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AIRCRAFT RASH,.yRIVA IDSI

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20. (Continued)

Volume I - Design Criteria and Checklists; Volume II - Aircraft Crash Environment and Human ToleranceVolume III- Aircraft Structural Crashworthiness,-

Volume IV - Aircraft Seats, Restraints, Litters, and Padding. K

Volume V - Aircraft Postcrash Survival,

This volume contains concise criteria drawn from Volumes II - V,

supplemented by checklists intended to assist designers in implementation of the criteria.

Aceession For

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I!

PREFACE

This report was prepared for the Safety and Survivability Technical 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 Sur1Va 1id6essiGUhiide; pif s-hie•....-

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. Ap- preciation is extended to the following organizations for providing accident case histories leading to the establishment of

the impact conditions in aircraft accidents:

* U. S. A.rmy Safety Center (USASC), Fort Rucker, Alabama.

"* Civil Aeronautics Board, Washington, D. C.

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

"* 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, Z. H. Robertson,

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

Volume I is a compilation of criteria and checklists for the

design of crashworthy aircraft. The criteria have been assembled in this one volume for the convenience of those involved

* in the design or evaluation of the overall aircraft and for

use as a concise criteria reference. Additional background

information is provided in Volumes II through V.

3

rkI
The design criteria contained in this volume are the result of

studies made and experience gained during design and manufacture of new, current Army aircraft.

4

* ; .i
TABLE OF CONTENTS

Section Title PageM R

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

LIST OF ILLUSTRATIONS .. ................ 10

LIST OF TABLES ............. .................... 14 1

INTRODUCTION..................................... . 15

CHAPTER 1. BACKGROUND DISCUSSION ....... .......... 18

CHAPTER 2. DEFINITIONS ....... ............... . 23

2.1 AIRCRAFT COORDINATE SYSTEMS AND

ATTITUDE PARAMETERS .............. .. 23

2.2 ACCELERATION-RELATED TERMS ....... 25

2.3 VELOCITY-RELATED TERMS . . . . . . ... 26

2.4 FORCE TERMS ....... ............... .. 27

2.5 DYNAMICS TERMS .. .............. 29 j

2.6 CRASH SURVIVABILITY TERMS .......... .. 30

2.7 OCCUPANT-RELATED TERMS ........... 30

2.8 SEATING GEOMETRY .............. .... 33

2.9 STRUCTURAL TERMS .... ........... 35

2.10 FUEL, OIL, AND HYDRAULIC SYSTEM TERMS. . 37

2.11 IGNITION SOURCE CONTROL TERMS ...... 37

2.12 INTERIOR MATERIALS SELECTION TERMS . . . 38

2.13 DITCH AND EMERGENCY ESCAPE TERMS . . . . 39

CHAPTER 3. AIRCRAFT CRASH ENVIRONMENT AND HUMAN

TOLERANCE ....... ................ 42

3.1 INTRODUCTION. . ........... . . 42

3.2 DESIGN CONDITIONS FOR IMPACT . . . 42 

3.2.1 General. ............... . 42

3.3 HUMAN TOLERANCE TO IMPACT. .............. 46

3.3.1 General. . . . . . .............. 46

3.3.2 Whole-Body Tolerance ... .......... 46

3.3.3 Head Injury Criteria ......... ... 48

3.3.4 Spinal Injury Criteria ........... 51

3.3.5 Leg Injury Criteria ..... ...... ... 52

3.3.6 Tolerance of Other Body Parts ...... 52

3.4 HUMAN BODY DIMENSIONS AND MASS

DISTRIBUTIONS ..... .............. .. 52

3.4.1 General. .............................. 52

3.4.2 Anthropometry....................... . 53

3.4.3 Inertial Properties.... ............ 54

3.5 CRASH TEST DUMMIES .... ........... 56

544 1
TABLE OF CONTENTS (CONTD)

Section Title page

CHAPTER 4. AIRFRAME STRUCTURAL CRASHWORTHINESS. . . 58

4.1 INTRODUCTION ..... ........... . . . 58

4.2 AIRFRAME CRASHWORTHINESS ........ 59

4.2.1 Longitudinal Impact. . . . 62

4.2.2 Vertical Impact. . . . . . 64

4.2.3 Lateral Impact . . . . . . 65

4.2.4 Rollover Impacts. . . . . 65

4.2.5 Wings and Empennage ............... 68

4.2.6 Engine/Transmission Mounts ... . 68

4.2.7 Shape of Fuselage Cross Section .... 69

4.2.8 Landing Gear. . .......... . 69

4.3 ANCILLARY EQUIPMENT RETENTION. . . . .. 71

4.3.1 Strength ................ . .. 73

4.3.2 Emergency and Survival Equipment

Stowage Location . . . . . . . . . . . . 73

4.3.3 Retention Devices Release for Emergency

and Survival Equipment . . .......... 73

4.4 INTERFACE OF OCCUPANT AND CARGO

RETENTION SYSTEMS WITH AIRFRAME. . . . . 73

4.5 CARGO RETENTION .... ......... . . . . 74

4.6 TESTING. . . . . . . . . ......... 78

4.6.1 Aircraft System Testing. . . . . . . . . 78

4.6.2 Landing Gear Crash Testing . . . . . . . 78

4.6.3 Cargo Restraint..... . . . . . . . . 78

4.6.4 Ancillary Equipment Retention. . . . . . 78

4.7 DESIGN CHECKLISTS ...... . . . .. . . . . 79

4.7.1 Landing Gear Design Checklist. . . . . . 79

4.7.2 Airframe Design Checklist. . . . . . . . 79

CHAPTER 5. AIRCRAFT SEATS, RESTRAINTS, LTTTERS,

AND PADDING ..... ............... ... 83

5.1 INTRODUCTION . . ............. 83

5.2 PRIMARY DESIGN CONSIDERATIONS. . . . 83

5.2.1 General .... ..... ...... . . . . . . 83

5.2.2 Design Conditions and Envelopes . . . 83

5.2.3 Structural Distortion. ......... . 84

5.3 DESIGN PRINCIPLES FOR SEATS AND

LITTERS . . . . . . 84

5.3.1 Seating System Orientation ...... 84

5.3.2 Litter Orientation ..... ......... . . 85

5.3.3 Materials. . . . . . . . . . . . ... 86

5.4 STRUCTURAL CONNECTIONS . . . . . . . 86

5.4.1 Bolted Connections ................ . 86

5.4.2 Riveted Connections. . . . . . . . . 87

6
r

TABLE OF CONTENTS (CONTD)

Section Title Pace

5.4.3 Welded Connections . . . . ........... 87

5.4.4 Seat Attachment . . . . . . ...... 87

5.4.5 Joint Deformation.. .......... . . . . 88

5.4.6 Material . . . . . . . . . . . . . . . 92

5.4.7 Restraint System Anchorage . . . . ... 93

5.5 ENERGY-ABSORBING DEVICES . . . . . . .. 94

5.6 SEAT CUSHIONS. . . . . . . . . . . . .. 95

5.6.1 General. . . . . . . . . . 95

5.6.2 Requirements ...... . . . 96

5.6.3 Energy-Absorbing Cushions. .1..... 96

5.6.4 Net-Type Cushions ..... ........... . 97

5.6.5 Seat Back Cushions . . . . ..... .......... 97

5.6.6 Headrests. . . . . . . . . . . 0 . . 97

5.7 DESIGN PRINCIPLES FOR PERSONNEL

RESTRAINT SYSTEMS .... ............ ... 97

5.7.1 General. . . . . . . . ..... . . .............. 97

5.7.2 Types of Systems . . . . . . . ... . 98

5.7.3 General Design Criteria .............. 106

5.7.4 Webbing and Attachments ........... ... 113

5.7.5 Inertia Reels, Control, and

Installation .................. . .. 118

5.8 SEAT STRENGTH AND DEFORMATION DESIGN

REQUIREMENTS . . . . . ........ ... 119

5.8.1 Recommended Occupant Weights for Seat

Design ............ ................. 119

5.8.2 Strength and Deformation .......... .. 121

5.8.3 Other Seats. . . . . . . .... . . ....... 127

5.9 PERSONNEL RESTRAINT HARNESS TESTING... 128

5.10 STRUCTURAL SYSTEM TEST REQUIREMENTS. .. 128

5.10.1 Static Test Requirements ........ .. 129

5.10,2 Dynamic Test Requirements. . ....... 135

5.10.3 Data Acquisition and Reduction . . . .. 140

5.10.4 Seat Component Attachment .......... . 142

5.11 LITTER STRENGTH AND DEFORMATION

REQUIREMENTS......... . 142

5.11.1 General. .. ........ ......... . 142

5.11.2 Recommended Occupant Weights for Litter

Design .. .. .... . . . . . . . 142

5.11.3 Vertical Loads ......... ....... . . 143

5.11.4 Lateral and Longitudinal Loads . . . . 145

5.11.5 Litter Restraint Harness Testing . . .. 146

5.11.6 Litter System Test Requirements . ... 146

5.12 DELETHALIZATION OF COCKPIT AND CABIN

INTERIORS. . . . . . . . . ............. . 148

5.12.1 General .......... ... . . . ........ 148

5.12.2 Occupant Strike Envelopes. . . . . . .. 148

5.12.3 Environmental Hazards ............. ... 150

7

,.~~ ..........•

Is I I ; l i I I I] J ] • • , •, • , I ¢ , , ,• I I I I I I
TABLE OF CONTENTS (CONTD)

Section Title Page

5.12.4 Head Impact Hazards. . . . . . . . . . . 151

5.12.5 Instrument Panel Structure Proximity . . 156

5.12.6 Rudder Pedal Configuration ....... .. 156

5.12.7 Controls and Control Columns . ..... 157

5.12.8 Sighting and Visionic Systems . ..... 157

5.12.9 Energy-Absorbing Requirements for

Cockpit and Cabin Interiors ....... ... 158

5.13 DESIGN CHECKLISTS ..... ..... .... 166

5.13.1 General Design Checklist.. . ......... . 166

5.13.2 Seat Strength and Deformation

Checklist ............ ............ . . 167

5.13.3 Seat Cushions Checklist. . . . . . . . . 168

5.13.4 Litter Strength and Deformation

Requirements Checklist ......... ... 168

5.13.5 Restraint System Design Checklist. . . . 169

5.13.6 Protective Padding Checklist . . . . . . 170

5.13.7 Cockpit Controls and Equipment

Checklist ...... ............... ... 171

CHAPTER 6. AIRCRAFT POSTCRASH SURVIVAL ....... ... 172

6.1 INTRODUCTION .... .............. ... 172

6.2 FUEL SYSTEM DESIGN CRITERIA ....... ... 172

6.2.1 General ........ ............... ... 172

6.2.2 Fuel Tanks ..... ............... .. 172

6.2.3 Fuel Lines ..... ............... .. 173

6.2.4 Frangible Attachments ............ ... 177

6.2.5 Self-Sealing Breakaway Valves ...... .. 179

6.2.6 Fuel Drains ..... ............... ... 182

6.2.7 Filler Units .... .............. .. 182

6.2.8 Fuel Boost Pumps ............ 182

6.2.9 Fue* Filters and Strainers ....... .. 183

6.2.10 Fuel Valves .............. ............ 183

6.2.11 Fuel Quantity Indicators ........... . 183

6.2.12 Vents .......... .............. ... . 184

6.3 OIL AND HYDRAULIC SYSTEM DESIGN

CRITERIA ........ . . . . . . . . . . 185

6.3.1 General ................ ........... 185

6.3.2 Oil and Hydraulic Fluid Reservoirs 185

6.3.3 Oil and Hydraulic Lines. ......... 185

6.3.4 Oil and Hydraulic System Components. 187

6.3.5 Oil Coolers ..... ............... ... 187

6.4 IGNITION SOURCE CONTROL CRITERIA .... 187

6.4.1 Electrical Systems .... ......... .. 187

6.4.2 Shielding ..... ................ .. 189

6.5 INTERIOR MATERIALS SELECTION CRITERIA. . 190

i8
I

I

TABLE OF CONTENTS (CONTD)

Section Title Page 1

6.5.1 General. .... 190 1

6.5.2 FAR 25.853 Flmmability Requirements 190

6.5.3 UMTA Flammability and Smoke Emission

Guidelines ......... ........... .. 191 • I

6.6 DITCHING CRITERIA.... .............. 193

6.6.1 General ........ .............. .. 193

6.6.2 Emergency Exits. ........... ........ 193

6.6.3 Underwater Emergency Lighting. ..... .. 193

6.6.4 Helicopter Flotation Systems ...... 194

6.6.5 Ditching Equipment . ... ........... . 194

6.7 EMERGENCY ESCAPE DESIGN CRITERIA . . . 194

6.7.1 Emergency Exits ..... .......... .. 194 1

6.7.2 Emergency Lighting . . .. . . .... . ... 199

6.7.3 Emergency Exit Markings .......... .. 200

6.7.4 Crew Chief Stations ............ ... 200

6.7.5 Alarm Systems................ . . ... 200

6.8 CRASH LOCATOR BEACON DESIGN CRITERIA . 200

6.8.1 General ...... ................ .. 200

6.8.2 Crash Sensors ..... ............ . . 201

6.8.3 Transmitters ..... ........... .. 204

6.8.4 Antennas ....... ............... .. 204

6.8.5 Power Supplies .... ............. .. 204

6.9 DESIGN CHECKLISTS ..... ........ .. 205

6.9.1 Fuel System Design Checklist. .... ... 205

6.9.2 Oil and Hydraulic System Design

Checklist...... 211

6.9.3 Ignition Source Control Checklist. ... 214

6.9.4 Interior Materials Selection

Checklist ............ ......... .. 215

6.9.5 Ditching Provisions Checklist. ....... 216

6.9.6 Emergency Escape Design Checklist. . 217

6.9.7 Crash Locator Beacon Checklist .... .. 220

"REFERENCES ............ ..................... .. 222

BIBLIOGRAPHY ............ .................... 229

INDEX ............... ....................... 265

9
LIST OF ILLUSTRATIONS

Figure Page

1 Aircraft coordinates and attitude

directions ............ . . . . 23

2 Typical aircraft floor acceleration

pulse .......... ................ . . . 27

3 Terminology for directions of forces on

the body . . . ............. . 31

4 Seating geometry . . . . . . . . . . . . 34

5 Design velocity changes - off-axis

requirements ....... ........... . . . 44

6 Duration and magnitude of spineward

acceleration endured by various

subjects .................. .... 47

7 Initial rate of change of spineward

acceleration endured by various

subjects ....... ................ .. 48

8 Duration and magnitude of headward

acceleration endured by vari')us

subjects .... ............ . . . . . 49

9 Initial rate of change of headward

acceleration endured by various

subjects ................ ........ 50

10 Conventional seated anthropometric

dimensions ................ .... 54

11 Structural layout for occupant

protection in a crash environment. . . . 61

12 Low angle impact design conditions

(simulated approach with antitorque

loss under poor visibility) ....... .. 63

13 Nose section design conditions ..... 64

14 Rollover, roof impact design

condition .... ............ ........ 66

15 Rollover, side impact design

condition ...... ................ .. 67

10
LIST OF ILLUSTRATIONS (CONTD)

Figure Pg

16 Load-displacement requirements for

energy-absorbing cargo restraint

systems (forward loading of rotary- I

wing and fixed-wing aircraft). . . . . 76

17 Cargo lateral load-displacement [ requirements .......... ... 77

18 Static test floor warpage requirement

to improve the probability of seat Sretention in crashes . .. .. .. . .. 89

19 Bulkhead in-plane warping. . . . . . . . 90

20 Pin joint releases oriented to allow

rotation around an aircraft roll axis.. 91 I

21 Basic aircrew restraint system ..... 99 I

22 Aircrew restraint system, including

reflected shoulder straps. . . . . . . . 100

23 Aircraft troop/passenger restraint

systems ................ . . . . . . 102

24 Gunner restraint system. . . . . . . . . 104

25 Inflatable body and head restraint . . . 105

26 Buckle fitting attachment and motion

angles . ..................... 107

27 Lap belt anchorage geometry .......... 110

28 Shoulder harness anchorage geometry . . 111

29 Stitch pattern and cord size . . . . . . 116

30 Wrap radius for webbing joints . . . . . 117

31 Webbing fold at metal hardware

attachment ............. . . ....... 117 -

32 Seat forward load and deflection

requirements for all types of Army aircraft (forward design pulse). . . . . 122

• 11

4l ~

I I I: I" I •1 I • I 'r•'• 1 •1 -I-1 i~ r - ' I-i -
LIST OF ILLUSTRATIONS (CONTD)

Figure Page

33 Lateral seat load and deformation

requirements for all types of Army

aircraft . ........ .... . . . . . 127

34 Suggested method of applying floor

warping for static testing of seats. . . 131

35 Suggested method of applying bulkhead

warping for static testing of seats. . . 132

36 Static load application point and

critical body block pelvis geometry . . . 133

37 Requirements of additional dynamic

tests if substituted for static tests.. 136

38 Dynamic test requirements for

qualification and for research/

development testing. . . . ........... 137

39 Example of input pulse for seats having

less than 12 in. of stroke . . . . . . . 139

40 Graphic approximation example. . . . . . 141

41 Litter downward load and deflection

requirements . . . . . . . . . . . . . . 144

42 Litter forward or lateral load and

deflection requirements for all

types of Army aircraft . . . . . . . . . 145

43 Full-restraint extremity striKe

envelope - side view . . . . . . . . . . 149

44 Full-restraint extremity strike

envelope - top view. . . . . . . . . . . 150

45 Full-restraint extremity strike

envelope - front view. . . . . . . . . . 151

46 Lap belt-only extremity strike

envelope - side view . . . . . ..... 152

47 Lap belt-only extremity strike

envelope - top view. . . . . . . . . . . 153

12
LIST OF ILLUSTRATIONS (CONTD)

Figure Page

48 Lap belt-only extremity strike

envelope - front view. . . . . . . . . . 154

49 Wayne State Tolerance Curve for the

human brain in forehead impacts

against plane, unyielding surfaces . . . 155

50 Antitorque, or rudder, pedal geometry

to prevent entrapment of feet. . . . . . 156

51 Crushable eyepiece concept . . . . . . . 159

52 Recommended stress-strain properties

for padding material for head contact,

with cushion thickness of at least 1.5 in .. . . . . . . . . . . . . . . . .163

53 Effect of density on stress-strain

curves for polyurethane-foamed plastic . . . . . . . . . . . . . . . . . 164

54 Hose assembly test modes . . . . . . . . 175

55 Sample frangible attachment separation

load calculation . . . . . . . . . . . . 178

56 Typical method of breakaway load

calculation for fuel tank-to-line

breakaway valve. . . . . . . . . . . . . 180

57 Static tensile and shear test modes

for self-sealing breakaway valves. . . . 181

58 Proposed specification for longitudinal

crash force sensors in rotary- and light

fixed-wing aircraft. . . . . . . . . . . 202

59 Proposed specification for vertical

crash force sensors in rotary-wing

aircraft . . . . . . . . . . . . . . . . 203

13
LIST OF TABLES

Table Page

I Crashworthiness criteria for the

preliminary design process . . . . . . . 22

2 Summary of design conditions for rotaryand light fixed-wing aircraft. . . . . . 45

3 Summary of anthropometric data for

U. S. Army aviators. . . . . . . . . . . 55

4 Summary of anthropometric data for

soldiers . . . . . . . . . . . 56

5 Performance requirements for structural

crashworthiness. . . . . . . . . . . . . 60

6 Cargo restraint loads and displacement

requirements . . . . . . . . . . . . . . 75

7 Aircraft cargo categories. . . . . . . . 77

8 Restraint harness components load-elongation design and test

requirements...... . . . . . . . . 114

9 Restraint webbing characteristics. . . . 114

10 Minimum webbing width requirements . . . 115

11 Typical aviator weights. . . . . . . . . 119

12 Troop and gunner weights . . . . . . . . 120

13 Seat design and static test

requirements . . . . . . . . . . . . . . 130

14 Litter system static test

requirements . . . . . . . . . . . . . . 147

15 Summary of ASTM test methods and

specifications for flexible

cellular plastics. . . . . . . . . . . . 161

16 Required minimum individual loads for

standard hose and hose-end fitting

combinations . . . . . . . . . . . . . . 176

14
INTRODUCTION

For many years, emphasis in aircraft accident investigation was

placed on determining the cause of the accident. Very little

effort was expended in the crash survival aspects of aviation

safety. However, it became apparent through detailed studies

of accident investigation reports that large improvements in

crash survival could be made if consideration were given in the

initial aircraft design of 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 can 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

i safety the Aviation and survivability. Safety Engineering Through and a Research series of Division c(ntracts (AvSER) with

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

with occupant survival in aircraft crashes were studied to determine specific relationships between 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 consclidated the design criteria

presented in these reports into one technical document suitable

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

other interested personnel. The document was to be a summary

*Now the Applied Technology Laboratory, U. S. Army Research

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

15
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

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,

TR-71-22, was the basis for the criteria contained in the Army's

aircraft crashworthiness military standard 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.

16

AQ I&
IIA

* - 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

Postcrash fire, ditching, emergency escape, crash locator

beacons, retrieval of accident information.

In this volume (Volume I), Chapter 1 introduces and explains

the intended use of the material contained herein. Chapter 2

contains definitions of terms used in the Design Guide. Chapters 3, 4, 5, and 6 contain the criteria and checklists extracted from Volumes II, III, IV, and V, respectively. The

reader of this volume is strongly encouraged to familiarize

himself with the material in the other volumes, at least in his

particular area of responsibility (e.g., seats and restraints

or fuel systems), in order to more fully appreciate the limitations of the criteria.

The criteria are supplemented by checklists that are intended

for use by aircraft designers in the original design stages and in the design review. These checklists should help the deapply the S-signerecessary criteria in a comprehensive and orderly manner during the development of crashworthy designs, and

provide a rapid and positive means of determining that none of

the criteria have been overlooked. The responses on the checklists also should aid the designer in determining the strengths

and weaknesses of an existing or proposed design. -.1

After the designer has finished reviewing a system design, each

item on the applicable checklists should have a check mark in

one of the spaces following the item. Those items marked "NO"

should be examined to determine the reason for noncompliance

with the design criteria. Unless the reason involves a conflicting, overriding requirement, the design should be revised

to meet the crashworthy criteria. Those items marked "N/A"

should be carefully reviewed to be sure that the item is truly

* not applicable to the system under consideration.

17
II

1. BACKGROUND DISCUSSION

The overall objective of designing for crashworthiness is to

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 crashworthy features is a wise investment (References 2 through 13). Consequently, new generation Army 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 protection afforded to vehicle occupants. When an aircraft impacts

2. ENGINEERING ANALYSIS OF CRASH INJURY IN ARMY OH-58 AIRCRAFT, USASC Technical Report 79-1, U. S. Army Safety

Center, Fort Rucker, Alabama, January 1979.

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. 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., University

Press of Virginia, 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

Technology Corporation; USAAMRDL Technical Report 75-59A,

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

Development Laboratory, Fort Eustis, Virginia, May 1976,

AD A026246.

18
* I

the ground, deformation of the ground absorbs some energy.

This is an uncontrolled variable since the quality of the impacted 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 J

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

Sstructure 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.

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

STRUCTURAL DESIGN INVESTIGATION, The Boeing Vertol Company; 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 THE 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 TANUARY

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 CRASHWORTHINESS AND FLIGHT SAFETY

DESIGN FEATURES IN ATTACK HELICOPTERS, USAAAVS Technical

Report 77-2, U. S. Army Agency for Aviation Safety, Fort

Rucker, Alabama, June 1977.

19

i.i
I'i

Ideally, it would seem most efficient to simply specify human

tolerance requirements and an array of vehicle crash impact

conditions, then develop the aircraft as a crashworthy system

with a mixture of those crashworthy features that are most efficient for the particular vehicle 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 functions; for example, aircraft attitude at impact may prevent the landing gear from absorbing its share of the impact crash energy. 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. Therefore,

to achieve the overall goal, minimum levels of crash protection

are recommended for the various individual subsystems.

Current aircraft crashworthiness criteria do require that a new

aircraft be designed as a system to meet the vehicle impact

design conditions recommended in Volume I1. Also, minimum criteria are specified for a few crash critical components. For

example, strengths and minimum crash energy-absorption requirements for seats and restraint systems are specified. All

strength requirements presented in this volume are based on r the crash environments described in Volume II. Testing reKi quirements 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 rotaryor light fixed-wing aircraft should conform to the requirements

of MIL-S-58095(AV) (Reference 14).

Although much higher levels of crashworthiness can be achieved

in 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 guide. This can even be

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

CRASHWORTHY, NON-EJECTION, AIRCRAFT, GENERAL SPECIFICATION

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

1971.

20

IT ______
achieved while expanding the combat effectiveness of the aircraft. Examples of this are the successful program to retrofit all U. S. Army helicopters with crashworthy fuel systems (Reference 15), and the U. S. Navy program to retrofit the

CH-46 with crashworthy armored crewseats (Reference 16).

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 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 between the affecting 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

preliminary 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 Industries; USAAMRDL

Technical Report 71-47, U. S. Army Air Mobility Research

and Development Laboratory, Fort Eustis, Virginia, November 1971, AD 739567.

16. Domzalski, L. P., et al., U. S. NAVY DEVELOPMENTS IN

CRASHWORTHY SEATING, Naval Air Development Center; Proceedings 1978 SAFE Symposium, Survival and Flight Equi-u

ment Association, Canoga Park, California, October 1978.

* 21 A

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22
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 1. When referring to an aircraft in any flight attitude, it is standard prac- tice to use a basic set of orthogonal axes as shown

in Figure 1, with x, y, and z referring to the longitudinal, lateral, and vertical directions, respec- tively.

+z

Yaw

++

Roll y

Figure 1. Aircraft coordinates and attitude directions.

23
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 1 as X, Y, Z.

* 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 1).

Aircraft pitch is the angle between its longitudinal

axis and a horizontal plane. Pitch is considered

positive when the nose of the aircraft points above

the horizon and negative when it points below the

horizon. Yaw is measured between the aircraft's longitudinal axis and the flight path. Roll is the

angle between an aircraft lateral (y) axis and the

horizontal, measured in a plane normal to the aircraft's longitudinal axis.

"* Flight Path Angle

The angle between the aircraft flight 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 the 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.

24

.. .. - -° I S- -T. ...... .
I

Flight

path impact

vlight path angle

Sangle Horizontal

: ' • • Terrain

2.2 ACCELERATION-RELATED TERMSf

* Acceleration

The ra.z 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 iotational), 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 of

the floor of the fuselage.

9 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 commonly experienced in

an aircraft crash (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.

25

................... ... 4.. .... . ... - .______________ .
9 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 illustrate, it is customarily understood

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

161 ft/sec

2.3 VELOCITY-RELATED TERMS

* Velocity Change in Major Impact (Av)

The decrease in velocity of the airframe during the

major impact, 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 2, the

major impact should be considered ended at time t . Elastic recovery in the structure will tend to reverse the direction of the aircraft velocity prior

to t . Should the velocity actually reverse, its

diretion 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

Av - 30 - (-5) - 35 ft/sec

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

* 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 exper- ience 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.

26

I I I I I II I I I I I I , ,
r

A

Peak

II

o• Average-

, • I

I i

U

t| 2-tl 2t t2 t 3 Time L, I

t 2-tI

Figure 2. Typical aircraft floor acceleration pulse.

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.

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.

27

d ____________
"* 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)

axis.

"* Downward Load

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

• Pward Load

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.

o Combined Load

Loading consisting of components in more than one of

the directions described in Section 2.1.

* Crash Force Resultant

The geometric sum of horizontal and vertical crash

forces: horizontal and vertical velocity components

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 force r G Vertical

Resultant | v crash force

angle J

Gh Horizontal crash force

28

-- I@ .8-I---
• 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

longitudinal

Horizontal /••

Aircraft

S~~pitch /

force • / force = Resultant Pitch

angle , / angle angle angle

r Resultant Gvertical L crash forc

/ • •"angle

I i Ghorizontal :

! ~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.

e• 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.

2

29 1

, | I i " '• "
e 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 in a manner 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 occuring 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

e 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

30

I • m "• i • . . .. -• 
I

compiled the accelerative terminology table of equivalents shown in Figure 3 (Reference 17). Terminology 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

(sternumward) Ltrlrgt Tases Tailward - Eyeballs-up

(+G Transverse

y Lateral right - Eyeballsleft

Lateral left - Eyeballsright

Back to chest - Eyeballsin

Lateral left Chest Chest to back - Eyeballs-

(-C to back Out

y (spineward) Note:

Tailward -x) The accelerative force on

(-G ) the body acts in the same

direction as the arrows.

Figure 3. Terminology for directions of forces on the body.

* 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 such

as iliac crests, ischial tuLerosities, rib cages, etc.

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.

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

1961, pp. 1109-1111.

31

T II II _IiI-II I I_____lII I IIII
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 applied 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 section 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.

* Submarining

Rotation of the hips under and about the lap belt as

a result of a forward inertial load exerted by deceleration of the thighs and lower legs, accompanied by

lap belt slippage up and over the iliac crests. Lap

32

-
Fi

belt slippage up and over the iliac crests can be a

direct result of the upward loading of the shoulder

harness straps at the center of the lap belt.

0 Effective Weight

The portion of occupant weight supported by the seat

t with the occupant seated in a normal flight position.

This is considered to be 80 percent of the occupant weight since the weight of the feet, lower legs, and

part of the thighs is carried directly by the floor

through the feet.

e Iliac Crest Bone

The upper, anterior portion of the pelvic (hip) bone.

These "inverted saddle" bones are spaced laterally

about 1 ft apart; the lower abdomen rests between

these crest bones.

* Lap Belt Tiedown Strap (also Negative-G Strap,

Crotch Strap)

Strap used to prevent the tensile force in shoulder

straps from pulling the lap belt up when the restrained subject is exposed to -G (eyeballs-out)

acceleration.

2.8 SEATING GEOMETRY (See Figure 4 from Reference 18)

o Design Eye Position

A reference datum point based on the eye location

that permits the specified vision envelope required

by MIL-STD-850 (Reference 19), allows for slouch, and

is the datum point from which the aircraft station

geometry is constructed. The design eye position is

a fixed point in the crew station, and remains constant for pilots of all stature via appropriate seat

adjustment.

18. Military Standard, MIL-STD-1333A, AIRCREW STATION GEOMETRY

FOR MILITARY AIRCRAFT, Department of Defense, Washington,

D. C., 30 June 1976.

19. Military Standard, MIL-STD-850B, AIRCREW STATION VISION1

REQUIREMENTS FOR MILITARY AIRCRAFT, Department of Defense,

Washington, D. C., 3 November 1970.

-- 33
Design eye position

Horizontal

vision line

,13* desired

minimum 4

back angle

Thigh tangent line /n.

> WBack line tangent

/00, reference

1i

LI -?, line+

200 maximum

" " ( copters, Buttock reference point planes

50 minimum

Zfor others Heel rest line

(Not necessarily the floor)

Figure 4. Seating geometry. (From Reference 18)

e Horizontal Vision Line

A reference line passing through the design eye position parallel to the true horizontal and normal

cruise position.

* Back Tangent Line

A straight line in the midplane of the seat passing

tangent to the curvatures of a seat occupant's back

when leaning back and naturally compressing the back

cushion. The seat back tangent line is positioned

13 in. behind the design eye position measured along

a perpendicular to the seat back tangent line.

34

"--. . -.... . .......
F I

e Buttock Reference Line

A line in the midplane of the seat parallel to the

horizontal vision line and tangent to the lowermost

t natural 

sitting protrusion of a selected size of occupant i

on the seat cushion.

* Seat Reference Point (SRP)

The intersection of the back tangent line and the

buttock reference line. The seat geometry and location are based on the SRP.

* Buttock Reference Point

A point 5.75 in. forward of the seat reference point

on the buttock reference line. This point defines

the approximate bottom of an ischial tuberosity, thus representing the lowest point on the pelvic

structure and the point that will support the most

load during downward vertical loading.

o Heel Rest Line

The reference line parallel to the horizontal visionI

line passing under the tangent to the lowest point

on the heel in the normal operational position, not

necessarily coincidental with the floor line.

2.9 STRUCTURAL TERMS

e Airframe Structural Crashworthincss

The ability of an airframe structure to maintain a

protective shell around occupants during a crash and

to minimize accelerations applied to the occupiable

portion of the aircraft during crash impacts.

* 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.

9 Static Strength

The maximum static load that can be sustained by a

structure, often expressed as a load factor in terms

of G (see Load Factor, Section 2.4).

35

__ _ _ _ I , . ... ... • • , • • -- .•_•- *. •- • •,••,• .
* Strain

The ratio of change in length to the original length

of a loaded component. 4

* Collapse

I

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.

9 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.

* Specific Energy Absorbed (SEA)

The energy absorbed by an energy-absorbing device or

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

* Bottoming

The exhaustion of available stroking distance accompanied by an increase in force, e.g., a seat stroking in the vertical direction exhausts the available

distance and impacts the floor.

36
* 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 provided.

2.10 FUEL, OIL, AND HYDRAULIC SYSTEM TERMS

* Boost Pump

A fuel pump installed in the tank of an aircraft to

supply the main (usually engine-driven) fuel pump with sufficiently high inlet pressure to meet net

positive suction head (NPSH) requirements under all

flight conditions.

* Frangible Attachment

An attachment possessing a part that is constructed

to fail at a predetermined location and/or load.

* Fuel Valve

Any valve, other than a self-sealing breakaway valve,

contained in the fuel supply system, such as fuel

shutoff valves, check valves, etc.

* Self-Sealing Breakaway Valve

A fluid-carrying line or tank connection that will

separate at a predetermined load and seal at both

ends so that an absolute minimum of fluid is lost.

2.11 IGNITION SOURCE CONTROL TERMS

* Fire Curtain

A baffle made of fire-resistant material that is used

to prevent spilled flammable fluids and/or flames

from reaching ignition sources or occupiable areas.

* Fire-Resistant Material

Material able to resist flame penetration for 5 min

when subjected to 2000OF flame and still be able to

perform its intended function.

S37

... 'A Or_
I%

* Firewall

A partition capable of withstanding 2000*F flame over

an area of 5 in. for a period of 15 min without

flame penetration.

"* Flammable Fluid

Any fluid that ignites readily in air, such as hydrocarbon fuels and lubricants.

"* Flow Diverter

A physical barrier that interrupts or diverts the

flow of a liquid.

* Ignition Temperature

The lowest temperature at which a flammable mixture

will ignite when introduced into a specific set of

circumstances.

* Inerting

The rendering of an aircraft system or the atmosphere

surrounding the system incapable of supporting combustion.

2.12 INTERIOR MATERIALS SELECTION TERMS

9 Autoignition Temperature

The lowest temperature at which a flammable substance

will ignite without the application of an outside

ignition source, such as flames or sparks.

a Flame Propagation Index (I s) S|

A number calculated by combining two factors derived

from the radiant panel test for material flammability

(see Section 6.5.3). One factor is derived from the

rate of progress of the flame front and the other is

derived from the rate of heat liberated by the material under test.

* Flame Resistant

Material that is self-extinguishing after removal of

a flame.

38
9 Flashover

The sudden spread of flame throughout an area due to

ignition of combustible vapors that arm heated to

their flash point.

* Flash Point

,The lowest temperature at which vapors above a combus-

,tible substance ignite in air when exposed to flame.

* Intumescent Paint

A paint that swells and chars when exposed to flames.

* Optical Density (D

The optical density is defined by the relationship

D100 ,

s= log T (1)

where T is the percent of light transmission through

a medium (e.g., air, smoke, etc.).

2.13 DITCH AND EMERGENCY ESCAPE TERMS

9 Brightness

The luminous flux emitted per unit of emissive area

as projected on a plane normal to the line of sight.

Measured in foot-lamberts.

