USAAVSCOM TR-85-D-4

AD-A162 097
ADVANCED TECHNOLOGY HELICOPTER LANDING GEAR
PRELIMINARY DESIGN INVESTIGATION

D. Lowry
SIKORSKY AIRCRAFT DIVISION
United Technologies Corporation
Stratford, Conn. 06601

October 1985

Final Report for Period 1983 - 1984 j!

Approved for public release;

LI distribution unlimited.

C.2

Prepared for
AVIATION APPLIED TECHNOLOGY DIRECTORATE
US ARMY AVIATION RESEARCH AND TECHNOLOGY ACTIVITY (AVSCOM)
Fort Eustis, VA. 23604-5577
AVSCOM - PROVIDING LEADERS THE DECISIVE EDGE

_ 85 12 2 165
 AVIATION APPLIED TECHNOLOGY DIRECTORATE POSITION STATEMENT

This preliminary design effort is one cf two parallel contractual investigations to develop landing
gear structural configurations and determine the associated weight changes for increased crash-
worthy capabilities. The contractor developed three baseline landing gears; a retractable gear de-
signed to normal loading requirements, and two crashworthy configurations, fixed and retractable,
designed to meet MIL-STD-1290 crashworthy requirements. Weight sensitivity analyses were
conducted for the crashworthy designs to determine the energy absorbing component weights for
various sink speeds and aircraft attitudes. These results are used with other landing gear design
parameters to establish recommended crashworthy design requirements.

The results of this program represent a significant advance in the understanding of the parameters
which influence crashworthy landing gear and energy absorbing structure weight. These findings
will be integrated with the parallel contract and past efforts in landing gear weight sensitivity to
develop less costly, more efficient crashworthy systems.

Mr. Geoffrey R. Downer of the Aeronautical Technology Division, Structures Technical Area,
served as project engineer and Mr. Drew G. Orlino of the Aeronautical Systems Division, Safety
and Survivability Technical Area, assisted in this effort.

DISCLAIMERS

"The findingsin this rpport are not to be construed as an official Department of the Army position unless so
"a? designated by other authorized documents.

When Government drawings, specifications, or other data are used for any purpose other than in connection with a
definitely related Government procurement operation, the United States Government thereby incurs no responsibility
nor any obligation whatsoever; and the fact that the Government may have formulated, furnished, or in any way
supplied the said drawings. specifications, or other data is not to be regarded by implication or otherwise as in any
manner licensing the holder or any other person or corporation, or conveying any rights or permission, to manu-
facture, use, or sell any patented invention that may in any way be related thereto.

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

DISPOSITION INSTRUCTIONS

%
Destroy this report by any method which precludes reconstruction of the document. Do not return it to the
originator.

"'a
a.-. . . . . ... ., - .. .- . .- , .-. .
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"4PERFORMING ORGANIZATION REPORT NUMBER(S) S. MONITORING ORGANIZATION REPORr NUMBER S)

SER 510148 USAAVSCOM TR 85-D-4

6a. NAME OF PERFORMING ORGANIZATION * b OFFICE SYMBOL 7a. NAME O( MONITORING ORGANIZATION

"Sikorsky Aircraft Division (if appicabi*) Aviation Applied Technology Directore

6C ADDRESS (City, Staff, and ZIPCode) 7b ADRESS (City, State, arcd ZIP Code)
United Technologies Corporation U.S. Army Aviation Research and Technology
Stratford, Connecticut 06601 Activitv (AVSCOM)
Fort Eustis, Virginia 23604-5577
8& NAME OF FUNDING ISPONSORING Sb. OFFICE SYMBOL 9. PROCUREMENT INSTRUMENT IDENTIFICATION NUMBER
ORGANIZATION (if applicable)
DAAK51 -3-C-O040
8c. ADDRESS (City; State, and ZIPCode) 10 SOURCE OF FUNDING NUMBERS
PROGRAM PROJECT TASK IWORK UNIT
ELEMENT NO. NO. 1 L162- NO. ACCESSION NO.

62209A 209AH76 04 033

11 TITLE (Include Security Classification)

Advanced Technology Helicopter Landing Gear Preliminary Design Investigation

12 PERSONAL AUTHOR(S)
F .D. Lowry
13a. TYPE OF REPORT 13bTIME COVERED 14 DATE OF REPORT (Year. Month, Day) S. PAGE COUNT
Final I FROM 1983 TO 1984 IOctober 1985 I 122
16 SUPPLEMENTARY NOTATION

17 COSATI CODES 18 SUBJECT TERMS (Continue on reverse of necessary and identify by block number)
FIELD GROUP SUB-GROUP 't, Crashworthy> Shock Strut'
Drag Strut. Velocities, .
Landing Gear, Weight
19 ABSTRACT (Continue on reverse if necessary and identify by block number)
IA preliminary design investigation has been performed to develop weight and cost sensitivities of various land-
ing gear systems for a 10,000-pound class LHX helicopter. Weights are established for three baseline main
landing gear systems: a noncrashworthy retractable, a crashworthy retractable, and a - rashworthy fixed. Each
"system is capable of kneeling the LHX helicopter. Weights are based on preliminra t structural analysis of land
ing gear loads developed by a computer program KRASH. KRASH was used to obtain the design loads at
various sink rates, and pitch and roll attitudes. The design loads were used to size the landing gear structure
in order to develop the landing gear weight for the different impact conditions.
Main landing gear system costs are developed from the weight data established.,
An aerodynamic drag assessment was performed. Criteria for crashworthy designs are recommended. The
recommended criteria were developed frt-m che weight trends, costs, and UH-60A Class A Mishaps. . , -

20 DISTRIBUTION/ AVAILABILITY OF ABSTRACT 121. ABSTRACT SECURITY CLASSIFICATION

0 UNCLASSIFIED/UNLIMITED 0 SAME AS RPT 0 OTiC USERS Unclassified
22a NAME OF RESPONSIBLE INDIVIDUAL 22b TELEPHONE (Include Area Code) 22c OFFICE SYMBOL
G. Dowaer , (804)878-5476 1 SAVDL-ATL-ATS
DO FORM 1473,84 MAR .. APR edition may be used until exhausted. SECURITY CLASSIFiCATION OF TH:S PAGE
', All other editions are obsolete Unclassified
Ucasife

a-. -A
 PREFACE

This Advanced Technology Helicopter Landing Gear Preliminary Design

Investigation was conducted under contract DAAK51-83-C-0040 with the
Aviation Applied Technology Directorate, U.S. Army Aviation Research
and Technology Activity (AVSCOM), Fort Eustis, Virginia.
The work was performed under the general direction of Mr. G. Downer
of the Aviation Applied Technology Directorate. Sikcrsky Aircraft
principal participants were Mr. S. Garbo, Project Manager, Mr. D.
Lowry, Principal Investigator, Ms. S. Hess, Loads and Criteria,
Mr. M. Pramanik, Crashworthiness, and Mr. T. Obenhoff, Landing Gear
Design.

-Accesion For
NTI1 CRA&1
DTC TAB
Unannounced

Jutf t..........

By

Av3;&biity Codes

"'4 Dist
Dist I p,cal
Avaii LJ ;or ,..

iii
 TABLE OF CONTENTS
PAGE

PREFACE ............... ..................... iii

LIST OF ILLUSTRATIONS ....... .............. .. vii

LIST OF TABLES .............. ................. x

INTRODUCTION .............. .................. 1

BASELINE HELICOPTER ....... .................. 2

BASELINE LANDING GEARS .......... ............. 5
Design Loading Requirements ..... ......... 5
Ground Handling ... ............. . 5
Obstruction, Symmetric and
Asymetric Reserve Energy .... ....... 6
Taxiing ............. ................. 6
Supplemental Design Requirements . ... 7
Vertical Impact Design Conditions
Envelope Crashworthy Gears ... ...... 8
Additional Crashworthy Design
Requirements .......... ............. 8
Pitch, Roll and Velocities .... ....... 8
Noncrashworthy, Retractable Main Gear . ... 9
Crashworthy, Retractable Main Gear ... ..... 12
Crashworthy, Fixed Main Gear ... ........ .. 15

I
"Tail Landing Gear ....... .............. .. 19
Drag Beam - Main Gears .... ........... .. 21

LOADS ANALYSIS ................................ 22

Design Mass Properties .... ........... .. 22
Main Rotor Parameters ..... ............ .. 22
Landing Gear Parameters .... ........... ... 22
Main Gear ....... ................ ... 22
Tail Gear ....... ................ ... 22
Normal Operating Conditions ... ......... ... 27
Static Design (Landing) ... ......... ... 27
Static Design (Ground Handling) ..... ... 27
Xjrmal Landing RequirementsIs ....... .... 27
Crashworthy Design Requirements ......... ... 28

KRASH ANALYSIS .......... ................. .. 33

KRASH Program Input ..... ............. ... 33
KRASH Model Mass Properties ... ......... ... 33
KRASH Model Spring Properties ... ........ .. 36
Lift Forces ......... ................. ... 37

v
 TABLE OF CONTENTS (Cont'd)

PAGE

KRASH Program Results ..... ............ .. 37

Time of Zero Vertical Velocity at
Aircraft C.G ...... .............. ... 37
Landing Gear Strut Maximum Stroke
and Time of Occurrences ... ........ .. 39
Ground Load ......... ................. ... 41
Inertial Load Factors ..... ............ .. 47
Seat Stroking ....... ................ ... 51
Fuselage Crushing ...... ............. .. 54

PRELIMINARY STRUCTURAL WEIGHT ANALYSIS ..... .. 59

Structural Loads (Main Gears) ........... ... 59
Tail Gear ......... .................. ... 65
Baseline Landing Gear Weight ... ........ .. 65
Composite Landing Gear Components ........ ... 68

WEIGHT SENSITIVITY ANALYSIS .... ........... ... 72

Crashworthy Systems ..... ............. ... 72
Component Weight .......................... 82
Landing Gear System Cost .... .......... .. 87

AERODYNAMIC DRAG .......... ................ .. 91

RECOMMENDED DESIGN CRITERIA .... ........... ... 94
Attitudes and Velocities .... .......... .. 94
Detail Design Criteria for the Main
Landing Gear ............................ 96
DESIGN UPDATE ......... .................. ... 97
Shock Strut Structural Analysis ......... ... 97
Optimized Shock Struts - Baseline and
Updated Designs ....... .............. .100
Weight Substantiation ..................... .. 104
Aerodynamic Drag - Updated Design ......... .. 108
Fabrication Methods and Processes ........ .. 108
Reliability................................ * 108
Maintainability ....... ............... .. 108
CONCLUSIONS ........... ................... ... 109
SRECOMMENDATIONS .ii ........ 10
REOMEDTIN.............................1

REFERENCES ............ ................... .. 111

APPENDIX - Weight Sensitivity and KRASH
Analysis ....... .............. .. 113

vi~* * * -,~'~,.t.
 LIST OF ILLUSTRATIONS
Figure Page

1 Baseline LHX General Arrangement . . .. 3

2 Crashworthy Design Envelope .... ...... 8

"3 Retractable, Kneeling, Noncrashworthy,
Main Landing Gear ........ ........... 10

4 Shock Strut, Noncrashworthy Main

Landing Gear ....... ............ .i.i.. 11

5 Retractable, Kneeling, Crashworthy

Main Landing Gear ... ........... .. 13
%-4

-. 6 Shock Strut, Crashworthy Retractable

Main Landing Gear ... ........... .. 14
7 Fixed, Kneeling, Crashworthy Main
Landing Gear .... .............. ... 16

8 Shock Strut, Fixed, Crashworthy Main

Landing Gear ......................... 17
9 Baseline Tail Gear, Retractable . . . 20

!0 Drag Beam, Main Landing Gear ...... .. 21

*11 Main Gear Tire Load Vs. Tire Deflection. 25

12 Tail Gear Tire Load Vs. Deflection . . 26

13 Pitch and Roll Attitudes in Crash impacts 32

, 14 KRASII Model ...... .............. .. 34
15 Final Model ...... .............. .. 35

16 Typical Load Vs. Deflection Curve . . 36

17 Drag Load Time History (Typ.) ..... .. 44

18 Side Load Time History (Typ.) ..... .. 45

19 Vertical Load Time History (Typ.) 46

20 Vertical Acceleration Time History on High

'.. Mass Items ........................... 49

vii

m .J m . \*.**.
 LIST OF ILLUSTRATIONS (Cont'd)

