys and composite materials. Finally, it 
probably was stimulated by some urgent prodding by DoD 
and NASA personnel, as well as the aerospace community. 



The theory for crack propagation was available and the 
finite element method of analysis was developed in the last 
decade to confirm the theory and investigate the problem 
further. Also, most important, the desire to find out 
how to practically deal with the problem was present. 
Many large size panel and fuselage components were tested 
at considerable expense at Douglas for the DC8, DC9, and 
DC10 to develop crack tolerant structure. Stress corrosion 
has also been studied and corrosion resistant coating have 
been successfully developed when the problem was recognized 
and needed solving. 

* * * 

This suggests to me the classic situation that prevails in 
many disciplines . A period of little apparent progress 
during which research in related areas continues . And 
then a period of rapid growth made possible by the synthesis 
of several ideas that have incubated and reached maturity. 

* * * 



11-15 



During that period, objectives were not clear, research was 
done by all because it was fashionable and the Government 
paid for it -- achievement was not the goal used to measure 
success, only budget and schedule. 



State the real problem and ingenuity can find a solution. 
In this case with minimum subsidy. 

We may learn that progress is not so much dependent on 
magnitude of funding but on clear definition of problems 
which must be overcome and on time to solve them. 

* * * 



2e . In view of the small market in aerospace compared to the 
vast automotive market, is there a possibility to apply aerospace 
R&D toward reducing automotive manufacturing costs and to 
establish a broader base for development funding? 



We are discussing our methods with a leading automotive 
company and if we can "sell" them on our advanced methods 
we expect them to do more R & D in the use of applications 
common to both and thus broaden the base for development 
funding eventually. The first step, and only step we are 
now pursuing, is to teach them advanced methods of 
structural analysis. 

Remember that initial tooling and manufacturing costs are 
the big items, a pound of weight saved is not as important 
to an automotive concern. We must convince them that the 
pound saved will yield fuel economy, less braking, smaller 
brakes, and the ability to put the pound into items that 
will increase the life or comfort level. How much they 
are willing to pay for this pound is not known. It is not 
uncommon for a long life advanced aircraft to use a 
$300/pound breakeven value. Can an automotive firm pay 
that price? A 1500 pound automobile, if all items were 
assessed at $300/pound would cost $U50,000; a 1500 pound 
auto is more like $1.5 per pound. Therefore, it is only 
in the analysis, manufacturing, and tooling development we 
can help the automotive concern if they are willing to 
use materials we are familiar with -- aluminum, titanium, 
super alloy steels, and possibly composites. 



II-16 



Should the question "be turned around? Is there a possibility 
that automotive R&D with its broader base of development 
funding can be slanted to give benefit to aerospace? The 
area that is common to both is in the manufacturing 
process field; forming, metal cutting, joining, shaping, etc, 
I assume that manufacturing people in aerospace stay pretty 
close to developments in the automotive field. 

•* #■ ■* 

Unlikely, since our objectives do not coincide. However, 
we are each being forced to consider the others point of 
view, namely, reliability, light weight, and safety of the 
aircraft industry and the low cost, high automation, and 
prototyping of the automotive industry. This is an exchange, 
even now, of each others R&D developments, ECM, EB welding, 
reenforced plastics, turbine engines, numerical control, 
molded plastics, fasteners, adhesive bonding, etc. It 
does not appear that the automotive community would want 
to underwrite R & D in aerospace since it would not be 
oriented directly to their needs . 

■* -x- * 

Not probable because aerospace objectives are totally 
different than automotive, i.e., aerospace — high 
reliability, long life, low production, hostile environment; 
automotive — mass production, planned obsolesence, etc. 



The economics of the two industries is so different that 
this possibility seems very remote in a useful way to me. 

*- •* * 

Surel!! But who is doing it? I think the government should 
provide leadership. Our professional societies have 
struck out so many times. 



I do not think so. Requirements are generally too different. 

•* ■* * 



3. Regarding present trends in fatigue criteria 

3a. Have we challenged the designer to recognize problems and 
limitations applicable to fatigue, particularly with regard 
to processing and inspection capabilities? 



11-17 



Many good designers are aware of some of the important 
factors. However, often these designers are not as close 
to the detail design as they should be. Therefore, for the 
average designer a handbook should be prepared to give 
him the necessary process and inspection specifications 
to note on the drawing. 

-* * -* 

I think the designer is unaware, generally, of the potential 
variation that processing has on fatigue-crack propagation, 
though somewhat better informed on fatigue behavior. 

* -* * 

I think our designers are fairly mature with respect to 
recognizing the differences between paper designs using 
paper properties and the real world. This is not to say 
that a continuing effort is not required to maintain and 
enhance this awareness -- especially for the relatively 
new and inexperienced engineer. 

* * * 

Designers are aware of most of the tricks used to improve 
fatigue in structure, but they can only call out an existing 
specification to cover processing and inspection — this 
is an area that needs much work. 

* ■* * 

Designers have been challenged but it went over their 
heads. Much more needs to be done in this area. Most 
know fatigue problems are detail design and/or manufacturing 
deficiencies. 

* •* -* 
Not yet adequately. 



Designers will begin to recognize that flaws are 
characteristically present in structures and that precautions 
in material selection, processing and inspection are of 
paramount importance, particularly where high strength 
structures are concerned. Fatigue cracking rates and Kj c 
data are presently still research items, however, it is 
anticipated that this will reach the designer as a practical 
tool in the near future. 



3b. How much can be gained from materials application compared 
to materials research? 



11-18 



Material development needs both. Separating the materials 
research from the materials application is often done for 
convenience but the complete package is necessary to 
prevent the designer from going from the research results 
without the confirmation of an application result. 



* 



Materials application of improved state of the art 
materials (somewhat nonconventional) will point up the 
need for more research as materials and attendant 
manufacturing difficulties are encountered. 



•* * 



Both areas are complementary -- one without the other can 
only be a repetitive function, not an advancement. 



Research results will not be used unless there is an 
adequate materials application program. 

¥■ •* * 

New applications should be emphasized but not at expense 
of research. We must exploit what we have at hand and 
at the same time provide for the longer range future. 

#■ * * 

Wrong question. They should complement each other. 



3c. Would tightening of specifications and processing methods 
for titanium, for instance, result in consistently finer grain 
size and better fatigue properties? 



I am not completely certain that we can at present specify 
process controls for titanium and other alloy systems 
that will optimize fatigue properties. This is one of 
the research areas, particularly for titanium, that needs 
definite study. 



Trobably not -- only if finer grain size caused fatigue 
improvement and there is contradicting evidence here — what 
is needed is to define the parameters in the metal that 
control fatigue behavior before we can "shot gun" an 
improvement with any chance of success, not just increase 
cost. 

* * * 



11-19 



Yes. A tightening of specification and processing methods 
is desirable. However, the effect on cost should be given 
also so that the designer is given a choice depending on 
the material application. 



* 



Hopefully. I would prefer some better understanding of 
what causes what and then refining. 



-* ■* 



This is needed once processes have been defined and cost 
effectiveness is established. 



Not necessarily. Might tend to price material out of 
realm of practical utilization. Applied research and 
development in addition to more stringent specification 
requirements are needed. Changes in titanium specifications 
brought about by alpha segregation problems have led to higher 
prices. This type of defect has only contributed to 
rotating engine part failures, so unnecessary tightening 
can lead to higher prices and therefore decreased usage. 

-* * ■* 



3d. Realizing that fatigue analysis is not yet an exact 
science and that fatigue design has to be based on previous 
experience, how can this experience be evaluated systematically? 



We have a fatigue checkoff list which is needed before signing 
off a drawing. Also, we have put together fatigue problems 
in a fatigue course. Unfortunately, the fatigue book 
has not always been kept up to date with examples of 
failures from other manufacturers . Perhaps the military 
services should fund a project for collecting these fatigue 
examples which all manufacturers would contribute to 
periodically and the services would redistribute periodically. 

* *• * 

I assume you are asking how can we learn more from the sum 
total of fatigue failures that have been experienced. We 
all agree (I think) that we can learn more from our failures 
than our successes in material application engineering. 
But there is an understandable reluctance to publish the 
failures. The most interesting technical papers I have 
read have been those that were completely candid about the 
wrong turns, the mistakes, and the failures, prior to the 
successful solution. Let's encourage more of this through 
our professional societies. 



11-20 



One can write a book on this question. There are some 
simple elements to approach this problem, some or all of 
which are being pursued to some extent. At the design 
stage, one's past experience with a similar airplane 
provides a link between design loads and real loads, and 
between design response (deflections and stresses) and real 
response . This information coupled with careful history 
on fatigue initiation and propagation in the previous 
airplane, provides some semblence of how the new structure 
will behave. The ingredients involve evaluation of what 
happens in many numbers of a fleet of a given aircraft 
during service in order to characterize realistic stress 
spectra for the aircraft. Similarly, one monitors the 
fleet for fatigue damage. Based on developmental fatigue 
tests of structural components, on fatigue of the basic 
materials, and on the fatigue test of the entire airplane, 
one can compare, analyze, and evaluate real life behavior 
and the basic soundness of previous programs and design 
philosophy. 

#• ■* * 

First the variables affecting fatigue life must be properly 
identified. Then failures obtained in service or test 
would have to be defined using these same variables. These 
failures could then be used as data points from which 
relationships could be established between the variables . 
The relationships could then be used for predicting future 
experience. 



This could be the theme of a book and has been on numerous 
occasions. I have no handy solution. Design handbooks of 
not-to-do's come closest perhaps or, better, best-to-date 
approaches could be useful. 

* * * 

Probably not beyond what to do and what not to do. 

* •* * 

Previous experience mostly allows one to better evaluate 
stress concentrations and better define design load spectra, 
An attempt to classify Kt for practical structural con- 
figurations might be helpful. Also design load spectra 
that have been inadequate for certain systems should be 
corrected and applied for similar applications on the 
next design. Attention should also be paid to those 
systems where the design spectrum or Kt has been assumed 
too severe by reducing the requirement in the next design. 

* * * 



11-21 



4. Regarding current trends in fracture mechanics and fail-safe 
design 

4a. W. S. Hyler's presentation indicated how much scatter can 
exist for K^ c data, even at uniform F y . This reflects on 
uncertainties of materials manufacturing, i.e., "process 
capabilities". What practical conclusions can be drawn regarding 
processing and specifications? 



Further analysis at Lockheed of D6AC Kj c data indicates 
that the Kj c value is a function of the quench rate obtained 
during processing which in turn is a function of the 
thickness of the part processed. The tensile properties 
in this case were not affected by the quench rate or 
thickness. The obvious conclusion is then that Kj c type 
tests must be specified as part of the process control 
if the fracture toughness properties are the properties 
that must be controlled. It is becoming obvious that 
ultimate strength tests can not be relied on to adequately 
control material properties such as fatigue, fracture 
toughness, crack propagation, etc. 

■*- ■* •* 

After further research, a specification which indicates 
the quench rate, heat soak time for grain size, and the 
allowable impurities necessary, plus the inspection and 
sample testing necessary, should be specified versus the 
coefficient of variation versus the cost comparison for 
an A, B, and C type value. 



Mr. Goepfert of ALCOA, a number of years ago, indicated 
on the basis of a large amount of testing and statistical 
evaluation that a specification does not control a process, 
rather it is the process that leads to the specification. 
Therefore, the decision to establish in a specification a 
minimum Kj c level without sufficient knowledge of how a 
given process schedule influences the distribution of 
Kj c values for the product can only lead to erroneous 
values, usually too high. Specifications can provide 
target values toward which the producer may move by 
modifying his process. But if no process can be evolved 
to acceptably meet this target, then it is the process 
that controls . 



1. Don't give up — keep working; 

2. Don't depend on material quality alone; 

3. Recognize possible variation and design around. 



11-22 



Extensive work is required to research processing effects 
and to establish processes resulting in consistent properties 
It is obviously feasible to get improved toughness. 