* Candela (cd)

A unit of luminous intensify equal to 1/60 of the

luminous intensity of 1 cm of a black-body surface

at the solidification temperature of platinum. Also

called candle or new candle.

* Class A Exit

A door, hatch, canopy, or other exit closure intended

primarily for normal entry and exit.

9 Class B Exit

A door, hatch, or other exit closure intended primarily for service or logistic purposes (e.g., cargo

hatches and rear loading ramps or clamshell doors).

39

L .••-..a•& l. " •.• w.•• ~ ~ •-•• •"i-- l €--
Ii

9 Class C Exit

A window, door, hatch, or other exit closure intended

primarily for emergency evacuation.

e Cockpit Enclosure

That portion of the airframe that encloses the pilot,

copilot, or other flight crew members. An aircraft

may have multiple cockpits, or the cockpit may be

physically integrated with the troop/passenger section.

* Ditching

The landing of an aircraft on water with the intention of abandoning it.

9 Emergency Lighting

Illumination required for emergency evacuation and

rescue when normal illumination is not available.

a Exit Closure

A window, door, hatch, canopy, or other device used

to close, fill, or occupy an exit opening.

* Exit Opening

An opening provided in aircraft structure to facilitate either normal or emergency exit and entry.

* Exit Release Handle

The primary handle, lever, or latch used to open or

jettison the exit closure from the fuselage to permit

emergency evacuation.

* Foot-candle (fc)

A unit of illuminance on a surface that is everywhere

I ft from a uniform point source of light of I candela.

* Foot-lambert (fL)

A unit of photometric brightness or luminous intensity per unit emissive area of a surface in a given

direction. One foot-lambert is equal to l/7ncandela

per square foot.

40
_4

" Illumination

The luminous flux per unit area on an intercepting

surface at any given point. Measured in foot-candles.

4

r.

I

41
3. AIRCRAFT CRASH ENVIRONMENT AND HUMAN TOLERANCE

3.1 INTRODUCTION

Design criteria that can be extracted from information on the

aircraft crash environment and the response of the human body

to that environment are presented in this chapter. Principles, data, and analysis methods that influence the survivability of

aircraft occupants in a crash environment are summarized. The

reader is referred to Volume II for a more complete discussion

of factors from which these design principles are drawn.

3.2 DESIGN CONDITIONS FOR IMPACT

3.2.1 General

3.2.1.1 Application: In using the design data tabulated in

this section, it should be emphasized that the values given are

estimates for survivable accidents in pre-1978 aircraft. New

aircraft can be designed to permit survival during a much more

severe crash. Although improvements in crashworthiness can

be achieved in existing aircraft by retrofit systems, such as

energy-absorbing seats or crashworthy fuel systems, the improvements are limited and may result in prohibitive weight and

cost penalties if requirements are too severe. Retrofit decisions are made as the result of tradeoffs between the benefits

in survivability and the penalties of cost and weight. An aircraft should be designed as a system to provide the required occupant protection for the recommended velocity changes because

deceleration is a design variable, a function of the structural

stiffness of the fuselage. Consideration of crashworthiness

in design of the complete aircraft system eliminates many of

the limitations inherent in retrofit and makes possible the design for more severe environments without significant weight penalties.

3.2.1.2 Deceleration Pulse Shape: Experimental data obtained

in full-scale crash tests of helicopters, light fixed-wing aircraft, and fixed-wing transports indicate that the deceleration

pulse shape for major impact in accidents can be represented

to a satisfactory degree for most engineering purposes by a

triangle as shown in Figure 2. Energy-absorbing landing gear

on new aircraft will produce a lower-level deceleration plateau preceding the fuselage contact, thereby reducing the energy that must be absorbed by fuselage crushing. However, the shape of the deceleration pulse during fuselage contact with the

ground will still approximate a triangle.

4 2

42

.--.- i
3.2.1.3 Impacted Surface: Statistically, the crash surface

most frequently impacted is sod. It is recommended that sod

with a California Bearing Ratio (CBR) of 2.5 be accepted as the

* standard for crashworthy design. Trees are the second most

* frequently impacted obstacle; however, the secondary (in this

case, major) impact would still be with sod.

3.2.1.4 Impact Attitude and Velocity Change: Information con- 4

cerning impact attitude is extremely important to the adequate

design of crashworthy aircraft. Data that would permit a cornplete statistical definition of aircraft impact attitude are

not yet available. However, studies of crash data (from two

helicopter types, cargo and attack) were reviewed, as discus--

sod in Volume II, and the typical impact attitudes of rotarywing aircraft are:

Roll t20 degrees

Yaw (not determined)

Pitch + (nose up) 25 degrees - (nose down) 15 degrees

The design information available for each major axis must be

extrapolated to intermediate positions with the global coordinate system to provide guidance for the design of structure

subjected to combined loading (combinations of loads with components in the three different axis directions). Wherever criteria are presented in this document for the three major axes,

combinations of the conditions also apply for all intermediate

positions between axes. To make this very clear, the criteria

specified for the specific axes x, y, and z are not to be construed as constituting the only requirements. Consideration

of combinations of the specified loads or velocity changes between axes also is required, as illustrated in Figure 5.

For helicopters and light fixed-wing aircraft, the resultant

velocity change for combined longitudinal, vertical, and lateral components does not appear to exceed 50 ft/sec. The vertical or lateral components do not exceed the 95th-percentile

* values based on the specific axis directions; i.e., 42 ft/sec

vertically for all rotary- and light fixed-wing aircraft,

25 ft/aec laterally for light fixed-wing aircraft and attack

and cargo helicopters, and 30 ft/sec for other helicopters.

Figure 5 illustrates combined longitudinal, lateral, and vertical velocity changes for helicopters, to be used in determining intermediate velocity change components. For light

43

WOW,

-- __......I7
10.

.4 44

44

-4n to

u

C ~ 0 C C U 0 0 ~ ~ *- 1.

AU

0

0

dim N

No 0 0

45p

444
A fixed-wing aircraft and attack and cargo helicopters, Figure 5(b) will still be correct, but (c) and (d) must be altered for a lateral velocity change of 25 ft/sec instead of

30 ft/sec.

In general, the three components are related by the equation

2 2 2 V2 (2)

X y z R

where vx longitudinal velocity change, ft/sec

v * lateral velocity change, ft/sec L y

vz = vertical velocity change, ft/sec

vR = resultant velocity change, ft/sec

and the axes are those illustrated in Figure 1. The curves

have been terminated at 15 degrees, based on a study of accident reports discussed in Volume II.

Table 2 gives the velocity change Av in feet per second for the

triangular pulse shape of Figure 2. The pulses resulting from

the values given in Table 2 are recommended for design purposes for rotary- and light fixed-wing aircraft.

TABLE 2. SUMMARY OF DESIGN

CONDITIONS FOR

ROTARY- AND LIGHT

FIXED-WING AIRCRAFT

Velocity

change

Impact direction (ft/sec)

Longitudinal 50

Vertical 42

Lateral* 25

Lateral** 30

*Light fixed-wing, attack, and

cargo helicopters.

"**Other helicopters.

45

I 1'A
3.3 HUMAN TOLERANCE TO IMPACT

3.3.1 General

Results of research on tolerance of the human body to impact

forces are presented in Volume II, Chapter 4. Although numerous experiments have been conducted and a wealth of information

has been collected, very few criteria that may be useful in

system design have been developed and validated. In this chapter, those criteria that are generally accepted for practical

application in assessing the crashworthiness of an aircraft

system are presented. As discussed here, these criteria may

be used to determine the acceptability of an aircraft or components, such as seats and restraint systems, based on the results of dynamic testing with anthropomorphic dummies or com-.

puter simulations as discussed in Volume IV. Criteria are

presented here only if validated quantitative values have

been determined. Injuries to other body parts have also been

studied and are discussed in Volume II.

3.3.2 Whole-Body Tolerance

Tolerance of the human body to abrupt acceleration has been

shown to depend on the magnitude and duration of the applied

force, as well as the direction and rate of onset. Data presented by Eiband (Reference 20) for occupants having upper

torso restraint are summarized in Figures 6 and 7 for spineward

(-G ) acceleration and in Figures 8 and 9 for headward (+G)

acceleration. Human tolerance to lateral (G ) acceleration2has

not been extensively studied. However, basdd on the testing that has been conducted, a maximum lateral acceleration of 20 G at a duration of 0.1 sec is suggested for design.

An acceptable personnel restraint system for Army aircraft

should include upper torso restraint, regardless of seat orientation. However, for reference and for comparison with the

above values, a spineward (-G ) human tolerance level of 20 G

and a lateral (G ) level of ldcG are recommended for lap-beltonly restraint. YThese levels are based on experiments with human subjects' in which minor trauma were experienced.

Although Figures 6 through 9 indicate the regions of acceleration and rate of onset that may be considered acceptable for the

aircraft interior, they do not permit complete evaluation of

such protective systems as restraint systems, energy-absorbing

seats, or protective padding. Injury criteria for critical body

20. riband, A. M., HUMAN TOLERANCE TO RAPIDLY APPLIED ACCELERATIONS: A SUMMARY OF THE LITERATURE, NASA Memorandum

5-19-59E, National Aeronautics and Space Admiristration,

Washington, D. C., June 1959.

46
14'40

I rl ) 4

w 0

1.4

004) >4)

41

0 C

04 to>1 -4

> NW

0 1:6

144

0 0 q

47.
SUBJECT SUPPORT

0 Human Lap, shoulder, thigh,

and chest straps C Human Lap, shoulder, thigh,

and chest straps

SHuman Lap, shoulder, thigh, -

ano chest strapsJ

" Chimpanzee Military lap and shoulder

straps

SChimpanzee 3-in. cotton webbing, 5 Acceleration

horizontal, 2 vertical

straps

b00 _____- __-_j-* - 4 •.. -. . - _._l --- -.

180 so 4 -4-+- •

.~60

Definite No 81 No shock

40 Cardiovascular shock shock g(conunctival- Sins and retinal

Shoc-- hemorrhage)

S20

I I i' H

o , t i -t, 

> 00

S0 G

Figure rfepax Onset rate 

pt2 s a t i c n s

stoigdsaco hte gie shude elwebgha

.001 .002 .004 .006 .01 .02 .04 .06 .1 .2 .4 .6

STirme, sec

Figure 7. Initial rate of change of spineward

acceleration endured by various

subjects. (From Reference 20)

parts, such as the head and spinal column, must be employed in

order to answer Such questions as whether a seat has sufficient

stroking distance, or whether a given shoulder belt webbing has

acceptable stiffness.

3.3.3 Head Injury Criteria

Various criteria have been used as predictors of head injury.

Concussiva threshold values have been identified for four such

criteria; peak G, peak transmitted force, Severity Index,

and Head Injury Criterion. The Severity Index is defined as

48
E4

4-t-4..

. t

-,-,4 4-4

-4'1

04)

F=I

0

0 I I

o4 04 0 C: 1

> -4

ý49

-_ In, _ _ _ 

> 0
41 Si (ai

Fu 2

fa4 0

0 11)

A C, . 0I c4

m w 0) a 0 00'

I4 4-r4-

0004

'U

o ww we 4Zz'83TQ ;0 UOTIVI&SaiZ

Us
t

S SI f andt (3)

to

where SI = Severity Index

a - acceleration as function of time

n - weighting factor greater than 1

t a time

and the Head Injury Criterion (HIC) of Federal Motor Vehicle

Safety Standard 208 is calculated according to

HIC = max f adt (t 2 - t1 ) (4)

2 tI /

where a is the resultant head acceleration, and t1 and t 2 are

any two points in time during the crash event.

Aircrewmen have experienced concussive head injury from helmeted head impacts that exceeded the following values for the

four criteria; peak head accelerations that exceeded 150 G,

peak force levels transmitted to the head that exceeded 1500 lb,

Severity Index values that exceeded 600, and Head Injury Criterion values that exceeded 500. These values should be taken

as the limits of human tolerance to concussion when using these

criteria as predictors of head injury.

3.3.4 Spinal Injury Criteria

. Although the Dynamic Response Index (DRI), as illustrated in

Section 4.8.1 of Volume II, is the only model correlated extensively for ejection seat spinal injury prediction, it has serious shortcomings for use in accident analysis. It assumes the

occupant to be well restrained and erect, so that the loading

is primarily compressive, with insignificant bending. Although

such conditions may be assumed for ejection seats, they are

less probable for helicopter crashes, in which an occupant may

be leaning to either side for better visibility at the time of

impact. Further, the DRI was correlated for ejection pulses of

much longer duration than typical crash pulses.

K 1

• -• H I .*=.-'-.--* • ..... - -i -'' , l , , ,
A more detailed model of the spinal column would yield more

realistic results, but injury criteria for the more complex

responses have yet to be developed. Consequently, the DRI is

not recommended as the criterion for use in designing crashworthy seats. Rather, the data presented in Figure 8 are reccommended for, use until more comprehensive data and criteria

are developed.

3.3.5 Leg Injury Criteria

Femoral fracture due to longitudinal impact on the knee has

been studied extensively, probably because of the frequency of

this type of injury in automobile accidents. A criterion that

assesses the dependence of the permissible human knee load on

the duration of the primary force exposure has been suggested in Reference 21. The permissible peak knee load suggested for

design is given by

F a 5200 - 160 t, t < 20 msec

F - 2000, t > 20 msec (5)

where F is in pounds and t in msec.

3.3.6 Tolerance of Other Body Parts

Although some research has been conducted on the tolerance of

other body parts, such as the neck, thorax, and abdomen, welldefined, valid criteria have not been established. The results

of this research are discussed in Volume II,-Chapter 4.

3.4 HUMAN BODY DIMENSIONS AND MASS DISTRIBUTIONS

3.4.1 General

Anthropometric measurements are external dimensions of the human body that can be used to define aircraft requirements such

as seat height and width, eye height, or cabin height. A specialized type of anthropometric measurement is the "link length,"

or distance between joint centers, which can be used in locating control positions and is essential for the design of

21. Viano, D. C., CONSIDERATIONS FOR A FEMUR INJURY CRITERION, Proceedings, Twenty-First Stapp Car Crash Conference,

Society of Automotive Engineers, Inc., New York, 1977,

pp. 445-473.

52

I
mathematical or physical simulators of the human body. Finally,

the iner*ial properties of the body and parts of the body also

are requ ed in the design of human simulators.

3.4.2 Anchropometry

Two types of anthropometric measurements have been recorded,

and the use of both types in vehicle design has been summarized

in Reference 22.

In the first type, conventional dimensions of the body with

subjects in rigid, standardized positions are easily obtained.

Extensive col)ections of such data are used in clothing design

and may determine certain vehicle design parameters including

seat height and eye height. The anthropometric data of greatest potential usefulness, illustrated in Figure 10, for U. S.

Army aviators and soldiers of the 5th, 50th, and 95th percentiles are presented in Tables 3 and 4, respectively. Complete data can be found in References 23 and 24.

The second type of anthropometric data, which may be referred

to as workspace dimensions, is more difficult to obtain and can

be applied only to the specific workspace studied. However,

these workspace dimensions are es ..ntial in designing aircraft

interiors for maximum occupant protection.

Workspace dimensions must involve a consideration of body

joints, the distance between them, and their range of motion.

Dempster reported on an extensive study of workspace requirements for seated operators, in which he determined "link

lengths" between effective joint centers for major body parts

22. Roe, R. W., and Kyropoulos, P., THE APPLICATION OF ANTHROPOMETRY TO AUTOMOTIVE DESIGN, SAE Paper No. 700553, Society of Automotive Engineers, Inc., New York, 1970.

23. Churchill, E., et al., ANTHROPOMETRY OF U. S. ARMY AVIATORS - 1970, Anthropology Research Project; USANL Technical Report 72-52-CE, U. S. Army Natick Laboratories,

Natick, Massachusetts, December 1971, AD 743528.

24. White, R. M., and Churchill, E., THE BODY SIZE OF SOLDIERS:

U. S. ARMY ANTHROPOMETRY - 1966, USANL Technical Report 72-51-CE, U. S. Army Natick Laboratories, Natick, Massachusetts, 1971, AD 743465.

53

, j

* * - * "*** 
II

Sitting height

S~ Eye height :

Functional

reach Shoulder breadth S

Shoulder-elbow length

Elbow-fingertip

Buttock-knee length Knee height

Popliteal height Hip breadth

Buttock-popliteal

length

Figure 10. Conventional seated anthropometric dimensions.

(References 25 and 26). These link lengths have a number of

crashworthiness-related applications: first, in developing

or expanding the strike envelopes shown in Chapter 5 of Volume II,; second, in designing crash test dummies; and third,

in providing numbers for mathematical simulators. Skeletal

joint locations and ranges of motion are presented in Section

6.2.2 of Volume II.

3.4.3 Inertial Properties

Anthropometric dummies and mathematical simulations require inertial properties of body segments, specifically moments of inertia, mass, and center-of-mass locations. Several studies of

these properties have been made using live human subjects and

25. Dempster, W. T., SPACE REQUIREMENTS FOR THE SEATED OPERATOR, WADC Technical Report 55-159, Wright Air Development Center, Wright-Patterson Air Force Base, Ohio, 1955,

AD 087892.

26. Dempster, W. T., and Gaughran, G. R. L., PROPERTIES OF

BODY SEGMENTS BASED ON SIZE AND WEIGHT, American Journal

of Anatomy, Vol. 120, 1967, pp. 33-54.

54
TABLE 3. SUMMARY OF ANTHROPOMETRIC DATA FOR

U. S. ARMY AVIATORS (Reference 23)

Percentiles (in.)

Measurement 5th 50th 95th

Weight (ib) 133.0 171.0 212.0

Stature 64.6 68.7 72.8

Seated height 33.7 35.8 37.9

Shoulder breadth 17.0 18.7 20.3

Functional reach 28.8 21.1 34.2

Hip breadth, sitting 13.2 14.8 16.7

Eye height, sitting 29.0 31.0 33.1

Knee height, sitting 19.3 20.8 22.6

Popliteal height 15.1 16.6 18.3

Shoulder-elbow length 13.3 14.4 15.6

Elbow-fingertip length 17.6 19.0 20.3

Buttock-popliteal length 17.7 19.3 21.0

Buttock-knee length 22.0 23.7 25.4

cadavers, and such data as have been obtained should be integrated into the design of any anthropometric dummy or mathematical simulation. Results of several of these studies are

summarized in Reference 27.

27. Singley, G. T., III, and Haley, J. L., Jr., THE USE OF

MATHEMATICAL MODELING IN CRASHWORTHY HELICOPTER SEATING

SYSTEMS, in Models and Analogues for the Evaluation of

Human Biodynam c Reponse, Performance and Protection,

AGARD-CP-253, NATO Advisory Group for Aerospace Research

and Development, Neuilly sur Seine, France, June 1979.

55

F,, , '"-w-'-•{•- 4 
II

TABLE 4. SUMMARY OF ANTHROPOMETRIC DATA

FOR SOLDIERS (Reference 24)

Percentiles (in.)

Measurement 5th 50th 95th

Weight (ib) 126.0 156.0 202.0

Stature 64.5 68.7 73.1

Seated height 33.3 35.7 38.1

Shoulder breadth 16.3 17.8 19.6

Nip breadth, sitting 11.9 13.0 14.5

Eye height, sitting 28.6 31.0 33.3

Knee height, sitting 19.6 21.3 23.1

Popliteal height 16.0 17.5 19.2

Shoulder-elbow length 13.3 14.5 15.7

Elbow-fingertip length 17.4 18.8 20.4

Buttock-popliteal length 18.0 19.6 21.3

Buttock-knee length 21.6 23.4 25.3

3.5 CRASH TEST DUMMIES

All of the recently developed dummies were designed for automotive testing and are based on the anthropometry of a 50thpercentile U. S. civi.lian male. In dynamic testing of an

energy-absorbing seat, design for aircraft occupant weight can

play a critical role. It would be desirable to evaluate a

seat for a range of occupant sizes. A 95th-percentile dummy would verify the strength of the seat structure and restraint

system as well as the adequacy of the energy-absorbing stroke.

Testing with a 50th-percentile dummy would demonstrate the performance of the system for an occupant of average height and

weight. A 5th-percentile dummy would probably experience accelerations of higher magnitude and would establish the severity of a given set of impact conditions for the smaller occupant. However, both the expense of dummy purchase and the cost

56

'I ~'* ~~& -
I I

of conducting dynamic tests may make such a test program impractical. An alternative procedure might be to establish the

occupant protection capability of a seat design by analysis and to conduct a dynamic test with a 95th-percentile dummy to

verify system strength.

There are two additional factors that should be considered in

dummy selection for aircraft seat testing. First, some designs

are more suitable than others for testing with a headward (+G )

acceleration component. None of the dummies have been designid

for accurate response to vertical impact. The spinal column,

which is a critical region of human tolerance to aircraft crash

loading, has been designed to simulate response to -G loading

t rather than the more critical *G direction. However, the reinforced rubber cylinder used ai the lumbar spine in a dummy designed in accordance with the specifications in the Code of

L Federal Regulations, Title 49 (49 CFR) Part 572 (Reference 28) permits more consistent positioning than the steel ball-andsocket configuration used in some other dummies. Instability

in the latter type could affect response of the upper torso

with concomitant penalties on test repeatability. Another advantage uf the Part 572 dummy for aircraft seat testing is a

humanlike pelvic structure, which should result in load distribution on the cushion close to that for a human. Secondly,

if the results of tests conducted at different facilities are

Sto be compared, standardization of dummies and test procedures , is mandatory.

At present, it seems that use of the Part 572 dummies, modified

to improve their simulation accuracy to impact loading in the

+G direction and sized to 5th-, 50th-, and 95th-percentile

veisions of the U. S. Army aviator, provides the best available

simulation and is, therefore, the recommended approach.

V

28. U. S. Code of Federal Regulations, Title 49, Chapter 5,

Part 572: ANTHROPOMORPHIC TEST DUMMY, Government Printing Office, Washington, D. C., (Rev.) 1978.

I 57
4. AIRFRAME STRUCTURAL CRASHWORTHINESS

4.1 INTRODUCTION

Salient features required in the definition of a crashworthy

structure are summarized in this chapter. The user is referred

to Volume III for additional information concerning the criteria or their sources.

In a crash situation, the basic requirements for occupant survival of impact hazards are:

e The maintenance of a protective structural envelope.

e The attenuation of impact forces to maintain a survivable acceleration environment.

To achieve the desirable occupant environment, the following

basic design requirements must be considered as an integrated

problem and a practical solution must be obtained. Such design

requirements should be included in new aircraft, and existing

designs could be improved by incorporating these features where

possible.

"* The basic structural envelope surrounding occupied

areas must be designed to maximize its energy absorption capacity.

"* The structure that makes initial contact with the

ground must be designed to minimize the probability

of earth gouging and scooping of soil. This will

minimize the acceleration and force levels to which

the structure is subjected.

"* All items attached to the structure must, where possible, be retained in a survivable crash environment.

These items include large masses, such as transmissions, engines, and rotor systems; internal cargo

and on-board equipment racks; externally mounted

components, such as fuel tanks, wings, and external

stores; and the empennage and landing gear. In the

past, shedding of large-mass items has been considered advantageous in a crash environment. This is

true from the viewpoint of reducing the energy content of the aircraft and, hence, the loads acting on

the structure in resisting aircraft postimpact motions. However, it is possible that penetration of

occupied areas could occur, and during the postimpact

motions, the aircraft could traverse shed objecte

58
w~I

causing high loading on the structure. It is, therefore, better to maintain a known mass if an optimum

acceleration profile is desired for occupant survival.

Thus, mass retention and landing gear integrity are

required for optimum crashworthiness and occupant

environment.

* In the case of helicopters, certain areas of the

cockpit and cabin structure must be reinforced to

withstand loads induced by blade strikes, impacts

with external objects such as trees, and rollover.

In addition, if overhead-mounted crashworthy seats

are used, the deflection of the overhead structure

relative to the floor must be minimized.

* Unoccupied areas of structure, such as the underfloor, nose, and tail areas, must be designed to deform in a controlled manner to absorb as much energy

as possible. Such deformation must be consistent

with the safety requirements of other installed systems such as fuel cells or seats and should not intrude into adjacent occupied areas.

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 rotaz/-wing aircraft

are summarized in Table 5. Any light aircraft designed to similar criteria would exhibit improvements in crashworthiness.

C A summary of desirable features for overall crashworthiness is

shown in Figure 11 for a single-rotor helicopter. Similar features must be implemented in all designs, whether fixed or rotary wing, to provide a survivable environment for all occupants.

When a more severe 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 provide such

protection, the design must permit 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 unac-

"* ceptable weight penalty, but on impact, high G levels that

compromise occupant survivability may be generated.

4.2 AIRFRAME CRASHWORTHINESS

The aircraft structure should provide a protective shell for

vehicle occupants in crashes of the severity cited in Table 2.

.59

,-:i L -li:--]i __ __ I _ .. ___iT r,, ___....I; 

1 g~j1~t.~. ... i ~ ~I
TABLE S. P3trIORMANCE REQUIREMENTS ?OR STRUCTURAL CRASHWORTHINESS

Velocity Vehicle Percentage

impact Impacted difforential attitude volume Other Date

direction surface* Ift/ioc) limits reduction requirements source

Longitudinal Rigid 20 No hazard Does not impede postcrash Volume II to pilot/ ogress

copilot

40 15 max. Inward buckling of side MIL-STD-1290

length ro- wells should not pose Volume II

duction for hazards

pass./troop

compartment

Lateral Rigid 30 f20* Yaw IS Max. Lateral collapse of oc- MIL-STD-1290

width cupled areas not haterd- Volume It

reductior. Ous. NO entrapment of

limbs.

Vortical Rigid 42 -25/-*l5* 15 max. G loads not iniurious to MIL-STD-1290

Pitch height red. occupants Volume I1

220, Roll in pass./

troop conpartmont

Resultant Rigid So Combination As above Max. velocity changesi MIL-STD-1290

for varicus long. * 5C ft/Soc Volume II

components vert. - 42 ft/sec lat. . 30 ft/seCb8

25 ft/Sec

Rollover Earth 90' sidewArd minimal forward fusol1ge buried to MIL-STO-1290

or 10S in- 4door depth of 2 in. finver-ed or

vetted or hatchos atc. on side). Load uniformly any inter_ assumed to distributed over forward

mediate anglo be non-load 250 of ocrupied fusolage

csrrying) length. Can sustain 4 C without injury to seated

and restrained occupants. SAll loading directions be- tween normal and parallel

to skin to be considered.

Rollover (post- Rigid Two 360 IS max. MIL-STO-129O

impact) rolls (msax.) volume ro- duction (St

* I desired)

Earth plowing Earth Preclude plowing when for- MIL-STo-1290 A scooping ward 2S& of fuselage has

tlongitudinal) uniformly applied vertical

load of 10 G and rearward

load of 4 G or the ditch- ing loads of MIL-A-00S6SA, whichever is the greatest.

Landing gear Rigid 20 lO1' Roll None. Pies- Aircraft deceleration at MIL-STD-1290

t10' Pitch tic deformas- normal G.W. for impact

tion of gear with no fuselage to ground

and mounting contact. All other A/C system al- Structural parts, except

lowable blades, should be flightworthy 'ollowing crash.

Landing gear Sod 100 long.€ -5* Fitch 15 max. No rollover, or if rollover MIL-STD-1290

14 vart. 110" poll volume re- occurs, two 360" rolls

120* Yaw duction (S5 without fuselage crushing Volume 1I desired)

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

b) Other helicopters.

C) Velocity at impact, not differential,

60

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These velocity changes are for the major impact, assumed to

occur on a rigid surface and with a triangular accelerationtime pulse shape. 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 that characteristically resist sparking caused by abrasion. Unless

otherwise stated herein, the aircraft basic structural design

gross weight (BSDGW) should be used for the vehicle weight in

the analyses described below. Directions are assumed with respect to the aircraft (Figure 1) unless otherwise stated.

4.2.1 Longitudinal Impact

4.2.1.1 Impact Conditions: The basic airframe should be capable 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 engine(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 should

be capable of impacting 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. 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 low angle, missed approach; the impact conditions of this type accident are illustrated in Figure 12. 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 accident experience. The volume of the cockpit for the occupied passenger/

troop compartment should not be reduced oy mce- than 15 percent (5 percent desired) for these conditioi

Should the aircraft turn over, the 'c .lu -ion of the

fuselage should maintain structural *..egrity to. a minimum of

two 360-degree rolls. The static 1L ads to be >- sidered for

rollover analysis are described in Section 4.2.4.

4.2.1.2 Earth Scooping: Design features for reducing the

earth scooping effects encountered in longitudinal impacts

should include the following:

62

S.. ... .. .. .. i
• "• IMPACT CONDITIONS

i. Soil of California Bearing Ratio = 2.5

F 2. Aircraft pitch (0) = 5 degrees nose down -

3. Aircraft roll (6) +10 degrees

4. Aircraft yaw (7) + +20 degrees 5. Flight path angle (a) - 8 degrees

6. Impact airspeed =60 knots re

Ground level

Figure 12. Low angle impact design conditions

(simulated approach with antitorque loss under poor visibility).

* A large, relatively flat surface should be provided1 in those areas which could otherwise gouge or plow, thereby increasing the aircraft's tendency to slide

over the impact terrain.

* Inward buckling of the fuselage nose or engine nacelle should be minimized 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 13.

4.2.1.3 Fuselase Deformation: To minimize hazards to person- nel created by buckling or oter deformation of the structure,

the aircraft should be designed to; 4

9 Provide sufficient strength of structure to prevent 4 bending or buckling failure of the fuselage in accord

with Table 5.

63

_ _ _ _ __, ;V li
G t

10 G

Figure 13. Nose section design conditions.

•Position personnel away from likely fuselage frac- :

ture areas.

"• Buckle the fuselage outward, if at all possible, rather than inward into living space when its col- :

lapse strength has been exceeded.

" Provide sufficient strength and rigidity in structure surrounding exits to ensure their postcrash operability in accordance with the criteria presented

in Chapter 6.

•Include cargo tiedowns that will restrain cargo

should fuselage bending failure occur.

4.2.1.4 Floor and Bulkhead: The floor structure should possess sufficient strength to carry, without failure, loads apS~plied by the occupant and cargo restraint systems in impacts

Sof the severity cited in Table 2. Considerations imust be made

for the specific loads and moments applied by these item6 to

the supporting structure in the warped conditions described in

Chapter S.

4.2.2 Vertical Impact

4.2.2.1 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. i

64

i .,. 5 •, ,.
4.2.2.2 Design Application: Design applications for accom-

"plishing the above goal should include the following:

o To the greatest extent feasible, locate massive items

in lower areas of the fuselage rather than in the upper areas.

* Increase cockpit and cabin vertical strength and

stiffness to prevent the structure from crushing the

occupants.

e Provide crash-force attenuating structure beneath

cockpit/cabin flooring.

* * Provide load-limiting landing gear capable of absorb-'

ing as much of the crash energy as practical.

4.2.3 Lateral Impact

The aircraft should have the capability to withstand lateral

impacts into a rigid barrier/wall of 25 ft/sec for light fixedwing and cargo and attack helicopters and 30 ft/sec for other

rotary-wing aircraft without reducing the width of the 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.

4.2.4 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 in soil to a depth of 2.0 in.

for the inverted attitude, and that the load is uniformly distributed over the forward 25 percent of the fuselage length.

It should also be assumed that the forward fuselage side is

buried in soil to a depth of 2.0 in. for the sideward attitude, and that the load is uniformly distributed over the forward 25 percent of the 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 14 and 15 respectively, without permitting sufficient deformation to cause

65
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injury to seated, restrainnd occupants. For both cases in Figt ures 14 and 15, 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 (i.e., shear loading). When designing for

this condition, it should be assumed that all doors, hatches,

transparencies, and similar openings cannot carry #ny loading.

4.2.5 Wings and Empennage

As discussed in Section 4.1, the wings and empennage structure

should remain attached during a crash. However, in the event

of high concentrated loads where failure is inevi'.able, their

structures should be designed to ensure that failure occurs

outside the occupant-protecting section of the fuselage.

The adjusted position of control surfaces such as flaps should

not block doors or other escape routes from the aircraft.

4.2.6 Engine/Transmission Mounts

For light fixed-wing airc , mounts on the engine and on the

supporting structure sh.. 'e designed to keep the engine attached to the basic sup, .ing structure under the crdsh conditions cited in Taole 2, even if considerable distortion of

the mounts and supporting 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 casings 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 manner hazardous to the occupants during the following impact con- ditions:

* 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-ir.-diameter rigid

cylinder.

Unless otherwise specified, all engines, transmlissions, rotor

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

be designed to withstand the following ultimate load factors

(G) and remain restrained:

68
* Applied Separately

Longitudinal ±20

Vertical +20/-10 Lateral ±18

* Applied Simultaneously

Design Conditions I

1 2 .3

Longitudinal t20 ±10 ±10

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

Lateral t0 t9 ±18

4.2.7 Shape of Fuselage Cross Section

The shape of the fuselage has an inherent influence on its re- I

sponse to the crash environment. 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 elli.psoidal fuselage will result in lower rollover

loads than would a flat-sided fuselage under identical condi- I tions. Even though operational considerations may prevent the

use of an exact ellipsoid-shaped fuselage, an approach to this

shape is a worthvhile design goal.

4.2.8 Landing Gear

The landing gear geometry should be such that no abnormal characteristics result from aircraft taxis, takeoffs, and landings

at the basic stiuctural design gross weight on terrain with

slopes of up to 12 degrees, or from landing sideways on a 15-

degree siupe undcr zero wind. The eink speed should not exceed 6 ft/sec for the above slope conditions. A differential kieeling landing system should not be utilized to satisfy this requiremnnt. These requirementL should be met regardless of the

orientation of the sloped site relative to the aircraft. 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

AMCPT06-201 (Reference 29).

29. ENGINEERING DESIGN HAYDBOOK, HELICOPTER ENGINEERING, Part

One, PRELIMINARY DESIGN, AMC Pamphlet 706-201, U. S. Army

Matcriel Command, Alexandria, Virginia, August 1974.

69

n. n :• a :• --i.-.- - -r - J ,,,, , n •
r

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 trucks).

4.2.8.1 Tail Bumper: Tail bumper wheels or skids should be

provided as necessary. Skids should have a simple, hardenedsurface, replaceable shoe to absorb the wear and damage of im-

* pact.

4.2.8.2 Ground Clearance: The ground clearance, with aircraft 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, The clearance should not be less than

6 in. with the aircraft in any of the following attitudes:

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

and all tires flat.

e 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.

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

and nose gear at maximum extension. The longitudinal

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.

4.2.8.3 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 failing within the

crash conditions of Table 5. If this cannot be accomplished

70
by location, the gear should be designed to break away under

longitudinal impact conditions, with points of failure located I

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.

4.2.8.4 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

percent 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 statically tested to failure, or which are not procured to MIL-A-21180. The yield factor of safety should be 1.0.

4.2.8.5 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 I conditions defined in Table 2. Forward and aftward motion of

the wheel in wheel-type landing gear of the trailing-arm 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 

face without 

impact 

allowing 

velocity 

contact 

of 20 

of 

ft/sec 

the fuselage 

onto a level, 

proper 

cigid 

with 

surthe

I

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 degrees roll and pitch.

4.3 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

71

ý 6.
equipment includes emergency and survival equipment, aircraft

subcomponents, and miscellaneous equipment. Typical items in

each of these categories are:

"* Emergency Equipment

Oxygen bottles

Fire extinguishers

First aid kits

Portable searchlights

Crash axes

"* Survival Equipment

Survival kits

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

* Miscellaneous Equipment

Navigation kits

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.

72
4.3.1 Strength

* Restraint devices and supporting structure for ancillary equipment should be designed. to restrain applicable items when exF posed 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 equipment. Loadlimiter stroking should not allow equipment to enter an occupant strike envelope.

