, USARTL-TR-79-22C

AD A089104

AIRCRAFT CRASH SURVIVAL DESIGN GUIDE

VOLUME III - AIRCRAFT STRUCTURAL CRASHWORTHINESS

SIMULA INC.

2223 SOUTH 48TH STREET

TEMPE, ARIZONA 85282

AUGUST 1980 r

FINAL REPORT

 PREPARED FOR

APPLIED TECHNOLOGY LABORATORY

U. S. ARMY RESEARCH AND TECHNOLOGY LABORATORIES (AW/RADCOM)

FORT EUSTIS, VIRGINIA 23604

i Ac-prove for public releasne; Jistribtition unlirniot,.

- ... 80 9 10 005
APPLIED TECHNOLOGY LABORATORY POSITION STATEMENT

This revised edition of the Crash Survival Design Guide was

prepared for the Applied Technology Laboratory by Simula Inc.

under the terms of Contract DAAJ02-77-C-0021. The original

Crash Survival Design Guide was published in 1967 F.s USAAVLABS

Technical Report 67-22 and subsequent revisions were published

as USAAVLABS Technical Report 70-22 and USAAMRDL Technical Report 71-22. This current edition consists of a consolidation

of design criteria, concepts, and analytical techniques developed through research programs sponsored by this Laboratory

over the past 20 years into one report suitable for use as a

designer's guide by aircraft design engineers and other interested personnel.

This document has been coordinated with USAAVRADCOM, the U. S.

Army Safety Center, the U. S. Army Aeromedical Research Laboratory, and several other Government agencies active in aircraft

crashworthiness research and development.

The technical monitors for this program were Messrs. G. T.

Singley III, R. E. Bywaters, W. J. Nolan, and H. W. Holland of

the Safety and Survivability Technical Area, Aeronautical Systems Division, Applied Technology Laboratory.

Comments or suggestions pertaining tn this Design Guide will

be welcomed by this Laboratory.

DISCLAIMERS

Th findings In this report are not to be construed as in official Department of the Army position unles so designated by bther authorized documents.

When Government drawings, specifications, or other data are used for any purpose other than in connection

with a definitely related Government procurement operation, the United States Government thereby Incurs no responsibility nor any obligation whatsoever: and the fact that the Government may have formulated, furnished, or in any way supplied the said drawings, specifications, or other date is not to be regarded by implication or otherwise as In any manner licensing the holder or any other person or corporation, or conveying any rights or permision, to manufacture, use, or seil any patented Invention that rmy in any way be related thereto.

Trade names cited in this report do not constitute an official endorsement or approval of the use of such

commerolal hardware or software.

DISPOSITION INSTRUCTIONS

Destroy this report when no longer needed. Do not return it to the originator.

NOTE: This is a revised edition of the Crash Survival Design

Guide (formerly USAAMRDL Technical Report 71-22). All

previoas editions are obsolete and should bu destroyed.

4)
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SECUMITY CLASSIFICATION OF THIS PAGO(MM Del. 3,M..4

20. (Continued)

Volume I - Design Criteria and Checklists

Volume 11 - Aircraft Crash Environment and Human Tolerance

Volume III -Aircraft Structural Crauhworthineuu

Volume IV - Aircraft Seats# Restraints, Litters, and Padding _Volume V - Aircraft Postcrash Survival

This volume-.4olume 1IIý contains information on the design of

aircraft structures and structural elements for improved crash

survivability. Current requirements for structural design of

U. S. Army aircraft pertaining to crashworthiness are discussed.

Principles for crashworthy design are presented in detail for the

landing gear and fuselage subject to a range of crash conditions,

including impact. that are primarily longitudinal, vertical, or

lateral in nature and those that involve more complicated dynamic

conditions, such as rollover. Analytical methods for evaluating

structural crashworthiness are described

I Jutf

UNCLASSXFIED

SECURITY CLASSIFICATION OF THIS PASEKWMO. Ddie Znflh.,d)
I PREFACE

This report was prepared for the Safety and Survivability Tech- nical Area of the Applied Technology Laboratory, U. S. Army Research and Technology Laboratories (AVRADCOM), Fort Eustis,

Virginia, by Simula Inc. under Contract DAAJ02-77-C-0021, initiated in September 1977. The Department of the Army Project Number is 1L162209AH76. This guide is a revision of USAAMRDL

Technical Report 71-22, Crash Survival Design Guide, published

in October 1971.

A major portion of the data contained herein was taken from

U. S. Army-sponsored research in aircraft crashworthiness conducted from 1960 to 1979. Acknowledgment is extended to the

U. S. Air Force, the Federal Aviation Administration, NASA, and

the U. S. Navy for their research in crash survival. Apprecia- tion is extended to the following organizations for providing

accident case histories leading to the establishment of the impact conditions in aircraft accidents:

o U. S. Army Safety Center, Fort Rucker, Alabama.

@ Civil Aeronautics Board, Washington, D. C.

* U. S. Naval Safety Center, Norfolk, Virginia.

a U. S. Air Force Inspection and Safety Center, Norton

Air Force Base, California.

Additional credit is due the many authors, individual companies, and organizations listed in the bibliographies for their

contributions to the field. The contributions of the following authors to previous editions of the Crash Survival Design Guide are most noteworthy:

D. F. Carroll, R. L. Cook, S. P. DesJardins, J. K. Drummond, J. L. Haley, Jr., A. D. Harper, H. G. C. Henneberger,

N. B. Johnson, G. Kourouklis, W. H. Reed, S. H. Robertson,

L. M. Shaw, Dr. J. W. Turnbow, and L. W. T. Weinberg.

This volume was prepared by Dr. D. H. Laananen of Simula Inc.,

G. T. Singley, III of ATL, A. E. Tanner of the Boeing Vertol

Company, and Dr. J. W. Turnbow of Arizona State University. Technical review and comments were provided by S. P. DesJardins

of Simula Inc. Information on cargo restraint was provided by

J. Shefrin of the Boeing Vertol Company.

3

A.
TABLE OF CONTENTS

Section Title

PREFACE .. . . . . . . . . . . . . . . . .. . . . . 3

LIST OF ILLUSTRATIONS .............. 8

LIST OF TABLES. . . . . . . . . . . . . . . . o o 14

INTRODUCTION. . . . . . o . . . o 16

CHAPTER 1. BACKGROUND DISCUSSION. . . . . . . . . . 19

CHAPTER 2. DEFINITIONS. . . . . . . . . . . o 25

2.1 AIRCRAFT COORDINATE SYSTEMS AND ATTITUDE

PARAMETERS o . e s # a. . . . . . . e . 25

2.2 ACCELERATION-RELATED TERMS . . . . . . . 27

2.3 VELOCITY-RELATED TERMS . .. . . . . . . 28

2.4 FORCE TERMS. . . . . . . . . . . . . . s 29

2.5 DYNAMICS TERMS . . . . * . . . . . . . . 31

2.6 CRASH SURVIVABILITY TERMS...... . . 32

2.7 OCCUPANT-RELATED TERMS . . * . . . . . 32

2.8 STRUCTURAL TERMS . . . . . . . . . . . 34

CHAPTER 3. GENERAL DESIGN CONSIDERATIONS. . . . . . 37

3.1 AIRCRAFT TYPES - THEIR MISSIONS AND

DESIGN RESTRAINTS. . . . . o o s o e 37

3.1.1 Helicopters.. . . . . .. .. . . . . . 37

3.1.2 Fixed-Wing Aircraft. . . . . . . . .o. . 38

3.2 APPLICABILITY OF CRITERIA. . . . . . . . 38

3.3 ACCEPTANCE OF STRUCTURES . . . , * o , . 38

3.4 SELECTION OF STRUCTURES. . .o. . . . . . 41

3.5 TESTING. . . . . . . . . . . . . . . 45

CHAPTER 4. CRASH ENVIRONMENT. . . . o . . o o o 46

4.1 INTRODUCTION .. .... . .. 46

4.2 STRUCTURAL DAMAGE WHICH FREQUENTLY

RESULTS IN OCCUPANT INJURY . . . o . . . 51

4.2.1 Longitudinal (Crushing) Loads on

Cockpit Structure.s "..o e s s s 9 52

4.2.2 Vertical (Crushing) Loads on Fuselage SShell.. . . 52

4.2.3 Lateral (Crushing) Loads on'Fuselage Shells s. . .. e 9 9 9 s 53

4.2.4 Transverse (Bending) Load. on Fuselage

5

I ~kIR pas
TABLE OF CONTENTS (CONTD)

Section Title Page

4.2.5 Deformation (Buckling) of Floor

Structure* 53

:•Shell,. 4.2.6 Landing Gear Penetration . . . of . .Fuselage . 54

4.2.7 Helicopter Lateral Rolvr. . .... 54

4.2.8 Rupture of Flammable Fluid Containers.. 54

CHAPTER 5. DESIGN REQUIREMENTS. . . . . . . . . . . 55

5.1 INTRODUCTION . . . . . . . . . . . . . . 55

5.2 GENERAL REQUIREMENTS . . . . . . . . . . 55

5.3 DETAIL REQUIREMENTS. . . . . . . . . . . 57

5.3.1 Airframe Crashworthiness . . . . . . , 57

5.3.2 Ancillary Equipment Retention . . . . . 66

5.3.3 Occupant Retention... . . . . . . . . 68

5.3.4 Cargo Retention. .. . .. .. .. .. 69

5.4 TESTING. . , o . . . . • . , . . . . . . 69

5.4.1 Aircraft System Testing. .... . . . . 69

5.4.2 Landing Gear Crash Testing . .o . . . , 70

5.4.3 Cargo Restraint. o . 6 .' 0 " " 0 ." " 71

5.4.4 Seat and Restraint System, . . . . . . . 71

5.4.5 Fuel System. . 0 . . 1 0 . . . . . . 71

5.4.6 Ancillary Equipment Retention. . . . . . 71

CHAPTER 6. AIRFRAME PRINCIPLES AND CONCEPTS . . . . 72

6.1 INTRODUCTION . o . . . . . . . . o 72

6.2 GENERAL DESIGN CONSIDERATIONS. . . . . . 72

6.2.1 Initial Layout .... .... . . . 73

6.2.2 Analysis and Simulation. 74

6.2.3 Design Review and Final Aircraft Format

Selection.... - . . . . . . 74

6.3 MATERIALS AND STRUCTURAL PROPERTIES. . 74

6.3.1 General, .o .# . . • . . . . . ' . 76

6.3.2 Material Strength'and Elongatio

Characteristics. o . 0 0 . 0 0 a a . . 0 77

6.3.3 Failure Modes Controlling Material

Selection. . . . . . . . . . . . 77

* 6.3.4 Composite Materias...... . . 81

6.3.5 Spark Generation . ........... 96 P 6.4 ENERGY-ABSORPTION DEVICES. . . . . . . . 99 p 6.5 LANDING GEAR . . . . . . . . . . o . . . 101

6.5.1 Wheel-Type Gear. . • • . , . . . . . . 103

* 6.5.2 Skid Gear. . . . . . . . . . . . . 109

6.6 FUSELAGE . . . . o o . o . . . . . . . . , i

6o6.1 General. . . . . . . . . . . o . 0 111

6

...i .
TABLE OF CONTENTS (CONTD)

Section Title Page

6.6.2 Energy-Absorption Capacity of Forward

Fuselage .. . . . . . . .... 112

6.6.3 Reduction of Earth Plowing . .. .... 116

6.6.4 Reduction of Aircraft Mass . . . . . . . 119

6.6.5 Design Concepts for Improved

Crashworthiness............. 123

SCHAPTER 7. ANALYTICAL METHODS . . . . . . . . . .. 169

7.1 INTRODUCTION . . . . . . . . . . . . . 169

7.2 DYNAMICS OF THE CRASH ENVIRONMENT. . . . 169

7.2.1 Kinematic Relationships . . 169

7.2.2 Energy Absorption During Deceleration. . 174

7.2.3 Stopping Distance . .... . . 181

7.2.4 Energy Content of Aircraft at Impact . . 189

7.2.5 Postimpact Energy Dissipation. . . . . . 193

7.3 LANDING GEAR ANALYSIS...... o . . 0 195

7.3.1 Wheel Landing Gear .. ......... 195

7.3.2 Skid Landing Gear.. .. . 201

7.4 SEMIEMPIRICAL ANALYSES OF AIRFRAME

STRUCTURAL CRASHWORTHINESS . . . . . . . 203

7.4.1 Analysis Procedure . 6 0 a 0 a 0 & . 9 6 204

7.4.2 Substructure Test Results. . * e e e 9 o 207

7.4.3 Correlation of Test and Analysis

Results... - . . . . . • . 208

7.4.4 Example of Analysis for Vertical !i Impact . 210

7.4.5 Example of Analysis for Longitudinal

Impact . * * * . a . . . . . . . . . 220

7.4.6 Lateral Impact . . . . . . . . . . . . . 224

7.4.7 Rollover . , 225

7.5 STRUCTURAL CRASHWORTHINESS SIMULATION

COMPUTER PROGRAMS. . . . . . . . . . . . 225

7.5.1 Program KRASH . . . . . . . . . .... 230

7.5.2 Program DYCAST * . . .a ..... 233

7.5.3 Program WRECKER, . a 6 .. . . 238

7.6 POTENTIAL SOURCES OF BASIC STRUCTURAL

DATA FOR CRASHWORTHINESS ANALYSIS. . . . 241

7.6.1 Estimates. .. . . . . . . 242

* 7.6.2 Aircraft Accident Data . 242

7.6.3 Controlled Crawh Testing . . . . . . . . 243

7.6.4 Scale Model Testing.. ......... 245

REFERENCES. . . . . . . . . . . . . . . . . . . . . . 249

BIBLIOGRAPHY. . . . . . . . . . . . . . . . . . . . . 259

INDEX . . . . . . . . . . . . . . . . . . . . . .. . 263

7

-- -- -- -
LIST OF ILLUSTRATIONS

Figure Page

1i Procedure for evaluation of structural

designs with respect to crashworthi- hess . o . . ° . ° ° . . . . . ° ° ° 23

2 Aircraft coordinates and attitude

directions . . ............. 25

3 Typical aircraft floor acceleration pulse . .. . . . . . . . 29

4 Terminology for directions of forces on

the body . . . . . . . . . . . . . . . . 33

5 Side elevations of typical U. S. Army utility and cargo helicopters . . . .. 39

6 Side elevations of typical U. S. Army

observation and attack helicopters . . . 40

7 Side elevation of typical U. S. Army

training helicopter. . . . . . .. 41

a Selected helicopter crashworthy

features . . . . . . . . . . . . . . . . 42

S9 Side elevations of typical U. S. Army

fixed-wing aircraft. . . . . . . . . . . 43

10 Design velocity change - off-axis

requirements .............. 49

11 Low angle impact design conditions

(simulated approach with antitorque

loss under poor visibility). . . . . . . 58

12 Nose section design conditions . . . . . 59

13 Rollover, roof impact design

condition. * * * * 9 o * * a * a 61

14 Rollover, side impact design

condition. .*............. 62

15 Structi-ral layout for occupant protection in a crash environment. . . . . . . 75

!i 8
LIST OF ILLUSTRATIONS (CONTD)

Figure page

16 Compression buckling at base of

buttline beam structure on a medium

cargo helicopter following vertical i•impact . . . . . . . . . . . . . . . . . 79

17 Typical buckling collapse of vertical

bulkhead and buttline beam on a medium

cargo helicopter following vertical impact . . . . . . . . . . . . . . . . . 80

18 Typical frame failure away from

occupied space in the cockpit area

of a medium cargo helicopter following vertical impact..... ..... . 81

19 Typical frame failure away from occupied

space in the cabin area of a medium

cargo helicopter following vertical impact . . . . . . .•••• 82

20 Failure modes resulting in jagged

elements protruding into occupied

space in the cabin area of a medium

cargo helicopter following vertical impact . . . . . ...... 83

21 Rotation joint failure with compression

and bending in a medium cargo helicopter

following vertical impact.. ...... 84

22 Failure of frame member and joint with

fragmentation, compression, and bending in a medium cargo helicopter following vertical impact . . . . . . ..... 84

23 Stress-strain relationship for aluminum

alloy (7075) and 0 degrees graphite/

epoxy composite ............. 85

24 Overall fuselage concepts .. ... .. 89

25 Energy-absorption concepts - beams and

bulkheads (vertical impact) ....... 91

26 Energy-absorption concepts - tubular

construction (oblique vertical

impact). 92
LIST OF ILLUSTRATIONS (CONTD)

Figure Page

27 Fuselage sidewall concepts - lateral impact . . . . . . . . . . . . . . . . . 94

28 Antiplowing concepts - longitudinal impact .. . . . . .. . . 95

29 Crashworthy joint concepts . . . . . . . 97

30 Examples of energy-absorbing devices . . 100

31 Landing gear characteristics ...... 105

32 Landing gear configuration . . . . . 107

33 Nose landing gear crash energy

absorption .1............. 108 o

34 Helicopter skid gear and improved

pitch-interconnected concept . . . . . . 10

35 Observation helicopter - layout of

crashworthy features ......... 113

36 Impulsive aircraft acceleration as a

function of velocity and ratio of

accelerated mass of earth to aircraft

mass (based upon assumed time, At, for

acceleration of earth mass) . . . . ... 120

37 Method of reinforcing nose structure to

provide increased resistance to vertical

loads and to reduce earth scooping . . . 125

38 Typical underfloor canted frame and

longitudinal beam member to minimize

nose plowing . . .. . .. . . . . . .. 126

39 Features of helicopter nose section to

prevent nose plowing . . . . . . . . . . 127

40 Two methods of reducing earth scooping

in engine-mounted areas. . . . . . . 128

41 Longitudinal impact and blade strike

"protection: nose section. The figure

delineates those members requiring

careful design consideration . . . . . . 130

10

* i
LIST OF ILLUSTRATIONS (CONTD)

SFigure aq

42 Longitudinal impact protection in cabin section . . . . . . . . . . . . . . . . . 131

43 Typical longitudinal beams adjacent to

cockpit for longitudinal continuity of

overhead structure . . . . . . . . . . . 132

44 Typical full-depth longitudinal beams

for overhead support of large mass

items and longitudinal continuity of

structure . . . . ... . . .. . . . . . 132

45 Fuselage structure with overhead large

mass items (idealization). . . . . . . . 133

46 Fuselage structure without overhead

mass items (idealization). . . . . ... 134

47 Cap-and-web combination beam design

with potential energy-absorbing

capability . . . . . . . . . . . . .. . 136

48 Conceptual structural configurations to

absorb maximum energy for sideward,

longitudinal, and vertical impact

forces . . . . . . . . . . . . . . . . . 138

49 Crushable belly deforms inward without

buckling floor or reducing living

space . . . . . . . . . . . . . . 139

50 Examples of closed box-beam frame

sections . .* . . .* . . . . . . . . . . . 140

51 Typical cockpit overhead longitudinal

beam member. ......... .... 141

52 Floor acceleration comparisons

(30-ft/sec vertical impact) . . . . . . 142

53 Overhead mass protection and rollover

protection in cabin section . . . . . . 143

54 Rollover protection in nose section.

The figure delineates those members

requiring careful design considera- i•tion . . . . . . . .144

11
LIST OF ILLUSTRATIONS (CONTD)

Figure Page

55 Typical structural sections for nose

section. . . . . . . . . . . . . . . .. 145

56 Externally mounted blade strike

deflectors . . . . . . . ....... 146

57 Frangible main rotor blade tip . . . . . 148

58 Typical interior support structure

for fuel cell. ... . . . .... 153

59 Typical crashworthy seat

configuration. ... ......... 155

60 Aircraft floor longitudinal pulses for

rotary- and light fixed-wing aircraft. 160

61 Load-displacement requirements for

energy-absorbing cargo restraint

systems (forward loading of rotarywing and fixed-wing aircraft). . . . . . 162

62 Cargo lateral load-displacement

requirements . . . . . . . . . . . . . . 163

63 Results of integrated cargo restraint/

crash simulation tests using energy

absorbers and low-elongation tension

members and forward longitudinal loaddisplacement requirements. .. . . . . .. 164

64 Definition of position and displacerent

for crashing aircraft. ......... 170

65 Assumed relationship for illustration

of crash kinematics. .......... 172

66 Definition of work. . . . . .. . ... 175

67 Illustration of impact reducing

aircraft kinetic energy. ........ 176

68 Illustration of deceleration by

constant force . . . . . . . . . . . . . 177

12

""Po
LIST OF ILLUSTRATIONS (CONTD)

Figure Page

69 Force-displacement curve for honeycomb

materials..,.. ..... ..... 178

70 Theoretical stopping distance as a

function of velocity change and

average deceleration level . . . . . 182

71 Deceleration, velocity, and distance

as functions of time for five pulse

shapes . . . . . . . . . . . . . . . 190

72 Comparison of stopping distances for

various deceleration pulse shapes. . . . 191

73 Shock strut schematic.......... 196

74 Free-body diagram of aircraft on

ground . . . . . . . . . . . . . . . . . 199

75 Landing gear loads ........... 199

76 Lower fuselage bulkhead and stiffener

arrangement. .... . .. .. .. . 205

77 Predicted-versus-test load-deflection

curves for a representative specimen . . 208

78 Comparison of test and analysis results

for "like" specimens . . . .... . . . 209

79 Types of load distribution acting on

helicopter structure for impact with

soft and hard surfaces . .* e . . 211

80 Typical load-displacement characteristics'for airframe structure. ..... 212

01 Simplified load-deflection characteristics for a statically loaded stringer

skin combination ... . . . . . 213

82 Typical frame section of side element

of a medium cargo helicopter center

fuselage . . . . .... 214

13

C,
LIST OF ILLUSTRATIONS (CONTD)

Figure Page

83 Crippling allowables for typical

aluminum alloy extrusions ...... . . 216

84 Crippling allowables for aluminum alloy formed sections other than simple zees

and channels . . . . . . . . . 217

85 Effective area of skin for aluminum

alloy stringer-skin combinations . . . . 218

86 Estimated dynamic energy-absorption capability of typical fuselage

structure . . . . . . .. . . . . . .. . 221

87 Structiiral crashworthiness features for

longitudinal impact resistance . .... 223

88 Model of existing utility helicopter . . 235

89 Model of existing medium cargo

helicopter . . . . . . . . . . . . . . . 236

90 Simulation of CH-47 nose section drop test . . . . . . . . . . . . . . . . . .237 ..

91 Cost ratio (scale model test cost/fullscale test cost) versus scale factor

for structures weighing less than

10,000 lb. ........ 247

92 Stiffened panel specimen .... . . o . 248

LIST OF TABLES

Table

1 Crashworthiness Criteria for the

Preliminary Design Process ....... 24

2 Summary of Crash Impact Conditions for

Helicopters and Light Fixed-Wing Aircraft Design 47

3 Performance Requirements of Structural

Crashworthiness ........... 56

14
LIST OF TABLES (CONTD)

3 ~Table Pg

4 Aircraft Cargo Categories. . . . . . . . 70

5 Ultimate Strength and Elastic Moduli

for Typical Structural Composites. . 86

6 Thermal Coefficients of Expansion for

Composite and Metallic Materials . . . . 87

7 Minimum Conditions Under Which Certain Abraded Metal Particles Will Ignite. . . 98

8 Cargo Restraint Loads and Displacement

R•equirements . . . . . . . . . . . . . . 166

9' Cargo Net Materials - Relative

Characteristics. * . . . . * • . • * • * 167

10 Program 1RASH Experimental

Verification 234

WA

,i'I.

15

0.
INTRODUCTION

For many years, emphasis in aircraft accident investigation

was placed on determining the cause of the accident. Very little effort was expended on the crash survival aspects of

aviation safety. However, it became apparent through detailed

studies of accident investigation reports that significant improvements in crash survival could be made if consideration

were given in the initial aircraft design to the following factors that influence survivability:

1. Crashworthiness of Aircraft Structure - The ability

of the aircraft structure to maintain living space

for occupants throughout a crash.

2. Tiedown Chain Strength - The strength of the linkage

preventing occupant, cargo, or equipment from breaking free and becoming missiles during a crash sequence.

3. Occupant Acceleration Environment - The intensity

and duration of accelerations experienced by occupants (with tiedown assumed intact) during a crash.

4. Occupant Environment Hazards - Barriers, projections,

and loose equipment in the immediate vicinity of the

occupant that may cause contact injuries.

5. Postcrash Hazards - The threat to occupant survival

posed by fire, drowning, exposure, etc., following

the impact sequence.

Early in 1960, the U. S. Army Transportation Research Command*

initiated a long-range program to study all aspects of aircraft

safety and survivability. Through a series of contracts with

the Aviation Safety Engineering and Research (AvSER) Division

of the Flight Safety Foundation, Inc., the problems associated

with occupant survival in aircraft crashes were studied to determine specific relationships among crash forces, structural

failures, crash fires, and injuries. A series of reports covering this effort was prepared and distributed by the U. S.

Army, beginning in 1960. In October 1965, a special project

initiated by the U. S. Army consolidated the design criteria

presented in these reports into one technical document suitable

for use as a designer's guide by aircraft design engineers and

other interested personnel. The document was to be a summary

*Now the Applied Technology Laboratory, Research and Technology

Laboratories of the U. S. Army Aviation Research and Development Command (AVRADCOM).

16
of the current state of the art in crash survival design, using

not only data generated under Army contracts, but also information collected from other agencies and organizations. The

Crash Survival Design Guide, first published in 1967, realized

Al,. this goal.

Since its initial publication, the Design Guide has been revised several times to incorporate the results of continuing

research in crashworthiness technology. The last revision,

published in 1971, was the basis for the criteria contained

in the Army's military standard dealing with aircraft crashworthiness, MIL-STD-1290(AV), "Light Fixed- and Rotary-Wing

Aircraft Crashworthiness" (Reference 1). This current revision, the fourth, contains the most comprehensive treatment of

all aspects of aircraft crash survival now documented. It can

be used as a general text to establish a basic understanding of

the crash environment and the techniques that can be employed

to improve chances for survival. It also contains design criteria and checklists on many aspects of crash survival and thus

can be used as a source of design requirements.

The current edition of the Aircraft Crash Survival Design Guide

is published in five volumes. Volume titles and general subjects included in each volume are as follows:

Volume I - Design Criteria and Checklists

Pertinent criteria extracted from Volumes II through V,

presented in the same order in which they appear in those

volumes.

Volume II - Aircraft Crash Environment and Human Tolerance

Crash environment, human tolerance to impact, military

anthropometric data, occupant environment, test dummies,

accident information retrieval.

Volume III - Aircraft Structural Crashworthiness

Crash load estimation, structural response, fuselage and

landing gear requirements, rotor requirements, ancillary

equipment, cargo restraints, structural modeling.

1. Military Standard, MIL-STD-1290(AV), LIGHT FIXED- AND

ROTARY-WING AIRCRAFT CRASHWORTHINESS, Department of Defense, Washington, D. C., 25 January 1974.

17 ii• i.1-
Volume IV- Aircraft Seats, Restraints, Litters, and Padding

Operational and crash environment, energy absorption, seat

design, litter requirements, restraint system design,

occupant/restraint system/seat modeling, delethalization

of cockpit and cabin interiors.

Volume V - Aircraft Postcrash Survival

i 2. beacons. Postcrash fire, ditching, emergency escape, crash locator

This volume (Volume III) contains information on aircraft

structural crashworthiness. Following a general discussion of

aircraft crashworthiness in Chapter 1, a number of terms commonly used in discussing the crash environment and aircraft

structures are defined in Chapter 2. Chapter 3 presents a general discussion of aircraft types, principles for the design of

crashworthy vehicles, and testing for evaluation of crashworthiness. Although Volume II discusses the crash environment in

detail, a summary of information pertinent to aircraft structural design is presented in Chapter 4. Structural design and I testing requirements for improved crashworthiness are described

in Chapter 5. Principles and concepts for improving crashworthiness in aircraft structures are described in Chapter 6.

Chapter 7 discusses analytical tebhniques for evaluating structural crashworthiness.

18

5;L
1. BACKGROUND DISCUSSION

* The overall objective of designing for crashworthiness is to

U! •eliminate unnecessary injuries and fatalities in relatively mild impacts. A crashworthy aircraft also reduces aircraft

crash impact damage. By minimizing personnel and material

losses due to crash impact, crashworthiness conserves resources, is a positive morale factor, and improves the combat

effectiveness of the fleet. Results from analyses and research

during the past several years have shown that the relatively

small cost in dollars and weight of including crashworthiness

features is a wise investment (References 2 through 13). Consequently, new generation aircraft are being procured to stringent, yet practical requirements for crashworthiness.

To provide as much occupant protection as possible, a systems

approach to crashworthiness must be followed. Every available

subsystem must be considered in order to maximize the protec- . tion afforded to vehicle occupants. When an aircraft impacts

2. ENGINEERING ANALYSIS OF CRASH INJURY IN ARMY OH-58 AIRCRAFT, USASC Technical Report, U. S. Army Safety Center,

Fort Rucker, Alabama (to be published).

3. ENGINEERING ANALYSIS OF CRASH INJURY IN ARMY CH-47 AIRCRAFT, USAAAVS Technical Report 78-4, U. S. Army Agency

for Aviation Safety, Fort Rucker, Alabama, June 1978.

4 "4. ENGINEERING ANALYSIS OF CRASH INJURY IN ARMY AH-l AIRCRAFT, USAAAVS Technical Report 78-3, U. S. Army Agency

for Aviation Safety, Fort Rucker, Alabama, March 1978.

5. Carnell, B. L., CRASHWORTHINESS DESIGN FEATURES FOR ADVANCED UTILITY HELICOPTERS, in Aircraft Crashworthiness,

K. Saczalski, et al., eds., University Press of Virginia, Charlottesville, Virginia, 1975, pp. 51-64.

6. Bainbridge, M. E., Reilly, M. J., and Gonsalves, J. E.,

CRASHWORTHINESS OF THE BOEING VERTOL UTTAS, in Aircraft

Crashworthiness, K. Saczalski, et al., eds., UnTiversty

of Virginia, SPress Charlottesville, Virginia, 1975,

pp. 65-82.

7. Rich? M. J., INVESTIGATION OF ADVANCED HELICOPTER STRUCTURAL DESIGNS, Volume I, ADVANCED STRUCTURAL COMPONENT DESIGN CONCEPT STUDY, Sikorsky Aircraft, Division of United

Technologies Corporation; USAAMRDL Technical Report 75-59A,

Eustis Directorate, U. S. Army Air Mobility Research and

Development Laboratory, Fort Eustis, Virginia, May 1976,

AD A026246.

19 Ag%

T1*
the ground, deformation of the ground absorbs some energy. This is an uncontrolled variable since the quality of the impact3d surface usually cannot be selected by the pilot. If

the aircraft lands on an appropriate surface in an appropriate

attitude, the landing gear can be used to absorb a significant

amount of the impact energy. After stroking of the gear,

crushing of the fuselage contributes to the total energyabsorption process. The fuselage must also maintain a protective shell around the occupant so the crushing must take place

outside the protective shell. The functions of the seat and

restraint system are to restrain the occupant within the protective shell during the crash sequence and to provide additional energy-absorbing stroke to further reduce occupant decelerative loading to within human tolerance limits. The

structure and components immediately surrounding the occupant

must also be considered. Weapon sights, cyclic controls, glare

shields, instrument panels, armor panels, and aircraft structure must be delethalized if they lie within the strike envelope of the occupant.

S. Hoffstedt, D. J., and Swatton, S.ý ADVANCED HELICOPTER

STRUCTURAL DESIGN INVESTIGATION, The Boeing Vertol Companyp USAAMRDL Technical Report 75-56A, Eustis Directorate, U. S. Army Air Mobility Research and Development

Laboratory, Fort Eustis, Virginia, March 1976, AD A024662.

9. Hicks, J. E., AN ANALYSIS OF LIFECYCLE ACCIDENT COSTS FOR

THE ADVANCED SCOUT HELICOPTER, U. S. Army Agency for Aviation Safety, Fort Rucker, Alabama, January 1977. }

10. McDermott, J. M., and Vega, E., THE EFFECTS OF LATEST

MILITARY CRITERIA ON TIHE STRUCTURAL WEIGHT OF THE HUGHES

ADVANCED ATTACK HELICOPTER YAH-64, Journal of the American

Helicopter Society, Vol. 23, No. 4, October 1978, pp. 2-9.

11. Haley, J. L., Jr., CRASHWORTHINESS VERSUS COST: A STUDY

OF ARMY ROTARY WING AIRCRAFT ACCIDENTS IN PERIOD JANUARY

1970 THROUGH DECEMBER 1971, paper presented at the Aircraft Crashworthiness Symposium, University of Cincinnati,

Cincinnati, Ohio. October 1975.

12. Hicks, J. E., ECONOMIC BENEFITS OF UTILITY AIRCRAFT CRASHWORTHINESS, USAAAVS Technical Report 76-2, U. S. Army Agency for Aviation Safety!, Fort Rucker, Alabama, July 1976.

13. THE ECONOMIC BENEFITS OF CRAS]{WORTHINE•S AND FLIGHT SAFETY

DESIGN FEATURES IN ATTACK HELICOPTERS, USAAAVS Technical I Report 77-2, U. S. Army Agency for Aviation Safety, Fort

Rucker, Alabama, June 1977.

20

. , 
I

Ideally, it would seem most efficient to simply specify human

tolerance requirements and an array of vehicle crash impact conditions and then develop the helicopter as a crashworthy

system with a mixture of those crashworthy features that are

most efficient for the particular helicopter being designed.

Unfortunately, the validated structural and/or human tolerance

analytical techniques needed to perform and evaluate such a

maximum design freedom approach to achieving crashworthiness

are not available. Furthermore, testing complete aircraft sufficiently early in the development cycle to permit evaluation

of system concepts in time to permit design changes based on

the test results is not practical. The systems approach dictates that the designer consider probable crash conditions

wherein all subsystems cannot perform their desired functionsu

for example, an impact situation in which the landing gear cannot absorb its share of the impact crash energy because of aircraft attitude at impact. Therefore, to achieve the overall

goal, minimum levels of crash protection are recommended for

the various individual subsystems. A balance must be struck

between the two extremes of: (1) defining necessary performance on a component level only, and (2) requiring that the aircraft system be designed for an array of impact conditions

with no component design and test criteria.

Current helicopter crashworthiness criteria require that a new

aircraft be designed as a system to meet the vehicle impact

design conditions recommended in Volume III however, minimum

criteria are also specified for a few crash-critical components. For example, strengths and minimum crash energy-absorption requirements for seat and restraint systems are specified. All

strength requirements presented in this volume are based on the

crash environments described in Volume 11. Testing requirements are based on ensuring compliance with strength and deformation requirements. Mandatory minimum crashworthiness design criteria for U. S. Army light fixed- and rotary-wing aircraft

are stated in MIL-STD-1290(AV) (Reference 1). All pilot, copilot, observer, and student seats in either rotary- or light

fixed-wing aircraft should conform to the requirements of MILS-58095(AV) (Reference 14).

Although much higher levels of crashworthiness can be achieved

during the development of completely new aircraft designs, the

crashworthiness of existing aircraft can be significantly improved through retrofitting these aircraft with crashworthy

components adhering to the design principles of this design

14. Military Specification, MIL-S-58095(AV), SEAT SYSTEM:

CRASHWORTHY, NON-EJECTION, AIRCREW, GENERAL SPECIFICATION

FOR, Department of Defense, Washington, D. C., 27 August

1971.

21 a
guide. This can even be achieved while expanding the combat

effectiveness of the aircraft. An example of this is the successful program to retrofit all U. S. Army helicopters with a

crashworthy fuel system (Reference 15).

In an initial assessment, the definition of an adequate crashworthy structure may appear to be a relatively simple matter.

In fact, many influencing parameters must be considered before

an optimum design can be finalized. A complete systems approach (as summarized in Figure 1) must be employed to include

all influencing parameters concerned with the design, manufacture, overall performance, and economic restraint on the aircraft in meeting mission requirements. Tradeoffs among the

parameters must be made in order to arrive at a final design

that most closely meets the customer's specified requirements.

It must be remembered that for ,each type of aircraft, different emphasis will be placed'in the parameter mix. Table 1 summarizes major crashworthiness criteria that must be considered

during the prelimfnary design definition phase.

15. Cook, R. L., and Goebel, D. E., EVALUATION OF THE UH-lD/H

HELICOPTER CRASHWORTHY FUEL SYSTEM IN A CRASH ENVIRONMENT,

Dynamic Science, Division of Marshall Industriesi USAAMRDL

Technical Report 71-47, Eustis Directorate, U. S. Army Air

mobility Research and Development Laboratory, Fort Eustis,

Virginia, November 1971, AD 739567.

22

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2. DEFINITIONS

2.1 AIRCRAFT COORDINATE SYSTEMS AND ATTITUDE PARAMETERS

* Aircraft Coordinates

Positive directions for velocity, acceleration, and

force components and for pitch, roll, and yaw are

illustrated in Figure 2. When referring to an aircraft in any flight attitude, it is standard practice to use a basic set of orthogonal axes as shown

.in Figure 2, with x, y, and z referring to the longitudinal, lateral, and vertical directions, respectively.

+z

:' +

Yaw

F e A a o n a a u iiPitchin

,25

:;) ~~~~Rolliy . ...

¢.• +X

iX

Figure 2. Aircraft coordinates and attitude directions.

25
However, care must be exercised when analyzing ground

impact cases where structural failure occurs, aircraft geometry changes, and reaction loading at the

ground plane takes place. In the simulation of such

impacts, it is often necessary to use more than one

set of reference axes, including the earth-fixed system shown in Figure 2 as X, Y, Z.

e Flight Path Angle

The angle between the aircraft flJght path and the

horizontal at the moment of impact. The algebraic sign of the Flight Path Angle is positive if the aircraft is moving downward immediately prior to impact. The sign is negative if impact occurs while the aircraft is moving upward.

* Terrain Angle

The angle between the impact surface and the horizontal, measured in a vertical plane. The algebraic sign of the Terrain Angle is positive when the direction of flight is uphill, and negative when thd direction of flight is downhill.

* Impact Angle

The angle between the flight path and the terrain,

measured in a vertical plane. The impact angle is

the algebraic sum of the flight path angle plus the

terrain angle.

Flight

path Flight path Impact angle angleHorizontal

26
e Attitude at Impact

The aircraft attitude in degrees at the moment of

initial impact. The attitude at impact is stated in

degrees of pitch, yaw, and roll (see Figure 2).

2.2 ACCELERATION-RELATED TERMS

0 Acceleration

The rate of change of velocity. An acceleration is

required to produce any velocity change, whether in

magnitude or in direction. Acceleration may produce

either an increase or a decrease in velocity. There

are two basic types of acceleration: linear, which

changes translational velocity, and angular (or rotational), which changes angular (or rotational) velocity. With respect to the crash environment, unless otherwise specified, all acceleration values

are those at a point approximately at the center cf

the floor of the fuselage.

* Deceleration

Acceleration which produces a decrease in velocity.

* Abrupt Accelerations

Accelerations of short duration primarily associated

with crash impacts, ejection seat shocks, capsule

impacts, etc. One second is generally accepted as

the dividing point between abrupt and prolonged accelerations. Within the extremely short duration

range of abrupt accelerations (0.2 sec and below),

the effects on the human body are limited to mechanical overloading (skeletal and soft tissue stresses),

there being insufficient time for functional disturbances due to fluid shifts.

e The Term G

The ratio of a particular acceleration to the acceleration due ýo gravitational attraction at sea level

(32.2 ft/sec ). In accordance with common practice,

this report will refer to accelerations measured

in G. To illuistrate, it is customarily understood

that 5 G rfpresents an acceleration of 5 x 32.2, or

161 ft/sec

27

• • i'•
2.3 VELOCITY-RELATED TERMS

e Velocity Change in Major Impact (gv)

The decrease in velocity of the airframe during the ajor impat, expressed in feet per second. The

maJor impact is the one in which the highest forces

are incurred, not necessarily the initial impact.

For the acceleration pulse shown in Figure 3, the

major impact should be considered ended at time t

Elastic recovery in the structure will tend to

verse the direction of the aircraft velocity prior

to t . Should the velocity actually reverse, its

dire~ion must be considered in computing the velocity change. For example, an aircraft impacting downward with a vertical velocity component of 30 ft/sec

and rebounding with an upward component of 5 ft/sec

should be considered to experience a velocity change

4v - 30 - (-5) = 35 ft/sec

during the major impact. The velocity change during impact is further explained in Section 7.2.

* Longitudinal Velocity Change

The decrease in velocity during the major impact measured along the longitudinal (roll) axis of the aircraft. The velocity may or may not reach zero during

the major impact. For example, an aircraft impacting

the ground at a forward velocity of 100 ft/sec and

slowing to 35 ft/sec before rebounding would experience a longitudinal velocity change of 65 ft/sec

during this impact.

* Vertical Velocity Change

The decrease in velocity during the major impact measured along the vertical (yaw) axis of an aircraft.

The vertical velocity generally reaches zero during

the major impact.

* Lateral Velocity Change

The decrease in velocity during the major impact measured along the lateral (pitch) axis of the aircraft.

28
Peak

SAverage

00

I ,

ti 2- i ----

2.4 FOC TEM

'-4 I

A cvrash foc can be exrseda...lileo h

I I

II

i ti ----- 2tlt2t '. Tm

t 2 -t1

Figure 3. Typical aircraft floor acceleration pulse.

2.4 FORCE TERMS

* Load Factor

A crash force can be expressed as a multiple of the

weight of an object being accelerated. A load factor, when multiplied by a weight, produces a force

which can be used to establish static strength (see

Static Strength). Load factor is expressed in units

of G.

0O 29

S,-',.. . . .l. , 
* Forward Load

Loading in a direction toward the nose of the aircraft, parallel to the aircraft longitudinal (roll)

axis.

, * Aftward Load

Loading in a direction toward the tail of the aircraft, parallel to the aircraft longitudinal (roll)

Saxis.

i, * Downward Load

Loading in a downward direction parallel to the vertical (yaw) axis of the aircraft.

e Upward Loadd

Loading in an upward direction parallel to the vertical (yaw) axis of the aircraft.

, Lateral Load

Loading in a direction parallel to the lateral (pitch) axis of the aircraft.

e Combined Load

Loading consisting of components in more than one of

the directions described in Section 2.1.