* 



As fracture concepts become more prevalent, Kj c criteria 
will become a standard spec, item for high strength 
materials. This should lead to processing methods for 
optimizing fracture toughness. Vacuum melting along with 
fabrication techniques designed to minimize anisotrophy 
effect will become requisites for obtaining consistently 
high Kj c /Fty properties . 



•* 



Ub. How can material be characterized clearly for fracture 
mechanics with respect to processing as well as thickness? 



A significant amount of testing is required in order to 
establish process capability for fracture just as for 
establishing process capability for F^- u and F+ . 

* * -* 

The fracture toughness properties must be characterized 
according to the variables that can affect the properties, 
e.g., quench rate and thickness for D6AC steel. More work 
needs to be done on materials to determine what effect 
processing variables have on the material properties, so 
that better processing specifications can be written. 
Sufficient tests need to be conducted so that MIL-HDBK-5A 
type allowables can be established. 



The averave K-j_ c value can be tested to determine its value 
versus thickness (using a close range in processing and 
inspection for the samples tested). From the sample 
size the confidence level can be given for each specific 
thickness . 



By process specification and more exploratory work in 
mixed mode fracture and ultimately design oriented testing, 

* ■* * 

We don't know yet, but it must be through experimental 
correlation and improved analytical tools. 

* * * 



11-23 



4c. What guidelines can be established for trade-offs between 
higher strength and lower IC ? 

The Kj_ c value influences the rate of crack growth and the 
residual strength versus crack size. A chart which shows 
the residual strength, rate of crack growth versus K^ 
with constant values of one "g" stress, desired life, 
and range of stress or, (olVl)? coefficient of variation 
of the loading, can show the designer the compromise he 
must make on the value of K^ c and Fy (which influences 
the initial residual strength). 

* •* -* 

Material and component tests, if done properly, should 
yield precise results. 

* * * 

There are many parts of a structure where one need not be 
concerned with fracture. Compression surfaces are one of 
these. High strength, moderate to low fracture toughness 
probably provides no problems in these areas and one has 
the advantage of high strength. Areas in certain aircraft 
structure offering multiple load paths may be another 
place that can dictate tradeoffs between strength and 
toughness, depending upon considerations of risk and > 
reliability. 

* * * 

I should think that highly redundant structures as well 
as those employing fail-safe design practices could lean 
more toward higher strengths at some sacrifice in K]_ c . 
But as noted earlier, so much depends on the application 
and the acceptable risk. 

*■ * # 

The tie-in here is the allowable flaw size that can be 
tolerated in the structure. If NDI capabilities are such 
that the flaw size existing in the structure is very 
small, then F^ n or fatigue properties may design the part. 
However, as the flaw size existing in the material increases, 
fracture toughness properties become all important. More 
work needs to be done to improve NDI techniques and the 
reliability of these techniques. 



For each design situation (load spectrum, initial flaw 
size, and required service life) the relative importance 
of ultimate or yield strength, fracture toughness and 
crack growth resistance can be determined and the extent 



11-24 



to which each might profitably "be increased or decreased 

established. Unfortunately little such analysis has 

been undertaken and none of it was evident at the symposium. 



#• * ■* 



Question whether there is a real trade-off involved. 
Design around! 



•* 



A good fracture control plan covering all aspects from 
basic material to in-service inspection. 



* 



The impact of reducing the strength of a material to 
improve K lc would be less if fatigue life properties were 
retained. 



1. Establishment of minimum critical crack sizes and 
resulting Kj c /Fty requirements . 

2 . Evaluation of Kj c versus Fty data to determine how 
rapidly Kj c falls with increasing Fty and practicability 
of controlling material within required Kj c /Fty range. 

3. Correlation of cyclic crack growth and Kj scc properties 
with critical crack size and NDT inspection. 



kd. Can any conclusions be drawn with respect to fail-safe 
design? 



Probably, to me, the most unsettling factor in damage 
tolerant design is its use with thick section materials, 
where the critical crack size may be borderline with 
regard to the probability of detecting a crack. NDI 
techniques are reputed to be capable of finding quite 
small flaws. However, the probability of finding such 
small flaws may be equally small in practical situations 
Consequently, in damage tolerant design, it makes some 
sense, to me at least, to base propagation life and 
inspection intervals on a crack length eminently capable 
of being found most of the time, rather than a length 
within the minimum bounds of the inspection device. 



11-25 



Fail-safe design is a design that according to the civil 
specification can sustain 80 percent of limit load at any 
time during its life time. To insure this, the inspection 
interval must be short enough to repair any cracks "before 
the residual strength decreases to the 80 percent level 
or the critical crack length occurs (fast fracture occurs). 
With transport vehicles, the requirements also desire 
that a full bay crack can be sustained before failure. 
Now, Ki c influences the residual strength and the rate of 
crack growth. It would be desireable that Kj_ c have a 
coefficient of varation small enough so that the value 
used for predicting the residual strength and inspection 
interval, (before a full bay crack occurs), is reliable. 
As shown, in Mr. Fischler's presentation, the decrease 
in the coefficient of variation, (a s /u- s ), increases the 
probability of failure greatly especially for a V/STOL 
aircraft with a high load spectra coefficient of variation, 
(oj/iij). Therefore, specifications which insure values 
of K^ c within a narrow range are desireable. 

-* * * 

By fail-safe design, it is implied that cracks, if they 
occur in the structure, will be found before they become 
catastrophic. Therefore, heavy reliance is placed on 
finding these cracks during routine inspections. To 
exploit the higher strength potential of metallic materials 
will require a substantial improvement in inspection 
capabilities to achieve the same reliability as with 
current materials at lower strength levels. 

* # # 

Fail-safe design has many merits, but it was not discussed 
in any depth at the symposium. 

* * * 

I sure do not agree with Fig. C-Uo regarding fail-safe 
problems. It reminds me of my dad's philosophy that "you 
only get out of things what you put in". If one sets 
out to prove that fail-safe designs can be devised that 
have problems, he can sure do that. If he intentionally 
configures so as to minimize "possible" problems, they 
will not arise. 

* * •* 

Fail-safe design should be employed wherever feasible, 
where not fracture control by use of tough materials, 
proof testing, etc., should be employed. 

* * * 



11-26 



There is more inherent fail-safe capability in a structure 
than will ever be calculated. On military aircraft there 
is great danger of going overboard with fail-safe require- 
ments unnecessarily. This has been shown by aircraft 
that have returned with significant battle damage in 
components that provide no visible fail-safe load path. 
On military aircraft the fail-safe concept should be 
utilized so long as it does not increase airframe weight. 



5. Regarding evaluation and application of structural metals 

5a. A report NMAB-2U6 was published in 1970 under the title: 

An Approach for Systematic Evaluation of Materials for Structural 

Application. It contains an outline for a data information 

system requiring data banks on material properties, material 

evaluation techniques, and applications analysis. Is there 

a follow-on area where fruitful concepts could be developed, 

e.g. 

failure analysis to develop new test techniques; 

case history development to support new test needs; 

or possible contributions of information centers? 



Yes / a follow- on area where failure analysis theories can 
be confirmed by tests would be desireable. I would like 
to test materials, find out their F-j- u , F^ , and K values 
from coupon tests, obtaining enough tests to get a high 
confidence level for their coefficient of variation. 
Then, using parts of the same sheet as the coupons come 
from, make up a test component of structure. I would 
test some with and without an initial crack at different 
mean stress levels and different stress ranges, periodically 
checking the residual strength by failing some of the 
specimens. The specimens remaining should be continuously 
cycled till every one fails. The actual life, residual 
strength, and rate of crack growth should be compared to 
the predicted values to determine the accuracy of the 
theories . Adjustments to the theories should be suggested 
to account for the slowdown in crack growth when the 
cracks reach attachments . 

* * * 

I'd like to see an outfit like the Batelle DMIC try out 
the ideas contained in NMAB-2U6 on a selected type of 
structure or component. This should not be a big deal but 
could serve as a pilot operation. 



11-27 



The Case History approach is good and might he a way of 
motivating more reports on failures, as for fatigue. 
Perhaps Don Shinn can spark plug a Failure Reporting 
System. I don't look for much results from failure 
analysis per se except as it contributes to the Failure 
Report . 



Tests in laboratories do not simulate real applications 
and environments. Some sort of technique is required to 
define the relationship that exists (assumed) between 
laboratory data and the real service behavior. 

* * ■* 

Development of systematic materials evaluation techniques 
should be continued and integrated into structural design 
procedures with the goal of developing a highly automated 
and interactive vehicle design system. 

* •* -K 

We feel steps should be taken to implement the NMAB suggestion, 
Good accelerated service simulation tests for time -dependent 
phenomena such as corrosion, radiation, creep, etc. are 
musts. Case histories are good but only half of the 
story — predicting and intercepting new problems is the 
other half. 

Regarding information centers, a major problem is how to 
get people to overcome the old NIH factor. A more modern 
version seems to be "I'd rather do it myself". 

* *■ *■ 

Data bank idea for materials is a good idea. 

-* •* •# 



5b. Is "full-scale" demonstration a barrier for new materials? 



Yes. The cost of full-scale demonstration is usually so 
high that no funds are available for new material tests . 
Full scale demonstration tests accomplish little. Usually 
they load the vehicle with one spectrum critical for only 
one component. Therefore, it has only limited use at a 
tremendous cost. Many new material tests of coupons and 
components could be made for the same cost. 

■* * -* 

Yes — cost. 

* # # 



11-28 



I prefer to think that a full-scale demonstration is a goal, 
a successful milestone to be reached rather than a barrier. 
It also serves as a way to pick up material behavior and 
processing difficulties that may not otherwise be revealed 
by small element tests . 



Experience with full-scale applications is essential to 
develop the confidence required for acceptance of a new 
material. 

*• * -K 

No! Lack of incentive and opportunities are real barriers 

* * # 

No, lack of cost data and service experience are. 

#■ *■ -* 

"Full-scale" demonstration need not be a barrier for new 
materials . The 5-year delay between the inceptual stage 
of a new material and its first flying application is 
largely attributable to the "no one wants to be first" 
philosophy. R&D support for the user who is willing to 
be first would encourage more risk taking, since present 
fixed price contracts penalize risk taking. 

#■ -x- -* 



5c. How is the materials -structures interface controlled? 
How can or how should it be? 



Usually by poor coordination between some of the necessary 
parties. It should be controlled by a high level management 
group with all the necessary parties represented. Because 
fracture mechanics is so important, the Structural Mechanics 
representative should be chairman of the group. 



Several ways: (l) Material process specifications; 
(2) Structures Design Hiilosophy documents at outset of 
new program; (3) Structures Design Manuals (company) 
which set forth design allowables, exceptions, qualifying 
assumptions re. usage of materials; (h) Joint Materials 
and Structures participation on Material Discrepancy 
Review Boards and on Structural failure analyses; (5) Joint 
review of lay-outs and detail design drawings . 



11-29 



Prefer to think of the interface as a dynamic boundary 
tending to shift toward structures when well established 
materials are used in design -- and towards materials 
when new or unusual uses or environments are encountered 
with established materials or wholly new systems are used 
for the first time. Materials should characterize each 
new alloy or material system, particularly regarding 
properties other than strength such as: corrosion resistance, 
embrittlement susceptibility, weldbondability, fracture 
toughness, and protection system requirements. 



One way is to have the materials man sign the Engineering 
Drawing for approval of materials and processing in 
addition to the Stress Engineer's sign-off. 



By hard work and recognition of potential problems and 
solving them in advance. 



Mainly by the structures design group with coordination to 
materials and manufacturing. 



5d. Are there any subjects related to the preceding topics 
which an NMAB Committee could usefully tackle? 



Yes. The committee could decide what analysis and what 
testing needs to be done to increase the reliability of 
new materials for new advanced vehicles. 