4.3.2 Emergency and Survival Equipment Stowage Location

Equipment should be: (1) located close to the primary crew

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

and passengers; 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 potential adverse effects of extreme temperature, abrasion, and uncleanliness.

4.3.3 Retention 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 tools 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

lifted free. If equipment is stowed in an enclosure, no more

than 5 sec should be required for opening the enclosure and removing the equipment. Aircraft attitude should not adversely affect release device operation. It should be possible to 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 light 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.

4.4 INTERFACE OF OCCUPANT AND CARGO RETENTION SYSTEMS WITH

AIRFRAME

Both seats and cargo tiedowns require structural attachments

capable of withstanding the applied loads without failure or

excessive deformation. Although additional seat design and

installation requirements are discussed in Chapter 5 of this

73

&
volume, there are several important points to be considered *

where structural interface occurs. For example, the basic

floor structure should evenly distribute loading to the underfloor frames and longitudinal members. All seat and cargo attachment fittings should be attached through the floor to primary underfloor structure; i.e., either the heavy, full-depth

longitudinal beams or substantial underfloor frame elements.

The elements should be compatible with the types and magnitudes of crash loading applied by the seat or cargo attachments. This includes reaching the loads and moments applied

by the seats or cargo with deformed floor and bulkhead structure.

The tiedown points must be designed for the worst case combination of cargo weight, center-of-gravity height above the floor,

and G environment during the crash.

If energy absorbers are used for the seat or cargo attachments,

the attachments and their fasteners should be designed to the

limiting load condition, considering the effects of angular

displacement relative to the floor. To ensure structural integrity, all seat attachments must be designed to withstand or

attenuate computed maximum loads with consideration for bottoming, or exhausting of available stroke. In the case of tiedown rings, which usually are rated to a certain load capability such as 5,000 lb, the attachments and structures must be

capable of withstanding the worst case, angled load without

yielding. Although cargo tiedown energy absorbers may be used,

if a choice exists between energy-absorbing and nonenergyabsorbing tiedowns, the design criteria must be for the worst

case, which will likely be the nonabsorbing equipment.

Structure surrounding an energy-absorbing seat must be designed

to allow clearance for seat operation. Elastic deformation

should be added to the envelope of seat stroke in determining

the required clearance. If a well is provided in the aircraft

floor to allow additional stroking distance, at least a 2-in.

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.

4.5 CARGO RETENTION

Cargo restraint should:

* Be as light in weight as possible.

* Require minimum storage space when not in use.

* Be easy to install and remove.

74

L
* Be easily and reliably adjustable for different r sizes and shapes of cargo.

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

the severity described in Table 2.

0 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. Complete load and displacement

requirements are presented in Table 6, and the requirements for the longitudinal and lateral directions are illustrated in

Figures 16 and 17.

TABLE 6. CARGO RESTRAINT LOADS AND

DISPLACEMENT REQUIREMENTS

Load direction

Item (with respect) Restraint Controlled

no. to floor) load displacement

1 Forward See Figure 16 See Figure 16

2 Aftward 5 G No requirement

3 Lateral See Figure 17 See Figure 17

4 Downward 16 G No requirement

5 Upward 5 G No requirement

6 Forward ) See Figure 16 See Figure 16

and Combined

Lateral 4 G No requirement

Nets used to restrain small bulk cargo should be constructed of

material with low-elongation characteristics in order to reduce

dynamic overshoot to a minimim. Restraining lines without load

limiters used for large cargo, as defined in Table 7, for longitudinal restraint should be so arranged that maximum loadelongation characteristics are not used on the same piece of

75

I-- ... i i li____III__liliI -- ---- .-- , .. ..
Permissible

controlled

X displacement

S~Aircraft

Controlled

oroor'X

Aircraft displacement floor devices

Net restraint Line restraint

20 I 1 1

Practical cargo displacement limit--

18 (depending on aircraft)

16

0 Acceptable failure area S14 

0

I.iw acceptable lodcurve

10 SSample EA

8 Restraint curves

-Lower(bs)c

4 . .

S.... Acceptable performance

1 5 10 0 15 20 25 30

Controlled forward cargo displacement (X), in.

Figure 16. Load-displacement requirements for energy-absorbing

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

76

S- T -A..,
t7 Failure load .--- Unacceptable performance

- Acceptable performance

• 10 Minimum Practical cargo

acceptable displacement

0 8 

o • load curve -- I on limit aircra (depending ft)-----p.

SAcceptable failure area

............... ........... .. ....

0 5 10 15 20

Controlled lateral cargo displacement, in.

Figure 17. Cargo lateral load-displacement

requirements.

TABLE 7. AIRCRAFT CARGO CATEGORIES

Small bulk cargo Large rigid cargo

(net restraint) (line restraint)

This class includes all boxes This class insludes all rigid

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

* mately 3 ft or less in size.

SExamples: 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 maintenance 5. Special weapona

components (priority cargo)

*77

k7.
cargo. If load limiters are used, restraining lines should be

metal cables with low-elongation characteristics to ensure the

most efficient energy absorption.

4.6 TESTING

4.6.1 Aircraft System Testing

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

(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.

4.6.2 Landing Gear Crash Testing

Instrumented 4rop tests should be conducted: (1) to verify

landing gear trash force attenuation and crash loading strength

characteristics analytically predicted and (2) to substantiate

the capabilitf of the aircraft :anding gear to meet the criteria of Section 4.2.8. Drop testing of wheel and skid landing gear shou d be conducted in accordance with paragraph 9-2.3

of AMCP 706-2 3 (Reference 30) and should include demonstration

of complianceiwith the reserve energy and crash impact requirements of Sect on 4.2.8. The 20-ft/sec sink speed drop test

should be con ucted with the landing gear oriented in a 10-

degree nose dbwn 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.

4.6.3 Cargo Restraint

Design loads are specified in Section 4.5. 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.

4.6.4 Ancillary \Equipment Retention

Design loads are pecified in Section 4.3. Static tests to

these loads are renmmended.

30. ENGINEERING DESIGN HANDBOOK, HELICOPTER ENGINEERING, Part

Three, QUALIFICATION ASSURANCE, AMC Pamphlet 706-203,

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

1972.

78
Yes No NAL

4.7 DESIGN CHECKLISTS

4.7.1 Landing Gear Design Checklist

I. Will the gear withstand an impact velocity

of up to 42 ft/sec without catastrophic

failure?

2. Will the gear prevent the fuselage from

contacting the ground in a 20-ft/sec impact?

3. Will the gear survive a 10-ft/sec impact

without structural damage?

4. Will the gear remain attached to the

fuselage after impact? -

5. Is the gear located to prevent penetration

of occupied areas during the energyabsorbing stroke or in the event of gear failure?

6. Has the gear been designed to absorb the

maximum energy consistent with available

stroke?

7. Is the gear located to prevent rupture of

fuel cells?

8. Is every blow-off valve located where fluid

will be confined or ejected outside the

aircraft?

9. Has the gear been designed to avoid interference with the stroke of energy-absorbing

seats?

S 4

4.7.2 Airframe Design Checklist

4.7.2.1 Fuselage

1. Are forward bulkheads canted aftwards below

the floor to prevent earth scooping?

2. Are the forward lower skin panels made of

tough, yet ductile, material to minimize

tearing?

79

-00-,
Yes No N/A

3. Are the forward lower skin panels shingled

aftward to prevent scooping?

4. Will the nose structure cupport an upward

load of 10 G and an aftward load of 4 G applied over the forward 25 percent of the

fuselage without failure that would increase

earth scooping tendencies?

5. Is the underfloor structure designed for

energy-absorbing crush under upward loading while remaining intact under longitudinal impact conditions?

6. Is structure designed to transfer loads due

to overhead masses to floor level without

hazardous crushing of the occupied volume?

4.7.2.2 Wing and Empennage

7. Will the loss of wings occur in a manner

that does not endanger the occupants and

that does not destroy the usable volume?

4.7.2.3 Rollover Structure

8. Will the forward fuselage roof support a

4-G load?

9. Are the side frame members designed for high

load capacity to prevent collapse during a

rollover-type impact?

4.7.2.4 Blade Impact Protection

10. Are overhead longitudinal members extended

continuously over cockpit areas?

11. Are upper surfaces smooth and is lateral

structure angled to deflect passing blades

rather than allow penetration?

80

L

i . . . . . '' ' - .•, , . ,,. -- - = -, . . .
p.I

Yes No N/A

4.7.2.5 Heavy Mass Support

13. Are the supports for massive overhead components designed to v'ithstand the following

loads:

±18 G lateral?

t20 G longitudinal?

+20/-10 G vertical? -

14. Will the supports for massive ovc. ic6

components withstand the following combinations of loads:

±20 G long., +10/-5 G vert., ±0 G lat.?

±10 G long., +20/-10 G vert., ±9 G lat.?

±10 G long., *10/-5 G vert., ±18 G lat.?

15. Do the engine mounts and fittings, integral

to the engine as well as the aircraft structure, have sufficient strength to remain intact until after failure of major structural

supporting members?

4.7.2.6 Fuel Cell Installation

16. Are fuel cells locat-d above floor level and

away from possible impact surfaces?

17. Are fuel cells located as far from occupiable areas as reasonably possible?

18. Is fuel containment assured for all anticipated survivable impacts?

9. Is the structure that supports fuel cells

smooth and clean of projections to provide

uniform support and avoid puncture?

20. Are frangible and self-sealing couplings

used in fuel lines where relative displacements of structure may occur?

21. Are fuel cells located outside the likely

landing gear motion envelope?

22. Have checklists of C•.•ptcr 6 oeen r- ',Lrr,!d

to for fuel systkm .A '1?

81
Yes No N/A

4.7.2.7 Seat and Cargo Installation

23. Is structure around'seats designed to avoid

interference with seat stroking and has suf- ficient clea~rance been allowed to enable efficient seat design (see Volume III)?

24. Are seat and cargo attachment fittings

secured through the floor to primary structural members?

25. Are tiedown points designed for the worst

case combination of cargo weight, center of

gravity height above the floor, and direc- Stions of loading and structural deflection? •

26. Have checklists of Chapter 5 been referred

to for seat system design? -

4.7.2.8 Emergency Egress

27. Has the structure surrounding emergency exits been designed for minimum distortion?

28. Have the egress checklists of Chapter 6 been

referred to for emergency egress requirements?

t!

82

_____________
F

5. AIRCRAFT SEATS, RESTRAINTS, LITTERS, AND PADDING

5.1 INTRODUCTION

This chapter summarizes the criteria for including crashworthiness into the design of aircraft subsystems that interface

directly with the occupants. These subsystems include restraint systems, seats, litters, cockpit controls, and padding

materials. The user is referred to Volume IV for additional

information concerning the criteria and their sources.

It is important to remember the basic operational difference

between passenger seats and crewseats. The primary function of

passenger seats and litters is to provide a place for aircraft

occupants to sit or lie during their transport, while the crewseats must provide the comfort, adjustments, and features that

aid crew members in accomplishing their operational responsibilities. These functional requirements obviously are of highest priority; however, crashworthiness and the ability of the

subsystems to help protect the occupant during crashes are also

of extreme importance and can be accomplished without significant degradation of comfort and operational aspects.

5.2 PRIMARY DESIGN CONSIDERATIONS

5.2.1 General

Occupant protection and survival in aircraft accidents should

be a primary consideration in the design, development, and

testing of aircraft seats and litters. All operational requirements as specified in other design guides should also be

met. Adequate occupant protection requires that both seats

and litters be retained generally in their original positions

within the aircraft throughout any survivable accident. In

bldition, the seat should provide an integral means of crash

load attenuation, and the occupant's strike envelope should be

delethalized.

Several environmental and operational factors other than those

associated with crashworthiness affect the design of an adequate seating system. They are very important in overall design, and are discussed in Section 3.2 of Volume IV.

5.2.2 Design Conditions and Envelopes

The design impact conditions for light fixed- and rotary-wing

aircraft are presented in Volume II and are repeated in Chapter 3, Table 2 of Volume I. All seats, restraint systems, and

litters should be designed to provide the desired performance

in the design claoh environments. It must be remembered that,

83
r

to produce a truly crashworthy design, systems analyses must

consider likely combinations of loadings, including potential

losses of energy-absorbing strvcture, such as landing gear.

5.2.3 Structural Distortion

SStructural distortion of the airframe and its resulting loading of the seat must be considered in the design. A major consideration in providing crashworthy seating systems is the possibility of a local distortion in that part of the aircraft to

which the seat is attached.

In ceiling-mounted seats the efficiency of use of the available

stroke distance must be considered. Energy-absorbing stroke

should be provided to maximize usage of the available space,

but the effective stroke of a seat considered to be rigidly attached (no energy absorbers between the seat and roof) to the

roof must be considered. The roof may deflect downward at

loads too low to make efficient use of the available stroke, a

particular concern for retrofit applications to older aircraft.

A systems analysis should be used to evaluate the advisability

of using ceiling-mounted seats in this situation and if so, establish the correct combinatioi, of variables.

A considerable amount of the downward motion of an aircraft

ceiling may be elastic. It would be advantageous to eliminate

from the occupant and seat the rebound due to recovery of this

elastic distortion. Consideration should be given to a device that allows vertical downward motion of the seat but restrains it from following the roof during its elastic rebound.

Adequate support of the ceilinq to support the applied loads

with low deflections eliminates the problems mentioned above,

and efficient use of ceiling-mounted seats can be achieved in

aircraft with such features.

Considerations for seats mounted on the floor, bulkhead or

sidewall, including requirements necessary for the attachments

to survive fuselage warpage, are presenced in Section 5.4.5,

Joint Deformation.

5.3 DESIGN PRINCIPLES FOR SEATS AND LITTERS

5.3.1 Seating System Orientaticn

There are several types of Army aircraft seating systems: pilot, copilot, crew chief, gunner, observer, student, medical

attendant, troop, and passenger. Cockpit seats are typically

forward-facin&i however, cabin seats may face in any direction.

84

I• ••
Most are single-place seats, but in a few aircraft, two-,

three-, and four-occupant cabin seats are provided. A singleoccupant seat is the preferred configuration in order to avoid

situations in which the energy-absorbing systems of multi-unit

seats are rendered ineffective due to partial occupancy (insuf-

* ficient weight to activate the energy-absorbing mechanisms at

loads within human tolerance limits). To the maximum extent

practical, seats should be interchangeable to enable standardization. It is desirable that all seats face in the same direction so that the seat backs protect occupants from loose

equipment which can become projectiles during crash impact.

The rearward-facing seat is optimal for providing maximum support and contact area in longitudinal impacts. The only critical impact sequence for the rearward-facing seat is one that

involves a severe lateral component that allows sideward movement of the occupant prior to application of the longitudinal

or vertical pulse. However, lateral torso movement can be minimized by use of an adequate restraint system of much lighter

weight than that required for other seat orientations. When

practical, the rearward-facing seat should be used.

Those crew members required to face forward in the conduct of

their duties can be afforded adequate protection by the use of

a restraint system consisting of shoulder straps, a lap belt,

and a lap belt tiedown strap as discussed in Section 5.7. Lapbelt-only restraint-is undesirable, as noted inthe human tolerance section of Volume II. If all forward-facing passengers

are provided with adequate upper- ani lower-torso restraint,

forward-facing seats are acceptable as a second choice to

rearward-facing seats. If a single, diagonal upper-torso restraint is used, it should be placed over the outboard shoulder of the occupant to provide restraint against lateral pro- trusion of the occupant outside the aircraft or impact with

the sidewall.

Previously, side-facing seats have been provided with lap belt

restraint only. This arrangement is considered completely inadequate for providing crash protection. Even with the addition of a shoulder harness or diagonal chest strap, the tolerance to abrupt acceleration is minimal. The use of side-facing

seats is least desirable from the crash safety standpoint;

however, when no reasonable alternative to their use exists,

adequate restraint must be.provided. If a single, diagonal,

upper-torso restraint is used, it should be placed over the

forward-facing shoulder (relative to the aircraft).

5.3.2 Litter Orientation

Litters should be installed laterally to provide more positive

restraint for expected combined crash forces. A lateral litter

85

Sml =• •• • > •:•| • • " _ _ _ __,_-_ _ _ l
orientation also will prevent detachment of the litter from its

supports, which may occur as explained in Reference 31. The

litter must withstand all of the environments previously described for seats.

5.3.3 Materials

Designers should select materials that offer the best strengthto-weight ratios while still maintaining sufficient ductility

to prevent brittle failures.

The degree of ductility needed in a seat's basic structural

elements is highly dependent upon whether the seat structure

is designed to absorb energy by the use of a separate loadlimiting device or whether large plastic deflections of the

basic structure are required. As a general rule, a value of

10-percent elongation is a rough dividing line between ductile

and nonductile materials. The 10-percent value is recommended

as a minimum for use on all critical structural members of

nonload-limited seats because the exact peak load is unpredict- able due to pulse shape, dynamic response of the system, and

velocity change. A minimum elongation of 5 percent in the

principal loading direction is suggested for use on critical

members of load-limited seats because the loads and strains

are more predictable. Also, castings are not recommended for

use in primary structural load paths.

The effects .f stress corrosion must be considered, as well as

hydrogen embrittlement due to heat treating or various proces- sing steps such as pickling. In short, adherence to all the

normal engineering design principles must prevail.

Flammability and toxicity retardation requirements are discussed in Chapter 6. Upholstery padding and other materials used

in seats should meet the specified requirements.

5.4 STRUCTURAL CONNECTIONS

5.4.1 Bolted Connections

For the manufacture of basic aircraft structure, most aircraft

companies recommend 15- and 25-percent margins of safety for

shear and tensile bolts, respectively. The margin of safety

for shear and tensile bolts located in load-limited portions of

31. Weinberg, L. W. T., AIRCRAFT LITTER RETENTION SYSTEM DESIGN CRITERIA, Aviation Crash Injury Research (AvCIR), Division of Flight Safety Foundation, Inc.; USAAVLABS Technical Report 66-27, U. S. Army Aviation Materiel Laboratories, Fort Eustis, Virginia, April 1966, AD 632457.

86
f

the seat where loads can be predicted accurately, can be reduced to 5 and 10 percent, respectively. Also, good aircraft

engineering practice dictates that bolts less than 0.25 in. in

diameter should not be used in tensile applications because of

the ease with whicb these smaller bolts can be overtorqued. Because of the obvious advantages of structure being able to

distort while maintaining load-carrying ability, fasteners of

maximum ductility for the application should always be selected.

Where possible, fasteners such as bolts and pins should have

a minimum elongation of 10 percent. A bolt loaded in shear

should have a shank of sufficient length to prevent application of the shear load on the threaded portion of the bolt.

5.4.2 Riveted Connections

Guidelines for riveted joints are presented in MIL-HDBK-5, and

it is recommended that these guidelines be followed (Reference 32).

5.4.3 Welded Connections

Welded joints can be completely acceptable and even superior

to bolted or riveted joints. However, strict inspection procedures should be used to ensure that welded joints are of good quality. The cross-sectional area of the basic material in the

vicinity of a welded joint should be 10 percent greater than

the area needed to sustain the design load. Welding processes

are discussed in Military Specifications MIL-W-8604, -6873, -45205, and -8611; these specifications should be used as

guides to ensure quality welding.

5.4.4 Seat Attachment

Acceptable means of attaching seats to the cabin interior are

listed below. (Refer to Section 3.3.3 of Volume IV for a discussion of ceiling-mounted seats and ceiling support stiffness.)

1. Suspended from the ceiling with energy absorbers,

and wall or bulkhead stabilized.

2. Suspended from the ceiling with energy absorbers,

and floor stabilized.

3. Wall or bulkhead mounted with energy absorbers.

32. Military Handbook, MIL-HDBK-5C, METALLIC MATERIALS AND ELEMENTS FOR AEROSPACE VEHICLE STRUCTURES, Department of Defense, Washington, D. C., 15 September 1976.

87
gri

4. Floor mounted with energy absorbers.

5. Ceiling and floor mounted (vertical energy absorbers

above and below seat).

Suspension or mounting provisions for all seats should not interfere with rapid ingress or egress. Braces, legs, cables,

straps, and other structures should be designed to prevent

snagging or tripping. Loops should not be formed when the restraint system is in the unbuckled poptation. Cabin seats must

often be designed so that they may be qu;-Kly removed or folded

and secured. Tools should not be required for this operation.

The time required by one person to disconnect each single occupant seat should not exceed 20 sec. The time required by one

person to disconnect multi-occupant seats should not exceed 20

sec multiplied by the number of occupants. All foldable seats

should be capable of being folded, stowed, and secured or unstowed quickly and easily by one person in a period not to exceed 20 sec multiplied by the number of occupants.

5.4.5 Joint Deformation

To prevent seat connection failures induced by fuselage distortion, structural joints should be capable of large angular displacements in all directions without failure. A floor-mounted

seat designed properly for structurally integral load limiting

would also satisfactorily accommodate floor buckling and warping under crash conditions. Figure 18 illustrates recommended

limits of floor warping or buckling that must be withstood by

all floor-mounted seat designs. The mounts should be capable

of withstanding a ±10-degree warp of the floor, as well as a

tl0-degree rotation about a roll axis of a single track. The

angles are based on distortions that have been noted in potentially survivable accidents.

The same general principles that apply for floor-mounted seats

also apply for bulkhead-mounted seats except that the deflection and degree of warping of the bulkhead appear to be less

than those of the floor. A possible bulkhead distortion configuration is shown in Figure 19. The recommended angular deflection requirement for bulkhead-mounted seats is a 5-degree

rotation in the plane of the bulkhead. To accommodate local

deformation, each attachment of the seat to the bulkhead should

be released to permit ±10-degree rotations in any direction.

Sidewall-mounted seats require the same considerations as

bulkhead-mounted seats. The sidewalls of aircraft tend to bow

outboard during impacts with high vertical loading. Therefore,

it is advisable that these seats be designed to accept relatively large distortions without failure.

88
Fi I~i

S-+I± 0* m in •

Iy

FFigure 18. Static test floor warpage requirement to improve

• the probability of seat retention in crashes.

Seats that are mounted totally on the sidewall should not

create a problem, as they will simply move with the sidewall.

Extremely flexible seats also should be inherently immune from

•.. these problems. However, rigid seats mounted to both the floor •

8

.- .. _ :

89-

1Y
Bulkhead

FSeat attach

points (4 places)

Lower fuselage

(a) Initial configuration

Bulkhead

5 0 ,

earth, • or other

obstacles

(b) Postcrash configuration

Figure 19. Bulkhead in-plane warping.

I9

90
and the sidewall will require special design considerations.

One way to provide the flexibility needed is to include releases such as pin joints, oriented to allow rotation around

an aircraft roll axis. An example is shown in Figure 20. The

attachments should be designed to permit the angle E to reach

25 degrees at the maximum dynamic deflection.

Initial fuselage

configuration

Deflected

fuselage

configuration

2 , 21._3

s? \ 

2 5 max 0 ...

Figure 20. Pin joint releases oriented to allow

rotation around an aircraft roll axis.

9

1 91

' 4: + • .• _ i _ :p1 _ + . _ . _ _ _ _ __. • _ _ _ IIi . . . . . • .'

Ip u~u~I
The underfloor, bulkhead, or sidewall structure must be designed to be compatible with the seat. For example, the design

of structural releases between the seat and the track may enable the seat to maintain its attachment during large floor

deformations but may add to the torsional responsibilities of

underfloor beams. If a large downward load is applied to the

floor structure through a joint that does not carry moment %.ileased), then the underfloor beams must resist any moment that

may be developed without assistance from the seat structure.

5.4.6 Material

5.4.6.1 General: An elastic stress analysis, as used in the

design of airframes and aircraft components subjected to normal

flight loads, is inadequate for the study of all the structure

in a crash situation. For normal flight loads, keeping the

stresses well below the material yield stress to avoid permanent deformation is necessary because of fatigue problems and

other considerations. In a crash situation, however, where

only one application of the maximum load is expected, fatigue

is not a factor, and the final appearance of a structural component or its subsequent operational use need not be considered.

Consequently, the load-carrying capacity of components deformed

beyond the elastic limit should be considered in determining the ultimate seat strength. For certain items in the load path

it is advisable to use the rupture strength as listed for many

materials in MIL-HDBK-5 (Reference 32). The concepts of limit

analysis or, in some circumstances, large deformation analysis

may be employed to make the best use of materials in certain

components.

5.4.6.2 Limit Analysis Concepts: Where ductile materials are

used, strain concentrations do not produce rupture prior to

significant plastic deformation. If the geometric configura- tion of the structure permits only small elastic deflections,

a "rigid-plastic" mathematical model may be used. This per- mits the use of a limit analysis, which assumes no defcrmation of structure until sufficient plastic hinges, plastic extensors, etc., exist to permit a geometrically admissible collapse mode.

Limit analysis is concerned with finding the critical load sufficient to cause plastic collapse with the physical requirements of static equilibrium, yield condition for the materials,

and consistent geometry considerations. Two useful principles

are mentioned here: the upper and lower bound theorems. The

upper bound theorem for the limit load (collapse load for

"rigid-plastic" structure) states that the load associa-4d

with the energy dissipated in plastic deformation will form an

upper bound for the limit load. The lower bound theorem states

92
-- ---

that the load associated with a statically admissible stress

* distribution, which at no point exceeds the yield conditions,

forms a lower bound for the limit load. Use of the upper and

lower bound theorems to bracket the limit load for a given

structure makes it possible to obtain a realistic evaluation

of the structure's load-carrying capacity.

5.4.6.3 Ldrge Deformation Analysis: If a structure contains

elements that will permit large, stable elastic deformations

when under load, the equilibrium of the deformed state must be

considered in evaluating ultimate strength. For example, if a

suitable attachment is made to a thin, flat sheet rigidly fixed

at the edges so as to load the sheet normal to the surface, a

diaphragming action will occur. The equilibrium and stressstrain (elastic-plastic) relations for the deformed state would

determine the load-carrying capacity. An example of this situation is a seat pan in which membrane rather than flexural

stresses are important.

5.4.6.4 Strain Concentrations: Handbook stress concentration

factors will provide sufficiently accurate data to allow the

designer to modify the structure in the vicinity of stress concentrations. When large deformations at high load-carrying capacity are desired, as in energy-absorbing seats, these areas

frequently become strain concentration points, and rupture occurs due to excessive etrain in areas with little deformation

and energy input. Large amounts of energy can be absorbed in

the structure only if large volumes of material are strained

uniformly.

5.4.7 Restraint System Anchorage

The seat designer must consider the effect of the anchorage of

the restraint system on the characteristics of the seat design. If possible, the restraint system should be anchored to the

seat rather than to basic structure.

If the harness is anchored to basic structure, a desirable reduction of loads on the seat frame results; however, the restrailit system must be designed to permit the energy-absorbing

dcformation of the seat during an impact. For example, if a

load-limited seat strokes vertically and the seat belt is anchored to the floor, loosening of the belt permits the occupant

to either submarine or move laterally under the belt. When the

harness is anchored to the seat structure, the problem of maintaining a tight harness is reduced.

93
i

5.5 ENERGY-ABSORBING DEVICES

The seat structure, in order to perform its intended retention

function, must possess either (1) the capability of sustaining,

without collapsing, the maximum inertial forces imposed by the

deceleration of the occupant and the seat, or (2) sufficient

energy-absorption capacity to reduce the occupant's relative

velocity to zero before structural failure occurs.* The first

alternative may result in an excessive strength requirement because the input pulse shape and the restraint system and

cushion elasticity can result in a large dynamic overshoot.

Computer simulation and experimental investigation have shown

that overshoot factors range from 1.2 to 2.0. This would necessitate a seat design strength requirement of 24 G to 40 G

to accommodate an input floor pulse of 20 G.

The second alternative of using collapse behavior (load limiting) appears to offer the more practical approach to most seat

design situations. With this option, the seat structure would

begin plastic deformation when the acceleration of the occupant

and seat mass reaches a level corresponding to the critical

structural load; the seat must absorb enough energy without

failure to stop the motion of the occupant relative to the aircraft. This energy must be absorbed at force levels within human tolerance limits to provide the intended protective function. The energy can be absorbed either by plastic deformatiozi

of basic structure or by the introduction of mechanical loadlimiting devices. Energy-absorbing motion of the seat can be

provided in all three directions as well as for all combinations of directions; however, it is absolutely necessary for

the vertical direction. A properly restrained occupant can

withstand the loads associated with the design environment in

the longitudinal (x) and lateral (y) directions but cannot sustain the loads in the vertical (z) direction without injury.

Therefore, the requirement for load reduction through use of

energy-absorption devices is mandatory for the vertical direc- S~tion.

Energy-absorbing mechanisms in aircraft structures which transmit crash forces to the occupant should stroke at loads tolerable to humans and should provide stroke distances consistent with these loads and with the energy to be absorbed.

*The term "failure" implies a rupture of restraint linkage,

while the term "collapse" pertains to a state of activu deformation with restraint integrity maintained.

94

|-|d--~--- - .. ___
Desirable features of energy absorbers are as follows:

"* The device should provide a predictable force-versusdeformation characteristic.

"* The rapid loading rate expected in crashes should

not cause unexpected changes in the force-versusdeformation characteristic of the device.

* The assembly in which the device is used should

have the ability to sustain tension and compression.

(This might be provided by one or more energy absorbers, or by the basic structure itself.)

* The device should be as light and small as possible.

* The specific energy absorption (SEA) should be high.

* The device should be economical.

* The device should be capable of being relied upon to

perform satisfactorily throughout the life of the

aircraft (a minimum of 10 years or 8000 flight hours)

without ziuiring maintenance.

* The device should not be affected by vibration, dust,

dirt, or other environmental effects. It should be

protected from corrosion.

* The device(s) should decelerate the occupant in the

most efficient manner possible while maintaining the

loading environment within the limits of human tolerance. A multiple-limit-load device, adjustable for

occupant weight, is desirable.

5.6 SEAT CUSHIONS

5.6.1 General

The seat bottom and back with which the occupant is in constant

contact should be designed for comfort and durability. Sufficient clearance between fabric backs and bottoms or sufficient

cushion thickness of the appropriate material stiffness should

be provided to preclude body contact with the seat structure

when subjected to either the specified operational or crash

loads. Seat bottoms made of fabric should be provided with

means of tightening to compensate for sagging in use.

For seat cushions, the problem is one of developing a compromise design that will provide both acceptable comfort and

safety. The optimum aircraft seat cushion should:

95
* Be extremely lightweight.

e Possess flotation capabilities.

* Be nonflammable.

9 Be nontoxic; not give off fumes when burned, charred, I

or melted.

* Be tough and wear resistant.

* Be easily changeable. I

9 Provide comfort by distributing the load and reducing

or eliminating load concentrations. I

* Provide thermal comfort through ventilation.

* Provide little or no rebound under crash loading.

* Allow an absolute minimum of motion during crash

loading.

5.6.2 Requirements

For seats of light movable weight (less than 30 lb), cushions

should be used for comfort only. The maximum uncompressed

thickness for a properly contoured cushion should be 1-1/2 in.,

unless it can be shown through analysis or through dynamic

tests that the cushion design and material properties produce

a beneficial (reduced force transmissibility) result.

For seats of greater movable weight, such as integrally armored seats, every effort should be made to design a cushion 1

that minimizes relative motion between the occupant and the

seat and that acts as a shock damper between the occupant and the heavy seat mass. Again, dynamic analysis and/or testing should be conducted to demonstrate that the cushion design

produces a desirable system result over the operational and .

crash environmental range of interest.

5.6.3 Energy-Absorbing Cushions

The use of load-limiting cushions in lieu of load-limiting

seats is undesirable. The only justifiable use of energyabsorbing cushions instead of load-limited seats might be in

retrofit circumstances where, because of limitations in existing aircraft, another alternative does not exist.

96

.4

Sl l _ _ _ _ _ _ _ I
5.6.4 Net-Type Cushions

This type of cushion serves the same purpose as the filled

cushion; however, a net material is stretched over a contoured

seat frame, and the body is supported by diaphragm action in

the net rather than by deformation of a compressible material.

The net-type cushion might more properly be called a net support. If a net support is used in the seat, its rebound characteristics should be capable of limiting the return movement

from the point of maximum deformation to 1-1/2 in. Net supports should not increase the probability of occupant submarining or dynamic overshoot.

5.6.5 Seat Back Cushions

The back cushion should be of a lightweight foam material or

net. The foam can be a standard furniture type that meets the

other requirements listed in Section 5.6.2. Lumbar supports,

particularly those that are adjustable by the occupant, are

desirable for comfort and because a firm lumbar support that

holds the lumbar spine forward in extension increases the tolerance to +G loading. z

5.6.6 Headrests

A headrest should be provided for occupant head/neck whiplash

protection. Headrest cushions arc used only to cushion head

impact and prevent whiplash injury due to backward flexure of

the neck. The cushioning effect can be provided by a thin pad

and a deformable headrest or a thicker cushion on a more rigid

headrest. For a rigid headrest, the provisions of Section

5.12 should be applied and at least 1.5 in. of cushion should

be provided if possible within the space limitations of the

application.

5.7 DESIGN PRINCIPLES FOR PERSONNEL RESTRAINT SYSTEMS

5.7.1 General

Restraint harnesses for personnel should provide the restraint

necessary to prevent injuries to all aircraft occupants in

crash conditions approaching the upper limits of survivability. Appropriate strength analysis and tests as described in Section

5.9 should be conducted to ensure that a restraint system is

acceptable.

97

¶j
Qualities that a harness should possess are listed below:

* It should be comfortable and light in weight.

e It should be easy for the occupant to put on and take

off even in the dark.

9 It should contain a single-point release system, easy

to operate with one (either) hand since a debilitated

person might have difficulty in releasing more than

one buckle with a specific hand. Also, it should be

protected from inadvertent release; e.g., caused by

the buckle being struck by a cyclic control or by inertial loading.

e It should provide personnel with freedom of movement

to operate the aircraft controls. This requirement

necessitates the use of an inertia reel in conjunction with the shoulder harness.

* It should provide sufficient restraint in all directions to prevent injury due to decelerative forces in

a potentially survivable crash.

* The webbing should provide a maximum area, consistent

with weight and comfort, for force distribution in

the upper torso and pelvic regions and should be of

low elongation under load to minimize dynamic overshoot.

5.7.2 Types of Systems

5.7.2.1 Aircrew Systems: The existing military lap belt and

"shoulder harness configuration with a center tiedown strap as

shown in Figure 21 is the minimum acceptable harness for use

by U. S. Army pilots. The configuration shown in Figure 22 is

preferred because it provides improved lateral restraint due

to the addition of the reflected shoulder straps. This syctem

resulted from the investigation reported in Reference 33. Details of the hardware in these systems are discussed in Section 7.5 of Volume IV.

33. Carr, R. W., and Desjardins, S. P., AIRCREW RESTRAINT SYSTEM - DESIGN CRITERIA EVALUATION, Dynamic Science, Division of Ultrasystems, Inc.; USAAMRDL Technical Report 75-2, Eustis Directorate, U. S. Army Air Mobility Research

and Development Laboratory, Fort Eustis, Virginia, February 1975, AD A009059.

98
r

Item identity 3A

1. Buckle A. Single-point assembly I

release buckle 3B

B. Tiedown strap

C. Tiedown anchor

2. Lap belt assembly

A. Lap belt

B. Adjuster

3. Shoulder harness

assembly

A. Inertia reel

B. Inertia reel strap

C. Lower shoulder D

strap

D. Adjuster 3

2V/I LI

Figure 21. Basic aircrew restraint system. •

•i 99

II

1I

- " I l l l I, I i I t • I i' •1 | .u.•• _I
Item identity

1. Buckle assembly

A. Single-point 4C

release buckle

B. Tiedown strap

C. Tiedown anchor B

2. Lap belt assembly

A. Lap belt B

B. Retractor

3. Shoulder harness

collar assembly

A. Pad

B. Roller fitting

C. Adjuster

D. Lower shoulder strap

4. Inertia reel assembly

A. Reflected strap

B. Anchor 3D

C. Inertia reel

(dual-spool)

I• 2B

ard

Figure 22. Aircrew restraint system, including

reflected shoulder straps.