* Crash Force Resultant

V The geometric sum of horizontal and vertical crash

forces: horizontal and vertical velocity components

4 i at impact, and horizontal and vertical stopping distances. The Crash Force Resultant is fully defined

by determination of both its magnitude and its direction. The algebraic sign of the resultant crash

force angle is positive when the line of action of

the resultant is above the horizontal, and negative if the line of action is below the horizontal.

Resultant G

crash orce r Vertical

R crash force

angle

h Horizontal crash force

30'.
4. Crash Force Angle

"The angle between the resultant crash force and the

longitudinal axis of the aircraft. For impacts with

little lateral component of force, the crash force

angle is the algebraic sum of the crash force resultant angle plus the aircraft pitch angle.

Aircraft

Sjq ~axis... longitudinal

fHorizontal fRt

anglee / anle angle angle

CrasResultant ical ,tc

crash force +

angl a eangle

,horizontal

2.5 DYNAMICS TERMS

• Rebound

Rapid return toward the original position upon release or rapid reduction of the deforming load,

usually associated with elastic deformation.

, Dynamic Overshoot

The amplification of decelerative force on cargo or

personnel above the floor input decelerative force

(ratio of output to input). This amplification is

a result of the dynamic response of the system.

31

0l•:. .2
0 Transmissibility

The amplification of a steady-state vibrational input

amplitude (ratio of output to input). Transmissibilities maximize at resonant frequencies and may increase acceleration amplitude similar to dynamic

overshoot.

2.6 CRASH SURVIVABILITY TERMS

, Survivable Accident

An accident in which the forces transmitted to the

occupant through the seat and restraint system do not

exceed the limits of human tolerance to abrupt accelerations and in which the structure in the occupant's

immediate environment remains substantially intact to

the extent that a livable volume is provided for the

occupants throughout the crash sequence.

* Survival Envelope

The range of impact conditions--including magnitude

and direction of pulses and duration of forces occurring in an aircraft accident--wherein the occupiable

area of the aircraft remains substantially intact,

both during and following the impact, and the forces

transmitted to the occupants do not exceed the limits

of human tolerance when current state-of-the-art restraint systems are used.

It should be noted that, where the occupiable volume

is altered appreciably through elastic deformation

during the impact phase, survivable conditions may

not have existed in an accident that, from postcrash

inspection, outwardly appeared to be survivable.

2.7 OCCUPANT-RELATED TERMS

* Human Body Coordinates

In order to minimize the confusion sometimes created

by the terminology used to describe the directions of

forces applied to the body, a group of NATO scientists

32

)
compiled the accelerative terminology table of equivalents shown in Figure 4 (Reference 16). Terminol- 2.: ogy used throughout this guide is compatible with the

NATO terms as illustrated.

Headward (+G) Direction of

z accelerative force

Vertical

Back to chest Headward - Eyeballs-down

Tailward - Eyeballs-up (sternumward) Lateral right

(+(+G ) Transverse

y Lateral right - Eyeballsleft

Lateral left - Eyeballsright

Back to chest - Eyeballsin

Lateral left Chest to back - Eyeballs-

(-Gy) to back out

(spineward) Note:

Tailward (Gx) The accelerative force on

(-GZ) the body acts in the same

direction as the arrows.

Ik ,

Figure 4. Terminology for directions of forces on the body.

9 Anthropomorphic Dummy

A device designed and fabricated to represent not

only the appearance of humans but also the mass distribution, joint locations, motions, geometrical similarities such as flesh thickness and load/deflection

properties, and relevant skeletal configurations much

as iliac crests, ischial tuberosities, rib cages, etc. All Attempts are also made to simulate human response of

major structural assemblages such as thorax, spinal

column, neck, etc. The dummy is strapped into seats

or litters and used to simulate a human occupant in

dynamic tests.

16. Gell, C. F., TABLE OF EQUIVALENTS FOR ACCELERATION TERMINOLOGY, Aerospace Medicine, Vol. 32, No. 12, December

1961, pp. 1109-1111.

33

0'.,: ..... , • .,!
e Human Tolerance

For the purposes of this document, human tolerance is

defined as a selected array of parameters that describe a condition of decelerative loading for which

it is believed there is a reasonable probability for

survival without major injury. As used in this volume, designing for the limits of human tolerance refers to providing design features that will maintain

these conditions at or below their tolerable levels

to enable the occupant to survive the given crash

environment.

Obviously, the tolerance of the human body to crash

environments is a function of many variables including the unique characteristics of the individual person as well as the loading variables. The loads ap- plied to the body include decelerative loads imposed by seats and restraint systems as well as localized

forces due to impact with surrounding structures.

Tolerable magnitudes of the decelerative loads depend

on the direction of the load, the orientation of the

body, and the means of applying the load. For example, the critical nature of loads parallel to the

occupant's spine manifests itself in any of a number

of spinal fractures, but typically, the fracture is

an anterior wedge, or compressive failure of the

front surface of a vertebra. Forces perpendicular

to the occupant's spine can produce spinal fracture

through shear failures or from hyperflexion resulting,

for example, from jackknife bending over a lap beltonly restraint. The lap belt might inflict injuries to the internal organs if it is not retained on the

pelvic girdle but is allowed to exert its force above

the iliac crests in the soft stomach region. Excessive rotational or linear acceleration of the head

can produce concussion. Further, skull fracture can

result from localized impact with surrounding structure. Therefore, tolerance is a function of the method of occupant restraint as well as the characteristics of the specific occupant. Refer to Chapter 4

of Volume II for a more detailed discussion of human

tolerance.

2.8 STRUCTURAL TERMS

e Airframe Structural Crashworthiness

The ability of an airframe structure to maintain a

protective shell around occupants during a crash and

34

. .
to minimize accelerations applied to the occupiable

portion of the aircraft during crash impacts.

e Structural Integrity

The ability of a structure to sustain crash loads

without collapse, failure, or deformation of sufficient magnitude to: (1) cause injury to personnel,

or (2) prevent the structure from performing as intended.

* Static Strength

The maximum static load that can be sustained by a

structure, often expressed as a load factor in terms

of G.

"e Strain

The ratio of change in length to the original length

of a loaded component.

* Collapse

Plastic deformation of structure to the point of loss

of useful load-carrying ability. Although normally

considered detrimental, in certain cases collapse can

prove beneficial as a significant energy-absorbing

process, maintaining structural integrity.

* Failure

Loss of load-carrying capability, usually referring

to structural linkage rupture.

* Limit Load

In a structure, limit load refers to the load the

structure will carry before yielding. Similarly, in

an energy-absorbing device, it represents the load at

which the device deforms in performing its function.

* Load Limiter, Load-Limiting Device, or

Energy Absorber

These are interchangeable names of devices used to

limit the load in a structure to a preselected value.

These devices absorb energy by providing a resistive

force applied over a deformation distance without

significant elastic rebound.

35

6|

i .
i

0 Specific Energy Absorbed (SEA)

The energy absorbed by a energy-absorbing device or

structure divided by its weight. SEA is usually presented in inch-pounds per pound.

* Bottoming

With respect to energy-absorbing structure, bottoming is a condition in which the deforming structure

or material becomes compacted and the load increases

rapidly with very little increased deformation.

* Bulkhead

A structural partition extending upwards from the

floor and dividing the aircraft into separate compartments. Seats can be mounted to bulkheads instead of the floor if sufficient strength is pro- vided.

* Basic Structural Design Gross Weight (BSDGW)

The structural design gross weight is cited in the

MIL-STD-1374, Part I, "Group Weight StatementDimensional and Structural Data", and is further

explained in the detail system specification for

the aircraft.

36

- ,'i*.. . . . • , ,:.. 
3. GENERAL DESIGN CONSIDERATIONS

AIRCRAFT TYPES - THEIR MISSIONS AND DESIGN RESTRAINTSS3.1 

The mission of Army aviation is to augment the capability of

the Army to conduct prompt and sustained combat incident to oporations on land.

Army aviation supports the Army's ground combat function in the

following areas:

9 Command, Control, and Communications. Army aviation

support includes courier and liaison missions, control of vehicular columns, message drop and pickup, command and control of airmobile operations, wire

laying, and radio relay.

* Intelliaence. Army aviation is an important means

of gathering intelligence. It provides aerial "eyes"

over the battlefield and conducts missions in support of aerial survey operations, aerial radiological survey, and target acquisition.

SMobilit. By airlifting troops and combat equipment, Army av ation provides an additional means of maneuver to the ground commander. Using Army aviation's

airmobile capability, weapons may be emplaced rapidly and troops may be carried quickly over obstacles.

* Firepower. Army aviation, which provides aerial adjustment of indirect fires, is expanding the use of

Army aircraft as weapon platforms to fill the gap

between the support provided by conventional ground

fire support and close air support.

9 Combat Service Support. Army aviation supports logistical operations by providing delivery of

troops and equipment and evacuation of casualties

and damaged equipment within the Army combat zone.

3.1.1 Helicopters

The U. S. Army inventory of helicopters can be divided into

five types by mission:

* Observation (OH) * Cargo (CH)

* Attack (AH) * Training (TH)

* Utility (UH)

37
Figures 5, 6, and 7 show side elevations for typical helicopters to indicate comparative sizes and layouts. When considering a design for occupant protection, it becomes apparent

that each type of helicopter poses different problems. The

size, proximity of occupied areas to the ground impact plane,

distribution of mass items and external stores, and the location of transparencies and cutouts for exits combine to provide the designer with major challenges in the efficient allocation

of primary structure.

Figure 8 shows a helicopter whose design considered the parameters listed above. Notations indicate several major design features that must be incorporated into a primary structural

matrix to satisfy the crashworthiness requirements of MIL-STD1290(AV) (Reference 1).

3.1.2 Fixed-Wing Aircraft

Side elevations of these aircraft types are shown in Figure 9.

All, except for the OV-l, are military versions of commercial

aircraft. The maximum capacity of any listed aircraft is a

crew of two, with 20 passengers.

3.2 APPLICABILITY OF CRITERIA

Information presented in this volume is generally applicable

to all types and sizes of aircraft. The different operating

characteristics and mission requirements of different aircraft

types provide some likelihood of different crash conditions, which should be considered in designing for crashworthiness;

however, experience in survivable accidents indicates that the

impact environment--considering acceleration pulse shape, magnitude, and direction--is similar for all types of existing

light fixed-wing and rotary-wing aircraft except for lateral

impact. For lateral impacts, cargo and attack helicopters are

grouped with fixed wing, while other helicopters appear to have

a somewhat more severe environment. Where aircraft size or

weight is an important factor, its effect is discussed. Whenever quantitative data concerning likely crash conditions or

design requirements for improving crashworthiness in either

existing or new aircraft are available, they are presented.

3.3 ACCEPTANCE OF STRUCTURES

A standard airframe structure requirement has always been that

the structure be capable of withstanding all loading conditions

that may reasonably be expected to occur during any flight or

ground handling operation. As understanding of various factors

affecting airframe requirements, such as metal fatigue, has increased, the design criteria have been changed to reflect this

38
UH- 1

UH- 60

JI

CH-47

0 0 0

CH-54

Figure 5. Side elevations of typical U. S. Army,

utility and cargo helicopters.

39
OH- 6

OH-58

AH- 1

YAH- 64

ti Figure 6. Side elevations of typical U. S. Army

observation and attack helicopters.

40
TH-55

. Figure 7. Side elevation of typical U. S. Army training helicopter.

new knowledge. Often, in designing aircraft structure, stringent requirements must be met even though it is understood

that every available means will be used to avoid those design

conditions. For example, aircraft are designed to withstand

severe gust loads, even though it is never intended that they

be deliberately flown into severe turbulence.

Both experience and reason indicate that as long as aircraft

are flown, there will be accidents, and that these accidents

will impose conditions that seriously threaten occupant survival. Although much effort will be directed toward avoiding

crashes, acceptable aircraft structures should always provide

the greatest possible degree of occupant protection from crash

conditions. All available information should be considered in

designing new aircraft structures to ensure that new designs

will be acceptably crashworthy, as well as airworthy.

"3.4 SELECTION OF STRUCTURES

The final design of a particular aircraft is the result of a

series of compromises with respect to aerodynamics, strength,

simplicity of fabrication, economics, etc. The additional re-

* quirement that crashworthiness become an important structural

consideration will bring a need for further compromise and for

good judgment in arriving at the compromise position. As more

attention is directed toward airframe crashworthiness throughout the design stages, methods and techniques of construction

will improve so that crashworthiness can be achieved without

prohibitive weight and performance penalties.

41 giii
High-strength

retention of

main transmission

High-strength

retention of htrength

rotor bead retention of

rotorheadengine (2)

assembly

Rollover

structure

Jettisonable

external stores

(wing mounted)

Crashworthy

Crashworthy landing gear "pilot seat

(armored)

Crashworthy

gunner seat

(arm%ýred)

Figure 8. Selected helicopter crashworthy features.

When there is a choice between two designs which both adequately meet normal structural requirements, then the design

which offers better crash protection should be chosen. Often,

the upgrading of crashworthiness will not result directly from

an increase in strength. Consideration of the deformations

which are likely to occur, and ensuring that several parallel

load paths are available to keep the structure intact even

though localized damage occurs, will also improve crashworthiness.

42
U-10

U- 3

IJ-21

If /1 7777 ý 777 11 1// 1 1 11111771711 12

Figure 9. Side elevations of typical U. S. Army

fixed-wing aircraft.

43
C-12

AIL

Iii

LJV-18

ov-1

Figure 9 (contd). Side elevations of typical U. S. Army

fixed-wing aircraft.

44

I -,Wig.
Excessively strong airframe structure is no more acceptable for crashworthiness than understrength structure, Not only

will unnecessary strength impose a highly undesirable weight 1 ., penalty, but also an excessively strong structure jan develop high loads which may produce high accelerations of the aircraft

during impact. These high-magnitude accelerations place severe

demands upon occupant restraint systems. If these demands are

not met, the extra structural strength can result in reduced

survivability.

Structures which have insufficient strength, on the other hand,

can permit occupant injuries due to loss of the protective

shell.

3.5 TESTING

The present-day state of the art in analysis of structural behavior under crash impact conditiors is such that accurate pre- diction of location and modes of collapse and failure and prediction of impact forces, accelerations, and deformations are

not possible. For this reason, the use of full-scale dynamic crash testing to complement and substantiate analytic determination of airframe behavior is highly recommended. Such

crash tests should be conducted under conditions similar to a

severe survivable crash which could be expected to occur in

service. When full-scale crash tests are conducted, test results should be carefully studied to provide information for

design improvement and for background to improve future dosigns. During the development of new aircraft, testing should

be conducted to demonstrate that the required design strengths

are met.

In many instances, static and/or dynamic testing of only one

component of the structure is necessary for validating certain

aspects of crashworthiness of the airoraft; however, full-scale

tests of the complete aircraft are recommended for proof of

compliance for fuselage and related structures such as landing

gear, engines, transmissions, and seat tiedown provisions.

45
4. CRASH ENVIRONMENT

4.1 INTRODUCTION

Aircraft impact conditions can vary considerably depending on

the vehicle's attitude and velocity components, and on the

characteristics of the impacted terrain.

A statistical analysis of accident data, summarized in Volume

II, was used to define velocity changes occurring in the major

impact for the 95th-percentile potentially survivable crashes.

In addition, estimates of acceleration pulses experienced at

the -abin floor level near the center of gravity of the aircraft were made for use in the definition of seat-occupant and

cargo design environments. The data were obtained from investigations of light fixed- and rotary-wing aircraft accidents

from 1960 to 1965 and from 1971 to 1976. The aircraft involved

did not contain crashworthy structures, energy-absorbing seats,

or orashworthy landing gear, thereby tending to increase occupant injury potential due to high accelerations at the floor

plane in conjunction with seat collapse or breakaway. These

data are summarized in Table 2.

The resultant velocity change for combined longitudinal, vertical, and lateral components of the 95th-percentile survivable

accident of rotary- and light fixed-wing aircraft does not exceed 50 ft/sec. The vertical or lateral components do not exceed thoir individual 95th-percentile values, i.e., 42 ft/sec

vertically, and 25 and 30 ft/sec laterally for light fixed-wing

aircraft and attack and cargo helicopters, and for other helicopters, respectively.

Since deceleration loads are a function of the strength of the

structure, a systems analysis should be performed to establish

the distribution of energy-absorbing properties to the gear,

fuselage, seats, etc. For new aircraft, the elements can then

be designed to provide the required optimum properties. For

older aircraft, certain elements in the chain cannot be changedl

thus those that can be changed should be designed to make up

for the deficient elements. For these older aircraft, which

have not been subjected to a systems analysis, and for those

in which certain key energy-absorbing elements, such as landing gear, might be lost before the element has absorbed its

allotment of the crash energy, standard design environments

have been established. These design environments are defined

in terms of triangular acceleration pulses having the required

peak accelerations, durations, and velocity changes and are

presented in Table 2. For the situations described above,

46

i i.. . ...... i i ~ ii .... ~ ~~~~~~~iOl U-l•"i ik
TABLE 2. SUMMARY OF CRASH IMPACT CONDITIONS FOR HELICOPTERS

AND LIGHT FIXED-WING AIRCRAFT DESIGN

Impact Pul e

direction Velocity Peak duration,

(aircraft change, Av acceleration At

axes) (ft/sec) ( G___ (sac) Comments

Longitudinal 50 30 0.104 Triangular (Cockpit) deceleration pulses

Longitudinal 50 24 0.130

(Cabin)

Vertical 42 48 0.054

Lateral 2 5 a 16 0,097 At calculated from

3 0 b 18 0.104 known or

assumed

values for

G and Avi

At -

peak

a) Light fixed-wing aircraft, attack and cargo helicopters. b) Other helicopters.

theme pulses should be used for the design of restraint my.-

teoms, seats, cargo restraint, and other items inside the air-

"craft.

Certainly in the case of retrofit, the maximum crashworthiness

possible within the restraints of the existing hardware should

be provided. This may mean that a less severe design environment might be necessary for the achievement of maximum crashworthiness in a practical vehicle.

47
Figure 10 shows plots of combined longitudinal, lateral, and

vertical velocity changes for helicopters to be used in determining intermediate velocity change components. For light fixed-wing aircraft and cargo and attack helicopters, Figure 10(b) will still be correct, but (a) and (d) must be altered

for a lateral velocity change of 25 ft/sec instead of 30 ft/sec.

Figure 10(a) represents a three-dimensional display of the resultant velocity changes defined in 10(b), 10(c), and 10(d).

In general, the three components are related by the equation

Vx2 +vy 2 + .v2 V 2 (2V)

where vx - longitudinal velocity change, ft/sec

vy a lateral velocity change, ft/sec

vz - vertical velocity change, ft/sec

v - resultant velocity change, ft/sec

and the axes are those illustrated in Figure 2. The curves

terminate at 15 degrees above the x-y plane of the aircraft.

As mentioned in Section 5.3.1.4, impact accelerations for design of the overhead structure have been determined. However,

these conditions are intended for aircraft inverting after impact, not those impacting in an inverted attitude. Therefore,

no design velocity changes are presented here for impact on the

upper parts of the aircraft.

It should be remembered that human survival is determined by acceleration magnitudes, durations, and rates of acceleration

change actually experienced by the body, rather than by the

aircraft velocity change. The preceding values for velocity change limitations act only as a guide for establishing crash

energy content. A combination of structure, landing gear, and

seat that allows a long stroking distance to absorb energy

may well provide survival capability in crashes with velocity

changes in excess of those shown.

The velocity change, acceleration, and pulse duration are useful in defining occupant impact loading environments in a crash

sequence. However, other structural factors that can influence

human survival in a crash impact must also be considered:

48

S. . . .. . . .
Lon'

C4.

f FA

4.4

rd

CI

01M

49N
a Structural collapse and support of large mass items.

: Structural elastic deformation that may intrude into

occupied areas.

• Structural penetration by aircraft components or external agents.

* Structural strength to ensure postcrash operation of

emergency exits with adequate egress potential.

a Adequate structural strength to provide support for

the seat and restraint system combination, as well

as the landing gear.

Helicopters are often used in areas where fixed-wing aircraft

cannot function and, of necessity, spend time maneuvering between obstacles at low forward velocities. Under such conditions, power failure at low altitude results in a predominantly

vertical impact. Blade or fuselage impact with ground obstacles, such as overhead wires, trees, or buildings, can generate

aircraft rotation with ground impacts occurring at roll attitudes as severe as 180 degrees; that is, completely inverted.

(See Volume II, Section 3.4 for further information pertaining

*1 to helicopter attitude upon impact.)

Fixed-wing aircraft can impact with high vertical velocity if

a stall occurs close to the ground plane with insufficient altitude to regain control. Impacts with a predominantly longitudinal velocity vector occur when aircraft are flown into inclined surfaces, mountains, ground obstacles, or when ground

impact occurs with the aircraft in an extreme nose-down diving

attitude. When an impact occurs with an obstacle located near

the ground, i.e., overhead wires, trees, or buildings, the subsequent rotational motions can result in ground impact occuring at almost any attitude.

After tho initial impact, subsequent aircraft decelerations,

rotations, and secondary impacts are generally less severe for

occupants. However, during this phase of the crash sequence,

items such as rotor blades, transmission assemblies, engines,

and external stores may become detached from the aircraft.

These items may represent subsequent hazards if they impact occupied aircraft se-jtions or cause postcrash fires by releasing

flammable fluids in the presence of an ignition source. During

the crash sequence, external agents that may poinetrate occupied

aircraft sections represent additional potential hazardsl these

include trees, ground equipment, metallic structures, etc.

50 S!)

, .I;
Therefore, occupants must be protected not only from initial

impact pulses, but also from secondary impacts, rotations, 4K ! penetrations in critical sections, fire, and water hazards.

4.2 STRUCTURAL DAMAGE WHICH FREQUENTLY RESULTS IN OCCUPANT

INJURY

Fixed-wing conventional aircraft and vertical takeoff and landing aircraft tend to experience similar structural loading for

similar impact conditions, although the distribution of accident types may vary considerably between these two generic

groups. The structural damage that produces occupant injury

is generally the same for both types of aircraft. Structural

damage in severe accidents cannot be avoided. However, improvements in airframe structure and optimization of element

distribution can work to control the manner in which structural damage occurs so that a survivable environment is more

likely to be maintained.

Usually, the structure that first contacts the impact surface

is the first to begin to deform. This localized deformation

continues until the kinetic energy of the aircraft is absorbed

at low loads over relatively large distances, or until there

is enough structure involved in the deformation to produce a

significantly high decelerative force on the aircraft mass. If

the quantity of kinetic energy to be absorbed is small, structural damage may be minor, and the aircraft may simply come to

rest without endangering occupants. When the initial kinetic

energy is high, there is a grea.er likelihood that structural

damage will be severe and that forces will build up until total

aircraft decelerative forces become large. Once these high

decelerative forces are reached, buckling throughout the aircraft may occur. The protective shell is then compressed between the impact surface and masses aft or above the protective

shell. It may become desirable to reduce the cabin deformation

by allowing parts of the aircraft, such as wings and tail sections of fixed-wing aircraft, to break free from the cabin section during the impact. However, while this may reduce cabin

deformation, it may produce no significant reduction of crash

forces. This would mean that a higher acceleration level might

be experienced within the cabin. The effects of mass reduction

upon energy-absorption requirements for protective shell structure and its effects upon acceleration levels within the cabin

are discussed further in Sections 6.6.5.5 and 7.2.5.

The following subsections define the general response of an

"aircraft for different impact conditions and injury-causing

events that have frequently occurred.

51

, ..
4.2.1 Longitudinal (Crushing) Loads on Cockpit [-:ructure

During longitudinal impact against soft earth, the aircraft

nose structure is sometimes deformed in such a way that it

forms a scoop that picks up earth as the aircraft slides along

in contact with the impact surface. When this occurs, the

scooped earth must be quickly accelerated to the velocity of

the aircraft. This impulsive acceleration of the scooped-up

earth mass produces momentary high forces that must be reacted

by aircraft structure near the impact point. Often, the forward cockpit bulkhead must support this load, resulting in collapse of cockpit structure, entrapment of occupants, and injury

to occupants' lower extremities. In addition, the high forces

produce high aircraft accelerations, resulting in high loads on

personnel and cargo restraint systems.

Sometimes in high-velocity accidents, the combination of nose

•,structure crushing and friction between the structure and the

teiiain (particularly in "long-nose" aircraft) causes the forwaid structure to be pulled beneath the rest of the aircraft.

This type of damage also causes rupture of the cockpit floor,

and results in higher longitudinal acceleration than would be

experienced if a smooth skid were maintained under the nose.

Longitudinal crushing also occurs during a crash in which a

high angle of attack exists between the aircraft and the obstacle against which it crashes. This can result from a shallowangle impact with terrain features such as a hillock or bank,

or from a steep-flight angle impact with respect to relatively

flat terrain. In these crashes, the aircraft nose usually

crushes a sufficient distance to destroy the occupied section

of the aircraft, thus providing a very poor chance for occupant

survival.

4.2.2 Vertical (Crushing) Loads on Fuselage Shell

Collapse of the protective shell due to vertical loading often

occurs in high-sink-rate accidents or rollover accidents. The

collapse is often aggravated by the attachment of large masses

to fuselage structurel these masses might be engines, transmissions, and rotor mechanisms in rotary-wing aircraft and high

wings in fixed-wing aircraft. This damage results in loss of

occupiable volume and crushing injuries or entrapment of occupants.

Also, if insufficient energy-absorbing structure is provided,

or if the underfloor structure crushes only at excessively high

loads, these high loads are transmitted to the occupants, re- sulting in compressive spinal and associated injuries.

52
4.2.3 Lateral (Crushing) Loads on Fuselage Shell

Lateral impact of utility helicopters occurs frequently and

produces a hazardous environment. A 1971 study showed that

over half of severe utility helicopter crashes result in rollover or side impact (Reference 17). Eyewitnesses of crashes

in which landing or hovering helicopters caught rotor blades

on trees or other obstacles report that the helicopters tend

to flip on their sides and rise to a height of approximately

15 ft before crashing. Since the sides of the fuselage are not

usually designed for crash protection, severe injuries can result from relatively minor accidents.

Occupants are placed close to the sides of the fuselage, and

often their restraint systems, such as lap belts used alone or

with gunner tethers, are not adequate to restrain the occupants

laterally. On many occasions, the doors have been removed previously or are lost during the crash. The occupant is then exposed to a variety of hazards, including being bodily ejected

and crushed, having protruding body parts entrapped and mangled, or being violently impacted against the side of the

fuselage. Injuries and fatalities then can result from crushing, dismemberment, loss of blood, excessive G loads, entrapment or debilitation, and exposure to fire and other postcrash

hazards.

4.2.4 Transverse (Bending) Loads on Fuselage Shell

Rupture or collapse of the protective shell often occurs due

to the high bending loads during rapid pitch change associated

with longitudinal crashes at moderate-to-high impact angles.

Rupture of the protective shell exposes occupants to injury

through direct contact with the impact surface, contact with

jagged metal, and loss of restraint. Miscellaneous equipment

also may strike occupants after breakup of the aircraft.

4.2.5 Deformation (Buckling) of Floor Structure

Breakup of ±loor structure often indirectly accounts for occupant injury. In most aircraft, odcupant and cargo restraint

depends heavily upon the integrity of the floor structure.

When this structure fails, restraint is lost. Often floor

failure is caused by crushing of underfloor supporting structure. Localized damage is frequently caused when fuselagemounted landing gear are driven into the floor structure.

17. Haloy, J. L., ANALYSIS OF EXISTING HELICOPTER STRUCTURES

TO DETERMINE DIRECT IMPACT SURVIVAL PROBLEMS, U. S. Army

Board for Aviation Accident Research, Fort Rucker, Alabama, 1971.

53

C),
4.2.6 Landing Gear Penetration of Fuselage Shell

Landing gear failures often result in personnel injuries,

either directly (as mentioned above) or indirectly, through

fire exposure caused by rupture of ignitable fluid lines and

Lanks. Attention must be focused on a landing gear design and

location that will prevent penetration into livable space or

into any part of flammable fluid systems.

4.2.7 Helicopter Lateral Rollover

In helicopter accidents, rollover invariably causes the mai-n

rotor blades to strike the ground. This contact involves two

potential hazards: displacement of the transmission and intrusion of blades into occupied areas. "

Transmission displacement is controlled basically by the

strength of its mounts. The problem has been found to be reduced in aircraft with fully articulated main rotor hubs, where

the blades tend to destroy themselves without transferring excessively high crash loads to the transmission. However, a

high inertia main rotor affords a pilot some additional margin

during a powerless autorotation, thereby reducing the frequency

and severity of accidents. The majority of Army helicopters

have a high inertia main rotor, so it must be assumed that

blade impact loads will be transferred to the transmission and

reacted by :.3s mounts.

Hazardous blade intrusion is minimized by using a combination

of longitudinal and lateral beams in the cockpit overhead

structure. The lateral beams can act to deflect the rotating

blades while the longitudinal members provide a continuous support.

4.2.8 Rupture of Flammable Fluid Containers

Rupture of structure surrounding ignitable fluid containers or

transfer lines is often an indirect cause of occupant injury

as a result of postcrash fire. Structural members surrounding

flexible (bag) fluid containers should not fracture in a manner

that causes penetration of the container. Penetration resistance may also be improved by the use of a foam liner, as described in Section 6.6.5.8.

54

)
0£

5. DESIGN REQUIREMENTS

5.1 INTRODUCTION

A crash can involve a wide range of dynamic conditions, from a

simple unidirectional impact to a complex combination of rotational and multidirectional impact conditions. The current requirements for Army light fixed- and rotary-wing aircraft are

summarized in Table 3. Any light aircraft designed to similar

criteria would exhibit improvements in crashworthiness.

"When a crash does occur, the service life of the aircraft is

usually ended, and the only structural requirement is to provide occupant protection. In order to accomplish this, structure must be allowed to crush and deform in a controlled, predictable manner so that forces and accelerations imposed upon

occupants will be minimized while still maintaining the pro- tective shell. This means that any anaP' ',is for crashworthiness must consider the large deflections of structural members

and joints as well as loading in the plastic range of stress.

Excessively strong airframe structure is no more acceptable

than understrength structure for crashworthiness. Not only will unnecessary strength result in an unacceptable weight penalty, but on impact, high G levels that compromise occupant

survivability may be generated.

* 5.2 GENERAL REQUIREMENTS

Aircraft systems should be designed to prevent occupant fatalities and minimize the number and severity of occupant injuries during crash impacts of the severity of up to those defined in

*: Chapter 4 while minimizing, to the maximum extent practical, aircraft damage. This should be achieved by selecting the most

* effective mix of the crashworthiness factors listed below and

by complying with the detail requirements cited in Section 5.3.

Design impact velocity changes are presented in Table 2. These

velocity changes are for the major impact, assumed to occur on

a rigid surface and with a triangular acceleration-time pulse

shape.

Probability of occupant survival during crash impacts will be

increased if proper attention is given to the following features that influence crash survivability during the initial design stages of the aircraft system:

* Airframe protective shell

* Tiedown strength

e Occupant acceleration environment

55 ,tc'
YASLE 3, PEPORNMANCI URQUXPJNINST FOR STRUCTURAL CRA5MWORTKXNI|S

VelooLt. Vehicle Percentage d, at mpacted differential attitude volume Ot~her Dkta

ta urface -.. laUUcL. limits reduction reauirs"nt ... source

Longitudinal W.gid 20 No haszard Does not Lmped posteorash Volume 11

Sto Tilot/ egress Oop lot

40 15 max, Inward buckling of side NIL-ITD-100

length re- wall* should not po:e Volume is

duat4ion for hazards

pas.l/troop

compartment

Lateral Rigid 30 *200 Yaw 15 max, Lateral collapse of cc- NIL-ITD-1200

Width cupied armIa not hazard- Volume 1I

reduction Gul, l. entrapment of limbs,

Vertical Rigid 42 42S5/-151 15 max, 0 loads not injurious to NIL-ITD-1290 Pitoh height red. occupants Volume It *20, Roll In page./

troop oam- partment

Resultant Rigid So Combination As above Max, velocity changesl NIL-ITD-1290 for various long, s 90 1t6te0 Volume It components vex', V 42 ft/lea 5

lat. 3 30 ft/Ceob 25 ft/sec.,

Rollover Nor'th - 000 ideward Minimal forward fuSelage buried to NfL-STD-12IO or l10" in- (door depth of 2 in. Jinverted or

vrted or hatches eta. an side). Load uniformly an inter- ammumed to distributed over forward meiate angle be non-load 251to0 ocoupLed fiselage

carrying) length. Can eustain 4 0

without injury to stated and restrained occupants.

All loading dirgetione be- tween normal and parallel to skin to be coitndered.

Rollover <post- Rigid two 140, Is mAx. NIL-5Fb-12SO impact) roll. (max.) volume reduction (15 desired)

*rth plowing Earth Preclude plowing when for- NIL-ITD-13*0 a macaping ward 2s% of fuselage has (longitudinal) uniformly applied vertical

load of 0 Oend rearward load of 4 a or the ditching load of NIL-A-OOIIS5A, whichever is the greatest. Landing gear Rigid 20 *l0* Roll None, Plea- Airoraft deceleration at MIL',ITD-1290 t101 Pitch tic deforms- normal U.N. for impact

tion of gear with no fuselage to ground and mounting contact. All other A/C system al- etructural parts, except

lowable blades, ehnild be flightworthy following crash.

Landing gear $od 100 long, 0 -50 Pitch 15 max. No rollover, or it rollover NIL-BTD-1I00 14 vprt. *10' Roll volume re- occur•, two 3600 rolle

*20' Yaw duction (15 without fuselage crushing Volume It

desired)

al Light fixed-winy aircraft, atteck and cargo helicopters,

b other heliaopterm. a) Velocity at opact, not differential,

56

I/
e Occupant environment hazards

a Postcrash hazards.

5.3 DETAIL REQUIREMENTS

5.3.1 Airframe Crashworthiness

The aircraft structure should provide a protective shell for

vehicle occupants in crashes of the severity cited in Table 21

moreover, the structure should allow deformation in a controlled, predictable manner so that forces imposed upon the occupant will be minimized while still maintaining the protective

shell. In structural areas where large structural deformations.'

are anticipated, joints and attachments should be designed to

withstand large angular deflections and/or large linear displacements without failure. All exterior surfaces and all

structures which could be exposed to contact with the impact

surface should be constructed of materials which characteristically resist sparking as a consequence of abrasion. Unless

otherwise stated herein, the aircraft basic structural design

gross weight (BSDGW) should be used for the ve'icle weight in

the analyses described below. Directions are assumed with respect to the aircraft (Figure 2) unless otherwise stated.

5.3.1.1 Longitudinal Impact

5.3.1.1.1 Impact Conditions: The basic airframe ahould be

capab)le of impacting longitudinally into a rigid abutment or

wall at a contact velocity of 15 ft/sec without crushing the

pilot and copilot stations to an extent which would either preclude pilot and copilot evacuation of the aircraft or otherwise be hazardous to the life of the aircraft occupants. For

such an impact, the angine(s), transmission, and rotor system

for helicopters should remain intact and in place in the aircraft except for damage to the rotor blades. The basic airframe's capability to impact longitudinally into a rigid abutment or wall at a contact velocity of 40 ft/sec without reducing

the length of the passenger/troop compartment by more than 15

percent shall be demonstrated analytically. Any consequent

inward buckling of walls, floor, and/or roof should not be hazardous to the occupants and/or restrict their evacuation. The

aircraft should also be designed to withstand impact as in a

"A, low angle, missed approachl the impact conditions of this type

accident are illustrated in Figure 11. These impact conditions

in plowed soil can result in a rollover, and rollovers can be

critical for inward crushing and/or separation of the fuselage

as shown by past experience. The volume of the cockpit for the

occupied passenger/troop compartment Lhould not be reduced by

more than 15 percent (5 percent desired) for these conditions.

57

.... . .
IMPACT CONDITIONS

1. Assume loose plowed soil

2. Aircraft pitch (#) 5 degrees nose down

3. Aircraft roll (a) = +10 degrees 4. Aircraft yaw (7) = +20 degrees

5. Flight path angle (a) - 8 degrees

6. Impact airspeed - 60 knots

x

X Ground level

Figure 11. Low angle impact design conditions

(simulated approach with antitorque loss under poor visibility).

Should the aircraft turn over, the fuselage container should

maintain structural integrity for a minimum of two 360-degree

rolls. The static loads to be considered for rollover analysis

are described in Section 5.3.1.4.

5.3.1.1.2 Earth Scooping Effects: Design features for reducing the earth scooping effects encountered in longitudinal

impacts should include the following:

a Provide a large, relatively flat surface in those

areas which could otherwise gouge or plow, thereby

increasing the aircraft's tendency to slide over the

impact terrain.

e Minimize inward buckling of the fuselage nose or engine nacelle for the purpose of maintaining skid surface integrity.

* The nose section should be designed to preclude any

earth plowing and scooping tendency when the forward

25 percent of the fuselage has a uniformly applied

local upward load of 10 G and an aft load of 4 G, as

shown in Figure 12.

58 

I
tI

4G "

l0OG

Figure 12. Nose section design conditions.

5.3.1.1.3 Buckling Effects: To minimize hazards to personnel

created by buckling of the structure, the aircraft should be

designed to:

9 Provide sufficient strength of structure to prevent

"bending or buckling failure of the fuselage.

e Position personnel away from likely fuselage fracture areas.

, *o Have the fuselage buckle outward, if at all possible,

rather than inward into living space when its collapse strength has been exceeded.

, Include cargo tiedowns that will restrain cargo

should fuselage bending failure occur.

3A

5.3.1.1.4 Floor: The floor structure should possess sufficient strength to carry, without failure, loads applied by the

occupants and cargo restraint systems in impacts of the severity cited in Table 2. %i;I

5.3.1.2 Vertical Impact

5.3.1.2.1 Impact Conditions: The aircraft should possess the , capability to withstand an impact velocity of 42 ft/sec vertically, with respect to the ground, without reducing the height

of the cockpit and passenger/troop compartments by more than

15 percent and/or causing the occupants to experience injurious

accelerative loading. For this analysis, the aircraft orientation (attitude) upon impact should be any attitude within

÷25/-15-degrees pitch and +20-degrees roll.

59

-..... .'r/' .i"
5.3.1.2.2 Design Application: Design applications for accomplishing the above goal should include the following:

* To the greatest extent feasible, locate massive items

in lower fuselage locations rather than in upper

fuselage locations.

• Increase cockpit and cabin vertical strength and

stiffness to prevent the structure from crushing

occupants.

• Provide crash-force attenuating structure beneath

cockpit/cabin flooring.

* Provide load-limiting landing gear capable of absorbing as much of the crash energy as practical.

5.3.1.3 Lateral Impact: The aircraft should have the capability to withstand Tateral impacts into a rigid barrier/wall of

25 ft/sec for light fixed-wing and cargo and attack helicopters

and 30 ft/sec for other rotary-wing aircraft without reducing

the width of occupied areas by more than 15 percent or permitting the lateral collapse of occupiable portions of the aircraft to an extent that would be hazardous to life. Precaution

should be taken during design of the vehicle to minimize the

chance of the occupant or his extremities being trapped between

the structure and any impacting surfaces following failure of

doors, canopies, or hatches.

5.3.1.4 Rollover Impacts: The aircraft should be designed to

resist an earth impact loading as occurs when the aircraft

strikes the ground in either a 90-degree (sideward) or 180-

degree (inverted) attitude. A rollover accident should not

cause an injury due to structural intrusion into occupied

areas. It should be assumed that the forward fuselage roof

is buried to a depth of 2.0 in. in soil for the inverted attitude and that the load is uniformly distributed over the forward 25 percent of the occupiable fuselage length. It should

also be assumed that the forward fuselage side is buried to a

depth of 2.0 in. in soil for the sideward attitude and that the

load is uniformly distributed over the forward 25 percent of

the occupiable fuselage length. The fuselage should be capable

of sustaining a 4-G (i.e., 4.0 x aircraft BSDGW) load applied

over the area(s) described for either the inverted or sideward

attitudes shown in Figures 13 and 14 respectively, without permitting sufficient deformation to cause injury to seated, restrained occupants. For both cases (Figures 13 and 14), the

4-G distributed load should be analyzed for any angle of load

application ranging from perpendicular to the fuselage skin

(i.e., compressive loading) to parallel to the fuselage skin

60
04

N 0

04

or >4v

r4

10

V.o

4,

614- 0A I'~~~4 _ __ . .. P-4 _ 4
V) I ryj 

.uI t

N~ 

0i 

4

.141

45'4

Iý4 0

frO4-4

-HO 0

diho ~ ~ ..

a) wb) I 00

N i

CD

Z P62
(i.e., shear loading). When designing for this condition, it

should be assumed that all doors, hatches, transparencies,

canopies, and similar openings cannot carry any loading.

5.3.1.5 Wings and Empennage: Wing design should possess frangible characteristics to allow wings to break free from the

fuselage under high longitudinal inertia loads for distributed

impact loads caused by striking a barrier such as an earth

mound. Empennage structure should be designed to collapse or

break away during longitudinal crash impact.

5.3.1.6 Engine/Transmission Mounts: For light fixed-wing aircraft, engine mounts, both on the engine and on the supporting

structure, should be designed to keep the engine attached to

the basic structure supporting the mount under the crash conditions cited in Table 2, even though considerable distortion of

the engine mount and support structure occurs. The basic

structure supporting the engine should fail or separate before

engine mount failure occurs. Engine mounts and supporting

structures, including firewall bulkheads, should be designed to

minimize earth scooping. Engine casing design should be compatible with these requirements.

Transmissions and rotor masts of helicopters should be designed

to prevent potentially hazardous displacement or tilting under

the crash conditions cited in Table 2. The transmission, rotor

mast, rotor hub, and rotor blades should not displace in a mannor hazardous to the occupants during the following impact conditions:

* Rollover about the vehicle's roll or pitch axis on

sod.

* Advancing and retreating blade obstacle strikes that

occur within the outer 10 percent of blade span assuming the obstacle to be an 8-in.-diameter rigid

cylinder.