#■ * ■* 

I'm not really current on what the NMAB Committees are 
covering as of today. But — I'd like to see some 
government-industry group address itself more completely 
to the following: 

(1) The need for and possible ways of implementing 
prototype material applications. (On materials 
research vehicles? More full-scale laboratory 
demonstrations? ) 

(2) A look at the total aerospace and related industrial 
materials R&D in the U. S. (both government and 
private) with the objective of identifying imbalances 

(3) Consider and propose several realistic approaches to 
risk evaluation and probability of failure and what 
types of data must be generated to make this feasible 

* ■* * 



11-30 



See latest NMAB report on Accelerated Use of New Materials. 
It has many good ideas that should be followed up. 



6. Regarding more general design problems 

6a. Are the concepts of probability of failure and risk 
evaluation bound to become a routine part of structural analysis? 



Yes. Eventually. However, new tools are resisted for 
long periods of time. For example: 

(a) Using Power Spectral Analysis as a recognized tool 
took about ten years. Even now, certain segments of 
the industry will not recognize it as a respectable 
tool, and will not allow specifications to be written 
which include it as a criteria. 

(b) The six degrees of freedom cross coupling analysis 
was not accepted until a fighter aircraft's vertical 
tail came off because of the additional cross coupling 
loads from three to six degrees of freedom. 

(c) Flutter analysis, fatigue analysis, sonic fatigue, 
supersonic panel fatigue, creep, and computer analysis 
was resisted until they were able to weather the 
storms of protest of being called "inexact", "costly", 
"unnecessary burden", etc. 

-* -* * 

I would hope so. But this will require an extensive 
education campaign. 

* -* •* 

Yes — This is what the ASIP, now in revision, will require, 

* * ■* 

Yes. 

* * # 

They've always been. If you mean quantification and 
documentation in depth, I wonder as to the worth. 



Yes, to a limited extent on components selected for fracture 
control and only for tradeoff studies of significant factors 

* # # 



11-31 



If they do, we are going to be kidding ourselves and 
others. I would have no confidence in the numbers. There 
are just too many variables . 



6b. How can qualitative considerations be quantified for 
tradeoff and risk evaluation? 



The qualitative considerations must be put into a cost 
effective analysis to determine what quantitative minimum 
values are necessary. The risk for a new material, must 
be less for new aircraft. To achieve this lower risk, 
adequate testing must be done to at least obtain the 
coefficient of variation of strength at time zero and the 
residual strength after subjecting a primary structural 
component to the expected spectrum for the desired service 
life. 



I doubt if they can, other than by the usual expedient of 
assigning weighting factors. 

* * * 

Develop case histories over a period of time, so some 
feedback is obtained on what is obtained for certain 
qualitative considerations. 

¥r #■ * 

Depends upon the case. This is called operations research 
or analysis a la Rand, etc. 

* ■* ■* 

By application studies and tests . 

* * * 

Waste of time to try. 

* * * 



6c. What can be learned from the past to make systematic use 
of available experience? 



By doing analysis first, followed by controlled material 
and small component tests to confirm the analysis, then 
followed by only two experimental aircraft (one for immediate 
flight and another for proof testing and detailed instrumental 
measured flight) new materials with temperature inputs can be 
understood before larger scale production is initiated. 

* •* ■* 

11-32 



First we must document our failures so that we have a 
"past" to learn from. 

* * * 

Books could be written on this. 

* * * 

The $6U,000 question. 

* *■ -* 



6d. What can be done to educate materials engineers in the 
problems of structural design and structural designers in the 
problems of materials and processing? 



By the services providing institutions and large corporations 
with funds for this purpose. By making it mandatory for the 
stress signout of drawings (at some later date -- 5 years 
from now) to obtain a certificate that they have completed 
these courses before they can signout any drawings for 
military aircraft. 

* * •* 

The question implies that this is not being done today. In 
aerospace — in my own experience, at least — I think there 
is a very good appreciation by the structural designers and 
the materials engineer of their mutual problems . This may 
not be so when the men first come out of school but then 
at this point the men have no identity anyway as either 
structural design engineers or materials engineers. This 
is an on-the-job "graduate" training. 

*• *• ■* 

Each must learn enough of each others discipline so they can 
effectively communicate with each other. 



Have them work together on a design team. 

*• * * 

Age old engineering problem. Colleges and universities fall 
short here. Companies seldom have time to handle basic 
problem. Individuals must recognize and do job for themselves 
National societies could help if they really wanted to. 

* * •* 

More development time and closer teamwork. 

* •* -* 



11-33 



They can work closer allowing each to get more involved in 
the others day to day problems . Each can be rotated into 
the others area for a period of time. Periodic meetings 
or seminars can be held within the company and during 
working hours to discuss materials and structures problems . 



Coordination between materials engineers and structural 
designers should begin in the R&D planning stages. Awareness 
of each others thinking from this early point through 
development and application eliminates the pitfalls 
encountered when each discipline just goes its own way. 

* -* * 



6e. What possibilities exist to break out of the ever-increasing 
complexities of our situation? 



The complexities will continue to increase. The only hope 
is to automate as much of the detail as possible and 
integrate all the needed technology into a design system. 
Large digitial computers make progress in this direction 
possible. 



1. Less wasteful practices of the military to buy vast 
numbers of aircraft simulateous with development 
testing. Development testing should be done first 
followed by pre-production prototypes. 

2. Funding for educating engineers in multi-disciplines. 

3. Symposiums similar to that in Monterey to share our 
problems . 

•# * * 

The computer by the storage of data and permiting the 
technique of interactive graphics will help us at early 
design stages to visually and numerically be able to 
determine the effect of a material change on the weight, 
shape, cost, and projected life of our design. 

* -* •* 

None. If continued improvement is desired or required, it 
will get more and more difficult to achieve an improvement, 
More factors become involved as the structural efficiency 
is increased and the structural weight is decreased. 



11-3^ 



I do not agree that the complexities are ever -increasing 
nor that we have a problem of "breaking out" from them. 
Certainly, it is true that viewed in total there are many 
more materials, more environments, more sophistication in 
the determination of design conditions and test evaluation. 
But when you get down to the one for one relationship of 
the individual engineer to the specific job, the situation 
hasn't changed so very much from that of ten or twenty 
years ago. The basic approach is essentially the same. 
I am an optimist about man's ability to cope. 



-* •* 



Just call for and finance new designs and solve problems 
on an orderly, continuous basis. Otherwise they'll 
accumulate to become bigger than we all are together. 



* 



6f. Where do we stand with respect to a clear definition of 
test requirements for structural airworthiness (component and 
full-scale tests for static and dynamic conditions related to 
program development) and for aerodynamic performance (prototype)? 



I'd say in an excellent if not overly burdensome position. 
Prototype per se are not an answer, especially for time- 
dependent phenomena. 

■* * * 

Commercial vehicles built at Douglas have had long life 
without primary structural failure. We have been successful 
because we have relied on thorough analysis, development 
and component testing during the design stage, with large 
component fatigue tests and carefully instrumented flight 
testing to confirm the stresses and expected loads. Full 
scale ground tests are too costly for the limited gain 
expected. We have used five aircraft to develop 1500 hours 
of flight to test the structure and aerodynamic performance. 
Therefore we can conclude that military aircraft could use 
the same techniques and save a considerable amount of 
funds which could be used for new material development. 



I interpret this question to mean: What are the essential 
differences between a structural air-worthiness verification 
program for a production aircraft as compared to a prototype 
aircraft? This is being rather thoroughly debated in 
industry right now. The trend is probably toward reducing 
or eliminating static tests to failure as well as fatigue 
tests for those prototype aircraft where essential purpose 
is aerodynamic and flight research. 



11-35 



ASD-TR-66-57, "Air Force Structural Integrity Program 
Requirements," January 1968 gives the most complete 
description of test requirements for airworthiness 
throughout the life of the vehicle. What is needed is 
more correlation between testing and service experience. 
Testing often does not simulate service experience, 
particularly with regard to environmental exposure . More 
work needs to "be done to properly simulate environmental 
exposure effects in the laboratory so that corrosion 
and fatigue problems can be identified before the aircraft 
gets in service. Lack of budget usually limits the 
amount of service experience correlation that can be done. 

* •* ■# 

Generally clear definition on structure. Systems funding 
should never be committed before prototype evaluations 
except in national emergencies. 

* * * 

Airframe problems that are a result of improper structural 
tests usually can be tied back to an improper definition 
of the environment for design as well as test, i.e., buffet 
loads, fatigue spectra, airload distribution, etc. These 
factors need defining ASAP so that testing of the airframe 
can be done ASAP to minimize the impact of any resulting 
changes . 



7. What other questions are considered pertinent to design 
problems ? 

1. Should load alleviation and mode stabilization be 
further developed to reduce loads? Would the funds 
used for this project reduce failures more than 
spending funds on structure development? 

2. Should more funds be used to develop optimization 
procedures with other disciplines? 

3. Should items (l) and (2) be developed concurrent with 
further structural development? 

•* * * 

a. Should the DOD encourage application of new materials 
offering significant improvement by assuming more of 
the risks? 



11-36 



b. Should the DOD encourage the use of a material like 
composites on a new major weapon system by paying the 
premium for the higher materials cost even when the 
usual cost-benefit analysis does not favor the use of 
the more costly material? 



See the latest NMAB report on Accelerated Use of New 
Materials . 

* * ■* 



11-37 



SECTION III 



SOME BASIC CONSIDERATIONS AND CONCLUSIONS 



SECTION III 
SOME BASIC CONSIDERATIONS AND CONCLUSIONS 

1. SYNOPSIS OF PRESENT SITUATION 



1.1 Introduction 

The objective of this last section is to consider some 
fundamental aspects of design problems and to arrive at practical 
conclusions. 

The present situation regarding structural design of 
aircraft is quite unique. Two different and rather contradictory 
sets of facts are coming together and causing considerable concern. 
On one hand there are recent developments in high-strength alloys, 
filament composites, fracture mechanics, finite element analysis, 
automated design, computer graphics, and other fields which give 
rise to high hopes and expectations for greatly increased 
structural efficiency. On the other hand, there have been 
unexpected difficulties lately regarding detail design and 
materials applicatication. They have occurred in fields which 
nad been considered thoroughly explored and tested -- except for 
some seemingly minor modification which may have caused major 
trouble . 

The result is a disturbing realization that the gaps in our 
understanding of traditional materials and conventional design 
practices are wider than we thought they were. This has dampened 
our previously so confident assurance in tackling new developments 
and has made us wary of extrapolating past experience. In spite 
of an impressive mastery of sophisticated techniques we are 
developing a humble readiness to occupy ourselves with some very 
fundamental considerations about these unexpected difficulties. 

The combination of future promises and present problems 
characterizes the situation in structural design. Promising 
developments of the future depend on finding solutions for 
recent difficulties. The difficulties have been caused basically 
by the increasing complexity of technological developments and 
it is this rather general aspect which assumes specific importance. 

The presentations of Section I give ample evidence for the 
pervasiveness of design problems. The unsolved questions and 
various comments of Section II indicate their range. The basic 
considerations of this Section III will draw attention to the 
fundamental character of design problems and to the need for a 
systematic and practical approach toward their solution. 



III-l 



1.2 Survey of Other Findings Concerned with Present Situation 

During the past two years several highly qualified committees 
addressed themselves to specific aspects of recent problems in 
aircraft structures and in materials application. References 1, 
2, and 3 are closely connected with design problems and it is 
fortunate indeed that they became available Just during the 
final drafting of the present report. 

If it would not be for these references, many of their 
conclusions would have had to be deduced in the present report. 
Since there is good agreement on all essential points, the 
established conclusions of these references may serve as the 
premises for this report which has been somewhat condensed 
accordingly. A brief summary of these references is given below. 

Reference 1 summarizes the lessons learned from structural 
problems in connection with the F-lll development. It states 
and discusses the conclusions of the SAB Ad Hoc Committee on 
the F-lll. The following points are particularly pertinent 
with respect to design problems : 

Application of fracture mechanics as a design tool holds 
great promise. Its present limitations must be recognized 
and a handbook with relevant data and analytical 
techniques should be published and periodically updated. 
Formal "Fracture Control Plans" are recommended as part 
of development programs . 