100
5.7.2.2 Troop Systems: Considerations in the selection of a

troop or passenger seat restraint system are different from

those for an aircrew system. First of all, the seat may face

forward, sideward, or aftward. Secondly, the restraint system

must be capable of being attached and removed quickly in an

operational environment by troops encumbered by varying types

and quantities of equipment. Also, whereas a pilot probably

uses the restraint system in his aircraft so frequently that

its use becomes a matter of habit, troops and passengers can

be expected tc be unfamiliar with the system. The effects of

this lack of familiarity would probably become more pronounced in a combat situation when the risk involved in not using the restraint system becomes even higher. Therefore, hardware

should be uncomplicated arid, if possible, resemble the familiar, such as automotive hardware. Finally, the need to quickly

remove and stow the seats requires compact and ightweight restraint systems.

Two systems that resulted from the investigation reported in

Reference 34 are shown in Figure 23. The Type II troop restraint system is preferred and consists of a two-strap shoul- der harness anc a lap belt assembly. The two shoulder straps

are attached to two single inertia reels. They extend forward

and down over the occupant's upper torso and are connected into

the single-point release, lift-lever buckle. The lap belt assembly includes left- and right-hand belts, with adjusters,

that are connected together at the lap belt buckle. The Type

I troop restraint system is acceptable and differs from the

Type II restraint by having a single shoulder strap that passes

diagonally across the occupant's upper torso. For side-facing

seats it should pass over the shoulder closest to the nose of

the aircraft. If the Type I system is used in either a forwardor aft-facing seat, the diagonal shoulder strap should pass

over the outboard shoulder to restrain the occupant from protrudirg outside the aircraft during lateral loading.

5.7.2.3 Crew Chief and Door/Window Gunner Systems: Restraint

systems for crew chiefs and door/window gunners are similar to

troop systems; however, they must allow the crewmember to move

out of the seat to perform duties such as maneuvering the gun

or observing tail rotor clearance while landing in unprepared

areas. The system should restrain the occupant to the seat the

instant he returns to the seat and provide adequate restraint

during a crash. The system should maintain the lap belt buckle

34. Carr, R. W., HELICOPTER TROOP/PASSENGER RESTRAINT SYSTEMS

DESIGN CRITERIA EVALUATION, Dynamic Science, Division

of Ultrasystems, Inc.; USAAMRDL Techniical Report 75-10,

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

Development Laboratory, Fort Eustis, Virginia, June 1975,

AD A012270.

101

"• i• i•[ • • • 1 ...... .....----.. T-- -- 7 -,7 " ./ ••--i- -.. -! . I i 
1 -i

2

2

SA ÷

i ,/,

3.

'3

6 -5

Type I Type II

Item identiLy

1. Inertia reel

2. Shoulder strap 3. Lap belt anchor

4. Buckle with shoulder strap connection

5. Lap belt

6. Adjuster/fitting

Figure 23. Aircraft troop/passenger restraint systems.

1I

, 102
in the proper relationship to the gunner, preventing the shoulder straps from pulling it up or the lap belt from pulling it

sideways. Such a system has been described in Reference 35

and is shown in Figure 24. It consists of a lap belt with inertia reels on each side of the seat and two shoulder straps

connected in an inverted-Y arrangement to a single inertia reel

strap. The lap belt with thigh strap attachment is easy to put

on and prevents the lap belt from riding up during operation of

the gun. The lap belt is plugged into the two seat pan inertia reels when the crewmember is to be seated or is standing

in front of the seat. The shoulder harness and lap belt with

thigh straps may serve as a "monkey harness" when the crewinember disconnects the two lap belt plug-in fittings from the inertia reels. The resultant configuration permits the crewmember more extensive travel within the cabin while still being

connected to the shoulder harness inertia reel, thereby restraining the crewmember from falling out of the aircraft.

5.7.2.4 Inflatable Systems: An automatically inflatable body

and l-Wad res-traint system, IBAHRS, for helicopter crewmen has

been jointly developed and tested by the Naval Air Development

Center and the Applied Technology Laboratory. As illustrated

in Figure 25, this system provides increased crash protection because it provides automatic pretensioning that forces the

occupant back in his seat, thereby reducing dynamic overshoot

and reducing strap loading on the wearer when the inflated restraint is compressed during the crash. The concentration of

strap loads on the body is reduced because of the increased

bearing surface provided by the inflated restraint, and both

head rotation and the possibility of whiplash-induced trauma

are also thus reduced.

Although more complex and costly than conventional belt systems, such a system may be justified because of its occupant

protection potential. Development of the system and results

of testing are documented in References 36 and 37.

35. Reilly, M. J., CRASHWORTHY HELICOPTER GUNNER'S SEAT INVESTIGATION, The Boeing Vertol Company; USAAMRDL Technical Report 74-98, Eustis Directorate, U. S. Army Air Mobility Research and Development Laboratory, Fort Eustis,

Virginia, January 1975, AD A005563.

36. Schulman, M., and McElhenney, J., INFLATABLE BODY AND HEAD

RESTRAINT, NADC-77176-40, Naval Air Systems Command, Department of the Navy, Washington, D. C., September 1977.

37. Singley, G. T., III, TEST AND EVALUATION OF IMPROVED AIRCRAFT RESTRAINT SYSTEMS FOR COMBAT HELICOPTERS, Paper No.

A.18, presented at NATO/AGARD Aerospace Medical Panel,

Aerospace Specialist's Meeting on Aircrew and Survivability, North Atlantic Treaty Organization, Bodo, Norway,

May 20-23, 1980.

103

' L 7 7 7iiiIi '~ __
I3

5

6 -

Item identity

1. Inertia reel

2. Shoulder strap

3. Shoulder strap adjuster 8

4. Attachment release buckle

5. Lap belt

6. Lap belt inertia reel (\

7. Thigh straps

8. Thigh strap adjuster

9. Lap belt plug-in fitting

Figure 24. Gunner restraint system. (From Reference 35)

104

-- - - -- ------.---
411

U)I

4.4

1051
5.7.3 General Design Criteria

5.7.3.1 Comfort

Comfort must not be compromised by crash-survival requirements

for obvious reasons. The main comfort consideration for restraint harnesses is the absence of rigid hardware located over

bony portions of the torso. Also, hardware assemblages that

are too wide or large, or are not configured efficiently to fit

the desired location on the body could be uncomfortable. Webbing that is too wide or too stiff could also cause discomfort

through creasing of the webbing or perspiration due to reduced

ventilation.

5.7.3.2 Emergency Release Requirements: From a crash survival

point of view, it is mandatory that a shoulder harness/lap belt

combination have a single point of release that can be operated

by one (either) hand to make it easier for debilitated occupants

to quickly free themselves from their harnessing after a severe

crash because of the dangers of postcrash fire or sinking in

water. The force required to release the harness with only one

finger should fall between 20 and 30 lb on the basis of existing requirements for military harnesses. Further, the release

should be possible with the weight of the occupant hanging in

the restraint system after experiencing the full crash loads.

The release forces for the inverted case should be minimized

and, in any case, should not exceed 50 lb applied with only

one finger. It should be possible to produce the torque necessary to release rotary buckles by applying a load at a single

point on the handle as described above.

In restraint systems other than the Type I of Figure 23, if a

lift latch or similar type buckle is used, the restraint system design should ensure that the latch lifts from left to

right on all installations. This will reduce the possibility of reverse installations and their resultant hazard.

The release device must either have the capability to withstand the bending moments associated with deflections and motions during loading, or it should contain features that allow

the fittings to align themselves with the loads, thereby reducing or eliminating the moments. If belt loading direction is

such as to cause the strap to bunch up in the end of a slot,

failure can occur through initiation of edge tear. The fitting

and motion angles illustrated in Figure 26 are recommended.

If the integrity of the attachment of the fitting within the

buckle can be compromised by rotation, then rotation must be

completely eliminated. Eliminating fitting rotation in the

flat plane of the buckle during loading may prove to be difficult in lightweight systems. Experience has shown that it is

106

S fl lilii1i:I i~i] -l r- .- "----'.----'----- ..--.. ....
350

max

Shoulder harness

fitting

Lap belt7

e leangle t igfitting

in h itng agimtoottion. o xmlarudpni

typical)

20a 0 hl 0r w b p 20t

S~Single-point

release buckle a l

Figure 26. Buckle fitting attachment and motion angles. o

better to design the attachment of the fitting within theasil buckle to be insensitive to rotation than to rely on restraining the fitting against rotation. For example, a round pin in

a round hole would be preferable to a flat-faced dog which must

seat on a flat face of a slot. In the latter case, a small

amount of rotation can cause point loading of a corner of the

dog against one end of the slot. The point loading can easily

increase the stress applied at the contact point to its ulti- mate bearing strength. This will result in metal deformation

and the formation of a sloped surface which then can act to

cam open the attachment mechanism.

Further, the release mechanism (buckle) should be protected against accidental opening. Neither decelerative loading of

components nor contact with aircraft controls, such as cyclic

controls, should open the device. It was mentioned earlier

in this volume that required cockpit dimensions should be reviewed. It appears that the occupant can be placed too close

107

_ _ _ _ Af
to the cyclic control in helicopters and that a fully retracted

cyclic head can contact the buckle. The buckle release mechanism should be protected against inadvertent release either

during operation or in a crash. It should be emphasized that,

if contact between the cyclic control and the buckle is possible in an operational mode, a considerable overlap can exist

during crash loading when the restraint system is deformed forward several inches.

5.7.3.3 Lap Belt Anchorage: The actual anchorage point for

the lap belt can be located either on the seat bucket or on

the basic aircraft structure, although it is usually desirable

to locate it on the seat. If the anchorage is located on basic

aircraft structure, consideration must be given to the movement of the seat when load-limiting means are used so that the

lap belt restraint remains effective regardless of seat position. Longitudinal load limiting of the seat serves little

purpose if the lap belt is attached to the basic structure.

However, careful consideration must be given to the belt assembly strength since the belt must restrain the motion of the

seat, as well as the occupant.

The lap belt should be anchored to provide optimum restraint

for the lower torso when subjected to eyeballs-out (-G

forces. One of the anchorage variables which has an influence

on restraint optimization is the location of the lap belt anchorage in the fore-and-aft direction. The important characteristic is the angle in a vertical fore-and-aft plane between

a projection of the lap belt centerline and the buttock reference line, or plane. This angle defines the geometrical relationship between the longitudinal and vertical components of

the belt load. A small angle provides an efficient path for

supporting longitudinal loads while a large angle provides an

efficient system for supporting large vertical loads. Thus,

for supporting large forward-directed loads, a small angle

would be desirable, but for reacting the large vertical loads

imposed on the lap belt by the loaded shoulder harness a large

angle is required. The compromise for location of the anchorage must consider all the variables including the tendency for

the occupant to submarine under the lap belt.

In order to avoid the increased possibility of both spinal and

abdominal injury, a properly designed restraint system should

not allow submarining to occur. Still, an efficient angle

should be maintained to limit the forward motion of the occu- pant.

Comfort is another concern in lap belt anchor location. A pilot must raise and lower his thighs during operation of rudder

pedals or antitorque pedals. If the lap belt anchor is too far

108

.........--.-- -------.---
forward, the lap belt will pass over the pilot's thighs forward

of the crease between the thighs and the pelvis and thus may

interfere with vertical leg motion. It is important, therefore, to position the lap belt anchorage so that it provides

optimum restraint while not interfering with the pilot's operational tasks. A forward location of the anchor does not negatively influence the comfort of passengers since passengers

are not required to perform operations with their legs.

In order to satisfy comfort and crash safety requirements, the

vertical angle between the lap belt centerline and the buttock

reference line as installed on the 50th-percentile occupant

should not be less than 45 degrees and should not exceed 55

degrees, as shown in Figure 27(a). Further, it is desirable

to locate the anchor point at or below the buttock reference

line to maximize comfort and performance. If the anchor point

must be located above the buttock reference line, as on most

armored seats, the anchor point should be positioned to ensure

that the belt angle lies within the desired 45- to 55-degree

range. For a system having a lap belt tiedown strap to counteract the upward force of the shoulder harness (e.g., in pilot

seats), the lap belt anchors should be positioned so that the

centerline of the lap belt passes through the seat reference

point as shown in Figure 27(b). If the restraint system does

not have a tiedown strap (e.g., in passenger seats), the lap

belt anchor should be positioned so that the belt centerline

passes through the buttock reference line 2 to 2-1/4 in. forward of the seat reference point as shown in Figure 27(c).

This position provides sufficient vertical load components to

help counteract the upward force of the shoulder straps. For

positioning anchors that do not fall on the buttock reference

line, the angle between the lap belt centerline and the buttock

reference line can be assumed to be 45 degrees for systems with

tiedown straps and 55 degrees for those without.

For seats that limit lateral motion of the occupant with structure, such as in armored seats, the anchorage point and hardware should possess sufficient flexibility and strength to sustain design belt loads when the belt is deflecteu laterally

toward the center of the seat through an angle of up to 60 degrees from a vertical position. The side motion of fittings on

other seats should also be capable of supporting design loads

with the lap belt deflected laterally away from the center of

the seat through an angle up to 45 degrees from the vertical.

These recommendations are made to ensure that lateral loading

on the torso will not result in lap belt anchorage failure.

5.7.3.4 Shoulder Harness Anchorage: The shoulder harness or

inertia reel anchorage can be located either on the seat back

structure or on the basic aircraft structure, although it is

usually more desirable to locate it on the seat. In placing

109
a) All systems b) Systems with lap belt c) Systems without tiedown

tiedown straps straps

45-55* 4"5 55*

Seat reference

ZButtock point (typical) r eference

line 2.0 in. + .25 -0

Figure 27. Lap belt anchorage geometry.

the inertia reel, strap routing and possible reel interference with structure during seat adjustment or energy-absorbing

stroke of the seat must be considered. Locating the anchorage

on the basic aircraft structure may be the only practical approach for improving crashworthiness in light aircraft, particularly in retrofit applications. It will relieve a large

portion of the overturning moment applied to the seat under

longitudinal loading. However, due consideration must be given

to the effect of seat bucket movement in load-limited seats.

Vertical movement of the seat pan can be provided for by placing the inertia reel aft of the seat back shoulder strap guide

a sufficient distance so that seat vertical movement will

change the horizontal posit.¢,n and the angle of the straps very

little.

Shoulder straps should pass over the shoulders in a plane perpendicular to the back tangent line or at any upward (from

shoulders to pull-off point) angle not to exceed 30 degrees,

as illustrated in the upper-left sketch in Figure 28. A shoulder harness pull-off point should be at least 26 in. above the

buttock reference line.

The shoulder harness anchorage or guide at the top of the seat

back should permit no more than 0.5-in. lateral movement (slot

no more than 0.5 in. wider than strap) to ensure that the seat

occupant is properly restrained laterally. The guide should

provide smooth transitions to the slot. The transition contour

should be of a radius no less than 0.25 in. and should extend

110
Back tangent lineCorc

Torso carries small

Maximum vertical load

shoulder ...

* angl

Shoulder ..... strap load

300 IP

............ (Force diagram)

26. 0 to . . . ........

2 7. 0 ...............

. . . . . . . . .

. . . . . .. . . . . . . .

-4

Sneorec reeec

Figuckref2. Shuderanesanhrgegomty

lin111

TTT''.~~hi ~ -4 -. -.
II

completely around the periphery of the slot to minimize edge

wear on the strap and reduce the possibility of webbing failure

due to contact with sharp edges under high loading.

5.7.3.5 Lap Belt Tiedown Strap Anchorage: A lap belt tiedown

strap is required for forward-facing crewmembers. It is recommended that the tiedown strap be located on the seat pan centerline at a point 14 to 15 in. forward of the seat back. For

shorter seat pans, the anchor must be placed as far forward

as possible.

5.7.3.6 Adjustment Hardware: Adjusters must carry the full

design load of the restraint system subassembly of which they

are a part without slipping or crushing the webbing under

items such as locking cams. In extremely highly loaded applications, this may require that the strap be doubled in a manner

that requires the adjuster to carry only one half of the strap

assembly load. The force required to adjust the length of

webbing should not exceed 30 lb in accordance with existing

military requirements for harnesses. Insofar as possible, all

adjustments should be easily made with one (either) hand. Adjustment motions should be toward the single-point release

buckle.

5.7.3.7 Location of Adjustment and Release Hardware: This

hardware must not be located directly over head points of the

skeletal structure, such as the iliac crests of the pelvis or

the collarbones. The lap belt length adjuster should be located either at the center of the belt near the release buckle

or at the side of the hips below the iliac crests, preferably

the latter. The shoulder strap adjusters should be located as

low on the chest as possible in order to avoid concentrated

pressure on the collarbones.

5.7.3.8 Webbing Width and Thickness Requirements: Webbing requirements are discussed in detail in Section 5.7.4.

5.7.3.9 Hardware Materials: All materials used for the attachment of webbing (release buckles, anchorages, and length

adjusters) should be ductile enough to deform locally, parti- cularly at stress concentration points. A minimum elongation

value of 10 percent (as determined by standard tensile test

vpecimens) is recommended for all metal harness-fitting mater4.als. There are obviously some components that, for operational purposes, rely on hardness. These components should be

designed to perform their necessary function but be made from

materials as nearly as possible immune to brittle failures.

112
5.7.3.10 Structural Connections

5.7.3.10.1 Bolted Connections: The safety margins for shear

and tensile bolts in restraint systems should be 5 and 10 percent, respectively. Also, bolts less than 0.25 in. in diameter

should not be used in tensile applications. Wherever possible 1

the bolts should be designed for shear rather than tension.

5.7.3.10.2 Riveted Connections: The guidelines presented in

MIL-HDBK-5, Reference 32, are recommended for restraint system ! hardware design.

5.7.3.10.3 Welded Connections: Acceptable welding processes

are discussed in Military Specifications MIL-W-8604, -6873,

-45205, and -8611; however, strict inspection procedures should

be used to ensure that all welded joints are of adequate quality. (Other provisions presented in Section 5.4.3 also apply.)

5.7.3.10.4 Plastic Strength Analysis: Plastic analysis methods should be used for strength determination wherever applicable in order to obtain maximum-strength hardware at the lowest possible weight.

5.7.4 Webbing and Attachmerts

5.7.4.1 Properties: The main advantage of a single-strength

harness (only one restraint harness in the inventory) would be

the assurance that harnesses could be interchanged between

load-limited seats and nonload-limited seats without fear that

an understrength harness might be installed on a nonloadlimited seat. On this premise, the design strength of all

forward-facing and side-facing restraint harnesses should be

equnl. The design loads for the various harness components

attached to the seat are listed in Table 8. The elongation of

all webbing used in the harness must be minimized to decrease

overshoot. Table 8 shows that the shoulder strap elongation

is restricted to 1.5 in., while the lap belt is restricted to

2.0 in. of total end-to-end stretch or 1.0 in. of loop elongation. Restraint systems for the new generation of Army helicopters use a low-elongation polyester webbing, the characteristics of which are listed in Table 9.

5.7.4.2 Width and Thickness Requirements: Minimum webbing

width requirements are specified in Table 10. All webbing used

for restraint harnesses must be thick enough to ensure that

the webbing does not fold or crease to form a "rope" or present

a thin sharp edge under high loading that will cause damage to

soft tissue. A minimum thickness of 0.055 in. is considered

acceptable.

113

1= ____________....
TABLE 8. RESTRAINT HARNESS COMPONENTS LOAD-ELONGATION

DESIGN AND TEST REQUIREMENTS (MIL-S-58095(AV))

Maximum elongation

Minimum load (design goal)

Harness components lb)(a) (in.)(b)

Inertia reel strap(s) 6 0 0 0 (c)

Shoulder harness strap(s) 4 0 0 0 (d)

Lap belt 4000 2.0

Lap belt tiedown strap 4500 0.5

NOTES: (a) Applied in straight tensi-n.

(b) Total length of harness comnlonent tested must be

the same as when installed on the seat and adjusted for a 95th-percentile clothed occupant.

(c) This represents the total load from all shoulder

straps. A single diagonal shoulder strap should

carry 6000 lb.

(d) This represents the minimum load that one of two

shoulder straps should carry.

(e) This applies only to the shoulder harness and inertia reel strap outside the reel (exclu74ive of

the webbing wound on the spool of the inertia

reel).

TABLE 9. RESTRAINT WEBBING CHARACTERISTICS

Nominal Minimum

Restraint webbing Webbing breaking

system width thickness strength Elongation*

component (in.) (in.) (lb) (percent)

Inertia

reel 1-3/4 0.057 6980 6.9 @ 3000 lb

Shoulder

straps 2 0.057 7800 7.6 @ 4000 lb

Lap belt 2-1/4 0.057 8880 7.8 @ 4000 lb

Lap belt

tiedown

strap 1-3/4 0.057 6980 6.9 @ 3000 lb

*Based on 10-in. gage length.

114
TABLE 10. MINIMUM WEBBING WIDTH REQUIREMENTS

Minimum w.,ch

Webbing identity (in.)

Lap belt 2-1/4*

Shoulder strap 2

Tiedown strap 1-1/2

*A greater width (up to 4 in.) cr pad is desirable in the center abdominal area.

5.7.4.3 Webbing Attachment Methods

5.7.4.3.1 Stitched Joints: The strength and reliability of

stitched seams must be ensured by using the best known cord

sizes and stitch patterns for a specified webbing type. The

stitch patterns and cord sizes used in existing high-strength

military restraint webbings appear to provide satisfactory performance. The basic stitch pattern used in these harnesses

is a "W-W" configuration for single-lapped joints. The 50-lb

strength No. 6 cord at 4-1/2 to 5 stitches per inch is recommended, as illustrated in Figure 29, for use on MIL-W-25361

webbings. The use of the 50-lb cord and an 80-percent efficiency results in a minimum strength of 160 lb/in. (4 stitches

x 50 lb/stitch x 80 percent) for a single-lapped joint or

320 lb/in. for a looped joint. Thus, the total stitch length

needed can be determined by the total required load.

It has been shown recently that the heavier thread is not compatible with the new low-elongation polyester webbing (Reference 38). For these webbings, a smaller diameter cord offers

the advantages of reduced webbing fiber damage and the ability

to be used with automatic sewing machines and is therefore acceptable.

The use of a 30-percent increase in the total stitch length

required is recommended to offset the normal aging strength

decrease as well as the possible abrasion strength decrease.

Covering the stitched joints with cloth to provide wear protection for the cords is also recommended.

38. Farris, L., HIGH STRENGTH STITCHING FOR AIRCRAFT PERSONNEL RESTRAINT SYSTEMS, Pacific Scientific Co.; Proceedings, 1978 SAFE Symposium, Survival and Flight Equipment

Association, Canoga Park, California, October 1978.

115 ii5

__ __ __ __
4-1/2 to 5 stitches per inch MIL-T-7807B, No. 6 nylon cord Metal .(Type I or II, Class I, 50 ib)

Mardwl r or equivalent seam strength har.ware \with smaller diameter cord for

the new thinner webbing

0.12 in. -_ - 25-in.

minimum

spacing

[~ul.5_in -bJ4---one stitch minimum minimum

Figure 29. Stitch pattern and cord size.

mized 

The size of the overlapped and stitched area should be mini- to reduce weight, reduce the stiffened section of the

webbing, and provide more room between fittings for adjustment.

of 

5.7.4.3.2 Webbing Wrap Radius: The wrap radius is the radius the fitting over which the webbing is wrapped at buckles, anchorages, 

0.06 2

and adjusters, as illustrated in Figure 30. The -in. minimum radius should be carried around the ends of the 

webbing 

slot as shown in Figure 30 to preclude edge cutting of if the webbing should be loaded against the slot end.

5.7.4.3.3 Hardware-to-Webbing Folds: A possible method of re- ducing is to fold fitting the width webbing at 

as 

anchorage, shown buckle, or adjuster fittings in Figure 3l. This reduces the 

also 

weight and size of attachment fittings; however, it can

plied 

cause 

by premature webbing failure because of the force ap- the top layer of webbing compressing the lower against the 

to 

fitting slot edge. If this technique is to be used, tests demonstrate integrity are recommended. Also, for configura- tions 

adjuster 

that require two load paths, such as lap belts, where an

looped 

cannot hold the required 4000-lb load, the webbing is

strap. 

through a full-width slot which halves the load in each An adjuster is then included in one strap. Adjustment requires that the webbing be freely drawn through the fitting, a requirement that folded webbing cannot meet.

116

qmi
A

~0. webbing 25-in.

minimum

typical

Wrap radius (0.062-in. minimum)

should be carried completely round ends of slot as shown by

he heavy line above

DETAIL A

Figure 30. Wrap radius for webbing joints.

w

E

Figure 31. Webbing fold at metal hardware attachment.

5.7.4.3.4 Surface Roughness of Fittings: A surface roughness

of no more than RMS-32 is recommended to prevent fraying of

the webbing due to the frequency of movement over the metal.

117

4, - . • ' • T u, l • • * • • • • - • • • • • • , •
5.7.4.4 Energy-Absorbing Webbing: Energy-absorbing restraint

system webbing has been considered for limiting loads on the

occupant. However, primarily because of the increased potential for secondary impacts of occupants, energy-absorbing webbing is not recommended for use in seating systems. The limited space available in aircraft requires that the strike envelope be minimized. Therefore, the use of the lowest elongation available is specified.

5.7.5 Inertia Reels, Control, and Installation

The inertia reel should give the crewmember full freedom of

movement during normal operating conditions while automatically

locking the shoulder harness during an abrupt deceleration.

The design requirements specified in MIL-R-8236 (Reference 39)

are compatible with the other restraint harness requirements

listed in this chapter, and it is recommended that the use of

this specification be continued with the single exception of

strap load. The specification requires reels to exhibit a minimum ultimate strength of 4000 lb applied to the spool. Since

6000 lb is required (Table 8), strap friction must be relied

on to reduce the loads to tolerable values at the reel. Therefore, higher strength reels should be developed to eliminate

the weakness.

In addition to the MIL-R-8236-type reel, which has the function

of Freventing further strap extension, there are power-haulback

reels that rapidly retract slack to apply a tensile load to

the belt. Generally, these systems, some of which use a basic

MIL-R-8236 inertia reel, are powered by a gas generator and

must be manually actuated prior to impact. Automatic actuation by an acceleration sensor is not recommended because human tolerance considezations limit the haul-back velocity. By

the time the crash could be sensed, there would not be time to

complete the haulback within tolerable accelerative limits.

It is recommended that the rate-of-extension type reel be used

on all aircraft types to assure locking regardless of load direction.

The inertia reel may be anchored to the seat back structure or

to the basic aircraft structure with the same reservations previously mentioned in Section 5.7.3.4. The shoulder straps must

be maintained within the acceptable angle range as presented in

Figure 28. If an anchorage to basic structure is used, consideration must be given to the possible seat bucket motion so

that the shoulder strap angle or length does not change by a

39. Military Specification, MIL-R-8236D, REEL, SHOULDER HARNESS, INERTIA LOCK, Department of Defense, Washington,

D. C., 19 December 1975.

118

Ir-
significant 

should be mounted 

amount 

and 

during 

the 

energy-absorbing 

webbing routed so 

stroke. 

that the 

The 

webbing

reel f

does not bear on the reel housing.

5.8 SEAT STRENGTH AND DEFORMATION DESIGN REQUIREMENTS

5.8.1 Recommended Occupant Weights for Seat Design

The 95th- and 5th-percentile occupant weights are recommended

for the upper and lower limits of occupant weights to be considered in seat design. Ideally, seat stroke limits should be

sized for the 95th-percentile occupant, while the occupant acceleration limits should be determined for the 5th percentile.

occupants, while the stroke lengths also would be adequate for

all. In most situations, sufficient stroke distance will not

be made available in the aircraft to permit using the ideal

approach; therefore, compromises will have to be made. Specific criteria for these cases are presented in this chapter.

5.8.1.1 Crewseats: The design weight should be based on the

typical weight of the seat occupant, not the extremes. This

means that the aviator weight recommended for crewseat design

should not include combat gear. Typical weights are presented in Table 11.

TABLE 11. TYPICAL AVIATOR WEIGHTS

95th- 50th- 5thpercentile percentile percentile

weight weight weight

Item (lb _(lb) .(lb

Aviator 211.7 170.5 133.4

Clothing 3.4 3.1 3.1

Helmet 3.4 3.4 3.4

Boots 4.1 4.1 4.1

Total weight 222.3 181.1 144.0

Vertical

effective

weight 175.2 142.3 112.6

119
I

Variable-load energy-absorbing systems are highly desirable

to maximize efficiency and provide protection in limited space.

Therefore, they should be incorporated in seat designs whenever possible.

5.8.1.2 Troop and Gunner Seats: The same percentile range of

occupant sizes should be considered for tLjop and gunner seat

designs. A greater variation of clothing and equipment is used

by troops than by aviators; troop seats should be designed to

accommodate them. The 95th-percentile occupant should be considered heavily clothed and equipped, while the 5th-percentile occupant should be considered lightly clothed and equipped.

The typical weights of seated troops in aircraft are as shown

in Table 12.

TABLE 12. TROOP AND GUNNER WEIGHTS

95th- 50th- 5thpercentile percentile percentile

weight weight weight

Item (lb) (lb) (lb)

Troop/Guni.er

weight 201.9 156.3 126.3

Clothing

(less boots) 3.0 3.0 3.0

Boots 4.0 4.0 4.0

Equipment 33.3 33.3 33.3

Total weight 242.2 196.6 166.6

Vertical

effective

weight

clothed 163.9 127.4 103.4

Vertical

effective

weight

equipped 197.2 160.7 136.7

120

?, Lt
II

5.8.2 Strength and Deformation

5.8.2.) Forward Loads: For a load-limited system, a minimum

displacement must be achieved if the system is to remain in

place during a given decelerative pulse. Actually, all systems

are load limited, although not necessarily through original intent. The inherent load-deflection curve for any system imposes a definite limit on the system's ability to resist impulsive loading. The objective of intentionally load-limited seat

systems is to make the best use of the space available for relative displacement of the seat and occupant with respect to the

airframe, while maintaining loads on the occupant consistent

with the type of restraint system used and the occupant's capacity to survive the loads imposed.

Design curves for the forward direction are presented in Figure 32, where it is estimated that the requirements are not

conservative for the input pulses selected for design purposes.

These are a 30-G peak triangular pulse of 50-ft/sec velocity change in the cockpit and a 24-G peak with 50-ft/sec velocity change in the cabin area.

The st 4 ic loads that the seat must withstand are obtained by

multij ag the load factors (G) shown in Figure 32 by the sum

of the total weight of the 95th-percentile crewmember or passenger plus the weight of the seat and any armor or equipment

attached to or carried in the seat. For crewseats, the weight

of combat gear is not included (see Section 5.8.1.1).

Longitudinal displacement of approximately 6 in. for cockpit

seats and 12 in. for cabin seats measured at the seat refer- 4

ence point (the seat reference point may be projected to the

outside of the seat pan for measurement convenience) is the practica2 limit for seats in existing Army aircraft. Since there is typically more room available in cabins than in cockpits, the advantages of longer energy-absorbing strokes can

usually be achieved. Longer strokes permit the absorption of

equivalent energy at lower loads and thus can serve to reduce I

seat weight and increase the level of protection offered over

a wider occupant weight range.

In viewing Figure 32, it can be seen that for cabin seats

12 in. of stroke enables the minimum limit load to be reduced

to 15 G, whereas for cockpit seats a 20-G minimum limit load

is required with only 6 in. of stroke. The 15-G and 20-G minimum limit loads fix the G levels of the base curves for the

cabin and cockpit seat, respectively. The available stroke will

be unique for each specific aircraft, and the energy-absorbing

mechanisms in the seats should be compatible with the available

stroke distances. If forward or sideward motion threatens to

121

-_ __
45

35

'~30

25 1Sample curves Efn

I5 J E :

o20

-40- 0 Seat failure load __3_.......... ,,5 l 6• 

15 ----Unac eptabple peformac

bAcceptable performancef Ine fficient performance

Ci 0Ccki sejetontre

0 S 0.25 i 2 3 4 5 6

.4--

15Acceptable failure area0 ReCection area i lo

20 d__

• 1, 5 ' B a se; c ', r ve| : " " T7 " "' :i

112

5 .... " - - - Cabin seats _.k:z k. :•ZL::.,-,:.•::

0 0.5 2 4 6 8 10 12

Total controlled deformation (x), in.

measured at seat reff rence point

Figure 32. Seat forward load and deflection requirements

for all types of Army aircraft (forward design

pulse). 122 )

122

__ __ __ _ I

Li
limit the effectiveness of the vertical energy attenuating system or increase the possibility of severe injury caused by secondary impact of the occupant with items in the aircraft, then

energy-absorbing stroke in directions other than vertical

should not be used. The 6 in. and 12 in. allowed by the

curves of Figure 32 should be viewed as maximum distances

which are subject to limitations of available space in each

specific aircraft and location in the aircraft.

The initial slope of the cockpit seat base curve to 1.0 in.

of deflection allows for elastic deformation consistent with a

relatively rigid crewseat while the lighter weight and more

flexible troop/gunner seat requires a lesser slope. The 30-G

and 35-G upper cutoffs reflect consideraticn of human tolerance limits, load variations between cockpit and cabin locations, and practical limitations of seat weight and excessive

airframe loading.

5.8.2.2 Use of Design Curves: To be acceptable, a seat design

must have a characteristic load-deflection curve that rises to

the left and above the base curves of Figure 32 and extends

into the region beyond the upper curve. This requirement also

applies to the lateral strength and deformation requirements

discussed in Section 5.8.2.6. In Figure 32, curves A, C, and

E are acceptable curves, but curve B is unacceptable because

it does not reach the required ultimate strength. Curve D reveals inefficient use of seat deflection by intruding into the

base area. The seat is deflecting at too low a load, thus absorbing less energy than desirable.

5.8.2.3 Aftward Loads: Large aftward loads seldom occur in

fixed-wing aircraft accidents but may occur in rotary-wing accidents. A capability to withstand 12 G is recommended for

aftward loads for all seats. This value will usually be automatically met by all seats meeting the forward load requirements. Occupant weight should be the total weight of the 95thpercentile crewmember or trooper as presented in Section 5.8.1.

5.8.2.4 Downward Loads: Human tolerance to vertical impact

limits the acceptable forces in the vertical direction for all

airzraft seats. The maximum allowable headward acceleration

(parallel to the back tangent line) for seated occupants, is on

the order of 23 G for durations up to approximately 0.006 sec.

Therefore, the 48-G design pulse imposes the requirement for

energy absorption in the vertical direction by some form of

load limiting.

The effective weight in the vertical direction of a seat occupant is approximately 80 percent of the occupant's total

weight because the lower extremities are partially supported

123

I I'I" i• '" I • i _II'i__....._|__I______-___.... "__' .. ....... -- - -
by the floor. The effective occupant weight may be determined

by summing the following:

"* Eighty percent of the occupant's body weight.

"* Eighty percent of the weight of the occupant's clothing (less boots).

"* One hundred percent of the weight of any equipment

carried on the body above knee level. Combat gear is not included in the effective weight of the pilot

or copilot (see Section 5.8.1.1).

The dynamic limit load for the load-limiting system should be

established by use of a load factor (G ) of 11.5. The dynamic

limit load is determined by multiplyikg the summation of the

effective weight of the seat occupant, and the weight of the

movable or stroking portion of the seat, by 11.5. The resulting dynamic limit load includes the total force resisting the

vertical 'movement of the seat in a crash; the dynamic limit

load of the energy-absorption system, simple friction, and

friction due to binding, etc. This requirement is difficult

to satisfy with a sliding guidance system because the frictional load varies with contact load which, in turn, varies

with the impact load vector direction. A relatively frictionfree rolling mechanism or collapsible structure is therefore

recommended.