Figure Lage

21 Longtudinal Acceleration Time History

on High Mass Items ... ........... ... 50

22 Vertical "G" Pulse On Seat ....... ... 52

23 Sketch of Subfloor Structure ...... .. 56

24 Fuselage Crushing ... ........... .. 57

25 Crushing Energy - Kevlar Members . . .. 58

26 Shock Strut - Noncrashworthy, Kneeling,

Retractable Main Landing Gear ..... .. 61
27 Main Gear Drag Beam ... .......... .. 62

28 Shock Strut Assembly, Crashworthy,

Retractable, Kneeling, Main Landing Gear 63

29 Fixed Kneeling, Crashworthy Main

Gear Strut ...... ............... ... 64
30 Tail Gear Structure ... .......... .. 56
31 Tail Gear Drag Strut ... .......... .. 67

32 Composite Main Gear Drag Beam Compared

to Steel ........ ................ .. 69
33 Pilot Seats Vertical Load - Gears Retracted 74
34 Fuselage Crushing Results .. ....... .. 76
35 Vertical Load Factors - High Mass Items. 78

36 Crashworthy Systems (Reference 2). . . . 79

37 Main Gear Component Loads/Drag Beam

Weight .......... ............... .. 81

38 Weight of Affected Structure as a

Function of Pitch and Roll Angles. . . . 85

39 Landing Gear System Weight as a

Function of Pitch and Roll Angles. . . . 86

viii

4
 LIST OF ILLUSTRATIONS(Cont'd)

Figure No. Page

40 Landing Gear Parametric Cost Study . . . 89

41 Landing Gear System Cost as a Function

of Pitch and Roll Angle ... ......... ... 89

42 Weight and Cost Trends for Landing

Gear System ....... ............... ... 90

* - F

5..,

ix
t5,.
 LIST OF TABLES

Table
1 Weight Summary ........... .............. 4

2 Mass Properties ...... ............. .. 23

3 Aircraft Properties and Parameters . . .. 24

4 Normal Landing Gear Design Conditions

and Ground Loads ..... ............. ... 29

5 Ground Handling - Gear Loads ......... ... 30

6 Design Conditions for Crash Impacts . 32
7 Time of Zero Vertical Velocity at
Aircraft C.G ....... ............... ... 38
8 Summary of Strut Actions ... ......... ... 40
9 Ground Loads Data ...... ............ .. 43
10 Summary of Vertical Inertia Load Factors 48
11 Summary of Cockpit Seat Stroking ..... ... 53

12 Summary of the Understructure Crushing 55

13 Ground Loads and Strut Loads - Main Gear 60

14 Landing Gear Weight Summary .. ....... .. 70
15 Landing Gear Weiqht Comparison ........ .. 71

16 Crew Seat/Troop Seat System Weight .... 73

17 Airframe Weights ..... ............. ... 75

18 Main Gear Shock Strut Weight Summary . 84

19 Landing Gear Parametric Study .. ...... .. 88

20 UH-60A Data Comparison Class A Mishaps
(At Major Impact) ....... ............ .. 95
21 Shock Strut Components - Margins of
Safety - Baseline Crashworthy Designs 99

-. . . . . . . . . . .~ ~-' ----. -~.- -. -k- - - -- '- - .. -. -

 LIST OF TABLES
Table Page
22 Baseline Shock Strut Weighc Reduction
For Minimum Margins of Safety ......... .. 103
23 Shock Strut Components, Wall Thickness 104

24 Main Gear Shock Strut Weight Summary,

* Adjusted Weight ........ ............. .. 105
25 Updated Landing Gear Shock Strut Weight 106
26 Updated Main Gear Shock Strut Weight
Summary ............ ................. .. 107
A-I Weight Sensitivity - Shock Struts and
Fittings - Fixed Gear - (42 FPS) ....... .. 114
A-2 Weight Sensitivity - Shock Struts and
Fittings - Fixed Gear - (36 FPS) ....... .. 115

A-3 Weight Sensitivity - Shock Struts and

Fittings - Fixed Gear - (30 FPS) ....... .. 116
A-4 Weight Sensitivity - Shock Struts and
Fittings - Retractable Gear - (42 FPS) 117

A-5 Weight Sensitivity - Shock Struts and

Fittings - Retractable Gear - (36 FPS) 118

A-6 Weight Sensitivity - Shock Struts and

Fittings - Retractable Gear - (30 FPS) 119
A-7 Mass Properties and Locations ......... .. 120

A-8 Node Point Locations .... ........... .. 120

" A-9 Beam Data .......... ................ .. 121
A-10 Material Properties ...... ........... .. 121
A-lI Beam End Fixity Data .... ........... .. 122

A-12 Spring Data .......... ............... .. 123

A-13 Oleo Strut Input Data ...... ............ 124

xi
 INTRODUCTION
Energy absorbing landing gears play a key role in meeting
helicopter crashworthiness design goals of reduced crash
injuries, fatalities, and material losses. However, the
ability to absorb large amounts of energy in the landing gears
is not achieved without paying strict attention to the energy
absorption requirements in the design effort. Furthermore, ex-
perience has shown that this capability cannot be realized
without some increase in gear weight and complexity, which in
turn has an adverse effect on the aircraft performance and
cost. These effects will, of course, vary with the level of
crash energy which must be absorbed and with the severity of
the crash attitude requirements.

The Army's most recent production helicopter, the UH-60 BLACK

HAWK, is designed to meet some of the requirements of MIL-STD-
1290 (AV) "Light Fixed and Rotor Wing Aircraft Crashworthi-
ness". Also, under the Advanced Composite Airframe Program, a
landing gear has been designed to meet most of the crash re-
quirements of MIL-STD-1290. However, these aircraft systems
are designed using a nonretractable landing gear. It is,
therefore, appropriate to conduct a preliminary design investi-
gation of a retractable landing gear, especially in light of
the Army's increased emphasis on increased forward flight
performance, in order to evaluate the applicability and effect
of crashworthiness design requirements.
This report documents the results of a three-phase effort of a
preliminary design investigation for a retractable and non-
retractable landing gear system designed to crashworthy re-
quirements.

-------------------------
 BASELINE HELICOPTER
Sikorsky's Advanced Blade Concept (ABC t m ) LHX baseline configur-
ation developed for the Advanced Technology Helicopter Landing
Gear Preliminary Design Study was derived from the utility
"variant air vehicle provided in Reference 1. However, the
Landing Gear Study aircraft was configured more in accordance
with the LHX requirements outlined in RFQ DAAK51-83-Q-0061,
dated July 1, 1983.
"The principal physical characteristics of the ABC LHX baseline
configuration are shown in Figure 1, the Baseline LHX General
Arrangement drawing. The accommodations for six troops and a
crew of two determined the cabin and cockpit size, and there-
fore, the body length.
The takeoff gross weight of this utility helicopter, summarized
in Table 1, is 10,000 pounds. The counter-rotating main rotors
and the auxiliary propulsor are driven by twin advanced tech-
nology engines (ATE's). The engines are installed behind the
main gearbox enclosed by easy opening cowlings to allow for
access, inspection, and/or maintenance.
A shroudeC pusher propellei is used to provide efficient and
quiet cruise thrust. The shrouded configuration will protect
ground personnel and minimize damage to the propeller while
operating on, or close to, the ground.

The Sikorsky ACAP diamond-shaped cross section was sele2ted for

the baseline body shape to provide a reduced body radAr return
at a minimal weight penalty. In addition, the shape of the
aircraft is used to control the deflection of the structure.
The body shape was iterated on as the landing gear design
progresses to insure that optimum airframe landing gear inter-
face and load carrying paths are established.

The landing gears are part of a crashworthy system. A crash-

worthy system include: crashworthy crew and troop seats, fuel
system, tub structure, and retention of high mass items. The
fuel system and the high mass items (engines transmission and
rotor head) are located Lehind the occupied area of he cabin
and cockpit. Locating the fuel system and high mass items
behind the occupied areas rtduces the possibility of peretra-
tion of tha systems into the occupied areas during a crash.
The structure supporting the fuel system and high mass items
becomes the major airframe structure.

2
 -~ ~- * 'L~.W?~VT.J W -w---

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Crash dynamic loads, normal operating loads, and weights
evaluation were developed relative to the baseline aircraft.
Prior to analysis of loads, landing gear configurations were
established. The main gear:- are located as close to the
helicopter'z forward center of gravity as feasible to utilize
the major airframe structure or load carrying members for the
landing loads. Main landing gears forward of the center of
gravity require a tail gear.

TABLE 1. WEIGHT SUMMARY

Weight (lb)
Total Structure 2272
Total Propulsion 2129
Total Equipment 3088

Total Weight Empty 7489

Useful Load 2511

Gross Weight 10,000

4
 BASELINE LANDING GEARS
The landing gear configurations described herein are compatible
to the LHX-utility class helicopter in the 6000- to 10,000-pound
gross weight range. This version is configured for two crew
members and six troops located forward of the heavy mass items
such as rotor head and transmission. The aircraft is con-
figured with a tail wheel landing gear arrangement as shown in
Figure 1.
To comply with the crashworthy requirements and to meet th, air
transportability requirement, the main landing gear is Located
outboard of the troop cabin area with the shock strut posi-
tioned aft of the cabin door frame. The mounting point for the
shock strut upper stage is provided by the aft cabin bulkhead
which also supports the heavy mass items such as transmission
and rotor head. To simplify the kneeling and/or retraction
geometry, a trailing arm (or "drag beam") concept is utilized,
thus allowing landing gear movement in a vertical plane
throughout its stroke without obstructing the cabin door or
infringing upon the space occupied by per-onnel. In the event
of a crash impact this arrangement allows the gear to stroke
without penetrating the fuel cells or occupied areas. AMCP
706-202 was used in determining the required angles defined in
Figure 1.
3 All shock strut configurations are air-oil types designed in
F accordance with MIL-L-8552 and using either a variable
orifice
assembly or metering pin. Both static and dynamic seals,
sealing surfaces, and seal grooves are designed in accordance
with MIL-G-5514. For the baseline concepts, the various
components comprising the shock struts and mounting or support-
ing structure are manufactured from heat treated, high strength
steels (4340, 300M) or aluminum alloys (7075, 7175).
The tire sizes were selected based upon the landing gear static
reactions taken at the helicopter design gross weight and to
meet the requirement for a CBR of 2.5, equivalent to the H-60
series helicopter. The tires meet the requirements of
MIL-T-5041, whereas wheels and brakes comply with requirements
of MIL-W-5013. All main and tail landing gears described
herein use the same wheels, tires, and brakes.
DESIGN LOADING REQUIREMENTS

Each landing gear system is designed to meet the following

requirements:

Ground Handling
a. Towing - The towing requirements shall be in accordance
with MIL-A-8862, and the basic design gross weight shall apply.

%k
b. Jacking - Jacking requirements shall be in accordance with
MIL-A-8862, and the basic design gross weight shall apply.

c. Mooring - Mooring requirements shall be in accordance with

MIL-A-8862 except that a 70-kt horizontal wind shall apply.

d. Transport - Transport requirements shall be in accordance

with AMCP 706-201 requiring a limit vertical load factor of
2.67 at the basic design gross weight.

Obstruction, Symmetric and Asymmetric Reserve Energy

"a. Obstruction - Obstruction landing requirements shall be in

accordance with MIL-S-8698 at the basic design gross weight.

b. Symmetric and Asymmetric - Symmetric and asymmetric

landing conditions for tail wheeled gear configurations shall
be in accordance with AMCP 706-201 except that landing surfaces
shall be flat and firm. Limit landing conditions shall also be
in accordance with AMCP 706-201 at the basic design gross
weight.

Taxiing
a. Two-Point Braked Roll - The requirements of MIL-A-8862
shall apply for the two-point braked roll except that the
vertical load factor at the center of gravity (CG) shall be 1.2
for all gross weights.
b. Three-point Braked Roll - The requirements of MIL-A-8862
shall apply to the three-point braked roll of a helicopter with
nose wheel landin9 gear except that the vertical load factor
at the CG shall be 1.2 for all grcss weights.

c. Reverse Braking - The requirements of MIL-A-8862 shall

apply.
d. Wheel, Biakes, and Tire Heating - The requirements of
MIL-W-5013, MIL-T-5041, and MIL-B-8584 shall apply.

e. Turning - The turning requirements of MIL-A-8862 shall

apply.