A damage -tolerant structural concept is considered to be 
an objective of vital importance. Periodic structural 
configuration audits should be accomplished during the 
development program for all primary structure to identify 
all aspects of damage tolerance . 

To exploit the promise of advanced materials while 
minimizing the application risks, research is required to 
formulate effective data collection, storage, extraction 
and presentation methods . A management procedure should 
be instituted to ensure that materials selection, 
processing and manufacturing -- the translation from 
engineering to production — is under a strict control 
and audit schedule . 

Quality assurance and nondestructive inspection must be 
judiciously defined and rigidly applied. Inspectability 
and human engineering factors must be adequately considered, 

Proof -test inspection should not be considered a desirable 
replacement for quality control and nondestructive 
inspection. 



III-2 



• Fatigue testing and analysis are the accepted approach 
for substantiation of fatigue strength. Efforts must be 
directed to the time compression of such tests, and to 
the understanding of fatigue damage mechanism. 

• Accelerated service testing of two or more aircraft from 
an early production lot is recommended. 

• Technology demonstrator programs, such as the Advanced 
Metallic Structure ADP, can strengthen confidence in 
emerging technologies . 

Probabilistic statements about risk can contribute to 
sounder command decisions regarding development and 
utilization of weapon or logistic systems. 

Among the critical problems repeatedly encountered during 
aircraft development and operation, the following are 
listed: 

misapplication of structural material; 

improper material purchase specifications; 

improper manufacturing processes; 

improper detail design with respect to fatigue; 

poor quality control and reliability: 

inadequate auo ing of subcontractors; 

improper analyses and assumptions; 

deficiencies in control and stability; 

change in aircraft usage; 

unanticipated life extensions; 

inappropriate and/or untimely fatigue, static and 

flight testing. 

Regarding organization and human factors , the need for a 
high degree of realism during the procurement process is 
emphasized and more extensive use of independent advisors 
and advisory panels, with real freedom to speak their 
beliefs, is suggested. 

Reference 2 is a report on the Structural Integrity of 
Current and Future Air Force Systems . It is based on an extensive 
team study and the resulting recommendations are directed toward 
most effective planning, execution, and follow-up of the Aircraft 
Structures Integrity Program (AS IP) within the procurement 
agency. Special emphasis is given to the need for making 
realistic estimates, for understanding the impact of trade-off 
decisions, and for bridging communication and apparent technology 
gaps . The benefit which can be derived from independent expert 
teams and review groups is also stressed. 



III-3 



Reference 3 is concerned with the gap which has developed 
between the development of new materials and their application in 
the design of aircraft. It summarizes the findings of the 
National Materials Advisory Board's Committee on Accelerated 
Utilization of New Materials . The committee has investigated 
the causes for the existing delay in the application of 
promising new materials and concluded that it should be possible 
to prevent or minimize delays. Its principal recommendation is 
to establish a continuing function under the auspices of an 
interagency government organization 

to review the status of new materials and processes; 

to identify those with a potential for wide applicability 
which can benefit by coordinated support; 

to organize a cooperative program to assure timely 
application of selected materials. 

These three references serve the purpose to evaluate recent 
difficulties from three clearly defined viewpoints, namely: 
experience with the F-lll program, need for early identification 
of structural problems in Air Force systems , and experience with 
the introduction of new materials . Their recommendations are of 
special significance in view of the knowledge and experience 
brought together in the committees responsible for these reports. 



1.3 Purpose of Present Project 

The present report attempts to take a step beyond the reports 
summarized in the preceding Section 1.2. As this report forms the 
final part of the project Interface of Materials and Structures 
on Airframes, it is based on a more general approach which began 
with Basic Design Considerations (Ref . h) and continued with a 
particular concern about the Decision Process in Structural 
Design (Ref. 5). These general considerations will now "be 
merged with practical aspects . 

The present situation should be quite propitious for 
combining fundamental and practical aspects . There is a growing 
awareness that many of our present problems can be solved by 
improved coordination between the fields of materials, structural 
mechanics, and design. Such coordination represents a natural 
process which has been developing slowly over a long time. 
Much can be accomplished, however, by clarifying and accelerating 
this process. 

The purpose of this report is to evaluate recent design 
experience from an overall viewpoint and to arrive at practical 
conclusions. Such an overall viewpoint will coincide with the 



III-U 



viewpoint of the designer who is responsible for the design. It 
will combine the three separate viewpoints of References 1 to 3 
and will take into account that a practical solution for present 
design problems has to be found first, but also that a basic 
approach toward the solution of future design problems has to be 
established before these problems have grown beyond bounds. 

l.U Method of Approach 

Reference 5 recommends further consideration of the 
structural design process on an industry-wide basis. In a 
modified form and from a somewhat different perspective this was 
accomplished at the Monterey Symposium on Design Problems in 
Aircraft Structures. With its representative participation 
from industry, research, and government agencies, the symposium 
provided a balanced perspective for practical and theoretical 
aspects . 

The proceedings of the symposium, as given in Section I of 
this report, serve to substantiate the essential aspects of 
design problems. The talks were arranged so that considerations 
regarding high-strength steels, fatigue, fracture mechanics, 
fail-safe design, procurement policies, space shuttle, probability 
of failure, risk evaluation, and technology demonstration were 
brought together as basic ingredients of the overall picture. 

Resulting questions and comments, as given in Section II, 
are obviously only a small part of a very wide spectrum. They 
may help to stimulate thoughts and discussions among engineers 
and researchers working in these fields. They certainly have 
contributed toward putting the results of the symposium into a 
wider perspective . 

Discussion and conclusions, as given in this present 
Section III, are rooted both in the fundamental, but somewhat 
generalized, considerations of References h and 5 and in the 
practical aspects expressed during the symposium. The practical 
considerations developed in articles 3^5 to 3 >1 contain the 
essence of the discussion. 

As stated in the foreword, there is no consensus of opinions* 
in this field and the conclusions must not be construed as 
representing the attitude of the Navy Department. 



III-5 



OUTLINE OF DESIGN PROBLENB 



2.1 General Remarks 

A discussion of design problems should properly begin with 
establishing the goal of structural design. In its simplest form 
this can be stated in two words: Optimum structure. The word 
optimum causes, of course, considerable tribulations. Let us 
just keep in mind that it comprises all the compromises which are 
necessary to satisfy the specified strength and stiffness 
requirements as well as considerations of performance, weight, 
cost, risk, time schedule, growth potential, maintainability, 
repairability, inspectability, etc. After some more detailed 
reflection it will be possible in Section 2.6 to arrive at a 
more specific interpretation of design goals. 

Achievement of an optimum structure is still in the distant 
future. Optimization procedures loom at the horizon as a major 
long-range problem and considerable effort is exerted in this 
field. This long-range problem is inseparably connected with the 
short-range problem of avoiding the type of technological 
difficulties which have occurred in the recent past. 

Technological difficulties have existed as long as aircraft 
have been built and they have been overcome reasonably well by 
different methods at different times. One aspect, however, is 
new. Our technology has reached a state of complexity where 
every decision has far-reaching implications. Traditional 
methods become inadequate when a detail design decision can have 
tremendous financial consequences, quite irrespective of safety. 
The cost for rework of a problem of fatigue or stress corrosion 
encountered during the guarantee period of an aircraft may 
exceed the financial resources of a company. 



2.2 Basic Concepts 

For the following discussion it may be helpful to begin 
with establishing two basic concepts which will be used in 
connection with structural design. 

a. Overall Responsibility of Structural Design 

Structural design has many facets. It includes advanced 
design which is concerned with establishing design concepts and 
selecting materials and structural configuration. It also 
includes detail design where the decisions made in advanced 
design are translated into final working details and into full 
substantiation of airworthiness and other considerations. 



III-6 



The full concept of structural design must refer to 
everything connected with the load -transmitting structure. It 
includes selection and optimization of material, structural 
configuration, and design details; substantiation of airworthiness 
by analysis and tests; and determination of weight, cost, 
reliability, fabricability, inspectability, and maintainability. 

This indicates that structural design has to bear full 
responsibility for all the complexities and consequences of 
technology. Anticipation of cost overruns as well as difficulties 
in scheduling, materials processing, fabrication, inspection, and 
maintenance must be included in this responsibility. Much of it 
has previously been left to manufacturing without giving it 
proper representation during the decision-making process. 

This overall responsibility of structural design is the 
clear lesson learned from the experience of recent years. It is 
emphasized here as a basic concept because the full significance 
of this lesson is not yet completely appreciated within the 
engineering community at large. 

b. Team Work within Structural Design 

The second aspect is a consequence of this first 
consideration. It is generally not possible to combine all the 
knowledge, experience, and skill required for a major design 
component in a single individual. A team effort is required 
and responsibilities must be clearly delegated. There are three 
basic fields which are closely related but distinctly separated 
by educational background: 

Materials engineering will be responsible for materials 
properties and characteristics, processing, fabricability, 
inspectability, and maintainability as well as materials 
testing, evaluation, application, maintenance, and 
follow-up procedures, i.e. all aspects of materials 
behavior from cradle to grave; 

Structural mechanics will be responsible for static and 
dynamic analysis with respect to strength and stiffness 
as well as for testing of structural components, i.e. all 
aspects of airworthiness at minimum structural weight; 

Design will be responsible for the traditional field 
of detail design as well as all the considerations of 
overall concepts and trade-offs regarding cost, risk, 
time schedule, and the various "-ilities", i.e. all 
aspects of coordination and optimization. 



III-7 



These three disciplines are integral parts of structural 
design. There will always be some overlap along the border-lines 
of these interrelated fields but we have to realize that 
structural design is an entity and requires a team effort which 
is basically different from adversary confrontation of various 
disciplines. Overall responsibility for guiding this team effort 
must be clearly assigned but, beyond this, each member of the 
team must be aware of his responsibility as an integral part of 
the total effort. The need for such a spirit of common 
responsibility should be recognized as another basic concept. 



2.3 Design Problems in the Realm of Technology 

Having established the basic responsibility of the design 
team for every aspect of structural design, it becomes apparent 
that we cannot be satisfied with considering technological 
problems only. They may serve as a starting point and it will 
be practical to consider basic design problems as falling into 
three groups: those within the realm of technology, those 
beyond the realm of pure technology, and those which are problems 
of policy but have a direct effect upon design. 

Problems of a technical nature may be categorized as 
follows : 

a. Basic Mechanics of Failure 

There is still a fundamental lack of scientific knowledge 
and understanding regarding mechanics of failure. This is 
particularly noticeable in the fields of fracture mechanics, 
fatigue, and stress corrosion. The necessary research must 
take place in the field of materials science and is, although 
of basic importance to structural design, beyond the jurisdiction 
of the design engineer. 

b. Materials Processing, Manufacturing, and Inspection Methods 

Limited knowledge about fundamental aspects of failure 
has been a frequent cause for trespassing unwittingly into 
forbidden zones during processing and manufacturing operations. 
Many typical examples are discussed in Section I and many more 
could easily be found. Quenching rate, residual stresses, hole 
preparation, change of vendor, nondestructive testing methods 
are just some of the potential pitfalls . 

Extensive testing with a great number of parameters is 
usually required. The large test programs on basic fracture 
mechanics data and on spectrum/environmental effects mentioned 
in Section I in the presentation by W. C. Dietz for just one 
aspect of one material indicate the magnitude of the task. The 



III-8 



test data on K lc values shown in the presentation by W. S. Hyler 
indicate how much scatter in test results can occur within 
given specifications and how difficult it may become to interpret 
test data correctly. 

From the perspective of the materials specialist, clearly 
formulated questions about the characteristics of new materials 
can be answered by systematic testing. However, when the 
materials engineer becomes a member of the design team, his 
problem is to recognize all potential modes of failure and to 
anticipate any difficulties which may develop in manufacturing 
and quality control. A multitude of environmental and operational 
conditions, varying from one component to another, result in 
many combinations of temperature, exposure time, stress, 
sequence of cycling, corrosive conditions, etc. Slightly 
modified processing techniques may influence test results 
greatly. All this means that definition of significant test 
conditions as well as evaluation techniques assume major 
importance . 