The 11.5-G design criterion, taken from Reference 40 and modified to provide a tolerable deceleration of the 5th-percentile

occupant, considers the dynamic response of the seat and occupant. The factor of 11.5 was established to limit the decelerative loading on the seat/occupant system to less than. 23 G

for durations in excess of 0.006 sec (the tolerable level for

humans as interpreted from the Eiband data) in crashes that do

not exhaust the stroke of the aeat.

Crewseats should be designed to stroke a minimum distance of

12 in. when the seat is in the lowest position of the adjustment range. This distance is needed tc absorb the residual

energy associated with the vertical design pulse. Further,

the load-limiting system should be designed to stroke through the full distance available including the vertical adjustment

40. Desjardins, S. P., and Harrison, H., THE DESIGN, FABRICATION, AND TESTING OF AN INTEGRALLY ARMORED CRASHWORTHY

CREWSEAT, Dynamic Science, Division of Marshall Industriesn USAAMRDL Technical Report 71--91, Eustis Directorate, U. S. Army Air Mobility Research. and Development Laboratory, Fort Eustis, Virginia, Jan-lary 1972, AD 742733.

124

,, ..., .
distance. Since a vertical adjustment of ±2-1/2 in. from neutral is typically required by crewseat specifications, proper

design can provide up to 17 in. of stroke, depending on seat

adjustment position.

If it is absolutely impossible to obtain a minimum of 12 in. of

stroke, a lesser amount is acceptable, but in no case should

it be less than 7 in. The reduced stroke is acceptable for a

retrofit application or for use in small aircraft in which it

is simply impossible to find the space for a 12-in. stroke.

In such cases a systems analysis is mandatoryi the analysis

must show that occupant protection is equivalent to the system

in which the 12-in. stroke is available.

For retrofit applications, the maximum protection possible

should be obtained in any component being modified, i.e.,

seats, gear, etc. Separate test criteria have been established

for seats not having the required 12 in. of stroke and are pre- sented in Section 5.10.2.2 of this document.

Energy-absorbing systems should be designed for 11.5 plus I G

minus 0 G including the effect of the dynamic loading rate.

To obtain the static test loads, dynamic limit loads should be

reduced by the amount due to rate sensitivity of the particular device used. Further, in the design of the system the desired total resistive load on the seat should be obtained by summing the resistive load provided by the energy-absorbing

system and the resistive load resulting from friction and/or

other mechanisms unique to the particular system. Thus, the

resistive load of the energy-absorbing subsystem must be less

than the load required to decelerate the seat by the amount of

the other stroke-resisting variables.

If the energy-absorbing system is to provide only one force

setting, the effective weight of the 50th-percentile occupant Sfrom Tables 11 and 12 should be used for sizing it in order

to ensure a tolerable stroke for the majority of the occupants, not exceeding the stroke limitations of the seat. These

weights are 142.3 and 160.7 lb for pilot/copilot and troop and

gunner seats, respectively.

In order to use the stroke distance available at maximum efficiency, regardless of occupant weight, a variable-force loadlimiting mechanism is desirable. With an infinitely variable

force system, the deceleration levels can be maintained within

acceptable limits (if the stroke is not exhausted) for the full

range of occupant weights for either crew or troop seats while

using equal stroke lengths for identical pulses. A compromise

is possible for a seat design that uses a load-limiting device

rather than collapsing structure. The device can be designed

125

9 ** *.-

S i . r ... i '
to produce two or more limit loads that can be selected by the

seat occupant. The selection would be made on the basis of

seat occupant weight. For example, for a dual-limit-load device, the lowest force might be established by using the weight

of a 5th-percentile occupant. The second force might be designed for the weight of a 50th-percentile occupant. In operation then, the occupant would be required to select a limit

load by movement of a lever or dial upon entering the seat. It

is recommended that at least a dual-level load limiter (preferably three or more levels) be used to provide maximum protection over the complete occupant weight range.

The interaction between the occupant and the movable seat

masses increases with seat mass. Therefore, the movable seat

mass should be minimized.

A

Troop seats should be designed for the maximum stroke feasible

to maximize protection over the large weight range represented

by the fully equipped and lightly equipped occupant. It is

recommended that the full 17-in. seat pan height normally consid6red desirable from the human engineering standpoint be used

for energy-absorbing stroke. It is further recommended, as

a minimum, that the limit load of the system be sized using

the 11.5-G load factor and the effective weight of the 50thpercentile heavily equipped occupant (160.7 lb). Variablelevel load limiters sized as discussed previously are also de- sirable for troop seats.

5.8.2.5 Upward Loads: A capability to withstand a minimum

upward load of 8 G is recommended for all aircraft seats. Occupant weight should be that of the 95th-percentile crewmember

or trooper as presented in Section 5.8.1.

5.8.2.6 Lateral Strength and Deformation Requirements: The

lateral load and deformation requirements for forward- and aftfacing seats are presented in Figure 33. Two curves are presented. One is for rotary-wing aircraft, and the other is for

light fixed-wing aircraft. The deflections at the seat reference point should be measured. Occupant weight should be as

specified in Section 5.8.1 and should be that of the 95thpercentile aircrew member or trooper.

Lateral loading in the forward direction (aircraft reference

system) on side-facing seats should be the same as for forward

loading (Figure 32) except that load limiting should be employed.

For crewseats, the lateral deflection should be minimized; however, it is doubtful if any great stiffness can be achieved in

lightweight hardware. It is believed adequate, as a design

126
25

Lateral 23

Acceptable failure area L

(Example for forward- "0 or aft-facing seat) Sarea 15 Rejection e 20Oad Utility and

0 observation ý4 -helicopters

____ _ _#--Design goal

U 1.4Light fixed -wing,

4J |attack, and cargo

m -_"__ * Base curve -

0 0.25 1.0 2.0 3.0 4.0

Total controlled deformation (y), in.

measured at seat reference point

Figure 33. Lateral seat load and deformation requirements

for all types of Army aircraft.

goal, to attempt to limit the initial deflection to 1 in. with

a 2-in. requirement. Because of the possible loading rate sensitivity of the seat materials, it is felt to be acceptable to

allow analysis of test data to demonstrate compliance. This

analysis might include adjustments of static test data by use

of measured or known deflection and load data from dynamic

tests. Further, in cases where wells are provie.ed under the

seats to increase the available stroke distance, the deformation should be elastic. This will allow the seat to realign

itself with the well prior to entry after reduction of the lateral and longitudinal loads in those cases where the loads

are relieved soon enough.

5.8.3 Othez Seats

The requirements presented for crewseats and troop and gunner

seats also apply to passenger seats and any other seat installed

in the aircraft for any purpose. Unique seats installed for

special uses are not to be exempt.

127

nIMF. .. , -
i

5.9 PERSONNEL RESTRAINT HARNESS TESTING

SThe restraint harnesses are to be statically and dynamically

Stested along with the seat and/or structure to which they are

attached. However, the lap belt, shoulder straps, and tiedown

straps, including all hardware in the load path, should be statically tested separately to ensure that all components possess

adequate strength and to determine elongation. The strength

and elongation test requirements of restraint system subassemblies are specified in Table 8.

Specific component tests, including operational tests, are detailed in a draft military specification (Reference 41). However, all components and subassemblies should be statically

load tested. Each subassembly should be tested to its full design load to demonstrate its adequacy. Elongation characteristics should be measured to document these data for comparison with requirements and use in systems analyses.

5.10 STRUCTURAL SYSTEM TEST REQUIREMENTS

Both static and dynamic tests are recommended, and it is also

recommended that all seat and litter systems be tested as complete units. This is not to imply that component tests are

not useful; on the contrary, they can be extremely useful

and should be employed wherever possible to verify required

strengths. This practice is particularly valid where analyses, such as by finite element methods, have been used to accurately predict distribution of loads in redundant structures.

Upon acceptance of prototype systems tested under both static

and dynamic conditions, no further tests should be required

except for quality assurance. Major structural design changes

in the basic seat system will require static retesting of the

new system to ensure that no loss in strength has been caused

by the design changes. If the changes could affect the energyabsorbing, or stroking, performance of the seat, additional dynamic tests should also be conducted. Major structural design

changes are those changes involving principal load-carrying

members such as floor, bulkhead, or ceiling tiedown fittings,

structural links or assemblies, seat legs, or energy-absorbing

systems. Minor changes, such as in ancillary fittings, czn be

accepted without a structural test. A significant weigrit .icrease, however, such as the addition of personnel or se-at -rmor, would require additional testing. In summary, &i•lfe

that increase loading, decrease strength, produce signiticrnt

changes in load distribution, or affect the stroking mechaiisir

will require retesting.

41. Proposed Draft Military Specification, MIL-R-XXXX(AV),

RESTRAINT SYSTEM, AIRCREW, September 1974.

128
All testing is to be conducted with the seat cushions in place

and, for seats with adjustments, the seats should be in the

full-up and full-aft positions unless another position is shown

to be more critical. All tests should be conducted under simultaneous conditions of floor buckling and warping as illustrated in Figure 34 or bulkhead warping as illustrated in Figure 35. The combination of warping conditions should be that

which represents the most critical case for seat performance,

such as that most likely to impede seat stroking. For example,

considering the combined-load static test (No. 5 in Table 13)

of a seat such as that shown in Figure 34, if the lateral load

component were applied to the right, the right-hand track

should be warped upward at the forward end (+10 degrees) to

evaluate the possibility of interference with vertical stroke.

Also, the seat should be mounted for testing on actual air- craft hardware, i.e., tracks or bulkhead fittings.

If desired, dynamic tests may be substituted for static tests;

however, loading in all principal directions is required. Alternate dynamic tests are presented in Section 5.10.1.9.

5.10.1 Static Test Requirements

5.10.1.1 General: Table 13 presents the static test requirements for complete seat units. All static tests should be conducted under simultaneous conditions of floor or bulkhead buckling and warping as described above.

5.10.1.2 Unidirectional Tests: Where separate strength and

deformation requirements have been specified in Table 11 for

longitudinal, vertical, and lateral loading of seats, the loads

should be applied separately. Seats must demonstrate no loss

in structural integrity during these tests and should demonstrate acceptable energy-absorbing capacity.

5.10.1.3 Combined Loads: Seats must demonstrate no loss of

structural integrity under conditions of combined loading as

shown in Table 13 and should demonstrate ability to stroke in

the vertical direction with the transverse loads applied.

5.10.1.4 Load Application Method; The test loads should be

applied through a body block (see Section 5.10.1.5) restrained

in the seat with the restraint system. The loads are to be ap- plied at the expected center-of-gravity location of the occupant or occupants of each seat, as illustrated in Figure 36.

The loads calculated by multiplying the weight of the occupant

and equipment plus the weight of the seat by the required load

factor should be applied continuously, or in not more than 2-G

increments while the load-deformation performance of the seat

is recorded. Maximum loads need not be held for more than

129

,=--/,--, aF . . .- - '.. ' - ,. 
F

TABLE 13. SEAT DESIGN AND STATIC TLST REQUIREMENTS

I Percentile

Test Loading direction occupant Load/deformation

ref. ,with resp,:t to Load used in load

no. fuselage floor required determinatioO requirementsali

1 Upward 8-G minimum 95 No requirement

2 Downwardb'd 11.5 0 G See Section 1.-0 G 50 5.8.2.4

3 Aftward 12-G minimum 95 No requirement

4 Forward See Figure 32 95 See Figure 32

5 CombinedFrade,f

Forward See Figure 32 95 See Figure 32

Downwardc 11.5 +2.0 G 50 Same as Test 2 h

-1.0OG

Lateralf 9-G minimum 95 No requirements

6 Lateralg See Figure 33 95 See Figure 33

(a) The aircraft floor or bulkhead should be deformed as detailed in

in Figures 34 and 35, simultaneously with, or prior to the conduct of all static tests and kept deformed throughout load application.

(b) If more than one load-limiter setting is provided, a representative sample of settings spanning the range of loads should be

tested.

(c) If more than one load-limiter setting is provided, the highest

load should be used.

(d) Subsequent to the stroking of the vertical energy-absorbing device, cockpit seats should carry a static load of 25 G, based on

the effective weight of the 95th-percentile clothed and equipped

occupant per Section 5.8.1 plus seat without loss of attachment

to the basic structure except when the seat pan has stroked to

and is supported by the floor.

(e) In the event that no load-limiting device is used in the forward

direction, a 20-G load for cabin seats and a 25-G load for cockpit seats may be used for this combined loading.

(f) For seats employing vertical guides which could distort under

combined loi ding and cause binding, the maximum forward and lateral loads should be reached prior to initiation of stroking.

This sequence demonstrates whether the seat will stroke downward

after transverse loads are applied.

(g) The lateral loads should be applied in the most critical direction. In the case of symmetrical seats, the loading direction

is optional.

(h) Failure to meet the 11.5-G +2.0/-I.0-G static vertical load

limit should not be cause for seat rejection if the seat vertical

energy-absorbing system meets dynamic load requirements.

(i) Plastic deformation is permissible; however, structural integrity

must be maintained.

130

L
*IoI

Hydraulic Hydraulic

cyine cylinder

Or g n l'Original

VN position

-NI ' 10 .

Figure 34. Suggested method of applying floor warping for static testing of seats.

131

I",
S~ Hydraulic cylinder ~Test fixture structure

Test

structure

Seat

•Deflected.

Seat Oiginal •

attachments position

Bulkhead

Figure 35. Suggested method of applying bulkhead

warping for static testing of seats.

I sec. The maximum load reacheu, regardless of duration, is

to be used to assess compliance.

On integrally armored crewseats, care should be taken to assure

that the loads are applied proportionally to the proper assembly or test item to simulate the loads that would typically be

carried by the restraint harness and the seat support structure.

In other words, the portion of the load that could be expected

to be restrained by the restraint harness should be applied to

the body block as described above. The portion of the load

representing inertial loading of the movable assembly should

be applied separately at the center of gravity of the appro- priate substructure through another provision. For example, a

lever to proportion the load between the body block and movable section of the seat, and a sling to apply the appropriate

portion of the load to the bucket, can be used. For seats with

a relatively heavy frame, the inertial load of the frame can be

applied separately at its appropriate center of gravity. This

technique, although adding complexity to the test setup, assures that all components in the seat and restraint system assembly have been tested to their approximate static design

loads and that, as far as a static test simulatiin can be extended, performance and structural adequacy have been demonstrated. For lightweight seats (less than approximately 45 lb

foi total seat and restraint system), the total load can be applied to the body block.

132
• II

15.0 in.

Forward Aftward

• I10.0 i O.5 in.

6.0 ± 0.5 in.

3.0-in, radius

Seat reference

" Cpoint

Q) 0

Figure 36. Static load application point and critical

body block pelvis geometry.

5.10.1.5 Static Load Body Block: The static test loads must

be applied through a body block contoured to approximate a

95th-percentile occupant seated in a normal flying attitude.

The body block must co~ntain shoulders, neck, and upper legs,

and provide for passage of a lap belt tiedown strap between the

legs. The upper legs should be contoured to simulate the flattened and spread configuration of seated thighs and to allow

the proper location of the buckle. Critical pelvis dimensions

are shown in Figure 36. Buttock contours must be provided

to permit proper fit in a contoured seat pan. The leg stubs

should be configured to permit proper seat pan loading as the

body block rotates forward under longitudinal loading; i.e.,

the leg stubs should be only long enough to provide a surface

to react the lap belt load. The side view of the buttocks

should include an up-curved surface forward of the ischial tuberosities to allow the forward rotation of the body block

133

.~y
L --

while maintaining the primazy contact between the ischial tuberosities and the seat pan through the cushions.

5.10.1.6 Deflection Measurements: Deflection should be measured as close to the seat reference point as possible to eliminate seat structure rotational deformation from influencing

the test results. To simplify these measurements, the seat

reference point can be projected to the outside of the seat pan

or bucket.

Normally the restraint system will be attached to the seat.

However, if a unique situation should develop in which the only

option for increasing crashworthiness is to attach the system

(lap belt and shoulder harness) to the basic aircraft structure rather than to the seat, certain factors should be considered. First, the forward and lateral deflection requirements of Figures 32 and 33 need not be considered because the

restraint harness limits torso and seat deflection. Second,

the vertical deflection of the seat pan still must be considered since the downward movement of the seat pan could cause

excessive slack in the restraint harness, or the harness could

limit the stroke of the seat, depending on where the restraint

system is anchored. Neither of these conditions is acceptable

in the design.

5.10.1.7 Load Determination: The total load required for all

test directions, except downward, is determined by multi&.' ng

the required load factor from Table 3]3 by the weight of _ ' 95th-percentile cl~thed and equipped occupant from Table 11 or

12 (Section 5.8.1) plus the weight of each seat. The effective

weight of the 50th-percentile occupant should be used to calculate vertical components of loading (Test Nos. 2 and 5 of

Table 13) as discussed in Section 5.8.7.4; the effective weight

of the 95th-percentile clothed and equipped occupant should be

used for the bottomed test (Test No. 2(d) in Table 13). The

weight of that portion of the seat that strokes with the loadlimited portion of the seat must be added to the occupant

weight to determine the total required load in the vertical

direction.

5.10.1.8 Multiple Seats: Multiple-occupancy seats should be

fully occupied when tested. If it is determined that the most

adverse loading condition occurs in other than full-occupancy

situations, additional tests should be run for those conditions.

5.10.1.9 Substitution of Dynamic For Static Tests: It is recommended that static tests be conducted because they are more

economical to run than dynamic tests; their slow rate of load

application permits closer real-time observation of seat response to the loading, and static testing provides structural

response information which is more comparable to the static

134
analyses typically used in the development of present seat designs. lI the future, when dynamic analysis becomes more reliable, this latter point will no longer apply. A significant consideration in static-versus-dynamic testing is the cost of

the hardware. Static testing can be conducted with a minimum

number of seats because the condition of the seat can be monitored and judgments made as to its acceptability for continued

testing. If failures due to previous tests occur, parts can be

replaced and the test economically rerun. 9

If Zinal acceptence decisions are based on dynamic tests alone,

a consider&bly more rigorous dynamic test matrix is required to

enable testing in all the principal loading directions. Dynamic tests are usually more expensive than static tests and

the increased number of tests will also require additional

hardware. If new hardware is not used for each test, the results may be inconclusive. If the seat passes the test, the

results are acceptable; but if the seat fails the test, another 3

test must be run since it will not be apparent whether the I failure was due to damage inflicted during a previous test or

due to a basic design or manufacturing flaw.

If for any reason, dynamic tests are substituted for the static

tests previously described, then loading in all principal directions must be conducted. The dynamic test requirements are

presented in Figure 37. These three tests must be conducted

in addition to the two presented in Section 5.10.2 and all five

must be passed. These tests are to be conducted in accordance

with the same ground rules as those presented in Section 5.10.2

and are subject to the same testing parameters and evaluation

procedures. A 50th-percentile dummy should be used in Test 1

and a 95th-percentile dummy in the others, both of the type and

weight described in Section 5.10.2. Further, the static upload

of 8 G and the static aftward loading of 12 G must be imposed and satisfactorily passed.

5.10.2 fLnamic Test Requirements

5.10.2.1 Dynamic Test Requirements for Seats Having at Least

should 

12 in. of 

be 

Vertical 

dynamically 

Stroke: 

tested 

All 

to the 

U.S. 

two 

Army 

conditions 

prototype 

specified

seats 1

in Figure 38. A 50th-percentile anthropomorphic dummy comply- ing with the Code of Federal Regulations, Title 49, Part 572

specification for dummies (Reference 28) should be used to simulate the seat-system occupant for Test 1. A 95th-percentile

anthropomorphic dummy simulating as closely as possible the

features of the 50th-percentile dummy described above should

135

I.,m
r!

G max -

00

G rmin - [i/

C

to tl -2

Time, sec

Cockpit seats Cabin seat*

Test Configuration* Parameter Qualification R&D Qualification R&D

Du~r4y

uieitial tI sec 0.036 0.02C .050 .020

t 2 sec 0.051 0.051 .074 .074

min 46 46 32 32

0G fax 51 51 37 3'

A- mlin, ft/sec 42 42 42 42

2ahic[tars t1 S 035

" .' atL~ e io t r I sa 0.062 01036 .062C3

jjy 2) sec 0.104 0.104 .104 .100

in~ta G n~mn 14 16 14 16

G max 21 21 21 2

S•vml, t/sec 30 30 3 30

Sm LLI•ht rixed-wina,

2b Iar3 a .itta~ k ti ccc 0.067 0.033 .057 .033 helico~ters

i n e r t m r y t 2 s e c 0 .1 00 0 . 20 0 . I 0 7 .1 20

xnertia! G min 14 4 14 14

_E v min, ft/soc 25 21, 25 25

3 i u~r tl ^cc 0.066 C. 039 .081 ,046

adu nm

C min 28 28 22 22

G max 33 33 27 27

Av min, fL/aec 50 50 50 So

*All tests should be performed with ai:craft floor or bulkhead deformed as shown in Figure 34 or 35, respectively. The combination of warping

conditions should be that which represents the most critical case for

seat performance.

Figure 37. Requirements of additional dynamic tests

if substituted for static tests. -

136
I

Sz • max

00

Time, sec

Cockpit sc'ats Cabin seats

Test Configuration * Parai'eter Qualif-cation R&D Qualification fRLD

• -~Dumm iadnertial-i

Dusi nrilti sOC 0.043 0O.0 2 4 0.059 10.034

S2 SeC 0.061 0.061 C.087 0.087

G min 46 46 32 32

G max 51 Si 37 37

Av mln, ft/sec so 50 so so ' ] 10,

2 e tI sec 0.066 0.038 0.081 0.046

St 2 sec 0,100 0.100 0,127 0.127

G min 28 28 22 22

inertial G max 33 33 27 27

load ] v min, ft/Oec 55U 50 s 50

*Al1 tests should be performed with aircraft floor or bulkhead deformed as shown in Figure 34 or 35, respectively. The combination of warping

conditions should be that which represents the most critical case for

seat performance.

Figure 38. Dynamic test requirements for qualification and for research/development testing.

be used to simulate the seat-system occupant for Test 2. Total

weight, including instruments, of these two test dummies should

be:

50th percentile: Pilot/Copilot - 181.1 lb

Troop/Gunner a 196.6 lb

95th percentile: Pilot/Copilot = 222.3 lb Troop/Gunner - 242.3 lb

137

-s-]..- ~ ~ ~ " ' - --- --- ..--. - .--..- _____________________
Dynamic testing of multiple occupant seats should be performed

with the maximum number of occupants specified for the test

seat. Additional tests should be run if it is determined that

the most adverse loading condition occurs in other than fulloccupancy situations. For both tests of Figure 38, adjustable

seats should be adjusted to the full-aft and up position of

the adjustment range. Plastic deformation of the seat is permissible; however, structural integrity must be maintained in

all tests. For Test 1, the seat should limit the acceleration

as measured in the pelvis of the dummy to values which ensure

that the 50th-percentile clothed seat-system occupant (see Section 5.8.1) will not experience vertical, *G , accelerations in

excess of human tolerance as defined in Sections 4.3 and 4.8 of

Volume II (see Figure 8). The roll direction (10 degrees right

or left) for Test I should be the more critical loading, if ap- A

plicable, for the specific seat design.

When determining compliance of the achieved test pulse with the

dynamic test requirements of Figure 38:

1. Determine the maximum acceleration and construct the

onset slope for the test pulse by the method explained in Section 5.10.3.

Compare the achieved onset and peak acceleration of

the test pulse with those allowed and presented in

Figure 38. The achieved onset slope should lie between the minimum and maximum onset slopes using the

values of t1 and t listed in Figure 38 for the specific test conditions. The maximum acceleration

should also fall between the upper and lower limits

allowed.

3. Integrate the actual acceleration-time curve of the

test pulse and establish the achieved velocity change.

The velocity change achieved should be equal to or

greater than that tabulated for the specific test conditions.

5.10.2.2 Special. Dynamic Test Requirements for Seats Having Less Than 12 in. of Vertical Strokes In the event tMat the

systems approach permits the seat to have less than 12-in. minimum vertical stroke, additional requirements are made of the

dynamic testing. First, it would be desirable to perform a *

full-scale crash test with the test specimen, including all assemblies involved in the energy-absorbing process. This would

include a section of the fuselage, landing gear, and the seat

or seats. This approach is totally acceptable for demonstrating the dynamic response and acceptability of the system.

138
Since cost associated with the type of system testing described

above is usually prohibitive, a different approach is accept- able. This approach includes dynamically testing the seat only,

as is done for systems with at least 12 in. of stroke, but modifying the input pulse to represent the energy-absorbing processes of the gear and fuselage. An example of such a modified

test pulse is presented in Figure 39. The initial plateau (t1

to t') represents the acceleration-time history created by 1

stroEing of the landing gear. The sharp increase in acceleration at ti relates to fuselage impact, and the pulse beyond t'

represent 2 the crushing of the stiffer fuselage section. The

velocity change under the pulse should be the same as identified for the particular crash force direction for other established tests (50 ft/sec for Test No. 1 or No. 2 of Figure 38).

G2

0

o

1 Fuselag

Landing gear stroking icrushing

tt 1 t; t

Time, sec

Figure 39. Example of input pulse for seats

having less than 22 in. of stroke.

* The most comprehensive and rigorous analytical techniques, sup- ported by test data, should be used for determining the proper- ties of the fuselage. Since drop tests of landing gear are required, a much more accurate approach exists for obtaining the

landing gear influence on the pulse. Seat testing should await

completion of landing gear testi so that the results can be

used to ebtablish the initial plateau (or other shape) between

t and t; of the input pulse.

139

- ---------
I

Since ea'2h system may display different characteristics, it is

not appropriate to present in this document specific quantitative limits for use in evaluating the acceptability of the test

pulse. However, the same general approach and tolerances already presented for the standard pulse apply and should be

used. The technique described in Section 5.10.2.1 for establishing compliance with the required test pulse applies directly to the portion of the special test pulse following t•.

5.10.3 Data Acquisition and Reduction

Data acquisition and reduction should comply with the requirements of SAE J211 (Reference 42) for measurements on anthropomorphic dummies and structures.

Data should be presented in both analog and tabular form in compliance with the sign convention shown in Figure 3. Impact velocity should be determined and recorded for the test platform or vehicle. In the analysis of the data, velocity change

should be computed through either electronic means or graphically with a planimeter by integrating the area under the measured acceleration-time trace.

The method recommended for use in establishing the acceptability of the pulse (see Section 5.10.2) and to determine other

parameters associated with the data is similar to that presented in MIL-S-9479(USAF); see Reference 43. Parameters such

as rise time, onset slope, and acceleration plateau duration

may be obtained using the following graphic approximation technique as shown in Figure 40.

* Locate the calibration baseline.

• Determine the maximum (G p) acceleration magnitude.

e Construct a reference line parallel to the calibration baseline at a magnitude equal to 10 percent of

the peak acceleration (G ). The first and last in- tersections of this linepwith the acceleration-time

plot defines points 1 and 2.

42. SAE Recommended Practice, SAE J211b, INSTRUMENTATION FOR

IMPACT TESTS, SAE Handbook 1979, Part 2, Society of Automotive Engineers, Inc., Warrendale, Pennsylvania, 1979,

pp. 34.117-34.118.

43. Military Specification, MIL-S-9479, SEAT SYSTEM, UPWARD

EJECTION, AIRCRAFT, GENERAL SPECIFICATION FOR, Department

of Defense, Washington, D. C., 24 June 1973.

140
SG 5, 6

0.9 G 4

p

Calibration baseline 1.0 Gp

Figure 40. Graphic approximation example. (From Reference 43)

e Construct a second reference line parallel to the

calibration baseline at a magnitude equal to 90 percent of the peak acceleration. The first and last

intersections of this line with the acceleration-time

plot define points 3 and 4.

* Some practical judgment may be required for selection

of the first and last intersections depending on the

degree of noise, structural or electronic, apparent

in the data. Significant tendencies are important,

not noise.

e Construct the onset line defined by a straight line

through points I and 3.

* If desired, construct the offset line defined by a

straight line through points 2 and 4.

e If desired, construct a line parallel to the calibration baseline, through the peak acceleration. The

time interval defined by the intersections of this

line with the constructed onset and offset lines

(points 5 and 6) is the plateau duration (At).

141
I ft - r

e Locate the intersection of the constructed onset line

with toe calibration baseline (point 7). The time

interval defined by points 7 and 5 is the rise time

(t - t. Referring to Figure 38, the rise time

should greater than t but less than t when determintng compliance witW dynamic test re;iirements.

Point 7 is *he initial time t0 in Figure 38.

55.10.4 Seat ComponentAttachment

Since components that break free during a crash can become lethal missiles't it is.vrecommended that attachment strengths be

consistent with those specified for ancillary equipment (see

Section 6.6.5.9, Volume I11). Therefore, static attachmen t

strengthz for components, e.g., armored panels, should be as

follows:

Downward: 50 G

Upward: 10 G

Forward: 35 G

Aftward: 15 G

Lateral: 25 G

5.11 LITTER STRENGTH AND DEFORMATION REQUIREMENTS

5.11.1 General

The ultimate vertical strength of existing litters with a

200-lb occupant and a total system weight of 250 lb (see Section 5.11.2) is about 13 G. Since the desired decelerative

loads to be imposed on these litters exceed 13 G, special techniques must be used to limit the deflection and to support some

of the occupant load.

Lateral orientation in the aircraft is preferred because of the

characteristics of existing restraint systems used on litters

which provide more support when loaded laterally than when

loaded longitudinally.

5.11.2 Recommended Occupant Weights for Litter Design

The litter strength and deformation requirements defined below

are based on a 200-lb, 95th-percentile litter occupant with

20 lb of clothing and personal gear, a 10-lb splint or cast,

and 20 lb of litter and support bracket weight for a total

142

! .
L weight of 250 lb (the weight of a litter and patient as specified in MIL-A-8865 (ASG), Reference 44).

5.11.3 Vertical Loads

5.11.3.1 Downward Loads: In the case of litter systems, human

tolerance is not the limiting case in the vertical direction.

The loads would be applied in a transverse direction to the

body of a litter occupant. However, design to the 45-G human

tolerance level is impractical due to the strength requirements

for litters and for the basic structure to support the litter

systems.

Litters are either hung from the ceiling or supported at the

floor. In either case, the input deceleration pulses are the

same as for floor- or bulkhead-mounted seats (see Volume II). Litters should not be suspended from the overhead structure

unless it is capable of sustaining, with minimum deformation,

the downward loads from the tiers of litters. Therefore, in

the design of an efficient system, intentional load limiting

should be related to the floor pulse.

The vertical strength and deformation requirements for a litter

system are detailed in Figure 41. This curve is read in the

identical manner a: the seat load-deflection curve shown in

Figure 32. The load factors in units of G are based on the

summation of the weights of the occupant plus clothing, personal gear, splint or cast, and the weight of the litter and

attachment brackets for a total of 250 lb as described in Section 5.11.2. The curve of Figure 41 is based on the assumption

that 3 or 4 in. of vertical deflection will occur at the midpoint of the litter. In the unlikely event that a rigid litter

is used, an additional 2 in. of deflection should be added to

the curve. The deflection curve is limited to 6 in., because

a large deflection occurring on one corner of the litter due

to an asymmetric loading could cause ejection of the litter occupant. A larger energy-absorbing stroke can be used effectively if a mechanism is included in the system to control the

amount of tilt allowed. For example, a system mechanism could

be designed that forced all four corners of the litter to

* stroke the same distance (within elastic limits) thus achieving

this goal.

The additional problem associated with inadequate litter

strength must be dealt with in the design of litter systems. The curve of Figure 41 assumes a litter capable of at least

17 G with a maximum of 25 G. If the existing litter is used,

44. Military Specification, MIL-A-8865, AIRPLANE STRENGTH AND

RIGIDITY MISCELLANEOUS LOADS, Department of Defense, Washington, D. C., 18 May 1960.

143
Aircraft ceiling

Downward

I errs -rtS.- rf•F

I z

/777777

Aircraft floor

25

ACCEPTABLE

REJECTION FAILURE

ARE AREA

0 o15

BASE AREA

10

0

•*d. 5 - -,-.-.-

W..- 0246 0 1 2 3 4 5 6

Total controlled deformation (z), in.

*G value based on 250 lb per litter

position.

Figure 41. Litter downward load and

deflection requirements.

then a pan, net, or other device should be included under the

litter to catch and support the litter occupant if the litter

fails. Actually the device should limit the deflection to a

value less than that required to fail the litter and should

stroke with the litter. If all of these provisions are included, i.e., a rigid new litter or old litter with supporting

pan underneath, together with the tilt-limiting mechanims, then

144
-I

-• .- the stroke can be extended to 12 in. at a 17-G limit-load factor. The load-deformation curve of Figure 41 would be extended

at 17 G to 12 in. of stroke.

5.11.3.2 Upward Loads: All litter systems should be capable

of withstanding a minimum upward load of 8 G.

S5.11.4 Lateral and Longitudinal Loads

Litter systems for all aircraft should be designed to withstand

the load and deformation requirements indicated in Figure 42 in

all radials of the lateral/longitudinal plane. The litter lateral loads are made equal to the longitudinal loads because

the litters may be oriented in either direction depending upon

the aircraft.

30-kAcceptable failure area %

25 eptable

25 A

0 K Unacceptable

4 215 

Si/i • Base curve -

C-)

IC4J~

0

0 1 2 3 4 5 6

Total controlled forward** deformation of litter bed, in.

•*G value based on 250 lb per litter position.

**Forward is the direction towards the nose of the aircraft

regardless of litter orientation in the aircraft.

Figure 42. Litter forward or lateral load and deflection

requirements for all types of Army aircraft.

145

I' I " I _-_I_-_I f I -i -I I -. .. ...- 7 ... .... 'i I -- -
The 20-G acceptable load level indicated in Figure 42 is predicated on the tolerance to acceleration of an individual restrained by straps on existing "table top" litters. If litters

and allied restraint harnesses are designed for improved crashworthiness, the 20-G load should be increased to 25 G.

Acceptable or nonacceptable load-deformation characteristics

are read from Figure 42 in the identical manner as the readings from Figures 32 and 33 for- seats. The deformation is measured

with respect to the aircraft floor along the longitudinal axis

toward the nose of the aircraft, regardless of litter orientation.

5.11.5 Litter Restraint Harness Testing

The restraint used in existing military litters consists of two

straps wrapped around the litter. These straps should withstand a straight tensile minimum load of 2000 lb (4000-lb loop

strength). The maximum elongation should not be more than

3.0 in. under the straight pull (end-to-end) test on a minimum

strap length of 48 in. Elongation is restricted for litter

belts in order to minimize dynamic overshoot.

5.11.6 Litter System Test Requirements

5.11.6.1 Static Test Requirements

5.11.6.1.1 General: Table 14 presents the static test requirements for complete litter systems. Since previous studies have

shown that existing litters will not withstand the loads as

specified in this chapter, the assumption must be made that a

litter of sufficient strength will be developed prior to imple- menting these recommendations. If a pan or net to catch the

litter occupant is included in the system, it should also be

included in the static testing to demonstrate its adequacy.

5.11.6.1.2 Unidirectional Tests: The test loads for forward,

lateral, and downward loading of litter systems as presented in

Table 14 should be applied separately.

5.11.6.1.3 Combined Loads: Litter systems must demonstrate

no loss of system integrity under conditions of combined loads

as specified in Table 14.