Unless otherwise specified, all engines, transmissions, rotor

masts, armament systems, external stores, and rotor hubs should

be designed to withstand the following ultimate load factors

(G) and remain restrained:

* Applied Separately

Longitudinal ±20

Vertical +20/-10

Lateral ±18

63
a Applied Simultaneously

Design Conditions

2 3

Longitudinal t20 t10 i10

Vertical +10/-5 .20/-10 +10/-5

Lateral :0 k9 *18

5.3.1.7 Shape of Fuselage Cross Section: The shape of the

fuselage has an inherent influence on the inward loaddeformation properties of the fuselage. Both crash test experience and accident analysis indicate that an ellipsoidal shape is optimum for the fuselage. A cylindrical cross section inherently provides a curved surface to resist inward

crushing. In addition, an ellipsoidal fuselage will result in

lower rollover loads than would a flat-sided fuselage under

identical conditions. Even though operational considerations

may prevent the use of an exact ellipsoid-shaped fuselage, an

approach to this shape is a worthwhile design goal.

5.3.1.8 Landing Gears The landing gear geometry should be

such that no abnormal characteristics result from aircraft

taxis, takeoffs, and landings at the basic structural design

gross weight on terrain with slopes of up to 12 degrees, or

from landing sideways on a 15-degree slope under zero wind.

The sink speed need not exceed 6 ft/sec for the above slope conditions. A differential kneeling landing system should not

be utilized to satisfy this requirement. There should be no

restrictions imposed due to positioning the aircraft relative

to the sloped site. The landing gear should be capable of

ground taxi, towing, ground handling, takeoff and landing roll,

and landings including autorotative landings at design sink

speeds in accordance with AMCP 706-201 (Reference 18).

The gear system should be designed to minimize entanglement

with wires, brush, landing mats, and other obstructions and

should have provisions for attachment of flotation and ski devices to permit operation on snow, water, and marshy areas.

The gear flotation capability should be such as to allow the

aircraft, empty except for full fuel load and an additional

200 lb, to be towed across soil with a California Bearing Ratio

of 2.5 by vehicles normally assigned to aviation units (i.e.,

1/4-ton or 3/4-ton truck).

18. ENGINEERING DESIGN HANDBOOK, HELICOPTER ENGINEERING, PART

ONE, PRELIMINARY DESIGN, AMC Pamphlet 706-201, U. S. Army Materiel Command, Alexandria, Virginia, August 1974.

64

"If `1-1
5.3.1.8.1 Tail Bumper: Tail bumper wheels or skids should be

provided as necessary. Skids should have a simple, hardened- i •+m'+• surface# replaceahle shoe to absorb the wear and damage of imu,'.'•,,pact,

5.3.1.8.2 Ground Clearance: The ground clearance, with air- craft level, for the antitorque (tail) rotor (exclusive of tail

bumper wheel or skid structure), fairings, control surfaces,

and external stores should not be less than 16 in. It should

be assumed that the aircraft is at rest at BSDGW and that the

landing gear struts are in the normal position with normal tire

"pressure. Alternatively, with the aircraft in any of the following attitudes the clearance should not be less than 6 in.:

9 Three-point and, where applicable, four-point attitude with all shock absorber struts fully compressed

and all tires flat.

* Three-point attitude with main wheel shock absorber

struts and tires under static deflection, nose-wheel

shock absorber strut fully compressed, and nose-wheel

tire flat.

9 Tail down, rolled attitude with main wheel shock absorber strut fully compressed, main wheel tire flat,

and nose gear at maximum extension. The longitu-

"dinal attitude of the rotary-wing aircraft should correspond to that obtained by contact of the aft fuselage structure or tail bumper with the ground or

deck. The lateral attitude should correspond to that

obtained by rotating the aircraft 5 degrees about its

roll axis.

5.3.1.8.3 Landing Gear Location: The landing gear subsystem

location should minimize the possibility that a part of the

gear or support structure will be driven into an occupiable section of the aircraft, or into a region containing a flammable fluid tank or line, in any accident falling within the

crash conditions of Table 2. If this cannot be accomplished

by location, the gear should be designed to break away under

longitudinal impact conditions, with points of failure located

so that damage to critical areas is minimized.

_ * Failure of the landing gear should not result in a failure of

any personnel seat/restraint system or seat/restraint system

tiedown. Failure of the landing gear should also not result

in blockage of a door or other escape route, or prevent the

opening of any door or other escape route.

65

- -- -5"•I . • :'4 '+. •.

• . ., : . • . t .Z ' . • .. . . . .. ' . * ' ' • . ' r ' ' .. .. . . . . . . . . ' , ' . -, .. . . "•" • • - • ! I N ,T''I+
5.3.1.8.4 General Strength Requirements: Unless otherwise

specified, strength and rigidity requirements should be provided in accordance with MIL-S-8698. The limit sink speed at

the BSDGW should be 10 ft/sec (level ground) and 6 ft/sec on a

12-degree slope in any direction. The forward velocity for

level ground contact should be all speeds between 0 and 120

perceAt of the airspeed corresponding to minimum power required

for level flight and landing gross weight. The reserve energy

sink speed should be 12.25 ft/sec. The following paragraphs

of MIL-A-008862 should apply for ground loads: 3.3 (except

3.3.7), 3.4, (except 3.4.3), 3.5, and 3.6. An analytical casting factor of 1.25 should be applied for the design of all

castings which will not be static tested to failure, or which

are not procured to MIL-C-21180. The yield factor of safety

should be 1.0.

5.3.1.8.5 Vertical Crash Force Attenuation in the Landing

Gear: Landing gear, including the skid-type, should provide

maximum practical energy-absorption capabilities to reduce the

vertical velocity of the fuselage as much as possible under

the crash conditions defined in Table 2. Forward and aftward

motion of the wheel in wheel-type landing gear of the trailingarm type is allowable in meeting this requirement.

The landing gear should be of the load-limiting type, and

should be capable of decelerating the aircraft at BSDGW from a

vertical impact velocity of 20 ft/sec onto a level, rigid surface without allowing contact of the fuselage proper with the

ground. Plastic deformation and damage of the gear and mounting system are acceptable in meeting this requirement; however,

the remainder of the aircraft structure should be flightworthy

after such an impact, with the exception of the main rotor

blades. The aircraft should be capable of meeting this requirement in accidents with simultaneous fuselage angular

alignment of ±10-degree roll and pitch.

5.3.2 Ancillary Equipment Retention

Ancillary equipment is a general term for all removable equipment carried inside the aircraft that could constitute a hazard to personnel if unrestrained during a crash. Ancillary

equipment includes emergency and survival equipment, aircraft

subcomponents, and miscellaneous equipment. Typical items in

each of these categories are:

66
* Emergency Equipment

Oxygen bottles

Fire extinguishers

First aid kits

Portable searchlights

Crash axes

e Survival Equipment

Survival kits

7• Life rafts

Life jackets Locator beacons

Special clothing

Food and water

* Subcomponents

Panel-type consoles containing control circuitry

Radio and electronic equipment

Auxiliary power units

Batteries

Special equipment

"s Miscellaneous Equipment

,. Navigation kits

S., Briefcases

Log books

Flashlights

Luggage

Toolboxes

All ancillary equipment frequently carried aboard an aircraft

should be provided with integrated restraint devices or anchors

to the aircraft structure. Restraint devices or anchors should

ensure retention of the equipment during any survivable crash

of the severity cited in Table 2. Stowage space for nonrestrained items that are not regularly carried aboard an aircraft should be provided in all aircraft. This space should

be located so that the items stored in it cannot become hazards to personnel in a survivable crash.

5.3.2.1 trgtht Restraint devices and supporting structure

for ancilla equipment should be designed to restrain applicable items when exposed to static loads of 50 G downward,

10 G upward, 35 G forward, 15 G aftward, and 25 G sideward.

Load-limiting devices are recommended for restraint of heavier

Sequipment. Load-limiter stroking should not allow equipment

to enter an occupant strike envelope.

67

'C•.I:•,,
5.3.2.2 Emergency and Survival Equipment Stowage Locations

Equipment should bet (1) located close to the primary crew

chief station, if applicable; (2) stowed in easy view of crew

and passengersl and (3) easily and reliably accessible in an

emergency. Equipment should not be placed in areas where cargo

shifting or fuselage distortion will prevent or impair access

to it. Equipment stowage location should minimize the potenzial adverse effects of extreme temperatures, abrasion, and

uncleanliness.

5A3.2,3 Retenti.on Devices Release for Emergency and Survival

Equipment, Retention devices used to restrain emergency and

survival equipment should be capable of quick release without

the use of tcols by one person using one hand. Release should

be effected by a single motion actuating one device and should

not require more than 5 sec from time of contact with the actuating device to the time when the equipment either falls free

or is 3ifted free. If equipment is stowed in an enclosure, no

more than 5 sec should be required for opening the enclQure

and removing the equipment. 'A.rcraft attitude should not adversely affect release device operation. It should be possible

%:o see the latch position (open or closed) of the release device. The release device actuating handle should be of a color

thAt contrasts with the surrounding area and be easily discernible in poor . '.ght or smoky conditions. No more than 30 sec

should be required for release of life rafts and their deployment outside the vehicle. Time should be measured from the moment when the operator takes a stand adjacent to the release

device or enclosure of the raft until the raft hits the water

uninflated.

5.3.3 Occupant Retention

Seating and litter systews should ensure that oucupants are retained in their precrash positions within the aircraft during

crashes of the severity cited in Table 2. Seating and litter

systems design should be coordinated and interfaced with the

design of the other aircraft areas to achieve a completely integrated and efficient crashworthy aircraft system design.

Seat and litter design should provide the greatest practical

amount of support and contact area for the occupants in the

directions of the most severe and likely impsicts. Seats should

provide an integral means of crash force attenuation. Occupant

comfort should aot be compromised to the extent that flight

safety and/or crew efficiency is adversely affected. Volume IV

contains a detailed discussion of occupant retention.

68

S N F _____T- --- _-- --.. - , •t , - - r i - "•- -• -
5.3.4 Cargo Retention

Cargo restraint should:

e Be as light in weight as possible.

e Require minimum storage space when not in use.

e Be easy to install and remove.

* Be easily and reliably adjustable for different

sizes and shapes of cargo.

I* Provide sufficient restraint of cargo in all directions to prevent injury to personnel in impacts of

the severity described in Table 2.

e Not permit cargo to shift in flight during turbulent

weather.

If the structure of the fuselage and floor is not strong enough

to withstand the cargo crash loads, load limiters should be

used to limit the loads transmitted to the structure. Cargo restraints should be capable of maintaining their integrity

under longitudinal loads of 16 G peak with a longitudinal velocity change of 43 ft/sec. Lateral and forward strengthdeformation charavteristics are discussed in Section 6.6.5.11.

Nets used to restrain small bulk cargo should be constructed of

material with low-elongation characteristics in order to reduce

dynamic overshoot to a minimum. Restraining lines without load

limiters used for large cargo, as defined in Table 4, for longitudinal restraint should be so arranged that maximum load

in all lines is reached simultaneously. Restraining lines

having different elongation characteristics should not be used

on the same piece of cargo. If load limiters are used, restraining lines should be metal cables with low-elongation

characteristics to ensure the most efficient energy absorption.

5.4 cESTING

5.4.1 Aircraft System Testing

Instrumented, full-scale crash test(s) should be conducteds

(1) to verify analyses performed and (2) to substantiate the

capability of the aircraft system to prevent occupant fatalities and minimize the frequency and severity of occupant injuries during crashes of the severity cited in Table 2.

69
TABLE 4. AIRCRAFT CARGO CATEGORIES

Small bulk cargo Large rigid cargo

(net restraint) (line restraint)

This class includes all boxes This class iniludes all rigid

or unpacked cargo of approxi- cargo of 3 ft or more in size.

mately 3 ft or less in size.

Examples: Examples:

1. Ammunition boxes 1. Wheeled or tracked vehicles

2. Foodstuffs 2. Aircraft engines

3. Medical supplies 3. Fuel barrels

4. Clerical supplies 4. Artillery pieces

5. Vehicle maintonance 5. Special weapons

compononts (priority cargo)

5.4.2 Landing Gear Crash Testing

Instrumented drop tests should be conducted: (1) to verify

landing gear crash force attenuation and crash loading strength characteristics analytically predicted and (2) to substantiate

the capability of the aircraft landing gear to meet the criteria of Section 5.3.1.8. Drop testing of wheel and skid landing gear should be conducted in accordance with paragraph 9-2.3

of AMCP 706-203 (Reference 19) and should include demonstration

of compliance with the reserve energy and crash impact requirements of Section 5.3.1.8. The 20-ft/sec sink speed drop test

should be conducted with che landing gear oriented in a 10-

degree nose down and 10-degree roll attitude and drop tested

onto a level, rigid surface with a sink speed of 20 ft/sec at

ground contact. Landing gear should also be drop tested in a

0-degree roll, pitch, and yaw attitude onto a level, rigid surface with a sink speed of 42 ft/sec at ground contact to demonstrate crash impact energy-absorption capability. Rotor lift

for all drop tests should not exceed two-thirds of the BSDGW.

19. ENGINEERING DESIGN HANDBOOK, HELICOPTER ENGINEERING, PART

THREE, QUALIFICATION ASSURANCE, AMC Pamphlet 706-203,

U. S. Army Materiel Command, Alexandria, Virginia, April

1972.

70

I'
5.4.3 Cargo Restraint

Design loads are specified in Section 6.6.5.11. Static tests

to these loads are recommended. All deformation measurements

are to be made at the floor level. Sufficient dynamic tests

should be made to assure that design predictions can be based

on static test results.

5.4.4 Seat and Restraint System

Testing requirements for seats and occupant retention systems

are described in Volume IV.

5.4.5 Fuel System

Testing requirements for fuel systems are described in Volume V.

5.4.6 Ancillary Equipment Retention

Design loads are specified in Section 6.6.5.9. Static tests

to these loads are recommended.

5.4.7 Static Structural Crash Load Testing

It is important that compatibility at the structural interfaces

between the airframe and all attached components is ensured.

The design of the fuselage structure, including the hard points

and load distribution structure around the hard points, must

be coordinated with the design of the attaching components.

Structural properties for all loading conditions and design features, such as structural releases, must be coordinated to

achieve the desired compatibility.

The importance of this aspect of system design requires that

compatibility be demonstrated. Static tests of components attached to the fuselage structure by their normal attachment

provisions should be performed to demonstrate compatibility.

Components such as seats, cargo, engines, transmissions, landing gear, and attachments for any ancillary or heavy equipment

located in an area which could create a hazard for the occupants if freed during crash foaling should be tested. The ultimate design crash loads should be applied in all principal

loading directions to demonstrate that the attachment points

as well as the load-bearing sections of the fuselage are capable of maintaining structural integrity during a crash. The

application of proof loads instead of ultimate crash design loads is an acceptable, although minimum, test condition.

71
6. AIRFRAME PRINCIPLES AND CONCEPTS

6.1 INTRODUCTION 4

In designing for airframe crashworthiness, there are certain

criteria that are applicable whether the results are approximate or precise. First, and most important, the structure surrounding occupiable areas must remain reasonably intact, without significantly reducing occupant living space. If occupants

are injured during a crash because the protective shell simply

collapses around them, then efforts to improve survivability

through such methods as improvement of occupant restraint or

reduction of postcrash hazards are futile. An aircraft which

does not provide the protective shell can never be crashworthy.

Ideally, for crashworthiness, the structure should be designpd

to minimize occupant accelerations to survivable levels in a

severe crash environment, while maintaining the required survivable volume and retaining large mass items, interior equipment, seats, and cargo. In addition, consideration should be

given to minimize the effects of rollover, earth plowing, and

cabin penetration. Obviously, other considerations must also

be addressed in defining the airframe structural configuration.

6.2 GENERAL DESIGN CONSIDERATIONS

Airframe structure should first be designed for normal airloads,

ground handling loads, and fatigue life, while considering the

details of the aircraft specification with respect to size,

range, performance, space envelopes, etc. After the basic

structural layout has been defined, the effects of crash loads

must be considered to determine where structural modifications

are needed to improve crashworthiness. Concurrent with this

process, space allocations must be made and locations for critical systems, landing gear, equipment racks, seats, cargo tiedowns, emergency exits, etc., must be determined to ensure an

integrated approach to the solution of the crashworthiness

2roblem.

To improve crashworthiness, design changes allowing the structure to remain more nearly intact through improved compliance

or improved progressive, yet predictable, deformation can sometimes be more effective than direct increases of strength.

Every attempt should be made to accurately determine the magnitudes and distribution of crash loads throughout the structure,

and the structure should be designed to these loads.

The use of safety factors of 1.0 with safety margins of 0.0

based upon the crash load factors presented elsewhere in this

72
volume will generally be considered satisfactory. Safety facAL) • tors may be increased by known amounts as special conditions

4WI) imay warranty however, caution is needed so that overdesigning

"of certain areas does not adversely affect the protection offered by the entire structural system. Safety factors and

safety margins used in the design of joints and connections

should be consistent with those discussed above.

In areas where large gross structural deformations are anticipated, joints should be analyzed and designed to reduce the

probability of failure under large angular deflections as well

as large linear displacements. The objective of these considerations is to keep structural sections intact and to ensure

that when massive crushing or large deformation occurs, maximum energy absorption also occurs. Consideration also must be

given to areas of structure where a predetermined failure sequence is deemed necessary to provide optimum occupant protection. Joint failure can be controlled so deformation occurs

away from occupied space, thus eliminating the problem of occupied space penetration by the ends of failed structural members.

Furthermore, if joints are designed to retain an effective pin

connection after failure of the majority of fasteners, a certain level of structural continuity can be retained. This can

result in a structure that retains the capacity to resist loading after the primary impact, thereby affording protection during rollover, pitchover, and intrusion of massive items.

Often, plastic yielding of materials will relieve stress

concentrations before the ultimate strength of members are

reached, making the consideration of stress concentrations,

from a strength standpoint, less important. However, very

often the presence of stress concentrations induces local fractures and eventual structure failure with low overall energyabsorbing capacity.

During a crash, rapidly increasing loads, not instantaneous

loads, are applied. In most cases, a minimum of 10 msec is

required for loads to reach maximum values. Under such conditions, inertia effects may be of importance, but strain rate

effects in materials are probably insignificant.

6.2.1 Initial Layout

Structurally, the aircraft must be capable of performing its

mission of carrying the required payload. Initial layouts must

consider the volume required to carry the requisite crew, passengers, and cargo after allocation of space for aircraft systems. The structure needed to carry and/or house the systems

and occupants must be laid out to adequately support all systems and to provide basic structural protection in a crash environment.

73

-'---I,. I I
Figure 15 shows a structure designed to protect the occupants

in crash conditions. Adequate space for allowing the structure

to stroke and/or collapse to absorb energy in the support of " large mass items must be provided. At the same time, seated

occupants and cargo must be restrained and G levels restricted

to provide a survivable environment. Crashworthy seats must be

supported by an integral part of the primary structure. To

preclude the seat bottoming on the floor of the aircraft, sufficient stroking distance must be allowed for impact conditions

defined in Table 2.

6.2.2 Analysis and Simulation

Once the design is defined, the next step is to simulate the

* preliminary concept. Computer simulations, such as described

in Chapter 7, are used to model primary structure, large mass

items, systems, occupants, and cargo. Then, potential impact

conditions are simulated to investigate the aircraft's dynamic

response, structural collapse, and acceleration environment.

An iterative process is used to optimize energy-absorption concepts, structural distributions, failure modes, mass retention

concepts, landing gear locations, etc. This process, conducted

concurrently with inputs from design, stress, and performance,

ensures design optimization. Further modifications are made

to control weight and weight distribution, producibility, maintainability, safety, and cost.

6.2.3 Design Review and Final Aircraft Format Selection

Several cycles of iteration may be required before the optimum

design, adequately satisfying the requirements of the basic

helicopter specification, evolves. The result is a final product that satisfies the major specified requirements with a minimum of compromise.

From the beginning, emphasis must be placed on the need to implement crashworthiness features. This provides an integrated

approach where all influencing constraints, particularly weight,

cost, and space are considered and optimized. Crashworthiness

cannot be added to a design without certain compromises, possibly even penalties, involving weight, cost, and structural

complexity. However, crashworthiness, when integrated with the

original design, can be achieved with a minimum and acceptable

level of these penalties.

6.3 MATERIALS AND STRUCTURAL PROPERTIES

Current helicopters are constructed primarily of metallic materials. Some secondary and fairing components are made from

74

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4 J40

41'

00

~ .4J

a0

0 b,

w.0)

-14'

41'

.ii ,44, 4

0 441

75
nonmetallic materials or composites, i.e., combinations of metallics and nonmetallics.

For crashworthy structures, the basic requirements are to maximite energy absorption while maintaining adequate volumes and

acceleration levels for occupant survival, and to minimize fire

risk.

Other requirements for a safe interior environment include the

followings

e Controlled collapse mechanisms.

e Material failure modes that do not produce projectiles.

* Joint designs and fastener selections that control

failure mechanisms and minimize the formation of

projectiles.

MIL-HDBK-17 (Reference 20) and MIL-HDBK-5 (Reference 21) contain basic design data for materials.

Subsequent sections discuss these requirements with reference

to both metallic and composite structures.

6.3.1 General

The kinetic energy of an aircraft subjected to a crash is absorbed in a number of ways, including deformation of aircraft

structure. Absorption of energy through structural deformation

can work effectively as an occupant protective device if the

structure surrounding the occupant protective shell is allowed

to deform, thus attenuating the forces transmitted to the occupiable section. The protective shell must remain intact during this process. The use of surrounding structure as a buffer can

be accomplished more efficiently if the impact causes crushing without complete rupture of structural members. Material prop- erties can greatly affect the degree to which this is actually achieved.

20. Military Handbook, MIL-HDBK-17, PLASTICS FOR AEROSPACE

VEHICLES, PART I, REINFORCED PLASTICS, Department of Defense, Washington, D. C., 1 January 1971.

21. Military Handbook, MIL-HDBK-5, METALLIC MATERIALS AND

ELEMENTS FOR AEROSPACE VEHICLE STRUCTURES, Department of

Defense, Washington, D. C., 15 September 1976.

76
Material ductility is required to ensure that crushing, twisting, and buckling can occur without rupture. The structural

shell should be able to deform without fracture insofar as possible.

All exterior surfaces and all structures which could be exposed

to contact with the impact surface should be constructed of

nonsparking materials to reduce the postcrash fire hazard. The

friction sparking problem and results of friction sparking ignition experiments using various typical aircraft materials

are discussed further in Section 6.3.5 and in Volume V.

Fabrication techniques and structural configurations, especially for areas where severe damage is probable, should be selected after consideration of the overall effects of structural

failure on occupant protection. Wherever possible, multiple

structural members should be used instead of larger single

structural members, so that localized impact damage will not

result in complete loss of structural integrity. Multiple load

paths also aid in maintaining uniform force transmission characteristics throughout structural collapse.

6.3.2 Material Strength and Elongation Characteristics

Material strength and elongation characteristics are described

in applicable military handbooks. In using these properties,

one should bear in mind that for some materials, such as steel,

J handbooks present only guaranteed minimum strength and elongation datai for other materials, such as aluminum, values are

"presented indicating both minimum guaranteed strength and the

strength values which statistics show will be met or exceeded

by 90 percent of the materials under consideration. The 90-

percent probability values are normally somewhat higher than

the guaranteed values. The particular application should be

considered in deciding whether to use minimum guaranteed values

or statistically based values. For design of structural members, such as members supporting heavy overhead masses, where

failure could result in a severe loss of occupant protection,

the use of minimum guaranteed values would be reasonable. For

design of structures likely to be subjected to massive crushing, such as airplane nose structures, the use of statistical

values would be justified. As in these two simplified examples, the. choice of which strength value to use should be based

* .upon the consequences of failure.

6.3.3 Failure Modes Controlling Material Selection

In a crash environment, structural failure modes should not

create potentially injurious conditions for cabin or crew section occupants. When occupant protection is considered, the

following failure modes should be avoided:

77
e Inward buckling structures, such as sidewalls, bulkheads, and floors.

0 Failures of members such as frames that result in

jagged, failed ends protruding into occupied space

or fuel cells.

e Fastener failures that may produce structural discontinuities and projectiles.

* Brittle fractures that suddenly unload, causing impulse effects in adjacent structures with potential

for progressive failures and generation of projectiles.

* Emergency exit surrounds that distort excessively and

preclude the opening or removal of doors or windows

after the crash sequence.

* Flammable fluid container penetrations with the potential for postcrash fire generation.

e Occupied area penetrations by failed structural elements or exterior agents.

Brittle structural failures can be avoided by choosing materials with good fracture toughness characteristics and a considerable degree of ductility before ultimate failure. The

choice of structural geometry can alleviate potential ingress

of failed structure into occupied space.

Also, predetermined failure points can be introduced into a design to help control structural response under dynamic loading

conditions. These points may be structural plastic hinges that

allow earlier plastic yielding and rotations in weaker sections.

Alternatively, joints may be designed to fail progressively to

allow rotation of structural elements with subsequent load redistribution. A final joint condition of effectively pin-ended

members will quite often be more desirable than member breakaway.

Typical examples of failure modes discussed in this section are

shown in Figures 16 through 22, demonstrating, in particular,

the requirements for failure away from occupied areas and prevention of potentially damaging projectiles and structural

elements. These photographs were taken after a vertical impact of a test vehicle onto a concrete ground plane at a velocity equivalent to at least a 95th-percentile potentially

survivable accident. The structural failures, in general, did

not result in unacceptable intrusions into occupied space.

78

!,'.
Figure 16. Compressive buckling at base of buttline

beam structure on a medium cargo helicopter following vertical impact.

The resulting sharp-edged elements could have been eliminated

by using materials with better fracture toughness properties

and/or increased dimensions.

It is interesting to note that, historically, aircraft fuselage

sections have been shaped to preclude inward collase mechanisms. Circular, elliptical, and contoured rectangular shapes have been used to accommodate the requirements for structural

efficiency and optimization of useful volume. The major factors degrading older designs were the lack of requirements to

withstand crash-induced loading and the lack of choices in materials. At the time, materials with the high strength, ductility, and fracture toughness properties now accepted as the

norm did not exist.

79
Ie

Figure 17. Typical buckling collapse of vertical

bulkhead and buttline beam on a medium

cargo helicopter following vertical

impact.

80

?I
SI

Figure 18. Typical frame failure away from occupied

space in the cockpit area of a medium

cargo helicopter following vertical impact.

6.3.4 Composite Materials

Composite materials offer certain advantages over metallics in

some areas of helicopter structure, primarily the advantage of

reduced weight. However, their high stiffness and lack of ductility require that composites be investigated very carefully

before being implemented in structural areas where high levels

of energy absorption are required and large stroking distances

are considered necessary. Figure 23 compares typical stressstrain curves for an aluminum alloy and a graphite/epoxy composite in tension. The shaded areas indicate the potential

energy absorption capabilities; the difference between the two

materials should be noted, i.e., the ratio A70 7 5 /AGE 12.3.

81[
Figure 19. Typical frame failure away from occupied

space in the cabin area of a medium cargo

helicopter following verticýal impact.

Table 5 presents values for ultimate strengths and elastic

moduli for several typical structural composites. These data

allow comparisons among various nonmetallic materials.

Another available method of energy absorption that may achieve

adequate performance is incorporation of filler materials, such

as honeycombs and structural foams. For longitudinal and lateral impacts, it is feasible to incorporate energy-absorpt-.ion

features of this type; but for the support of large mass items

with high energy content, other techniques are probably necessary. Of course, any limitations on the use of energy- absorption techniques will be dependent, to a degree, on the

size and overall layout of the aircraft necded to satisfy mis'-

sion requirements. For structural energy absorption adequate

82
F, F

Figure 20. Failure modes resulting in jagged elements

protruding into occupied space in the cabin

area of a mediumn cargo helicopter following

vertical impact.

83

IA
Figure 21. Rotational joint failure with compression

and bending in a medium cargo helicopter

following vertical impact.

Figure 22. Failure of frame member and joint with fragmentation, compression, and bending in a medium cargo helicopter following vertical impact.

84
200

Energy-absorption ratio A- = 12.3 •,,. •AGIE

M /

~~ 10 7075

LO dG/E

50

.H

0

0 0.02 0.04 0.06 0.08 0.10 0.1.2

Strain (in./in.)

Figure 23. Stress-strain relationship for aluminum alloy

(7075) and 0 degrees graphite/epoxy composite.

for occupant survival in the crash conditions of Table 2, optimization studies may indicate that some major primary structural elements need to be manufactured from metallic materials,

whereas composites may find use in other areas.

The effect of thermal mismatch is a major potential problem

area that must be fully investigated when using mixed construction techniques. The use of boron/epoxy or graphite/epoxy materials in close conjunction with other materials must be carefully examined in the conceptual design stage because of the

generation of internal stresses induced by differences in thermal expansion coefficients.

Combinations of composites with steel or titanium alloys can

be satisfactorily designed for most structural applications,

but combinations with aluminum alloys can lead to severe problems. If thermal curing of the bonding agent is also needed,

this can further aggravate the induced stress problem and may

result in warpage of the finished product.

85

h Al•
TABLE 5. ULTIMATE STRENGTH AND ELASTIC MODULI

FOR TYPICAL STRUCTURAL COMPOSITES

Ultimate Modulus of elasticity i tensile

stressi2Longitudinal stress.2 Transverse

Composite (b/in. (lb/in. (lb/in. 2

Boron/epoxy [0] 192,000 30 x 106 2.7 x 6 -30 x 106 -2.7 x 10

Graphite/epoxy (01 110,000 25 x 106 1.7 x 6

-5x106 1.7 0 (high modulus -25 x 1 -1.7 x 106

graphite)

Graphite/epoxy [0] 180,000 21 x 106 1.7 x 10

(high strength -21 x 106 -1.7 x 106

graphite)

Boron/aluminum 160,000 34 x 106 20 x i10

-30 x 106 -19 x i06

Borsic/aluminum 140,000 32 x 10 6 22 x I06

-30 x 106 -19 x i0

Thermal mismatch problems are less severe when the two materials are in the form of sandwich face sheets and core since

the flexibility of the core can absorb limited thermal displacements, particularly those normal to the core ribbon direction.

Boron/aluminum composites, on the other hand, are compatible with aluminum alloys, but their use with steel and titanium

alloys must be carefully considered.

Table 6 gives values of thermal expansion coefficients for com-

* posits and metallic materials. The table and other useful design information can be found in Reference 22.

A: If composites are considered for primary crashworthy structure,

the designer must be aware of the pitfalls (such as possible weight and cost penalties) that can render them less effective

22. Air Force Systems Command, ADVANCED COMPOSITE DESIGN GUIDE, Advanced Development Division, Air Force Materials Laboratory, Wright-Patterson Air Force Base, Ohio, January 1973.

86

I' .. •
TABLE 6. THERMAL COEFFICIENTS OF EXPANSION FOR

COMPOSITE AND METALLIC MATERIALS

(REFERENCE 22)

Coefficient of thermal

expansion

(106 in./in./*F)

Material Longitudinal Transverse

Boron filament 2.7 NA

Epoxy matrix resin 2.7 27

Graphite fiber -0.05 NA

E-glass filament 2.8 NA

Boron/epoxy [0) 2.3 10.7

Boron/epoxy [02/*45] 2.4 7.7

Graphite/epoxy (0] 0.3 14.4

Graphite/epoxy [0/±45/90) 1.9 1.9

E-glass/epoxy [0] 4.8 -

E-glass (181 style weave)/epoxy 5.5 6.7

PRD-49/epoxy [0] -6.0

PRD-49 (181 style weave)/epoxy 0.0

Aluminum 13

Steel 6

Titanium 5.6

than initially anticipated. Alternatively, if a composite

structure with an energy-absorption capability less than that

* of a metallic structure is used, a survivable deceleration environment must be created by the use of landing gear with

greater energy-absorbing capacity, energy-absorbing seats,

and possibly some forms of floor load attenuation other than

through primary structural deformation. Once again, such upgraded devices can introduce additional weight and may require

extra installation space to allow adequate stroking distances.

87
Reference 23 presents a survey of the crash impact characteristics of composite structures and suggests the possible design

concepts discussed below.

6.3.4.1 Overall Fuselage Concepts: Figure 24 illustrates an

approach to satisfy the crashworthy fuselage consideration

where the primary function is to maintain a protective shell

around the occupied area. This can be accomplished by provid- ing rollover strength in the form of elliptical sandwichstiffened ring frames on the top and sides of the fuselage and

a structural floor. Note that in Figure 24 the side fillets

are designed to provide better rigidity and corner continuity between the floor and the sidewalls, while they will also act to absorb energy and distribute loads from a side impact. The

foam-filled Kevlar tubes would be multifunctional in that they would act as floor beams to satisfy the normal airworthiness

criteria of strength and stiffnessl they would also function

as energy absorbers to react the vertical and lateral crash

impact forces. The heavy sandwich construction with local

strap reinforcement on the lower mold-line surface would act

as a shear and axial load path to satisfy the normal airframe

requirements, while functioning as a skid during a longitudi- nal impact to prevent plowing and gouging. A common post- crash hazard is entrapment of the occupants due to collapse of

the emergency exits. The hatches could be reinforced with a

filament-wound composite tube that would provide the strength

and stiffness required to prevent the distortion of the hatch

region. This local reinforcement could also serve as a framing member that would redistribute the airloads around the cutouts.

Local details, such as seat attachment points in the floor or

cabin roof, need to be in areas where the airframe distortion

during crash impact would not adversely affect the operation of

the seat stroking mechanism or the seat belts. Consequently, reinforcements in the form of a higher density core in the

sandwich construction or a filament-wound lug are shown in a

typical application.

6.3.4.2 Composite Material Concepts for Vertical Impact: In

order to reduce the inertia forces experienced by the occupant in a crash, energy-absorbing material must be placed between

the ground and the floor of the cabin. Normally, this location

is occupied by controls, wires, and life-support systems; consequently, the overall design must take this into account.

23. Cronkhite, J. D., et al., INVESTIGATION OF THE CRASHIMPACT CHARACTERISTICS OF ADVANCED AIRFRAME STRUCTURES,

Bell Helicopter Textroni USARTL Technical Report 79-11, Applied Technology Laboratory, U.S. Army Research and Technology Laboratories (AVRADCOM), Fort Eustis, Virginia, April 1979, AD A075163.

88

J A_ . -,' "
~' Reinforced exits

•" ~Strong rollover ; structure

Energyabsorbing

construction

under fuselage

Local strap Strong floor for

reinforcement seat attachment

Frangible fittings provide Seats and other

for jettisonable equipment nonbreakaway equipwhich will not impinge into ment provided with

crew compartment filament-wound

secondary equipment

Sandwich-stiffened

Sring frames prevent

•i Side filletsSide flletsinward buckling

prevent inward

buckling of

frames

Squared - -/

corners provide lateral energy 300 Vertical impact

absorption for impact

oblique impact

Figure 24. Overall fuselage concepts. (From Reference 23)

i8
The beam and bulkhead concepts shown in Figure 25 are designed

to react the vertical, longitudinal, and lateral impact loads/

however, the most efficient direction for providing a progressive collapse is vertically. In each of the four concepts

shown, it is assumed that the floor structure would also be

designed to react the crash impact forces without failure. The

concepts shown are directly applicable to many current fixedand rotary-wing aircraft because their structural arrangement

is similar. The first and second concepts would promote progressive crushing by stratifying the core and skins using the

less-dense core and the less-stiff skin at the lower ends.

The variation of skin stiffness would be accomplished by deleting plies. The stratification would have two effects: It

would promote the crushing et the lower surface, and it would

provide the increased strength of the core and skin adjacent

to the floor where it is needed for local reinforcements or

for seat or equipment attachments. Design support tests would

be used to select the materi&!, fiber orientations, etc.; however, a promising candidate for fulfilling both the airworthiness and crasshworthiness criteria is graphite because of its

high specific-energy absorption and its high strength and

stiffness.

The third concept features a frangible corrugated core that is

bonded to the face sheets at the nodes. This concept requires

the normal airworthiness loads to be shared by the face sheet

and corrugated core. During a crash impact, the face sheets

would act as short columns with a length equivalent to the corrugation node spacing. When the face sheets fail as a column,

the corrugated core would continue to crush and dissipate the

energy. Kevlar was chosen as a material candidate for the

face sheets based on a series of component tests. These tests

demonstrated that, although Kevlar has a lower compression allowable than other conposite materials, it also has a favorable

failure mode for progressive folding that is required for this

concept. The fourth concept uses the corrugated Kevlar face

sheets to carry the shear loads during normal operations.

These sheets would be continuously supported by the core in a

manner that would cause the crash impact loads to be shared by

the sheets and co..e. The design would have a peak crushing

load which would be determined by the column stability of the

face sheets, and an effective crush distance as governed by the

energy-absorption qualities of the core.

Figure 26 shows three concepts using tubular shapes which adSdress the combintu "ateral/vertical or oblique impacts. From

a manufacturing viewpGint, the tubular concepts would offer

the opportunity for 1ilament winding and may also offer a cost

advantage. Again, all of the concepts are predicated on the

design of a structura' .loor system of sufficient strength and

90
Beams and bulkheads must .iprovide progressive collapse

and energy absorption and

react vertical, longitudinal,

and lateral impact loads

(structural floor removed)

Stiffen

skins by

i /padding

plies

Most dense Dense -Kevlar Foam or

core or core face balsa

foam y sheet core

Least dense Less stiff Frangible foam or core skins corrugated

core

Figure 25. Energy-absorption concepts - beams and bulkheads

(vertical impact). (From Reference 23)

stiffness to react the crash impact forces. The heavy hatched

area in the floor represents local reinforcement for seat or

cargo attachments that coincide with the tube reaction points.

In these concepts, the energy-absorbing agent is the foam or

honeycomb structure. In the first concept, Figure 26(a), the

vertical forces are dissipated by the foam and reacted directly by the floor structure. For an oblique impact, the Kevlar

straps would resist the tendency of the outer tubes to roll

by reacting the side forces with a truss action. The straps would also replace the normal subfloor bulkheads to react the

normal flight loads. The truss is used instead of the bulkheads because it permits the tubes to expand freely in a lateral direction instead of being constrained by the bulkheads.

91
Kevlar straps maintain

structural integrity

and react side loads

Foam-filled Kevlar

tubes providc

vertical and

lateral energy 300 impact - absorption

SVertical impact

(a)

No foam in center

Corrugated Kevlar section for controls

semi-tube provides routing

vertical and lateral

energy absorption Outer tubes may be

foam filled for an

i Iadditional absorption imp capability

Vertical impact

(b)

Filament-wound sandwich

double-tube substructure

around crushable core provides

vertical and lateral energy absorption

Honeycomb or

foam provides

additional

I vertical and

lateral energy ~absorption

'"'''":"f•300 impact

Vertical impact

(c)

Figure 26. Energy-absorption concepts - tubular

construction (oblique vertical impact). (Prom Reference 23)

92

4 1 4A
The second concept, shown in Figure 26(b), uses the center tube

to route the controls, wires, and hydraulic lines. The tubes

would be stiffened circumferentially by corrugations to increase their load capacity for both the normal airworthiness

and crash impact design conditions. It is felt that the stiffened tubes would dissipate a portion of the crash impact energy working in conjunction with the foam, but tests would be re- quired to determine the relative efficiencies of both filled

and unfilled tubes.

The third concept, shown in Figure 26(c), uses the filamentwound sandwich tubes to form a housing for routing controls,

etc., down the center of the aircraft. The vertical legs of

the tubes are supported by a core of honeycomb or foam so they

would progressively crush in a vertical impact and not fail

catastrophically. The side impact forces would be reacted by the honeycomb or foam, which would distribute it to the floor.

6.3.4.3 Fuselage Sidewall Concepts: The fuselage sidewall

concepts shown in Figure 27 are essentially designs which are

intended to preserve the protective shell. The concepts illustrated in Figure 27(a) through (c) make use of the geometrical advantage of a circular cross section to reduce load concentrations while providing a strong sidewall to prevent failure.

The sidewalls may contain multiple frames or use sandwich construction to react the crash impact forces. However, the sandwich construction would also provide some energy dissipation

in addition to its structural integrity. The fourth concept, Figure 27(d), enhances this premise by applying energy- absorbing material to the sidewalls in a manner similar to the

subfloor concepts.

6.3.4.4 Longitudinal Impact Antiplowing Concepts: In a longitudinal Iimpact, the main considerations are to ensure the

integrity of the structural shell, to minimize earth scooping

or plowing of the lower fuselage, and to provide some energyabsorbing material forward of the occupied area to dissipate

the crash impact forces.

Figure 28 shows the strong floor structure required for the lateral and vertical impact conditions fairing into a sled-like

nose that will act as a landing skid and as a backup structure

to react the longitudinal impact forces. The current functions

for the nose structure are to house electronic equipment, batteries, radar units, and ballast, and to act as a storage location for the nose gear. The design concept must be modified

to satisfy these requirements while also integrating crushable

energy-absorbing material to react the impact force. The belly

skin would be fabricated in one section, similar to a large

93
44 4) si

00

00

- 0

'- 4.l

"4

r-44

(n 0C4

0

~t4J

0

r4

4-) M94
'.1 :Sled-like

strong floor

structure

Longitudinal energyabsorbing materialCrushes in a continuous manner

to prevent plowing/scooping

Figure 28. Antiplowing concepts - longitudinal

impact. (From Reference 23)

bathtub, to minimize joints that could fail and present a protrusion gouging into the earth's surface. This method of construction eliminates parts and has been suggested as a means

for reducing cost of composite components. A sandwich construction using Kevlar face sheets is a material choice that

offers strength in shear and a resistance to tearing or puncture. All panels and hatches would be designed to have a leading edge that laps under the skin so that it would not act as

a scoop during the large deformations that occur in a crash.

The one-piece skins in conjunction with the energy absorbers,

which could be foam-filled Kevlar tubes, would be encouraged

to collapse in a continuous manner by including longitudinal

straps under the absorbers to add bending strength and to act

as a skid.

6.3.4.5 Joint Concepts: The joints and attachment fittings of a crashworthy airframe structure must be able to perform the

following functions:

* Withstand large deflections without failure in those

areas where large deformations are anticipated.

e Connect the large overhead mass items to the fuselage such that failure or separation of the major structural members supporting the items occurs before the

joint fails.