The sheer magnitude of required materials data for 
application of a new material is immense. The corresponding 
problems are well known to the materials community and further 
detail discussion would go beyond the scope of this report. 

c. Application of Recent Technology and Techniques 

Newly developed high-strength materials and corresponding 
manufacturing technologies pose innumerable problems with hard-to- 
predict consequences regarding crack initiation, hydrogen 
embrittlement , stress corrosion, etc. Typical problems in this 
field are discussed in several presentations of Section I. 

W. H. Sparrow shows illustrative examples for high- 
strength steel parts which failed unexpectedly. J. C. Ekvall 
starts with typical problems in fatigue and surveys the present 
state of the analytical airt, after a decade and a half of 
intensified development in this field. W. C. Dietz presents 
typical problems in fracture mechanics and gives an outline of 
present techniques in this field which has come into its own 
only very recently. 

W. S. Hyler shows another problem which is a typical 
example for the need of full coordination between new technologies 
in materials engineering, structural mechanics, and design. 
Slightest variations in materials processing can cause large 
scatter of K lc values, affecting structural analysis and 
fundamental aspects of fail-safe design. 



III-9 



A large part of the present research effort is directed 
toward further development of fracture mechanics as an important 
analytical tool and toward far-reaching application of damage 
tolerance as a basic concept in structural design. From a wider 
perspective this appears as the latest, but certainly not the 
last, of a long line of design problems which have included 
stress corrosion, fatigue, integral structures, thin-sheet 
design, etc. 

d. Future Technology 

Technological problems of a new type will have to be 
faced in connection with high -temperature applications . The 
presentation by F. F. W. Krohn is concerned with the new tasks 
which may confront us for a space shuttle in the fields of 
materials, structural mechanics, and design. Full evaluation 
of the experience gained with recently developed technology 
and a methodical approach will be a prerequisite. 

There is also the field of filament composites. This 
has not been included in the presentations at the symposium to 
avoid distraction from basic issues. Much specialized development 
work is required but systematic progress is made since the 
importance of this field has been generally recognized. 

Another field which should be mentioned in connection 
with future technology is structural optimization. This aims 
at the very core of the designer's problem, namely how to 
obtain an optimum structure. Considerable difficulties exist 
in the fields of mathematical programing and search methods and 
work proceeds along various lines . 

e. Communication Within a Discipline 

Technological information to keep abreast of newest 
developments is not easily accessible, particularly to the 
engineer fully involved in everyday work. Frequently there is 
a flood of information but research results are published in 
many different places, practical experience of others becomes 
known belatedly and in rather incoherent form, and clear 
conclusions are disseminated only slowly. Competitive 
considerations can also have a retarding influence. 

Although interest in new fields grows rapidly -- e.g. 
fatigue one-and-a-half decades ago, stress corrosion in the 
early 1960's, or fracture mechanics now — it is a slow and 
uneven process to arrive at accepted standards. Again this 
is pointed out quite clearly in several presentations of 
Section I. 



111-10 



W. H. Sparrow illustrates how experience with high- 
strength steels had to be accumulated the hard way. J. C. Ekvall 
shows that determination of fatigue life is not yet an exact 
science, in spite of a tremendous effort spent on it, and 
depends on experience with design details on previous structures 
and with previously established stress levels. W. C. Dietz 
demonstrates in the field of fracture mechanics how practical 
experience is related to the development of new methods. 
W. S. Hyler indicates the role of experience in materials processing 
Everywhere costly mistakes could be avoided if systematic 
information about previous experience were available. 

The problem of communication within a technical 
discipline can be considered to be a technological problem. 
However, it becomes apparent that it cannot be solved on the 
level of the engineering specialist. This points toward 
further problems which are beyond the realm of pure technology. 



2 .k Design Problems Beyond the Realm of Pure Technology 

The preceding problems in the realm of technology can be 
recognized easily. There are other problems, however, which 
are not so clearly visible and which are in an ill-defined 
region beyond pure technology. They may be considered in the 
following problem areas : 

a. Communication Between Technical Disciplines 

The difficulties of communicating and keeping informed 
within one's own technical discipline were discussed under 2.3e. 
Design, however, involves several different disciplines and 
communication between them can become formidably difficult. 
Materials engineering, structural mechanics, and design are 
closely related but the difficulty of communication between them 
is emphasized strongly in several of the presentations in 
Section I. 

W. H. Sparrow concludes his presentations about 
high-strength steels with the warning that there is no substitute 
for communication between engineering, manufacturing, and quality 
control. W. C. Dietz shows throughout his considerations about 
fracture mechanics how important it is to have a full exchange 
of specific information between engineering, manufacturing, 
and inspection. W. S. Hyler begins his comments on fail-safe 
design with pointing out the need for interrelation between 
designer and materials and process engineers for any consideration 
of fracture and fatigue crack propagation. 



III-ll 



Considering a large number of typical design problems, 
it appears that a lack of inter-disciplinary communication can 
be found either as their cause or, at least, as a contributing 
factor. Such a statement should not be taken lightly. It 
indicates that no amount of specialization can solve present 
design problems unless the specialist develops an understanding 
for interaction between his field and others . 

There is nothing new about the need for communication. 
It has existed as long as there has been specialization. Some 
progress has certainly taken place during the last decade but 
this has not overcome the basic fact that different specialists 
have different concerns and are not readily aware of each 
other's line of thinking. Team work in structural design requires 
full communication and understanding of interaction to a degree 
which has not yet been developed. 

b. Risk Evaluation 

New developments in aircraft structures involve some 
risk. There is, first of all, the technological uncertainty 
which may be expressed as probability of failure. Beyond this, 
there is the risk of exceeding cost estimates and time schedules. 
All these aspects have to be incorporated in risk evaluation. 
Some of the inherent difficulties are discussed in Section I. 

J. E. Fischler substantiates an approach to use 
probability of structural failure for the comparison of 
different designs. Developments in fracture mechanics make 
it possible to determine the probability of structural failure 
as a function of several design parameters. This concept can 
be used to compare materials and types of construction on an 
equal basis . 

W. E. Ellis indicates some tentative steps toward 
overall risk evaluation. Much can be learned from procedures 
which have been developed by operations analysis . It will be 
a long way to transform this into a useful tool for structural 
design. A first and very important step is to make the 
designer aware of the line of systematic thinking which has 
been developed in this field and which can supplement his 
engineering techniques in a significant way. Beyond this, 
much thought will have to be given to the expression of 
qualitative considerations in quantitative terms. 

c. Ideas and Decisions 

Each of the problems in structural design provokes 
creative thinking and, because several solutions are generally 
possible, calls for decisions in the face of uncertainties. This 
requires a systematic approach in addition to technological 
expertise. Structural design has to be rooted in both technology 
and methodology. 



111-12 



The decision process in structural design has been 
discussed in Ref. 5. It will have to be based on analytical 
models which combine the considerations of airworthiness and 
optimization. Models as shown in Fig. B-l, C-23, and C-2U 
are typical steps in this direction. Development of such a 
decision process is the goal of design methodology. It has to 
take place in step with the solution of the other basic problems 
in structural design. A fundamental need is to provide 
visibility and clarity for any design decision. 



2.5 Policy Problems Affecting Design 

The design problems discussed on the preceding pages give 
full regard to the engineering viewpoint even if some of them 
cannot be solved on an engineering level. There are other 
decisions, however, which may have a far-reaching influence on 
structural design without giving full cognizance to the 
engineering viewpoint. These decisions are usually made on a 
management level where engineering considerations are balanced 
against various aspects of funding, timing, and general policy. 
The two subjects of procurement policies and test programs 
deserve special attention. 

a. Procurement Policies 

Procurement policies can have a pronounced effect on 
structural design and the following discussion may help to 
clarify some of these aspects. 

Throughout the 1960's there was a trend toward 
increasingly rigid contracts at fixed price, with a total 
package incorporating R&D, production, performance, and time 
schedule. It was only around 1970 when some fallacious 
reasoning in this trend became apparent. Procurement agencies 
were driving toward unrealistic requirements to obtain maximum 
performance while bidders made over-optimistic estimates in 
a highly competitive environment. The risks of an evolving 
technology were not appreciated properly and resulted in huge 
R&D and production costs to solve unforeseen difficulties . 
Structural design conservatism was frequently squeezed thin 
between technological need and available budget or had its 
flexibility for trade-offs curtailed by detail specifications. 

Important lessons have come from this type of 
experience and may be summarized as follows: 

Pre -contractual assessments require a high degree of 
technological and budgetary realism and objectivity 
both on the side of procurement agency and prospective 
contractor. 



III-13 



Weight, cost, and time schedule require closest 
control but should not be considered as fixed 
quantities by themselves . They are means to an 
end and trade-offs must be encouraged in order to 
obtain an optimum total design rather than a 
design which is best in one aspect at the expense 
of another. 

Costs for research and development must be separated 
from production costs and only the latter should be 
subjected to penalties and incentives. 

P^re -production concept proof should be emphasized 
but it must be realized that this is not as simple 
as just having a prototype . Time dependent 
phenomena or production feasibility will not 
necessarily be demonstrated in prototype programs. 

In case there is concurrent development and 
production, the inherent risks must be clearly 
recognized and accepted. 

Due to this interrelation between structural design and 
procurement policies, future policy developments will have to be 
scrutinized closely. The trend seems to be toward providing a 
cost-reimbursement basis for the R&D phase with its higher risk 
and a fixed -price-incentive basis for the production phase. Much 
emphasis will be on milestone demonstrations as well as cost and 
schedule control. 

The future will probably also hold a shrinking market 
with fewer but more sophisticated projects. There might be a 
recourse to prototypes by several companies with the production 
contract awarded after technical demonstration. 

From the viewpoint of structural design there are two 
particularly important aspects to be emphasized in connection with 
any procurement policy: 

The risk of new technological developments must be 
clearly recognized and potential difficulties should 
not come as surprises; 

The program for structural testing has to be prepared 
carefully not only with respect to its scope but 
also with respect to its timing. 

Risk evaluation and a reliable estimate of structural 
testing requirements are important aspects of a procurement 
policy. 



III-lU 



b. Structural Testing 

Structural testing is, of course, a straightforward 
engineering function. However, it is listed as a policy problem 
because the necessary funding depends on policy decisions. 
There will always have to be a compromise between the engineer's 
desire for verification by testing and the manager's reluctance 
to provide the considerable funding. Merely to set up a 
consistent program for structural tests has become a task which 
is not easily done. There is also the additional aspect that 
expenditures for testing can result in significant economic 
advantages in production as well as in improved analytical 
techniques . 

Structural testing is an integral part of structural 
design. With increasing emphasis on environmental conditions 
and on damage tolerance the amount of testing can grow excessively. 
Yet new technology requires extensive and systematic testing in 
order to reduce risk to an acceptable level. Agreement between 
analysis and experiment is the basis for our confidence in the 
integrity of a structure. Much engineering work is still 
required to obtain closer coordination between analysis and 
testing and to save time, money, and talent as analysis may 
eliminate some testing. 

Component testing is caught in the dilemma that it can 
start only after the component has been designed and manufactured 
but that tests should be finished in time so that any modifications 
do not affect detrimentally the production process. In view of 
the advanced planning necessary for production, the necessary 
compromises have to be made within the context of overall policy. 

Within this need for compromising, the structures 
engineer has the responsibility to recognize clearly what kind 
of information is required from what type of testing in order to 
substantiate airworthiness for all operational and environmental 
conditions. In this connection it is important to evaluate 
critically any prototype testing to see whether it plays an 
important role within the structural testing program. 

These few remarks may suffice to show the close 
relationship which has to be established between structures 
engineering and management, starting at the very beginning of a 
project. Recent experience has shown that funding is always 
available if a panic situation should develop but that a 
comparatively small amount of expenditure may prevent such a 
situation. 