5.11.6.1.4 Point of Load Application: The loads should be applied through a body block that simulates a supine occupant.

5.11.6.1.4.1 Forward (Longitudinal) - Lateral Tests: For systems using the existing litter, a rigid simulated litter may be

substituted for the actual litter. This will enable application of equal loads at all attachment points between the litter /-

146

S i l Ii i 'iiI
V

t •TABLE 14. LITTER SYSTEM STATIC TEST REQUIREMENTS

Test Loading direction

ref. with respect to Deformation

no. fuselage floor Load required requirements

I Forward See Figure 42 See Figure 42

2 Lateral See Figure 42 See Figure 42

D3 ownward See Figure 41 See Figure 41

4 Upward 8 G No requirement

5 Combined loading

Downward plus See Figure 41 See Figure 41

transverse load

along any radial

in the x-y-plane of the aircraft See i gure 42 See Figure 42

and the suspension system and allow testing of the suspension

system. The rigid litter substitution does not apply if the

litter system has adequate strength to take the loads.

5.11.6.1.4.2 Downward and Upward Tests: Downward and upward

loads may be applied to each vertical suspension point separately. If the suspension system has the tilt-limiting features, and the litter strength is adequate, then the load

should be applied at the center of gravity of the body block.

5.11.6.1.5 Deflection Measurements: Downward, forward (longitudinal), and lateral deflections should be measured at the

bracket attaching the litter to the suspension system.

5.11.6.1.6 Load Determination: The test load should be determined by multiplying the required load factor (G) as specified

in Table 14 by 250 lb.

5.11.6.2 Litter System Dynamic Test Requirements: A single

test to evaluate the vertical load-limiting system is required.

Litter systems with 95th-percentile anthropomorphic dummies and

30 lb of additional weight (250-lb total) in each litter should

be subjected to a triangular acceleration pulse of 48-G peak

and 0.054-sec duration (42-ft/sec velocity change).

147
The same test pulse tolerances, data, handling, and processing

requirements as presented for the seats in Section 5.10.2 apply.

At least three accelerometers should be placed in the dummy;

one in the head, one in the chest, and one in the pelvic region. The instruments should be positioned to sense accelerations in the vertical directions (x-axis of the supine

occupant, z-direction relative to the aircraft). The input

acceleration-time pulse also should be measured. It is advisable to use redundant accelerometers to sense the input pulse

to assure acquisition of the needed impact environment data.

5.12 DELETHALISATION OF COCKPIT AND CABIN INTERIORS

5.12.1 General

The kinematics of body action associated with aircraft crash

impacts are quite violent, even in accidents of moderate severity. The occupant's immediate environment should be designed

so that, when the body parts do flail and contact rigid or

semirigid structures, injury potential is minimized.

Several approaches are available to alleviate potential secondary impact problems. The most direct approach, which should

be taken if practical, is to relocate the hazardous structure

or object out of the occupant's reach. Such action is normally

subject to tradeoffs between safety and operational or human

engineering considerations. If relocation is not a viable alternative, the hazard might be reduced by mounting the offending structure on frangible or energy-absorbing supports and applying a padding material to distribute the contact force over

a larger area on the body member.

5.12.2 Occupant Strike Envelopes

5.12.2.1 Full Restraint: Body extremity strike envelopes are

presented in Figures 43 through 45 for a 95th-percentile Army aviator wearing a restraint system that meets the requirements

of MIL-S-58095(AV) (Reference 14). The restraint system consists of a lap belt, lap belt tieduwn strap, and two shoulder

straps. The forward motion shown in Figures 43 and 44 was obtained from a test utilizing a 95th-percentile anthropomorphic

dummy subjected to a spineward (-G ) acceleration of 30 G. The

lateral motion is based on an extfapolation of data from the

same 30-G test. In positions where an occupant is expected

to wear a helmet, the helmet dimensions must be added to the

envelope of head motion.

148

S6 JW
10 in.

Heel rest line for cockpit.

Aircraft loor me for

troop compartment.

Figure 43. Full-restraint extremity strike

envelope - side view.

5.12.2.2 Lap Belt-Only Restraint: Although upper torso restraint is required in new Army aircraft, strike envelopes for

a 95th-percentile aviator wearing a lap belt-only restraint are

presented in Figures 46 through 48 for possible use. They are

based on 4-G accelerations and 4 in. of torso movement away

from the seat laterally and forward. In positions where an

occupant is expected to wear a helmet, the helmet dimensions

must be added to the envelope of head motion.

5.12.2.3 Seat Orientations The strike envelopes of Figures

43 through 48 apply to all seat orientations.

149

_71

!•-•M •,•,-, :,, , , :'"'•-" • " ' ,'r"f
i - " I_ o _ -I

S• i10 in.

L / 

-\! ,

Figure 44. Full-restraint extremity strike

envelope - top view.

5.12.3 Environmental Hazards

5.12.3.1 PrimryHaards: The primary environmental hazards

are those rigid or seminrigid structural members within the extremity envelope of the head and chest. Since the upper torso,

and particularly the ivnad, is the most vulnerable part of the

body, maximum protectioii must be provided within its strike

envelope.

5.12.3.2 Secondary Haiards: Sec"..ndary environmental hazards

are those that could result A'r t.rapping or injuring the lower

extremities to the extent that one's ability to rapidly escape

would be compromibed. Areas within the lower extremity strike

envelope must also include ample protective design.

150
II

iI, .. .. , ,__

1 .0 i0 n!

10in

* o° "

_ _ _ ........... ., el rest line for cockpit.

Aircraft floor line for

troop compartment.

Figure 45. Full-restraint extremity strike

envelope - front view.

5.12.3.3 Tertiary Hazards: Tertiary environmental hazards are

those rigid and semirigid structural members that could cause

injury to flailing upper limbs to an extent that could reduce

an occupant's ability to operate escape hatches or perform

* other essential tasks.

5.12.4 Head Impact Hazards

5.12.4.1 Geometry of Probable Head Impact Surfaces: Typical

contact hazards in the cockpit area include window and door

* frames, consoles, controls and control columns, seat backs, electrical junction boxes, glare shields, and instrument panels.

Contact hazards commonly found in aircraft cabin areas include

window and door frames, seats, and fuselage structure. Use of

suitable energy-absorbing padding materials, frangible break-

.• away panels, smooth contoured surfaces, or ductile materials

151

LA',A
r

!-

10 in.

I...........

___._, Heel rest line for cockpit. Aircraft floor line tor

troop compartment.

Figure 46. Lap belt-On.' extremity strike envelope - side view.

in the typical hazard areas mentioned is recommended to reduce

the injury potential of occupied areas.

5.12.4.2 Tolerance to Head Impact: Protection of the head in

the form of protective helmets and energy-absorbing structure

and padding in the occupant's immediate environment is essential.

Tolerance levels for head impact are discussed in detail in

Volume II, and the reader should refer there for an understanding of the problem. However, for the case of forehead impact

on a flat surface, which is pertinent to the discussion of this

section, the most widely accepted collection of tolerance data

is represented in the tolerance curve of Figure 49. • )

152

_________
r - I, 404

>11

153
Il

0 044

00

4 0)

_____ 4 ..

41 .'

0)

54

414

coI

154
600

600 T

S6 0 I I

II

0

0 400

Dangerous to life

o 300

> 200

.4.

44 100

44 W Not to life dangerous I•}- A

0 2 4 6 8 10 12 30 100

Time duration of effective

acceleration, msec

Figure 49. Wayne State Tolerance Curve for the

human brain in forehead impacts

against plane, unyielding surfaces.

(From Reference 45)

5.12.4.3 Test Procedures: The simplest test procedure for

evaluating the effectiveness of protective structure and padding in preventing serious head injury makes use of an instrumented headform. The headform, equipped with an accelerometer,

can be propelled by a ram, dropped, or swung on a pendulum to

impact the surface to be evaluated. This procedure is described

in SAE J921 (Reference 46). The measured acceleration pulse can

be averaged for comparison with the Wayne State Tolerance Curve,

or integrated to compute a Severity Index, as discussed in Sec-

"tion 4.4.1 of Volume II.

45. Patrick, L. M., Lissner, H. R., and Gurdjian, E. S., SURVIVAL BY DESIGN - HEAD PROTECTION, Proceedings, Seventh

Stapp Car Crash Conference, Society of Automotive Engineers, Inc., New York, 1963.

46. SAE Recommended Practice, SAE J92lb, MOTOR VEHICLE INSTRUMENT PANEL LABORATORY IMPACT TEST PROCEDURE - HEAD AREA,

SAE Handbook, 1979, Part 2, Society of Automotive Engineers,

Inc., Warrendale, Pennsylvania, 1979, pp. 34.133-34.134.

155
II

5.12.5 Instrument Panel Structure Proxiijt

In most aircraft cockpits, the instrument panel and its supporting structure are placed directly above the pilot's lower

legs. The danger of impact from this proximity dictates that

designers consider using suitable energy-absorbing padding materials, frangible breakaway panels, or ductile panel materials

for structure within the lower leg strike envelope.

5.12.6 Rudder Pedal Configuration

Rudder pedals should be capable of supporting both the ball of

the foot and the heel, and provide a surrounding structure of

sufficient strength to prevent crushing and trapping of the

lower limbs. The geometry required by MIL-STD-1290(AV) (Reference 1) to prevent entrapment of feet is illustrated in Fig- J

ure 50.

A -- B C

Dimesion A, B, andC us eite

less than 2 in. or more than 6 in.

Figure 50. An~titoque, or rudder, pedal geometry

to pr'event entrapment of feet.

156

I"-'-
5.12.7 Controls and Control Columns

It is recommended that control columns be designed so that

fracture due to the occupant's striking the column will occur

at a point no more than 4 in. above the pivot point. The failture should occur in the form of a clean break, leaving no jagged or torn edges. Control columns that pass longitudinally

through the instrument panel are not recommended since these

tend to impale the crewmembers in severe longitudinal impacts.

However, where they are used they should be equipped with a

frangible or energy-absorbing section similar to automotive

steering columns.

5.12.8 Sighting and Visionic Systems

Delethalization of the copilot/gunner (CPG) station of an attack or scout helicopter equipped with a weapon sighting optical relay tube (ORT) can present a difficult design problem.

"The cockpit should be designed to minimize the probability of

the CPG head/neck striking the ORT and minimize injury if the

CPG should strike the ORT, for both the "head-up" and "headdown" CPG positions. Some of the options available to the designer given this task are:

* ORT Eyepiece Relocation - Consideration should be

given to reducing occupant strike hazards by moving

the ORT further away from the CPG.

* Restraint System - The restraint system of Figure 22

would offer improved upper torso restraint, particu- larly when combined with the power-haulback inertia

reel.

9 Inflatable Restraint - Consideration should be given

to the inflatable restraint system (IBAHRS) discussed

in Section 5.7.2.4. This type of restraint harness can prevent injury to the CPG in both the erect and

S~head down position by reducing slack, supporting the I

head, ard increasing the surface area of the body

over which the harness reacts.

a Frangile/Breakaway Features - ORT or ORT components

designed to be frangible should break away at a total

force not to exceed 500 lb. For the frangible ORT,

this force should be applied along any direction of

loading within the plane normal to the axis of the

ORT, as well as along the axis of the ORT. Breakaway point(F) of the ORT should be outside the head

strike envelope.

157

S*~i ... / 1 I -I IIIII. .-
A

o Collapsible Features - If the ORT is designed to col- 4

lapse in order to avoid injuring the CPG, the collapse load along the axis of the ORT should not exceed 500 lb. Figure 51 illustrates one crushable

sight eyepiece concept (from Reference 47). Two advantages of the crushable sight eyepiece are that it 4

is always available and it should function regard- less of head location. A helmet crash-absorber pad

would attenuate crash loads to the helmet when available crushing is expended.

e Power-Haulback Inertia Reel (PHBIR) - On the basis of

Air Force testing accomplished for the development of

PHBIR, the retraction time is 0.3 to 0.4 sec, which

is too slow for effectiveness in most crashes. If

this time were reduced, the retraction velocity of

the torso would have to be increased considerably

over the current limit of 9 ft/sec. A retract4.on

velocity greater than this is not recommended due

to the lack of human tolerance data on this type of

loading. In a crash with a single pulse of 30-G

peak and 50-ft/sec velocity change, the retraction

velocity should be approximately 25 ft/sec; therefore, the known tolerance limits would be exceeded

at the higher velocity. In summary, the PHBIR, as

currently qualified under both Air Force and Navy

military specifications, requires excessive time to

position the torso by crash sensing. To be fully

effective, the system should move the torso into po- sition in approximately 0.06 sec, but the resulting

acceleration would exceed known human tolerance

limits. The primary crashworthiness advantage of

the PHBIR would be as a manually activated tightening device for the head-up CPG position; the PHBIR offers only limited advantage for the head-down CPG

position.

5.12.9 Energy-Absorbing Requirements

for Cockpit and Cabin Interiors

5.12.9.1 General: To minimize occupant injury, the acceleration experienced during secondary impacts of the occupant with

surrounding structures must be reduced to a tolerable level.

The areas of contact to be considered for energy absorption

47. Fox, R., Kawa, M., and Sharp, E., DESIGNING CRASHWORTHINESS INTO THE YAH-63, paper presented at the Aircraft

Crashworthiness Symposium, University of Cincinnati, Cincinnati, Ohio, October 1975.

158
F]

Normal operation

SPH-4 helmet

Sreference

Helmet crashenergy absorber Pressure support

e y rarea anproximately

Crash energy- 6 in.

absorbing material

Preloaded

slip rings Inner contours

Optical relay Crush stroke,,ofc ms

tube Outer contour

"face mask

20 mun

eye relief

After crash stroke

(- Helmet crash-

?1 energy absorber

Yielded

material /

Deformed tube

, Figure 51. Crushable eyepiece concept. (From Rpcerence 47)

include instrument panels, glare shields, other interior surfaces within the occupant's strike envelope, and seat cushions.

A padding material should not only reduce the decelerative

"force exerted on an impacting body segment, but should distribute the load in order to produce a more uniform pressure of

safe magnitude.

In order to prevent head injury, matezials must be carefully

selected to absorb and attenuate the energy of impact. The

159

ii
material must reduce the level of acceleration, the rate of

onset, and the amount of energy transmitted to the head.

5.12.9.2 Paddiaq Material Properties: The selection of a

foam material for vehicle energy-absorbing applications involves an evaluation of its processability; its mechanical,

thermal, and chemical properties; as well as its cost. Along

with the primary foam materials, the characteristics of adhesives and surface coatings must be considered, particularly

with respect to emission of smoke and toxic vapors. The characteristics of suitable materials for such use are listed below:

"* Adaptability and ease of processing

"* High energy dissipation

"* Effective load distribution

"* Low rebound

"* Temperature insensitivity i

* Low water absorption

e Resistance to chemicals, oil, ultraviolet radiation,

and sunlight

* Nontoxic fume generation

* Favorable flammability rating

* Minimal smoke generation

* Durability and long life j

* Cost competitive

o Aesthetically acceptable

5.12.9.3 Standard Test Methods: ASTM standard test procedures I

are widely used by manufacturers to specify various properties

of a particular type of material. Table 15 summarizes ASTM

test methods and specifications for flexible cellular plastics

that provide a basis for comparison of materials. Here it may

be noted that most ASTM tests involve simple tests, whereas

the operational environment involves dynamic loading and more

complex conditions.6

160
TABLE 15. SUMMARY OF ASTM TEST METHODS AND

SPECIFICATIONS FOR FLEXIBLE

CELLULAR PLASTICS (Reference 48)

D1564-71" Testing Flexible Cellular MaterialsSlab Urethane Foam

D1667-76* Specification for Flexible Cellular

Materials - Vinyl Chloride Polymers

and Copolymers (Closed-Cell Sponge)

D1565-76* Specification for Flexible Cellular

Materials - Vinyl Chloride Polymers

and Copolymers (Open-Cell Foam)

DIC55-69" Specification for Flexible Cellular

(1975) Materials - Latex Foam

D1056-73" Specification for Flexible Cellular

Materials - Sponge or Expanded Rubber

D3575-77 Testing Flexible Cellular Materials

Made From Olefin Plastics

D1596-64* Test for Shock-Absorbing Characteristics

(1976) of Package Cushioning Materials

D2221-68* Test foz Creep Properties of Package

(1973) Cushioning Materials

D1372-64* Testing Package Cushioning Materials

(1976)

D696-70" Test for Coefficient of Linear Thermal

Expansion of Plastics

E143-61* Test for Shear Modulus at Room Temp-

(1972) erature

D412-75" Tests for Rubber Properties in Tension

D1433-76' Test for Rate of Burning and/or Excent

and Time of Burning of Flexible Thin

Plastic Sheeting Supported on a 45-

degree Incline

D1692-76 Test for Rate of Burning and/or Extent

and Time of Burning of Cellular Plastics Using a Speciman Supported by a

Horizontal Screen

*Indicates that the standard has been approved as

American National Standard by the American National

Standards Institute.

161
In particular, ASTM D 1564-71 describes "Standard Methods of

Testing Flexible Cellular Materials - Slab Urethane Foam" (Reference 48). Among other tests, there are compression-set and

load-deflection tests.

The above tests provide results that specify the material, but

do not necessarily portray its performance under actual impact

situations. A simple dynamic drop test, such as ASTM D1596-64

(1976), "Standard Test Method for Shock-Absorbing Characteristics of Package Cushioning Materials" (Reference 49), more

closely simulates actual impact conditions.

Other standard test procedures include SAE J815, "Load Deflection Testing of Urethane Foams for Automotive Seating" (Reference 50), which points out the factors of interest in testing

materials for vehicle seat cushions: the thickness of the padding under the average passenger load, a measurement that indicates the initial softness, and a measurement that indicates

resiliency.

Also, SAE J388, "Dynamic Flex Fatigue Test for Slab Urethane

Foam" (Reference 51), describes procedures for evaluating the

loss of thickness and the amount of structural breakdown of

slab urethane foam seating materials.

SAE J921, "Motor Vehicle Instrument Panel Laboratory Impact

Test Procedure - Head Area," describes a test procedure for evaluating the head impact characteristics of such areas as instrument panels (Reference 46).

48. ASTM D 1564-71, STANDARD METHODS OF TESTING FLEXIBLE CELLULAR MATERIALS - SLAB URETHANE FOAM, 1977 Annual Book of

ASTM Standards, R. P. Lukens, et al., eds., American Society for Testing and Materials, Easton, Maryland, 1977,

Part 38.

49. ASTM D 1596-64, STANDARD TEST METHOD FOR SHOCK-ABSORBING

CHARACTERISTICS OF PACKAGE CUSHIONING MATERIALS, 1977 Annual Book of ASTM Standards, R. P. Lukens, et al., eds.,

American Society for Testing and Materials, Easton, Maryland, 1977, Part 30.

50. SAE Recommended Practice, SAE J815, LOAD DEFLECTION TESTING OF URETHANE FOAMS FOR AUTOMOTIVE SEATING, SAE Handbook

1979, Part 2, Society of Automotive Engineers, Inc., Warrendale, Pennsylvania, 1979, p. 34.31.

51. SAE Recommended Practice, SAE J388, DYNAMIC FLEX FATIGUE

TEST FOR SLAB POLYURETHANE FOAM, SAE Handbook, 1979,

Part 2, Society of Automotive Engineers, Inc., Warrendale,

Pennsylvania, 1970, pp. 34.28-34.30.

162
S 5.12.9.4 Acceptable Stress-Strain Characteristics: Energyabsozbing materials with stress-strain curves that fall between the limits shown in Figure 52 will offer reasonable 7 survival potential for head impacts at velocities of up to

22 ft/sec where a padding thickness of 2.0 in. is used. The

impact surface is assumed to be flat; the data from which

r Figure 52 was developed were obtained for simulated head impacts on flat surfaces with energy levels up to 84 ft-lb, i.e., l1.2-1b head weight x 7.5-ft drop height. The acceleration

of the head should not exceed 120 G at an impact velocity of

20 ft/sec (or greater) while a higher level of acceleration

can be sustained at lower velocities (shorter pulse duration).

This accounts for the different stress-versus-strain values

shown in Figure 52, i.e., a higher G or crush stress is acceptable at the lower design velocity expected for the tnin padding.

60 6J

40L • • •Acceptabbe

• U ACCIpt sbl . 4 ZA&Clptabl, I unacceptable

(A)For p&Jding th: ck:o:a: greater th*0 one inch () ror padding th;Ckn,.8 le thAn one

i l0 1 1 I , , ,

29 4.1 3 . ; 2' 40 6 in"

Strain, percent

Figure 52. Recommended stress-strain properties for padding

material for head contact, with cushion thickness

of at least 1.5 in.

The criteria of Figure 52 are to be satisfied by the padding

material over the entire anticipated operating temperature

rapge if the potential for survival is to be maintained. Practical considerations and risk analysis, however, may reduce

the temperature range requirements. Temperature sensitivity

must be considered as a padding material selection criterion.

Other padding material evaluation methods are discussed in

Section 10.9.4 of Volume IV.

Stress-strain curves for several polyurethane-foamed plastics

are shown in Figure 53. The curves show that a density of

163

S,-- -ij., ,~ ~--= -- _ _ _ _ _ _
>f5

70 I 4

70 :0  4V

• 660

"'• 50

60 LI

so 0

a) ' FIX/

2 0

U,-

V&

0 20 40 60 80 100

Compressive strain, percent

Figure 53. Effect of density on stress-strain curves

for polyurethane-foamed plastic.

164
3@

l3 b/ft or less will satisfy the criteria of Figure'52 (superimposed as a crosshatched area) over at least part of the oper- ational temperature range.

5.12.9.5 Application of Padding Material: In the absence of

data for extremity impacts, it is assumed that padding material that is suitable for head impact protection will be suitable also for protecting extremities. Extremity impacts are

not likely to have the potentially severe effectb of head impacts. It is suggested that areas within the extremity strike

envelope having radii of 2 in. or less be padded and that such

padding have a minitnum thickness of 0.75 in.

Caution must be exercised in padding sharp edges and corners.

Padding installed in a manner that allows it to be broken away

from the corner or cut through by sharp edges offers no pro- tection. It is recommended that edges and corners to be padded

have a minimum radius of 0.5 in. prior to padding. A definite

volume of the padding must be crushed to absorb the initial

kinetic energy of the head and protective helmet.

5.12.9.6 Ductile Materials: In cases where the use of padding material iS impractical or the thickness allowed is inadequate

to provide the necessary protection, ductile energy-absorbing materials or frangible breakaway panels should be used where

possible. Window and door frames, control columns, electrical junction boxes, etc., should be designed with large radii

(1 in. or more) rather than with sharp edges and corners.

Swearingen concluded in Reference 52 that at impact velocities

of 30 ft/sec against rigid structure padded with materials even

6 in. thick, unconsciousness, concussion, and/or fatal head

injuries will be produced. Where possible, a combination of

deformable structure and padding material should be considered

to absorb the impact energy and to adequately distribute the

forces over the face. Surfaces to which this combination

should be applied are instrument panels, seat backs, bulkheads, and any other structure that the head may impact during the

crash sequence.

52. Swearingen, J. J., EVALUATIONS OF VARIOUS PADDING MATERIALS FOR CRASH PROTECTION, FAA Technical Report AM 66-40,

Federal Aviation Administration, Civil Aeromedical Institute, Oklahoma City, Oklahoma, December 1966, AD 647048.

165

l -n . , . .,. . . .. -, .. - , - ,~ *~t , , *--~- .. ... .-- . - ___________
Yes No N/A

5.13 DESIGN CHECKLISTS

5.13.1 General Design Checklist

1. For load-limited seats, do all materials in

critical structural members possess a minimum elongation of 5 percent in the principal load direction?

2. For nonload-limited seats, do materials in

critical structural members possess a minimum of 10 percent elongation?

3. Is there adherence to the flammability and

toxicity requirements of Chapter 6?

4. In load-limited portions of the seat, where

loads can be predicted accurately, are minimum margins of safety for shear and tensile

bolts 5 and 10 percent, respectively?

5. In nonload-limited portions of the seat, are

minimum margins of safety for shear and tensile bolts 15 and 25 percent, respectively?

6. In the vicinity of welded joints, have

cross-sectional areas been increased by 10 percent to account for uncertainties,

stress concentrations, etc.?

7. Have seat attachments been designed so that

neither buckling nor warping of the floor or

bulkhead will interfere with seat operation

or seat integrity in a crash?

8. Has the restraint system anchorage been

designed so that the restraint system will

function effectively as the seat strokes?

9. Is the use of castings avoided in the primary seat structure?

10. If castings are uscd, are they sufficiently

ductile, or does the design allow for realistic seat deformation during crash load

application without failure of the castings?

166 L _____

, i I I % 
Yes No N!_A

11. Do nonmetallic materials comply with FAR 25?

12. Can troop seats be removed in 20 sec per

occupant position?

5.13.2 Seat Strength and Deformation Checklist

1. Does the seat meet the longitudinal load- H

deformation requirements of Figure 32?

2. Will the seat withstand a 12-G aftward load?

* 3. Is the vertical energy-absorption system designed for a load factor of 11.5 G based on

the effective weight of the 50th-percentile

aviator or trooper? A

4. Does the crewseat possess a minimum vertical

stroke distance of 12 in. (from the lowest

vertical adjustment position)? -

S. Has the use of a variable-force energy absorber been considered?

6. Does the troop seat possess a minimum of

17 in. of vertical stroke?

7. Does the seat have a capability of withstanding an upward load of 8 G?

8. Does the seat meet the lateral loaddeformation requirements of Figure 33?

9. Are the static attachment strengths for a

components mounted on the seat, such as

armored panels, based on the following load

factors?

S* Downward: 51 G

o Upward: 10 G

o Forwards 35 G

e Aftwardi 15 G * Lateral: 25 G

167
Yes No &A

5.13.3 Seat Cushions Checklist

1. Are seat cushions of the type that minimize

dynamic overshoot in vertical deceleration?

2. Is the thickness of the compressed seat

cushion between 0.5 and 0.75 in., or has it

been demonstrated that the cushion design

and material properties produce a beneficial

result?

5.13.4 Litter Strength and Deformation

Requirements Checklist

1. Does the litter system possess the vertical

strength-deformation capability of Figure 41,

based on an occupant weight of 250 Ib?

2. Does the litter system possess the capability of withstanding an upward load of 8 G?

3. Does the litter system meet the lateral

load-deformation requirements of Figure 42?

4. Can the litters be loaded laterally into the

aircraft?

5. Can the complete set of litters be loaded

and unloaded to flight readiness in 10 sec

or less in an emergency situation?

6. Does the litter system eliminate need for

special mounting hardware that remains

attached to the aircraft?

7. Can the standard cargo tiedown system be

used as the primary litter system attachment to the aircraft structure?

8. Will the litter installation accept the current standard military litter?

9. Does the installation support the litter in

such a manner as tc develop the maximum

i..: tarrying capibil.'.y of the standard

litte'?

168
Yes No N/A

E 10. Would the litter installation be adaptable

to a new and improved military litter design?

11. Does the litter installation, when removed

from the aircraft, leave the aircraft free

of all protuberances, brackets, ard other

objectionable operational hazards?

5.13.5 Restraint System Design Checklist

1. Are the lap belt anchor points located so

that a maximum angle of 55 degrees and a miniimum angle of 45 degrees exists between the

lap belt and the buttock reference line, as

illustrated in Figure 27?

2. Is the point where the shoulder harness is

attached to or passes through the seat back

between 26 and 27 in. above the seat reference point?

3. Does the shoulder harness anchorage or guide

on the seat back permit no more than 0.5-in.

lateral clearance?

4. Does the shoulder harness guide on the seat

back have a 0.25-in. minimum radius as illustrated in Figure 30?

5. Is the lap belt tiedown strap (crotch strap)

attached to the seat pan centerline at a

point 14 to 15 in. forward of the seat back?

6. Are the forces required for adjustment of

all vebbing item lengths no greater than

30 lb?

S7. Are the lap belt adjusters located so as to

not exert pressure on the iliac crests?

8. Are the shoulder strap adjusters located low

enough on the chest to avoid concentrated

pressure on the collarbones?

9. Do the restraint harness subassemblies meet

the minimum load and maximum elongation requirements of Tables 8 and 9?

169
Yes No N

Have the stitched S10. joints in the restraint

harness been designed according to the criteria discussed in Section 5.7.4.3 and do

the joints have a 30-percent margin?

11. Is a minimum webbing thickness of 0.055 in.

used on all restraint harncss components?

12. Do the restraint harness components meet the

following minimum width requirements?

* Lap Belt - 2-1/4 in.

* Shoulder strap - 2 in.

e Tiedown Strap - 1-1/2 in.-

13. Do all webbing fittings, over which webbing is wrapped, possess the 0.062-in. minimum

radius illustrated in Figure 30?

14. Does the restraint harness have a singlepoint release system that can be released

after being exposed to design crash loads

by exerting a 30-lb force with one finger

or a 50-lb force with one finger when supporting the entire weight of the occupant?

15. Is the single-point release protected fron,

inadvertent release?

5.13.6 Protective Padding Checklist

1. Are all areas within the extremity strike

envelope, having radii of 2 in. or less,

padded with a minimum thickness of 0.75 in.?

2. Do padded corners of edges have a minimum

unpadded radius of 0.5 in.?

3. Are ductile energy-absorbing supports used

where possible under padding, particularly

where head impact is likely?

170
Yes No N/A

5.13.7 Cockpit Controls and Equipment Checklist

1. Are rudder pedals separated from each other

and from adjacent structure by less than

2 in. or more than 6 in., as illustrated

r in Figure 50?

2. Are controls and control columns designed

so that fracture due to an occupant's striking the column will occur at a point no

more than 4 in. above the pivot point, and 4 •so that the failure will be clean without

jagged or torn edges, or are they equipped with an energy-absorbing section?

171

____________________________i4

I.S

................-
6. AIRCRAFT POSTCRASH SURVIVAL

6.1 INTRODUCTION

This chapter presents the criteria that are to be applied in

designing postcrash survival into an aircraft. Although initial crashworthy considerations, such as maintaining structural

integrity around the occupant and reducing the crash forces

transmitted to the occupant, are of primary importance in survival, hazardous postcrash conditions must be prevented or reduced if the occupant is ultimately to survive. The threat of

postcrash fire must be minimized and adequate escape and rescue

provisions must be incorporated into the aircraft.

The criteria presented in this section include those for designing fuel, oil, and hydraulic systems to minimize the occurrence of postcrash fires; for selecting less flammable interior

materials; for selecting provisions that increase survival

chances during aircraft ditchings; and for designing emergency

escape provisions and crash locator beacons. The user is referred to Volume V for more complete information and reference sources.

6.2 FUEL SYSTEM DESIGN CRITERIA

The following criteria are applicable to all auxiliary fuel

systems, such as ferry systems and extended range systems, as

well as to the primary aircraft fuel system.

6.2.1 General

The fuel system must be designed to minimize fuel spillage during and after all survivable crash impacts. It also must be

designed to prevent spillage of fuel through the vents during

a rollover or in any other adverse attitude. Spillage that

cannot be avoided, such as during the functioning of selfsealing breakaway couplings, must be precluded from ignition

by controlling ignition sources (see Section 5.5 of Volume V).

6.2.2 Fuel Tanks

6.2.2.1 Fuel Tank Location: The location of fuel tanks in an

aircraft is of considerable importance in minimizing the postcrash fire hazard. The location must be considered with respect to occupants, ignition sources, and probable impact

areas. The fuel tanks should be located as far as possible from probable impact areas and from areas where structural deformation might cause crushing or penetration of the tank. If

possible, fuel tanks should not be installed:

172
a Immediately adjacent to occupiable areas.

* Immediately adjacent to engine compartments.

o Immediately adjacent to electrical compartments.

* Under heavy masses, such as transmissions and engines.

e Near the bottom of the fuselage.

* Over landing gears.

e In leading edges or anticipated failure areas of

wings.

6.2.2.2 Fuel Tank Construction: Fuel tanks should have smooth,

regular shapes, with the sump area contoured gradually into the

tank bottom. All concave corners should have a minimum radius

of 3 in., and all convex corners a minimum radius of I in.

All fuel tanks must be fabricated from crash-resistant material

which meets or exceeds the requirements of MIL-T-27422 (Reference 53). All fuel tank fittings must have a tank pullout

strength that meets or exceeds that specified in MIL-T-27422.

A self-sealing, breakaway, tank-to-tdnk coupling should be used

wherever two tanks are connected directly with no intervening

fuel line.

6.2.3 Fuel Lines

Fuel lines should be constructed and routed so as to withstand

all survivable crash impacts. This may b(% done by allowing

the lines to elongate or shift with deforming aircraft structure rather than being forced to carry high tensile loads.

6.2.3.1 Fuel Line Construction: All fuel lines that could be

readily damaged in an accident of severity up to that indicated

in Table 2 should consist of flexible hose with a steel-braided

outer sheath, where possible. The hoses should be capable of

elongating 20 percent without the hose assembly spilling any

fuel. If "stretchable" (20-percent minimum elongation) hoses

are not used, all hoses should be a minimum of 20 percent

longer than necessary to provide added length for structural

displacement.

53. Military Specification, MIL-T-27422B, TANK, FUEL, CRASHRESISTANT, AIRCRAFT, Department of Defense, Washington,

D. C., 13 April 1971.

173

S. '_ . 'i4 " • r , . .. •=" _ _ _i 
When the hose assemblies are subjected to pure tension loads

or to loads applied at a 90-degree angle to the longitudinal

axis of the end fitting, as shown in Figure 54, hoses must not

pull out of their end fittings, nor the end fittings break, at

less than the minimum loads shown in Table 16. Loads must be

applied at a constant rate not exceeding 20 in./min.

The number of fuel line couplings should be held to a minimum.

Wherever possible, a single, one-piece hose should be run

through a bulkhead opening rather than being attached to the

bulkhead with rigid fittings. The opening should be I in. larger in diameter than the hose diameter, with the hose stabilized by a frangible panel or structure. A grommet must be installed in the opening to preclude wear on the hose. Selfsealing breakaway couplings must be used whenever a line goes

through a firewall so that the line will seal if the engine

is displaced during crash impact. Breakaway ..)uplings will

not be required if the engine is tied down to a strength level

of 20 G , 20 G , and 18 G and if the engine is located so that

crushina of thi lines andYfittings is not likely in any survivable accident.

All fuel line-to-fuel tank connections must consist of selfsealing breakaway couplings. These couplings must be recessed

into the tank so that the tank half does not protrude outside

the tank wall more than 1/2 in. after coupling separation. The

shape of the tank coupling half must be basically smooth to

avoid snagging on adjacent structures or cutting the tank wall.

An acceptable substitute for a breakaway valve is a hose constructed of material identical to that of the tank with an

end fitting strength equal to 80 percent of the tank tear-out

strength (MIL-T-27422, Paragraph 4.6.5).

6.2.3.2 Fuel Line Location: Fuel lines should be located as

far as possible from probable impact areas and areas where

structural deformation can cause crushing, penetration, or excessive tensile loading of the lines. When fuel lines must be

routed through areas of probable large displacement, such as

wing-to-fuselage attachment points, self-sealing breakaway

couplings must be incorporated into the lines to allow for complete line separation with a minimum of fuel spillage.

Fuel lines should not be routed in the following areas:

"• Near the bottom of the fuselage.

"* Over landing gears.

"* Uniter, in front of, or at the sides of heavy masses,

such as engines and transmissions.

174

1,IIIi
V

LodLoad S~Load

I;

Tension tests

Load Load

II

I!:

9 0 -degree tests

Figure 54. Hose assembly test modes.