95

,-L t i.
Of the two considerations, the first imposes the greater constraint on a joint constructed of composite materials or, for

that matter, on a metallic fitting manufactured from a casting

or a forging since such configurations generally have a lower

strain to failure than a sheet-metal-type joint. Composite

joint concepts should have redundant or back-up load paths to

satisfy the above considerations -while relying on the airframe

T,! structure to absorb the energy and redistribute crash loads.

Most of them contain some metal and therefore are considered

Y'l hybrids.

In the concept shown in Figure 29(a), the primary load path is

the bolted connection between the fitting and the bulkhead.

The filaments around the bolts would be designed such that they

would not begin to accept any load until the primary load path

in the joint has yielded.

The joint concept shown in Figure 29(b) uses a primary load

path through the attachment hole and the bondline between the

fitting and the skin. A secondary load path exists between the

filaments that are wound in a "racetrack" fashion around the

fitting and the skin.

The concept shown in Figure 29(c) depends upon the bond joint

between the fitting and the skin for its primary load transfer.

The secondary load path uses the mechanical connection provided

by the rivets to transfer the load from the fitting to the metallic back-up plate to the composite face sheets. The rivets

also provide a tension capability to resist peeling forces

which are present due to the eccentric load paths.

The flattened filament-wound cone concept shown in Figure 29(d)

provides a progressive failure mode which causes the joint

to contract circumferentially as it elongates longitudinally.

After failure of the matrix, the joint would still retain some

residual strength due to the fibers contracting around the metal insert.

6.3.5 Spark Generation

Two types of* sparks should be considered potential ignition

sources. The friction spark is a particle abraded from a parent material through contact with a moving surface. Initially, the particle is heated by friction. If the friction is great

enough, the material's combustion temperature will be reached,

causing the particle to ignite. Electrostatic sparks result

from the discharge of an electrostatic charge accumulated on

parts during normal operation. The discharge may be triggered when crash forces cause separation of the parts. This section

is concerned only with material selection to minimize spark

generation.

96
Filaments bonded around fitting

provide, secondary progressive

failure mode

Metal fitting bolted to

bulkhead

ill i[•Hybrid joint Filaments bonded around

(a) fitting provide secondary

progressive failure mode

Attachment hole

Metal or composite fitting bonded in place

Honeycomb panel

Hybrid or all composite joint

(b)

Rivets and metal backup plate

provide secondary failure mode

Aluminum fitting \ I bonded in place Composite face Flattened filamentsheets wound cone provides

progressive failure

Composite Metal. backup plate md

filler

Hybrid joint

(c)

insert

Hybrid joint

(d)

Figure 29. Crashworthy joint concepts. (From Reference 23)

"97

:C•.
Friction sparks become possible ignition sources when portions

of aircraft structure scrape along the ground. While all common metals can be abraded, not all sparks are sufficient to ignite spilled fluids; rather, ignition is dependent on the thermal energy of the spark.

Thermal energy is a function of the following parameters:

e Bearing pressure of structure on ground.

e Sliding velocity of structure relative to ground.

* Metal hardness.

e Temperature at which metal particles burn.

Table 7 gives some results of research conducted by NASA to determine the minimum conditions under which friction sparks from

metallic materials typically used in aircraft construction will

ignite.

TABLE 7. MINIMUM CONDITIONS UNDER WHICH CERTAIN ABRADED

METAL PARTICLES WILL IGNITE (REFERENCES 24 & 25)

Minimum

bearing pressure Drag speed

Metal .(lb/in. 2 ) (mi/h)

Titanium 21-23 less than 5

Chrome-molybdenum steel 30 10

Magnesium 37 10-20

Stainless steel 50 20

Aluminum 1455* 40

*Ignition was not obtained with aluminum.

24. Pinkel, I. I., et al., MECHANISM OF START AND DEVELOPMENT

OF AIRCRAFT CRASH FIRES, Lewis Flight Propulsion Laboratory; NACA Technical Note 2996, National Advisory Committee for Aeronautics, Cleveland, Ohio, 1953.

25. Campbell, J. A., APPRAISAL OF THE HAZARDS OF FRICTIONSPARK IGNITION OF AIRCRAFT CRASH FIRES, Lewis Flight Propulsion Laboratoryl NACA Technical Note 4024, National

Advisory Committee for Aeronautics, Cleveland, Ohio, May

1957.

98

, . . ) | ' .• .. .. •.. .......... ' :' " ' '• . •" ' "•" • . ... i'.••'''', .... .. :• ..•" • ' "•,•"'••,• ''",'," '";" •• , ,• '• • ,•:•,+.••. .,-•, ,,,,•, ,.• ,,, ,% :,,,•.•i• • .,;*; :, -
Two approaches can be used to alleviate the hazards posed by

V friction spark ignition. One approach is to select materials ir that possess high ignition thresholds or will not ignite at

all for those structures apt to strike the ground plane. The

other approach is to prevent a loss of flammable fluids in a

crash environment. Perhaps the most practical approach to the

problem is to minimize the probability of ignition and fuel

loss for a given design configuration.

Aluminum alloys are the least likely to cause ignition of

spilled fuel; however, the abrasion rate for aluminum is high,

which can result in rapid wear and subsequent tearing. This,

in turn, exposes other structures that may be manufactured from

materials prone to generate sparks. To minimize the probabil-

"ity of belly skin loss, relatively thick skins made of ductile

materials are recommended.

Particular attention should be given to attachment points for

hoists, landing gears, boarding steps, and other components

located in anticipated impact areas. Also, particular attention should be given to steel nuts, bolts, and washers that

can contaminate otherwise spark-free areas.

Composites included in aircraft structures have not been considered as potential ignition sources. However, if the aircraft belly primary structure is constructed from composites,

the high bearing pressures and sliding velocities may generate sufficient heat in the form of hot spots to ignite spilled flammable fluids. The relevant properties to ensure that ignition temperatures cannot be developed must be fully investigated for any composite located in such an area of pressure and

velocity.

6.4 ENERGY-ABSORPTION DEVICES

Several methods of absorbing energy have been investigated in

the past. Some may be included in the primary aircraft structure to help control the deformation sequence during a crash;

however, none are applicable for use as major structural members, such as beams. Several possible concepts for accomplishing this aim are illustrated in Figure 25, and a major program

to develop this type of structural member is now in progress.

Energy-absorbing devices that operate under uniaxial tension

or compression and utilize plastic deformation of metal are

described in detail in Chapter 5 of Volume IV, particularly

with respect to application in seats. Certain of these devices

such as those in Figure 30, are also applicable in cargo restraint systems, to be discussed in Section 6.6.5.11. Since

the amount of energy absorbed is a function of the area under

99

J,
Outer tube -Inner tube

Rollers

/Wire Slot in inner tube

(a) Wire bending - absorbs energy by plastic bending of wire over rollers

Outer tube

\ \ Plastically

Thin-walled metal roll tube deforming

(roll) region

(b) Inversion tube - absorbs energy by inverting

a thin-walled tube

Wire helix Outer tube~k

Outer tube

(c) Rolling torus - absorbs energy by rolling wire

helix between concentric tubes

A -Deformed

housing

AA

(d) Tensiontpulley - absorbs energy \• _••// by plastc spreading of the

pulley housing

Figure 30. Examples of energy-absorbing devices.

100

V,,MASK... '?~ I. ~ . ......... 
the stress-strain curve traversed by the absorbing device, it

is advantageous to strain the material well in excess of its

yield poiaL. Care in designing such feaLures into primary

structure must be taken since element failure, or a sequence

of failures, could reuce a perfectly good protective struc-

( tural envelope into a collapsible mechanism resulting in a dangerous loss of living space.

Shock struts used in crashworthy landing gear absorb energy by

the flow of oil through an orifice, where the pressure drop is

sensitive to the stroking velocity. Operation of such a gear is discussed further in Sections 6.5.1 and 7.3.1.

6.5 LANDING GEAR

The design of landing gear for improving crashworthiness presents two definite problems. First, the landing gear must protect the fuselage against contact with the impact surface to

the greatest extent possible. Thus, the landing gear must possess certain strength characteristics and energy-absorbing cap- abilities. The second problem axises once the maximum strength and energy-absorbing capabilities of the landing gear are

reached. The designer must attempt to ensure that the landing

gear does not fail, and in the extreme case, if failure does

occur, to ensure that occupant injury does not result from the

failure.

The energy-absorbing capabilities of the landing gear are

important primarily in the vertical impact, because maximum

occupant protection can be obtained only if every inch of

available stopping distance is used to provide a controlled

deceleration of the fuselage section. Since both rotary- and

fixed-wing aircraft may sometimes impact with relatively large vertical velocities, the landing gear should provide maximum

energy absorption to reduce the velocity of the fuselage cabin

as much as possible before it contacts the ground.

The relationship between energy-absorbing stroke, peak load factor in G, and velocity change can be obtained from the following energy balance:

mv2 = fdS fFd 1max ma (2)

'JJ

- 4i•

,i101
where m w mass of aircraft being decelerated

v - impact velocity

F - maximum force exerted by the structure during

the stroke

S u stroke distance

n efficiency of landing gear system.

Substituting the relationship

max/g= Gmax (31)

where g a gravitational constant

Gmax a peak load factor

amax a maximum deceleration

produces the desired relationship:

2 S - v /2gG (4) "Max (

As an example, an 18.25-G peak load-limited gear with a stroke

of 18 in. and 100-percent efficiency would decelerate the aircraft completely from an impact velocity of 42 ft/sec, which

is the total vertical velocity change of a 95th-percentile survivable rotary- or light fixed-wing aircraft.

This stroke can be achieved by single-stage struts. In practice, the ground clearance and underfloor structure are often

determined by other design requirements, such as ease of ingress and systems installation, that result in a ground-toaircraft belly clearance of less than 18 in.

MIL-STD-1290(AV) (Reference 1) requires prevention of fuselage

ground contact for a 20-ft/sec vertical impact velocity through

landing gear energy absorption and a capability to withstand a

42-ft/sec impact without failure of the gear or attachments.

These conditions can be met with a two-stage strut.

The first stage provides protection up to 20 ft/sec. This requires a strut cylinder design that prohibits bursting when

the internal pressure reaches the level consistent with the

102

, I I I I I I I I I I I I I I I I I I
42 ft/sec impact velocity. After initial stroking and ground impact occur, the second stage and crushing of the underfloor

S:.s structure occur in parallel.

The benefits that can be gained from effective vertical energyabsorbing landing gear design extend even further. The fuselage is protected from impacts until much of the energy associated with vertical velocity is dissipated. With vertical

energy absorption in seats remaining the same, the limits of

survivability are extended. Also, less energy is available

for deformation of fuselage structure. This means that floor

structure remains more continuous, emergency exits are more

likely to operate, and flammable fluids can be more easily

• contained energy-absorbing in fuselage landing tanks. gear for Thus, high provision sink rates of effective yields improvement of all survivability factors. These benefits apply

equally to conventional fixed-wing aircraft and vertical takeoff aircraft.

In view of the variation in efficiency of various energyabsorption devices, a landing gear stroke of at least 18 in.

is recommended. Fuselage structure strength must be compatible with landing gear strength to assure maintenance of occupant

living space during gear stroking if the gear is mounted adjacent to the cabin or cockpit areas.

The gear should provide the specified protection in accidents

that involve an aircraft attitude as shown in Section 5.1,

Table 3.

6.5.1 Wheel-Type Gear

Crashworthy landing gear consists of controlled motion devices

that absorb the kinetic energy of the vertical velocity of the

aircraft by stroking without rebound; this is, in effect, plastic deformation. The energy-absorbing performance of the total

landing gear system is dependent on its load-stroke characteristics and its attachment to primary structural members.

In severe longitudinal impacts typical of fixed-wing aircraft,

kinetic energy levels are so great that effective use of landing gear for energy absorption does not appear to be practical.

In these crashes, sufficient stopping distance to keep accelerations at relatively low levels is usually available, and landing gear failure usually occurs if the gear strikes an abrupt

obstacle, such as the lip of a runway. For longitudinal impact

at velocities up to 160 mi/h, it has been found that less than

103

£ ..
1 percent of the total kinetic energy at impact is absorbed by failure of the landing gear (Reference 26). 4

The problem in designing landing gear for longitudinal impact, then, is not to provide energy absorption capable of protecting

the occupiable portion of the aircraft, but to design landing

gear systems in which failure of the gear does not produce increased danger for occupants., The hazards involved include

rupture of flammable fluid tanks and lines and local penetra- tion of the fuselage shell, particularly in the occupiable

areas.

Some haLards caused by failure of the structure supporting the

landing gear may be prevented by locating the landing gear away

from flammable fluid systems and occupiable areas. If this approach is impractical, landing gear components should be designed so the gear is carried away on impact, with the points

of failure located where minimum damage to critical aircraft

areas will occur. The particular problems associated with a

given aircraft will dictate the method to be used.

Crash impacts may occur with combined longitudinal and vertical

velocities. Provisions allowing the landing gear to be driven

upward and rearward, either into supporting structure or by stroking the energy absorbers, must be made without increasing

impact hazards. One method of accomplishing this is to leave

an open bay behind and above landing gear locations to permit displacement. Another is to place landing gear elements at

extrcme outside corners of the structural enclosure where displacement will occur outside the fuselage profile.

One concept for crashworthy landing gear struts employs valves

that bypass fluid and allow high impact velocities to be tolerated without causing hydraulic overpressure in the cylinders.

Figure 31 shows the load-stroke characteristics for different

landing gear concepts. Curve A shows the response of a fixedorifice strut that is not equipped with a pressure blow-off

valve. As illustrated, the curve is steep, the strut failure

point is reached without absorbing a great amount of energy,

and the energy is proportional to the area under the curve.

26. Reed, W. H., and Avery, J. P., PRINCIPLES FOR IMPROVING

STRUCTURAL CRASHWORTHINESS FOR STOL AND CTOL AIRCRAFT,

Aviation Safety Engineering and Research (AvSER), Division

of Flight Safety Foundation, Inc.; USAAVLABS Technical Report 66-39, U. S. Army Aviation Materiel Laboratories,

Fort Eustis, Virginia, June 1966, AD 637133.

1.04
,,-Load without

attenuating valve

(fixed orifice) (A)

Failure

pointUltimate load limit

Dsg load limit (omloeain

S•Load at design

velocity (B)

-Load after

attenuating valve

operation

(variable orifice) (C)

Landing gear stroke

Figure 31. Landing gear characteristics.

The best concept in Figure 31 for an energy-absorbing strut is

one having a variable-orifice valve. Such a design allows for

maximum operational conditions# as shown in Curve B, and for

impact, as S~high-velocity shown in Curve C, In both instances, ,il loads approaching the maximum allowable within the design coni straints for normal operation, or proof condition, and crash

or ultimate condition, S~impact, are generated.

•i A major problem in designing such a strut is the variable6'orifice valve. Such a valve system must be able to sense the

• high velocities associated with crash impacts and respond to

105 fill

Deign loa limit (normal... oper.....t

.9"i.;.,,•.
change the orifice size in the very short time span available.

The response rate must be very rapid to preclude excessive cylinder pressure and possible rupture.

Variable-orifice valves are not available at present due to the

major problem of producing a reliable device with a response

rate compatible with the short-term transients associated with

crash impacts. Current research into designing variableorifice valves should soon yield a producible strut system

that works to its energy-absorbing limit for both normal operational landings and crash impacts.

Figure 32 presents a helicopter landing gear configuration designed to meet the requirements of MIL-STD-1290(AV) (Reference 1). It incorporates two-stage struts using fixed-orifice

valves and blow-off devices, as indicated, to provide the required level of energy absorption at stroking rates of up to

42 ft/sec. Operation of such a gear is discussed quantitatively in Seation 7.3.1. For comparison with the desirable

strut performance, using a variable orifice, Figure 33 shows

the load-stroke characteristics for a state-of-the-art landing

gear for a helicopter such as that shown in Figure 32. As illustrated, the cylinder pressure, or load, increases rapidly

until the blow-off valves open. This results in a drop in

load-carrying capability, which then remains approximately

constant until the full stroke condition is reached. Then, as

shown, strut motion stops, and the load builds up once again

due to strain energy effects. The energies involved in the

process are indicated, and it should be noted that after the

20 in. or so of travel, the aircraft velocity is 36.3 ft/sec

when ground impact of the fuselage occurs.

A major problem to overcome when designing landing gear instal- lations is the effect of side loads. Helicopter crashes often

occur with some roll attitude (see Section 3.4 of Volume II)

and when impact occurs with trees or overhead wires, a purely

lateral impact can be initiated. When considering crash loads,

it is important to assess the effects of side loading and design landing gear elements and structural attachments to withstand such loads. In the past, crash hazards have been compounded due to loss of landing gear integrity when side loads

caused failure as initial ground impact occurred. When a crash

occurs laterally or with a roll attitude, the aircraft will

often rotate to a sensibly upright position after initial

ground contact. After this, energy can be absorbed by vertical stroking of the gear struts only if complete gear integrity is maintained during the transition from lateral to vertical loading.

1064
P-4 4J 4) t

%~ " .fIu -' Hut

o cl 0') )1 ~O4J ) -W 444 0 *-r4 MP4 4

N 0 41

-1 .- 1 9

'4 4q1 4

Q c~n,. 41 r

0 -1P4u

\\n 13Ifal 4

\\ 4)\k

ro

~Z

0 -H .0

10
Energy absorbed during stroke = 411,000 in.-lb

Strain energy absorbed = 32,000 in.-lb

at end of stroke

Total energy absorbed = 443,00( n.-lb

Final velocity = 36.3 ft/sec

Valve blow-off

X 3

0

H, Strain

energy

4.)

w

,-Wheel and tire effect

0 10 15 20 25

Mass travel after ground contact, in.

Figure 33. Nose landing gear crash

energy absorption.

108

* .'- . r'' '' '•i ''' i........t .. i.•'i ......... ............ ..... - -....... i •. ...
6.5.2 Skid Gear

49 The simplest form of skid gear, which offers advantages of low

cost and weight, is the bent tube attached to structural frame

members. The skid gear can also provide improved support in

landings on soft or marshy terrain. This type of gear provides

a limited level of energy attenuation in heavy landings. The

skid system is basically a nonlinear structure providing a constant spring rate for small deflections only. Such a design

provides a low-cost means of creating small elastic deflections

during normal landings while providing energy dissipation efficiencies comparable to those of an oleo strut. However, in

a heavy landing or crash the skids deflect appreciably and the

nonlinear plastic characteristics of such a system prevail.

During plastic bending the moment arm increases, lowering the

loads as the deflection increases until the fuselage contacts

the ground. A further problem manifests itself when lateral

impact conditions occur due to sideslip velocity or roll attitude in a primarily vertical impact. Then, digging-in of the

skid can cause rollover and/or collapse of the skid under the

belly of the aircraft, conditions that are not desirable for

energy attenuation or crash survival.

Improvements to the fixed-skid concept have been designed in an

attempt to provide some level of energy absorption with rate

damping. In addition, such a system can be tuned to minimize

the effects of dynamic excitations that may produce ground resonance. Figure 34 shows the basic skid system used on one observation helicopter and an improved, pitch-interconnected concept (Reference 27). The improved design was initiated to

reduce the nose-down pitching response of the aircraft and,

hence, minimize the incidence of blade-tail boom strikes. Although tail boom clearances were improved, the new concept did

not satisfy the MIL-STD-1290(AV) requirement for no ground contact by the fuselage for a 20-ft/sec vertical impact (Reference 1). Only 17.5 ft/sec was reached before fuselage contact.

27. Logan, A. H., ANALYTICAL INVESTIGATION OF AN IMPROVED

HELICOPTER LANDING GEAR CONCEPT, Hughes Helicopters;

USAAMRDL Technical Report 76-19, Eustis Directorate, U. S.

Army Air Mobility Research and Development Laboratory,

Fort Eustis, Virginia, August 1976, AD A029372.

109
Fuselage

oleo

dampers

Cross tubes

Drag brace

Modified oleo

damper/sleeve

assembly

pitch initerconnectidoncet

110
6.6 FUSELAGE

S6.6.1 General

Efforts to improve aircraft crashworthiness begin with improvements in fuselage design, since the fuselage provides the occupants' protective shell. in this section, principles for improving fuselage structure and methods for accomplishing these

improvements are presented.

In order to conduct a meaningful program aimed at improving overall structural crashworthiness, a means of measuring crashworthiness is essential. Two indices of crashworthiness that

have been proposed are:

* The degree of cabin collapse, under standard crash

conditions, chosen through consideration of the respective aircraft operating characteristics.

o The level of acceleration experienced by occupants

during the crash.

Whenever structural design, or changes in design, brings about

a reduction in the degree of collapse that may be expected in

a crash and/or in the level of accelerations experienced by occupants, the probability of occupant survival is increased.

Acceptance of these indices as structural crashworthiness criteria permits a crashworthiness comparison of various config- urations subjected to similar crash conditions and leads to

consideration of several methods which offer promise ur increasing the probability of occupant survival in aircraft accidents. These include:

1. Increase in the energy-absorption capacity of the

structure surrounding occupiable areas to provide added protection for these areas.

2. Alteration of the structure that makes initial contact with the ground to reduce gouging and scooping of soil, hence lowering accelerations and transmitted

forces.

3. Modification of structures such as the antitorque

rotor, empennage boom, support system for the external stores, and, in certain cases, the landing gear

to ensure that if these parts fail, they fail safely. Component breakaway during a crash can be used to effect a reduction in the mass of the aircraft, hence

reducing the strength requirement for the cabin

structure.

.... . ......... S.. 1 1 1
4. Reinforcement of cockpit and cabin structure to enable the structure to withstand greater forces without collapse.

5. Modification oZ fuselage structure to allow increased

deformation or collapse of structure in unoccupied regions, thus permitting additional structural energy absorption.

The shape of the fuselage has an inherent influence on all

these methods except 3. Rectangular cross sections can be

designed to provide the same crashworthy characteristics as

spherically, cylindrically, or elliptically shaped fuselages/ however, in practice they are not. The circular or elliptical

cross sections normally are stronger structures. Also, the

cavities between curved fuselage skin and flat floors or essentially flat inner walls provide volume for the inclusion of

energy-absorbing material. The result has been that curved

fuselage configurations are generally more crashworthy than

rectangular ones. Figure 35 shows a light observation helicopter shape and component layout advantageous for providing crash impact protection (Reference 28).

To determine the potential contributions of all possible methods of improvement, it will be necessary to evaluate each aircraft on the basis of its own structural characteristics and

the anticipated operating characteristics which indicate prob- able modes of impact.

An understanding of the potential influence upon arashworthiness of methods 1, 2, and 3 above will be helpful in selecting methods of structural improvement which will effectively and

efficiently increase the probability of occupant survival.

6.6.2 Energy-Absorption Capacity of Forward Fuselage

From a consideration of conservation of energy, the initial

kinetic energy of an impacting aircraft must be accounted for

in energy dissipated during the deformation of both soil and

structure. Therefore,

mA(Vo2 - vf2

2 UG + Us (5)

28. Smith, H. G., DESIGNING HELICOPTERS FOR IMPROVED CRASH

SURVIVABILITY, paper presented at NATO/AGARD Aerospace Medical Panel Specialist Meeting, Oporto, Portugal, 23, 24,and 26 June 1971.

112
I

04J 'o

4j -H ~ r44H

4) 000 a

'U

00

4) 0

A 41)

4J ~30 IP4 u r-I 0 4 Q) 4

to 44 -Ho - Aa r404

00 0

P4.

ri WU

-r-4 0 1

0 4U

41 ) .0 U 1

to N4 0)I

HQ U)4

I 4J 4113
where mA - mass of aircraft

V0 * initial impact velocity of aircraft

Vf a velocity remaining after impact

UGa energy dissipated in soil deformation and ground

friction

U energy dissipated in structural deformation.

This equation states that the reduction in aircraft kinetic

energy must be equal to the energy absorbed by deformation of

earth and structure and is simplified by the assumption that

there is no change of aircraft mass during the impact.

As a simplified model, the structural energy absorption, Us,

may be expressed as

u Ps u + (6)

where P *average force developed in collapse of structure

e aforward of the cabin

S = linear deformation (reduction in length) of

structure forward of the cabin

4! u;a deformaticn energy in structure other than in

the cabin or structure forward of the cabin

U - energy to be absorbed in cabin deformation.

The cabin deformation energy, U , represents the quantity of

energy absorbed by deformation Sf cabin structure, and may be

obtained from Equations (5) and (6)t

Fm (v02 Vf 2 1

UC 2 UG] - (PavS + Us) (7)

This equation for cabin deformation energy is valid if conditions reach or exceed the point of onset of cabin deformation.

Assuming a fixed mass aiAd velocity and ignoring control over

energy dissipated outside the aircraft, the factors that are

controllable in Equation (7) are Pav' S, and U

114

li. .
Consequently, U , the energy which must be absorbed in cabin

collapse or coliapse of the protective shell, may be reduced 3 by.

Increasing P , the average crushing force acting

during collaXe of structure forward of the cabin.

P may be increased for a given maximum collapse

fSyce by providing forward structure which will

maintain a force as nearly uniform as possible during collapse. In addition, P may be further in- , creased by admitting an increaf in the maximum force

applied to the forward structure. This latter option

is limited, however, by the existing strength of the

cabin. If, for example, the maximum collapse force

for forward structure were to exceed the cabin collapse strength, then the energy-absorption objective

would be defeated, as the cabin deformation would

commence prior to full collapse of forward structure

and, therefore, prior to full energy dissipation in

"the forward structure.

Another point to be considered is the effect of forward structural modifications upon the second crashworthiness index, occupant acceleration. If the

maximum collapse force is increased, aircraft acceleration will increase, adversely affecting this

second index. The tradu-cff between the energy- absorption benefits of increasing P by increasing the maximum force and the detrimentalveffects of in-

"creasing maximum aircraft acceleration inust therefore be considered.

e Increasing the available deformation distance, S,

which would also permit greater energy absorption in

the forward structure. (This could be accomplished

without increasing the maximum collapse force.)

This factor should definitely be considered in original design. It could be accomplished by placing the

cabin as far aft as practical, as on modern agricultural aircraft.

e Increasing the deformation energy absorbed in air-

"craft structure other than forward structure or cabin

structure. This would further contribute to a lower

cabin deformation energy requirement. Structural design which permits plastic deformation of structure

at selected points away from occupiable areas could

accomplish this.

115 ie
Application of any of these principles to the airframe design,

with a given cabin structural configuration, will make it pon- sible for the aircraft to withstand impact at increased velocity without collapse of the protective shell, thus extending

the limits of survivable conditions and improving survivability in crashes under less severe conditions.

6.6.3 Reduction of Earth Plowing

Under certain conditions of impact and structural deformation, the forward section of an impacting aircraft, particularly a

helicopter with a blunt nose, deforms to become a scoop, picking up a mass of earth and "driving" it to the velocity of the

aircraft. This process is accomplished in a very short time

intervall therefore, the principle of conservation of momentum

may be Applied to the system, which includes the mass of soil

that must be accelerated, referred to as the effective mass of

earth, and the aircraft mass. Accordingly, conservation of

momentum leads to the relationship

mAvo - (mA + mE)v (8)

where mA - mass of aircraft

mE - effective mass of accelerated earth

vo - initial impact velocity of aircraft

v = velocity of combined system immediately after

impact.

Solving Equation (8) for v, we obtain

v - mA +mEA vo (9)

To find the interaction force involved in the momentum exchange,

an impulse-momentum relationship may be applied to the earth

mass as a free body.

t2

f - mEv (10)

tl

116
where F a interaction force

'Vt - time just prior to impact

t 2 - time directly following momentum interchange.

By definition of the average force,

t2

] Fdt FavAt 11)

ti

where At - 2 - I

Substituting Equations (9) and (11) into Equation (10) yields

Fa (mA'2mE) V (12) Fay m• A m E At 12

Consequently, the average acceleration of the aircraft mass due

to the acceleration of the earth mass is

mEvv 

a au- ME - a (13) mA m -A + m, At

Also, mE * KAV At (14) E 0

where K is a constant and A is the cross-sectional area of the

gouge in the earth.

Thus,

, ~KA'o2

a A mA + KAvoAt (15)

117
which indicates that the deceleration of the' aircraft varies

with the square of the initial velocity (where the scoop effect

is a dominant factor and At is small). Thus, at high impact

velocities, the scoop phenomenon assumes a greater significance.

Experimental evidence substantiates the fact that impact ac-

"celerations do increase with the increasing impact velocity;

however, maximum aircraft acceleration due to impulsive acceloration of earth is limited by the strength of the aircraft

structure. Experimental evidence also indicates that, under

high rates of loading, the force transmission ability of structures increases due to the stabilizing effect of lateral inertia during buckling of the structural members.

As an example of the use of Equation (15), assume that the following conditions exist:

0 "v 140 ft/sec

At 0.02 sec

ME 0.185mA

Then (a mE 0v ,

A mA + mE At

aA E)

a /0.1 85\ (140

aA 1092 ft/sec2 .34 G

If the impact velocity, v , and the time interval, At, remain

unchanged but the effective earth mass, mi, is reduced to

0.10 m., the average impulsive acceleratioE of the aircraft becomek

aA 0.1) (14 637 ft/sec 2 19.8 G

Increasing the impact velocity, v , to 160 ft/sec (under the

aforementioned assumption that accbleration varies with velocity squared), when m 0.185m., results in an aircraft acceloration Ers n ic

S11 8 f, ,,

• I.'• , V
1160\2 f/ec 2 

a A = (A 1092 1430 ft4 4

And if a mass of earth equal to 0.lm is accelerated under an

impact velocity of 160 ft/sec, the aarage acceleration is computed to be 25.8 G.

Figure 36 shows a family of curves relating impulsive aircraft

acceleration to the ratio of effective earth mass to aircraft

mass for various impact velocities. The duration of the impulsive loading varies with impact velocity, although the

curves are based on a single assumed value of 0.02 sec for the

140-ft/sec impact.

In addition to the force associated with momentum exchange,

soil penetration by projecting structure gives rise to a drag force, sometimes called the "plowing effect." This force adds

to other soil reactive forces. The plowing force should be

distinguished from the impulsive force associated with momentum exchange, because it is a steady-state force depending upon velocity, soil strength and density characteristics, and projected area of interference. It should be noted, however; that

a design which effectively reduces the scoop effect also helps to reduce the plowing effect contribution.

6.6.4 Reduction of Aircraft Mass

As suggested earlier, it is possible to reduce the aircraft

mass through planned breakaway of portions of the aircraft during a crash. An analysis of the influence of the reduction of

aircraft mass is presented here.

The expression for cabin deformation energy, U , which was obtained from the solution of Equations (5) and •6), is repeated below:

v 2 f UGi " avs + Uv) (7)

Changes in effective aircraft mass during a crash would leave

the energy absorbed in collapse of forward structure, P S,

and the energy absorbed in plastic deformation of other %ircraft structure, U , essentially unaffected. The energy dissipated outside thn aircraft in soil deformation, U , would,

on the other hand, be influenced by the aircraft mags. This

soil deformation energy assumes several forms: principally, the energy associated with superficial friction and that involved in the plowing, scooping, and compression of soil.

119
a) 0

00 K Ow K,0 0

.F44

0

'N7

VW M

44 4

00

fI4-4

0 41

0-4 0

120
With a reduced aircraft mass, contact forces would tend to

be less, and energy dissipated in friction would be reduced.

Also, with reduced aircraft mass (assuming that the force

transmitted through deforming structure is controlled by structural collapse strength and is therefore constant with respect

to mass change), less time would be required to accomplish a

given velocity change. This would allow less time to dissipate energy in soil deformation. Consequently, a reduction in

mass of the ai-craft would also serve to reduce the energy absorbed at or within the ground. As a plausible approximation,

in the absence of a developed soil dynamics study, the magnitude of UG was assumed to be proportional to the aircraft mass.

Denoting the effective aircraft mass after breakaway of por-

"tions of the aircraft by m ', an expression may be written for

the cabin deformation energy with reduced mass:

av5Aus (16)

where U denotes the soil deformation energy obtainable without redortion in mass.

A numerical example will serve to illustrate the influence of

mass reduction more clearly. In this case, it is assumed that

the soil deformation energy for a given accident environment

is equal to 70 percent of the initial kinetic energy, or

[m_( 2

UG 0.7 [ ° J (17)

"Then the equation becomes

I (V 2 - vf2] - (PavS + U•) (18)

121
r.

and, for the following impact conditions,

Aircraft weight - 8000 lb

mA -250 slugs

v 0 -140 ft/sec

vf -80 ft/sec

PavS - 300 x 10 ft-lb

U, - 50 x 103 ft-lb

The cabin deformation energy is

m (495 x 103) - (350 x 103) ft-lb

C mA

Consequently, if there is no reduction of the mass of the aircraft during the crash, the cabin deformation energy is

U - (145 x 103) ft-lb

For a mass reduction to 0.85 of the original mass, the requirement for cabin deformation energy is reduced to

(Uc) - 0.85 (495 x 103) - (350 x 103) ft-lb

0. 8 5mA

(UC) - (71 x 103) ft-lb 0.85mA

This illustration shows the reduction of cabin deformation energy which is possible through a small reduction of aircraft

mass.

122
EF 6.6.5 Design Concepts for Improved Crashworthiness

6.6.5.1 General: In the foregoing discussion, it was noted

that improvement of airframe crashworthiness is accomplished through improvement in either one or both of two survivability It 'factors: the ability of the protective container to maintain

living space for occupants during a crash, or the attenuation

of accelerations experienced by the occupants during a crash.

When considering any design concept for improving structural

crashworthiness, the survivability factor to be improved must

be kept in mind and, additionally, the energy-dissipation characteristics of the crash must be understood.

There is a basic difference between the dissipation of kinetic

energy in crashes which are primarily longitudinal (high velocity, low angle) and those that are primarily vertical.* In

longitudinal impacts onto relatively flat surfaces, a high

percentage of the initial kinetic energy of the aircraft is

dissipated in compressiton and acceleration of masses of earth

and in friction between the aircraft and the earth. Consequently, in this type of longitudinal crash, a relatively low

percentage of the initial kinetic energy is absorbed by structural deformation. In primarily vertical impacts, longitudinal impacts into barriers, and lateral impacts, much more of the

initial kinetic energy must be absorbed by the structure.

This leads to separate consideration of design concepts for

improving crashworthiness under primarily longitudinal impact

conditions and under other impact conditions.

6.6.5.2 Improvement of Crashworthiness in Longitudinal

Impacts: Since, in primarily longitudinal crashes, the compression and scooping of earth contribute a major part of the

floor accelerations experienced within the cabin, the first

topic presented is a discussion of methods for reducing the

scooping or plowing of earth.

When the forward sections of an aircraft deform so that a scoop

is formed and earth is impulsively accelerated, two adverse effects may occur. First, high acceleration of the aircraft may

occur. Second, the large forces required to accelerate the

earth mass may be concentrated in a small area and cause local

*As used here, "longitudinal" refers to the orientation of the

velocity vector with respect to the impact surface, i.e.,

parallel to the impact surface. "Vertical" impact refers

primarily to the autorotation type impact for rotary-wing

aircraft.

123
collapse of the cockpit protective shell. Reduction of earth

scooping will thus lower fuselage floor acceleration levels and

tend to prevent the buckling of the protective container.

Reduction of earth scooping can be accomplished by structural

design which eliminates those surfaces that can gouge or dig

into terrain. The structural design should provide a large,

relatively flat surface so that the aircraft skids along on

top of the terrain.

One design method involves the prevention of inward buckling

of fuselage lower nose skin. This would be similar to preventing a thin spherical shell from buckling under excessive uniform loading. A method of accomplishing this is shown in Fig-.

ure 37. The design method shown may not increase the longitudinal strength of the nose section, but it will reduce the

tendency to buckle inward due to vertical loads distributed

over the lower nose surface. If this design is to be effective,

the lower skin should be a ductile, tough material with enough

thickness to resist tearing. The skin should remain continuous

to provide a skidding surface. It is recommended that the forward fuselage belly skins on aircraft weighing up to 3,000 lb

be capable of sustaining loads of 1,500 lb/in., over 3,000 lb

but under 6,000 lb, 2,400 lb/in.j and over 6,000 lb, 3,000

lb/in. The above running loads are to be applied over the forward 20 percent of the basic fuselage length.

This method for reducing earth scooping increases the deformation strength of the underfloor structure, thus increasing the

deceleration levels at the floor level due to vertical velocity components. Therefore, the reinforcement should not be

continued back any further than necessary under occupied sections of the fuselage. Modifications to reinforcing structure

under occupied regions are presented in Section 6.6.5.3.

Strong structural crossmembers can present abrupt contour

changes during deformation, thus forming the lip on the scoop

that tends to trap earth. Forward underfloor frame members

may be canted aftwards at the bottom to provide an upward load

component on the aircraft that tends to prevent, or limit, digging in of the structure. The longitudinal strength of the

nose section can be increased by the use of strong continuous

structural members running fore and aft in the underfloor section of the aircraft. These beams can be used to support the

crossmembers and act as skids to further reduce scooping or

digging-in tendencies. Figure 38 illustrates this type of construction.

124

3L.

'.ii.£
4=f4

Figure 37. Method of rainforcing nose structure to provide

increased resistance to vertical loads and to

reduce earth scooping.

125
I.,

Figure 38. Typical underfloor canted frame ana

longitudinal beam member to minimize

nlose plow~ng.

Additionally, the underbelly skins, made from thick sheet material, are shingled in an aftward direction to preclude their

picking up at the front edge. These features are illustrated

in Figure 39.

These typical structural features of a passenger-carrying helicopter may be used also for other types of aircraft, at least

in principle. obviously, certain oporational and structural

features require special attenti.on when crashworthy primary

structure in designed.

often, in aircraft with the eng±ihe mounted in the nose, or in

the design of engine nacelles, otructural bracing as discussed

W above is not practical because the structure beneath the enNgine is a very light secondary structure made up of removable

doors and cowling. The engine and engine mc*ýnts, however, may J, be strong enough to support a skidding surface if a filler is

provided between the engine and the lower skin. The filler

may be made of lightweight plastic foam or honeycomb material,

contoured to fit the lower surface and to fill as much space

as possible between the skin and the engine# as illustrated in

the lower sketch of Figure 40. The attachments of removable

1* 126
Overhead

buttline, beams

Windshield

r posts

Cockpit Cabin

Longitudinal

floor beams

Thick-shingled

belly skins

Figure 39. Features of helicopter nose section

to prevent nose plowing.

127

Y
Normal configuration

Design of firewall to

r remove the abrupt surface

Addition of energy-absorbing

filler to prevent scoop from

developing under impact forces

Al

Figure 40. Two methods of reducing earth scooping

in engine-mounting areas.

access doors and cowling should logically be strong enough to

sustain the impact loads, even when considerably deformed.

In multiengine aircraft, the engine nacelles may present as

much of an earth scoop as the nose of the fuselagel and, since

the engines are often attached to the strong, rigid wing conter section, the forces produced by engine earth scooping are

transmitted to the fuselage cabin.

128

.. .........
Use of the design concepts discussed above can be helpful in

reducing the harmful contribution of engine and fuselage nose

earth scooping. The results of full-scale experimental crash

tests indicate that longitudinal aircraft accelerations produced by earth scooping can be significantly reduced by the

application of these methods.

Most aircraft crashes occurring at impact angles up to 30 degrees involve a rapid change in pitch attitude to quickly align the aircraft fuselage with the impact surface. The resulting dngular acceleration produces a fuselage bending moment which usually causes a compression of upper members of

the forward fuselage. This compression is combined with compression of the fuselage due to the longitudinal forces of

impact. The result is a compressive buckling failure of fuselage structure. When the failure occurs in personnel areas

along the fuselage, the protective shell is compromised.

It is sometimes possible to provide sufficient strength to

prevent fuselage bending failure. If this is not practical, it is desirable to determine the probable failure points and

to position passengers away from those locations to minimize

the risk of injury in a crash. Cargo tiedown attachments

should be designed to prevent loss of cargo restraint, should

fuselage bending failure occur.

The design of floor structure and the floor support structure

also has a considerable effect upon the protection offered by the cabin structure as a protective shell. The floor structure should be strong enough to carry the loads which will be

applied to it by passenger and cargo restraint systems (see Sections 6.6.5.10 and 6.6.5.11) without the need for supports which carry through to the lower fuselage skin and stiffeners.

Such supports transfer crash loads applied at the outer fuselage surface directly into the floor. Buckling of the floor

produced by such a load transfer can reduce the bending strength of the fuselage and also reduce the effectiveness of

restraint systems which depr id upon floor integrity. Evacuation of the aircraft can also be impeded by interruptions in

the floor surface.

MIL-STD-1290(AV) requirements specify the following longitudi- nal impact protection: *At 15 to 20 ft/sec impact velocity, no

hazard shall be posed to the pilot or copilot, and at 40 ft/sec,

a 15-percent reduction in volume of the troop/passenger compartment must not be exceeded. Additionally, inward buckling

of side walls shall not pose a hazard and postcrash egress must

129

--
not be impeded." The major structural elements of a helicopter

designed to satisfy these crashworthiness requirements are emphasized in Figure 41 for the nose section, and in Figure 42

for the cabintcenter section. The nose section structure is

basically a propped cantilever with the underfloor structure

being the cantilever and the props being the windshield posts

that extend upward to the transmission-support buttline beams.

Longitudinal loading is reacted primarily by underfloor beams

that resist bending and secondly by windshield posts that

transfer loads to the transmission buttline beams. Underfloor

frame members and the outer skin assemblies provide lateral

stability for the longitudinal members.

Overhead

Longitudinal Beams

Windshield

Posts

Longitudinal

Floor Members

Floor Crossmembers

Figure 41. Longitudinal impact and blade strike protection:

nose section. The figure delineates those

members requiring careful design consideration.

In the past, other proposed methods of absorbing energy have

suggested the use of masses of crushable material. Honeycombs,

expanded foams, or similar materials share the major disadvantage of using relatively large volumes of installation space.

Such space usage can result in system installation compromises,

enlarged profiles, and increased primary structure weight to

accommodate such energy attenuation methods. They do, however,

still provide a method warranting consideration and trade-off.

130 )
Longitudinal beams

!.ff (high compressive

load capability)

tudinal floor beams

load capability)

Figure 42. Longitudinal impact protection in cabin section.