111-15 



2.6 Designer's Viewpoint 

The preceding broad-brush treatment of problem areas in 
structural design requires, of course, much amplification before 
it can be thoroughly interpreted. However, in spite of its 
briefness, it seems to be adequate for the purpose of clarifying 
and illustrating two basic aspects: Design problems are 
rooted in technological difficulties but they branch out into 
wider fields . 

After having considered separately those problems which are 
of a specialized technological nature and those which go beyond 
pure technology, it can be recognized that each of these two 
categories has several facets. Instead of expressing the two 
categories as a field of technology versus another field which 
is beyond pure technology, we may think in terms of airworthiness 
versus optimization, or specialization versus interrelation. 
Each of these terms implies a different aspect, and indicates 
the many-sidedness of each type of problems. 

Technological problems can be seen clearly after they have 
developed -- even if it may take a post mortem in extreme cases. 
These problems have a direct effect on airworthiness, and steps 
toward their solution are taken quickly. Responsibilities 
are distributed among well-defined disciplines or organizational 
groups and steady progress toward the solution of these problems 
can be expected. There seems to be no need to pursue them any 
further in this report. 

The picture is different, however, for problems reaching 
beyond the realm of pure technology. These problems are 
ill-defined and at the same time most pervasive and elusive. 
Although they are of a very different kind than the well-defined 
technological problems, they may easily permeate any of them. 
It became quite apparent from the presentations by W. H. Sparrow, 
W. C. Dietz, and W. S. Hyler how the work of the specialist 
has to be correlated with inter-disciplinary communication, 
and from the presentations by J. E. Fischler and J. W. Ellis how 
probability of failure and risk evaluation may play an important 
role in the decision making process. This type of problems 
reaching beyond traditional aspects of pure technology forms 
the very essence of design because design cannot be satisfied 
with just finding a technical solution. It has to strive for 
an optimum solution considering all circumstances. 

Much general consideration has recently been given to 
design objectives and methods. Several books on design 
methodology have been published just during the past few 
years. Combining quotations from various authors, design must 
not be confused with art or science or a form of mathematics, 



111-16 



but it is a hybrid activity which depends on a proper blending of 
all three. We may think of it as a creative, goal -directed, 
problem- solving activity which depends on decision-making in 
the face of uncertainties and is concerned with all aspects of 
a problem. 

Although design is a creative activity, it has to submit 
itself to a rigorous logic . A large number of ideas have to be 
analyzed and evaluated. Both the climate for encouraging 
new ideas and the decision process for evaluating the implications 
of these new ideas have assumed much importance for the solution 
of design problems. 

This kind of considerations leads toward a more specific 
interpretation for the goal of structural design. The dual 
nature of design problems -- specialized technology versus 
optimization and complex interrelations -- is of basic significance, 
We have to realize that most of our present design problems can 
no longer be solved by merely concentrating our efforts on 
specialized technology. Full consideration must also be given 
to the interrelation between specialized fields -- both for 
optimization and for assurance that no aspect of airworthiness 
has been overlooked. 

The following part considers these not so obvious yet 
completely vital aspects of structural design problems. 



111-17 



3- UNEXPLORED ASPECTS OF DESIGN PROBLEMS 

3-1 General Remarks 

The preceding consideration of design problems from the 
designer's viewpoint drew attention to the need for going beyond 
the traditional concern about specialized technology. Four 
questions have particular significance: 

How can communication between various specialists be 
improved? 

How can risk be evaluated? 

How can the decision-making process be clarified? 

How can available information be made more accessible? 

Answers to these questions cannot be found on a purely 
technological level. They require full consideration of 
educational and organizational aspects. 

The recurrent theme contained in these questions is the need 
for providing full communication and mutual understanding among 
all members of the design team. This would sound like a self- 
evident and rather superfluous statement if it were not for the 
overwhelming evidence -- expressed in many presentations of 
Section I -- which indicates how hard it is to accomplish this 
communication and mutual understanding and how close this point 
comes to the roots of many or even most of our design problems. 

The implications of this statement must be recognized. The 
problem cannot be solved on paper. The objective will be 
accomplished only when attitudes and actions of engineers express 
that they think not as engineering specialists but in terms of 
the overall design project. Our conventional engineering 
education has not prepared us to do this . 

The following considerations begin with some basic aspects 
of engineering curricula, continuing engineering education, and 
engineering professionalism and lead up to a practical approach 
regarding some very fundamental features of design problems. 

3-2 Engineering Curricula 

Engineering is based on science and the scientific approach 
consists of analytical and experimental techniques applied to a 
clearly defined problem. Education along these scientific methods 
has been the foundation for outstanding technological achievements 



111-18 



but unfortunately also for a lack of success in translating 
specific achievements into an overall entity. We have to learn 
how to coordinate diverse and frequently contradictory 
requirements from many specialized subjects in order to obtain 
an optimum overall design. 

Contrary to the typical problems in science, a typical 
design problem is not clearly defined. It is the responsibility 
of the designer to recognize all significant parameters and to 
define the problem before the scientific process of analyzing 
it can take place. Usually there are several possible solutions 
and the designer is responsible for determining which is the 
optimum among them. Our engineering education in the past two 
decades has done exceedingly well in preparing the student for 
the analytical process of problem solution but has usually 
neglected the basic design aspects of problem definition and 
optimization. 

A first reaction to the long trend of putting so much 
emphasis on scientific specialization became visible in the 
mid-1960 's. Development of inter-disciplinary graduate courses 
was pioneered at MIT and Stanford and sponsored further by NASA. 
These courses contributed much to a growing awareness for the 
need of coordination between academic disciplines. 

However, these courses were mostly oriented toward advanced 
design concepts and were not particularly concerned with the 
interaction of materials and structures, i.e. structural design. 
The AIAA Aircraft Design Committee, among others, has been 
concerned about the vanishing of design from aero curricula 
(Ref. f ) . More than a quarter of the aero curricula seem to have 
eliminated design courses entirely and only less than a quarter 
devote a minimum of four semester hours to design. The need for 
much closer cooperation between aerospace industry and academic 
community is an obvious conclusion. 

The present situation is full of contradictions. In spite 
of the gloomy outlook of Ref. ^, there are a good number of very 
promising starting points. Most of them are outside the field 
of aeronautics . The Engineering Development Program of the 
University of California at Los Angeles and the methods of Case 
Studies as developed at Stanford and Berkeley may be mentioned. 

There seem to be three basic difficulties in making our 
engineering curricula responsive to the needs of industry. 
Firstly, any changes in academic life take place at a slow pace. 
Secondly, no clear guidelines have been established in spite of 
much interest in the subject. Thirdly, particularly in aeronautics, 
engineers in industry and faculty members at universities who may 
have recognized the problem have been absorbed so much in their 
specialized fields that they did not find the necessary time to 
do something about this overall aspect. 



IH-19 



3-3 Continuing Engineering Education 

Continuing education is based on the recognition that 
technological developments make an engineering education 
obsolescent after not too many years. Evening classes, 
company-sponsored courses during working time, lectures and 
meetings, and full-time short courses are most frequently used 
to keep the engineer abreast of technological developments . 

All this is being done in many modifications , depending 
very much on special conditions. The spectrum ranges from 
courses which are given to obtain an academic degree to others 
which are directed purely toward professional development. 
Educational media are increasingly employed. 

For instance, in regions where several aerospace companies 
and a university are located, lectures given on the campus are 
brought by TV into classrooms inside the companies during 
working hours, frequently with two-way communication between 
instructor and each individual. Such courses are usually part 
of an academic curriculum and they are taken for credits as the 
participants are working toward an advanced degree. 

On the other hand, full-time courses which may extend over 
a few days or a few weeks usually are directed toward a subject 
of specific professional interest of an advanced nature. Such 
courses are given by a group of specialists from a viewpoint of 
sharing information on recent developments and the participants 
are experienced engineers . The purpose is clearly continuing 
education in the basic sense of the concept. 

It can easily be seen that there is much more flexibility 
in this type of continuing education than in courses which are 
part of a formal curriculum. Academic credits lose their 
significance, courses can be tailored much more readily to the 
needs of industry or special developments, and instructors may 
be chosen in accordance with their specific competence, whether 
they come from faculties or industry or research. 

Mich is going on in this field of continuing education. 
Contrary to the slow changes in well-established curricula of 
formal engineering education, everything in continuing education 
is in a fluid state of early development. There seems to be a 
unique opportunity to apply some of these developments in the 
educational field to the design problems in aircraft structures. 



3.*+ Engineering Professionalism 

Engineering graduates going into industry have generally a 
clear analytical mind and are well equipped to solve problems in 
their fields of specialization. Some of these young engineers 
are outstandingly bright but the process of advancing into 



111-20 



engineering positions with broad responsibilities is slow and 
tedious. Usually it takes the route of proving their excellence 
in a specialized field and even for capable engineers it may take 
a good deal more than 10 years before they begin to develop an 
understanding for the all-important interrelation between 
various fields of specialization. Basically they are left to 
their own devices how they go about it. 

There are, of course, highly competent engineers in industry 
with many years of experience who have grown beyond their field 
of specialization and have learned to communicate with adjacent 
disciplines. They are in responsible positions and generally 
overburdened with work. However, for each one of them there are 
a great number of others in various stages of the laborious 
process of trying to accumulate some experience beyond their 
own field in order to broaden their horizon and make them more 
valuable engineers . Would it not be plain common sense to help 
accelerate this important process? 

Let us look just at the typical problem of communication 
between structures and materials engineers . When the structures 
engineer obtains a K-l c value from the materials engineer, does he 
understand sampling techniques and processing tolerances on which 
the value is based? When the materials engineer proposes the use 
of a certain heat treat, does he understand all implications with 
respect to residual stresses, environmental and operational 
conditions which may occur? The specialist's jargon frequently 
expresses concepts which are not readily explained. Any attempts 
of explanation may leave some essential detail misunderstood due 
to differences in viewpoints and lines of thinking. Section I 
contains many examples for design problems which were caused by 
this type of difficulties in communication between specialists. 

Management is, of course, vitally interested in improving 
this communication process. So is the individual engineer 
because it enhances his professional growth. No engineer worth 
his keep wants to be just a cog in a big machine. A pure 
specialist might too easily be unemployable as soon as he becomes 
unemployed . 



3.5 Practical Considerations about Educational Aspects 

We have seen that our background has preconditioned us to 
think as specialists rather than in terms of an overall design 
project. On the other hand there is a basic need and a willing 
readiness to recognize the role of complex interactions and to 
develop new methods to deal with them. The question is how this 
can be accomplished. 



111-21 



An answer can be found along the lines of continuing education. 
There is, however, one important qualification. Contrary to 
other fields, the subject matter for this type of course cannot 
be prepared on the usual academic level of research and special- 
ization. It rather has to be based on a systematic evaluation of 
recent practical experience in structural design. 

This experience has been gathered in industry. Therefore, 
the subject has to be evaluated and written up by engineers 
who are thoroughly familiar with recent design experience and 
who understand the far-reaching implications. Nothing of this 
sort has been done because it requires a joint effort of 
considerable magnitude to prepare the necessary outline, 
subject matter, and text material for such a course. 

If the development of a course program and the corresponding 
text material is sponsored by a government contract, it can be 
done thoroughly and will be available to the entire industry. 
The project is too important and too urgent to be left to a 
somewhat haphazard development on local levels . On the other 
hand, when a well-prepared text is available, qualified instructors 
can be found locally. The text will serve as the backbone for 
courses or seminars in continuing engineering education given 
throughout, industry or for self -study. A loose-leaf textbook --to 
be kept up to date -- may be the most practical format, but some 
other methods of educational media may deserve consideration. 

Such a project requires the cooperation of several highly 
competent and motivated engineers . It will have to start with 
systematically extracting, describing, documenting, and evaluating 
recent experience. This has to be translated into a form which 
can serve to prevent repetition of past mistakes, to provide an 
introduction into newly developing fields , and to direct 
attention toward new methods . 