175

h"/i///1///
TABLE 16. REQUIRED MINIMUM INDIVIDUAL LOADS FOR STANDARD

HOSE AND HOSE-END FITTING COMBINATIONS

Minimum Minimum

Hose end Fitting tensile load bending load

fitting type size* _(Ib) (lb)

STRAIGHT -4 575 450

Tension: -6 600 450 a

-8 900 700

-10 1250 950

-12 1900 1050

Bending: -16 1950 1450

-20 2300 1600

-24 2350 2750

-32 3500 4000

900 ELBOW -4** 575 800

Tension: -6** 600 850

900 1250

-10 1250 575

-12 1900 675

Bending: -16 1950 1200

-20 2300 1250

-24 2350 2025

-32 3500 3500

450 ELBOW -4** 575

Tension: -6** 600 425

-8** 900 425

-10 1250 425

-12 1900 600

Bending: -16 1950 1000

-20 2300 1600

-24 2350 2400

-32 3500 3700

*Fitting size given in 1/16 in. units, i.e., -4 = 4/16 or

1/4 in.

**Elbow material is steel.

176

_ _ I
-M -_ - , - - 7

* In the leading edges of wings.

* In anticipated areas of rotor blade impact.

* Adjacent to electrical wiring.

Fuel lines must not be routed through electrical compartments

or occupiable areas unless they are shrouded or otherwise designed to prevent spillage.

In order to protect the lines from impact damage, fuel lines

should be routed along heavier -ic structural members wherever possible. All fuel lines be adequately supported by

frangible clamps attdched to other structure.

Fuel lines should be grouped together and exit a fuel tank in

one centralized location. This location shou'l be in the area

of the tank that is least vulnerable to anti. pated crash loads

and structiral deformations. However, ballistic vulnerability

considerations may modify this requirement.

The number of fuel lines in tne engine compartment should be

minimized. When more than one line enters an engine compartment, the lines should be grouped together and pass through the

firewall in a protected location unless the structural integ- rity of the firewall would be compromised.

6.2.4 Frangible Attachments

Frangible structurpq or frangible bolts must be used at all

attachment points between fuel tanks and aircraft structure to

prevent fuel tank components from being torn out of the tank

wall during impact. Frangible attachments should be used at

other points in the flammable fluid systems where aircraft

structural deformation could lead to flammable fluid leakage.

The load required to separate a frangiole attachment from its

support structure must be between 25 and 50 percent of the

minimum load required to fail the weakest component in the attached system, as illustrated in Figure 55. (The failure load

of the attached system components may be determined either by

analytical computations or by testing methods based upon the

failure modes most likely to occur during crash impact.) To

prevent inadvertent separation, failure loads must be at least

five times normal operational and service loads at the frangible attachment location.

A frangible attachment must separdte whenever the required load

(as defined above) is applied in th- mod,-s moat likely to occur

177

--- "-NNW",,m•• ,
Aircraft •Tank wall structure

Metal tank fitting

Frangible bolt

Aircraft

structure--

Flange---- Shear plane

ITEM LOWEST FAILURE LOAD (lb)* FAILURE MODEJ

Aircraft

structure 4000 Shear

Tank fitting 3000 Pull out of

tank

Flange 5000 Shear

Franl4ible bolt Not more than Break

3000 (tension-shear)

=1500

Not less than

3000

*Loads may or may not be representative; values are for

explanatory purposes only.

Figure 55. Sample frangible attachment separation

load calculation.

during crash impact. These modes--whether tension, shear, compression, or combinations thereof, such as bending Itensionshear)--must be determined for each attachment by analyzing

the surrounding aircraft structure and probable impact forces

and directions.

178

04 -

rI
)d

- , All frangible devices must be statically tested in the three

most likely anticipated modes of separation. Test loads must

be applied at a constant rate, not exceeding 20 in./min, until

failure occurs. In addition, all frangible attachments must

be proof tested under dynamic loading conditions in the three

most likely anticipated modes of operation. The test load must

be applied in less than 0.005 sec, and the velocity change experienced by the loading jig must be 36 ± 3 ft/sec.

6.2.5 Self-Sealing Breakaway Valves

Self-sealing breakaway valves should be installed at all fuel

tank-to-fuel line connections, tank-to-tank interconnects, and

at other points in the fuel system where aircraft structural

deformation could lead to system failure. The valves must

allow only a minimal amount of spillage upon separation and

should permit no external leakage when partially separated.

The load required to separate a breakaway valve must be between

25 and 50 percent of the minimum load required to fail the

weakest component in the attached system, as illustrated in

Figure 56. To prevent inadvertent actuation during flight and

maintenance operations, the separation load must be greater

than five times normal operational and service loads at the

coupling location. To avoid complete or partial breakaway

coupling separation during maintenance operations, the separation load must never be less than 300 lb, regardless of the

fuel line size.

A breakaway valve must separate and seal whenever the required

load (as defined above) is applied in the modes most likely to

occur during crash impact. These modes, whether tension, shear,

compression, or combinations thereof, must be determined for

each coupling by analyzing the surrounding aircraft structure

and probable impact forces and directions.

A

All breakaway valves must be subjected to static tensile and

shear loads to establish the load required for separation, nature of separation, leakage during valve actuation, general

valve functioning, and leakage following valve actuation. The

rate of load application must not be greater than 20 in./min.

Tests to be used where applicable are shown in Figure 57.

In addition, all breakaway valves must be proof tested under

dynamic loading conditions. The valves must be tested in the

three most likely anticipated modes of separation. The test

configurations should be similar to those shown in Figure 57.

The load must be applied in less than 0.005 se%, and the velocity change experienced by the loading jig must be 36 f 3 ft/sec.

179

'll ! i l i t ll ...J ill .. .Ii '] •-' I• - .. . ..
[I

Airframe structure

Tank

Hose end

Flex hose

Metal tank fitting

Breakaway valve

Frangible section

ITEM TLOWEST FAILURE LUAD (lb)* FAILURE 14CDE

Flex hose 3000 Tensile breakage

Flex hose 1500 Pull out of end

fitting Tank fitting 7500 Pull out of tank A

Hose end coupling 1650 Break (bending)

Breakaway valve 2500 Pull out of tai.k . fitting

Breakaway valve Not more than Break at fran1500 = 750 gible section

Not less than

1500-- 375

*Loads may or may not be representative; values are for

explanatory purposes only.

Figure 56. Typical method of breakaway load calculation

for fuel tank-to-line breakaway valve.

All breakaway valves must incorporate positive proVisions for

ascertaining that the valve is locked together during normal

installation and service. In addition, all breakaway valves

must incorporate provisions in their design to prevent uncoup- ling due to operational shocx.s, vibrations, accelerations, etc.

180
II S~Load

il i •Load

Load 1

Static tensile test Static shear test I

Load

Load

Hose

Load

Static bending Static shear test (tensile-shear) test (tank-to-tank coupling)

Figure 57. Static tensile and shear test modes for self-sealing breakaway valves.

181

S .. ......... 4.n.. Ia.. *-I i 14 I in - ,,--I . . J I Illl• "]
6.2.6 Fuel Drains

All fuel tank drains must be recessed into the tank so that

part of the drain protrudes outside the tank wall. All attachments of fuel drains to aircraft structure must be made with

frangible fasteners.

The number of fuel line drains should be held to a minimum by

designing the fuel system to avoid low points in the lines.

If drain lines are necessary, they must be made of low-strength

materials.

Drain valves for tanks and lines must be designed to be positive locking in the closed position. Fuel drain actuation

must not require the operator to lie down under the aircraft.

Drains should be located where discharged fuel will not cause

an added fire hazard.

6.2.7 Filler Units

The filler unit must be fastened to the structure with a frangible attachment, and filler caps must be recessed into the

tank wall to ensure that the cap remains with the tank if the

tank moves relative to the aircraft structure.

Long filler necks should be avoided if possible. If they must

be used, they should be fabricated from frangible materials

and designed so that the filler cap remains with the tank and

does not snag on the aircraft structure during impact.

Tank fillers must not be located adjacent to engine intakes or

exhausts where flammable vapors could be ingested and ignited.

6.2.8 Fuel Boost Pumps

Boost pumps should be selected according to the following order

of preference:

1. Suction system, engine-mounted pump.

2. Air-driven, tank-mounted or in-line pump.

3. In-line electric pump.

4. Electrically operated tank-mounted pump.

Pumps mounted within the fuel tanks should be rigidly bolted

to the fuel tank only. If the pump must be supported or attoched to the aircraft structure, a frangible attachment should

be used.

182

----.~. .4,-
The state of the S.k" art in fuel system design has shown that electrically driven boost pumps can be eliminated. Air-driven

boost pumps and engine-mounted suction-type boost pumps now in

operation are much less hazardous alternative solutions.

If electric boost pumps are used, the electrical wires must

contain 6 in. of extra length at the pump connection to accommodate crash-induced structural deformation. The wires also

must be shrouded to prevent their being cut during crash impact. Nonsparking breakaway wire disconnects may be used in

lieu of the extra wire length.

6.2.9 Fuel Filters and Strainers

Fuel filters and strainers should not be located within the

engine compartment or adjacent to engine intakes or exhausts,

if at all possible.

Filters and strainers should retain the smallest possible quantity of fuel.

Filters and strainers must have a structural attachment capable

of withstanding a 30-G load applied in any direction.

Self-sealing breakaway valves should be used to attach fuel

lines to fuel filters and strainers in those locations where

structural displacement is likely to cause a separation of

those components.

6.2.10 Fuel Valves

* The number of fuel valves should be kept to a minimum.

Large valves (e.g., fuel shutoff valves) must have a structural

attachment capable of withstanding a 30-G load applied in any

direction. Self-sealing breakaway couplings should be used at

the valve-fuel line connections. Small valves (e.g., check

valves) must be fastened to the aircraft structure with frangible attachments.

If electrically operated valves are used, they should be

mounted on bulkheads so that the electrical wires are on one

side of the bulkhead and the valves and lines are on the other

side.

6.2.11 Fuel Quantity Indicators

Fuel counters and float-type quantity indicators are preferred

over rigid capacitance probes to preclude puncture of the fuel

tank during impact. If a capacitance probe must be used, it

183

i
5 -

should be fabricated from material possessing as low a flexural

rigidity as is consistent with operational requirements. A

slightly rounded shoe should be incorporated at the probe bottom end to avoid any tank-cutting tendency. Consideration 31

should be given to the use of frangible low-flexural rigidity

curved probes to reduce the danger of puncturing the tank during crash impact. The probe may also be mounted frangibly or

at an angle.

If tank-mounted quantity indicators must be attached to the

aircraft structure, frangible attachbrrnts must be used.

6.2.12 Vents

Vent systems must be designed to prevent fuel flow through the

vent lines regardless of aircraft attitude or vent line failure. For this reason, high-strength fittings should be used

between the metal insert in the tank and the vent line. If

the vent outlet must be supported, it should be supported by

frangible attachments. The vent line should be made of wirecovered flexible hose and should be routed so that it cannot be

snagged in displacing structure during a crash. Self-sealing

breakaway valves must be used at the tank-to-line attachment

if there is danger of the tank being torn free of the supporting structure.

Vent lines should be routed inside the fuel tank in such a manner that, if rollover occurs, spillage cannot continue. This

can be accomplished with siphon breaks and/or U-shaped traps

in the line routing.

Antispillage vent valves inside the fuel tank are particularly

advantageous during rollover accidents and can be used in lieu

of flexible lines, breakaway valves, and all other alternate

considerations. These valves must be designed and tested to

demonstrate that:

"* The vent will remain fully open during all normal

flight environmental conditions.

"* The vent valves will close in extreme attitudes such

as would occur in a rollover.

"* The vent valves will possess adequate venting capability under critical icing conditions in flight.

If the fuel system is to be pressure refueled, a bypass system

for tank overpressurization must be used. However, care must

be taken to ensure that spillage resulting from overpressurization due to tank compression during a crash is released away

from aircraft occupants and ignition sources.

184

SI1' II ' 1' II'[1 I1:1i; ii , ; -----
6.3 OIL AND HYDRAULIC SYSTEM DESIGN CRITERIA

6.3.1 General

Even though oil and hydraulic fluids are carried in relatively

small quantities, they are easily ignited and can serve, in

turn, as ready ignition sources for fuel. Therefore, oil and

hydraulic fluid spillage must be prevented at all reasonable

cost. The crashworthy design criteria presented in Section 6.2

for fuel systems also are generally applicable for oil and hydraulic systems.

6.3.2 Oil and Hydraulic Fluid Reservoirs

Oil tanks and hydraulic reservoirs must not be located where

spilled or sprayed fluid can readily be ingested into the engine or ignited by the engine exhaust.

Oil tanks and hydraulic reservoirs must not be located in the

following areas:

* Near the bottom of the fuselage.

* In or above engine compartments.

* In electrical compartments.

* In occupiable areas.

* Under, in front of, or at the side of heavy masses,

such as engines and transmissions, nor above landing

gears.

Reservoir construction and mounting must be able to withstand

30-G forces applied in any direction.

Oil tanks should be constructed from flexible, crash-resistant

materials that meet or exceed the strength and tear resistance

required in MIL-T-27422 for fuel tank material.

Alternatively, a metal tank can be used if it is in a relatively safe area and is shielded and coated to prevent leakage

in the event of a tank rupture.

6.3.3 Oil and Hydraulic Lines

6.3.3.1 Construction: Oil and hydraulic lines should consist

of flexible hoses with steel-braided outer sheaths, where possible. If the hoses cannot elongate 20 percent without the

hose assembly spilling fluid, 20 percent extra length should

185

hi 

~.2~- ___ 

_____

A
be provided to compensate for structural displacement during

a crash. All hose assemblies must meet the requirements of

Table 16 when tested as shown in Figure 54 (Section 6.2.3.1).

Where high-temperature operational requirements preclude the

use of flexible hose, coiled metal tubing should be used in

areas where large crash deformation is expected.

The number of line couplings must be kept to a minimum. Wherever possible, a single, one-piece hose should be routed

through a bulkhead opening rather than attached to the bulkhead with a rigid connection. The opening should be I in.

larger in diameter than the hose diameter, with the hose stabilized by a frangible panel or structure. However, selfsealing breakaway valves must be used wherever a line goes

through a firewall so that the jine will seal if the engine is

displaced during crash impact.

Self-sealing breakaway valves must be used to connect flexible

hoses to engines, oil tanks, hydraulic reservoirs, and system

components, if enough structural deformation to cause line

elongation to the breakage point is probable.

When hydraulic or oil lines must be stabilized, they should be

attached to the aircraft structure with frangible fasteners.

6.3.3.2 Routing: Hydraulic or oil lines must not be routed

in electrical or occupiable areas unless they are shrouded to

prevent spillage. Hydraulic or oil lines should not be routed

in the following areas:

* Near the bottom of the fuselage.

e Over landing gears.

* Under, in front of, or at the sides of heavy masses,

such as engines and transmissions.

* In the leading edges of wings.

"* In areas of anticipated rotor blade impact.

"* In any area where flammable fluids could be spilled

or sprayed onto hot surfaces or ingested into the

engine.

"* Above electrical wiring.

186
" ~I

The number of hydraulic and oil lines in the engine compartment must be kept to a minimum. The lines should be grouped

together and enter the engine compartment in a protected loca- 4

tion.

6.3.4 Oil and Hydraulic System Components

System components (e.g., pumps, valves, filters, actuators)

must not be located in electrical compartments or occupiable

areas. Components should not be located near the bottom of

the fuselage or in the leading edges of the wings.

Components located in the engine compartment should be restricted to those absolutely necessary for engine operation.

For example, oil filters must not be located there unless they

are an integral part of the engine.

The construction and mounting of all system components must be

able to withstand 30-G forces applied in any direction.

6.3.5 Oil Coolers

Oil coolers must not be located in the engine compartment and

should not be located under the engine or transmission, or in

any area where oil could be spilled or sprayed onto hot surfaces, or ingested into the engine.

The oil cooler should be located as far as possible from anticipated impact areas.

The oil cooler mounting(s) must be able to withstand 30-G

forces applied in any direction.

6.4 IGNITION SOURCE CONTROL CRITERIA

6.4.1 Electrical Systems

6.4.1.1 Wiring: Electrical wires should be routed along

heavier structural members of the airframe wherever possible.

Structural openings for wire passage should be 8 to 12 times

larger in diameter than the wire. Sharp metal edges must be

protected by grommets to prevent chafing. Wire bundles must

be supported at frequent intervals along their lenqth by frangible attachments to the aircraft structure.

Wires that must pass through areas of anticipated structural

deformation should be approximately 20 to 30 percent longer

than necessary. The extra length should be accumulated in the

torm of loops or S-shaped patterns and located at the areas of

anticipated structural deformation.

187

" •ti |_ a I i _ _ 
Wires should be routed above or away from flammable fluid

lines, and they should never be closely spaced between cuter

skin and fuel lines. Wires must not be routed near flammable

fluid tanks unless the wires are shrouded to prevent arcing.

Wires should not be routed in the following areas:

"* Near the bottom of the fuselage.

"* Over landing gears.

"* In the leading edges of wings.

"* In areas of anticipated rotor blade impacts.

"* In areas of anticipated fuel spillage.

"* Immediately adjacent to flammable fluid lines and

vent openings.

Electrical wiring and components must be kept to a minimum in

flammable fluid tank areas.

Nonsparking breakaway connectors should be used in areas where

excessive tensile loads may be applied, such as the wing-to- fuselage joint. All wire connectors must be of the shielded,

nonsparking type.

6.4.1.2 Batteries and Electrical Accessories: Batteries and

electrical accessories should be located as far as possible

from flammable fluid tanks.

Batteries and accessories should be housed in compartments

built into the airframe. These compartments should be lined

with flexible, nonconductive, fire-resistant panels as specified in Section 6.4.1.5.

Electrical wires must exit the batteries and inverters on their

least vulnerable side. There must be one full 6-in.-diameter

loop of extra wire at the battery and inverter connections to

accommodate crash-induced structural deformation.

The battery and accessory mountings must withstand a force of

30 G applied in any direction.

6.4.1.3 Generators and Magnetos: If generators and magnetos

are not engine mounted, they should be installed in compartments built into the airframe. These compartments should be

located fairly high in the structure and as far as pcosible

from flammable fluids. The compartments should be lined with

panels as specified in Section 6.4.1.5.

188
Electrical wires must exit the generators and magnetos on their

least vulnerable side regardless of their location. The generator and magneto mountings must withstand a force of 30 G applied in any direction.

6.4.1.4 Lights and Antennas: Lights and antennas should be

located as far as possible from flammable fluids. Lights

should be located as high as possible on the airframe structure. Landing lights must not be located in front of wing

fuel tanks.

The wires that attach to the lights should contain a 6-in.-

diameter loop near the connection to accommodate crash-induced

structural deformation.

6.4.1.5 Liners and Snrouds: Nonconductive paneling must be

used as a liner for all electrical compartments. The paneling

materials must possess a minimum tensile strength of 250 lb/in.

of width and allow a minimum elongation of 200 percent.

Nonconductive material must be used to shroud all electrical

wiring that could be cut by deforming aircraft structure during

crash impact. The shrouding material must meet or exceed a

tensile load of 250 lb/in. of width and must possess a minimum

elongation capability of 200 percent.

6.4.2 Shielding

Shielding should be used wherever necessary to prevent spilled

flammable fluids from reaching potential ignition sources or

occupiable areas.

6.4.2.1 Spillage Barriers: Fuel tanks must be isolated from

the occupants by a minimum of two spillage barriers. These

barriers may consist of the normal tank cavity chafing liner

and the surrounding airframe structure. If the chafing liner

is considered as a barrier, it must be continuous structure

completely encasing the fuel tank.

6.4.2.2 Firewalls: Firewalls must be designed to withstand

all survivable crash impacts without losing their structural

integrity or sealing ability.

6.4.2.3 Fire Curtains: Fire curtains made from fire-resistant

cloth may be used to protect occupiable areas or ignition

sources from flammable fluid spillage. Fire curtains may be

installed in addition to but not in place of the spillage barriers required in Section 6.4.2.1.

18

189

, - * I I I I I I~ I I I I . I *
4

1•

6.4.2.4 Flow Diverters: Drainage holes must be located in

all flammable fluid tank compartments to prevent the accumulation of spilled flammable fluids within the aircraft. Drip

fences and/or drainage troughs should be used to prevent the

gravity flow of spilled fuels from reaching ignition sources

such as hot engine areas or electrical compartments.

6.5 INTERIOR MATERIALS SELECTION CRITERIA

6.5.1 General

All aircraft interior materials such as seat fabrics and cushions, interior wall insulations, and nonmetallic structural

components must be flame resistant and produce the least amount

of smoke and toxic gases possible. Interior materials in all

U. S. Army aircraft must meet the flammability criteria specified in Federal Air Regulation (FAR) 25.853 (Reference 54);

these requirements are summarized in Section 6.5.2. Passengercarrying aircraft should meet the flammability and smoke emis--

sion criteria guidelines issued by the Urban Mass Transportation Administration (UM¶A) (Reference 55); these criteria are

surtmarized in Section 6.5.3.

6.5.2 FAR 25.853 Flammability Requirements

Materials used in each compartment occupied by the crew or passengers must meet the following requirements:

* Ceiling panels, wall panels, partitions, structural

flooring, etc. Must be self-extinguishing when

tested vertically by applying a 1550OF flame to the

lower edge of the specimen for 60 sec. Average burn

length not to exceed 6 in.; average flame time after

removal of test flame not to exceed 15 sec. Drippings may not continue to flame more than an average

of 3 sec.

54. U. S. Code of Federal Regulations, Title 14, Chapter 1,

Part 25, Section 853: COMPARTMENT INTERIORS, Government

Printing Office, Washington, D. C., (Rev.) 1980.

55. Transportation Systems Center, PROPOSED GUIDELINES FOR

FLAMMABILITY AND SMOKE EMISSIONS SPECIFICATIONS, (Unofficial) U. S. Department of Transportation, Cambridge, Massachusetts.

190
I

* Floor coverings, textiles (including upholstery),

seat cushions, paddings, insulations (except electrical insulation), etc. Must be self-extinguishing

when tested vertically by applying a 1550*F flame to

the lower edge of the specimen for 12 sec. Average

burn length not to exceed 8 in., average flame time

after removal of test flame not to exceed 15 sec.

Drippings may not continue to flame more than an

average of 5 sec.

* Acrylic windows, signs, restraint systems, etc. May

not have an average burn rate greater than 2.5 in./ I

min when tested horizontally by applying a 1550OF

flame to the specimen edge for 15 sec.

See Reference 54 for the complete text of the regulations and

test requirements. I

6.5.3 UMTA Flammability and Smoke Emission Guidelines

Combustible materials used in transit systems are required to

possess the following flammability characteristics:

* Seat cushions and insulations (except electrical

insulation). Must pass ASTM E 162-76 (Reference 56)

Radiant Panel Test with a flame propagation index

(I ) not exceeding 25, with the added provision that

there shall be no flaming, running, or dripping.

o Wall and ceiling panels, seat frames, partitions,

etc. Must pass ASTM E 162-76 Radiant Panel Test with

a flame propagation index (I ) not exceeding 35, with

the added provision that thire shall be no flaming dripping.I

* U holstery Materials. Burn length must not exceed

6 when tested-y FAR 25.853 vertical test. Average flame time after removal of flame source may not

exceed 10 seconds. Flaming dripping not allowed.

* Carpeting (tested with its padding). Must pass NBS

flooring Radiant Panel Test, NBSlR-74-495 iith a minimum critical radiant flux of 0.6 watts/cm-.

56. ASTM E 162-76, STANDARD TEST METHODS FOR SURFACE FLAMMABILITY OF MATERIALS USING A RADIANT HEAT ENERGY SOURCE,

1977 Annual Book of ASTM Standards, R. P. Lukens, et al.,

eds., American Society for Testing and Materials, Easton,

Maryland, 1977, Part 18.

191

U

- ... . . ..-
* Plastic windows. Must pass ASTM E 162-76 Radiant

Panel Test with a flame propagation index (I ) not

exceeding 100. s

* Flooring. Must withstand requirements of ASTM E

119-76 (Reference 57) when underside is exposed to a

flame up to 1400*F for 15 min.

* Elastomers. Must pass the requirements of ASTM

C542-76 (Reference 58), with the added requirement

of no flaming dripping.

When tested in accordance with the National Fire Protection Association Standard No. 258-1976 (Reference 59) in both flaming

and nonflaming modes, combustible materials should meet the

following smoke emission requirements:

"* Upholstery, air ducting, insulation (except electrical insulation). Optical density (D ) must not exceed 100 within 4 min after start of %est.

"* All other materials, (except foam seat cushioning,

electrical insulation, and carpeting). Optical density (D ) must not exceed 100 within 90 sec after

start of test, nor exceed 200 within 4 min after

start of test.

The UMTA guidelines and the NFPA standard are being voluntarily

used by several transit authorities and manufacturers although

the guidelines are not Government standards and have no official status. See References 55 and 59 for the complete text of

the regulations and test requirements.

If fire-retardant coatings are used for fabric and trim materials, the effects, if any, of routine maintenance and cleaning

procedures must be assessed. If the coatings can be removed

57. ASTM E 119-76, STANDARD METHODS OF FIRE TESTS OF BUILDING

CONSTRUCTION AND MATERIALS, 1977 Annual Book of ASTM Standards, R. P. Lukens, et al., eds., American Society for

Testing and Material3, Easton, Maryland, 1977, Part 18.

58. ASTM C 542-76, STANDARD SPECIFICATION FOR LOCK-STRIP GASKETS, 1977 Annual Book of ASTM Standards, R. P. Lukens,

et al., eds., American Society for Testing and Materials,

Easton, Maryland, 1977, Part 18.

59. NFPA 258-1976, STANDARD TEST METHOD FOR MEASURING THE

SMOKE GENERATED BY SOLID MATERIALS, National Fire Codes,

1979, National Fire Protection Association, Boston, Massachusetts, 1979.

1I

192
by routine cleaning procedures, the flammability and smoke/

toxic fume tests should be repeated after a representative number of cleaning cycles.

6.6 DITCHING CRITERIA

6.6.1 General

Occupant survival during a ditching is highly dependent on

egressing rapidly from the aircraft before it sinks. This is

especially true in helicopters, which tend to roll inverted

and sink very rapidly. Disorientation and poor underwater visibility further hamper successful egress. Available escape

times from helicopters range from a few seconds to a few minutes. The availability of emergency exits, adequate emergency

exit lighting, and helicopter flotation provisions can all increase the available escape time. Adequate and easily deployed

ditching equipment increases the probability of survival after

successful egress.

6.6.2 Emergency Exits

All U. S. Army aircraft must meet the criteria for emergency

exits contained in Section 6.7. Passenger-carrying helicopters operating over water environments, however, should contain more and larger emergency exits than might normally be

provided. Additional escape exits should be provided in the

overhead, deck, and tail sections.

Explosively created exit systems should be considered because

of their rapid initiation times and immunity to the crash environment. Linear-shaped charges should be placed around and

extend beyond existing windows and hatches to preclude the

problem of jammed or stuck exits. Strategically placed shaped

charges in the overhead, deck, empty bulkhead spaces, etc. can

provide the additional emergency exits required in the ditching

environment. Criteria for these types of systems are contained

in Section 6.7.

6.6.3 Underwater Emergency Light4ng

Emergency exits must be lighted with high intensity lights if

they are to be seen underwater. The required brightness of

the lights depends on the turbidity of the water, the distance

between the observer and the light, and the threshold sensitivity of the observer's eyes.

The escape hatch lights must have a minimum brightness of

120 fL. However, higher brightness levels of light, if possible, should be employed for underwater escape lighting.

193
6.6.4 Helicopter Flotation Systems

An adequate number of helicopter flotation devices should be

provided. Combinations of flotation w1ethods, such as sponsons

in conjunction with flotation bags, sealed hulls, etc., should be used.

Sponsons can help stabilize a helicopter in relatively calm

seas. However, they must be quite large to be of any value in

providing flotation to counteract the inherent instability due

to a helicopter's high center of gravity. Calculated aircraft

stability must be verified by data from tests performed on the

aircraft or on a scaled model thereof.

The calculated stability afforded by flotation bags also must

be verified by test data. To achieve maximum effectiveness,

the bags must inflate simultaneously prior to or upon water

contact at slow speeds. Reliability of a flotation bag system

is of prime importance.

6.6.5 Ditching Equipment

Tiedown or stowage locations must be provided for life rafts,

life preservers, survival kits, and miscellaneous ditching

equipment. Restraint devices and supporting structures must

be designed to restrain the equipment to static loads of 50 G

downward, 10 G upward, 35 G forward, 15 G aftward, and 25 G

sideward. All survival equipment must be readily available

and easily released from restraining devices after ditching.

Life raft mountings and restraining devices must be located

and designed so that rafts can be removed and deployed outside the aircraft within 30 sec from the time the release or removal action is initiated.

When exterior installations for life rafts or other survival

equipment are provided, the mountings and restraining devices

must be designed to prevent inadvertent release or damage in

flight or when ditching. Such equipment must be recoverable

from an exit intended for use in ditching. Release mechanisms

must be designed to minimize the possibility of jamming due to

structural deformation incurred during ditching.

6.7 EMERGENCY ESCAPE DESIGN CRITERIA

6.7.1 Emergency Exits

6.7.1.1 General: Exits of sufficient size and number must be

provided to ensure that all occupants can evdcuate the aircraft

before postcrash conditions become intolerable, even if half

194
I

of the exits are blocked. If a crash-resistant fuel system is

not installed, the maximum number of personnel to be carried

must be able to evacuate the aircraft within 10 sec. The allowable evacuation time can be extended to 30 sec if a crashresistant fuel system is installed in the aircraft. The emergency exit criteria presented in this chapter are predicated

on a 30-sec evacuation time.

6.7.1.2 Types of Exits: A Class C exit constitutes the minimum requirement for an emergency exit. (A Class C exit is a

window, door, hatch, or other exit intended primarily for emergency evacuation). Class C exit closures must be capable of

being removed from the exit opening within 5 sec regardless of

the aircraft's attitude.

A Class B exit consists of a door, hatch, or other exit intended primarily for service or logistic purposes (e.g., cargo

hatches and rear loading ramps or clamshell doors). Class B

exits may be used instead of Class C exits if adequate emergency releases are installed. A Class A exit (doors, hatches,

etc., intended primarily for normal entry and exit) generally

may be used in lieu of a Class C exit; however, if either

Class B or Class A openings are used in place of Class C exits,

they must meet the 5-sec opening requirement.

6.7.1.3 Size of Exits: All exits must be sufficient in size

and shape to allow 95th-percentile combat-equipped troops and

aviators to pass through the exit at a rate of 1.5 sec per man

or less. Therefore, Class C exits must be a minimum of 22 in.

in diameter, or 22 in. square, with 6-in. radius corners, although larger exits are recommended. Other shapes may be used

if the minimum dimensions are met or exceeded.

6.7.1.4 Number of Exits: Each flight crew member must have

access to at least one usable emergency exit regardless of the

attitude of the aircraft after impact. When sliding or clamshell canopies are used, Class C exits must be provided for

crew escape in case the postimpact attitude of the aircraft

prevents jettisoning of the canopy.

A minimum of two Class C exits (or equivalent,' must be provided

in troop/passenger sections, one on each side of the fuselage.

Cockpit exits may not be counted toward this requirement. Additional exits must be provided whenever the ratio of seats to

passengers exceeds the 1-to-10 ratio (e.g., if the capacity is

21, three exits are required). These requirements also apply to

cargo compartments if the compartments have a capability for

troop transport.

195

T--- --• " - " - I --.... :I m I • •! m =•l m '
6.7.1.5 Location of Exitsz Emergency exits must be equally

divided between both sides of the aircraft to provide alternate

means of escape if, for any reason, the exits on one side become blocked. If feasible, in order to prevent crowding during

evacuation, side exits should not be located directly across

from each other. At least one exit on each side must be well

above the anticipated waterline during a ditching.

If the width of the fuselage between sidt exits is 5 ft or

more, at least one additional Class C exit must be provided

overhead so that easy access to an exit is available when the

aircraft comes to rest on its side. If more than 20 occupants

can occupy the troop/passenger section, o',e overhead exit must

be provided for every 20 occupants. If overhead exits are not

feasible, bottom or fore and/or aft exits may be provided instead. Alternatively, side exits may be located where interior

aircraft structures or components can be used as steps to gain

access to the upside exits. Such component-steps must be able

to support at least 300 lb. They must also maintain their

structural integrity and attachment to the aircraft when exposed to static loads of 50 G downward, 10 G upward, 35 G forward, 15 G aftward, and 25 G sideward.

Emergency exits should not be located in the following areas:

e In close proximity to the main landing gear.

o Under heavy components, such as engines and transmissions.

9 In any area where it is necessary to move equipment,

cargo, etc., to gain access to the exit.

* In any area where external components, such as engines or armament, will interfere with occupant escape.

* Near potential fuel spillage areas.

* Near major ignition sources, such as hot engines.

6.7.1.6 Operation of Exits: The method of releasing and opening an emergency exit must be simple, obvious, and natural to

all personnel carried in the aircraft. All emergency exits

must be capable of being completely opened within 5 sec after

the person initiating the action first places his hand on the

release handle.

196
Exit release mechanisms must permit release handle actuation

and exit opening by one person using one hand. The releasing

action must be natural to the position of the operator initiating the action and must be a continuous motion from start to

finish without sharp changes in direction. Secondary operations must not be necessary. The final motion of the release

handle should contribute to the opening of the exit.

Release handles must be located on the exit closures them- S

selves, or immediately adjacent to the exit openings, so that

they are readily ac'.essible. However, the handles must not obstruct the removal of the exit closure or impede escape through

the exit opening. Release handles in cockpits and troop compartments must be located so that crew members need not unlock

their shoulder harnesses in order to actuate the release me- chanism.

Accidental release of exits in flight must be prevented. Release mechanisms must be designed so that improper or incomplete closing of the exit closure will be obvious. Easily removable protective covers may be used to prevent inadvertent 1

actuation of exit release handles.

It is essential that all emergency exits be capable of being

opened by rescue personnel from outside the aircraft. Internal

and external release mechanisms must be capable of being actuated simultaneously without interfering with each other. Means

to prevent icing of the outside release mechanisms and handle

mounts must be provided.

Once the release mechanism has been actuated, only the single

operation of pulling or pushing the exit closure into the clear

should be necessary. All emergency exit closures must be designed to fall free or be easily pushed outward if the aircraft " not pressurized. In pressurized aircraft, exit closures must be removed inwardly, but, if possible, should then

be canted at an angle and pushed out the exit opening. "Push

out"-type Class C exits also must be capable of being pushed

in from the outside by rescue personnel.

Emergency exits must be designed to permit removal of the exit

closure in spite of seal vulcanization, ice accumulation, and

moderate fuselage deformation. A peripheral clearance of at

least 0.20 in., provided between the exit closure and its

frame, will help accomplish this goal.

6.7.1.7 Explosively Created Exits: Explosive systems for cutting emergency exits through existing doors and windows and

through fuselage structures should be considered. These systems provide the advantages of extremely rapid release times,

197
simplicity of operation, and immunity to jamming. If an explosive exit system is incorporated into the aircraft, the following design criteria apply.

The arming/firing system must be designed for simple and rapid

actuation of the explosive system, yet provide maximum safety

against inadvertent actuation. Arming and firing must be accomplished in two separate and deliberate actions, with the

arming function always under the control of the flight crew.

The safe/arm mechanism must remain in its chosen position

(armed or disarmed) until a deliberate action to change its

position is initiated. The safe/arm mechanism must not change

positions due to system failure, or due to any environmental

or crash inputs. Disarming capability must be provided to permit safing the system when normal safing modes are inoperable.

The firing mechanism must be independent of any external energy

source. Firing mechanisms should be located adjacent to each

emergency exit so that each exit can be opened independently,

from both inside and outside the aircraft.

The linear shaped charges used to cut the exit openings must

be held securely in position against the aircraft structure.

The size of the exit openings must conform to Class C requirements. The jettisonable section must be ejected outward.

Energy-absorbing backup material must be placed behind the

shaped charge to control the backblast of the explosive.