6
 f. Pivoting - The pivoting requirements of MIL-A-8862 shall
apply.

g. Taxiing - The taxiing requirements of MIL-A-8862 shall

apply.

h. Special Tail Gear Conditions - The special tail gear

conditions of MIL-A-8862 shall apply to a tail gear helicopter,
including a tail gear obstruction condition.
Supplemental Design Requirements
a. The horizontal speed requirement shall include all speeds
from zero up to 50 knots at limit sink speed on level ground
and from zero un to 40 knots at reserve energy sink speed on
level ground at th basic design gross weight.

b. Consideration siall be given to both crew-initiated and

automatic retractable wheeled landing gear during transition
to/from NOE flight.
c. Design fatigue loads shall be considered.
d. A landing gear shall have kneeling provisions for trans-
portability.

O.f,
Vertical Impact Design Conditions Envelope Crashworthy Gears

The crashworthy landing gear systems are designed to sink speed

and altitudes given in Figure 2.

40 40

200 20

42 IT/SIC / .. . . __\ 36 FT/S1C

.10

Figure 2. Crashworthy Design Envelope

Additional Crashworthy Design Requirements

1. Retracted or extended gear shall not, upon crash impact,
protrude into occupied id fuel cell areas.

2. A retracted gear shall provide some vertical crash impact

energy attenuation.
3. Fixed and extended/retractable gears shall remove their
design incremental energy at sink speeds up to z2 fps.

Pitch, Roll and Velocities

1. No fuselage-ground contact shall occur, and there shall be
minimal or no effect upon dynamic components during a crash
impact with the landing gear extended. For vertical velocities
of 15 and 20 fps, all combinations of the following pitch and
roll attitudes shall be considered:

Roll (deg) Pitch (deg)

0 -5
+5 0

8
 +10 + 7.5
+15 +15

2. Injurious loadings, as defined in MIL-STD-1290, shall not

be transmitted to the crew with the gear retracted, assuming
energy absorbing subfloor structure and stroking seats. For
vertical velocities of 25, 30, and 35 fps, all combinations of
the following pitch and roll attitudes shall be considered:

Roll (deg) Pitch (deg)

0 -5
+5 0
+10 + 7.5
+15 +15

3. Injurious loadings, as defined in MIL-STD-1290, shall not

be transmitted to the crew with the gear extended, assuming
energy absorbing subfloor structure and stroking seats. For
vertical velocities of 30, 36, and 42 fps, all combinations of
the following pitch and roll attitudes shall be considered:

Roll (deg) Pitch (deg)

0 -5
+5 0
+10 + 7.5
+15 +15

NONCRASHWORTHY, RETRACTABLE MAIN GEAR

The normal main landing gear designed to the operating condi-

tions as specified in the Design Loading Requirements is a
noncrashworthy, retractable landing gear with kneeling capa-
bilities to meet the air transportability requirements. The
upper structural member, or retraction brace, and hydraulic
retraction/extension actuator are mounted to the aft cabin
bulkhead as shown in Figure 3. In addition to retracting and
extending the landing gear, the hydraulic ;actuator also con-
trois the kneeling function of the gear. Uplocks and downlocks
within the hydraulic actuator locks the gear in its respective
position during the retraction/extension cycle. The lower
stage is an air-oil shock strut, as shown in Figure 4, designed
to withstand the normal landing loads. The strut consists of a
piston inside a cylinder housing, both components manufactured
from 7075-T73 aluminum alloy. Internally, the strut assembly
contains an aluminum alloy metering pin and a floating piston

X.
 which separates the air from the hydraulic fluid. This assem-
bly acts as a shock strut capable of withstanding normal
landing loads at 10 feet per second (fps) sink speeds with a
reserve energy capability to 12.25 fps. Total wheel stroke
from extended to compressed position is 12.00 inches, which
results in a shock strut stroke of 11.90 inches per data from
the KRASH computer program. The retraction brace connects the
hydraulic actuator and the shock strut assembly to the aft
cabin bulkhead. This member, as depicted in Figure 3, is made
from 7075-T73 aluminum alloy. Mounting hardware is comprised
of standard self-lubricating bushings and bearings, and is
attached with high strength bolts.

A hydraulic system consisting of appropriate pump, valves, and

reservoir with hydraulic fittings and hose assemblies is
incorporated into the helicopter for providing power to the
hydraulic actuator. An electrically operated system activates
the hydraulic cycling of the landing gear and also includes the
respective limit switches to indicate uplock and downlock
positions.

RETRACTED POSITION
HYDRAULIC

ACTUATOR
1A

I\ ~RETRACTION
BRACE
I
il _\ -. -

all ~
'
1Kf~
IPIVOT
/ FUEL CELL
") POINT

300
SHOCK STRUT
WL
43.1
STATIC GROUND LINE , . / WL 35

iSTA
259.25

Figure 3. Retractable, Kneeling, Noncrashworthy

Main Landing Gear
10

"N
- Y," - ' " -"- - . '-% ' + % " - ' ' - _ .".- " ' ' 'r -. ' '._. ." -. " ' - ", ". " "
 HYDRAULIC
BLEED PORT SERVICE/

AIR
OIL

AIR VALVE FLOATING PISTON METERING PIN

REBOUND DAMPER VALVE

Figure 4. Shock Strut, Noncrashworthy, Main Landing Gear

4.4,

* -,'. "- , % " = , % " . ' ' , . _ . .. . .. . . , . . . , . ,

CRASHWORTHY, RETRACTABLE MAIN GEAR

The main landing gear configuration depicted in Figure 5 is a

ciashworthy, retractable landing gear designed to normal
operating requirements and the crashworthy requirements for the
vertical impact design conditions. This concept is a trailing
arm design with a universal-mounted, two-stage, air-oil strut
designed to absorb 65 percent of the energy at a sink speed of
42 fps at a pitch and roll of zero degrees. In addition to
meeting the crash impact criteria, the strut acts a6 a hy-
draulic recraction and extension actuator which also controls
the kneeling action required for air transportability.
event of crash impacts, the strut strokes in a vertical In the
plane
as the drag beam pivots about the fuselage attachment point,
thus preventing the landing gear from protruding into the fuel
cell or occupied areas. Also in a crash condition, the drag
beam does not obstruct the cabin door opening, thereby allowing
rapid evacuation of personnel. With the landing gear retract-
ed, some vertical crash impact energy absorption is provided
through the compression of the tire, wheel deformation, and
controlled failure of the upper housing attachment fitting on
the rear cabin bulkhead.

Both upper and lower shock strut stages are air-oil types
designed to the requirements previously specified but utilizing
a variable orifice assembly in lieu of a metering pin, as shown
in Figure 6. With a metering pin, gear retraction would be
hindered because of the interference between this pin and the
floating piston. The lower stage is designed to withstand
landing loads up to 12.25 fps reserve energy sink speeds within
14.50 inches total strut stroke. This assembly contains a
piston inside a housing of which the upper portion is machined
as the piston of the upper stage. Both components are made
from 4340 heat treated steel. The upper housing which mounts
to the airframe fitting is made from 7075-T73 aluminum alloy.
Separating the air from the hydraulic oil in both stages is an
aluminum alloy floating piston. Each stage contains an orifice
assembly, split bearings, a lower housing bearing, and aluminum
alloy retraction/kneeling piston. To retract, or kneel, the
landing gear, hydraulic oil pressure is added to each respec-
tive housing port. At the same time,ports in the upper portion
of the housings are opened to allow hydraulic fluid displaced
during the retraction cycle to be bled into a reservoir. As
the hydraulic fluid is pumped into the strut, both stages
retract, resulting in gear retraction or aircraft kneeling. An
external uplock system shuts the retraction system off when the
gear is fully retracted. Reversing the cycle extends the gear.

12
 Full strut travel is 27.00 inches which corresponds to an axle
vertical travel of 29.00 inches, the travel required to fully
retract or kneel the main landing gear. Mounting hardware
consists of self-lubricating bushings and spherical bearing
with high strength bolts.

BL SHOCK STRUT STA

' I34.25 275.25

.108.25
ti WL
-. +

4 33.60 270
F REF

" ~UEL
,NK

"
KN E ELE..DD .7.87--9RETRACTED/KNEELED
-E - ]"-
POSITION B COMPRESSED
BL BL , -" STA SL
10 24 200 26.42' 4_ 1.W.

- - - 3 EXTEND E
BL DRAG BEAM STA
0 6L STA 300.0
41.25 259.25

Figure 5. Retractable, Kneeling Crashworthy

Main Landing Gear

13

-- p . - - K V A .. 3, k A
 v~r:~2~v~A'.T ~'W'J
z I'JJ' v- V 7 ir- 7-7 wz4 - 7.7W~ -:r~WT2 Wq-_ r,.j ~ . ~"-IPY7 V VJT

~\ L.

v~N
C,

wOw
'u
2 oi
-J -o0
-.

0. c -I-)

LUU

-4 >

w >

oM ~ rd (a

0. 0~ E65 d

0 Lf -i :c
LL I W rI. t c

0 - cc W COw 4-JJ
LLU

LU LUa-C
I- LLU,
I LUj 40 '. H

CL ,

0 I~ LU

-J 4
0 0U
U
-.1 CIO jL, c
>w t
"C F

CD 14

z- -- 0: -*
C)- *U
The main landing gear, in conjunction with the tail landing
gear, has an energy absorption capability to prevent the
fuselage from contacting the ground at a 20 fps sink speed.
Results from the KRASH Computer program indicate that each main
gear absorbs approximately 44% of the total energy for this
sink speed rate. To meet this criteron, the upper and lower
variable orifice assemblies are designed in such a way as to
allow the shock strut to absorb energy at crash impacts result-
ing from 20 fps to 42 fps sink speeds and still react the loads
due to normal landings up to 12.25 fps reserve energy sink
speed.
The hydraulic system for controlling landing gear retraction,
extension, and kneeling cycles is part of the helicopter
overall hydraulic system and contains the appropriate shuttle
valves, check valves, pumps, and accummulator or reservoi-,
with fittings and hose assemblies. An electrically operated
system activates the hydraulic cycling of the landing gear and
also includes the respective limit switches to indicate uplock
and downlock positions.
To comply with the Supplementa2 Design Requirements specified,
it is recommended that the landing gear be automatically
retractable with pilot override. The aircraft's flight speed
indicator and radar altimeter will provide signal indications
for automatic retraction and extension, so that below (to be
determined) knots and below (to be determined) feet from the
ground, the gear will always be in the extended position unless
the pilot chooses to override the indicators. Automatic
extension/retraction reduces the pilot's work load.

CRASHWORTHY, FIXED MAIN GEAR

The main landing gear concept shown in Figure 7 is a crash-

worthy gear complying with the normal operating requirements of
page 6 and the crashworthy requirements of page 8. This con-
figuration is also a trailing arm design with the shock strut
mounted to the lower end of the arm and to the upper fitting at
the rear cabin bulkhead. This arrangement allows the strut to
stroke in a fore and aft vertical plane as the drag beam pivots
about the fuselage attachment point in the event of crash
impacts, thus preventing the gear from puncturing the fuel cell
or protruding into the occupied area.