The cabin structure, by virtue of specific aircraft mission requirements for rapid troop ingress and egress, contains a large

area of cutouts necessitating a longitudinal beam system for

carrying major loads. In the upper vertical structure, two

buttline beam. continue into the cockpit section to provide

continuous paths for crash loads in the longitudinal direction, Figures 43 and 44 show details of typical beams. Loading at and below the floor level is transferred from the cookpit through the cabin by longitudinal floor beams.

6.6.5.3 ieomntof Crashworthiness in Vertical Imiacts:

For a primarTiyvrtical impact (or for the vertical component

of any impact), structural energy-absorption requirements differ appreciably from those of a longitudinal impact. In a ver7 tical impact, th~ere exists no possibility of low-force-level

(and hence low acceleration level) energy absorption external

to the aircraft comparable to the frictional energy absorption

in a longitudinal skid. The velocity change in the vertical

131

yl .
Figure 43. Typical longitudinal beams adjacent to cockpit for longitudinal continuity of overhead

structure.

Figure 44. Typical full-depth longitudinal beams for overhead

support of large mass items and longitudinal continuity of structure.

132

. . . .. . . .... . S... • ',,,,,...'..
direction must be accomplished in a short time interval. Consequently, when the vertical velocity component is high,,

crashes are generally characterized by large structural deformation and high accelerations at aircraft floor level.

Several studies have treated methods of reducing the effects of

the high floor accelerations upon occupants. These studies

have resulted in recommendations for providing energy-absorbing

passenger and crew seats to protect occupants in crashes at energy levels which are survivable from the standpoint of general

cabin collapse (see Volume IV). It is also apparent that increasing the energy-absorbing stroke of the floor through vertical strength reduction can reduce floor acceleration levels.

"In order to evaluate potential improvements in crashworthiness

of cabin structure for vertical impacts, two idealized extreme

configurations are presented that serve to point out problem

areas. First, consider a fuselage section in which the aircraft mass is concentrated at the top of a fuselage section.

This model is schematically illustrated in Figure 45.

Mass

j,•' .•. Initial ,::! '

velocity :Nonlinear

ve iFloor

Impact surface

Figure 45. Fuselage structure with overhead

large mass items (idealization).

•'A.. 133
The fuselage acts as a nonlinear spring which is intially elastic and remains so for a moderate deformationi thereafter, the

spring force reaches a critical value that produces a plastic

collapse of the support structure.

For such a model, a vertical impact requires that substantially

all of the kinetic energy of the mass be converted to deformation energy of the structure. If this kinetic energy is large,

complete collapse of the structural support, i.e., the cabin,

can occur, depending, of course, upon its strength.

This model of the fuselage can reasonably be extended to include a crushable region below the cabin floor which would deform plastically for forces below the critical load for the

upper fuselage sidewalls. This crushable region would then

absorb energy, providing an increased buffer against general

collapse.

The opposite extreme of this fuselage model consists of a

structure of negligible weight with the aircraft mass concentrated at or near the bottom of the fuselage. Figure 46 illustrates this configuration schematically. Upon vertical impact

with this second configuration, without a crushable subfloor

structure, the kinetic energy of the mass would be largely dissipated in the soil with perhaps high acceleration levels occurring at the floor. The light upper structure in the cabin,

however, would not collapse even for high-energy impacts.

Lightweight

upper structure

Initial

velocity

Mass

M- Irmpact surface

Figure 46. Fuselage structure without overhead

mass items (idealization).

134

i ' I I II I
If crushable subfloor structure were included in the model,

then a sizeable portion of the kinetic energy could be dissipated in subfloor deformation. The subfloor energy absorption

could serve to attenuate the floor acceleration.

Actual conditions generally fall between the two idealized extremes presented above. To the extent that large masses are

secured to the upper fuselage in typical helicopters or high

fixed-wing aircraft, cockpit/cabin collapse presents a serious I problem. To the extent that mass is concentrated at cockpit/ cabin floor level, energy absorption in the subfloor structure,

and the associated acceleration attenuation, may assume significant beneficial proportions in reduction of floor accelerations.

Considering again the two proposed crashworthiness indices -

(1) extent of cockpit/cabin collapse and (2) cockpit/cabin

(floor) acceleration level - an evaluation may be made of potential improvements in crashworthiness for vertical impacts

for a given design.

The threat of general cockpit/cabin collapse under vertical

impact may be reduced in any of several ways: First, to the

extent that it is feasible, either in original design or by

modification, or in cargo and equipment tiedown, a transferral

of mass from the top of the fuselage to the floor would be beneficial. Second, a general strengthening of cabin structure

may be effected so as to increase the structure's resistance

"to vertical collapse. Localized strengthening at locations of

large concentrations of mass attached to upper structure will

provide an increased resistance to general vertical collapse. Third, modifications in fuselage structure that increase elastic energy absorption or provide for plastic energy absorption

at loads less than the general collapse load help to maintain

the primary cockpit/cabin integrity. Fourth, any increase in

energy absorption in the subfloor structure realizable at load

levels below the fuselage collapse load further helps to protect the cockpit/cabin against collapse. The threat of high

vertical acceleration at the floor may also be reduced by providing subfloor crushable structure. The crushing resistance

of the subfloor structure may be optimized to meet the conditions of the anticipated vertical impact velocity and the mass

associated with the cockpit/cabin floor. The conceptual cap- and-web beam design illustrated in Figure 47 (Reference 29) is

29. Haley, J. L., Jr., HELICOPTER STRUCTURAL DESIGN FOR IMPACT

SURVIVAL, paper presented at the Joint Symposium on Environmental Effects on VTOL Designs, Arlington, Texas,

November 1970.

135
Low-density, highi i strength filler foam

Formed, high-strength,

low-ductility sheet

Before impact

After impact

Figure 47. Cap-and-web combination beam design with potential

energy-absorbing capability. (From Reference 29)

136

)J

.......................... "~
shown used under a strong floor structure reinforced by conventional beams in Figure 48. The configuration results in the

desired combination of a nonscoop-forming keel design, a continuous strong floor, and an energy-absorbing understructure

that provides the crashworthy design concept illustrated in

Figure 49. The belly skin material should be tough and ductile

to permit deformation without tearing. Energy-absorbing thickness is shown in the side walls for lateral impact protection.

"Also, forward underfloor frame members may be canted aftwards,

as previously shown in Figure 38, Section 6.6.5.2, to provide

an upward load component on the aircraft that tends to prevent,

or limit, digging in of the structure.

A 42-ft/sec impact vertical velocity component, with varying

degrees of pitch and roll, is specified in MIL-STD-1290(AV)

(Reference 1). Allowable vertical deformation of the cockpit/

cabin for this condition is limited to 15 percent. However,

for overhead-mounted crashworthy seats, adequate stroking distance must be ensured after consideration of plastic and elastic deflections of the structural elements.

Figure 15, Section 6.2.1, shows a typical structural layout

for a single-rotor, utility-type helicopter with large mass

items mounted overhead and large door cutouts for rapid ingress

and egress. To prevent excessive intrusion into the occupied

cabin, the support of large mass items is of primary iniportance. This is achieved by using deep longitudinal buttline

beams and built-up box-frame members. The box-frame structure

is achieved by adding skin to the inboard and outboard profiles

of two adjacent frame members, thus producing a closed-.box

structure. Such a section is stronger and more stable under

compressive loading conditions than two individual frame sections acting together. Figure 50 shows typical sections

through box frames of this typea section variation can be tailored to the anticipated loading conditions at any point in

the structure. When using this type of construction, care

should be taken to provide adequate inspection capability in

enclosed areas where corrosion may occur and cause potentially

hazardous structural strength degradation. Alternatively, adequate protective coating must be used and wet assembly techniques employed to provide protection against corrosion for

time periods equivalent to major structural overhaul intervals.

Widely spaced frames -that are skinned on the inside profile may

be fitted with detachable skin panels to allow the location of

equipment and system elements in the space between the frames.

However, only nonhazardous equipment should be located in such

areasl combustibles and potential ignition sources are not to

be installed in such locations.

137, 
View A-A

Continuous formers

(high strength and

elongation)

Ductile floor sheet

Conventional high- strength beams

High-strength

sheet

Keel beams Belly skin

(.see Figure 47) tough, ductile,

and thick

ON'

•.Crushable belly A

Crushable nose

Figure 48. Conceptual structural configurations to

absorb maximum energy for sideward, longi- tudinal, and vertical impact forces.

The underfloor structure, landing gear, and crashworthy seats

are other structural elements providing a measure of occupant

protection from the effects of a vehicle impact.

The optimum energy-absorption capability of underfloor structure is a direct function of the available stroking distance

between the cockpit/cabin floor and underside of the aircraft.

This distance varies for different aircraft types and results

138

LIJ A,
Velocity

',r'

Continuous high-strength

{• keel beams prevent floor •+•. •jcollapse

Figure 49. Crushable belly deforms inward without

buckling floor or reducing living space.

in differing landing gear-to-structure-to-seat energy-absorption

distributions. This emphasizes the need for a crashworthy seat

designed as an integral part of a helicopter and suggests that

a common seat design cannot be used efficiently for several

types of aircraft.

With a possible exception for small single- or twin-seat helicopters, large mass items and their support should be designed

to eliminate overhead threat. If transmissions do break loose

and pivot forward, some overhead protection against the subsequent movement is necessary. Protection against blade impact

and/or impact by large mass items that have pivoted forward or

completely broken loose is required. Figure 51 shows details

of the overhead protection provided by structural continuity of

the windshield posts and longitudinal beams mating with the

cabin section.

Finally, energy-absorbing landing gear can be used to reduce

the severity of cabin accelerations for less severe impacts.

To illustrate the reduction in floor acceleration that can be

139

;I• • .,
I

Sd outer skin

Formed

frame section

Diaphragm Inner skin

.•• Outer skin

Extruded Formed

angle frame member

Formed angle --Inner skin

Diaphragm

Figure 50. Examples of closed box-beam frame sections.

140
II

Figure 51. Typical cockpit overhead

lonyitudinal beam member.

achieved through proper design, a trade-off was made of a specific helicopter using the analytical model described in Reference 30. Figure 52 shows graphically the reduction in peak

floor accelerations achieved by the redesign of landing gear,

fuselage belly, and both landing gear and fuselage belly. The

gear redesign included an increase in strength from 4 to 8 G

and an increase in stroking distance from 9 to 15 in. A fuselage belly strength of 19 G was assumed with the capability of

a 15-in. stroke and 0-G/in. plastic slope as opposed to 22 G

with a plastic slope of 20 G/in. for the standard design. Decelerative forces were calculated for the fully loaded configurations Note that a 20-percent reduction was realized for

the landing gear improvement and 45 percent for the fuselage.

Both improvements provide a total reduction in floor G level

of 65 percent for the helicopter involved in a 30-ft/sec vertical impact.

30. Gatlin, C. I., et al., ANALYSIS OF HELICOPTER STRUCTURAL

CRASHWORTHINESS, Volumes I and II, Dynamic Science, Division of Marshall Industriesy USAAVLAHS Technical Reports

70-71A and 70-71B, Eustis Directorate, U.S. Army Air Mobility Virginia, Research and Development Laboratory, Fort Eustis, January 1971, AD 880680 and AD 880678.

141

S- . - •-',.*--.*-*u- u**".,I***.-I'-I--I- I-,I I I l 1
4J 1

improved fuselage

50 100

Reduction in peak acceleration, percent

Figure 52. Floor acceleration comparisons (30-ft/sec

vertical impact). (From Reference 30)

The principles discussed in this section for providing increased resistance to vertical collapse and reducing acceleration levels for vertical impacts are quite similar to the principles discussed earlier for longitudinal impacts. Therefore,

no further discussion of the relative merits of the use of each

method is necessary. However, it is important to realize that

these principles for improving fuselage crashworthiness for

longitudinal, vertical, and lateral impacts are applicable to

the design of both conventional takeoff and vertical takeoff

aircraft. Consideration of these aspects of crashworthiness

becomes more important as the normal approach sink rates of

conventional aircraft increase and as helicopters develop

greater forward approach speeds.

6.6.5.4 Lateral Impact and Rollover Protections The loadcarrying members indicated for vertical and rollover protection in Figure 15, in general, provide protection against lateral impacts and must be designed for this function.

Rollover protection must be provided to withstand loading applied to the structure during the dynamic stage when rotation

is occurring and after the aircraft has come to rest in any

142
attitude. In addition, sufficient emergency exits must be accessible and operable with the aircraft in any postcrash attitude. The required occupant protection for this type of impact is specified in Table 3, Section 5.1.

Figure 53 indicates one approach to providing the primary loadcarrying structure needed to resist the effects of a postimpact rollover. The structural elements needed for similar

protection in the cockpit area are emphasized in Figure 54.

Longitudinal beams

(high bending capability)

Box-beam frame members

(high bending and compressive load capability)

Figure 53. overhead mass protection and rollover

protection in cabin section.

S * Rollover protection in the cabin section is provided by the

buttline beams, longitudinal floor beams, and main box frames.

The cockpit of the helicopter shown in Figure 15 is protected

by a mesh of structural elements, the majority of which are

143
Overhead

longitudinal beams

Contoured

Windshield posts frames

Longitudinal

floor members

Floor beams

Figure 54. Rollover protection in nose section. The

figure delineates those members requiring careful design consideration.

door and transparency support members. These are shown in Figure 541 details of the sections and materials used are shown

in Figure 55. Reference 31 contains more detailed information

on lateral rollover protection.

Prevention of main rotor blade entry into the cockpit is an

extremely difficult task due to the conflicting requirement of

good pilot visibility. Structural members that can deflect a

main rotor blade will have to be large in size, thus restricting the pilot's visibility. A roll cage cockpit that could

minimize blade intrusions would include large structural beams

overhead and on the outboard sides. High-strength windshield

posts and door posts that attach the overhead structure to the

floor beams provide the remainder of the roll cage.

High-strength deflector beams might be installed forward of

the windshield rather than as windshield posts, as shown in

Figure 56. Moving the beams further away from the pilot will

minimize the pilot's visibility restriction. Such external deflector beams could also deflect wires up over the cockpit.

31. Fox,, R. G. LATERAL ROLLOVER PROTECTION CONCEPTS, Bell

Helicoper Textron, USAAVRADCOM Technical Report 80-D-1,

Applied Technology Laboratory, US Army Research and

Technology Laboratories (AVRADCOM), Fort Eustis, Virginia, January 1980, AD A081420.

144

•. 'i
inin

0: w 0 

IA I-In

hi -

14 0 04

ý4.

145-
I..

IAi

Oft

•~ 4,

r 

0

tU

.44
' 6.6.5.5 Rotor Blade, Wing, and Empennage Design: Plastic

• - deformation of aircraft elements outside the occupied area is

an additional method of absorbing aircraft energy. Elements

such as rotor blades, wings, tail sections, and even external ,~, ~,,stores can be bent, battered, or destroyed in the process lilt of bringing the occupied section to a successful halt. When

trees, ground obstacles, or uneven terrain are involved, impacts with blades, wings, etc., often occur. Energy is expended in breaking the impacted object, or failing the aircraft

alement. Rotor blades on helicopters can significantly contribute to energy absorption if the aircraft impacts into

trees. The rotating blades incrementally chop the trees until rotor motion stops, or until blade separation occurs. In

either case, an appreciable quantity of energy can be absorbed,

providing that the attachment strength to the basic fuselage is

adequate to prevent the blades from being detached, or the rotor mast and transmission system being torn out of the aircraft.

Contrary to the case of the fuselage structure, where indiscriminate use of composites produces poor energy absorbers,

composite helicopter rotor blades offer a desirable degree of

energy absorption. For example, when a composite blade impacts

and severs tree limbs, damage occurs to cause progressive delamination between the axial fibers due to adhesive failure.

This process, called brooming, does not destroy the longitudinal load-carrying elements in the bladel these elements can

continue to flail the tree and transfer energy from the aircraft to the tree.

An additional property of composites that can be used to advantage is the high modulus of elasticity achievable by using

certain fibers such as boron. A stiffer blade that reduces

deflections and reduces the probability of a blade strike on

the fuselage crown can thereby be designed.

When rotor blades with metal spars are subjected to tree im-

"pacts, successful energy absorption takes place until a tree

element too strong for the blade to sever is impacted. Then

the blade will fail by plastic bending or by completely detaching from the rotor. During these failure sequences, fuselage impact may occur as a result of blade motions or impact by detached blade pieces. Such occurrences are unavoidable, and

overhead fuselage structure must be designed to preclude pene- tration.

Blade tips can be made frangible to crush upon impact. Deformation of the tip should be as controlled as possible. For example, if a frangible tip with tip weight fractures and leaves

the blade, the resultant main rotor centrifugal unbalance may

147
cause severe pylon damage. Therefore, blade tips should retain 7,

their weights during the blade strike. A replaceable frangible

blade tip concept (from Reference 32) is shown in Figure 57.

Tough tip spars fold on impact

14 in.- 2 in.

6 in.

21 in.

II •.

I I/

SCG of undeflected tip CG of

deflected tip /

Crushable tip stays

on rotor after damage

to prevent large rotor

imbalance. Tip is

replaceable.

Figure 57. Frangible main rotor blade

tip. (From Reference 32)

32. Gupta, B. P., HELICOPTER OBSTACLE STRIKE ANALYSIS, Bell

Helicopter Textron; USARTL Technical Report 78-46,

Applied Technology Laboratory, U.S. Army Research and

Technology Laboratories (AVRADCOM), Fort Eustis, Virgi.n1,

April 1979, AD A069877.

148
"For future main rotor blades, it may be possible to design a

blade that can accept localized destruction during a blade

strike such that harmful blade strike loads are not transmitted

to the mast. The blade should progressively fail inward starting at the blade tip. Filament-reinforced composite materials

could accept the fai lure of the resin material, yet retain some

of the load-carrying capability of the fibers (Reference 31).

Some means for blade balance weight retention should be ineluded.

Analysis of strength and energy-absorption characteristics of

typical wing structures for various modes of failure indicates

that the energy absorbed by wing deformation and failure does

not significantly reduce the total longitudinal kinetic energy.

Wing failures typically account for only 5 percent, or less,

of the total kinetic energy which may be present in a crash of

moderate velocity (Reference 33). The use of the wing structure as an effective deformable energy absorber, then, does not

appear to be promising. Increasing this energy-absorption cap- ability to a significant level would involve adding material to

structural members which can already adequately withstand normal loading. This approach would be doubly inefficient in that

nonstructural weight would have to be added to improve energy

absorption, and, in all likelihood, most of the added material

would not be used effectively in any particular crash.

On the other hand, it is reasonable to design the wings so that

the wing sections outboard of the fuselage break free from the

fuselage structure under high longitudinal impact forces. This

could account for a considerable reduction in mass, especially if the wings contain fuel. As discussed earlier in the fuselage design section, this reduvtiow in mass can effectively reduce the requirements for cabin structure energy-absorption capability.

Wing impacts usually occur in one of two ways. Either a tree,

pole, or similar object is hit, producing highly concentrated

loading, or the wing strikes a barrier, such as an earthen

mound or a dike, which produces loading that is more evenly

distributed along the wing's leading edge. Crushing and shear

strength, for typical wings, will allow trees or poles to cut

into the wing as the aircraft moves forward, until the wing is

cut off or until the pole breaks. The fore-and-aft loads under

these conditions are low in terms of their effect on fuselage

accelerations. In fact, even thc more evenly distributed loads

can produce fuselage accelerations of perhaps only 5 G.

33. Greer, D. L., CRASHWORTHY DESIGN PRINCIPLES, Convair, Division of General Dynamics Corporation, San Diego, California, SepLember 1964.

149

r II.
Wing removal also can provide the possibility of removing flammable fluids to areas farther away from the fuselage and occupants. It is understood that other design considerations may make these goals difficult to attain; however, the possible benefits make design for wing loss in severe accidents definitely worthy of consideration.

Empennage structure or, for that matter, any structure aft of

the occupiable cabin, provides no beneficial effects during a

longitudinal crash. Instead, the mass tends to increase the

compression loads which must be supported by cockpit/cabin

structure. Therefore, if empennage structures can be designed

to collapse during a longitudinal impact, the requirements for

cookpit/cabin structural strength and energy-absorption capability will be reduced.

6.6.5.6 Design of Engine Mounts and Structural Support of

Overhead Masses: Engine mounts should be designed to keep the

engines attached to the basic structural members supporting S~the mounts (nose section, wing, aft fuselage section, etc.)

throughout a survivable crash, oven though considerable distortion of the engine mounts and/or support structure may occur,

This will reduce the fire hazard and the localized damage to

other structure which occurs when engines break free and are

traversai by the aircraft.

The strength of engine mounts and fittings, including both

those integral to the engine and the interfacing mounts on the

supporting structure, should be such that failure or separation

of the major structural members supporting the mounts occurs

before engine mounts or fittings fail under any anticipated

crash conditions. Structural support of massive components located overhead, such as the transmission and rotor mast on helicopturs, should be designed to withstand the following loads:

lateral, 18 G; longitudinal, 20 G; and vertical, 20 G. These

strengths are necessary to assure that overhead comporLents do

not penetrate the occupants' protective shell.

Engine mounts and supporting structures, including firewall

bulkheads, should be designed to minimize earth scooping, as

discussed in Section 6.6.5.2.

6.6.5.7 Mmergency Exit Structure: The structural framing

around all emergency exits should be rigid enough to prevent

deformation to such a degree that emergency exits are inoper- able under crash conditions. In addition, the structure should

be designed to withstand at least a 5-G load based on the maximum gross weight of the aircraft when it is inverted. The

structure should also be able to withstand a transverse load

of 5 G applied between the floor and ceiling of the aircraft.

150
Locations of emergency exits should be considered in the overall design of aircraft. These exits must be placed in locations which are favorable for rapid egress. If components near

exit locations are likely to be damaged to the extent that the

exit will be blocked# then this possibility must be considered

and allowances must bp made to compensate for it. An example

of this situation is the location of an emergency exit near a

"landing gear attachment in a helicopter. Upon vertical impact,

the landing gear could be driven upward into the floor structure, causing severe distortion of the floor structure, thus

impeding escape.

6.6.5.8 Fuel Tank Installation: Fuel systems are designed to

store and transmit flammable fluids to their respective energy

generators. Hazardous conditions can occur if fuel is spilled

and ignited. MIL-T-27422 (Reference 31) and Volume V define

the requirements for the design and .. astallation of crashresistant aircraft fuel systems.

Location of a flammable fluid-carrying tank in an aircraft is

an important factor in minimizing postcrash fire hazard. The

location must be considered with respect to occupants, ignition

sources, and probable impact areas. Creater distance between

occupants and the fuel supply tends to increase potential escape time if there is a fire because it reduces the likelihood

of fuel entering the occupied area. The tank also should be

kept away from probable ignition sources. While this is not

always feasible, tanks should not be installed in or over the

engine compartment, the battery, or other primary ignition

sources. An extremely important consideration is the location

of tanks with respect to probable impact damage. The effects

of penetration by aircraft hardware, such ac landing gear, or

ground obstacles require that the tank be located where there

is minimal probability of penetration with the subsequent fire

risk. Fuel containment is of primary importance, even if proximity of the tank to occupiable areas and ignition sources must

be accepted as the only alternative.

The installation of fuel tanks within the primary structure

matrix must be carefully determined in conjunction with the

following design considerations.

34. Military &pecificati n, MIL-T-27422, TANK, FUEL, CRASH

RESISTANT, AIRCRAFT, Department of Defense, Washington,

D. C., 13 April 1971.

151
9 Fuel containment, with or without crashworthy fuel

cells, for high G impacts.

* Minimization of structurally generated jagged edges, etc., which may cause fuel cell penetration.

9 Maximization of distances from potential ignition

sources.

* Sufficient distance from belly of aircraft to minimize potential for tank wall penetration by rocks,

tree stumps, posts, etc., located on impacted surface.

* Sufficient distance from landing gear to preclude tank penetration due to gear failure.

9 Large mass items located away from fuel tanks wherever possible to prevent excessive compression of

tank volume and/or penetration.

* Occupied areas segregated from and located as far

away as possible from tanks.

* Fuel cell supports and system hardware attachments

designed to allow major structural deformation without tearing of cells and/or attachment points which

result in fuel spillage.

* Fuel tanks regular in shape, cylindrical or rectangu- lar, to minimize the effects of internal pressure due

to structural compression and to reduce the propen- sity to snag on structural elements.

e Structural deformation not exceeding the compression capability of the fuel cell under survivable crash

impact conditions.

in summary, the fuel tank installation should be, wherever possible, mounted well above the bottom of the aircraft and situated as far as possible from large mass items, landing gears, occupied areas, and potential ignition sources.

Adequate support of fuel cells is required to minimize the

probability of penetration. Figure 58 shows a typical interior

support structure for a fuel cell. Closely spaced frame members are interconnected by metallic strips to provide a uniform support matrix tc minimize cell comprossion and distortion under severe impact conditions.

152
Figure 58. Typical interior support

structure for fuel cell.

Onn method of improving the resistance to penetration by frac-

,A tured metal is the use of a plastic foam as a tank cavity liner.

A foam liner would accomplish two objectives: (1) it would assist in preventing fuel cell penetration or cutting by sheet

metal components, and (2) it could fill a void space between

internal rib stiffeners on some tank cavities, thus removing

the possibility of fuel vapor collection in the voids.

Attention to these requirements, and use of the latest crashworthy fuel system techniques, will result in a satisfactory

installation.

6.6.5.9 Ancillary Equipment Retention: Objects that are unrestrainer or dislodged from their attachment during a crash

behave as missiles during the crash sequence. They can cause

injuries by striking the occupants directly or, in the case of

larger masses, by striking seats or other retention system components, causing failures of these items. Therefore, all ancillary equipment carried within the occupiable portion of the

aircraft must be restrained so that it cannot become dislodged

during a survivable crash.

"153

S. ........
All ancillary equipment frequently carried aboard the aircraft

must be provided with integrated restraint devices or anchors

to the aircraft structure. Stowage space for nonrestrained

items that are not regularly carried aboard the aircraft should

be provided in all aircraft. This space must be located where

items stored in it cannot become hazards to personnel in a survivable crash.

To provide adequate protection during severe but survivable

crashes, ancillary equipment restraint must equal or exceed

the load capability of the aircraft seats. As stated in Section 5.3.2.1, restraint devices and supporting structure for

ancillary equipment must be designed to restrain the equipment

during exposure to the following static loads:

Downward: 50 G

Upwards 10 G

Forward: 35 G

Aftward: 15 G

Sidewardt 25 G

It is a relatively simple task to restrain small items of

equipment to withstand the static loads specified above. For

larger items, however, significant weight penalties may be incurred or the available supporting structure may not be capable of withstanding the anticipated loads. For these reasons,

"load-limiting devices are recommended for the restraint of

heavier equipment. However, load-limiter stroking must not

allow any equipment to enter an occupant strike envelope. (Occupant strike envelopes are discussed in Volume II.)

6.6.5.10 Seat Installations: Three major catagories of seats

are used on U. S. Army helicopters: crew (pilot/copilot), gunner, and troop seats. Crashworthiness specifications in MILSTD-1290(AV) require that each seat occupant be provided with

a survivable environment when the aircraft is subjected to a

95th-percentile potentially survivable impact (Reference 1).

To meet this requirement, energy must be absorbed and living

space maintained as the total aircraft system is decelerated.

"Ideally, each occupant must be brought to a state of rest without incurring debilitating injury that might preclude timely

egress after a crash impact. Various energy-absorbing seat

concepts are illustrated in Figure 59.

Seats are installed to act as part of the decelerating system

and complement the landing gear and deforming fuselage structure. Whichever seat type is used, the total system must be

considered when the seat-to-structure interface is evaluated.

154

4I
He~d~~t rameHeadrest

SouLder, inertia reel shoulder ¶ hharness -Blucket attach fitting harness Inartia reel linear bearing assembly I Vertical energy ' etcleeg

A Amrdabsorber Armored Vesrtibleer g ordbucket Tilt back side panel ' bore

Lap belt control

handonle r-ulkhead B 

Cuho C~ushion

Fore-and- Latera haftcndrle enocn Armored buok.Lt - Guide tube

af cntrl* straps

loc Fore-and-aft ins

(a) Crewseat, armored bucket, (b) Crewseet, armored bucket, 1

floor-mounted bulkhead-mounted

YShoulder

Ymel harness an neria real

TubularFabric backr

ieia reel 

LaSbl

Lap belt~. a bnryabebn les

n" rongiturena

Lap er ltasrbn a ml Iryasibn 

Fiegu bsreo 59Lyia rswrtysa ofgrtons.udna

low ~~~~~~energy and seat absorptobneesr. Atraielyri

thela fingmea aosrcind strutur cobnt onwere etreel rgi

alayd theseleathatopvie the majoraf porwthion ofe theiva protek tion to the occupant, a much longer stroking distance would be

needed to absorb enough energy to fully decelerate the occupant

within tolerable load limits.

155
When assessing the seat installation, it is necessary to know

the design impact conditions, the total energy content, and the

predicted energy-absorption capability of the landing gear and

primary structure. Whether the crashworthy seats are floor

mounted or suspended from the overhead structure can be incorporated into the computations regarding energy absorber load

limitations, stroking distance, etc.

The question of occupant restraint must also be considered.

Restraint systems are designed to allow each occupant a sufficient degree of movement to complete assigned tasks in relative comfort and to provide adequate restraint in a crash environment. Restraint system loads must be transmitted to the

primary structure via the seat structure to ensure correct restraint if the seat strokes.

Where crashworthy stroking seats are used, an additional factor must be considered when defining the geometry of the structure. This factor is the increased strike envelope derived

from the seat moving from its position for normal flight opera- tions to its final postimpact location. Modification of the

strike envelope is dependent upon whether the seat has the

capability for uniaxial, biaxial, or triaxial stroking.

Because energy-absorbing stroke in the vertical direction is

mandatory, items that surround the seat# such as consoles and

collective controls, should be placed to allow for lateral and

longitudinal deflection of the seat without blocking the vertical stroke. The required clearance will vary depending on the

specific seat design and whether the lateral and longitudinal

motion results from intentional energy-absorbing stroke or from

elastic or plastic deformations of the structure. Three inches

of dynamic lateral deflection is common even in an elastic

structure designed for vertical stroke alone. Of course, this

amount of deflection is not likely to be retained throughout

the entire vertical stroke; however, as much clearance as possible should be provided.

A minimum of 12 in. of vertical seat stroke is recommended

(from the lowest vertical adjustment position), as discussed in

Section 8.3.1 of Volume IV. Because of the desired positioning of the seat in the aircraft, 12 in. may be difficult to provide

between the seat bucket and the floor. In these cases a hole,

or well, in the aircraft floor should be provided to allow additional stroking distance for the seat. At least 2 in. of

clearance should be maintained between the outer edges of the

bucket and the innermost hardware extension on the sides or

front of the well including the tracks.

156
i• •! • It •''i!'• !•

• •!i The underfloor or bulkhead structure, depending on where the :•

seat mounts, and the aircraft attachment hardware should be ,•I

•" ,, designed to withstand the loads and moments generated by the 'if'i• 'i•! seat. The loads and moments should be those applied by the

specific seat for all design loading conditions and for all pc- !

i •: sitions in the aircraft. In other words, the design of the

'• aircraft structure and the seat should proceed concurrently,

•, with reaction loads and moments calculated for the specific •i• seat used in designing the aircraft structure.

i!•I: 6.6.5. i i Cargo Retention

I• 6.6.5.11.1 General: Cargo carried within transport vehicles

•c" •:i," should be restrained to whatever degree is necessary to ensure

•' survival of the crew and passengers in a potentially survlv-

•: able crash. In order to determine the types of cargo restrain-

• ; ins devices needed for U. S. Army Aircraft, At is important to

!i : consider the following related factors:

• ype of aircraft. Si, 

Tiedown backup structure and likely crash modes S• .

P • Type of cargo being carried.

iv • Crew and passenger locations relative to cargo.

i • Aircraft and cargo tiedown provisions.

• Cargo/personnel clearance envelopes.

• Type of restraint devices available.

The type of aircraft and its predominant crash mode should dictate the selection of appropriate restraint criteria; e.g.,

forward, lateral, aft, and up. Helicopter and flxed-wing aircraft crash modes differ, and the simplest, most effective tiedown arrangement should be sought for the specific needs of

each aircraft.

i, The types of cargo to be flown and the restraint criteria will

i determine the available tiedown clearance. An awareness of

)" aircraft ,truatural response to impact will help to identify I ' realistic clearance envelopes for specific personnel locations.

i, Consideration should also be given to restraint integrity in

the event of local structural discontinuities and/or secondary . or tertiary impacts.

I, Aircraft and cargo tiedowns and restraint devices in use have ii been to various criteria, of which obsolete. Sdesigned some are

I'•% 15 7

i

• 1•iI• . .... . ..... .... "'(' iii,. ,f:lll.,1,, ,tl,,•ll.*h*lI ....... i,1,i,,,I • ,,,1t1/1•1 i, illll, lr ........ •llf[ i;ii ill . ..... ,11,,. •,,I ,Ij , . ... ,*• h,], ,;,;hl,*l,,,llt.'ll, , • .,I •h • •', ,l', '11;; •:
As passive devices, they should not be subject to timedependent or environmental deterioration. In many cases, the

tiedowna will prevent maximizing restraint effectiveness or result in inadvertent failure. Restraint devices may possess the

required structural strength but may be unsuitable if not optimized for crash pulse loading, i.e., if they are too soft or

too stiff. The devices may not be equipped to maintain the

cargo captive at all times during aircraft movement (e.g.,

rollover) occurring after the main impact has been survived.

New, broad-based mi'litary requirements call for multimodal bulk

cargo handling. Palletized cargo will become more prevalent

ir use. A single cargo package moved in this manner between

source and deatination offers speed and economy.

Some U. S. Army aircraft will be dedicated to cargo handling,

if only on a special mission basis, to extend retail delivery

to the battle zones. For this reason, the restraints must be

adaptable to and compatible with the USAF 463L System for protection of the aircraft crew.

p Cargo netted to 463L pallets (References 35 and 36) are restrainable to USAF flight load criteria (Reference 37). The

USAF cs.ý,gO tiedown criterion requires 3 G forward to minimize

the portion of mission effort devoted to cargo restraint.

The designer should be aware, however, of the basis for the

differing USAF philosophies and practices. The USAF and commercial operators of large cargo aircraft have accumulated over

40.4 million hours of flight time between January 1960 and July

1976. Analyses of major and minor accidents over this time interval have shown that tho risk of passenger fatality above the

3-G forward restraint criteria is statistically rare for USAF

operations (Reference 38).

35. Military Specification, MIL-P-27443, PALLET, CARGO, AIRCRAFT, TYPE HCU-6/E, HCU-12/E, AND HCU-10/C, Department

of Defense, Washington, D. C., 24 February 1967.

36. Military Specification, MIL-N-27444, NET, CARGO TIEDOWN,

PALLETS, HCU-7/E, HCU-15C, HCU-ll/C, AND HCU-16/C, Department of Defense, Washington, D. C., 8 May 1969.

37. Military Specification, MIL-A-8421, AIR TRANSPORTABILITY

REQUIREMENTS, GENERAL SPECIFICATION FOR, Department of Defense, Washington, D. C., 25 October 1974.

38. CARGO AIRCRAFT AND SPACECRAFT FORWARD RESTRAINT CRITERIA,

USAF Technical Report 76-30, Aeronautical Systems Division,

Wright-Patterson Air Force Base, Ohio, December 1977.

158 ,
Army warehouse pallets provide no restraintl cargo may be loose 'W or only lightly secured to the pallet by banding (a I-G restraint).

Both pallet systems must be given additional restraint to

meet U. S. Army crashworthiness criteria, albeit minimizing

restraint time is also a necessity (Reference 39).

6.6.5.11.2 Cargo Categories: In the selection of restraint

devices and procedures, it is important to define the probable

types of cargo that will be carried. U. S. Army cargo has traditionally been divided into two categories, based primarily on

size, as listed in 3Table 4 (Section 5.4.1). Small cargo (up to

approximately 3 ft in size) when routinely handled is banded

to wooden Army pallets to expedite handling by fork lifts.

These loads can best be restrained by nets attached to aircraft

floor fittings. Nets specifically designed for this purpose do

not exist. They should be of a lightweight, low-elongation

material that can be conveniently sized and compactly stowed.

Presently, netting of such loads means using heavy, cumbersome

equipment normally used for handling external cargo. The exceptions to this are the adjustable size nets used with 463L

pallets, which are sufficiently large to accommodate a large

number of small packages stacked within a volume as large as

108-in. length x 88-in, width x 96-in. height. These nets are

restrained to structural tiedown rings integral with the pallets. The deficiency of the present 463L net is its elasticity, which is incompatible with U. S. Army crash survival criteria.

Large or bulky cargo can best be restrained individually by

cables, ropes, or chains that are attached to the floor or

sidewalls of the aircraft.

It is assumed that both categories of cargo (including, for example, packse rifles, rations, and wheeled or tracked vehicles)

will sometimes be carried along with troops. If troops are to

be transported along with both types of cargo, crash protection

should be provided for troops as well as crewmembers. Where

aircraft are dedicated to cargo shuttle missions, cargo crash

restraints should be provided primarily for flight crew protection.

6.6.5.11.3 Strength: It is recommended in Volume IV that the

design strength of troop seats be sufficient to retain personnel in 95 percent of survivable crashes. Should cargo carried

39. Shefrin, J., et al., INTEGRAL HELICOPTER CARGO RESTRAINT

SYSTEM, The Boeing Vertol Company; USAAVLABS Technical

Report 69-68, U. S. Army Aviation Materiel Laboratories,

Fort Eustis, Virginia, October 1969, AD 864899.

159

S. ...... ... . ...... .. ...: ,,: . . • ., , .. ... . .•, .. ... .. . .. ,.,,........p! ,
alongside, forward, or aft of troops also be restrained to the

same level? Perhaps it should bel however, the weight of a restraint system capable of keeping a maximum cargo load in place

at such load levels would be high. Several factors other than

the expected floor pulses, indicated in Figure 60, affect the

retention strength level required for cargo. In particular

these include the type of cargo, where it is located, and hQw

it is restrained in the aircraft.

Percentile-1- 95th 90th 85th 50th

Cockpit area

30 G 20G 16G G

50 A 43 AA 3 At 39 2

AV e A f. src ftosec ftosec pulses

cabin area

4 G 6 13 G 4 iG

_ 0_ )28\ /4e AV ftO. ecAftsec ftsac Af/e

Fiue60. Aircraft floor longitudinal pulses for

rotary- and light fixed-wing aircraft.

Most small cargo (see Table 4 of Section 5.4.1) probably will

be restrained by a net stretched over several items. It is

also probable that the same net will be used to secure any number of boxes. Therefore, statistically, a net designed for a

given load will be loaded to a lower value in most accidents.

Furthermore, failure of a cargo's net restraint is not likely

to be as large a threat to human survival as the failure of an

occupant's seat restraint would be. For example, failure of a

single net midway in a row of nets would not become lethal until other nets also failed. Therefore, for practical reasonsg

i.e., the low actual probability of injuries caused by displaced cargo and the structural problems involved in restraining cargo to levels equal with personnel restraint, it is concluded that the cargo restraint level can reasonably be less

than the 95th-percentile level.

160

• ' ,6.1! iJ V, iI "'J, 4",, ...... , 
Of course, in the case of large cargo carried behind troops,

• the above reasoning does not apply. However, a detailed cargo restraint study by the Boeing Vertol Company (Reference 39)

indicates that most items of large cargo are carried outside

the helicopter by sling.

On this basis, it is recommended that cargo nets and lines be

designed to sustain slightly lower loads, e.g., the 90thpercentile crash pulse rather than the 95th-percentile standard required for human restraint systems. Floor longitudinal acceleration pulses are shown for the 50th-, 85th-, 90th-, and

95th-percentile cases in Figure 60. The crash severity is reduced significantly at the 90th-percentile value over that for

the 95th percentile. Selection of the 90th-percentile value

* reduces the peak deceleration from 24 G to 16 G and the velocity change from 50 ft/sec to 43 ft/sec.

6.6.5.11.5 Energy Absorptioni As in crewseat energy absorp- tion, the use of load limiters for cargo restraint is recommended. In the cargo application, the load limiter controls

the cargo restraint load level and physical movement into space not occupied by personnel. For the aircraft cabin pulse shown

at the 90th-percentile level in Figure 60, calculations have

been made as developed for seat and occupant displacement (Volume IV) to show the required restraint G level and the corresponding stroke for energy-absorbing or load-limiting restraint

systems. The required G-level-versus-displacement relationship, shown in Figure 61, results from minimizing the restraint

systems' structural strength (and hence, their weight) and establishing an acceptable load displacement. These results neglect any restraint system elasticity or longitudinal friction.

NASA and AvSHR crash data indicate that a vertical floor acceleration is always present to some degree simultaneously with

the longitudinal acceleration. The presence of the vertical

(normal) force allows the development of some (aiding) longi- tudinal friction forcel however, the friction force may be offset by the dynamic overshoot due to elasticity in the system,

or by cargo/floor interface rollers.

Recommended performance goals for such load-limiting restraints

are shown in the forward longitudinal load-versus-displacement

curve, Figure 61, and the lateral load-versus-displacement

curve, Figure 62.

To meet the necessary energy-absorption requirements, restraint

displacement must rise rapidly to the left of and above t."

lower base curve, terminating its displacement above the mi.,

mum acceptable load (ultimate strungth) curve.

161

.. . .
Permissible

controlled

X displacement

Aircrat dispacemen

floor devices

N~et restraint Line restraint

20 -1 1

Practical cargo displacement limit--*

18. (depending on aircraft)

10 Sample EA

404 Restraint curves

4J

2 ' ~ -. Una ceptable performance

0 0 1 5 10 15 20 25 30

controlled forward uargo displacement (X), in.

Figure 61. Load-displacement requirements for energy-absorbing

cargo restraint systems (forward loading of rotarywing and fixed-wing aircraft).

162

'Iq
CDFailure load

--.. Unacceptable performance - Acceptable performance Minimum Practical cargo

S8 acceptable displacement 8 load curve -- limit (depending'

on aircraft)-

S6

A able failure area

1'2

0

0 5 10 15 20

Controlled lateral cargo displacement, in.

Figure 62. Cargo lateral load-displacement requirements.

Similar load-limiter ca.]lrulations have been made for the lateral direction, and these values are shown in Figure 62. The

lateral strength requirements are based on a 90th-percentile

velocity change of 21 ft/sec and a 10-G triangular pulse.