All explosives used in the system should possess as high a

thermal limit as possible. The system must be able to function when exposed to ambient air temperatures up to 400*F, yet

not function during brief exposure (30 to 60 sec) to postcrash

fires. The system must be designed to minimize the possibility of system actuation igniting any spilled fuel. Thus, the

amount and duration of any exposed flame must be minimal.

6.7.1.8 Access to Exits: Access from aisles to all exits

must be provided so that exits are not obstructed by any aircraft structures or components that would impede escape. The

width of aisles at any point between seat rows must allow unobstructed movement of 95th-percentile troops with full combat

equipment. Therefore, the aisle width must be at least 17 in.

Where it is necessary to pass through seat rows to gain access

to emergency exits, the longitudinal spacing between the rows

must be sufficient to permit these troops to move at a rate

consistent with the capacity of the exit (1.5 sec per man or

less).

198

_ _ _

*- - ~ ___
6.7.2 Emergency Lighting

6.7.2.1 Interior Emergency Lighting: Interior emergency

lighting must provide sufficient illumination throughout cockpit and cabin areas to permit occupants to locate emergency

exits and survival equipment, perceive escape paths, and avoid

obstacles while moving toward the exits. Minimum average illumination in clear air along passageways leading to each exit

and in front of each exit must be 0.05 fc measured 20 in. above

the floor (excluding canopy aircraft).

6.7.2.2 Emergency Exit Lights: Supplementary emergency lighting units, with adequate brightness to permit occupants to

identify exits, read exit operating instructions, and actuate

exit release mechanisms during reduced visibility conditions

(darkness, smoke, etc.), should be provided at or near each

emergency exit. All passenger/troop-carrying aircraft must

contain internally illuminated exit signs with a minimum

brightness of at least 25 fL, although brighter lights are

strongly recommended. Aircraft whose mission requirements

include troop transport over water should contain exit sign

lighting meeting the requirements specified in Section 6.6.3.

Canopy aircraft may be excluded from these requirements.

6.7.2.3 Exterior Emergency Lighting: For noncombat missions,

exterior emergency lighting should be considered to illuminate

the ground near each exit and in areas where escape and survival equipment will be deployed. The light intensity on the

ground should be 0.02 fc minimum.

6.7.2.4 Structural Requirements: All emergency lighting units

must be self-contained, explosion-proof, operable under water,

and accessible for periodic maintenance. To ensure structural

integrity and continued operation after a crash, the lighting

system must be capable of withstanding the following crash

loads: 50 G downward, 10 G upward, 35 G forward, 15 G aftward,

25 G lateral. The crash environment is more fully defined by the velocity changes presented in Table 2. Except for those

lights directly destroyed by the crash, breakup of the fuselage

must not render any portion of the lighting system inoperative.

6.7.2.5 Power Sources: All units must be capable of operating

independently of the main aircraft lighting system. Emergency lighting power sources must be independent of the main power

source of the aircraft. They must contain power sufficient to

provide effective illumination for a minimum of 15 min.

6.7.2.6 Actuation of Lighting Units: Emergency lighting units

should be actuated automatically in as many survivable accidents as possible. This can be accomplished by using inertia

199 I IC
sensors capable of sensing lower-severity accidents. Sensor

criteria should be identical to those specified for crash locator beacons in Section 6.8. An override switch to nullify

the automatic feature when desired must be provided. Manual

actuating switches must be provided so that emergency lights can be turned on prior to a crash if desirable.

6.7.3 Emergency Exit Markings

Emergency exits must be clearly marked both inside and outside the aircraft. In addition, instructions for releasing

the exits must be clearly marked beside the exit release mechanisms.

All U. S. Army aircraft must be painted and marked according

to the requirements of TB 746-93-2 (Reference 60). Although

these requirements are summarized in Volume V of this guide,

the reader is referred to TB 746-93-2 for complete details.

6.7.4 Crew Chief Stations

At least one crew chief station must be located in each troop

compartment. The station should be located as near the main

or emergency exits as possible and should provide complete surveillance of the troop compartment.

6.7.5 Alarm Systems

Aircraft with passenger or troop compartments should be equip- I

ped with an audible emergency alarm device that can be heard

over the highest decibel noise level expected in the aircraft.

Consideration should be given to providing visual as well as

audible warnings.

6.8 CRASH LOCATOR BEACON DESIGN CRITERIA

6.8.1 General

Crash locator beacons may be fixed, portable, or deployable, a

as specified by the procuring activity according to its aircraft mission requirements.

Fixed equipment is permanently mounted in the aircraft. Al- I though the transmitter, antenna, and power supply need not be

contained in one package, their close proximity to each other

will reduce the chances of connecting circuitry being damaged during crash impact.

60. Technical Bulletin, TB 746-93-2, PAINTING AND MARKING OF

ARMY AIRCRAFT, Department of the Army, Washington, D. C.,

10 August 1978.

200
Portable and automatically deployed beacons must contain the

transmitter, antenna, and power supply in one package. Portable beacons must be easily removed from their installations

by crew members, yet their installations must be secure enough

to protect them from impact damage.

Automatically deployed beacons must be designed to withstand

ground impact forces following their ejection. They must also

be buoyant, self-righting, and stable when floating in water,

and not adversely affected by immersion in fresh or salt water

for the life of the power supply.

Crash locator beacons may be either manually or automatically

activated. Since automatic activation requires no previous

action on the part of the crew, it is the preferred method.

However, an arming switch must be provided so that automatic

activation can be used or not, depending on the aircraft mission. A manual activation switch also must be provided so

that the beacon can be activated if the arming switch is not

on, or if, for any other reason, the beacon is not automatically activated.

6.8.2 Crash Sensors

Although different types of crash sensors might be used, the

current state of the art is such that inertia sensors are the

preferred choice. Regardless of the type of sensor used, the

sensor must be responsive to the majority of survivable aircraft accidents, including those accidents in which the crash

forces and damage are minimal. At the same time, the sensor

must ignore normal vibrational loads and flight loads up to

the limits of maneuverability.

In order to sense 75 to 80 percent of rotary- and light fixedwing accidents, an inertia sensor must have a sensing threshold

of 2 G. Although the 2-G threshold level is below the accelerations sometimes experienced during flight, the inertia sensor

can be designed to filter out vibration and flight loads if it

also must detect a velocity change typical of crash rather than

operational conditions before it actuates.

Since most fixed-wing aircraft accidents have a major longitudinal component of velocity and force, a unidirectional inertia sensor mounted with the active axis forward in the direction of the longitudinal axis of the aircraft is sufficient.

A longitudinal inertia sensor should be designed to actuate at

a threshold of 2-G acceleration and a minimum velocity change

of 3 ft/sec. These specification limits are shown in Figure 58.

201
10

tV 3. 0+0.3 ft/sec

00.

0

,-1

U

Sensor must activate

AU

4

0

• t h = 2 . 0 + . 2

00.

:Sensor must not ... .... ... ... ..• .... i~ i !i i : : •!:!: • : ... .. .... .. ....- ., , . -

0 20 40 60 80 100

Pulse duration, msec

Figure 58. Proposed specification for longitudinal crash

force sensors in rotary- and light fixed-wing

aircraft.

The above specifications are also satisfactory for rotary-wing

aircraft in the longitudinal direction. However, since helicopters often have large vertical crash forces with minimal

longitudinal forces, a vertically oriented crash sensor must

be employed in addition to a longitudinal sensor. The vertical sensor should be designed to actuate at a 2-G acceleration

level when the velocity change is 10 ft/sec or more. The specification limits for this case are shown in Figure 59.

202

. S" .. ,
40 .% 10.0 + 1.0 f t/sec

- .

ke

o

S30

* a

Sensor must - activate

20 ;Y

>

C.)

S10

*Sensor must+0.

not activate Gt . .

0

0 20 40 60 So 100 120 140 160

Pulse duration, msec

Figure 59. Proposed specification for vertical crash

force sensors in rotary-wing aircraft.

The sensor must be able to withstand impact forces associated

with severe survivable crashes and still function. Thus the

sensor must withstand shock pulses equal to or greater than

those listed in Section 6.7.2.4.

The inertia sensor criteria presented above are based on crash

forces typical of those experienced in the occupant compartment

during survivable crashes. Therefore, the sensor must be located in an area that will experience crash forces representa- tive of those in the occupant compartment. The sensor must,

of course, be protected from possible impact damage.

The sensor must be mounted to rigid structure to prevent the

amplification or attenuation of flight or crash loads that

can occur with flexible structures. For the same reason, soft

mounting materials, such as flexible straps or Velcro fasteners, must not be used.

203

. .- .- •.....*.-..-. ---. . -- ... -. . . ... . .. . .. .
II

6.8.3 Transmitters

Operating frequencies and transmitter ranges (power) must be

determined by the procuring activity according to its own

needs.

The transmitter must be capable of being either manually or

automatically activated. An arm switch must be provided so

that automatic activation can be selected or not, as desired.

A cockpit warning light must be provided to alert the crew to

inadvertent transmitter activation.

The transmitter must be located in an area that is not subject

to impact damage. The transmitter and its mounting must be

designed to withstand the impact forces of a severe survivable

accident without compromising the operation of the transmitter.

6.8.4 Antennas

The antennas, except for those used in portable and deployable

beacons, are usually mounted outside the aircraft. The antennas must be located away from anticipated impact areas, such

as the front or bottom of the aircraft, wing or tail surfaces

likely to impact trees, etc., and those portions of helicopters apt to experience rotor blade strikes during impact. The

antenna mounting must be able to withstand the decelerative

forces of severe survivable impacts.

6.8.5 Power Supplies

The crash locator beacon must have its own independent power

supply so that it is not dependent on aircraft power for its

operation. The power supply must be capable of providing the

necessary power for optimum transmitter operation over the time

peciod and under the environmental con4itions specified for the

particular aircraft.

If the power supply is not integral with the transmitter, it

should be mounzed to the aircraft in a location away from anticipated impact areas and must have an attachment strength

equal to that of the transmitter.

All electrical wiring between components of the system must be

protected from impact damage unless the components are packaged

together. Protection can be accomplished by following the criteria in Section 6.4.1.1.

204

L _ _ _---
Yes No N/A

6.9 DESIGN CHECKLISTS

6.9.1 Fuel System Design Checklist tI

6.9.1.1 Fuel Tanks

1. Are the fuel tanks located as far as possible

from anticipated impact areas, occupiable

areas, large weight masses, and primary ig- I

nition sources?

2. Are the fuel tanks located as high up in the

structure as possible?

3. Are the fuel tanks located where there is no

danger of puncture by a collapsing landing

gear?

4. Are the fuel tanks located so that transmissions, engines, and similar massive components will not crush the tanks during a

crash?

5. Are the fuel tanks relatively safe from

penetrative damage by structural stringers

and stiffeners?

6. Can each fuel tank displace in the airframe

structure without tearing or inducing leaks

around the filler area, the fuel line entry

and exit, the quantity indicator, and the

tank-to-structure attachment points?

7. Do the fuel tanks have smooth, regular

shapes, with the sump gradually contoured

into the tank bottom?

8. Do all fuel tank concave corners have a minjnimum radius of 3 in., and all convex corners

a minimum radius of 1 in.?

9. Do all fuel tanks meet or exceed the requirements of MIL-T-27422?

10. Do all fuel tank fittings meet or exceed

the tank pullout strength specified in

MIL-T-274227

205
Yes No N/A

6.9.1.2 Fuel Lines

11. Are all fuel lines made from flexible hose

with a steel-braided outer sheath?

12. Do all hose assemblies meet the strength requirements listed in Table 16, Section

6.2.3.1?

13. Can all hoses elongate 20 percent without

the hose assemblies spilling fuel?

14. Do fuel lines exit the fuel tank in one protected location?

15. Has the number of fuel lines in the engine

compartment been kept to a minimum?

16. Are fuel lines routed along heavier structural members wherever possible?

17. Is as much of the fuel line as possible

routed through the fuel tanks?

18. Are fuel lines routed as far as possible

from occupiable areas and electrical compartments?

19. Are fuel lines routed as far as possible

from all electrical equipment and wires?

20. Are fuel lines routed away from areas where

large structural damage is likely during a

crash?

21. Are fuel lines routed away from the exhaust

system and high-temperature heating ducts?

22. Are the fuel system lines designed with as

few fittings as possible?

23. Are tae fuel system lines designed so that

uncut hoses are run through bulkheads rather

than attached to the bulkheads with fittings?

24. Are self-sealing breakaway valves used wherever a fuel line goes through a firewall or

bulkhead or is attached to the bulkhead?

206
Yes No N/A

25. Are lines entering and exiting in-line boost

pumps made of flexible hose that is approxi- nmately 20 percent longer than necessary?

26. If fuel lines are not longer than necessary

for in-line boost pumps, are self-sealing

breakaway valves used in the lines near the

boost pump?

27. Are self-sealing breakaway valves used at

all points in the fuel lines where aircraft

structural deformation could lead to line

failure?

28. Are fuel line supports frangible to ensure

release of the line from the structure

during crash impact?

29. Will the frangible supports meet all operational and service loads of the aircraft?

30. Are all continuous lines running through

bulkheads stabilized by frangible panels?

6.9.1.3 Frangible Attachments

31. Are frangible attachments used at all attachment points between the fuel tanks and

aircraft structure?

32. Do the specified frangible tank attachment

separation loads exceed all operational and

service loads by a satisfactory margin?

33. Are the specified frangible attachment

separation loads between 25 and 50 percent of the loads required to fail the attached

system or components?

34. Will the frangible attachments separate whenever the required loads are applied in

all possible modes likely to occur during

crash impacts?

207
Yes No NIA

6.9.1.4 Self-Sealing Breakaway Valves

35. Are breakaway valves installed in all fuel

tank-to-fuel line connections, tank-to-tank

interconnects, and at other points in the

fuel system where aircraft structural deformation could lead to system failure?

36. Are the shapes of the breakaway valves re- I

maining in the fuel tank basically smooth?

37. Are the breakaway valves recessed into the

tank wall so that the tank half does not

protrude outside the tank wall more than

1/2 in. after valve separation?

38. Do the zpecified breakaway valve separation loads exceed all operational and service

loads of the aircraft?

39. Are the specified breakaway valve separation loads between 25 and 50 percent of the loads

required to fail the attached components or

lines?

40. Are the breakaway valves required to separate whenever the required loads are applied in the modes most likely to occur during crash impacts?

6.9.1.5 Fuel Drains

41. Are all fuel line drain valves sWabilized

where necessary with frangible attachments?

42. Are all structural attachments of fuel tank

drains made with frangible attachments?

.• Are all fuel tank drains recessed into the

tank so that no part of the drain protrudes outside the tank wall?

6.9.1.6 Filler Units

44. Are filler units attached to the aircraft

structure with frangible attachments?

45. Are filler caps recessed into the fuel tank

wall?

208

W74 IL
Yes No N/A

46. Are long filler necks avoided?

47. If filler necks are used, are they made

from frangible materials and designed so

that the filler cap stays with the tank

after filler neck separation? - -

6.9.1.7 Boost Pumps

48. Can an engine-mounted, engine-driven boost

pump be used in the aircraft?

49. If an engine-mounted suction system cannot

be used, can an air-driven boost pump be

used?

50. Do in-line boost pumps have a structural

attachment capable of withstanding a 30-G

load applied in any direction?

51. Are tank-mounted boost pumps fastened to the

structure with frangible attachments?

6.9.1.8 Fuel Filters and Strainers

52. Are fuel filters and strainers mounted outside the engine compartment wherever possible?

53. Do all strainers and filters have a structural attachment capable of withstanding a

30-G load applied in any direction?

54. Do all strainers and filters retain as small

a quantity of fuel as possible?

6.9.1.9 Fuel Valves

55. Has the number of fuel valves been kept to

the minimum required for operation?

56. Are self-sealing breakaway valves used at

all valve-to-fuel line connections where

crash-induced line failure is likely?

57. Are all small in-line valves fastened to the

structure with frangible attachments?

209

* j
Yes No N/A

58. Do large valves have a structural attachment

capable of withstanding 30-G loads in any

direction?

59. Are fuel shutoff valves located outside the

engine compartment, either on the outside

face of the firewall or at the fuel tank

outlets?

6.9.1.10 Fuel Quantity Indicators

60. Can float-type quantity indicators be used

in this fuel system?

61. If probe-type indicators are used, are they

fabricated from material that either is

frangible or possesses as low a flexural

rigidity as possible?

62. Is a slightly rounded shoe incorporated at

the probe bottom end of all probe-type int dicators, or is the probe mounted at an

angle toward the rear of the aircraft?

63. Are frangible attachments used where it is

necessary to stabilize the indicator by

fastening it to the structure?

6.9.1.11 Vent Systems

64. Are high-strength fittings used between the

metal insert in the tank and the vent line?

65. If vent outlets must be supported, are they

supported by frangible attachments to the

structure?

66. Is the vent line made of wire-covered flexible hose?

67. Is the vent line routed so that it cannot

be snagged in displacing structure during

a crash?

68. Is a self-sealing breakaway valve used at

the tank-to-line attachment if there is

danger of the tank being torn free of the

supporting structure?

(

210
r

Yes No N/A

69. Are vent lines routed inside the fuel tank

in such a manner that spillage cannot continue after a rollover accident?

70. If an antispillage vent valve is used inside

the tank in lieu of the above items, will

the valve remain fully open during all normal flight conditions?

71. Will the vent valve close in the extreme attitudes that will occur during a rollover?

72. Will the vent valve possess adequate venting

capability under critical icing conditions

in flight?

73. If the fuel system is to be pressure refueled, is a bypass system provided in case

of tank overpressurization?

74. Is any spillage due to tank overpressurization released away from aircraft occupants

and ignition sources?

6.9.2 Oil and Hydraulic System Design Checklist

6.9.2.1 Oil Tanks and Hydraulic Reservoirs

1. Are the tanks and reservoirs located as far

as possible from anticipated impact areas,

occupiable areas, large weight masses, and

primary ignition sources?

2. Are the tanks and reservoirs located as high

up in the structure as possible?

3. Are the tanks and reservoirs located where

there is no danger of puncture from a

collapsing landing gear?

4. Are the tanks and reservoirs located where

transmissions, engines, and similar massive

components will not crush them during a

crash?

5. Are the tanks and reservoirs relatively safe

from penetrative damage by structural

stringers and stiffeners?

211
Yes No N/A

6. Can the oil tanks displace in the airframe

structure and still not leak around the

filler area, the fluid line entry and exit,

the quantity indicator, and the tank-tostructure attachment points?

7. Are the hydraulic reservoirs constructed

and mounted to withstand 30-G forces applied

in any direction?- -

6.9.2.2 Oil and Hydraulic Lines

8. Are all oil and hydraulic lines made from

flexible hose with a steel-braided outer

sheath wherever possible?

9. Do all hose assemblier, meet the strength

requirements listed in Table 16, Section

6.2.3.17

10. Can all hoses elongate 20 percent without

the hose assemblies spilling fluid?

11. Is coiled metal tubing used in areas where

flexible hose cannot be used, but large

structural deformations are expected?

12. Has the number of fluid lines in the engine

compartment been held to a minimum?

I: 13. Are fluid lines routed along heavier structural members wherever possible?

14. Are fluid lines routed as far as possible

from occupiable areas and electrical compartments?

15. Are fluid lines routed as far as possible from all electrical equipment and wires?

16. Are fluid lines routed away from areas where

large structural damage is likely during a

crash?

17. Are fluid lines routed away from the exhaust

system and high-temperature heating ducts?

212
S.,YesB Neo IL/_A

18. Are the fluid system lines designed with as

few fittings as possible?

19. Are the fluid system lines designed so that

continuous hoses are run through bulkheads

rather than attached to the bulkheads with fittings?

20. Are self-sealing breakaway valves used wherever a fluid line goes through a firewall

or a bulkhead or is attached to the bulkhead? I -

21. Are self-sealing breakaway valves used at

all points in the fluid lines where aircraft

structural deformation could lead to line

failure?

22. Are fluid line supports frangible to ensure

release of the line during crash impact?

23. Are uncut lines running through bulkheads

stabilized by frangible panels? -

6.9.2.3 Oil and Hydraulic System Components

24. Are all oil and hydraulic system components located as far as possible from anticipated impact areas, occupiable areas, and electrical compartments?

25. Are the components located in the engine

compartment restricted to those absolutely

necessary for engine operation?

26. Can the construction and mounting of all

system components withstand 30-G forces

applied in any direction without leakage?

6.9.2.4 Oil Coolers

27. Is the oil cooler located outside of the

engine compartment?

28. Is the oil cooler located as far as possible

from anticipated impact areas, (ccupiable

areas, and other potentially injurious components?

213
Yes No N/A

29. Can the oil cooler and connecting lines ex29. perience considerable deformation without

leaking?

30. Can the oil cooler mounting withstand 30-G

forces applied in any direction?

6.9.3 Ignition Source Control Checklist

6.9.3.1 Electrical Systems

1. Are wires routed as high up in the structure

as possible?

2. Are wires routed away from areas of anticipated structural damage, i.e., landing gear

failure, nose crush-in, etc.?

3. Are wires routed above or away from flammable fluid lines?

4. Are all wires routed through the structure

so that extensive structural collapse or

displacement can take place without breaking

wiring? -

S. Are wire bundles supported at frequent intervals by frangible attachments to the

aircraft structure?

6. Are wires shielded by felt or similar protective covers in areas where crushing is

likely?

7. Are wires to electrically operated boost

pumps 20 to 30 percent longer than neces- sary?

8. Is all electrical wiring going through the

fuel tank compartments shrouded?

9. Is wiring in the fuel tank compartment routed as high as possible in the compart- merit?

10. Are electrical wires in the fuel tank compartment 20 to 30 percent longer than necessary? - -

214
Yes No &_A

11. Are batteries, generators, and inverters

located in areas relatively free from structural collapse?

12. Are batteries, generators, and inverters

located as far as possible from flammable

fluids?

13. Are batteries and generators (unless engine

mounted) housed in compartments built into

the airframe?

14. Are battery, inverter, and generator mountings capable of withstanding a 30-G force

applied in any direction?

15. Are the wires connecting the generator,

battery, and inverter into the system located in relatively crush-free areas?

16. Are light bulbs and attaching wires on lower

airframe surfaces designed to readily displace, rather than remain stati.onary and be

broken?

17. Are all electrical compartments l.ined with

a tough, nonconductive paneling?

6.9.3.2 Shielding

18. Are fuel tanks isolated from the occupants

by a minimum of two spillage barriers?

19. Are firewalls designed to withstand all

survivable crash impacts without losing

their structural integrity or sealing

ability?

20. Are drainage holes located in all flammable

fluid tank compartments?

21. Is the hot metal of the engine shielded from

flammable fluid spillages?

6.9.4 Interior Materials Selection Checklist

1. Do all interior materials meet the flammability requirements specified in Federal Air

Regulation (FAR) 25.853?

215

__ _ _ _ __ _ wilý _ _ __ _ _•_ _ _ _ _

i l-i i i ii i/ - i--•l• II I 'l~ i i . ..-
I

Yes No N/A

2. Do all interior materials produce the lowest

possible amount of toxic gases? --- 5

3. Do all interior materials in troop/passengercarrying aircraft meet the flammability

guidelines recommended by UMTA? (See Section 6.5.3)

4. Do all interior materials in troop/passengercarrying aircraft meet the smoke emission

guidelines recommended by UMTA? (See Section 6.5.3)

6.9.5 Ditching Provisions Checklist

1. Are emergency exits larger and more numerous

than normally required to meet minimum

standards?

2. Are additional escape exits provided in

the overhead, deck, and tail sections?

3. Have explosively created exit systems been

considered?

4. Are emergency exits lighted with high intensity lights with a minimum brightness of

120 fL?

5. Even though escape lights meet the minimum

requirement, is the brightness level of

escape lighting the highest permitted by

other design conditions?

6. Has more than one aircraft flotation method

been provided?

7. Does the flotation bag system have a high

reliability?

8. Are tiedown or stowage facilities provided

for life rafts and other ditching equipment?

9. Are equipment restraint devices and support- ing structures designed to restrain the

equipment to loads of 50 G downward, 10 G

upward, 35 G forward, 15 G aftwird, and 25

sideward?

216

", i

----.. - • I ISk
Yes No N/A

10. Is all survival equipment readily available

and easily released after ditching?

11. Can life rafts be removed and deployed outside the aircraft within 30 sec?

6.9.6 Emergency Escape Design Checklist

6.9.6.1 Emergency Exits

1. Are the numbers, sizes, and locations of the

exits such that a full load of troops and

crew can evacuate in 30 sec when the aircraft is on its side?

2. Are all escape exits a minimum of 22 in. in

diameter, or 22 in. square with 6-in. radius

corners?

3. Can all c gency exits be completely opened

within 5 a,. after the person initiating the

action first places his hand on the release

handle?

4. Does each crew member have access to at

least one emergency exit regardless of aircraft attitude?

5. Are a minimum of two exits, onr: on each side

of the fuselage, provided in troop/passenger

compartments?

6. Is at least one exit provided for every 10

persons expected to occupy troop/passenger

compartments?

7. Are emergency exit locations equally divided

on each side of the aircraft?

8. If the width of the fuselage is 5 ft or

more, are additional exits provided in the

overhead, bottom, fore or aft sections of

the aircraft?

9. Are all exit release mechanisms of the

single motion type?

10. Is the number of different types of exit release handles held to a minimum?

217
Yes No N/A

11. Can all exits be opened from both the inside

and outside of the aircraft? 4

12. Can the exits be opened even if the fuselage

evidences considerable distortion?

13. Can the exits be easily operated when the

aircraft is on its side?

14. Will removed or opened exit covers inherently be positioned so as to not block the

exit openings nor interfere with occupant

egress?

15. Is the exit opening operation designed to

inherently resist jamming by loose objects?

16. Can an exit be opened easily when the operator is being pushed or crowded by other

occupants?

17. During emergency evacuation, do all passengers have essentially the same distance to

move during egress?

18. Are aisles between seat rows wide enough to

allow unobstructed movement of occupants

(at least 17 in. minimum)?

19. If occupants must pass through seat rows to

reach the exits, can they move to the exits

at a rate that permits one person to exit

every 1.5 sec or less?

6.9.6.2 Explosive Exit Systems

20. Are arming and firing accomplished in two

separate and deliberate actions?

21. Is the arming function under the control of

the flight crew?

22. Will the safe/arm mechanism remain in its

preselected position regardless of system

failure or environmental or crash inputs?

23. Is the firing mechanism independent of any

external energy source?

218

__. . _ ... ____I_ . ___ . ....
Yes No N/A

24. Can the exits be opened independently of

each other?

25. Are the explosive charges used to cut the

openings held securely in position against

ths aircraft structure?

b d e i a 26. Are energy-absorbing backup materials placed behind the explosive charges?

27. Can the system function in ambient air temperature up to 4000F, yet not function during

30- to 60-sec exposures to postcrash fires?

28. Are the amount and duration of any exposed

flames from explosive actuation minimal?

6.9.6.3 Emergency Lighting

29. Does the interior emergency lighting provide

sufficient illumination to permit occupants

to locate emergency exits, survival equip- i

ment, and escape paths? - -

30. Is there an average illumination in clear

air of 0.05 fc or greater, measured 20 in.

above the floor along passageways leading to

exits?

1;31. Are supplementary lighting units located at I: or near each emergency exit?

32. Do all internally illuminated exit signs

• i' •have a minimum brightness of at least 25 fL?

33. For noncombat missions, is exterior emergency lighting provided to illuminate the

ground near each exit and the areas where

escape and survival equipment will be deployed?

34. Is the exterior light intensity on the

ground at least 0.02 fc? -

35. Can the lighting system withstand the crash

conditions listed in Section 6.7.2.4, and

still function?

219
Yes No A

36. Is emergency lighting power independent of aircraft power systems?

37. Can the emergency lighting system be actuated both automatically and manually?

6.9.6.4 Emergency Exit Markings

38. Are emergency exits clearly marked both inside and outside the aircraft?

39. Are instructio..s for releasing the exits

clearly marked beside the exit release mechanisms?

40. Do all exit markings meet the requirements

of the Department of Army Technical Bulletin 746-93-2?

6.9.7 Crash Locator Beacon Checklist

1. Can the crash locator beacon be activated

both automatically and manually?

2. Is an inertia sensor used to automatically

activate the beacon?

3. Do the longitudinal inertia sensors in

fixed-wing aircraft meet the actuation

limits shown in Figure 58, Section 6.8.2?

4. Are both longitudinal and vertical inertia

sensors provided in rotary-wing aircraft?

5. Do the inertia sensors in rotary-wing aircraft meet the actuation limits shown in

Figures 58 and 59?

6. Is the inertia sensor mounted solidly to

rigid structure located in an area that will

experience crash forces representative of

those in the occupant compartment?

7. Are the transmitter and antenna located in

areas that are not subject to impact damage?

8. Can the transmitter and antenna withstand

the crash forces listed in Section 6.7.2.4?

220
P Yes No N/Ai

9. Does the crash locator beacon have its own

independent power supply?

10. Is all electrical wiring between system com- !

ponents protected from impact damage?

2211

Wi

tI

{M

1

I

it

)I

! I

g : I

S221 ': -

SI i I___________ l I I,• ___ ___ __ __ _
REFERENCES

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A

________ I
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L _ ___
REFERENCES (Continued)

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224

224 1
REFERENCES (Continued)

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225

._ 1
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226

'-.
F

~ •REFERENCES (Continued)

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52. Swearingen, J. J., EVALUATIONS OF VARIOUS PADDING MATERIALS FOR CRASH PROTECTION, FAA Technical Report AM 66-40,

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227

S i 41 .. ! ' 1 11 ' l '' t I I I
REFERENCES (Continued)

58. ASTM C 542-76, STANDARD SPECIFICATION FOR LOCK-STRIP GASKETS, 1977 Annual Book of ASTM Standards, R. P. Lukens,

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228
S_7

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23

i. 238 I _i

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S..... il-q I~ l I '£ 1 'I I I I ,-
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249

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Development Center, Warminster, Pennsylvania, February 1977.

Zinn, S. U., Jr., CRASH FIRE HAZARD EVALUATION OF JET FUELS,

Federal Aviation Administration; SAE Paper No. 720324, presented at National Business Aircraft Meeting, Society of Automotive Engineers, Inc., Wichita, Kansas, March 1972.

264
INDEX

Acceleration, definition 25

Aircraft coordinate system 23

Aircrew restraint 98

Ancillary equipment retention strength 71

testing 78

Anthropometry 52

Anthropomorphic dummy

definition 31

seat testing 57, 135

Autoignition temperature, definition 38

P Beacon, crash locator 200

Blade strikes 59, 68

Body block 133

Boost pumps 182

Brightness, definition 39

Buckling, fuselage 63, 64

Buttock reference line, definition 35

Cargo restraint

airframe interface 73

requirements 74

testing 78

Class exits

definition 39

location 196

size 195

Controls, frangible 157

Crash force angle, definition 29

Crash force resultant, definition 28

Crash sensor 201

Crashworthiness

cost 18, 42 k preliminary design 21 Crew chief restraint 101

Cushions 95

Deceleration pulse shape 42, 62

* Design conditions

Army requirements 59

landing gear 69

lateral impact 65

* longitudinal impact 62

rollover 65

seats 83

velocity 45

vertical impact 64

Ditching 193, 194

Dummy. See Anthropomorphic dummy Dynamic overshoot, definition 29

265

L ___;.)
It

INDEX (Continued)

Dynamic testing

data handling 140

litters 147

seats 134, 135

Earth scooping 58, 62

Effective weight

definition 33

seat design 123

Electrical systems 187

Emergency equipment release 73

Emergency exits 193, 194

Emergency lighting 193, 199

Empennage design 68

Energy-absorbing devices characteristics 94

load 125

stroke 119, 124, 138

Energy absorption

cockpit requirements 158

cushions 96

fuselage 59

seats 96

webbing 118

Engine mounts 68

Environmental hazards, definition 150 Exits

classes defined 39

location 196

size 195

Fire curtain, definition 37

Fire-resistant material, definition 37

Firewall, definition 38

Flame propagation index, definition 38

"Flame resistant, definition 38

Flammability

FAR 25.853 190

UMTA guidelines 191

Flight path angle, definition 24

Floor design 64

Frangible attachment controls 157 

- definition 37

fuel lines 174, 177

Fuel drains 182 Fuel lines

construction 173

location 174

266
'. • INDEX (Continued)

Fuel system

boost pump 182

design criteria 172

drains 182

self-sealing couplings 174

valves 183 -"

vents 184

Fuel tank

construction 173

location 172

Fuselage

Sblade strikes 59, 68

buckling 63, 64

earth scooping 58

Senergy absorption 59

floor 64

lateral impact 65

longitudinal impact 62

rollover 65

shape 69

vertical impact 64

wings 68

Gunner restraint 101

Head impact

hazards 151

injury 46, 48

Headrest 97

Hose fitting strength 174 I Human body coordinates 30

Human tolerance

design 20

head injury 48, 152

* leg injury 52

spinal injury 51

* whole body 46

Hydraulic system

construction 185

criteria 185

location 186

Ignition source control

electrical system 187

* •shielding 189

Ignition temperature, definition 38

Impact angle, definition 24

Impact attitude, rotary wing 43

Impact velocity 43

Impacted surface 43

267

I 1i
4

INDEX (Continued)

Inertia reel 118

Inflatable restraint 103

Injury

head 46, 48 leg 52

spinal 46, 51

Instrument panel padding 156, 159, 165

Interior materials 190

Landing gear design conditions 69

location 70

testing 78

Lap belt anchorage 108

Lateral impact design conditions 65

Leg injury 52

Lighting, emergency 193, 199

Litters

orientation 85

strength 142

testing 146

Longitudinal impact design conditions 62

Major impact, definition 26

Marking, exits 200

Materials

restraint hardware 112, 117, 118

seats 86, 92

MIL-S-58095(AV) 20

MIL-STD-850 33

MIL-STD-1290(AV) 16, 20, 60, 156

Optical density, definition 39

Padding materials 160

testing 160

Pitch

defined 24

impact conditions 43

vertical impact 64

Preliminary design 2.

Restraint system

adjustment hardware 112, 117

aircrew 98

anchorage 93, 108, 109

de-sign principles 97

gunner 101

inflatable 103

release 106

troop 101

268
INDEX (Continued)

98 S~types Swebbing 112, 113

Retrofit 20, 42

Roll angle I definition 24

impact conditions 43

vertical impact 64 t Rollover, design conditions 65

pedals 156 SRudder 

Seat

attachment 87

cushions 95

design conditions 83

energy absorption 94

reference point, definition 35

strength 119, 121

stroke 46, 56, 119,'124 testing 128, 135

types '84

Self-sealing hardware

couplings 174

valves 179

Shoulder harness anchorage 109

Sighting systems 157

Smoke emission

UMTA guidelines 191

Specific energy absorbed, definition 36

Spinal injury 46, 51

Static testing

litters 146

seats 129

Strike envelopes 148

Stroke, energy-absorbing 119, 124, 138

Submarining, definition 32

Survivable accident, definition 30

Systems analysis 20, 138

Terrain

angle, definition 24

impacted 43

Testing

cargo restraint 78

dynamic 135

landing gear 78

litter 146

restraint system 128

seat 88, 128

static 129, 146

269
INDEX (Continued)

Transmissibility, definition 30

Transmission mount 68 Troop restraint 101

Velocity, impact lateral 65

longitudinal 62

vertical 64

Vents, fuel system 184

Vertical impact design conditions 64

landing gear 71

Webbing, restraint system energy-absorbing i18

joints 115

strength 113

Weight

aircraft design 62

effective 123

energy absorber design 125

seat occupant 119

test dummy 137

Wing design 68

Yaw angle

definition 24

impact condition 43

270 I