15
V

V .- -i[ F

,[ CRUSHABLE SPLIT SLEEVE

FOR CRASH ENERGY

!iI\ ABSORPTION, REMOVEABLE

FOR KNEELING/

-aIFUEL CELL

DRAG BEAM "-

, . STRUT SHOCK
WL AIR-OIL

43.1 ;TA

STATIC GROUND LINE WL 35 300

_ /STA
259.25

Figure 7. Fixed, Kneeling, Crashworthy Main

Landing Gear

16

.4,Z
 AIR VALVE -LOWER STAGE HYDRAULIC PORT

I-A-

AIR-OIL AIROI
STRUT
SRU -CRUSHABLE SPLIT
= SLEEVE HYDRAULIC7PORT
FOR CRASH ENERGY FOR KNEELING
WITH VARIABLE
ORIFICE ASSEMBLY ABSORPTION, REMOVEABLE
FOR KNEELING

Figure 8. Shock Strut, Fixed, Crashworthy Main

Landing Gear

17

K" . ..
The shock strut, as shown in Figure 8, is a universal-mounted,
two-stage, air-oil strut designed to dissipate energy at sink
speeds up to 42 fps crash conditions. The lower stage con-
sists of a 4340 steel piston inside a housing made from
7075-T73 aluminum alloy. The upper segment of this housing is
machined to become the piston for the upper stage. Separating
the air from the hydraulic fluid is an aluminum alloy floating
piston. The variable orifice assembly is designed to allow
this lower stage to absorb energy up to 20 fps sink speeds
within a 14.50-inch stroke. The upper stage contains an air
and oil chamber which is bled off into a reservoir when kneel-
ing the gear for air transportability of the helicopter. The
upper outer housing is also made from 7075 aluminum alloy and
attaches to the airframe bulkhead.
Separating the two stages is a crushable split sleeve for crash
energy attenuation and is removable for kneeling the landing
gear. The sleeve has a tapered section designed to crush
progressively as the load increases caused by sink speeds in
the range of 20 to 42 fps. Materials for this tube could be
either steel, aluminum, or composites. To kneel the landing
gear requires this sleeve to be removed and the hydraulic fluid
in both the upper and lower stages to be bled off until the
desired height is obtained. Reservicing the strut will extend
the upper stage, allowing reassembly of the sleeve. All
mounting hardware is comprised of standard high strength
fasteners used with self-lubricating, spherical bearings and
bushings.
The stroking sequence of the shock strut during crash condi-
tions is such that as the sleeve crushes the upper stage
compresses, which forces the hydraulic fluid out the service
port to a reservoir. Blowout plugs could be designed into the
inner tube in such a way as to allow the hydraulic fluid to
enter the internal air chamber as the upper stage strokes, thus
containing the fluid within the strut.
TAIL LANDING GEAR

The tail landing gear sho~m in Figure 9 is designed to the

normal operating requirements and to the crashworthy require-
ments of a 20-fps sink speed. At 20 fps sink speeds, the shock
strut of the tail gear is capable of absorbing 12% of the total
energy of the aircraft within 12 inches of strut travel, which
prevents the fuselage from contacting the ground. KRASH
program data indicates that the total strut travel at 36 fps
and 42 fps crash impact conditions is 24 inches, resulting in
the deformation of the tail gear and tailcone structure. The
combination of the two deformations results in 12% energy
absorption capability distributed over a larger area.

The location for the tail landing gear was chosen in order to
minimize the strut stroke when preventing fuselage contact with
the ground at 20 fps sink speed conditions and to provide
adequate rollover characteristics. This location also allows
the landing gear to be retracted forward and to become com-
pletely enclosed within the tailcone, thus maintaining smooth
airflow through the ducted fan. From this retracted position,
the combination of gear weight and airflow would provide a
positive release and extension to a down and locked position
upon actuation of the emergency uplock release system. A
hydraulic system powers the actuator with uplocks and downlocks
which controls the retraction, extension, and kneeling of the
tail landing gear. An electrical indicating switch system is
also incorporated to indicate gear position.
The air-oil shock strut is a cantilever-mounted type designed
to absorb energy at sink speeds up to 20 fps. The strut
assembly consists of a steel piston and fork insidt an aluminum
alloy trunnion that mounts to the airframe structure. Separat-
ing the air from the hydraulic fluid is an aluminum alloy
floating piston. The assembly also contains a variable orifice
assembly. Mounted to the piston and fork is the axle, wheel,
and tire. The drag brace assembly connects the trunnion to the
airframe structure and pivots at the hydraulic actuator mount-
ing point for retraction and kneeling. Both sections of the
drag brake assembly are made from 7075 aluminum alloy. Stan-
dard fasteners and self-lubricating bearings and bushings are
used throughout.

For a nonretractable tail landing gear, the configuration and

components are the same as those fur the retractable concept
except for the removal of the hydraulic and electrical systems.

19
The hydraulic actuator in this concept would be used to kneel
the landing gear by means of an external hydraulic power
source. Smooth air flow through the ducted fan can be main-
tained by positioning a fixed statoi that supports the shroud
in line with the fixed tail gear.

7 Aircraft

Nr

Hydraulic Actuator
(Locked,
, Kneeling)

"-----" --- iBrace

___ - . ... /Shock

Locked___
retracted
Kneeled Position
WL 56.9"'".,
Comn-pressed Position.+ +.%41.9
_--W

Static Ground Line - WL 35 150--- Extended

Sta. 450

Figure 9. Baseline Tail Gear, Retractable

"20
 DRAG BEAM -MAIN GEARS
The trailing arm (or drag beam) geometry as shown in Figure 10
is used for all concepts of the main landing gear. In addition
to providing an attachment point for the lower stage shock
strut, it also mounts the axle, wheel, tire, and brakes. A
jack pad conforming to MIL-STD-809 and configured in accordance
with MS33559 is also incorporated and located to allow posi-
tioning of a jack under the drag beam with the tire flat.
Material of the drag beam for both normal and crashworthy
designs is 300M steel with the section thicknesses being the
same for the normal and crash conditions. To mount the drag
beam to the airframe attachment fittings requires the insertion
of the drag beam into the fitting at BL 10 and positioned 90
degrees to the ground line. A bayonet fitting on the end of
the beam locks into the airframe fitting when the beam is
rotated 70 degrees to a static position. The drag beam is
designed such that it can react the bending, torsional, and
side loads imposed upon it during normal and crash impact
landings.

BL 10
REF
BEARING JOURNALS

5 BL24
F REF

BRAKE MOUNTING
_ SHOCK STRUT FLANGE

ZBAYONET FITTING AXLE JACK PA

END BL SUPPORT
34.25

Figure 10. Drag Beam, Main Landing Gear

21
 LOADS ANALYSIS

DESIGN MASS PROPERTIES

The basic design gross weight is 10, 000 pounds and is employed
to develop design mass properties. The CG (center of gravity)
for the conf iguration is at fuselage station 288. 0 and the
corresponding moments and products of inertia are shown in
Tables 2 and 3.

MAIN ROTOR PARAMETERS

The aircraft assumed for this study is a modification of the

Sikorsky LHX air vehicle. The ABC main rotor is used and its
parameters are shown in Table 3.

LANDING GEAR PARAMETERS

Main Gear

The three baseline design configurations for the selected

aircraft landing gear are as follows:

1) A retractable gear designed to normal design operating

conditions (taxi-ing, ground handling, obstructions,
operating sink speeds, etc.).

2) A retractable gear designed to crashworthy requirements.

Plastic deformation of the gear and mounting system is
acceptable in meeting the,-e requirements.

3) A nonretractable gear designed to crashworthy requirements

as stated in paragraph 2 above.

The load-deflection curve for the main gear tires is shown in

Figure 11.

Tail Gear

The baseline tail gear design is a typical, retractable landing

gear. Tail landing gear is designed to 20 fps landings.
Figure 12 shows the load-deflection curve for the tail gear
tire.

22
 * c 41

14-
ci,
0
OK
~
,,-
.: O .
Q)ZNN4 CO
t LOCs 4)=

- 'L4 J I

0- w w
m : LI) X IN (1) N 04i
w C---o -o C 0cocoN N-

-0
0-
LI0-, ~ N C-
40
0 LI)')
IN~ 0 4 ri 0 ' N
2 <r 0 oNl o c

E-4 <ii N IN'N N

1,- rw 0~ co0 L
(1

ci- M 0
IN4C N-cv 0

4W

W. 11 a _j > r A.x t

0 (DC-flNX4I)z z _ ( z
4
I-) r-L CD
4-4 ~ ~ 0 'O0~ U) O(0N
Mi-i CO .J0 I ..
4-J I ci: ' CIN' tN I-I
E-4 Z 0- Z.0

U,) LA.I0.
Hl 13- CA 0 L
CC 40 L6
rCL CJ o 0 U)00 Lo) C N
4a. 4l 0 W -

I-)4 O N" 0 (2,NWX 0

(IIL NIT)WOO I-_4 O

0.: 4I)OO 0i>4-

JC0 0 1:1 -4r)1

%tI.L 0) U) 17 0!. 1
0' t I-

c0O NC'li CO4 C) r ri

23)
TABLE 3. AIRCRAFT PROPERTIES AND PARAMETERS

Weight-Lb i0,000.
Center of Gravity
XCG-in 288.0
YCG-in 0.0
ZCG-in 90.0

Moments of Inertia
Ix - lb-in-sec 2 38207.

IY - lb-in-sec 2 173924.

Iz - lb-in-sec 2 174969.

Ixz - lb-in-sec 2 11013.

MAIN ROTOR
Station (hub center line) 300.0
Buttline (hub center line) 0.0
Waterline (hub center line) 132.0

24
 4-7

3.0

2.5-

z 2.0-

z
0

w
-J
LL
w 1.5-
0
w

1.0-

.5

0
0 2000 4000 6000 8000 10000 12000

LOAD LB

Figure 11. Main Gear Tire Load Vs. Tire Deflection

25
 CD

1 %

00

E-4

'44

NI3diN0103-130

26o
NORMAL OPERATING CONDITIONS

Static Design (Landing)

Landing loads are developed for two categories of normal
landing conditions. Limit conditions are done at a sink speed
of 10 fps and a forward velocity of 50 knots. Reserve energy
conditions are done at a sink speed of 12.25 fps and a forward
velocity of 40 knots. Both limit and reserve energy conditions
are done at the basic design gross weight range using two-
thirds rotor lift during landing. The landing gear is designed
according to requirements set forth in MIL-S-8698, as well as
additional design criteria prescribed in specification ANC-2.
A landing gear vertical extension position of two-thirds from
fully compressed serves as the application point for determin-
ing maximum vertical reaction while a nine-tenths extended
position is used to evaluate spin up and spring back. Normal
landing requirements and the calculated ground loads to meet
the requirements are presented in Table 4.
Static Design (Ground Handling)
Loads data for braking, turning, pivoting, taxiing, towing, and
jacking conditions are developed in compliance with MIL-A-8862.
In every case, the gear static position serves as the load
application point. Ground handling conditions are done at the
basic design gross weight with the rotor lift equal to zero. A
list of the ground operations requirements and resulting loads
are presented in Table 5.

Normal Landing Requirements

For normal landings up to and including the reserve energy
condition, the energy absorption system depends largely on the
landing gear stroking and tire deflection to attenuate the
impact loads without permanent structural deformation or
malfunction of the landing gear. In this case, the airframe is
idealized as a rigid body and emphasis is placed on modeling
the landing gear characteristics. The actual KRASH model used
in this study is illustrated in Figures 14 and 15. An indepth
description of the mechanics of the KRASH program is provided
on page 33. For normal landing conditions, the KRASH program
is used to evaluate gear stroke and efficiency. Associated
values
Sikorskyof developed
stroke andprogram
efficiency are thcn used as ininputs
of energy equations orderto toa
develop specification required landing and ground handling
conditions. Loads corresponding to normal operating require-
ments are presented in Table 4.

27

""k -
 CRASHWORTHY DESIGN REQUIREMENTS
Crash conditions are evaluated for landing impacts of up to 42
fps using one-G rotor lift at basic design gross weight. For
vertical velocities up to 20 fps, the landing gear has been
designed to attenuate the total impact energy while disallowing
any fuselage ground contact or yielding of the airframe struc-
ture. This criteron has been met for simultaneous fuselage
angular alignments of 10 degrees roll and +15 to -5 degrees
pitch. The landing gear and mounting system, however, may
experience permanent deformation. For vertical velocities
above 20 fps and up to 42 fps, yielding of airframe structure
is acceptable. Envelopes of sink speed vs. fuselage angular
alignment with the ground, for which the crashworthy landing
gear system has been evaluated, are depicted in Figure 13.

Further design requirements for the crashworthy gear system are

as follows:
- Any gear upon crash impact shall not protrude into
occupied and/or fuel cell areas.
- A retracted gear shall provide some crash impact
energy attenuation.
- At vertical sink speeds of 20 fps to 42 fps, the
landing gear system shall provide for a limited 15%
reduction in cabin space.