While the troop/cargo mix situation has received the most atr " tention in this portion of the design guide, special cargo missions exclusively involving the aircraft crew may possibly require other design solutions; i.e., emphasis on forward longit4, tudinal nature, restraint involving and both crew crew isolation. and passengers# An extreme is the case Navy of C2-Athis

COD fixed-wing aircraft where extremely tight quarters, requirements for high utilization of the volume devoted to cargo, and a forward longitudinal restraint of 20 G prevail (Reference 40). These conditions have dictated the use of a structural locker for multidirectional restraint.

Recent research was conducted in energy management restraint

(low elasticity/load limited) systems applied to the Navy aircraft. The test results, Figure 63 (Reference 41), provide an

40. Hate, R. L., NAVY CARGO RESTRAINT CRITICAL REVIEW, NADC

Technical Report 74082-30, Naval Air Development Center,

Warminster, Pennsylvania, May 1974.

41. Shefrin, HARDWARE J., DEMONSTRATION OF ADVANCED CARGO RESTRAINT FOR COD AIRCRAFT, The Boeing Vertol Company;

NADC Technical Report 77154-60, Naval Air Development Center, Warminster, Pennsylvania, December 1978.

163
co IV

W N 41 Vmw )1

o~to

to1 roe' -$4 0

4J4J

4J J 4 t

E- -1E4W 0udH

4"

00

t, HH "4 to '4

17 ' 0 

04

0)0

3⁄4.~. cy, H '-n

0d~

03

0 N3 44 

164

T;
illustration of the performance of crashworthy cargo restraint

systems.

Shown are the plotted results of tests at two levels of acceleration with two load-limited systems, each system having different elastic deformation characteristics.

S'Peae tests illustrate a comparative acceptability relative to

the knee of the base curve, which is used as a target for limiting system deflection. The most efficient systems (Tests 1

and 3) were achieved with members possessing the highest stiffness. Thus, Systems 1 and 3 maximized the probability of pro- tecting nearby personnel, and minimized the problems of stored

energy (rebound) and dynamic overshoot.

In the tests, three identical metallic load limiters were used.

The first and third tests employed an aramid, low-elongation restraint device (tension member), and the second test used a

polyester restraint (approximately double the strength of the

aramid capability to achieve equivalent strain). All config- urations may be considered acceptable where ample displacement

space is available. However, minimizing the elastic displacement in trade for plastic deformation of the load limiter pro- vides more energy absorption for a given cargo displacement. The minimum cargo displacement could be achieved with the load

limiter alone (without the tension member). The displacement

curve for a load limiter alone would have fallen to the left

of the Test 1 curve, due to a higher stiffness preceding stroking. If extra load-limiter stroke capability is available, the system will provide an increased measure of performance at

percentile pulse levels above the 90th without pull-out or bottoming.

The selection of design load levels for load limiters is important to ensure stroking and to protect the aircraft floor

tiedowns from ultimate load failure, as evidenced by the results of the U. S. Army/NASA cargo experiment crash test of a

CH-47C helicopter (Reference 42, Test T-40, August 1976). Floor attachments failed as a result of two factors:

42. Burrows, L., Lane, R., and McElhenney, J., et al., CH-47

CRASH TEST (T-40) STRUCTURAL, CARGO RESTRAINT, AND AIR

CREW INFLATABLE RESTRAINT EXPERIMENTS, USARTL Technical

Report 78-22, Applied Technology Laboratory, U.S. Army Research and Technology Laboratories (AVRADCOM), Fort

Eustis, Virginia, April 1978, AD A055804.

165

rENE 5'

. . . . . . . . . . . . . .
9 Limiter design load level was less than 2 percent below fitting ultimate strength.

* Inadequate consideration was given to the variability in both the limiter loads and the tiedown fitting

strengths.

A load-limiter design that is too conservative, however, is

counterproductive to minimizing cargo restraints. Longer

stroke requirements are also a result of this approach. This

may mean that higher fitting strengths should be considered

for new aircraft designs And that old aircraft fittings should

be upgraded. The payoff for this approach is optimization of

the restraint problem and achievement of maximum productivity

in restraining internal cargo.

The design strength values for cargo along all axes of loading

are summarized in Table 8 on the basis of the 90th-percentile

pulse. The aftward and upward loads in Table 8 are based on

expected rebound loads of approximately 30 percent of the peak

input values. Hardware designs should not allow disengagement

of the restraint during a crash. Restraints usually slacken

in the time period between initial impact loading and rebound,

and a restraint hook without a keeper can become disengaged

during this short interval.

TABLE 8. CARGO RESTRAINT LOADS AND

DISPLACEMENT REQUIREMENTS

Load direction

Item (with respect Restraint Controlled

no. to floor) load displacement

1 Forward See Figure 61 See Figure 61

2 Aftward 5 G No requirement

3 Lateral See Figure 62 See Figure 62

4 Downward 16 G No requirement

5 Upward 5 G No requirement

6 Forward See Figure 61 See Figure 61

and Combined

Lateral 4 G No requirement

166 5iI

Ii•
The 16-G downward loading of Table 8 is recommended to ensure Z that the cargo does not crush the floor, and perhaps destroy

the longitudinal restraint for personnel or cargo, during the

initially high vertical impact loading. This requirement

should not be difficult to meet for the net-restrained bulk

cargo because this cargo will probably be stacked on pallets

that will assist in load distribution. Shoring, however, will

probably be required to prevent floor penetration by wheeled

vehicles. Combined forward and lateral loads are considered

realisticl Item 6 in Table 8 is included to ensure that the

system will not fail under this type of loading.

Displacement requirements are not suggested for the aftward,

downward, and upward loads in Table 8 because these loading

directions are not considered to be as potentially hazardous

as the loads in the forward and lateral directions. Upward

restraint is usually not a problem since all restraints used

act in unison for this direction.

For personnel located aftward of the cargo, a buffer spacing

to allow for restraint system elasticity under the 5-G loading

should be considered. Restraints are snugged-up, but very

little stretch is actually removed. Normal preload is in the

range of 150 to 350 lb. Therefore, cargo movement in response

to an aftward 5-G rebound load can be expected. The magnitude

of the total cargo motion will depend upon the type and characteristics of the restraints used and the elasticity and/or deformation of the backup structure.

6.6.5.11.5 Net Design: Nets should be designed to restrain

the small (bulk) cargo up to the specific loads in Table 8 with

a minimum of elongation to reduce overshoot. New webbing fibers such as polyesters and aramids should be used. Their relative elongation and weight advantages over nylon are shown in

Table 9 (Reference 41).

TABLE 9. CARGO NET MATERIALS - RELATIVE CHARACTERISTICS

E!Eiongation Breaking

@ 5000 lb Weight strength

Webbing fibers (percent) (lb) (lb)

Nylon 7.5 3.0 1.1

Polyester (Dacron) 2.0 5.9 1.8

Aramid (Kevlar) 1.0 1.0 1.0

167

.... .......... ..
Present disadvantages affecting the adoption of aramid fibers

for tension member application are cost, low abrasion resistance, and fabrication technique development.

Nets should be designed for the most critical loading condiVtions, i.e., a high stack with small boxes. Such a loading

would tend to cause a larger diaphragm action on the net as

the individual boxes slide over each other. More concentrated L loads may be encountered in restraining palletized loads on

floor rollers.

6.6.5.11.6 Cable, Rope, Strap, or Chain: Large cargo of the

type outlined in Table 4 (Section 5.4.1) can be secured most

easily by cables, straps, ropes, or chains (lines). The simultaneous use of these devices must be carefully considered, howeebecause their differing elastic moduli can result in premauefailure of the stiffer devices before the others are

ulyloaded (Reference 43).

Incases where the use of load-limiting devices is not possilretaining lines for the longitudinal direction should

be arranged to simultaneously attain the maximum load in all

lines. A combination of fabric ropes and steel cables should

not be used on the same piece of cargo because of the difference in their elastic stiffnesses. Also, symmetrical restraint

configurations should be used to avoid overloading individual

tiedowne and, eventually, failing the system.

With load-limiting devices, it is desirable that metal cables,

or other low-elongation tension members stressed below the

yield point, be used.' A low elongation of the lines used with

load-limiting devices is essential to ensure that the energy

is absorbed in the most efficient manner. If stretchable metal

cables are used as integral load-limiting devices, they will

be stressed above the yield point, and the low-elongation line

will not be app] icable. Metal cables possess handling qualities that discourage their use. Nonmetallic fiber materials,

which are lighter in weight as well as more flexible and stowable, should be used wherever possible.

43. Russo, A., Jr., CARGO RESTRAINT SYSTEM PHASE I REPORT,

All American Engineering Company, Wilmington, Delaware,

January 1964.

168

S. ...... .
7. ANALYTICAL METHODS

7.1 INTRODUCTION

Crashworthiness technology can be extended beyond simple concepts by simulation of the crash phenomena. As shown in MILSTD-1290(AV), there are several sets of crash conditions that

must be investigated in support of the design process. Additional sets of impact conditions may appear critical for certain specific systems, and these require examination as well.

These conditions may be simulated by analytical models, scale

models, and full-scale tests.

The objective of simulation is to provide a rational basis for

the sequence of events and the modes of failure of vital elements of stri.ucture during the crash. Complex interactions of

crash, inertial, and structural forces which contribute to the

structural distortion and the acceleration environment experienced in a crash can be observed. Dissipation of the potential

and kinetic energy of the aircraft can be studied for conditions that exist in the crash sequence. Structural distortion

with subsequent ruptures, volumetric reductions, and penetra- tion of occupied spaces can be assessed, and estimates of the

acceleration levels on critical components and occupants obtained.

In terms of fidelity, the dynamic testing of full-scale structures most closely approximates actual crash conditions, especially if velocity components and impacted surface conditions

can be realistically represented. However, during the early

design stages of a new aircraft, full-scale testing is untimely

and costly. In fact, the testing of full-scale airframes has

been confined to technology development, rather than design development and improvement.

In this chapter are presented discussions of kinematic and dynamic relationships as applied to the crash environment, analytical methods of crash simulation, and methods for obtaining

the properties required 2or such simulation.

7.2 DYNAMICS OF THE CRASH ENVIRONMENT

7.2.1 Kinematic Relationships

Relationships among the kinematic quantities of position, velocity, and acceleration form the basis for the study of dynamic phenomena. Although these relationships are generally

well known, use of their precise definitions facilitates an understanding of the crash event and its analysis. Therefore, a

discussion of these relationships as applied to the crash environment is presented below.

169
Consider first an aircraft impacting a rigid vertical barrier

as shown in Figure 64. The position of Point C fixed in the

aircraft relative to a point fixed on the ground is referred

to as X. A change in position is referred to as displacement S. In this case, assume that the displacement of C is

measured from its poeition at the time when the aircraft just

contacts the wall. If a change in displacement AS occurs in

a time interval At, then the average velocity of the aircraft

over that interval is

AS (19) At

Fixed reference

-X - Position (measured from fixed reference)

•-S Displacement (increment _ _ . in position during time t)

C/

Reference point on aircraft (such as

for mounting of accelerometer)

Figure 64. Definition of position and displacement

for crashing aircraft.

As the time increment is taken to be very small, Equation (19)

yields the instantaneous velocity

V lim AS dS (20) A -t--o0O At wt170 LVijz~

S. .• .. " " ,: • .. .. ,.,, •:•': ',' "! •: ... ":; • • j ' ,,;• ::• •/",,•:• ,:, ' ,,"•",' ,,, •,,,, ...... ... ,,.,,, ,• " "" "'.1.'
:( ~ Note that velocity has the units of length per unit time, i.e.,

"__ 'feet per second.

i If in the same time interval the aircraft undergoes a change

"in velocity Av, the average acceleration over the interval is ' ' ,igiven by

J - (21)

and the instantaneous acceleration by

lnm AV dv

am~~.m (22)

Note that acceleration has units of velocity per unit time,

i.e., feet per second per second.

'V

A graphical examination of the relationships among position,

velocity, and acceleration will aid in understanding their ap- plication in the crash environment. Referring to Figure 65,

assume that the aircraft of Figure 64 has a velocity v just

prior to contacting the barrier and that the velocit? of

Point C, fixed in the aircraft, varieN with time as shown in . . Figure 65(b).

First note that in Equation (22) defining acceleration that "a" is the height of the a-t curve and Av/At (for small At) is

the slope of the v-t curve in Figure 65. Thus Equation (22) indicates that the height of the a-t curve is equal in magni- tude and sign to the slope of the v-t curve, that is

a -dv (23)

"This is an invariant relationship and any data, whether experi- mentally or theoretically obtained, must meet this criterion

to be valid.

In a similar way Equation (20) indicates that the height of the

v-t curve is equal to the slope of the S-t curve.

171

,,Si

•.i'if'
UA Aea a- *At AVL

0a

44 TiMelbec Height of

a-t curve

S(Decel).- I

0

Q (a) Accel .era ition -time,4

slope- AV/at -a

venoitya V -Slope.

veloityal A

'At~ Slope of

v-t curve

_ ~Maximum slope of 'v curve -ama

rea i,

S)j' v. At - S I

0:' 1 eoct Height of

- v-t curve

rime, sec -(b) Veloct- time 'N" 1 av

At Slope, - AS/ft -v

I A ehce lpe 4 ofAS 

V A t v e h i c l ecS lve o

(c) Displacement -time Time, sec

Figure 65. Assumed relationship for illustration

of crash kinematics.

172
Rewriting Equation (23) and integrating, the change in velocity

over any time interval between t and t in

t

4v =vv v Jadt (24)

to

Therefore, the total change in velocity during a given interval

is equal to the area under the a-t curve in the interval. An

an example, consider the total oroashatched area between t w 0

and t t t in Figure 65(a). This area is equal to the initial

velocity 40, that is, the change in velocity between t a 0 and

t aU

Note that the acceleration in Figure 65 is negative (deceleration), thus producing a negative velocity change or reduction

in velocity from v to 0. A positive acceleration, on the

other hand, produco4 an increase in velocity.

The same condition exists between the velocity and displacement

curves, that is,

S~t

AS vdt ( (25)

to

or, in other words, the total distance travelled is equal to

the area under the v-t curve. Thus the maximum vehicle travel,

as in Figure 65(c), would be equal to the area shaded with dots

under the v-t curve of Figure 65(b).

The following important points may be noted in Figure 65:

9 The velocity is changing at its most rapid rate when

the acceleration (or deceleration) is maximum, at

time tI•

e The displacement reaches a maximum when the velocity becomes zero, time t 2 .

173
e The velocity need not necessarily be zero (time t 2)

when the acceleration is maximum (time t0). 2

* The area contained within the deceleration pulse

(from t to t ) is equal to the initial 'velocity

Dlus the rebou~d velocity.

0 The area under the deceleration curve between t and

t is equal to the rebound velocity.

7.2.2 Energy Absorption During Deceler4tion

According to Newton's second law of motion, the resultant force

(F) acting on a mass (m) produces an acceleration (a) according to

"---ndv

Applying Equation (20),

Sdv dS dv (

~-.•'"mv~ * (- 27)

Multiplying by the incremental displacement dS and integrating,

"% , •,

82 v2

f(1 2 1 2 (8 mv dv m "FdS mvv (28)

S v

which states that the work done' on a mass m by resultant

force moving through a change in displacement from S to S

changes the kinetic energy of the mass, which is defined ai

Kinetic Energy, T I mv2 (29)

174

I' S
Work done by force F during the change in displacement is de- , fined as

S2

Work - FdS (30)

Si

which is equivalent to the area under the curve of force versus

displacement between any two displacements, as illustrated in

Figure 66.

U Work f }2dS

0 S

4j (in iri.-lb)

i,,

1 2

Displacement (S), in.

Figure 66. Definition of work.

In application of Equation (28) to the crash environment, an

aircraft moving with initial velocity v can be slowed to a

reduced velocity v (which is zero when ¶he a~ircraft comes to

rest) by application of a force through a distance S

175

£

* *it***.**

. . . . . .. . . . . . . . . . . . . . . . . .*'.
S

JFdS I mVo 2 - * mv2 (31)

0

where the force acts in a direction opposing the velocity, as

shown in Figure 67. Referring to Figure 67, it might be said

that the kinetic energy of the aircraft is absorbed by crushing of its forward structure as its velocity is reduced.

S - displacement

Barrier

Crushable structure, whichYO

exerts forc. F to the left

(a) Idealization of impact into barrier

At displacement S1

the force F has

reduced the aircraft

velocity to vl given by

0 FdS -•-(o-Vi2 1 3 2 2

0 (vo

s1

Displacement, S

(b) Force versus displacement for barrier impact

Figure 67. Illustration of impact reducing aircraft kinetic energy.

176

....I
If a constant stopping force were applied, the velocity would

be reduced at constant deceleration, as shown in Figure 68(a)

and the force-displacement curve would be as shown in Figure 68(b).

, O Force -ma constant

Acceleration, a - constant

0

Time

(a) Constant acceleration versus time

0)

0

rL;. Energy change - FAS

Displacement, S

X (b) Constant force versus displacement

Figure 68. Illustration of deceleration by constant force.

177

;. .;
Often, for protection of occupants, it is desirable to limit

the decelerating force to some prescribed value. Given a maximum force, the most efficient energy-absorbing system would be

the one requiring the smallest displacement, AS, which would be

the constant-force system. Therefore, energy absorption by a

constant force is often referred to as "ideal" energy absorption. The devices shown in Figure 30 exert a nearly constant

force and thus act as nearly ideal energy absorbers. They are

desirable for use where concentrated loads are applied and the

transmitted force must be limited, such as in seats and landing gear.

Structural systems and certain materials, such as plastic foams

and honeycomb materials, approach the ideal force-displacement

curve to the extent shown in Figure 69.

Force Force

Loacding

A S

2unloading

IDisplacement, S

Material thickness, T ...

Figure 69. Force-displacement curve for honeycomb materlais.

Again the area under this curve represents energy absorbed.

The area can be divided into three important regions. If loading results only in reaching Point A in Figure 69, then unloading generally occurs dlong the elastic curve OA, and the energy indicated by Area "l" is restored to the system just as a

spring releases its energy when it is unloaded. Area "2" represents inelastic, or plastic, energy absorption and if loading

reached Point C the energy correspJnding to Areas "l" plus "2"

plus "3" is absorbed. However, as unloading occuirs, the energy

of Arua "J" is restozed in the form of rebound. Loading in the

178
region from B to C in the figure is often referred to as "bot-

!! toming," a condition in which the deforming structure or materi f • ial has become compacted and the load increases rapidly with

very little increased deformation.

The energy dissipated in locked-wheel skidding is of interest

in calculating, for example, the initial velocity when the

skidding distance is known. This is readily obtained from

Equation (31) if it is assumed that during skidding the sliding forces are constant and are the only significant forces

present. Thus, the work done by the sliding force is simply

the force times the distance and from Equation (31), for

Vfinal

FS mV2 (32)

or

V - (33)

The sliding force F ca4 be assumed to equal an average coefficient of friction times the normal force between the vehicle

and the sliding surface, or the weight of the vehicle if the

surface is horizontal, thus

F - IAN a uW (34)

and

V - -g•S .• (35)

where A - coefficient of kinetic (sliding) friction

W - weight of vehicle

S - stopping distance

• sec2

g - acceleration due to gravity - 32.2 ft/.ec

For the case vf 0 we can use the work-energy principle to obtain

z 179
Work - AKinetic Energy

-JAWS I W 2 vo2]

From which

Vf Vo - 2gpS (36)

and

V-4vf + 2gMS (37)

For sliding on a slope of angle a to the horizontal

V N-vf 2 + 2gS[A cos 6 1 sin 01 (38)

For sliding uphill the + sign applies; for sliding downhill,

the - sign.

Examplez The landing gear on an aircraft collapses and the

aircraft slides 1200 ft, at which point it was estimated to

have a residual speed of 50 kn. By dragging the wreckage with

a tank retriever through a load-measuring device it is found

that the required load to move the aircraft is 0, 7 times the

weight of the aircraft. To determine the speed when the gear collapsed

vf - 50 kn - 1.69 x 50 - 83 ft/sec

- 0.7

V 0 (83)2 + 2 (32.2)(0.7)(1200)

vo - 246 ft/sec

- 148 kn

180
7.2.3 Stopping Distance

For a complete aircraft system, or a subsystem such as seat and

occupant, average deceleration values (G) can be determined for

given velocity changes as demonstrated above, and from these

data an average stopping distance can be computed.

vo 2 vf 2 v 2gGS (39)

where vo - impact velocity

vf - final velocity (usually zero)

g - acceleration of gravity (32.2 ft/sec2 )

* deceleration in "G" units (average)

S - stopping distance.

Solving for S gives

V o02 - v f 2

s v= , (40)

2gG

This expression is useful for assessing the required stroking

distance for a seat when impact velocity and acceleration limitations are known; however, it must be noted that this rela-

,tion yields the total stroking distance, including deformations

of the impacted surface and deflection of the gear, fuselage,

and seat. Superposition of the seat, structure, and impacted

surface characteristics can be used to assess the net average

deceleration experienced by a neat occupant.

Figure 70 shows the variation of stopping distance as a function of G and velocity change (Cv) derived from the standard

Newtonian equations for assumed constant acceleration.

These characteristics assume 100 percent efficiency, but real

structure will react somewhat differently, requiring larger

stopping distances for given velocity changes and/or deceleration levels. The elapsed time values caln be used for assessing human tolerance potential when human injury criteria are

known as functions of G level and exposure time.

181

,•f"
Acceleration

equations

30- S v

2gG

t -2SI/

25- units: 0

fvt/sec i , /

g, 32.2 ft/sec2 /

20-/c,

.r0.1

~0/0

5- o

.02 .04 .06 .08 .10

Elapsed time (t), sec

Figure 70. Theoretical stopping distance as a function

of velocity change and average deceleration

level.

182
Figure 70 is useful for the initial assessment of motion enve- -•: lopes needed to provide occupants with a survivable decelerative environment and for the requirements of various structural

energy-absorption elements.

.t is interesting to compare several deceleration pulse shapes

(negative acceleration) to show the effect pulse snape has on

deceleration distance. Rectangular, triangular, and sinusoidal

• pulses are considered.

7.2.3.1 Rectangular Deceleration Pulse

i~~i Sl 1111 1f i f i t f , I# I /I I • S

Accel, a a (Deceleration)

• ~t ,

VC)

+

Velocity, V

Time

4- S lone - 0

44,

*0/

Position, S 0g

'AI dl I / I Ii I i 1i l

ime

From Equation (24) vo - at (41)

Then t - v /a

From Equation (25) S l 0o (42)

So that 1 0

S V. Vo

4 S~183
2 Or Vo -2aS (43)

And Vo2

S - a-E (44 )

The implications of Equations (43) and (44) are that for a

large v , a very large stopping distance would be required.

Or, for a small acceleration, a large stopping distance is required.

7.2.3.2 Triangular Deceleration Pulse

comparing two triangular pulses here, one with zero rise (onset) time and another with zero offset time,

Zero rise time Zero offset time

a

a

0 0

0

0I

2vo a t --2vo 2.1*a vot Sa at vt

1 2vt o

t2v0

184

0 o

II 2vo

,,:....,,I .. ... ,., .. . ... . :. ... v..: • • • • • ,i.• •, ,. .-. • •. 2r :.q V•,' ,.;• a.. •l , .• ,,•," i , 
20

S aa

Note that the decel.erationl pulses are equal in time (t *2v /a)

but that the zero-rise-time case requires twice the deceleAtion distance of the zero of fset time case.

7.2.3.3 symmetrical Trianaular Pulse

For a symmetrical (isoceles) triangular pulse,

a

V0

vO0

?t Mv 0

2v 2

S v 2v

aa

L 185
7.2.3.4 Summary for the Case of Final Velocity, vf * 0 A

Summaries for various shaped pulses are included below:

Deceleration time

V0

Velocity

time

Rectangular Pulse

Pulse Duration: v

Deceleration Level: Vo2

GStopping Distance: va

S

Or: S 32.2Gt 2

Decelera- G

tion

time

Valocity V0

time

186

~~/
Triangular Pulse No. 1

Pulse Duration: 2v.

tuTa

Deceleration Level: 2vo2

G*9636S

Stopping Distance: 2vo 2

Or: 32.2Gt 2 I.

O• ~ ~ - - 6 'i;

Deceleration

time

Velocity

time

Triangular Pulse No. 2

Pulse Duration: 2vo

2 Deceleration Level: vo

G -

2

Stopping Distance: vo

S 32.2G

Ora 32.2Gt 2

.44

187

kI ;

V ~ VI4 '.I..,.A U. UV11A1 - A. ~ V~V
Decelera- t

t ion

StO

0

Velocity

- time

Triangular Pulse No. 3

Pulse Duration: 2v0

t a 

V Deceleration Level: 4vo2 G " -= 0-.

Stopping Distance: 4v2

S 396.6G

Or: S 32.2Gt 2

3

Decelera- G

tion ,time

It

Velocity v

Half-Sine Pulse

Pulse Duration: 1.57v0

t - -1

Deceleration Level: 0.7854vo2 G 2

188

'. !i '.4.
Stopping Distance: 0.7854v 2

0

- 322G

Or: 32.2Gt

A plot of these equations is given in Figure 71. Relative

times and stopping distances are shown in Figure 72 for convenient visual comparison.

INote that the time to stop is equal for all three triangular

deceleration-time pulses, but that the stopping distances are

not. Minimum stopping distance is achieved with the rectangu- lar pulse and hence it is the most desired pulse shape from a

consideration of deceleration from maximum velocity at a given deceleration level in the shortest possible distance.

7.2.4 Energy Content of Aircraft at Impact

In a crash situation, immediately prior to impact, the complete aircraft system possesses a total energy which must be dissipated before the crash sequence is complete.

Sources of energy prior to impact include the following:

9 Translational kinetic energy, (K.E.)T

e Rotational kinetic energy, (K.E.)R

e Potential energy, P.E.

e Strain energy, S.E.

The total energy input during the crash sequence, T.E., is

T.E. - (K.E.)T + (K.E.)R + P.E. + S.E. (45)

7.2.4.1 Translational Kinetic Energy, (K.E.)T: This energy

component is a direct function of the aircraft mass and the

velocity of the mass center at impact.

189
100

W 4

3 ------ ----

200

800

-A

1000 1. "0 C) _V 30

~'120040 I 0

~1400 - '%o

1600 Vi 50

Figure 71. Deceleration, velocity, and distance as

functions of time for five pulse shapes.

190
t Pulse Time to stop Distance to stop

, V_ 2v v 2v 00 0 0

0 a a a a

aI

t

""1L

Figure 72. Comparison of stopping distances for

various deceleration pulse shapes.

if v~ G resultant velocity of mass center

X G n longitudinal component of velocity

§G lateral component of velocity

z vertical component of velocity

4191

-V.
(K.E.)T - mvG

me 2 .2 *2 (XG + G+ z) (46)

The total K.E. of the aircraft is the summation for all mass

elements of the aircraft.

7.2.4.2 Rotational Kinetic Energy, (K.E.)R: Rotational kinetic energy can take the form of total aircraft rotary motions

about the three basic axes of rotation of aircraft elements

such as engines or rotor systems. In a helicopter, two major

sources of rotational energy may exist: the rotor head system, including the blades and rotating swashplates, and the rotating machinery elements, such as engines and transmissions. During

a crash, the element with the energy content may impact external agents and/or the helicopter structure, thus imparting loading to the structural system. However, if, for example,

gas turbine engines are not unduly disturbed from their attitude at impact, it is likely that there will be no appreciable change in the angular velocity of the rotational elements during the time span of the crash impact and no subsequent structural deformation.

Selective assessment of rotational kinetic energy contributions

to the crash energy balance must be made.

Basically, (K.E.) R 2;-1 S1I•

52 + 2 ÷ 7 pI, •2 (47)

where 0 - angular velocity component in x-z plane (pitch)

angular velocity component in y-z plane (roll)Sa 

- angular velocity component in x-y plane (yaw).

and I1, I1, It, are the mass moments of inertia of the vehicle

with respect to pitch, roll, and yaw axes, respectively, at its

mass center.

7.2.4.3 Potential Energy, (P.E.): For each crash sequence, the total potential energy input into the system equals the

vertical displacement contribution of each mass from the time

of impact until the time of completion.

P.E. - 2(mgAZ) (48)

192

* .,
The effects of large mass items, such as transmissions and en-

* gilnes, are major contributors to the energy balance.

7.2.4.4 Strain Energy, (S.E.): Structural strain energy may

exist as a result of structural deflections to accommodate inflight loading. In addition, pressurized systems may represent a level of stored energy. However, such influences are

expected to be minimal during a crash sequence.

7.2.5 Postimpact Energy Dissipation

After the aircraft's initial contact with the impacted surface,

there are several ways of absorbing energy that bring the vehicle to a stop, while also providing a survivable environment

for the occupants.

Possible major contributors in the energy-absorption process

are the following:

1. Structural Deformation - Structural deformation provides the major means of energy absorption in a

severe crash environment. Compression, tension,

bending, torsion, and shear from low levels up to

ultimate conditions all contribute to energy absorption.

Vertical inertia effects are reacted by crushing of

the belly structure, while combined bending and compression of frame and bulkhead members provide support to the large mass items. Longitudinal resintance is provided by ground friction and nose plowing

for impacts on planar surfaces and by compression for

impacts into abutments.

2. Energy-Absorbing Landing Gear - The landing gear, an

integral part of the structure# provides some level

of energy absorption, depending on design criteria

for the particular aircraft involved.

3. Breakaway of Large Mass Items - The breakaway technique for shedding energy has the effect of an instantaneous mass change resulting in a corresponding

reduction in kinetic and potential energies of the

remaining items of concern. A major problem involved

is designing the vehicle to ensure clean breakaway

characteristics with adequate clearance between each

k free mass item and the occupied area. If potential

hazards from impacting occupied areas or generating

fires can be identified, it is better to retain the

large mass items and provide adequate structures to

support them.

* 193
4. Ground Friction and Nose Plowing - Longitudinal deceleration is provided by ground friction and/or

nose plowing, depending on the type of impacted surface. Structural design must preclude the digging in

of structure that increases the deceleration and also

may cause quantities of earth to be scooped inside

the aircraft, or cause the aircraft to pitch over.

Additionally, where hard surfaces are involved, materials that minimize the generation of sparks and

hot spots should be chosen to reduce the potential

for ignition if flammable fluids are spilled.

The ability to push a mass of earth by wedging action

effectively increases the mass of the system. Energy

is absorbed in accelerating the plowed material to

the same velocity as the decelerating aircraft, and,

as the earth mass incrementally increases, effective

braking action due to momentum exchange ensues. Section 6.6.3 pre~sents a more detailed discussion of

plowing.

Impact with ground obstacles such as rocks, trees,

and poles, can be included in this category of energy

absorption. However, the relative sizes of the aircraft and obstacle are significant in establishing

the quantity of energy absorbed. Usually these types

of obstacles absorb small amounts of energy.

5. Energy-Absorbing Seats

The seat is the final link between an aircraft occupant and the ground. The seat and its attachments

to the airframe structure must possess sufficient

strength to keep the occupant within the aircraft's

protective shell. Combined with the restraint system, the seat must prevent injurious contact between

its occupant and the aircraft interior. Because the

human body can tolerate relatively low deceleration

levels in the direction parallel to the spine, it is

imperative that some mechanism for energy absorption

in the vertical direction be included in the seat design.

Although crashworthy seat design is covered in detail

in Volume IV, the energy-absorbing stroke of the seat

must be considered in design of surrounding structure

to prevent interference with seat operation. In fact,

probable longitudinal and lateral deflections must be

added to the strike envelope of the seat in laying

194
out surrounding structure to assure that the critical

vertical nekrgy-absorbing stroke of the seat is not

blocked.

7.3 LANDING GEAR ANALYSIS

.Methods of analyzing performance of both wheel- and skid-type landing gear were surveyed and described in Reference 44. Applicable techniques are summarized below,

7.3.1 Wheel Landing Gear

The methodology for designing a strut-wheel landing gear delineated below is that contained in Reference 27 with some modification.

The typical oleo'.ptrut-wheel landing gear is essentially an

air-oil hydraulic cylinder as shown in Figure 73, which schematically represents one stage of the landing gear illustrated

in Figure 32. The cylinder is pressurized with an air pressure

that acts to balance the static loads of the vehicle and the

dynamic loads during taxi. The air trapped within the cylinder

follows the laws governing compressibility of a gas in a closed

container that are simply described by

P V n P V n (49)

where 2 P = pressure of the gas (lb/in.

V = speciiic volume (in. 3 /lbm)

The subscripts I and 2 define the initial and final states of

the gas, respectively, and the exponent n defines the nature

of the process between states I and 2.

"During taxi, the vehicle rides on an air cushion. Since the

temperature of the air within the cylinder remains essentially

constant during taxi, the process can be considered isothermal

and n is approximately 1. However, during impact conditions

the small time available for heat transfer from the rapidly

compressed air ensures a nearly adiabatic process, for which

n = 1.4.

44. Phillips, N. S., Carr, R. W., and Scranton, R. S., CRASHWORTHY LANDING GEAR STUDY, Beta Industries, Inc.; USAAMRDL

Technical Report 72-61, Eustis Directorate, U.S. Army Air

Mobility Research and Development Laboratory, Fort Eustis,

Virginia, April 1973, AD 765489.

195
Outer . -. %%~

cylinder - rfc

SealsInner

cylinder

Floating

piston

Air4

F Applied force

Figure 73. Shock strut schematic.

196
The hydraulic portion of the cylinder functions to limit loads

during impact conditions. The high stroking rate of the gear

is limited by the pressure generated in the oil as it is forced

through the orifice, rather than by air pressure. As the fluid

is forced through the orifice, the pressure in the cylinder is

defined by Bernoulli's principle for an ideal fluid. The hydraulic force becomes

ph (S) 2

Sr2(50)2 (GdAn)

where p = density (lbm/in. 3 ) Am

j7Ah hydraulic area of the piston (in. 2 )

= stroke velocity (in./sec)

Gd - orifice coefficient

d2

An = orifice area (in.2

This basic equation indicates why landing gear units have

failed in the initial instants of impact without absorbing

significant energy. If the vehicle impacts with a high velocity, the relative velocity across the strut is very large, and

• k the force is proportional to the square of the relative velocity. The strut quickly becomes a rigid link between the

ground and the fuselage, and the extreme forces separate the

IT landing gear from its attachment points.

This applies to a single-stage strut without any blow-off capability. If a blow-off valve is incorporated, as discussed in

Section 6.5, by sizing the orifice, the strut can be designed

to stroke at high velocity levels and high load values. The

incorporation of a variable orifice offers even more control

over the load-stroke relationship and allows more energy to be

absorbed, as illustrated in Section 6.5. The landing gear cylinder also resists motion through bearing frictional forces

that act upon the piston.

The cylinder is supported to resist lateral loads associated

with both operational and crash conditions.

All of the landing gear systems reviewed have the design features mentioned. Many variations are possible. The orifice

usually is combined with a metering pin to adjust the orifice

area with stroke length. Orifice and relief valve combinations

197

...,.. . .. ....... ...... ....... ... .. ...," ......r .. ..
are used to introduce orifice variation as a function of force.

Some liquid springs have been used where the function of the

air pressure is replaced by compression of a fluid. Theae a~rr

a few of the possible variations that produce desirable -.finements of the response but do not alter the basic characLeristics of the landing gear.

The approach to the design of the particular landing gear is

discussed here to demonstrate the various steps leading to a

finished piece of hardware. The major airframe manufacturer

generates a set of criteria for the landing gear design subcontractor. These are the appropriate military specifications,

preliminary weight estimates, moments of inertia, center-ofgravity locations, landing gear stroke requirements, and vehicle attitudes.

If the energy relation is used, the sink rate, gross weight,

and strut efficiency are needed to calculate a load factor.

This is calculated for forward and aft centers of gravity, as

well as for selected attitudes such as level two-point, level

three-point, and tail down. The attitude is important because

it modifies the stroke of the strut. It is assumed that the

vehicle falls vertically, but the strut compresses along its

axis. The output from the energy equation is the load factor.

The load factor is multiplied by the gross weight, and a summation of forces and moments is calculated for each landing condition. For the level three-point condition with maximum verti-

* cal reaction, it is assumed that each landing gear will have

its dynamic vertical force apportioned according to the static

distribution. The drag forces are one-quarter of the vertical

forces, and the side forces are zero.

Therefore, referring to the diagram of Figure 74,

Fv VA + VM + IW NzW W 0 (51)

2ý ZM0 - VA(a + b) - NzW(b) - Tc + *Wb + Ma - 0 (52)

ZD - DA + DM T - 0 (53)

where MS B moment contribution due to landing gear drag

Lu 2

D I

I VA

T - residual thrust

198

--. **L 'F
La

a ,

:• VA -IM

"Pigure 74. Free-body diagram of aircraft on ground.

The system is solved for the reactions VA, VM, DA, and DMO

The next step is to calculate the loads in the landing gear

axes. This is done by equations of the form

Nz - Nvcos 0- N sin 0 (54)

where N and N are the load factors in the earth axis, and G

is the Ittitudi of the strut from configuration drawings of

the various landing conditions as shown in Figure 75.

NV

Figure 75. Landing gear loads.

199

•,g{

'4,
The strut load factors N , N , and N are then used to calculate the forces that eaN lAnding cear must carry. These

forces are tabulated and analyzed to determine critical condi- tions.

The airframe manufacturer then supplies the landing gear manufacturer with the critical design conditions along with gross

weight, maximum load factor, static load, stroke, piston diameter, attachment point locations, overall length, and tire type. The landing gear manufacturer examines each condition and calculates the reaction loads due to each.

The design is significantly influenced by pressure considerations. The air pressure within the cylinder must support the

vehicle in a static configuration. The design specification indicates that the piston is of a given diameter t: carry the

maximum forces and strokes from full extension to full compres- sion. Additional guidance is provided in MIL-L-8552 (Reference 45). From the various design conditions, the maximum

axial force is selected for static coritions. A reasonably low static pressure (200 to 500 lb/irý. ) is desired, although the specifications permit 2500 lb/in.. The difficulty is that

decreased pressures dictate very great stresses during full

compression and necessitate increased material thicknesses.

The cylinder must be designed to stroke a full stroke plus an

amount sufficient to alleviate bottoming. Assuming the additional stroke is sufficient to create a compression ratio of

four, and the static-to-full-stroke dimension is known, an additional stroke is calculated by

(Full stroke - Static stroke) + Additional stroke

Additional stroke 4.0

1. 1 4 The relation P V 4 P V4 is then used to calculate the full

extended pressbA, which must be sufficient to ensure full

extension, and to determine the fully compressed pressure, which is approximately a limit of 2500 lb/in.

The hydraulic operation of the strut is a function of orifice

size and metering pin design. There have been analytical investigations to define the orifice area-stroke relation based

upon a desired constant force response; however, it appears at

present that the design is more a function of previous exper- ience.

45. Military Specification, MIL-L-E552, LANDING GEAR, AIRCRAFT, SHOCK ABSORBER, AIR-OIL TYPE, Department of Defense, Washington, D. C., 10 December 1968.

200

•j! .,
Landing gear manufacturers have orifice and metering pin data

Ask, available with orifice coefficients that will enable them to

design a preliminary configuration. Because the configuration

is selected based upon a worst condition, it remains for test

results to indicate how the strut responds to various inputs.

That is, the design is known from empirical data to generate

the proper efficiency for a given payload and drop height.

Apparently, other conditions are not evaluated analytically

but only through test. By examining the measured data, the

metering pin is modified to reduce undesirable peak forces

and improve the efficiency.

At this point in the system development, the preliminary data

are compiled and the structural design is initiated. Overall lengths are known, externally applied forces are available, and

internal pressure can be specified. These characteristics collectively dictate the necessary wall thicknesses and pivot

point lug sizes. The remainder of the design becomes a stress

analysis problem to be solved using classical techniques.

7.3.2 Skid Landing Gear

As indicated in Figure 6, the skid-type landing gear has been

used for many years on lightweight helicopters. It is a lowcost means of creating small elastic deflections during normal

landings, while providing energy dissipation efficiencies comparable to those of the oleo strut at limit sink speeds of reserve energy impact velocities. The design problem associated

with skids is that they are nonlinear structural elements. The

skid has linearly related applied force and resulting deformation over the small amount of stroke that results in normal

landing impact velocities. Beyond that point it is necessary

to consider the plastic deformation of the skid.

The simplest design technique assumes the vehicle is supported

by tubular members that cross horizontally at the bottonm of the

fuselage. The vehicle impact is in a horizontal attitude and

dissipates all energy by the strain energy required for bending of the tubes.

The skid stiffness is idealized as a bilinear curve with slope

decreasing to a small value after yielding to duplicate the

load-deflection curve. This is derived from an idealized

stress-versus-strain curve for the particular material used.

Since the load-deflection curve is piecewise linear, it is

easily integrated to determine the potential energy as a function of skid deflection. The ealergy absorbed in the elastic

region of the gear is

201 '4

Z Z Z . . . ...
6

sy

E / P(6)d W C6 2 /2 (55)

0

where P(d) - applied load (lb)

6 - deformation of skid (ft)

S a deformation at elastic limit (ft)

C a slope of the elastic raage of the loaddeflection curve (lb/ft)

E1 n energy absorbed (ft-lb).

For the plastic portion of the curve the energy absorbed is

E2 f J P(6 )M

Sy

f [C6Y + C1 (6 - 6)] dS

6

Y

cyy)

C6 ( 6 + Q 2 (56)

where C1 = slope of plastic range (lb/ft).

By equating kinetic and potential energies, equations are developed to relate applied loads, elastic limits, vehicle masses,

weights applied per skid, and impact velocities.

For a perfectly plastic material,

2i

RW (Tv) 2 +(C - C M l) 

5. (57)

+(C - )

202
where R a I rotor lift

weight

W - effective weight (ib)

g a gravitational constant (ft/sec2

v - impact velocity (ft/sec).

Therefore, for varying impact velocities, rotor lift values, and desired effective weights on the skids, the deformation

can be calculated. The load factor applied is then the ratio

of applied force to effective weight:

F

The procedure requires approximating the stress-strain curve

of a material, then approximating the resulting load-deflection

curve for the integration process in the energy equations. The

results of this type of approach have produced reasonable results when compared with test data.