It appears that a course subject of prime importance will be 
Interaction of Materials and Structures. This requires the 
cooperation of engineers experienced in structural design, 
analytical methods, materials characteristics, processing methods, 
manufacturing, and inspection. Improved communication and 
interrelation between specialists is based on a basic understanding 
of underlying principles, applicable methods, significant aspects, 
and recent developments in adjacent fields. This means familiarity 
with each other's outlook and methods of approach. Each 
specialist has to put himself into the shoes of other specialists 
and has to explain to them some basic aspects of his own field. 
Emphasis will be on those aspects which have contributed to 
recent design problems and which have to be understood by other 
members of the design team to prevent interface problems. The 
text can be tailored to the practical needs of a design team. It 
is this type of information which is not available in a systematic 
form anywhere and which could prevent a large percentage of our 
typical design problems. 



111-22 



Other course subjects should be Risk Evaluation and Decision 
Making. These are strange fields to most engineers and communication 
is correspondingly aggravated. In the past these subjects have 
not played any role in design -- except in some of the more 
abstract aspects of parametric performance studies and fatigue . 
With increasing complexities a close interaction between engineering 
and operations analysis, extending to the level of structural 
design, cannot be avoided much longer. The engineer needs an 
introduction into operations analysis from his design viewpoint 
to familiarize himself with basic possibilities, methods of 
approach, and practical applications so that he does not violate 
elementary rules and can recognize when there is need for 
specialist advice. New possibilities can be explored only if 
the engineer is able to communicate with the operations analyst. 

Such educational efforts within the framework of continuing 
engineering education are aimed at engineers in industry. This 
should produce early results. For thorough results, however, the 
aeronautical curricula at universities have to be affected and 
closer coordination between universities and industry is 
necessary. It seems that an advisory group made up of members 
from universities and from industry could exert a very healthy 
influence. Such a group would have to provide guidelines 
firstly to bridge the gap which has developed between engineering 
needs of industry and scientific orientation of engineering 
curricula, and secondly to coordinate the efforts which will be 
required in the field of continuing education. Many implications 
are involved and the full spectrum of education for aerospace 
engineers must be taken into account. University curricula and 
continuing education have to be coordinated as two fields of 
fundamental importance. 



3.6 Practical Considerations about Information Systems 

After having considered educational aspects as a prerequisite 
for dealing with complex interactions, attention must also be paid 
to another side of the problem. Much waste and frustration occur 
when basically available information is not accessible for 
secondary reasons — which is a problem of organization. 

This has been a particularly blatant problem in the field 
of materials characteristics where test data are produced in 
many places but become meaningful only when correlated with other 
data and evaluated with respect to clearly specified test 
conditions . The enormous quantity of data being developed makes 
it imperative to have clearly assigned responsibilities for 
collecting, interpreting, storing, and disseminating this 
information. 



III-23 



The Defense Metals Information Center of Battelle represents 
a basic step in this direction. It incorporates the essential 
capability to interpret and evaluate data. Yet much additional 
effort and funding are required. Decentralization in accordance 
with available talent and facilities is quite feasible. Reference 
discusses some aspects of a materials information system and 
computerized methods, and Ref. 1 also points out the need in 
this field. The dominant need for such centers of information as 
a prerequisite for a healthy aerospace industry is easily 
apparent and it can only be hoped that the various obstacles will 
be overcome in the near future. 

Reference 2 recommends an analogous step by establishing a 
Structures Information and Analysis Center "to collect, process, 
investigate, analyze, evaluate, disseminate and advise on 
structural materials applications, analysis methods, test 
techniques and failure modes and causes". 

Another step consists of handbooks containing up-to-date 
techniques in newly developing fields . Most major aircraft 
companies have developed manuals of this type in fields like 
fatigue and stress corrosion. Much duplication of efforts could 
be avoided and more complete information and consistent 
application could be assured if such handbooks would be sponsored 
on an industry-wide basis. Reference 1 recommends particularly 
such a Handbook on Fracture Mechanics for Aircraft Designers, 
periodically updated as new data and experience become available. 

A further step should consist of Case Studies, describing 
the full history of significant failures and design problems 
which have been encountered and solved in industry. At present 
this is done to a certain extent. Important failures result in 
engineering reports which have restricted circulation. Some 
basic aspects eventually are filtered into technical papers or 
articles. Other aspects enter into a grapevine system. However, 
the full information should be available to all those engineers 
who may learn from this experience to avoid similar mistakes. To 
write up a comprehensive case study is a major task and Reference ' 
discusses the practical aspect of having this done by graduate 
students who can gain much insight into design problems by 
doing this . 



3.7 Practical Considerations about Overall Perspective 

The preceding considerations about educational aspects and 
information systems indicate some stimulating and far-reaching 
possibilities which are completely within practical reach. Their 
realization, however, requires an effort which can be exerted 
only when the importance of design problems is considered from a 
long-range perspective . 



III-2U 



Such a long-range viewpoint has been taken for the USAF 
program on Advanced Metallic Structures. This program is 
described by D. A. Shinn in his presentation in Section I. It 
is based on the recognition that there is no systematic approach 
available to structural design at increasing complexities and 
the program is directed toward establishing a practical approach 
for solving inherent problems. Special emphasis is given to an 
efficient system for distributing the resulting information to 
the entire technical community. 

This large-scale effort toward finding practical solutions 
for design problems represents an important step. To be fully 
effective, however, it must be coordinated with a corresponding 
effort regarding fundamental considerations. An educational 
program as outlined under 3.5 can prepare the ground for such 
fundamental considerations and an information system as outlined 
under 3.6 can remove basic obstacles in the path of solving 
design problems. 

It appears that an educational program and an information 
system along these lines should be considered as an important 
complement to the program on Advanced Metallic Structures. Such 
an approach toward three esential aspects -- hardware, software, 
and education -- together with the well-recognized need for 
technological research and development, would provide a logical 
and promising course of action. The necessary funding for the 
data information system will be considerable but can be spread 
out -- besides, any delay will increase the eventual costs. For 
all other aspects suggested under 3*5 and 3*6 only modest funding 
is required. 

As an encouraging omen it may be mentioned that the USAF 
program on Advanced Metallic Structures as well as the report 
by the Ad Hoc Committee (Ref . l) and the Study of Aircraft 
Structural Integrity (Ref. 2) emphasize the need for a systematic 
exchange of information. This recognizes an attitude that problems 
of structural design should be beyond competitive considerations. 
A structural failure in one aircraft hurts all others . The whole 
aircraft industry is in the same boat and everybody suffers when 
somebody contributes to a leak. 

Technical competitions will be governed by the quality of a 
design team and the corresponding probability for a successful 
design. To build such a team, to provide stimulating working 
conditions, and to instill a creative spirit takes a long time. 
Technological expertise can be acquired by hiring a few experts. 
However, an awareness of complex interactions must be developed 
methodically and still begs to be recognized as a fundamental 
aspect of design problems. 

* * * 



III-25 



There is an additional aspect which is worth mentioning as 
we are looking at design problems from an overall perspective. 
An important clue is provided in the conclusions of both the 
Ad Hoc Committee (Ref . l) and the Study of Aircraft Structural 
Integrity (Ref. 2) which emphasize very strongly the benefit 
which can be derived from an independent group of experts . 

Independent of these conclusions, the principal recommendation 
of the Committee on the Accelerated Utilization of New Materials 
(Ref. 3) consists of having the function of such a group in the 
field of materials. Correspondingly, practical considerations 
about educational aspects (see art. 3. 5 of this section) point 
toward an advisory group in the field of engineering education. 
Some "practical" people may say that this is wishful thinking. 
Yet in the field of aircraft structural integrity, where 
considerable obstacles of a competitive nature had to be 
overcome, an industry-wide group of experts has been in existence 
for a decade, steadily growing in importance. 

Much grief, frustration, and waste could be avoided and 
many potential design problems of the future could be prevented 
from developing if advisory groups of independent experts would 
be available to provide guidance in the three fields of materials, 
structural integrity, and engineering education. Even if this 
is only an advisory function, such panels can exert a great 
influence if they represent the proper blend of experience, 
realism, and vision. 



111-26 



k. CONCLUSIONS 



Design problems in aircraft structures can be considered as 
belonging to three different categories. Each of them requires 
efforts of a distinct kind in order to solve present and future 
difficulties . 

a. Technology 

On a technological level additional research and 
development is required regarding 

• mechanisms of failure, particularly in the fields of 
fracture mechanics and fatigue; 

• materials processing, manufacturing, and inspection 
methods ; 

application of recent and future technology and 
techniques . 

Necessary efforts in these fields have been generally 
recognized and identified. Systematic progress will depend on 
the available funding for research and development programs 
which have been outlined by cognizant agencies. 

b. Technological Organization 

Some basic design problems cannot be solved on an 
engineering level. They require an organizational effort by 
government agencies and top management. 

Data information systems are necessary to assure that 
available data become accessible to the engineering 
community. DMTC of Battelle and the planned 
Structures Information and Analysis Center of AFFDL 
are steps in this direction but much additional 
effort is required (see 3-6). 

Handbooks containing up-to-date techniques in newly 
developing fields should be sponsored on an industry- 
wide basis (see 3-6). 

Case Studies investigating all aspects of recent 
failures and complex problems which have been solved 
can serve as lessons to be learned by the industry. 
They should be written up systematically and 
circulated widely (see 3-6). 



111-27 



Interaction between management decisions and design 
problems requires particularly close attention in 
the fields of procurement and structural testing 
(see 2.5) . 

c . Engineering Education 

On an educational level the engineer needs a helping 
hand to grow beyond his field of specialization and to understand 
the complexities of technological problems . 

Early results can be achieved by sponsoring the 
development of a course outline and corresponding 
text material on Rroblems of Interaction between 
Materials and Structures . 

General awareness of unfolding new possibilities can 
be stimulated by additionally sponsoring the 
development of a course outline and corresponding 
text material on Risk Evaluation and Decision Making 
in Engineering. 

long-range results can be influenced by an advisory 
group on engineering education, representing both 
industry and universities. 

Details are discussed under 3^5- The funding required 
for this educational effort is small compared to the large 
effect it will have on the engineer's approach to complex design 
problems . 



Design problems can be prevented before they develop. This 
requires an approach where individual engineer, engineering 
community, management, and government agencies have to pool 
their competence, resources, and initiative. Increasing 
complexities may result either in challenging tasks which can 
be met or in frightful nightmares which are hopelessly entangled. 
There still seems to be a promising opportunity to influence 
these developments . 



III-28 



REFERENCES 



1. SAB Ad Hoc Committee on the F-lll, Special Report on Lessons 
Learned from the F-lll Structural Experience , August 1971. 

2. USAF-ASD, Report on the Study of Aircraft Structural Integrity 
of Current and Future Air Force Systems , July 1971. 

3. National Materials Advisory Board, Accelerating Utilization 
of New Materials , NMAB-283, May 1971 

h. U. Haupt, "Interface of Materials and Structures on Airframes, 
Part 1, Basic Design Considerations", Naval Postgraduate 
School, NP3-57HP9111A, November 1969. 

5. U. Haupt, "Interface of Materials and Structures on Airframes, 
Part 2, Outline of Decision Process in Structural Design", 
Naval Postgraduate School, NPS-57HP0121A, December 1970. 

6. J. M. Swihart, "Need for Design Classes in Aerospace 
Engineering Schools", Astronautics and Aeronautics, June 1971. 



III-29 



DISTRIBUTION LIST 

Defense Documentation Center 20 

Naval Air Systems Command, Washington, D. C. 10 

Library, Code 0212, Naval Postgraduate School, 2 
Monterey, California 

M. U. Clauser, NP3, Monterey, California 1 

Ulrich Haupt, Code 57Hp, NP3, Monterey, California 10 

J. W. Mar, Department of Defense, Washington, D. C. 1 

Jerome Persh, Department of Defense, Washington, D.C. 1 

W. P Raney, Department of the Navy, Washington, D.C. 1 

Nathaniel E. Promisel, Materials Advisory Board, 1 
Washington, D. C. 