:-4
7.4

28
 00 H~ 0 0o

LOL

E-4.

r- N
E-4 rX4 H
0 W4 >.
UzH0 0 .0 000

144- co ItnH.4 op
qt1, - N

44 0' 0 0 % -0
4- 00
~-H i-I H 01) - 0-H

H .QH
H 0
C) r-000 mc ,%

X- LO qc M- 0 CYO L 0) Q4
w "I C 4 -q HH HO)-

0H 0 E4
-.-- (0 0 (

ur4 U 294J
 (.- 0 N .t-t- t--
N C,0
N 4 -4 ) i Ir 1

L0 0 0D N00

0 10No

-, 0
zN

0 01 0 O0

53 N (P'D

C, 0 0 0'~D~

~ 0 .40000

0- .(

41r o a) O Co
D 0

00

o e

-4

V 0- CC
'-4 ) 3

);14 > 4
V~O.)i ~ E0 H

C3
 The light experimental helicopter (LHX) meets the objective of
providing a high level of protection for its occupants in
severe crash impacts with advanced landing gear systems and
fuselage energy absorption capability. These impacts, speci-
fied in MIL-STD-1290(AV) (Modified) "Light Fixed and Rotor Wing
Aircraft Crashworthiness", include impacts at 42 fps vertical
velocity with 25 fps longitudinal velocity. The fuselage
structure and landing gear has been demonstrated to protect the
occupants and their living space in vertical crash impacts when
roll angles are limited to 10 degrees or if the impact is at 36
fps with a 20-degree roll angle. The landing gear also meets
* crashworthiness requirements of decelerating the aircraft at
normal gross weight from an impact velocity of 20 fps onto a
level, rigid surface without allowing the fuselage to contact
the ground.
Crashworthy seats provided for both crew and passengers are all
equipped with improved restraint systems. The crew seats meet
the crashworthiness requirements of USARTL-TR-79-22A "Aircraft
Crash Survival Design Guide", and the troop seats shall be
designed to meet the above requirements also.
The crash impact conditions investigated using the KRASH
computer program are impacts on level rigiu jround. The basic
requirement of vertical impact design conditions envelope is to
demonstrate the capability of the aircraft to withstand verti-
cal impacts of 42 fps without either a reduction of cockpit or
cabin height of more than 15 percent or causing the occupants
to experience injurious decelerative loading. The envelopes of
pitch and roll angle for both the 42-fps and the 36-fps verti-
cal velocity are included in Figure 13. The crash impacts
investigated by the KRASH program for the baseline landing gear
designs are designated by an asterisk (*). For all these
impact conditions 25 fps longitudinal velocity has also been
included.
Tae KRASH computer runs for the design conditions are identi-
fied in Table 6.

N'. 31
 I,J, -

'7jBLE 6. DESIGN CONDITIONS FOR CPASH IMPACTS

IMPACT VELOCITY ROLL PITCH

DESIGNATION VERTICAL LONGITUDINAL ANGLE ANGLE
(fps) (fps) (deg) (deg)
LHX 11 42 25 0 0

LHX 03 42 25 10 10

LHX 12 42 25 5 15

LHX 13 42 25 10 -5

LHX 17 36 25 20 10

LHX 18 36 25 10 -10

LHX 19 36 25 10 20

LHX 30 20 - 10 15

LHX 31 20 0 0

PITCH ATT. DEG. FPS

"42 FPS
V0

ROLL ATT. DEG

-20 -15 -10 -5 5 10 .15 20

"Figure 13. Pitch and Roll Attitudes in Crash Impacts

-- ..-
- -5

32
 KRASH ANALYSIS

KRASH PROGRAM INPUT

The program requires the definition of a system of lumped
masses, structural beams, springs and node points that simulate
the mass distribution, the structural strengths and stiffness,
the landing gear, and fuselage underside crushing character-
istics as shown in Figure 14. Special means of simulating
seated occupants are included in the KRASH program. A 50th
percentile cockpit occupant has been simulated in the LHX KRASH
model. The final model for advanced landing gear study is
shown in Figure 15. The model is comprised of 15 masses, 15
springs, 16 beams, and 19 node points.

KRASH MODEL MASS PROPERTIES

The crash impact evaluation was conducted with the helicopter
at its basic structural design gross weight of 10,000 pounds
and a representative center of gravity at Fuselage Station
288.6 in., Butt Line 0.0 in.,and Water Line 94.7 in.

Mass properties for each of the 15 lumped masses and their co-
ordinates X (Fuselage Station), Y (Butt Line), and Z (Water
Line) are presented in the Appendix.

In order to more accurately represent the structural members of

the airframe and crushing of fuselage, use has been made of
node points. These are points in space, each related to a
specific mass, and which move with the mass in a fixed relation
to the centroid of that mass.
The masses, with associated node points, and the coordinates X
(Fuselage Station), Y (Butt Line), and Z (Water Line) of the 19
node points are presented in the Appendix.

The sixteen (16) beams of the model, the masses and node points
that they join, and their structural characteristics are
included in the beam data given in the Appendix.

The material properties included in the beam data are coded

with material code (MC) numbers. The properties associated
with each material code number are shown in the Appendix.

F
'.3 33
UU

0.0

m)

ci~

>'
2. 34
 C.)

LO 0

N I

a.

LL Wl 7 4

N-j

LN 0

> cc'

N(
UV L L' 17~
TI I 9.I
~ '.. i ~ ,i ~ ~ ..I,-
*.. ixi'.- Z<I i :i A~ l R I K 1 -'2 . I W2

KRASH MODEL SPRING PROPERTITS

The KRASH program requires that those portions of the aircraft

that come into contact with the ground be represented by springs.
The LHX model has 15 springs, each having stiffnesses (load-
deflection characteristics) representative of the portion of the
landing gear or fuselage underside that they simulate. The
springs define a combination of linear and nonlinear behavior.
Data which describes the load (F) and the deflection (S) charac-
teristics of crushable structures are used. The deformation of
crushable structure is such that a region is reached where the
confined crushing is very significant and the stiffness increases
substantially.
A typical load-deflection curve is shown in Figure 16; it repre-
sents a combined load-deflection of three different stiffness
characteristics in series.

SF 2
BOTTOMING

2j2 0. I LB/I
5I I j
w III

S1 SA SB SF
DEFLECTION - IN.

Figure 16. Typical Load Vs. Deflection Curve

The spring data, defining their load-deflection curves, is in-

cluded in the Appendix.

36
 It should be noted that the friction coefficient of 0.34 has
been used for all the spiings to represent deflated landing
gear tire and fuselage crushing.
In the KRASH model, the main landing gear simulation includes
the air-oil upper and lower oleo struts and the articulation of
the gear that occurs during gear stroking. The tail gear has
also been modeled as air-oil strut up to 20 fps vertical
velocity impacts. Beyond 20 fps it has been modeled as a
vertical spring.
The oleo input data given in the Appendix are used in the
KRASH model.
The pilot seat and the 50th percentile occupant have been
modeled in the KRASH analysis. The stroking seat has been
represented by a nonlinear member which deflects elastically
0.75 inch then strokes at 14.5 'g'. The occupant and seat were
included in the analysis to develop the seat stroking distance
and then determine whether the seat had exceeded a maximum
stroke of 12 inches.
LIFT FORCES

Masses in the KRASH model have lift loads applied to them equal
to their own weight. Thus, the vertical impacts are represent-
7 ative of impacts at constant velocity.
F
KRASH PROGRAM RESULTS
The information on the dynamic effects of a crash impact
available from a single KRASH computer program run results in
600-800 pages of printouts. Selected data from nine impact
conditions studied for the design of the landing gear are
summarized below.
TimE of Zero Vertical Velocity at Aircraft C.G.
The variation of beam -ads, inertial forces and fuselage
crushing is of significance until the vertical velocity at the
aircraft center of gravity has dropped to zero. The time at
which this occurs for each of the conditions studied is in-
cluded in Table 7. Any loads developed by the KRASH program
beyond the time of zero velocity is the result of the crushed
structure still behaving as elastic structure. This behavior
then causes excessive rebounding and additional loads.

"37
 C 00 00 0

04

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H 'LI) 00 N-

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38
 Landing Gear Strut Maximum Stroke and Time of Occurrences

The main landing gear for the advanced technology helicopter

consists of a two-stage air-oil oleo system. The lower stage
is capable of stroking 14.50 inches while the upper stage
strokes 12.5 inches. For normal landirg and vertical. impacts
up to 12.25 fps the lower stage is used for the energy absorp-
tion. For a crash impact up to 20 fps both the upper stage and
lower stage stroke with one set of orifice openirgs. For
vertical crash impacts up to 42 fps both stages stroke dgain
with added orifice areas. The maximum strut stroke and the
time of occurrences are shown in Table 8. The time of occur-
rence is the time before the pistons "bottom" in the cylinders.
Load data beyond the time of maximum stroke or bottoming is not
considered. The time at maximum stroking given in Table 8
corresponds to the time of zero vertical velocity at the
aircraft C.G. shown in Table 7.

4.'

39
 TABLE 8. SUMMARY OF STRUT ACTIONS

KRASH Time at Max Max.

Conditions Oleo Strut Stroke (Sec) Stroke(in)

LHX1l R.H. Upper 0.09 11.5

Velocity Vert. 42 fps L.H. Upper 0.09 11.5
Long. 25 fps R.H. Lower 0.1 14.59
L.H. Lower 0.1 14.59
00 Roll, 00 Pitch Tail Gear .1 23.34

LHX03 R.H. Upper .17 8.3

Velocity Vert. 42 fps L.H. Upper .16 11.1
Long. 25 fps R.H. Lower .16 14.01
L.H. Lower .16 14.54
100 Roll, 100 Pitch Tail Gear .1 26.64
LHX12 R.H. Upper .21 10.1
Velocity Vert. 42 fps L.H. Upper .20 11.06
Long. 25 fps R.H. Lower .20 14.28
L.H. Lower .20 14.51
5' Roll, 150 Pitch Tail Gear .1 26.2

LHX13 R.H. Upper .11 10.21

Velocity Vert. 42 fps L.H. Upper .10 11.21
Long. 25 fps R.H. Lower .10 14.4
L.H. Lower .10 14.6
100 Roll, -50 Pitch Tail Gear .14 15.16

LHX17 R.H. Upper .17 3.26

Velocity Vert. -CG fps L.H. Upper .17 9.32
Long. 2S fps R.H. Lower .19 9.79
L.H. Lower .14 14.05
200 Roll, 100 Pitch Tail Gear .12 23.85
LHX18 R.H. Upper .12 7.43
Velocity Vert. 36 fps L.H. Upper .12 10.89
Long. 25 fps R.H. Lower .12 13.56
L.H. Lower .10 14.33
100 Roll, -100 Pitch Tail Gear .19 15.5
LHX19 R.H. Upper .26 6.55
Velocity Vert. 36 fps L.H. Upper .27 10.56
Long. 25 fps R.H. Lower .27 13.27
L.H. Lower .25 14.33
100 Roll, 200 Pitch Tail Gear .11 21.5

40
 TABLE 8. SUMMARY OF STRUT ACTIONS (Cont'd)

KRASH Time at Max Max.

Conditions Oleo Strut Stroke, (Sec) Stroke (in)
LHX30 R.H. Upper .384 .98
Velocity Vert. 20 fps L.H. Upper .42 6.25
R.H. Lower .44 9.12
L.H. Lower .384 13.28
100 Roll, 150 Pitch Tail Gear .108 11.8

LHX31 R.H. Upper .22 4.13

Velocity Vert. 20 fps L.H. Upper .23 4.16
R.H. Lower .18 12.59
L.H. Lower .18 12.59
00 Roll, 00 Pitch Tail Gear .12 11.68

GROUND LOAD

The landing gear structure has been designed for the dynamic
load obtained from the KRASH analysis. Typical time histories of
ground loads are shown in Figures 16, 17 and 18. As the gear
articulates, the ground friction causes drag and side loads on
the gear. The ground friction coefficient of 0.34 has been
used in the analysis. Ground loads for the landing gear system
are shown in Table 16.
Figure 17 shows the main gear drag load at the ground as a
function of time. At zero time, the tail gear has contacted
the ground and the helicopter is pitched forward. At 5 milli-
seconds, the main gear contacts the ground. Between 50 and 60
milliseconds an aft drag load is developed as the main gear
articulates vertically and aft. As the assembly is moving aft,
friction between the tire and ground develops which produces a
forward acting drag load. The peak drag load is developed at
approximately 70 to 80 milliseconds after the tail gear has
contacted the ground. The next peak drag load occurs at
approximately 179 milliseconds, which is 10 milliseconds after
the helicopter's vertical velocity at the 6.6 is zero as shown
in Tables 7 and 8. Load developed by the KRASH analysis are
not considered after the C.G. velocity is zero.