The procedure shown has several distinct steps:

e Establish the stress-strain characteristics of the

tube material.

* Calculate the force-displacement characteristics of

the tube.

* Incorporate the force-displacement characteristics

into an energy relation or set of dynamic response

equations.

* Calculate the vehicle response.

7.4 SEMIEMPIRICAL ANALYSES OF AIRFRAME STRUCTURAL

CRASHWORTHINESS

Several computer programs have been developed recently to analyze crash conditions. However, the accuracy obtained depends greatly on the ability to represent the proper crushing characteristics of the structure. The use of analytical tools to

aid designers would be most beneficial during the preliminary

stages of design when meaningful tradeoffs can be made among

203
weight, cost, space, and crashworthiness capability. It is at

this time in the design process that potential changes can be

achieved with the least cost penalty to a project. At this

stage in the design, only a limited amount of data exists;

probably little more thaa basic weights, stiffnesses, strengths,

configurations, and sizing are available for analysis. To maximize the effectiveness of preliminary design analytical ` ols,

the crushing characteristics of the typical vehicle struk ire

must be obtainable with reasonable accuracy using relatively

simple techniques.

Reference 46 describes a procedure for prediction of crushing

characteristics of primary energy-absorbing structure, as summarized below.

7.4.1 Analysis Procedure

The general procedure for determining the total load-deflection

characteristics, including failure and postfailure behavior

for a given substructure, is discussed below. The procedure

was demonstrated for a test specimen built to represent a section of a utility helicopter lower fuselage. The location of

the segment in the lower fuselage and the four-edge support

representative of the manner in which loads are transmitted

from the fuselage to the transmission housing are shown in

Figure 76. Included in the procedure is a step-by-step process

with the following sequence:

1. Prediction of failure loads for stiffened panels.

2. Postfailure analysis of stiffened panels.

3. Main beam and bottom skin analysis.

4. Total substructure load-deflection curve.

46. Wittlin, G., and Park, K. C., DEVELOPMENT AND EXPERIMENTAL

VERIFICATION OF PROCEDURES TO DETERMINE NONLINEAR LOAD

DEFLECTION CHARACTERISTICS OF HELICOPTER SUBSTRUCTURES

SUBJECTED TO CRASH FORCES, Volumes I and II, LockheedCalifornia Company; USAAMRDL Technical Reports 74-12A,

74-12B, Eustis Directorate, U.S. Army Air Mobility Research

and Development Laboratory, Fort Eustis, Virginia, 1.974, AD 784191 and AD 784192.

204
44

00

H0 0

44

00

H 4J4

NN

H0

P14 0

44.

u $4.

944

(A

04

0.

'4:4

$4.

00

0 to 0 .1

205
In predicting the failure load of the model substructure, extensive use of the work of Needham (Reference 47) and Gerard

(Reference 48) was made. For the postfailure analysis, the

concept of D'Amato (Reference 49) was used as a base; however,

a somewhat different failure mechanism was employed. Subelement crushing characteristics were superimposed in a piecewise

sense to yield the total load-deflection curve. Depending

upon the rivet pitch, spacing, and strength, a given structure

may fail in one of several ways. Therefore, the procedure described herein takes into account monolithic, wrinkling, and

interrivet failure modes. The basic equations were developed

for a nondimensional slenderness ratio of L/p - 20. However,

column effects were included for L/p ratios greater than 20.

Although developed for angle-type stiffened panels, the procedure is reported to be applicable to a variety of panel types,

including T-type stiffeners, formed angle stiffeners, extruded

angle stiffeners, formed multicorner sections, hat-formed

"stiffened panels, and extruded-Y stiffened panels.

The failure loads were estimated using the semiempirical/

analytical methods described in References 47, 50, and 51 for

stiffened short panel elements. The analysis, as outlined, is

based on two assumptions: (1) at the threshold of the failure

load, full plastic hinges are developed at the constrained supports and the midsection of the panels, and (2) the free warping energy of the flange of the stiffening angles can be neglected.

As is discussed in Section 7.4.3 and in detail in Reference 46,

these assumptions are considercd on thc avcragc reasonable.

The analysis neglects the effect of strain hardening, the influence of the axial force on the plastic hinge mechanisms of

47. Needham, R. S., THE ULTIMATE STRENGTH OF ALUMINUM ALLOY

FORMED STRUCTURAL SHAPES IN COMPRESSION, Journal of Aerospace Science, Vol. 21, 11o. 4, 1954, pp. 217-229.

48. Gerard, G., HANDBOOK OF STRUCTURAL STABILITY, NACA Technical Note 3781-3785, National Advisory Committee for

Aeronautics, Cleveland, Ohio, 1957.

49. D'Amato, R., STATIC POSTFAILURE STRUCTURE CHPRACTERISTICS,

WADC Technical Report 59-112, Wright Air Development Center, Wright-Patterson Air Force Base, Ohio, 1959.

50. Chitner, A. H., THIN WALLED STRUCTURES, John Wiley, New

York, 1967.

51. Sechler, E. E., and Dunn, L. G., AIRPLANE STRUCTURAL ANALYSIS AND DESIGN, Dover, New York, 1963.

206
the stiffened panel, the changes in local failure pattern during the postfailure stage, and the effect of geometrical imperfections caused by manufacturing and/or damage. In light Sof the test results and correlation with analysis, these effects are probably not significant for this type of substructure.

7.4.2 Substructure Test Results

To verify the analytical results, experimental data were obtained during the same study by testing twelve specimens, representative of typical helicopter lower fuselage structure,

under the following conditions: static load, dynamic impact (e30 ft/sec), four-edge support, two-edge support, skin web

thickness from 0.025 in. to 0.064 in., number of angle stiffeners from 12 to 40, and specimen depth of 6.125 in. and

12.125 in. The test specimens were 46 in. long, included five

bays, and were 18 in. wide with a side panel at each side.

Each specimen was supported in its normal manner. All compara- tive static and dynamic tests on similar specimens were performed at equal energy input levels. Instrumentation consisted

of several strain gages to measure compression and bending, a

linear variable differential transducer to measure deflection, accelerometers mounted on the impact head and load cells installed between the test specimen support and the ground for

the dynamic tests.

The test program results yielded valuable information for future consideration. For example, for these types of structural elements, static testing to determine load-dcfloction

characteristics should yield sufficiently accurate results

when compared to dynamic test results, but in a more economical manner. A static test can provide the desired information

with regard to shape of the load-deflection curve, peak failure load (in the program described herein the dynamically obtained failure loads are between 9 and 24 percent higher than

the loads measured in static testing and the loads from predicted static calculations, respectively), deflection at which

failure load occurs, and energy-absorption capability of the

structure.

The test results show that during a dynamic test the amount of

springback from the maximum deflected value is immlediately evident. After a static test the structure will rulax slowly (up

to several days) to its permanently deformed position. The

springback in the dynamic test was as high as 40 percent of the

maximum deflected value for a test performed with an impact velocity of 30 ft/sec. This observation could be of consequence

in areas such as doors, where distortion of structure affects

4 occupant egress after a crash.

207

.," " ,
7.4.3 Correlation of Test and Analysis Results

Comparison of test and analytical data showed reasonable agreement with respect to peak failure load and energy-absorption

capability, particularly for the dynamic tests in which load

cells are installed. Figure 77 shows a comparison of the analytically predicted load-deflection curves with test data and

illustrates good correlation with regard to shape of the curve

as well as peak load and energy absorption. As can be noted

in Figure 77, the typical response of the type of structure

tested in this study is a relatively sharp load buildup in the

elastic region up to the peak failure load (at approximately

0.35 in. or less of deflection). After failure, the load decreases at a rate less rapid than the initial buildup until the

deflection reaches approximately 0.75 to 1.0 in. Thereafter,

the load decreases gradually until the structure being crushed

is confined and once again stiffens. This region, wherein confined crushing takes place, is very significant because the

stiffness at this point can greatly influence the response of

the upper masses (the transmission and engine of the test vehicle). The initial peak failure load, while of a substantial

level, is of short duration and as such may not greatly influence the upper mass response.

30.0

0.0

0.00 1.00 2.00 3.00

Deflection, in.

Figure 77. Predicted-versus-test load-deflection curves for

a representative specimen. (From Reference 46)

Figure 78 shows a comparison of test and analytical results

for three groups of specimens. A detailed discussion of the

correlation between analysis and test is presented in Referenice 46. The energy absorption was predicted within 8 percent

of the test results for three of the five specimens that were

208 9'
Peak load (lb) Ez~er¥y absorbed (in.-lb)

Specimens (+7.0%) Specimens 4012

2-4 26,500 2-4 39,975

87 420 110,380

specimens' (22%) Specimens

5-6 ( 2,500 5-6 (+17,6%)

68,500 129,800

Specimens 132,000 (Static) 144,000 IDynamic) 133.59n specimens '' . . .. .. . .. Specimens

136,000 (44.2 9) to (-4.2%) 7,8,123

Teat (values shown are the average of applicable tests)

Analysis

Figure 78. Comparison of test and analysis results

for "like" specimens.

tested dynamically with load cells installed. Correspondingly, the load was predicted within 7 percent. For the other two

specimens, the data are based on accelerometer time histories

rather than load cells. These traces contain considerable

noise, which makes it difficult to interpret the test data.

The results of the correlation studies indicate that it is

practical to use relatively simple techniques to determine

load-deflection characteristics of multiweb, angle-stiffened

panels, typical of the lower fuselage of an existing utility model helicopter. The overall crushing behavior of the substructure can be predicted within reasonable accuracy although

the failure mode of each individual stiffener cannot consistently be accounted for. The analytical procedure described

herein is capable of determining satisfactorily the two most

significant parameters for crushable structure, i.e., peak

failure load and energy-absorption capability. The approach,

* as developed in this study, has limitations with regard to accurately predicting the deflection at which failure occurs and

the deflections at which postfailure stiffening is encountered.

Although the predicted failure deflection is a consistently

lower value than that obtained from tests, it represents an extremely small percentage of the total crushing energy, to the

"extent that it is not significant as long as the peak failure

load and load-deflection characteristics are properly defined.

The analytical assumptions regarding the mode of failure in the

postfailure region and the hinge formation at failure are considered valid in view of the test results. For example, although test results indicate that failures can occur in both

209

. ......- .-. 
the asymmetrical and symmetrical modes, the predicted energy- absorption capability of the structure, which considers only ' the symmetrical mode, is still very close to test results. Thus, the assumption of a symmetrical failure is, on the average, reasonable.

7.4.4 Example of Analysis for Vertical Impact

The example given here is an analysis to predict major structural collapse for the crash of a medium cargo helicopter at a

velocity representative of a 95th-percentile potentially survivable accident. This analysis was performed for the full- scale crash test of a cargo helicopter conducted on 6 March

1975 at the NASA Langley Research Center's Impact Dynamics Research Facility, Hampton, Virginia (Reference 52). The impact velocity components selected for the test were 40 and 30 ft/sec for vertical and longitudinal velocities, respectively, produc- ing a 50-ft/sec resultant velocity.

The structural response and failure modes in a crash environment are not easy to predict using static analytical techniques. Initially, the underfloor structure absorbs energy by crushing, but the amount of energy absorbed is dependent upon the type of ground impacted. Soft ground can displace and provide a

reasonably uniform loading on the base of the structure, but a

rigid ground plane results in more direct loading of the frame

members. This difference is illustrated in Figure 79, where

possible groulnd impact types planes. of load distribution are shown for soft and hard

The helicopter occupants, assuming an otherwise survivable environment, must be protected from injury due to the collapse of

basic structure or the penetration of occupied areas by large

mass items. The following calculations demonstrate a semiem- pirical method of predicting overall reduction in the height of occupied areas.

Data obtained from past major accidents, either actual values

or estimated impact conditions, have been used in conjunction with deformation information to obtain a simplified relation- ship between the static and dynamic load-carrying capabilities for structural elements. Tie case considered here involved the absorption of 135,000 ft-lb of energy with a 42.2-in. structural collapse.

52. Singley, G. T., III, FULL SCALE CRASH TESTING OF A CH-47C

HELICOPTER, paper presented at 32nd Annual National V/STOL Forum, American Helicopter Society, Washington, D. C.,

May 1976.

210

3AL
Vertical member

1A .Load concentrated

towards vertical

ý ,members (no ground

deformation)

Load distribution for impact

on hard ground, concrete

S• Uniformly distributed

load (ground deforms

to accomodate shape

of vehicle bottom)

Load distribution for impact

on soft ground, wet sod

Figure 79. Types of load distribution acting on helicopter

structure for impact with soft and hard surfaces.

211

I, .. , I,.
The general nature of the energy-absorbing characteristics of

an element of airframe structure can be of the form shown in

Figure 80. When stability is not the limiting criterion, the

curve in Figure 80(a) is applicable; Figure 80(b) represents

the response for stability-limited load capability.

.1

Displacement.

(a) Load capability not limited by stability

Displacement

(b) Load capability limited by stability

Figure 80. Typical load-displacement characteristics

for airframe structure.

212
0.1 iii:•- In severe impact conditions, stability is the limiting criter-

"ion for many structural elements of interest. Figure 81 shows

simplified load-deflection characteristics for static load apY •, plication to a stringer-skin combination subjected to compressive loading (Reference 53).

100,

5: Coordinates

Load Deflection

( 9•1.0 0.0387

-4 0.7383 0.0757

0.51 0.0915

8. 0.2819 0.0968

0.1879 0.1239 0.1436 0.1479

P0• 107 . 0.0 0.0309 0.4225 1.0

60

s60

W.i 40O

0 10 20 30 40 50 60 70 80 90

Percent of maximum deflection

Figure 81. Simplified load-deflection characteristics

for a statically loaded stringer skin

combination. (From Reference 53)

53. Greer, D. L., Heid, T. L., and Weber, J. D., DESIGN STUDY

AND MODEL STRUCTURES TEST PROGRAM TO IMPROVE FUSELAGE

CRASHWORTHINESS, Convair, Division of General Dynamics

Corporationj FAA DS-67-20, Federal Aviation Administration, Washington, D. C., October 1967, AD 666816. 9 213
The area under a load-deflection curve determines the energyabsorption capability for the structure under consideration.

The basic section of a typical side frame member is shown in

Figure 82, The crippling stress is computed as follows:

S Skin (.025-in.

thick 2024-T4)

Harvey Aluminum Co.

11301-3

7075-T6 extrusion

0.040-in. thick

2024-T4

Alcoa 59574

7075-T6 extrusion

Figure 82. Typical frame section of side element of

a medium cargo helicopter center fuselage.

214

.... ... ... ... .
For Harvey Aluminum Company extrusion 11301-3,

A 0.210

0.094 + 0.125 - 859

,FC * 62,000 lb/in. 2 (From Figure 83)

For Alcoa 59574,

"A 0.289 - 2 = 5.67

t2 0.1882 + 0.1252

2 F = 70,000 lb/in.

cc

For the flanged web,

A (G.5)(0.040) 18.8

gt 2 (2)(0.04021

F 19,000 lb/in. 2 (From Figure 84) cc

The total compressive capability of the frame element is computed as follows:

Capability of stringer elements * (62,000)(0.210)

+ (70,000)(0.289)

+ (19,000) (1.5) (0.04)

* 34,390 lb (From Figure 85)

215
N N N

4 H 0

44,

f-44

ONH
Q

Q

0

Id 00

'H 

0

0 a) 0"*

0)4) M I I I I I -- r0

."I I 0Y iA .44

0

0 0 0 

r~r- Nc' 41

1E II I H

00 00 0 0 0 % '0 4 NH

0 wN c HTsX1( a)sseaqs Bu1TddT.ao e~voTT

217
0. 004"

0.006 - -1001- -

0.008"

0.010

0 )-

0.080 - -_

Allowabe striner trs ( , e

Cie

Figue Effctie 85 ara ofskinforalumnum llo

0tige-ki1o5iatos

21800
Capability of adjacent skin * (37,000)(0.034)

+ (37,000)(0.032)

+ (19,000)(0.024)

2,898 lb

Total compressive capability - 37,288 lb

This value is for one section through a typical frame member.

for a symmetrical fuselage section, the total compressive capability is twice this value, or 74,576 lb.

From the photographs of helicopter crashes under known impact

conditions, the average crushing of the typical frame section

was 42.2 in. for the impact velocity considered.

Using this load value and total displacement in conjunction

with the typical stringer-skin compression static characteristic shown in Figure 81, an estimation can be made for th, correction required for dynamic energy-absorption capability. "Using Figure 81 data, the energy absorbed is given by:

EA= P y Z A

A(74576) (,,)2 A (59)

where A - incremental area under curve in Figure 81

(74576)(3.5) [(1.0)(0 0-87) + ( 0)0.037

+ 0.7383 + 0.5100)00158 + (0.51 + 0.2819)0.0053

+ (0.2819 2 + 0.1879 )0.0271 (01879 0)18797+ + 0.1436) 0.1436 0240

÷ 0.1436 + 0.0309 0157

14 2 )0.2746 + (0.0309) (0.575

219

..... !
= 261,016 (0.01935 + 0.03216 + 0.00986 + 0.00210

+ 0.00637 + 0.00398 + 0.02396 + 0.00892)

= 27,850 ft-lb (STATIC)

This result shows that the static collaipse energy computation

does not agree with the empirical value of 135,000 ft-lb.

Assuming rectangular characteristics would produce an overly

optimistic energy-absorption value of 261,016 ft-lb. An assumption of triangular characteristics yields an energyabsorption capability of 130,508 ft-lb, and a comparison of

this value with the actual kinetic energy of 135,000 ft-lb

yields an error of only 3.34 percent, which is well within an

acceptable range for such an analysis. Figure 86 shows the

differences between the static and dynamic energy-absorption

capabilities of a typical structural element; the result is

based on actual test and accident data.

7.4.5 Example of Analysis for Longitudinal Impact

In the previous example, structural resistance to loading was

assumed to be in the vertical plane of the helicopter. Additional structural crashworthiness requirements for the primarily longitudinal-lateral impacts are specified in MIL-STD1290(AV) (Reference 1) and Section 5.3.1 of this document.

A similar method of anallsis can be employed after defining

the primary structural members assigned to resist the applied

loading and after computing load-carrying limitations and potential stroking distances.

As a further example, the nose section of a developmental utility ,elicopter will be used. In the initial design definition

phase, the requirements of MIL-STD-1290(AV) were specified for

both the 20- and 40-ft/sec longitudinal impact conditions (see

Table 3, Section 5.1). These requirements delineated the maximum acceptable intrusions into occupied space for the given impact velocities for longitudinal impacts into a vertical abutment. The basic design philosophy is that the kinetic energy

involved in a 20-ft/sec crash impact velocity is absorbed into

the 25 in. of structure forward of the rudder pedals. For a

40-ft/sec crash, energy is absorbed by the structure forward of

Station 78; the area aft of Station 78 constitutes the cabin

area.

220
Ideal energy

absorption

100

90

80

0 70

'I Assumed dynamic :,60 .capability (based

on empirical data)

50

,40 --Static

capability

p43 0

Is 20

10 l

0 10 20 30 40 50 60 70 80 90 100

Percent of maximum deflection

Figure 86. Estimated dynamic energy-absorption capability

of typical fuselage structure.

221 0,
7.4.5.1 20-ft/sec Impacts Figure 87 shows the basic structural elements to be considered; a summary of the analyses is

as follows:

Longitudinal impact velocity changes - 20 ft/sec

Basic structural weight - 15,587 lb

Kinetic energy to be absorbed M 1 m X2

.() ( ((202

- 96,813 ft-lb

Figure 87 shows the basic structural sections that resist longitudinal loading. It should be noted in Figure 87 that these

members are stabilized by formers located at approximately a

6-in. pitch.

Using standard methodology for computing the crushing strength

of each section yields:

Crushing strength of upper cap, BL 14.5 (left and

right sides) a 25,569 lb

Crushing strength of lower cap, BL 14.5 (left and

right sides) m 17,919 lb

NOTE: When these values are computed, the amount of

effective skin actively participating is depen-.

dent on whether interrivet buckling occurs.

If interrivet buckling is considered likely,

the effective width of skin must be reduced as

follows:

Fir '

Wcorrected -Weff Fir

2cs

222

1i•,
LAX0 LA %

_ LAM I LA

LAn

- 0 . f n

IILL

N- Ln44J

Ln m 00

ý4.

toJ 41

14-4

IN /4 . 0 4i 

NN

14$4

IWI

NO LA

04-

41 0

'I, k% r 0. 

C) E-4 223
Where Fir * interrivet buckling stress

FC. u crippling stress of stringer.

Summation of the strength capabilities of the four longitudinal

sections yields:

Total crushing strength, Station 13 to 38

* 2 (25568 + 17919)

. 86,976 lb

Energy-absorbing capability

- 0.5 x 25 x 86796

* 90,600 ft-lb

Equivalent Impact Velocity - 19.4 ft/sec

This value, within 3 percent of the desired 20-ft/sec capability, is acceptable when considering the assumptions made and

the use of an empirical dynamic factor of 0.5.

7.4.5.2 40-ft/sec Imgactt Similar computations are used for

this analyss, but additional structure is involved. After the

section from Station 13 to 38 has collapsed, members from Station 38 to 78 pick up the loads. The members involved are indicated in Figure 87, but the computations are not given since

the process is the same as for the 20-ft/sec case.

7.4.6 Lateral Impact

Once again, resisting structure is identified and energyabsorption capability computed for the frame member sections

in the crown and floor that resist volume reduction due to

loading in the lateral direction. Summation of the buckling

strengths, as demonstrated in Section 7.4.4, is used to determine volume reduction for the specified impact velocity to conform with the MIL-STD-1290(AV) requirements (Reference 1) summarized in Section 5.3.1.3.

224

I. .. .....
7.4.7 Rollover

Postimpact rollover design requirements for load applications

* and for structural areas that must withstand that loading

are specified in MIL-STD-1290(AV) and summarized in Section

5.3.1.4. Basically, a 4W requirement is definedy this can be

satisfied by normal static structural analysis techniques.

Since a rollover maneuver is a secondary effect occurring after

initial impact# an energy-absorbing analysis is unnecessary.

The 4W, postimpact static analysis is adequate for substantiating the rollover performance of the helicopter because normally the primary impact will absorb most of the energy.

7.5 STRUCTURAL CRASHWORTHINESS SIMULATION COMPUTER PROGRAMS

Crashworthiness analysis as a scientific discipline is in a

transitional stage between pure research and support of design

functions. A number of digital computer programs for analysis

of vehicle structures in a crash environment have been devel-,

oped by the research community, and several of these, generally the simpler, have begun to find use in preliminary design. The more significant programs have been critically reviewed

recently by several authors, including K. J. Saczalski (Reference 54), Hayduk, et al. (Reference 55), I. K. MoIvor (Reference 56), and M. P. Kamat (Reference 57). The programs vary

widely in their modeling characterization and mathematical

54. Saczalski, K. J., STRUCTURAL PROBLEMS ASSOCIATED WITH THE

PREDICTION OF VEHICLE CRASHWORTHINESS, in Surveys of Research in Transportation Technologyf AMD-Vol. 5, presented

at ASME Winter Annual Meeting, American Society of Mechanical Engineers, 11-15 November 1973.

55. Hayduk, R. J., et al., NONLINEAR STRUCTURAL CRASH DYNAMICS

ANALYSES, SAE Paper No. 790588, presented at Business

Aircraft Meeting, Society of Automotive Engineers, Inc.,

Wichita, Kansas, 1979.

56. Mclvor, I. K., MODELING, SIMULATION, AND VERIFICATION OF

IMPACT DYNAMICS, Volume I, EXECUTIVE REPORT, Highway

Safety Research Institute and Department of Applied Mechanics, University of Michigan, Ann Arbor, Michigan, October 1973.

57. Kamat, M. P., SURVEY OF COMPUTER PROGRAMS FOR PREDICTION

OF CRASH RESPONSE AND OF ITS EXPERIMENTAL VALIDATION, in

Measurement and Prediction of Structural and Biodynamic

Crash-Impact Responseg Winter Meetinq, ASME, American

Society of Mechanical Engineers, New York, 5-10 December

1976, pp. 33-48.

225
treatment of the model equations. They are easily grouped,

* however, into five main classes, namely:

1. Simplified spring-mass models.

2. Generalized spring-mass models.

3. Hybrid models. !i I

4. Frame-type models.

5. Finite element models.

Simplified spring-mass models represent the vehicle by one to

four lumped masses and up to eight nonlinear spring elements.

Generalized spring-mass models are similar to the above except that they generally treat a larger number of masses and spring

elements, which are typically more generally defined, thus

lending some flexibility to the modeling process. Cited examples of spring-mass models are those of Sato (Reference 58),

Emori (Reference 59), Tani and Emori (Reference 60), Miura and

Kawanturo (Reference 61), all in the simplified category, and

58. Sato, T. B., et al., DYNAMICAL CONSIDERATIONS ON AUTOMOBILE COLLISION, Journal of the Society of Automobile

Engineers of Japan, Vol. 20, No. 5 (1966) and Vol. 21,

No. 9 (1967).

59. Emori, R. I., ANALYTICAL APPROACH TO AUTOMOBILE COLLISIONS, SAE Paper No. 680016, presented at Automotive Engineering Congress, Society of Automotive Engineers, Inc.,

Detroit, Michigan, January 1968.

60. Tani, M., and Emori, R. I., A STUDY IN AUTOMOBILE CRASHWORTHINESS, SAE Paper No. 700175, presented at Automoti,,e

Engineering Congress, Society of Automotiva Engineers,

Inc., Detroit, Michigan, January 1970.

61. Miura, N. and Kawamura, K., AN ANALYSIS OF DEFORMATION

MECHANISM IN HEAD-ON COLLISION, SAE Paper No. 680484,

presented at Society of Automotive Engineers Mid-Year

Meeting, Society of Automotive Engineers, Inc., Detroit,

Michigan, May 1968.

226

IA 
those of Gatlin, et al. (Reference 62) and the Battelle Columbus Laboratories (Reference 63) in the generalized olass.

Hybrid models are those requiring experimentally det3rmined

component behavior data as program inputs. A program developed by M. M. Kamal at the General Motors Resaarch Laboratory is considered typical of this class (Reference 64). It uses

three lumped masses and eight nonlinear spring elements to simulate unidirectional, frontal, or barrier impacts of automobile

vehicles.

Frame-type models employ beam elements, instead of spring elements, and lumped or rigid body masses at the intersecti.ons

(nodes) of the beam elements in either two-dimensional or

three-dimensional configurations. Typical of this class are

the crash simulation programs developed by researchers at

Lockheed-California (Reference 65), Calspan (Reference 66),

62. Gatlin, C. I., Goebel, D. E., and Larsen, S. H., ANALYSIS

OF HELICOPTER STRUCTURAL CRASHWORTHINESS, Volume I. MATHEMATICAL SIMULATION AND EXPERIMENTAL VERIFICATION OF HELICOPTER CRASHWORTHINESS, Dynamic Science, Division of

Marshall Industries/ USAAVLABS Technical Report 70-71A,

Eustis Directorate, U. S. Army Air Mobility Research and

Development Laboratory, Fort Eustis, Virginia, January

1971, AD 880680.

63. Herridge, J. T., and Mitchell, R. K., DEVELOPMENT OF A

COMPUTER SIMULATION PROGRAM FOR COLLINEAR CAR/CAR AND CAR/

BARRIER COLLISIONS, Battelle Columbus Laboratories/ Report

No. DOT-HS-800-64S, U. S. Department of Transportation,

Washington, D. C., January 1972.

64. Kamal, M."M., ANALYSIS AND SIMULATION OF VEHICLE TO BARRIER IMPACT, SAE Paper. No. 700414, presented at the International Automobile Safety Conference, Society of Automotive Engineers, Inc., Detroit, Michigan, May 1970.

65. Wittlin, G., and Gamon, M. A., EXPERIMENTAL PROGRAM FOR

THE DEVELOPMENT OF IMPROVED HELICOPTER STRUCTURAL CRASHWORTHINESS ANALYTICAL AND DESIGN TECHNIQUES, Volumes I

and Il, Lockheed-California Company; USAAMRDL Technical

Reports 72-72A and 72-72B, Eustis Directorate, U. S. Army

Air Mobility Research and Development Laboratory, Fort

Eustis, Virginia, May 1973, AD 764985 and AD 764986.

66. Shieh, R. C., BASIC RESEARCH IN CRASHWORTHINESS II - LARGE

DEFLECTION DYNAMIC ANALYSIS OF PLANE ELASTOPLASTIC FRAME

STRUCTURES, Calspan Corporation: Report No. DOT-HS-800-781, V U. S. Department of Transportation, Washington, D. C.,

December 1972.

227
Philco-Ford (Reference 67), and Chrysler Corporation (Reference 68). Major differences among these lie primarily in the

conceptual representations of the highly nonlinear behavior of

the beam elements, and in the numerical methods used in the

time-history solutions. The Calspan program limits application

to two-dimensional structures, whereas the other three are capable of predicting three-dimensional response.

The finite element approach represents an attempt to make up

for the limitations inherent in the lumped-parameter analyses,

by employing more formal approximation techniques in the diecretization of the structure and a greater reliance on more

fundamental structural representations and properties. The

limitations of this approach are found in the inherent tendency

toward more complicated and expensive computations and the difficulties found in modeling the extensively complex phenomena

associated with the crash environment. Such phenomena include

large deflections and rotations in the deformed structure, regions of intense curvature (wrinkling), material strain rate

effects, and interference and contact among structural components during the response. These procedures are not totally

free of a reliance on testing and analytical judgment and, in

fact, can be viewed as a somewhat more formal lumped-paramneter

systcm and used in connection with the more empirical procedure

in hybrid, combination models.

A variety of finite element analyses directed toward the dynamic analysis of vehicle (primarily automotive) structures

have been reported previously. However, most have been limited

to beam-type elements and have, therefore, lacked the advantages of the general-purpose elastic analysis programs widely

used in industry today. These models have been grouped with

the frame-type models above. The response of vehicle structures under crash loadings is a complex process primarily involving:

61. Young, J. W., "CRASH": A COMPUTER SIMULATION OF NONLINEAR

TRANSIENT RESPONSE OF STRUCTURES, Philco-Ford, Subsidiary

of the Ford Motor Company; Report No. DOT-HS-09-1-125B,

U. S. Department of Transportation, Washington, D. C.,

March 1972.

68. Thompson, J. E., VEHICLE CRUSH PREDICTION USING FINITEELEMENT TECHNIQUES, Chrysler Corporation; SAE Paper No.

780157, Society of Automotive Engineers, New York, January 1973.

228

i, i,
* Transient, dynamic behavior.

e Complicated framework and shell assemblages.

9 Large deflections and rotations.

* Extensive plastic deformations.

Most attempts at a complete, formal analysis of this process have been only partly successful due to a variety of limitations which, in particular instances, have included inadequa- cies in element formulations, material representations, or

solution procedures. The field of nonlinear finite element

analysis is currently an extremely active area of research with

an extensive, related literature and a variety of methods and

approaches. Consequently, a complete review of the field as

background for this document is not attempted. Instead, major features of two finite element programs that have been specifi- cally tailored to the claso of problems inherent in vehicle

crash response and employ or extend current avenues in finite

element analysis that seem best suited to such problems are

discussed. These are Grumman's DYCAST (Reference 69) and

IITRI's WRECKER (Reference 70).

Reference 55 presents the results of three nonlinear computer

programs, KRASH, ACTION (Reference 71), and DYCAST, used to

analyze ths dynamic response of a twin-engine, low-wing airplane section subjected to a 27.5-ft/sec vertical impact velocity crash condition. The three distinct analysis techniques for nonlinear dynamic response of aircraft structures are

69. Winter, R., Pifko, A. B., and Armen, H., Jr., CRASH SIMULATION OF SKIN-FRAME STRUCTURES USING A FINITE ELEMENT

CODE, SAE Paper No. 770484, presented at Business Aircraft

Meeting, Society of Automotive Engineers, Inc., New York,

March 29 - April 1, 1977.

70. Yeung, K. S. and Welch, R. E., REFINEMENT OF FINITE ELEMENT ANALYSIS OF AUTOMOBILE STRUCTURES UNDER CRASH LOADING, Volume II, IIT Research Institute; Report No. DOTHS-803-466, U. S. Department of Transpoirtation, Washington, D. C., October 1977, PB 287589.

71. Winter, R., Pifko, A. B., and Cronkhite, J. D., CRASH SIMULATION OF COMPOSITE AND ALUMINUM HELICOPTER FUSELAGES

USING A FINITE ELEMENT PROGRAM, AIAA Paper 79-0781, prb- sentod at AIAA/ASME/ASCE/AHS 20th Structures, Structural

Dynamics, and Materials Conference, St. Louis, Missouri,

4-6 April 1979.

229

4'.. .. ..:-...:Z "w . .. •:_ . Z * .. . ..... ........... :...••i••:}
q,.

briefly examined and compared versus each other and the experimental data. The report contains brief descriptions of the

three computer programs, the respective aircraft section mathematical models, pertinent data from the test performed at NASA

Langley, and a comparison of the analysis versus test results.

Cost and accuracy comparisons among the three analyses are made

to illustrate the possible uses of the different nonlinear programs and their future potential.

The remainder of this chapter presents the structural models

of greatest potential use in aircraft structural crashworthiness.

7.5.1 Program KRASH

The computer program KRASH was originally developed by LockheedCalifornia Company under U. S. Army auspices to analyze the dynamic response of helicopters subjected to a multidirectional

crash environment (Reference 65). Subsequent development of

KRASH was sponsored by the Federal Aviation Administration

(FAA). The FAA's goal was to acquire an analytical tool with

minimal additional development that could assist in the performance of structural crash dynamic analyses of general aviation

fixed-wing airplanes. The general aviation version of KRASH

has been exercisod on four full-scale single-engine, high-wing,

aircraft crash tests performed at NASA Langley's Impact Dynamics Research Facility.

Program KRASH is a digital computer program which predicts the

structural response of vehicles to multidirectional crash environments. The program computes the time histories of N interconnected masses, each allowed six degrees of freedom, defined by inertial coordinates Xi, Y., Z and Eulerian angles

0, 0 0 0, whare i a 1,2,...N. Euler'l equations of motion

ate wlittin for each mass. The equations of motion are integrated numerically to obtain velocities, displacements, and

rotations. Gravitational forces, internal forces and moments,

and external forces are computed. For small deflections, a

linear analysis is obtained, and for large deflections, general

plastic deformation is allowed. The program provides for unloading and subsequent reloading along a linear elastic slope.

Program KRASH describes the interaction between a series of

massless interconnecting structural elements and concentrated

rigid body masses to which the structural elements are attached

at their ends with the appropriate end fixity (pinned, fixed).

The structural elements can be connected between node points

which are offset from, and rigidly attached to, selected mass

points. The interconnecting elements represent the stiffness

characteristics of the structure between the masses. The

230

I.J
masses can translate and rotate in all directions under the

4 influence of the external forces (i.e., gravity, aerodynamic, impact), as well as the constraining internal element forces.

The movement of the masses results in changes in the relative

distortion of the structural elements and, in turn, results in

a new set of element forces acting throughout the system.

Computer Program KRASH has the capability to:

e Define the response of six degrees of freedom at each

representative location, including three translations

and three rotations.

0 Determine mass accelerations, velocities, displacements, and internal member loads and deformations at

each time interval.

e Provide for general nonlinear stiffness properties

in the plastic regime, including different types of

load-limiting devices, and determine the amount of

permanent deformation.

* Define how and when rupture of an element takes place

and redistribute the loading over the structural elements involved.

9 Define mass penetration into an occupiable volume.

e Define the volume change due to structural deformations of an occupiable volume.

* Provide for ground contact by external structure including sliding friction and a nonrigid ground surface.

* Include internal structural damping.

o Include a measure of injury potential to the occupants, e.g., the probability of spinal injury indicated by the Dynamic Response Index (DRI).

* Determine the distribution of kinetic and potential

energy by mass item, the distribution of strain ant

damping energy by beam element, and the crushing and

sliding friction energy associated with each external

spring.

* Determine the vehicle response to an initial condition that includes linear and angular velocity about

three axes and any arbitrary vehicle attitude and

position.

231
9 Provide a measure of the airplane CG velocity by

means of translational momentum relationships.

* Analyze an impact into a horizontal ground and/or an

inclined slope.

* Provide a measure of the internal stress state of

internal beam elements.

e Analyze a mathematical model containing up to 80

masses and 150 internal beam elements.

* Treat up to 180 nonlinear element degrees of freedom.

A comprehensive discussion of the theoretical development of

program KRASH is presented in Reference 72. A discussion of

program KRASH input-output format as well as modeling techniques and applications are included in Reference 73. Reference 74 provides a discussion of design guidelines which can

be used in conjunction with program KRASH in the structural

crashworthiness analysis of general aviation airplanes. Information related to the program's system requirements and

functional organization to facilitate bringing the program to

an operationl status on an individual user's computer system is contained in Reference 75.

72. Gamon, M. A., GENERAL AVIATION AIRPLANE STRUCTURAL CRASHWORTHINESS USER'S MANUAL, Volume I, PROGRAM "KRASH" THEORY,

Lockheed-California Companyl FAA-RD-77-189-1, Federal

Aviation Administration, Washington, D. C.# February 1978.

73. Gamon, M. A., Wittlin, G., and LaBarge, W. L., GENERAL

AVIATION AIRPLANE STRUCTURAL CRASHWORTHINESS USER'S MANUAL, Volume II, INPUT-OUTPUT, TECHNIQUES AND APPLICATIONS

(Revised), Lockheed-California Company; FAA-RD-77-189-II,

Federal Aviation Administration, Washington, D. C., Sep- temuber 1979.

74. Wittlin, G., GENERAL AVIATION AIRPLANE STRUCTURAL CRASHWORTHINESS USER'S MANUAL, Volume III, RELATED DESIGN INFORMATION, Lockheed-California Companyl FAA-RD-77-189-11I,

Federal Aviation Administration, Washington, D. C., February 1978.

75. LaBarge, W. L., GENERAL AVIATION AIRPLANE STRUCTURAL

CRASHWORTHINESS PROGRAMMER'S MANUAL, Lockheed-California

Company; FAA-RD-78-120 (Revised), Systems Research and

Development Service, Federal Aviation Administration,

Washington, D. C., June 1979.

232

.. . .- ' I I I I mewI II-:- 
lob

The experimental validation of KRASH is summarized in Table 10.

Examples of helicopter structures modeled by KRASH are the

existing utility model shown in Figure 88 and a medium cargo model shown in Figure 89 (Reference 76). Also, demonstrating

the use of KRASH to model a structure in greater detail, Figure 90(a) illustrates a half-model of the cargo helicopter nose

section, where symmetry is utilized in modeling the structure

to the left of the aircraft mid-plane. Reference 77 describes

the drop test of a nose section from the cargo helicopter at an

impact velocity 33.3 ft/sec, as shown in Figure 90(b), and correlation of test data with KRASH predictions.

7.5.2 Program DYCAST

The computer program DYCAST (DYnamic Crash Analysis of STructures) is one module of the PLANS (Plastic and Large deflection ANalysis of Structures) system of nonlinear finite element

structural analysis computer codes. These programs have been

developed under contract to NASA Langley Research Center as

part of a joint NASA/FAA program in general aviation crashworthiness. The modules for static analysis of structures are

described in Reference 78.

As usual in finite element modeling, the structure is idealized

into natural structural components% stringers, frames, beams,

skin sheets, and bulkhead sheets, using the element library of

DYCAST. Some portions could be modeled crudely if in the judgment of the analyst detailed modeling was not necessary. The

element library consists of:

76. BadriNath, Y. V., SIMULATION, CORRELATION, AND ANALYSIS

OF THE STRUCTURAL RESPONSE OF A CH-47A TO CRASH IMPACT,

The Boeing Vertol Company; USARTL Techniual Report 78-24,

Applied Technology Laboratory, U.S. Army Research and

Technology Laboratories, Fort Eustis, Virginia, August

1978, AD A062643.

77. Tanner, A. E., Widmayer, E., STRUCTURAL CRASHWORTHINESS,

SIMULATION AND ANALYSIS, FINAL REPORT, The Boeing Vertol

Company; USARTL Technical Report 78-21, U. S. Army Research and Technology Laboratories (AVRADCOM), Fort Eustis,

Virginia, unpublished.

78. Pifko, A. B., Levine, H. S., and Armen, H., Jr., PLANS -.

A FINITE ELEMENT PROGRAM FOR NONLINEAR ANALYSIS OF STRUCTURES, VOLUME I, THEORETICAL VANUAL, NASA CR-2568, National

Aeronautics and Space Administration, Washington, D. C.,

November 1975.

233

.. "... . . .. ... ...
TABLE 10. PROGRAM KRASH EXPERIMENTAL VERIFICATION

Weight Impact velocities (ft/sec)

Aircraft (ib) Vertical Longitudinal Lateral

i. Rotary-wing, 8660 23 - 18.5 utility-type

2. Single-engine 2400 46 70 -

high-wing

3. Single-engine 2400 22 71.3

high-wing

4. Single-engine 2400 49 70

high-wing

5. Single-engine 2400 43 69.5 -

high-wing

6. Twin-engine 545 27.5

low-wing

substructure

*Test performed on soill all other tests on rigid surface.

9 Stringers - Two types of axial force members are

available, constant or linearly varying between

nodes.

0 Beams - Ten different cross sections are currently available for the 12-degree-of-freedom beam element.

Axial force, shear forces, and torque are uniform

along the length, with a linear variation of bending

moments.

0 Membranes - Triangular membrane elements are available with either constant strain or linear strain.

Hybrid elements permit mixed strain conditions, i.e.,

constant strain along an edge with linear strain

along the 3ther two sides.

* Springs - An axial force stringer with the spring

constants specified in tabular form has recently been added to the element library. It can be used

to simulate structural sections with known axial

234
25 masses

32 members

'F

Xj M

i'gr 8. Mde feisigutltHelicopter.cner/

oadgrveri l -eletotyavost epee a

iy i) j

dynmialy laddtrutu-s matria an emtrc h

Sij *_.----Beam ij

! ,,'" •" X11 X Beam ij mj

nonlinear mtil evo ebzki

stra Figure 88. Model of exis tic-pelity helicopter. i,' (From Reference 65)

•! if!load-versuas-def lection behavior; to represent an

elsi-energy-absorbing devicel or, as a gap element with

Th ureteezero stiffness over a certain range of deflection

.and nonzero thereafter.