Joseph R. Lane, Materials Advisory Board, Washington, D. C. 1 

A. M. Blamphin, Materials Advisory Board, Washington, D.C. 1 

B. A. Kornhauser, Materials Advisory Board, 1 

Washington, D. C. 

Richard R. Heldenfels, NASA Langley Research Center 2 

H. F. Hardrath, NASA Langley Research Center 1 

Joseph Maltz, NASA, Washington, D. C. 1 

George Deutsch, NASA, Washington, D. C. 1 

James Gangler, NASA, Washington, D. C. 1 

J. F. Blumrich, NASA, NSC, Hunts ville, Alabama 1 

William A. Langen, NADC Johns ville, Warminster, Pa. 2 

Murrary S. Rosenfeld, NADC Johns ville, Warminster, Pa. 2 

E. J. McQuillen, NADC Johns ville, Warminster, Pa. 2 

Jerome Persh, Department of Defense, Washington, D. C. 1 

Walter S. Hyler, Battelle Memorial Institute, Columbus, 
Ohio 

111-30 



Col. G. P. Haviland, ASD, WPAFB, Ohio 1 

Walter J. Crichlow, ASD, WPAFB, Ohio 1 

R. F. Hoener, AFFDL, WPAFB, Ohio 1 

H. B. Lowndes, AFFDL, WPAFB, Ohio 1 

J. R. Johnson, AFFDL, WPAFB, Ohio 1 

Keith I. Collier, AFFDL, WPAFB, Ohio 1 

Vince Russo, AFFDL, WPAFB, Ohio 1 

A. M. Lovelace, AFML, WPAFB, Ohio 1 
S. W. Tsai, AFML, WPAFB, Ohio 1 
George Peterson, AFML, WPAFB, Ohio 1 
Donald A. Shinn, AFML, WPAFB, Ohio 2 
W. P. Conrardy, WPAFB, Ohio 1 
Walter Trapp, AFML, WPAFB, Ohio 1 

B. R. Erarich, AFML, WPAFB, Ohio 1 

Don Forney, AFML, WPAFB, Ohio 1 

Stuart Arnold, Army Materials & Mechanics Research 1 
Center, Watertown Massachusetts 

R. S. Berrisford, Army Aviation Materials Laboratories 1 
Fort Eustis, Virginia 

Holt Ashley, University of Maryland, College Park, Md. 1 

Henry 0. Fachs, Stanford University, Palo Alto, Calif. 1 

Koichi Masubuchi, MIT, Cambridge, Massachusetts 1 

Earl R. Parker, University of California, Berkeley, Ca. 1 

P. F. Packman, Vanderbilt University, Nashville, Tenn. 1 

B. R. Noton, Washington University, St. Louis, Mo. 1 

R. C. Carlston, California State Polytechnic College 1 
San Luis Obispo, California 



111-31 



Alan V. Levy, Aerojet General Corp., Sacramento, Calif. 1 

George C. Martin, Boeing Co., Seattle, Washington 1 

Paul Sandoz, Boeing Co., Seattle, Washington 1 

William Gray, Boeing Co., Seattle, Washington 1 

John M. Swihart, Boeing Co., Seattle, Washington 1 

George E. Hughes, Boeing Co., Seattle, Washington 1 

Donald E. Strand, Boeing Co., Seattle, Washington 1 

Ken Dickenson, Boeing Co., Seattle, Washington 2 

R. A. Davis, Boeing Co., Seattle, Washington 1 

R. E. Pearson, Boeing Co., Seattle, Washington 1 

J. K. Fuller, Boeing Co., Seattle, Washington 1 

John Arnquist, Boeing Co., Seattle, Washington 1 

D. B. King, Brush Beryllium Company, Cleveland, Ohio 1 

Julian Glasser, Chemical and Metallurgical Research Inc., 1 
Chattanooga, Tennessee 

Wolfgang Steurer, General Dynamics, San Diego, California 1 

H. F. Rogers, General Dynamics, San Diego, California 1 

W. W. F. Krohn, General Dynamics, San Diego, California 2 

T. T. Tanalski, General Dynamics, San Diego, California 1 

W. C. Dietz, General Dynamics, Fort Worth, Texas 2 

W. K. Bailey, General Dynamics, Fort Worth, Texas 1 

C. F. Herndon, General Dynamics, Fort Worth, Texas 1 

D. C. Little, General Dynamics, Fort Worth, Texas 1 

Ira G. Hedrick, Grumman Aerospace Corporation 1 
Bethpage , L . I . , New York 

Warner Lansing, Grumman Aerospace Corporation 1 
Bethpage , L . I . , New York 



111-32 



T. C. Adee, Grumman Aerospace Corporation 1 
Bethpage , L . I . , New York 

R. W. Wood, Grumman Aerospace Corporation 1 
Bethpage , L . I . , New York 

Frederick T. Main, Grumman Aerospace Corporation 1 
Bethpage , L . I . , New York 

Jim Brennan, Grumman Aerospace Corporation 1 
Bethpage , L . I . , New York 

R. H. Wells, Lockheed -California, Burbank, California 1 

W. T. Shuler, Lockheed-Georgia, Marietta, Georgia 2 

L. W. Lassiter, Lockheed-Georgia, Marietta, Georgia 1 

Lloyd Nelson, Lockheed -California, Burbank, California 1 

John C. Ekvall, Lockheed-California, Burbank, California 2 

J. C. Wordsworth, Lockheed-California, 1 
Burbank, California 

J. H. Best, LTV- Aeronautics, Dallas, Texas 1 

W. H. Sparrow, LTV-Aeronautics, Dallas, Texas 2 

Alan Starr, LTV-Aeronautics, Dallas, Texas 1 

Leon Boswell, LTV-Aeronautics, Dallas, Texas 1 

Melvin Stone, McDonnell -Douglas, Long Beach, California 1 

Jerome E. Fischler, McDonnell-Douglas , Long Beach, Ca. 2 

H. C. Schjelderup, McDonnell-Douglas, Long Beach, Ca. 1 

Don G. Smillie, McDonnell -Douglas , Long Beach, California 1 

M. L. Ramey, McDonne 11 -Douglas , St. Louis, Missouri 1 

R. C. Goran, McDonnell -Douglas, St. Louis, Missouri 1 

0. B. McBee, McDonnell-Douglas, St. Louis, Missouri 1 

S. A. LaFavor, McDonnell-Douglas, St. Louis, Missouri 1 

Otto Acker, North-American Rockwell, Columbus, Ohio 1 



IH-33 



Paul Maynard, North -American Rockwell, Columbus, Ohio 1 

B. S. Coogan, North -American Rockwell, Columbus, Ohio 1 
J. W. Ellis, North-American Rockwell, Los Angeles, Ca. 2 
H. W. Sweet, North -American Rockwell, Los Angeles, Ca. 1 
R. V. Cimino, North -American Rockwell, Los Angeles, Ca. 1 
Robert Olsen, North -American Rockwell, Downey, California 1 
L. R. Fowell, Northrop Corp., Hawthorne, California 1 
Matthew W. Sagal, Western Electric Co., Princeton, N. J. 1 
William M. Duke , Whittaker Corp . , Los Angeles , California 1 
K. Berg, Whittaker Corporation, San Diego, California 1 
T. B. Owens, National Science Foundation, Washington, D.C. 1 
N. Perrone, Office of Naval Research, Washington, D. C. 1 
J. M. Crowley, Office of Naval Research, Washington, D.C. 1 
0. Remson, Naval Material Command, Washington, D. C. 1 
J. Mayers, Stanford University, Stanford, California 1 
J. P. Reese, AIA, Washington, D. C. 1 

C. P. Baum, Bethesda, Maryland 1 

Robert H. Brown, Natrona Heights, Pennsylvania 1 

Henry R. Clauser, Armonk, New York 1 

Dean of Research Administration, Naval Postgraduate 2 
School, Monterey, California 

R. W. Bell, NPS, Monterey, California 1 

D. W. Mathews, NPS, Monterey, California 1 
W. M. Coates, NPS, Monterey, California 1 
R. E. Ball, NPS, Monterey, California 1 
M. H. Bank, NPS, Monterey, California 1 



111-3^ 



J. A. J. Bennett, NPS, Monterey, California 1 

0. Biblarz, NPS, Monterey, California 1 

D. J. Collins, NPS, Monterey, California 1 

A. E. Funs, NPS, Monterey, California 1 

T. H. Gawain, NPS, Monterey, California 1 

R. A. Hess, NPS, Monterey, California 1 

G. J. Hokenson, NPS, Monterey, California 1 

C. H. Kahr, NPS, Monterey, California 1 

D. M. Layton, NPS, Monterey, California 1 
G. H. Lindsey, NPS, Monterey, California 1 
J. A. Miller, NPS, Monterey, California 1 
D. W. Netzer, NPS, Monterey, California 1 
M. F. Platzer, NPS, Monterey, California 1 
H. L. Power, NPS, Monterey, California 1 
M. H. Redlin, NPS, Monterey, California 1 
W. Schlacter, NPS, Monterey, California 1 
L. V. Schmidt, NPS, Monterey, California 1 
R. P. Shreeve, NPS, Monterey, California 1 
M. H. Vavra, NPS, Monterey, California 1 
R. D. Zucker, NPS, Monterey, California 1 
J. D. Esary, NPS, Monterey, California 1 
R. B. Leonesio, NPS, Monterey, California 1 



IH-35 



UNCLASSIFIED 



Security Classification 



DOCUMENT CONTROL DATA -R&D 

(Security classification ol title, body ol abstract and indexing annotation must be entered when the overall report is classified) 



Originating ACTIVITY (Corporate author) 

Naval Postgraduate School 
Monterey, California 939^0 



2a. REPORT SECURITY CLASSIFICATION 

UNCLASSIFIED 



2b. GROUP 



REPOR T TITLE 



Interface of Materials and Structures on Airframes, Part 3, Design Problems 
in Aircraft Structures, including Proceedings of Monterey Symposium 

descriptive NOTES (Type of report and.inclusive dates) 

Final Report 

au THORisi (First name, middle initial, last name) 



Ulrich Haupt 



S. REPOR T DATE 

October 1971 



7«. TOTAL NO. OF PAGES 



228 



7b. NO. OF REFS 



Sa. CONTRACT OR GRANT NO 

WR-2-6059 
b. PROJEC T NO 



»a. ORIGINATOR'S REPORT NUMBER(S) 



57HP7H11A 



9b. OTHER REPORT NO(S) (Any other numbers that may be assigned 
this report) 



10 DISTRIBUTION STATEMENT 



This document has been approved for public release and sale; its distribution 
is unlimited. 



II. SUPPLEMENTARY NOTES 



12. SPONSORING MILI TARY ACTIVITY 

Naval Air Systems Command 
Washington, D. C. 



13. ABSTR AC T 



The proceedings of the Monterey Symposium on Design Problems in 
Aircraft Structures provide a basic survey of design problems from the 
engineer's viewpoint. Further analysis of the present situation draws 
attention to some essential aspects which are not yet generally 
recognized. This leads to the conclusion that recent design problems 
cannot be solved on a technological level alone. An organizational 
effort is needed to disseminate available information. Beyond this, the 
complexity of interactions must be understood more thoroughly and this 
requires an educational effort on a broad basis. A practical and 
systematic approach toward the solution of these problems is developed. 



DD 



FORM 

I NOV •» 

S/N 0101 -807-681 1 



1473 



(PAGE 1) 



III-36 



UNCLASSIFIED 

Security Classification 



A-3140S 



UNCLASSIFIED 



Sci unly Classif nation 



KEY WORDS 



Materials 
Structures 

Design 
Airframes 



DD ;™r„ 14 73 ' 

S/N 010t-807-hM-| 



BACK) 



UNCLASSIFIED 



Seoirity Classification 



U14263 



DUDLEY KNOX LIBRARY • RESEARCH REPORTS 



5 6853 01060463