41

- - - - -
Figure 18 shows the main gear side load at the ground as a
function of time. The helicopter is assumed rolled to the
right. At approximately 60 milliseconds after the tail gear
has contacted the ground, a side load on the right main gear is
developed to the left. The aircraft is now rolling to the left
and a large side load is developed in the opposite direction at
90 milliseconds. Lateral rebounding then develops until the
helicopter's vertical velocity is at zero.

Figure 19 shows the vertical load of the main right gear at the
ground as a function of time. Again, 50 milliseconds after
tail gear contacts,
vertical loads. the main gear
At approximately 80begins to develop the
milliseconds, maximum
peak
load is obtained. The helicopter is rolling to the left and
the left hand gear contacts the ground. Again load data beyond
160 milliseconds is not considered.
Peak ground loads for the landing gear system are shown in
Table 9. The peak loads, drag, side and vertical, are over a
10 to 20 millisecond time period since the time response to
impact loads of all the materials in the gear vary. The loads
of KRASH condition LHX03, for example, were obtained from
Figures 17, 18, and 19 at 10 milliseconds for the vertical
load. A similar method was used for the other KRASH conditions
of Table 9.

42
 TABLE 9. GROUNE LOADS DATA

MAIN GEAR
KRASH Drag Load Side Load v7ertical Load
Conditions (+ Aft) Lb (+ Left) Lb (+ Up) Lb

LHXII
Velocity Vert. 42 fps -16,000.0 -5,000.C 55,000.0
Long. 25 fps
00 Roll, 00 Pitch

LHX03
Velocity Vert. 42 fps -20,000.0 -15,000.0 60,000.0
Long. 25 fps
100 Roll, 100 Pitch

LHX12
Velocity Vert. 42 fps -22,000.0 -10,000.0 70,000.0
Long. 25 fps
50 Roll, 15' Pitch

LHX13
Velocity Vert. 42 fps 21,000.0 2,000.0 62,000.0
Long. 25 fps
10' Roll, -5' Pitch

LHX17
Velocity Vert. 36 fps -6,000.0 -9,000.0 34,000.0
Long. 25 fps
200 Roll, 100 Pitch

LHX18
Velocity Vert. 36 fps 11,000.0 1,500.0 32,000.0
Long. 25 fps
100 Roll, -100 Pitch

LHX19
Velocity Vert. 36 fps -1,000.0 -4,500.0 46,000.0
Long. 25 fps
100 Roll, 200 Pitch

TAIL GEAR
LHX30
Velocity Vert. 20 fps -6,580.0 -1,390.0 20,015.0
100 Roll, 150 Pitch

LHX31
Pitch 20 fps
00 Roll, 00 Vert.
Velocity 5,880.0 0.0 17,300.0

43
 Lii- T ii , -r I iL, -1 - _o--~
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INERTIAL LOAD FACTORS

The mass number 3 in KRASH analysis represents high mass items.

This mass includes transmission mass, rotor mass and part of
the engine mass. Structures to retain these masses are de-
signed to the inertia load factors obtained from the KRASH
analysis. Since all these high mass items are located away
from the living space, in a severe crash impact it is very
unlikely that these masses will penetrate into the living
space. The maximum vertical load factors experienced by some
of the representative masses are surmarized in Table 10.
Typical time histories of accelerations of high mass item 3 are
shown in Figures 20 and 21.
Figure 20 is a plot of the vertical acceleration at high mass
item 3 as a function of time from the iime of initial ground
contact at a level attitude and 42 fps vertical, 25 fps forward
velocity (Cond. LHX 1). As the vertical ground load on the
gears are developing, vertical accelerations on the high mass
items are also being developed. The peaks in the acceleration
plot are caused by the structure flexing.

Figure 21 is a plot of the force and aft accelerations at high

mass item 3. This is due to the forward velocity at ground
contact which is driving the wheel aft, then friction loads
develop. This motion of the wheel causes a reversal of the
fore and aft accelerations.

Table 10 is a summary of vertical load factors at the cockpit

cabin, fuel for all areas, and the transmission for each crash
condition specified.

47
 H 0 0 0 0 N N- m N- m0 c
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m N H N
Nl 11 N II II
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-
 CONDITION LHX I

0
-10-

0-3

-20x
0 20 40 60 80 100 120 140

TIME - MS

Figure 20. Vertical Acceleration Time History

on High Mass Items

49
 CONDITION LHX I
i~i--
15-

y~-. 10-

0
,,\, 15

x
.

-10-

0 20 40 60 80 100 120 140

TIME-- S

k 5.0

[71 Figure 23. Longitudinal Acceleration Time History

on High Mass Items
. "2- 5
"SEAT STROKING
The vertical stroking capability of the cockpit seat incluled
in the KRASH model is simulated by beams that yield at a
constant load after a small initial elastic deflection. The
seat cushion has been represented as a soft elastic member, and
a 50th percentile man with a Dynamic Response Index (DRI) spine
has been used as an occupant. The seat is capable of stroking
12 inches at 14.5g. The seat, required in the 42 fps impact,
strokes 9.98 inches, which is consistent with the cockpit
capability. To find the level of protection provided to the
occupant, a comparison between KRASH floor pulse and seat
design pulse has been done. In all cases the velocity change
is less than 50 fps. A typical floor pulse with superimposed
seat design pulse is shown in Figure 22. Table 11 summarizes
the cockpit seat stroking data. The 9.98 inches of seat
stroking occurs during a 42-fps sink speed, 25 fps forward
speed, rolled 20 deg and pitched -5 deg, as shown in Table 11.
The stroking begins .0794 second after ground contact.
Maximum stroke is obtained .180 second after ground contact,
as shown in Table 11. The time of stroking is .180 - .0794 =
.100 second.

51
 I- N

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52
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53
-,oL

FUSELAGE CRUSHING

The amount of crushing of the fuselage underfloor structure is

predicted ly the KRASH computer program in terms of deflections
of the springs used to simulate the understructure crushing
I, characteiistics.
The frames and beams below the floor and fuel cells are con-
structed of graphite/epoxy. G-aphite/epoxy tape and fabric are
used in the frames and beants to react primary airframe loads.
The depth of the structure is approximately 7-3/4 inches.
Figure 23 is a sketch of the subfloor structure.

Kevlar epoxy could be used as a crushable structure and graph-

ite epoxy as primary structure, as shown in Figure 23; however,
a lower weight structure was obtained with full depth, vertical
sine wave graphite epoxy webs.

Each KRASH computer analysis contains a summary of spring

"loading and unloading, including maximum forces and deflec-
tions.
Table 12 is a summary of the understructure crushing of 12
springs. The locations of these springs in the airframe are
shown in Figure 15. The amount of fuselage crushing predicted
in two crash impact conditions is illustrated in Figure 23. At
42 fps level impact fuselage, crushing is less than 3-3/4
inches. When the crash zone cxceeds 3-3/4 inches, the graphite
upper structure will crush locally to a point beyond which
there will be local structural failure. In a 36-fps, 20-degree
roll,and 10-degree pitch condition, before the crush line goes
above the floor, there will be secondary structural failure of
lateral beams, "parallelo-gramming" of the fuselage, and the
remainingi structure will withstand the crash loads. Thus 85%
of the li.ving space shall be maintained in a severe roll
impact.
Figure 25 presents the results of crushing tests conducted on
beam members with webs of Kevlar epoxy or aluminum. From
Figure 25 a 4-ply Kevlar sine wave web is shown to have crush-
ing capabilities (load deflection) which are better than
aluminum webs or 4-ply Kevlar webs with beaded stiffeners.

54
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58
 PRELIMINARY STRUCTURAL WEIGHT ANALYSIS
A preliminary structural analysis was performed early in the
study for each of the three baseline landing gear systems. The
analysis was performed during the development of the KRASH
model to provide model input data and establish baseline
weights.

STRUCTURAL LOADS (MAIN GEARS)

The noncrashworthy main landing gea_ was analyzed for the Main
Gear Obstruction Ground Loads shown in Table 4. The shock
strut load was 20,083 pounds. Figure 26 presents the results.
Maximum bending moment on the drag beam was 460,000 inch-
pounds. Figure 27 presents the results and compares the
noncrashworthy drag beam with a crashworthy drag beam.
The shock struts and drag beams were sized based on hand
calculators. Loads were calculated for the crashworthy compo-
nents assuming the critical design condition was at 42 fps
vertical, 25 fps longitudinal, and 10 degrees roll and pitch.
The shock strut load developed was 50,390 pounds. Figures 28
and 29 present the results of those calculations. Maximum
bending moment on the drag beam was 887,881 inch-pounds.
Figure 27 presents the results.

The KRASH program was run using the data of the preliminary
crashworthy landing gear designs. The program developed design
data such as ground loads, shock strut loads, shears, and
moments in the drag beam, load factors, aircraft attitudes and
gear geometry during a crash sequence. Table 13 presents the
main gear ground loads, also given in Table 9, and the result-
ing shock strut loads. It can be shown from Table 13 that the
original hand-calculated strut load of 50,390 pounds was within
the range of loads obtained by the KRASH analysis. Maximum
load was 55,000 pounds, minimum of 47,900 pounds from the KRASH
analysis. The loads from the KRASH analysis substantiated the
original baseline design of the main gear.

59
 TABLE 13 GROUND LOADS AND STRUT LOADS - MAIN GEAR

KRASH Drag Load Side Load Vertical Strut

Conditions (+ Aft - (+ Left - Load Load

I.' ~PLHXtc
Lb) Lb) (+ Up-Lb) (Lb)

Velocity Vert. 42 fps -16,000.0 -5,000.0 55,000.0 47,900

Long. 25 fps
0' Roll, 0' Pitch
.LHX03
-' Velocity Vert. 42 fps -20,000.0 -15,000.0 60,000.0 48,900
Long. 25 fps
100 Roll, 100 Pitch

LHX12
Velocity Vert. 42 fps -22,000.0 -10,000.0 70,000.0 54,000
-- Long. 25 fps
Roll, 150 Pitch S5

LIA13
Velocity Vert. 42 fps 21,000.0 2,000.0 62,000.0 55,000
Long. 25 fps
100 Roll, -5 Pitch

Average Strut Load 51,450

- -1

60
 2008320083
(LBS i(LBS)

095

A 9416 A- 1222
I= 1173 1- 25638
30= 1116 419 144b
3250
D/i= 342 D/t- 441
10 do 7075-T73
F SECTION A-A
7075-T73
SECTION B-8

"Figure 26. Shock Strut - Noncrashworthy, Kneeling,

Retractable Main Landing Gear

K4

ii-61
 o 4 o
04-

, U, I d

z < .
. 62

:.m
nw" ; I-

62.
5q390 :16,9
f
(LB I

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08-A1 1 8645 2 7 A= 106 4
22(A2 96
50 444 0
p 112
1082th
~
320 07} D/r-3736 416 4007~'O 5-T73 ALUM

H
T180.0004340
4340 H T 2600000 SECTION G-G

SECTION D-D
SECTION C-C 087
D/t- 37 36
A- 8645

P10822

SECTION E-E

Figure 28. Shock Strut Assembly, Crashworthy, Retractable,

Kneeling Main Landing Gear

63
 -1"X7'LL-w w U-Y 17;l-: VW'U V V ( C ~ '~7 '"L ~ ~ ~ ) ~21 L-%4-' L'W -- o I 1 . 1-6

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64
 TAIL GEAR

The tail gear was analyzed for a level 20-fps sink speed and
normal landing conditions. The trunnion cylinder and drag
brace were analyzed for the 20-fps crash condition. The axle
and piston/fork were analyzed for the normal loads. Figures 30
and 31 present a summary of the tail gear structural analysis.
BASELINE LANDING GEAR WEIGHT

Table 14 summarizes the weights of the components for each

baseline landing gear. The weights for wheels, tires and
brakes were obtained from catalogs. Weights for the shock
struts, axles, drag beam, fittings, hardware and oil were
calculated from drawings. Table 15 compares the baseline
landing gear weight with the landing gear systems of other
Sikorsky helicopter mcdels.