D¥CAST accounts for two types of nonlinearities which occur in

dynamically loaded structuresi material and geometric. The

nonlinear material behavior, exhibited by metals yielding plastically, enters the simulation process through the stressstrain curves input by the analyst. Three types of stress-

!•,strain curves are permittede elastic-perfectly plastic#

elastic-linearly hardening, and elastic-nonlinearly hardening. The current element stiffnesses during a multi-stop analysis

I 235
Shell and pylon

Masses • 36

Beams • 62 34

Axial elements ,*

Tension only : 20 1,

Aft L.G. oleo of2 00, 1 3.

Landing gears External springs: 4 2§

Underfloor structure A

(Crush area) 2

External springs: 12

15

,,gre8. Model of exstngmeiu cag heiotr

/~ ~ . .0 ,8 •

00 00'

(From Reference 76)

are determined using the tangent moduli corresponding to the

current stress and strain, generalized for multiaxial states

at various points in the element.

00•

6 Geometric nozalinearities, due to large deformations of the ;,

structure, change the effective stiffness of the structure £nd .'

are treated in DYCAST by an incremental convect~ad coordinate

approach. After each time increment the structure is reformed , with straight elements between the displaced nodesi the previ- '

ous values of accumulated strain, stress, and internal loads

are carried forward as initial states in the reformed elements.

The incremental equations of motion for the system written in

matrix form for an increment of time are:x:' i

[m] •Atii + [K] l!•ul - IAP} + S

236

~% % .,,0 6 .. '
0)0

Ur 0

:ro

'0 m

4J 9)

u 0

14 9 0

400

hfi

44.

I~ ~ ~ l I.* I' 0 
where [m] is the consistent mass matrix, [K] the stiffness matrix, A and JA U1 the displacement and acceleration ýtcrements, A PI the incremental external load vector, and R the •,, vector of equilibrium corrections. Both material and geometric " 1:! nonlinear effects enter the system of incremental equations • ~through the stiffness matrix [K]. This system of equations in

&il; • integrated numerically to obtain the response of thq structural

"•'•" model,

Currently, both explicit and implicit algorithms are implemented in DYCAST. Central differences and modified Adams

predictor-corrector methods are the explicit algorithmsi and,

Newmark-0 and Wilson-6 methods are the implicit algorithms. All methods but central difference. permit a variable time step throughout the analysis.

DYCAST analysis results are printed histories of:

* Nodal displacements, velocities, and accelerations.

e Nodal forces and moments.

, Element strains and stresses through the crosssections.

Postprocepoing of saved data permits plotting of:

- Time histories of displacements, velocities, and acoelerations.

e Views of the deformed structure at any time and from

any viewing angle.

Further details of DYCAST can be found in References 69 and 78;

and, as mentioned above, Reference 55 presents a comparison of

DYCAST and KRASH modeling of the same section of fuselage structure.

7.5.3 Program WRECKER

Under the sponsorship of the National Highway Traffic Safety

Admini.straLion, a finite element program called WRECKER has, been developed by the IXT Research Institute for use in dy- namic analysis of vehicle structures, including sheet metal, in a crash environment. Tha program consists of the following

features:

238
. Large displacement, nonlinear static and dynamic, and

elastic and plastic including strain-rate effects.

e Plate, spring, rigid link, and three-dimensional

beam elements with a variety of beam end conditions.

* Options of using either the oxplicit or implicit time

integration procedures.

* Options of specifying stress, mass and center of

gravity, and energy oi-tput.

7.5.3.1 Solution Procedure: Two analysis procedures are used:

the explicit or implicit, timewise numerical integration of the

equations of motion for the node points (three translations and

three rotations per node). The relative merits of implicit and

explicit procedures and the importance of the mass formulation

has been studied by Krieg and Key (Reference 79). The original

version of WRECKER, documented in Reference 80, utilized only

an explicit solution procedure, but WRECKER II, described in

Reference 70, has both schemes.

An explicit procedure is exploited throughout the analysis

by using lumped nodal masses and by calculating the internal

forces at the nodes from direct integration of the element

stress fields without reference to element or assembled stiffness matrices for the structure. The choice of an explicit

integration procedure and a direct calculation of nodal forces

resulto in a program with minimum (but still substantial) computer storage requirements. This, in turn, is equivalent to a

capability of processing reasonably detailed and extensive

structural models with relative ease. An early capability for

handling relatively large problems is considered essential in

developing realistic simulations of actual crash events because of the complex geometry and construction found in real

vehicles.

The principal shortcomings of this integration procedure are

the fairly extensive run times and the difficu]ty of obtaining results for static or quasi-static situations in which the

79. Krieg, R. D., and Key, S. W., TRANSIENT SHELL RESPONSE BY

NUMERICAL TIME INTEGRATION, International Journal for

Numerical Methods in Engineering, Volume 17, 1973.

80. Welch, R. E., Bruce, R. W., and Belytschko, T., FINITE

ELEMENT ANALYSIS OF AUTOMOTIVE STRUCTURES UNDER CRASH

LOADINGS, IIT Research Institute; Report No. DOT-HS-801-

847, U. S. Department of Transportation, Washington, D. C.,

March 1976.

239
loading and response vary slowly with time. With regard to the

latter observation, it was noted that, although the vehicle

crash event is typically a high-speed phenomenon, low-speed or

quasi-static crush testing provides valuable data regarding the

performance of vehicle components and is an integral part of

crashworthiness studies.

The implicit integration procedure requires the formation of

the tangent stiffness matrix and the matrix inversion for the

solution' of incremental displacements at each time step. The

choice of this procedure results in a program requiring considerably greater computer storage. However, this procedure is capable of carrying out dynamic analyses at substantially

greater time steps than are admitted in the explicit version

and permits simulation of quasi-static crush phenomena that

involve slowly varying loadings and structural response.

7.5.3.2 Element Formulation: The treatment of large displacements and rotations employs a decomposition of the element displacement field into a rigid body rotation and translation

associated with the local coordinate system attached to and

moving with the element, and a remaining displacement field

which describes the deformation of the element relative to the

current position of the element axes. This transformation, in

effect, removes the average rigid body rotation of the element

and allows the use of small or moderate deflection element formulations in the calculation of element forces. In this manner,

extremely large rotations and deflections can be accommodated

by the analysis with accuracy depending primarily on the size

of the elements relative to the curvature of the structure.

Although a formal convergence theorem is lacking, the decomposition does hold for infinitesimal regions, and numerical

studies show excellent agreement with classical solutions.

The computer program at present includes low-order triangular

plate elements, three-dimensional beam elements, and spring

elements. A triangular membrane-hinge line element is also

available but is not presently compatible with the beam formulation. Hinges and sliding joints, or a combination of both,

are available in beam elements.

7.5.3.3 Material Properties: The computer program currently

uses simple elastic-plastic stress-strain lawsi a uniaxial relation for beam and spring elements, and a biaxial strain hardening Von Mises model for plates. Element forces and bending

moment for given strain fields are calculated by piecewise linear numerical integration of the stresses at selected points in

the cross section. Options of an explicit moment-curvature relationship and the strain rate effect are also provided. The

program is designed to accommodate other constitutive relations.

240

-i )
In fact, linear relationships are provided at this level which

result in more efficient computations for this class of problems.

7.5.3.4 Program Validation: The explicit version of WRECKER

II was validated through a series of test problems involving

beams, plates, and shells subjected to various types of loading. Also, as described in Reference 80, the front end of an

automobile was modeled in simulation of 30- and 44-mi/h barrier

crash tests. Reference 70 describes validation of the implicit

procedure by comparison with previously obtained explicit solutions or, in the case of static loading# with classical linear

and nonlinear solutions.

7.5.3.5 Computer Program Description: The program contains

spring, beam, and triangular-plate finite elements and treats

the large deflection phenomena by decomposing the element deformations into rigid body and deformation components before

computing element forces. The numerical integration technique

employed is a Newmark-0 method.

The program consists of approximately 4500 cards and normally

executes in 76 K words of core storage on the UNIVAC 1108 coinputer under the EXEC-8, version 27 operating system. A version of the program has been prepared and executed in 300 K

bytes on an IBM 370/168 system under the H compiler with no

optimization. The capacity of the program is approximately

150 nodes, 150 elements, 100 displacement boundary conditions,

5 different sets of material properties, and 10 distinct beam

cross sections. The solution procedure is completely in core

and makes use of a peripheral device for temporary storage

during execution. A restart capability and output data storage are provided via optional external files.

7.6 POTENTIAL SOURCES OF BASIC STRUCTURAL DATA FOR

CRASHWORTHINESS ANALYSIS

A major obetacle to analyzing structural crashworthiness is

the difficulty of obtaininq adequate data concerning the dynamic failure mechanisms of structural assemblies and elements.

Such data are necessary for the analysis techniques describel

earlier, with the possible exception of the finite A1ivlent

methzids. Data useful for structural analysis can bo obtained

from the several sources discussed within the following paragraphs.

241.
7.6.1 Estimates

An estimate of the dynamic capability of a structural element

or system can be obtained by using standard static failure criteria and multiplying the result by a dynamic correction factor. This method has limited usage since the factor usually specified for the overall failure sequence is representative

of the energy-absorption variation between static and dynamic

failure modes.

Consequently, a time history of element failure is not usually considered; but the net effect, in terms of energy absorption, is used to dbtermine the gross capability of the structure considered. Such analyses are useful in defining the gross crashworthiness contributions of structural layout prior to the use

of more complex analysis techniques.

Care must be exercised when using such a technique since structural elements exhibit differing failure modes, and these modes

may vary depending on the rate of load application. Additionally, the total load-carrying capability of a member also will

be dependent on the rate of load application.

7.6.2 Aircraft Accident Data

Aircraft accidents can provide useful data concerning structural response if the wreckage can be examined in detail before

it is disturbed from its postimpact location. Examination of

the wreckage can yield data concerning the failure modes and

sequences; the effectiveness of energy-absorbing features, if

any; and the occupant environment experienced in terms of

space, extremity flailing, acceleration levels, and injury

causes.

Such a relatively subjective review also requires other data

to enable the investigator to determine a complete crash scenario. Some of these data can be determined from the crash site

itself, particularly terrain hardness, ground obstacles, impact attitude, mass item breakaway, flammable fluid containment,

egress potential, etc. However, more important parameters that

are harder to obtain include the impact velocity components and the acceleration environment in the occupied areas. In

fact, these data probably will not be available in a recorded

format unless the crashed aircraft happened to be a test vehicle with on-board instrumentation. As a result, the velocity

components at impact are often estimated using the gross deformations oZ the structure and the motion of the aircraft relative to the ground. Simple energy and momentum techniques are

242
•:•'•i~iused to do this; and although the resultant velocity and accel- : •i , eration estimates are the best available, such an analysis can-

•:i'ii•'not apportion the energy absorption through the aircraft, i.e.,

•, in the landing gear, structure, seats, etc.

•: Reference 65 describes an examination of 3,657 U. S. Army

•:• rotary-wing aircraft accidents that occurred between 1967 and

'•ii:•:1971 and a detailed review of 32 of these. References 81 and

•.. 82 describe investigation of general aviation accidents and the

i~i :i•Iinformation that can be obtained.

•'• A major contribution to the quest for better aircraft crash-

, .i worthiness made by crash analyses is the delineation of the

.• good and bad features of a particular design. By compiling ac-

*' cident data, qualitative assessment of the desirable and unde-

;:, sirable features exhibited by various aircraft types can be

made. Once a reasonable data base is compiled, the desirable

features can be pinpointed and included in future designs or

redesigns. Undesirable features can be suitably advertised to

ensure their use is not continued.

•. 7.6.3 Controlled Crash Testing

S....Controlled testing has been done in the past on a variety of

aircraft to demonstrate the capabilities of structure, landing gears, fuel systems, etc. in typical crash environments.

'•'81. Snyder, R. G., CRASHWORTHINESS INVESTIGATION OF GENERAL '• AVIATION ACCIDENTS, SAE Paper No. 750537, presented at

•I " Busi~ness Aircraft Meeting, Society of Automotive Engineers, Inc., Wichita, Kansas, 8-11 April 1975.

"'• 82. Snyder, R. G., GENERAL AVIATION CRASH SURVIVABILITY, Uni--

•i versity of Michigani SAE Paper No. 780017, presented at i Congress and Exposition, Society of Automotive Engineers, ' Inc., Detroit, Michigan, 27 February - 3 March 1978.

24

:0;
(See, for example, References 42, 52, and 83 through 85). Initially, such tests were often performed using powered aircraft

that were remotely controlled or prealigned for impacts into

selected terrain conditions. Useful data were obtained from

many of these tests, although some resulted in postimpact fires

and information loss.

A more recent approach for vertical impacts has been to drop

test specimens from a fixed site or a moving carrier., Structural assemblies, small aircraft, and helicopters have been

tested in this manner. Representative impact conditions and

velocities can be achieved by adjusting the vertical drop

height and/or the longitudinal velocity of the carrier as well

as roll and yaw attitudes.

The largest facility used for full-scale crash testing of light

aircraft and helicopters is the Impact Dynamics Research Facility at NASA Langley Research Center (Reference 86). Here aircraft are allowed to swing on cables that are preset to determine the overall impact attitude. Velocity components are

controlled by varying the drop height, cable length, and cable

anchor locations. Immediately prior to impact, the aircraft

is released by pyrotechnic means. The ground impact and subsequent motions are then completely unrestrained,

83. Alfaro-Bou, E., and Vaughan, V. L., Jr., LIGHT AIRPLANE

CRASH TESTS AT IMPACT VELOCITIES OF 13 AND 27 M/SEC, NASA

Technical Paper 1042, NASA Langley Research Center, National Aeronautics and Space Administration, Washington,

D. C., November 1977.

84. Castle, C. B., and Alfaro-Bou, E., LIGHT AIRPLANE CRASH

TESTS AT THREE ROLL ANGLES, NASA Technical Paper 1477,

National Aeronautics and Space Administration, Washington,

D. C., October 1979.

85. Haley, J. C., Turnbow, J. W., and Walhout, G. J., FLOOR

ACCELERATIONS AND PASSENGER INJURIES IN TRANSPORT AIRCRAFT ACCIDENTS, Aviation Safety Engineering and Research

(AvSER), Division of Flight Safety Foundation, Inc.1

USAAVLABS Technical Report 67-16, U. S. Army Aviation Materiel Laboratories, Fort Eustis, Virginia, May 1967,

AD 815877L.

86. Vaughan, V. L., and Alfaro-Bou, E., IMPACT DYNAMICS RESEARCH FACILITY FOR FULL-SCALE AIRCRAFT CRASH TESTING,

NASA Technical Note D-8179, National Aeronautics and Space

Administration, Washington, D. C., April 1976.

244

, I II I
The primary advantage of testing full-scale aircraft is that

there is no need to interpret the data or attempt to extrap- olate the results to other structural formats. All data such

as velocities, attitudes, accelerations, and structural strains

are measured directly as functions of time from impact. In

addition, high-speed cameras can record displacements from various locations inside and outside the aircraft in order to provide visual time histories of structural and occupant response during the crash sequence. Postorash review of damage also

provides a direct indication of the aircraft's performance with

respect to occupied volume penetration, seat and landing gear

performance, large mass item retention, and flammable fluid

containment.

Instrumented anthropomorphic dummies positioned and restrained

in seats with typical restraint systems are used to investigate the potential for occupant survival. Seat occupant motions

are photographed using high-speed cameras, and the occupant's potential for impacting the aircraft interior is assessed.

The above description of full-scale testing may imply that this

is the only technique worth using to attain realistic conditions. However, it must be emphasized that such tests are expensive to run, and aircraft, especially new design prototypes,

are difficult to obtain. In addition, such tests require careful planning with a redundancy of recording equipment because

of the probability that some instrumentation channels may fail

to function.

7.6.4 Scale Model Testing

Scale model testing has been used extensively when investigating the aerodynamic characteristics of aircraft, bridges, buildings, etc. Scale model testing for structural strength and deflection verification also has been used where material

sizes allow. For evaluating crashworthiness, however, scale

model testing becomes a more difficult problem, especially when

severe plastic deformation and element rupture occur.

Scale modeling of major structural members may provide data

that can be used for crashworthiness studies; however, when

semimonocoque construction is considered, stringers and skin

are often made from relatively thin sheet material, measuring

from approximately 0.015 to 0.06 in. Depending on the structure being modeled, certain nondimensional parameters must be

satisfied for both the model and the aircraft. Examples of

these are:

xi * vt

L-- E', 6t,-Z

245

....................... .. .......
where X spatial coordinate

L - characteristic length

w strain

- stress

E - Young's Modulus

* strain rate

t t time

v W velocity

Some of these parameters involve the thickness of the miaterial,

and, for example, scaling 0.015 for a one-tenth scale poses

major problems in the manufacture, handling, and tolerance effects when using 0.0015-in. shim material.

Reference 87 presents a discussion of scale modeling techniques

applied to structural crashworthiness. Practical considerations in geometric scaling are discussed and illustrated using

barrier tests of two different automobile front-end structures

and an impulsively loaded section of semimonocoque cylinder

similar to an aircraft fuselage.

A conclusion of that study was that for prototype structures

in the 1,000-to-10,000-lb weight range, and for scale factors

of from 3 to 8, a model test can be performed at less than half

* the cost of a corresponding full-scale test, depending on the

scale factor, as demonstrated by Figure 91. It was concluded

that scale model tests can meet the same objectives and could

therefore replace many full-scale tests in a crashworthiness

research and development program. Model tests are particularly

useful for screening new concepts, for performing parametric

experimental studies, and for the initial optimization of a

given concept. Full-scale tests are still required for proving

the concept and for making detailed measurements such as measurements of occupant response.

87. HIolmes, B. S., and Colton, J. D.# APPLICATION OF SCALE

MODELING TECHNIQUES TO CRASHWORTHINESS RESEARCH, in Aircraft Crashworthinessr K. Saczalski, et al., eds., UnT---

versify Press of Virginia, Charlottesville, Virginia,

1975, pp. 561-582.

246
Full scale ; i"¢ .• iT -1- I ! I -T-T-- .1 . -.

,,'! '•1/2 •i 3 scale

S1/3 14J

4-i~ 1/4

So 1/8

0.1 K I 1LLLLI

1 2 4 6 8 10

Scale factor

Figure 91. Cost ratio (scale model test cost/full-scale

test cost) versus scale factor for structures

"weighing less than 10,000 lb. (From Refer-

* ,ence 88)

Also, in another recent study 1/3-scale models of various configurations of stiffened sheet panels, like that shown in Figure 92 (Reference 88), were tested in order to determine the

influence of the following parameters:

e Variation of the stiffener pitch (d) with respect

to the height (h).

e Variation of the stiffener section (S) with respect

"to the panel sheet section.

* Variation of the lightening hole diameters (0) with

respect to the stiffener pitch.

88. Mens, J., HELICOPTER BEHAV1OUR IN CRASH CONDITIONS, Aerospatiale, Helicopter Division; Paper 66 presented at

Fourth European Rotorcraft and Powered Lift Aircraft Forum,

Stressa, Italy, 13-15 September 1978.

(1~. 247 14
100'

h

Figure 92. Stiffened panel specimen. (From Reference 88)

* Type of stiffener (Z sections, angles, stiffening beads).

* Load influence (distributed load, load concentrated

at the ends).

* Bottom shape influence (flat bottom, curved bottom).

Comparison of model test results with those of full-scale tests

showed that the models exhibited the same failure modes and

predictable failure loads.

The techniques of dynamic testing become more involved as the

required quantity of data increases. For instance, if the final

deformation shape is the only result needed, the instrumentation and data handling requirements can be minimized. Since a

test is of short duration, even for full-scale specimens, specialized equipment is required if data requirements are expanded to include time histories of structural response, acceleration, stress, etc. This equipment must have rapid response

characteristics and maintain high fidelity measuring capability

when subjected to high G levels.

248 4

..... ....
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•"• 249 4
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250

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REFERENCES (Continued)

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251

*] ,
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252

*i: - '*' 4'> .l>.4.$4I'.,
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254 )

..................... ~~~ ".,....... ................... . ~ . . .
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t,! ,..I- t I | e e e • :i".r'
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256

4!
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74. Wittlin, G., GENERAL AVIATION AIRPLANE STRUCTURAL CRASHWORTHINESS USER'S MANUAL, Volume III, RELATED DESIGN INFORMATION, Lockheed-California Company; FAA-RD-.77-189-III,

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76. BadriNath, Y. V., SIMULATION, CORRELATION, AND ANALYSIS

OF THE STRUCTURAL RESPONSE OF A CH-47A TO CRASH IMPACT, The Boeing Vertol Company; USARTL Technical Report 78-24,

Applied Technology Laboratory, U.S. Army Research and

Technology Laboratories (AVRADCOM), Fort Eustis, Virginia, August 1978, AD A062643.

77L Tanner, A. E., Widmayer, E., STRUCTURAL CRASHWORTHINESS,

SIMULATION AND ANALYSIS, FINAL REPORT, The Boeing Vertol

Company; USARTL Technical Report 78-21, U. S. Army Research and Technology Laboratories (AVRADCOM), Fort Eustis,

Virginia, unpublished.

78. Pifko, A. B., Levine, H. S., and Armen, H., Jr., PLANS -

A FINITE ELEMENT PROGRAM FOR NONLINEAR ANALYSIS OF STRUCTURES, Volume I, THEORETICAL MANUAL, NASA CR-2568, National Aeronautics and Space Administration, Washington, D. C., November 1975.

79. Krieg, R. D., and Key, S. W., TRANSIENT SHELL RESPONSE BY

NUMERICAL TIME INTEGRATION, International Journal for

Numerical Methods in Engineering, Volume 17, 1973.

80. Welch, R. E., Bruce# R. W., and Belytachko, T., FINITE

ELEMENT ANALYSIS OF AUTOMOTIVE STRUCTURES UNDER CRASH

LOADINGS, IIT Research Institute; Report No. DOT-HS-801-

847, U. S. Department of Transportation, Washington, D. C.,

March 1976.

* 81. Snyder, R. G., CRASHWORTHINESS INVESTIGATION OF GENERAL

AVIATION ACCIDENTS, SAE Paper No. 750537, presented at

Business Aircraft Meeting, Society of Automotive Engineers, Inc., Wichita, Kansas, 8-11 April 1975.

82. Snyder, R. G., GENERAL AVIATION CRASH SURVIVABILITY, University of Michigani SAE Paper No. 780017, presented at

Congress and Exposition, Society of Automotive Engineers, Inc., Detroit, Michigan, 27 February - 3 March 1978.

fi I257

*.' -,, - I i. '*1'
REFERENCES (Continued)

83. Alfaro-Bou, E., and Vaughan, Vs L., Jr., LIGHT AIRPLANE

CRASH TESTS AT IMPACT VELOCITIES OF 13 AND 27 M/SEC, NASA

Technical Paper 1042, NASA Langley Research Center, National Aeronautics and Space Administration, Washington,

D. C., November 1977.

84. Castle, C. B., and Alfaro-Bou, E., LIGHT AIRPLANE CRASH

TESTS AT THREE ROLL ANGLES, NASA Technical Paper 1477,

National Aeronautics and Space Administration, Washington,

D. C., October 1979.

85. Haley, J. C., Turnbow, J. W., and Walhout, G. J., FLOOR

ACCELERATIONS AND PASSENGER INJURIES IN TRANSPORT AIRCRAFT ACCIDENTS, Aviation Safety Engineering and Research

(AvSER), Division of Flight Safety Foundation, Inc.1

USAAVLABS Technical Report 67-16, U. S. Army Aviation Materiel Laboratories, Fort Eustis, Virginia, May 1967,

AD 815877L.

86. Vaughan, V. L., and Alfaro-Bou, E., IMPACT DYNAMICS RESEARCH FACILITY FOR FULL-SCALE AIRCRAFT CRASH TESTING,

NASA Technical Note D-8179, National Aeronautics and

Space Administration, Washington, D. C., April 1976.

87. Holmes, B. S., and Colton, J. D., APPLICATION OF SCALE

MODELING TECHNIQUES TO CRASHWORTHINESS RESEARCH, iA Aircraft Crashworthiness, K. Saozalski, et al., eds.,;u n-nversity Press of Virginia, Charlottesville, Virginia,

1975, pp. 561-582.

88. Mens, J., HELICOPTER BEHAVIOUR IN CRASH CONDITIONS, Aerospatiale, Helicopter Divisionj Paper 66 presented at

Fourth European Rotorcraft and Powered Lift Aircraft

Forum, Stressa, Italy, 13-15 September 1978.

258
,

BIBLIOGRAPHY

Alfaro-bou, E., et al., SIMULATIONS OF AIRCRAFT CRASH AND ITS

VALIDATION, in Aircraft Crashworthiness, K. Saczalski, et al.,

ads., University Press of Virginia, Charlottesville, Virginia, 1975, pp. 485-498.

Armen, H., Jr., Pifko, A., and Levine, H., NONLINEAR FINITE

ELEMENT TECHNIQUES FOR AIRCRAFT CRASH ANALYSIS, in Aircraft

Crashworthiness, K. Saczalski, et al., eds., University Press

of Virginia, C7harlottesville, Virginia, 1975', pp. 517-548.

Avery, J. P., Ph.D., CARGO RESTRAINT CONCEPTS FOR CRASH RESISTANCE, Aviation Safety Engineering and Research (AvSER), Division of Flight Safety Foundation, Inc.1 USAAMRDL Technical Report 65-30, U. S. Army Aviation Materiel Laboratories, Fort

Eustis, Virginia, June 1965, AD 618493.

Belytschko, T. B., Welch, R. E., and Bruce, R. W., NUMERICAL

ANALYSIS OF SHEET METAL IN CRASH LOADING, in Aircraft Crashworthiness, K. Saczalski, et al., eds., University Press of

Virginia, Charlottesville, Virginia, 1975, pp. 549-560.

Brownfiald, H. A., and Rogers, D. 0., ANALYSIS OF 30 MPH FRONTAL BARRIWR UTILIZING HALF-SCALE METAL MODELS, SAE Paper No.

780366, presented at 1978 Congress and Exposition, Society of

Automotive Engineers, Inc., Detroit, Michigan, 27 February3 March 1978.

Carnell, B. L., REVIEW OF NAVY HELICOPTER CRASHWORTHINESS, Document Number SER-50941, Sikorsky Aircraft, Division of United

Technology Corporation, Stratford, Connecticut, September 1976.

Chernoff, M., ANALYSIS AND DESIGN OF SKID GEARS FOR LEVEL LANDING, Journal of the American Helicopter Society, Vol. 7, No. 1,

January 1961, pp. 33-39.

Davies, R. G., and Magee, C. L., THE EFFECT OF STRAIN RATE UPON

THE BENDING BEHAVIOR OF MATERIALS, Ford Motor Companyl Paper

No. 76-Mat.-FF, Journal of Engineering Materials and Technology, "Vol. 99, No. 1, January 1977, pp. 47-51.

Gamon, M. A., and Wittlin, G., ANALYTICAL TECHNIQUES FOR PREDICTING VEHICLE CRASH RESPONSE, in Aircraft Crashworthiness,

K. Saczalski, et al., eds., University Press of Virginia, Charlottesville, Virginia, 1975, pp. 605-622.

Fit

259

t.. 

. . 4
BIBLIOGRAPHY (Continued)

Hone, H. T., Huebner, W. E., and Baxter, D. J., DEVELOPMENT OF

CARGO SLINGS WITH NONDESTRUCTIVE CHECKOUT SYSTEMS, Sikorsky Aircraft, Division of United Technology Corporation; USAAMRDL

Technical Report 73-106, Eustis Directorate, U. S. Army Air

Mobility Research and Development Laboratory, Fort Eustis,

Virginia, February 1974, AD 777497.

Heubner, W. E., DESIGN GUIDE FOR LOAD SUSPENSION POINTS, SLINGS,

AND AIRCRAFPl HARD POINTS, Sikorsky Aircraft, Division of United

Aircraft Corporation; USAAMRDL Technical Report 72-36, Eustis

Directorate, U. S. Army Air Mobility Research and Development Laboratory, Fort Eustis, Virginia, July 1972, AD 747814.

Kruse, G. S., AN AUTOMATED PROCEDURE FOR PRELIMINARY DESIGN OF

PRIMARY STRUCTURE FOR TIRANSPORT AIRCRAFT, Convair, Division of

General Dynamics Corporation; Paper No. 76-WA/Aero-9, American

Society of Mechanical Engineers, New York, December 1976.

McIvor, I. K., A SIMULATION PROGRAM FOR LARGE DYNAMIC DEFORMATION OF VEHICLES, SAE Paper No. 770054, presented at 1977 International Automotive Engineering Congress and Exposition,

Society of Automotive Engineers, Inc., Detroit, Michigan, 28 Febr-.y-4 March 1977.

Melosh, R. J., CRASHWORTHINESS ENGINEERING OF AUTOMOBILES AND

AIRCPAFT: PROGRESS AND PROMISE, Journal of Aircraft, Vol. 14.

No. 7, July 1.977, pp 693-698.

Melosh, R. J., and Kamat, M. P., COMPUTER SIMULATION OF LIGHT

AIRCRAFT CRASH, Journal of Aircraft, Vol. 14, No. 10, October

1977, pp 1009-1014.

Ni, C-M., and Fine, D. S., PREDICTING CRUSH RESPONSE OF AUTOMOTIVE STRJCTURAL COMPONENTS, SAE Paper No. 780671, presented

at Passenger Car Meeting, Society of Automotive Engineers, Inc., Troy, Michigan, 5-9 June 1978.

Park, K. C., MODELING AND ANALYSIS TECHNIQUES FOR VEHICLE

CRASH SIMULATION, in Aircraft Crashworthiness, K. Saczalski,

et al., eds., University Press of Virginia, Charlottesville,

Virginia, 1975, pp 499-516.

Pifko, A., et al., PLANS - A FINITU ELEMENT PROGRAM FOR NONLINEAR ANALI!SIS OF STRUCTURES, Paper No. 74-WA/PVP-6, American

Society of Mechanical Engineers, New York, November 1974.

260

'. 1 ,

' I • Ii i I I ' , 
BIBLIOGRAPHY (Continued)

.• Robinson, D. C., CRUSH CHARACTERISTICS OF AUTOMOBILE STRUCTURAL

COMPONENTS, National Bureau of Standards; Report No. COM-75-

10464, National Highway Traffic Safety Administration, U. S.

Department of Transportation, Washington, D. C., January 1975.

Sacizalski, K. J., and Pilkey, W. D., eds., MEASUREMENT AND PREDICTION OF STRUCTURAL AND BIODYNAMIC CRASH-IMPACT RESPONSE, The

American Society of Mechanical Engineers, New York, 1976.

Saczalski, K. J., and Pilkey, W. D., TECHNIQUES FOR PREDICTING

VEHICLE STRUCTURE CRASH-IMPACT RESPONSE, in Aircraft Crashworthiness, K. Saczalski, et al., eds., University Press of

Virginia, Charlottesville, Virginia, 1975, pp. 467-484.

Simonian, S., and Hart, G. C., IDENTIFICATION OF STRUCTURAL

COMPONENT FAILURES UNDER DYNAMIC LOADING, SAE Paper No. 770958,

presented at Aerospace Meeting, Society of Automotive Engineers,

Inc., Los Angeles, California, November 1977.

Skogh, J., and Stern, P., POSTBUCKLING BEHAVIOR OF A SECTION

REPRESENTATIVE OF THE B-1 AFT INTERMEDIATE FUSELAGE, LockheedCalifornia Company; AFFDL Technical Report 73-63, Air Force

Flight Dynamics Laboratory, Wright-Patterson Air Force Base,

Ohio, May 1973.

Smith, H. G., and McDermott, J. M., DESIGNING FOR CRASHWORTHINESS AND SURVIVABILITY, Paper presented at 24th Annual National

Forum Proceedings, National Helicopter Society, Inc., Washingtion, D. C., May 1968.

Tani, M., and Funahashi, A., ENERGY ABSORPTION BY THE PLASTIC

DEFORMATION OF BODY STRUCTURAL MEMBERS, SAE Paper No. 780368,

presented at 1978 Congress and Exposition, Society of 4utomotive Engineers, Inc., Detroit, Michigan, 27 February-3 March

1978.

Townley, G. E., and Kiahs, J. W., DYNAMIC SIMULATION OF AN

AUTOMOBILE BODY UTILIZING FINITE ELEMENT AND MODAL SYNTHESIS

TECHNIQUES, SAE Paper No. 780364, presented at 1978 Congress

and Exposition, Society of Automotive Engineers, Inc., Detroit,

Michigan, 27 February-3 March 1978.

Weingarten, J. L., Mayrand, C. V., and Muller, G. E., AIR CARGO

RESTRAINT CRITERIA, ASD-Technical Report 73-17, Aeronautical

Systems Division, Wright-Patterson Air Force Base, Ohio, April

1973.

261
BIBLIOGRAPHY (Continuod)

Widmayer, Eo, APPLICATION OF KRASH TO THE SOAC ACCIDENT, in

Aircraft Crashworthinesu, K. Saczalski, et al., eds., Univercity Press of ViTrginia, Charlottesville, Virginia, 1975,

pp. 583-604.

Yeung, K. S., and Hollowell, T°, LARGE DISPLACEMENT, NONLINEAR

STATIC AND DYNAMIC ANALYSIS OF AUTOMOBILE SHEET METAL STRUCTURE,

SAS Paper No. 780367, presented at 1978 Congress and Exposition,

Society of Automotive Engineers, Inc., Detroit, Michigan, March

1978.

262

,* l L,
RiI

INDEX

Acceleration

definition 27, 171

N' : impact 46

Accident data 46, 242

Aitward load, definition 30

Aircraft

"coordinate systems 25, 57 types 37, 48

Analytical methods •i :"classes of 226

landing gear 195

Ancillary equipment

categories 64

retention requiremonts 67, 154

Anthropomorphic dummy 33

Attitude

definition 25

impact 48, 50

Bending, fuselage 129

Blade, rotor

damage 57, 147

impact protection 54, 139

penetration 50, 144

Bottoming definition 36

seats 74

"structure 179

Brittle fracture 78

Buckling

floor 53

fuselage 59, 78, 124

Cabin structure

collapse, prevention of 135

energy absorption 115

longitudinal impact 52, 130

rollover protection 143

vertical impact 135

Cargo categories 70, 159

Cargo restraint

cable requirements 168

k! design loads 161, 176

design requirements 69, 157

energy absorption 69, 161

floor requirements 165

' nets 69, 159, 167

strength requirements 161, 166

testing 71

263

L I ..................... ~ .- 

.

.

.

.

. .'
INDEX (Continued)

Castings, landing gear 66

Clearance for seat stroke 156, 194

Cockpit structure

longitudinal impact 52, 129

vertical impact 135

Combined load, definition 30

Composite materials

'considerations in use '81

cost effectiveness 19

energy absorption 82, 90

fuselage 88

joints 95

properties 82

rotor blades 147

thermal expansion 85

Coordinate system, aircraft 25

Corrosion 137

Crash environment 46

Crash testing 45, 69, 210, 243

Crashworthiness

criteria 22, 38, 72, 111

integration into design 21, 72

requirements 55

Crippling stress 214

Crush characteristics

estimation 203

testing 207

Design, preliminary 22

Downward load, definition 30

Ductility 77, 124

Dummy 33

DYCAST 229, 233

Dynamic overshoot

cargo restraint 161

definition 31

Earth scooping

deceleration in 52

design features to prevent 58, 123

fixed-wing aircraft 126

minimization of 116

Emergency equipment

retention 67

stowage location 68

Emergency exits

design loads 150

postcrash operation 50, 78, 143

264
INDEX (Continued)

Energy absorber

definition 35

devices 99

Energy absorption

ancillary equipment 67, 154

cargo restraint 69, 99, 161

composite materials 82, 88, 90

contributing processes 46, 51, 114, 123, 193

fuselage 112, 212

ground friction 194

ideal 178

landing gear 60, 66, 101, 109, 139

mass reduction 51, 119, 149, 193

nose plowing 119

plastic deformation 193

rotor blades 147

seats 68, 154, 194

skid gear 109

structural 212, 220'

underfloor 52, 60, 135

vertical impact 133

wing 149

Energy content of aircraft 189

Engine mounts 63, 150

Exits, location of 143, 151

Failure modes

controlling material selection 77

typical 78

Fasteners, failure of 78

Finite element analysis 228

Fixed-wing aircraft

Army, U. S. 38

earth scooping 126

impact conditions 50

nose plowing 52, 58

types 38

Flight path angle, definition 26

Floor

buckling 53

loads 59

structure 129

Forward load, definition 30

Fracture toughness 79

Frame-type models 227

Fuel cell

location of 151

protection of 54, 153

rupture by landing gear 65, 151 P support of 152

265

. • , . ... . .. .. ,, .. .. , . . . . ..... •
INDEX (Continued)

Fuel systems

protection of 54, 153

requirements for 151

testing 71

Fuselage

bending 129

earth scooping 116, 123

energy absorption 112

longitudinal impact 52, 57

penetration by landing gear 50, 54

shape 64, 79, 112

skin 124, 139

vertical impact 52

Ground friction 194

Helicopter

impact conditions 46, 50

types 37, 48

Human body coordinates 32

Human tolerance

definition 34

factors 48

Hybrid models 227

Hydraulics, land4hng gear 197

Impact angle, d finition 26

Impact conditi( 

s

helicopter F46, 50

landing geai 66

MIL-STD-1290(AV) 55

velocities: 46

Injuries 53

Interference, landing gear with seats 65

Joints

composite materials 95

controlled failure 57, 73, 76, 78

Kinetic energy

aircraft 51, 176, 189

definition 174

KRASH 229, 230

Landing gear

crashworthiness requirements 62

design problems 101

dynamic strut 11, 195

energy absorption 60, 66, 101, 139

fuel cell rupture 104

ground clearance 65

hydraulics 197

impact velocity 66

266 :, .iI
INDEX (Continued)

"interference with seats 65

location 65

longitudinal impact 103

MIL-STD-1290(AV) 64, 102

penetration of fuselage 54, 65, 104

response characteristics 104,

i•de load effects 106

skid-type 109, 201

soil characteristics for design 64

testing 70

"two-stage strut 106

"wheel-type 103, 195

Large mass items

breakaway 51, 111, 119

fuselage collapse 52

support 50, 54, 60, 63, 137, 150

Lateral impact

conditions depending on aircraft type 38, 60

fuselage crushing 53, 142

MIL-STD-1290(AV) 60, 224

Lateral load, definition 30

Life rafts 68

Limit load, definition 35

Load factor, definition 29

Load limiter. See Energy absorber.

Load paths, redundant 77, 96

Longitudinal impact

analysis 220

cabin structure 130

design requirements 52, 57

earth scooping 58, 123

landing gear 103

MIL-STD-1290(AV) 57

Major impact, definition 28

Margins of safety 72

Mass reduction 51, 119, 149

Materials

composite 81

energy absorption 76

requirements 74

spark generation 96

MIL-HDBK-5 76

MIL-HDBK-17 76

MIL-STD-1290(AV) 17, 21, 38

impact conditions 56

landing gear 64, 102

lateral impact 60, 224

267
INDEX (Continued)

longitudinal impact 57, 129, 220

requirements 56

rollover protection 60, 225

seats 154

vertical impact 59, 137

Model testing 245

Nets, cargo 69

Nose plowing

energy dissipation 119

minimization of 93

"Nose section

longitudinal impact 129

vertical impact 135

Palletized cargo 158

Pitch

definition 25

vertical impact 59, 66

Plastic deformation

energy absorption 193

joints 78

landing gear 66

relieving stress concentrations 73

skid gear 109, 201

Preliminary design 22

Properties, estimation 203

Protective shell 51, 53, 57, 72, 76

Rebound, definition 31

Retrofit 22, 47

Roll

definition 25

landing gear design 96

vertical impact 59, 66

Rollover protection cabin structure 143

cockpit structure 143

design requirements 60

fixed-wing aircraft 50

longitudinal impact 57

MIL-STD-1290(AV) 60

Rotor blades, See Blade.

Safety factors 72

Scale-model testing 245

Seat

attachment to structure 154, 157

clearance for stroke 155, 194

energy-absorbing 139, 154

MIL-STD-1290(AV) 68, 154

268
INDEX (Continued)

. stroking distance 155

testing 71

types 154

Shape, fuselage 64, 79, 112

"Side loads, landing gear 106

Simulation, objective of 74, 169

Skid gear 109, 201

Spark generation 77, 96

Specific energy absorbed, definition 36

Stability, affecting energy absorption 212

Static strength, definition 35

Stiffened panels 206, 247

Stopping distance 181

Strength properties 77

Structural analysis finite element 228

.semiempirical 203

Structural integrity

definitions 35

in rollover 58

Survivable accident, definition 32

Survival equipment 67

Systems analysis 21, 72

Tail wheel 65

Terrain angle, definition 26

Testing

component 207

full-scale 45, 69, 210, 243

scale model 245

Tiedown rings 147

Transmissibility, definition 32

Transmission, support 54, 63

Underfloor structure

beams 130

energy absorption 52, 60, 135

Upward load, definition 30

Variable-orifice valve 105, 197

Velocity

definition 170

impact 46

Velocity change 28

Vertical impact

analysis 210

cabin section 135

collapse 52

fuselage design 131

impact conditions 59

269

Ci

S~ .~i ~ 
INDEX (Continued)

landing gear 102

6noseection 135

Weight, aircraft 57

Wing design 63, 149

WRECKER 229, 238

Yaw, definition 25

270

el)

V•

I. L
DEPARTMENT OF THE ARMY

US ARMY RIESEARCH APPLIED AND TECHNOLOGY TECHNOLOGY LABORATORY LABORATORIES (AVRADCOM)

•' ~FORT EUSTIS. VIRGINIA 23604 "

READER'S SERVICE LETTER

(For use in submitting comments, recommendations, corrections, and

revisions for Aircraft Crash Survival Design Guide)

FROM:

I

TO: Director, Applied Technology Laboratory, US Army Research and Technology

Laboratories (AVRADCOM), ATTN DAVDL-ATL-ASV, Fort Eustis, Virginia

* 23604

REMARKS:

i

SIGNATURE: __________ DATE:__________

('I