The percentage of gross weight for the landing gear system of

the noncrashworthy desagn compares well with the percentage for
a noncrashworthy commercial S-76A design and a noncrashworthy
design for the HH-3E military helicopter. The crashworthy
designs of this study appear reasonable when compared to the
UH-60A. The crashworthy design for the ACAP (S-75A) appears to
"be out of line. The larger weight for the crashworthy S-75A is
in the shock strut and drag beam which is the result of the
geometry for a main gear aft of the c.g. and attached to the
narrower transition section.

The baseline crashworthy drag strut, shown in Figure 29, was

K:: calculated to weigh 76.6 pounds. This weight appeared too
heavy when compared to the drag strut of the UH-60A shown in
Table 15. A review of the detailed structural analysis of the
UH-60A drag strut showed that the strut had been designed for a
normal ground obstruction condition. It was then assumed that
the wheel, axle, and tire w#ould be destroyed during a crash
onto a rigid surface. The crash loads were then applied to the
A+ strut at the location of the base of the axle. The result was
that the obstruction condition designed the UH-60A drag strut.
Based upon the analysis of the UH-60A drag strut, the noncrash-
worthy drag strut is used for the crashworthy main landing
gears.

65

. - -"
 RETRACTING CYLINDER

ATTACH PINS (1"250 x 095)

S P3.5"N-FNG

7-T36 DIA
47
8 A
DRAG BRACE 410

475 AIA.

PISTON-FORK
b
--B SECION 255

80 ,A S LUGS

t-50

SECTION B-8 125 AXLE (1 492 x 0701 9" LONG

4340 HIT 200.000
-170
100R=130

L 40'90O35

a SECTION A-AE-E
SECTION D--
100O325
C C

100 f 0 30020
250
SECTION C-C12
-r---------SECTION E-E
PISTON -FORK26

300 M H T 280000Fl

SECTION F-F

Figure 30. Tail Gear Structure

66
 '- 36

R LOWER LUGL:

t= .5.6

4J4.

R=~R .625 RETACIN CYLNDE

">".:,,'::' /, UPPER LUGS:

""D-- .5625 RETRACTING CYLINDER

..;::.t7- .44
-''>AIRFRAME (REF.)

[--.180 S
(TYP')

I 7075-T73
,150

-2.0 1-

SECTION G-G (TYP.)

Figure 31. Tail Gear Drag Strut

67
COMPOSITE LANDING GEAR COMPONENTS
Four landing gear components which appeared to be good candi-
dates for weight savings using composite materials were select-
ed for analysis:
a. Main gear drag beam
b. Oleo extension fitting (noncrashworthy gear)
c. Upper fitting, main gear/fuselage
d. Tail gear drag brace
A preliminary analysis of the main gear drag beams for the
three gears resulted in a 16-percent weight savings. A compos-
ite drag beam is sketched in Figure 32.
A weight savings of 45 percent was obtained for the oleo
extension fitting compared to an aluminum fitting.
A composite upper fitting, which attaches the top of the main
gear to the airframe, was estimated to result in a 60-percent
weight saving compared to a steel fitting. The weight estimate
was based on the bearing strength density ratio (FBRU/P) for
125,000 psi heat treat steel and 450 graphite/epoxy.

A preliminary analysis of a composite tail gear drag brace

resulted in a 47% weight savings compar-d to an aluminum brace.
The lower weight savings for a composite main gear drag beam
was the result of providing torsional stiffness to the beam to
prevent excessive "scuffing" of the main gear tires.
A graphite/epoxy main gear wheel was considered but was not
feasible due to the heating of the wheel during braking. A
metal matrix main gear wheel appears feasible, but it is not
clear that the technology will have matured sufficiently to be
captured by an accelerated LHX acquisition strategy.
Although composite
landing gear materials
structures, designappear to must
problems reduce
be the weight The
overcome. of
design of end caps for pistons and cylinders may require stain-
less steel caps to be bonded to composite tube members.
Provisions for valves threaded into the oleo system must be
considered. Possible fiber breakage could cause leakage of the
system due to the high pressures in a cylinder. Repeated high
impact loads on epoxy matrix materials may result in matrix
splitting.

68
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000

c. co
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U, . h .- i
'.4 WEIGHT SENSITIVITY ANALYSIS

A preliminary weight sensitivity analysis was conducted to

evaluate the effect of various sink rates on the components of
a crashworthy system. The system included crashworthy crew/
troop seats, crushable fuselage understructure, and the Wain
landing gears. The weights of seats systems, post-crash fire
prevention, emergency egress, airframe crashworthiness and main
gears were investigated by Sikorsky under a Crashworthiness
Design Parameter Sensitivity Analysis study conducted for the
Applied Technology Laboratory under contract DAAK 51-79-C-0043.
The results of that study were reported in Reference 2. A more
detailed analysis was conducted on the main landing gears to
evaluate the effect of pitch and roll attitudes.
Ci'ASHWrWHY SYSTEMS
At present, creshworthy seats are designed independent of the
landing gear. This is ,lone to provide crew protection should
the heli-opter, for example, crash into the water. Under the
Reference 2 study, weiqnts for crew/troop seats were dex eloped
for various levels of zrashes up to 42 fps sink rates. Table
16 summarizes the results of that study.

The study of crew/troop seat weights showed very small weight

increases when going from 20 fps capabilities to 42 fps capa-
bilities. For the crew seat, the weight was 39 pounds at
20 fps 'nd 12.5 pounds at 42 fps, a change of 3.5 pounds. -For
the troop seat, the change was 7.73 pounds. At 20 fps, the
troop seat was 11 pounds. For the sink rates above 20 fps to
42 fps, the weight remained unchanged at 18.7 pounds. It
appears from the study that very small changes in the weight of
a crew/troop seat is obtained due to large increases in sink
rates. Tae weight chanjes are considered insignificant for this
landing aear study.

Reference 3 concludes that a properly restrained human can

withstand the lateral and longitudinal loads Fssociated with
;urvivable cras1 ' puls(s. Therefore, no load attenuation is
necessary in the.e directions. Based upon that conclusion, the
crew seat can be expected to perform in the crash attitudes
specified for tCe design of the larning gears. Also, since the
seat is designed indepepdent of the landing gear, the seat can
limit the load on the oocupants, with gear retracted, as shown
in Figure 33. Crew seat loads 'nd s..roking distances were ob-
tained from che KRASH program for crash conditions of 25, 30,
.ind 35 fps, level landings, gear retracted. As seern in Figure 33,

72
 the load on the crew seat is limited to approximately 3,000
pounds for each crash condition. The maximum seat stroke
during the crash conditions is 12 inches. The duration of the
seat loads are not injurious to the pilot. The peak load
factor is approximately 15 g's for a 200-pound occupant.
.,

TABLE 16. CREW SEAT/TROOP SEAT SYSTEM WEIGHT

Crew Seat System

Crash Level I II III IV
S-76 SH-3D Lamps
Seat Structure 19.8 lb 30.0 lb 33.5 lb
Cushions 7.0 lb 4.0 lb 3.9 lb
Restraint System 3.2 lb 5.2 lb 5.1 lb
Total 30.0 lb 39.2 lb 42.5 lb

Troop Seat System

Crash Level I II III, IV
CH-53 RH-53 UH-60A
- (Troop) (Crew Chief) (Troop)

Seat 4 8 lb 10.) lb 15.55 lb

Restraint System 1.2 lb .9 lb 3.20 lb

Total 6.0 lb 11.0 lb 18.75 lb

Crasi! Levels I = 0 - 10.5 fps; I I = 10.5 - 21 fps;

III= 21 - 31.5 fps; IV : 31.5 - 42 fps.

-.7

-T

73
N1 - 3MIOHIS IV3So

CL0

U-)
In'4, I

cTn or

I~~~~i I j
00
0L 0

I-A
c4-)

91~~
VX 4JV1
A crushable fuselage under structure is not requited for sink
rates at 20 fps or less based on the requirement of this study.
Beyond 20 fps, the amount of crushing structure depends on the
depth of the structure below the floor and the sink rate. For
this study, the depth of the structure below the floor is
approximately 7 inches. Three and three-quarter inches are
used to provide crushing structure. Four inches are required
for primary structure.
The weight of airframe crashworthiness was evaluated in Refer-
ence 2. Table 17 summarizes the results. For this study, the
weight of the crushable fuselage under structure is considered
part of earth plowing and longitudinal impact structure of
Table 17. It is estimated that 70 percent of the earth plowing
and impact structural weight can be attributed to crushable
structure. The aircraft weight of this study and that of
Reference 2 are comparable.
The KRASH program evaluated the effects on the airframe for
level landings, gear retracted. Figure 35 shows the result of
those conditions Figure 35 indicates that the airframe vertical
load factors (G) at the pilot's floor, and at the high mass items,
exceed the requirements of MIL-1290 (AV) Section 5.1.7.2. The
MIL-1290 (AV) requirement is an ultimate vertical load factor of
20 G's (average). Also from Figure 34 it is apparent that there
is insufficient depth of crushable structure for level 25, 30,
and 36 fps gear retracted crashes. It is noted that the loads
on the pilots seat limit the pilot to approximately 15 G's for
a 200-pound occupant as shown in Figure 33.
TABLE 17. AIRFRAME WEIGHT

Airframe Crashworthiness

CRASH LEVEL I II III IV

1) High Mass Item 0 10 30 35

Retention
2) High Strength/ 0 10 15 15
Equipment Attach.
3) Earth Plowing 0 5 10 15
Capability

4) Longitudinal 0 5 10 15
Impacts

TOTAL WEIGHT 0 30 70 85

75
 11 If

w>w wU > +
CL
<1. < L.
xC D )CD (D ;

00

(14

4-)

C(12

I!I 0)

r.4

CDl

0- IL <I u.
CD 0D
NE

76Y
xi1

~ ~ 4 -
 Vertical load factors for the high mass items (rotor head,
transmission, etc.) were developed by the KRASH program for the
crash condition with the gears retracted. Figure 35 presents
those load factors for a level condition. The 35 foot-per-
second sink rate results in a load factor (67 Gs) which is far
greater than MIL-STD-1290 (AV) requirements of 40 Gs peak, 20
Gs average Sink rates greater than 25 feet per second, gear
retracted, would result in the high mass items breaking loose
during the crash sequence. Load-limiting devices or heavier
structure would be required for the higher sink rate with the
"gear retracted. This requirement is beyond the scope of this
study.
A study of crashworthiness effects with the gear retracted must
include the fuselage structure for attachment of seats, high
mass items, fuel cells, backup structure for the landing gear,
and retention of cargo.

Reference 2 also evaluated the weight trends for crashworthy

fuel systems and main gearboxes. The results are shown in
"Figure 36.

12
F

77
 i~J X~U)
.T.Xd ~1h~J'l2
UVL~'. L' LW '~b\r ~~'~ ~~7' "' 7~ V'. 17 ' ~~ ~ 7' LWIn.

m C*

L) -4
012

10-H
0

II
I..

CO)d

Co ------- aVI-3I3

78-.
 CREW SEAT (ARMORED)

140 *NOTE: EACH LEVEL IS 25% OF 42 FPS

(120) (125 ,(2.)
120 -

100- (.1000
0 (975)
:_I. (960)
" 80 -
r
-
Sm (920)
60

SZ 900 (890)
40- TROOP SEAT

20 (11) (18.7)
(6)
0 11* 111* IV* 8001 1 11 1111 IV

CRASH LEVEL 42 FPS CRASH LEVEL 42 FPS

CREW/TROOP SEATS AIRFRAME

300- 560-
* (255) (550)
(247)
(221) (542)

0 200 0 540(
m50

,.- .I--
_0 0 (526)

: 100 5
520

(504)
0 Ioo I
1I 111 IV 500 1 11 I V
I II III IV

CRASH LEVEL 42 FPS CRASH LEVEL 42 FPS

FUEL SYSTEM MAIN GEARBOX

Figure 36. Crashworthy Systems (Reference 2)

79

LI'"" """""; "/ / " "' :'

Landing gear weights were evaluated in Reference 2 for various
sink rates up to 42 fps.

A more detailed weight sensitivity analysis was conducted for

this study to evaluate the effect of various sink rates at
various angles of pitch and roll on the landing gear shock
strut, drag strut, and fuselage fittings. The other components
of the landing gear system are not designed by crash load
conditions. The KRASH program was run to develop main gear
crash loads at 42, 36 and 30 feet per second sink rates at
combinations of pitch angles and roll angles given on pages 8
and 9. A pitch angle of -7.5 degrees was also used to deter-
mine where the peak loads in the shock strut would occur. The
loads on the shock strut were obtained from KRASH for a sink
rate of 42 fps and the helicopter pitch and roll angles shown
in Table A-1. The loads are plotted in Figure 37 and shown in
Table A-1. A peak load was obtained at a nose down pitch angle
of 5 degrees and a roll of 15 degrees. Shock strut loads were
obtained for sink rates of 36 and 30 fps. Again,maximum leads
were at a nose down pitch angle of 5 degrees and a roll of 15
degrees. The weight of the shock strut is determined based on
the thickness of the structural load carrying components in the
strut. The thickness required to resist the loads imposed are
based on the strength allowables for the materials or on
machine shop limits.

Figure 37 also presents the bending moment and weight of the

drag strut based on the KRASH program. Also shown in Figure 37
is the baseline moment used to design the baseline drag strut.
The KRASH program appears to develop the proper loads in the
shock strut but overestimates the bending moment of the drag
strut. The baseline drag strut was sized based on plastic
bending allowables for steel. The bending stiffn'ss of the
baseline drag strut was modeled assuming various sti..nesses.
It should be noted that the drag strut of the UTI-60A is de-
signed by plastic bending. The current KRASH program does not
analyze a plastic hinge.

80
 rrn J '60~
R 4ruv PS - -
P=- 50-

50 P 7.5r
BASELINEP

" 0- FPSP=150
-1 30
4cT 20 15-2

0 10
<ROLL =150 42 FPS

-5 0 5 10 is
60E=30 20P1 50 -

PITCH ANGLE "- DEG ROLL ANGLE ~- DEG

50 -

0-40
AXIAL LOAD ON SHOCK STRUT VS. PITCH AND ROLL ANGLE
FROM KRASH

60 -E =30 X i06 Psi so5

1 50 -40

J' I WEGH
WEIGHT - BASELINE
I
Z 40 --
LBS. N;O
I-/
z 30
3-- Z
6
E 3 X 10 PSI DESIGNED
0 20 BY
20 NONCRASHWORTHY

"
z
a
mu
10
10

ROLL =150
I
CONDITIONS1

0 0
-5 0 5 10 15 -5 0 5 10 15

PITCH ANGLE -' DEG PITCH ANGLE --DEG

ASSUMES:
PLASTIC BENDING DURING CRASH SEQUENCE; NO FRACTURE
BASELINE DESIGNED FOR GROUND OBSTRUCTION LOAD

Figure 37-., Main Gear Component Loads/Drag Beam Weight

S~81
COMPONENT WEIGHT
The components of the main landing gear system that are af-
fected by changes in loads are the shock strut and fuselage
fittings. Table 18 summarizes the weight of the components in
the retractable and fixed shock strut and those components
affected by load.

The shock strut is an axially loaded structure; therefore, the

weight of the affected components is assumed, in this study, to
be proportional to the shock strut lcads of Figure 37 and the
baseline load of 50,390 pounds. Figure 38 presents the weight
of the shock strut as a function of angle.
The weight of the fittings is assumed to be proportional to the
shock strut load.

Figure 39 summarizes the fixed landing gear system weight as a

function of pitch angle.
A review of the landing gear system weight shown in Figure 39
and Table 13 was performed to determine the factors that cause
the changes in weight. From the earlier study (Reference 2),
the weight of the landing gear was expressed as
^/DwG\8
WLG= 62 .86 DG KL
(100/0)
where DWG is the helicopter design gross weight and KLG the
landing gear weight factor. For various sink rates, the change
in the weight of the landing gear is primarily dependent on the
energy that it must absorb; the weight coefficient is expressed
by
b V2 LD
KLG =a+ 100
where a and b are constants and VLD is the landing gear design
sink speed.

The landing gear weight equation is now expressed as

DGw DGW by 2

82
The first expression gives the weight of the landing gear
components for wheels, tires, brakes and hardware. The con-
stant "all was found to be 0.6861. The second expression gires
the weight of the energy absorbing system and the constant "b"
was found to be 0,013'.. The weight of wheels, tires, brakes,
etc., for this study becomes

W= 62.86 .6861 = 272 pounds

This agrees closely with the weights shown iTi Table 13. For
example,
Noncrashworthy weight 244.9+73.3-26.6-22.2 = 269.4
pounds
Crashwcrthy, retracted weight = 322.0+73.3-92.9-22.2 =
280.2 pounds
Crashworthy, fixed weight = 293.4+73.3-99.4-22.2 = 249.1
pounds

The weight of the energy absorbing system for this study

becomes

= 10000)* .8 1942/0
WE 62.86 000 0139 422/100 = 97.25 pounds

This agrees closely with the weight of the shock struts shown
in Table 13 for the crashworthy landing gears.
It appears that the weight of the energy-absorbing system in
the landinq gear is affected more by velocity than by attitude.
The maximum load in the shock strut would occur when the shock
strut is perpendicular to the ground.

83
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(ad 0 r H r-i 444 1144 Ln

E-4)

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84
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81 -. HO13MW31SA

U) 86
LANDING GEAR SYSTEM COST

The cost of the landing gear system is based on actual cost

data of the UH-60A Lot I, fixed landing gear. Table 19 pre-
sents the cost of that landing gear. Figure 40 presents the
data of Table 19 for the cost as a function of weight.

Figure 41 presents the cost of the fixed landing gear system at

42 fps.

Based on the weight and costs shown in Figures 39 and 41;

trending data was developed for the weight and cost as a
function of sink rate. Tile trend data is shown in Figure 42.
Also shown in Figure 42 is a comparison of this study and the
results of a 11CRASHWORTHY DESIGN PARAMETER SENSITIVITY
ANALYSIS", conducted in 1980, for the SCOUT helicopter (Refer-
ence 2). The SCOUT landing gear system is based on a design
gross weight of 8683 pounds. Increasing the landing gear
system weight by the ratio of the LHX design gross weight
(10,000 pounds) results in a landing gear system weight very
close to this study.

The SCOUT landing gear system cost shown in Figure 42 is

updated to 1984 dollars assuming an annual inflation rate of 5
percent. The SCOUT cost is increased further by the ratio of
design gross weights. The result is a cost which is comparable
to this study.

87
 TABLE 19. LANDING GEAR PARAMETRIC STUDY
(FY '84 $, THOUSANDS)

Cum. Avg. Cost For

Weight 1000 Units $/Lb @ T1000
(lb (Thousands) (Dollars)

271.2 36.70 $135.3

296.2 39.63 134.8
321.2 42.53 132.4
346.2 45.40 Cost = .2793 131.1 $/lb 288.67
371.21 48.24 (WGT .8707 130.0 (WG) - .1348
396.2 51.05 r = 1.00 128.8 r = - .006
421.2 53.85 127.8
446.2 56.62 126.9
471.2 59.37 126.0

NOTES:
UH-60A Calibration Point (based on UH-60A Lot I Actual
Cost Data; 42 fps, fixed gear).
2 Costs represent Labor, Material and Factory Overhead, but
no General OH or Fee ( 1.3 x cost, if desired).
3 FY '84 $, Constant
Methodology used: RCA Price System using UH-60A Actual
Cost Data.

88
 70 -

U)
" COST =.2793 (wgt)**.8707I

0
0

U.

z
so - U./ . . .

0 40- - - - ----------

30
00300 400 500 600

WEIGHT - LB

Figure 40. Landing Gear Parametric Cost Study

54 t--- -- ~~
- -
52

48
" -

-46

S42

o4- I 100
ROLL 15 42 FPS

-5
'--, 0 5 10 15 0 5 10 15
PITCH ANGLE - DEG ROLL ANGLE - DEG

NOTE: COST .2793 (SYSTEM WEIGHT) .8707

Figure 41. Landing Gear System Cost As Function of Pitch

and Roll Angle

89

-- 7
rx,"C)U Wr ~~~~ W-V
-~T' -WL W
L- F7.P - 7V N .

C'J

C4d
an lp

0! LL. a)

vi - 4

> =

0!0

In CO) C1

cmi

0C-4
0 0

U).
- iA
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44
I::I
co)

C3 440

900
 AERODYNAMIC DRAG

An aerodynamic drag estimate was made for the retractable and

fixed/extended landing gear configurations. The power loss due
to drag is 7.2%/ft2. For the extended gears (main and tail)
"the power loss is 26.6%. Retracted, the loss is 0.7%. The
data in terms of square feet of drag area is as follows:
Landing Gear Type
Retracted Fixed/Extended
0.1 ft 2 2.3 ft2
Main Gear
(both sides)
2
Tail Gear 1.4 ft

Total 0.1 ft2 3.7 ft2

+3.6 ft 2
A Drag
(fixed-retracted)
The drag estimate ' for the components of the fixed/extended
main gear are as follows:
Aerodynamic Drag, main gear components
Shock strut = .17 ft 2 /side (15%)
Drag beam = .16 ft 2 /side (14%)
.28 2 /side
Tire = ft (24%)
Brake and axle = .15 ft 2 /side (13%)

Component total = .76 ft 2 /side (66%)

Interference drag
between shock strut
and drag beam = .39 fc 2 /side (34%)
Total drag 1.15 ft 2 /side (2.3 ft 2 main
gear system)

The percentages of drag of the components in the extended tail

gear are similar. The drag estimate for a "clean aircraft" is
11-12 ft2.
The drag estimates have been based on the following assump-
tions:

Aircraft weight 8000 to 10,000 lb

Forward velocity 200 knots
Power 1200 HP

91
The following is noted.
"" The main gear tires, brakes, and axles are the same for
all gear concepts, therefore the drag values will remain
unchanged.
* The outside diameter of the main gear drag beam is not
changed for each concept, therefore the aerodynamic drag
values will remain unchanged.
* The outside diameter of the lower piston of the shock
strut for each main gear concept is not changed, therefore
any changes to the wall thickness of cylinders will have a
negligible effect on the overall drag of the main gear.

The interference drag between the shock strut and the drag
beam remains constant.

92
 RECOMMENDED DESIGN CRITERIA
Based upon the sensitivity analysis the following design
criteria are recommended:
ATTITUDES AND VELOCITIES

A preliminary review of eleven Class A Mishaps with the UH-60

(reference Table 20) indicates that in six of the mishaps the
landing gear stroked at pitch angles between 20 degrees nose
down and 15 degrees nose up, and roll angles ranged from ze..o
degrees to 15 degrees. Sink rates ranged between 18 fps and
90 fps. Horizontal velocities ranged fiom zero to 34 fps.
However, at a 90-fps sink rate the crew and troop seats stroked;
there were three fatalities and no minor or major injuxies. At
a nose down pitch angle of 20 degrees (V = 20 fps, H = 0), the
crew seats stroked; however, the troop seats did not. The result
was three major injuries and no fatalities.

Four mishaps resulted from either very high horizontal veloci-

ties (152 fps) or upside down impacts (pitched or rolled).
One mishap, at 40 degrees nose up, caused the main gear to
partially stroke.
14 The average sink rate of seven mishaps was 44.4 fps, the average
horizontal velocity was 20
fps.
Based on the data of Table 24 and the baseline helicopter
studied, the criteria for attitudes and sink rates to develop
landing gear loads should be as follows:
* The peak design load in the shock strut occurred at a nose
down pitch angle of 5 degrees and a roll of 15 degrees for
all crash sink rates studied. It is therefore recommended
that the criteria of 5 degrees nose down to 15 degrees
nose up be used for crashworthy designs.
The bending movement in the drag beam, assuming elastic
bending, occurred at a pitch angle of 15 degrees and
rolled 15 degrees. It is recommended that the criteria
for a drag beam include plastic bending for roll angles of
0 to 15 degrees.

Based on the loads developed for 42 fps sink rates com-

bined with a longitudinal velocity of 25 fps, the criteria
of 42 fps vertical and 25 fps longitudinal appear reason-
able.

93

---------------------- --------. ~--... ..

 Final LI- helicopter definition will determine the value of
-added landing gear weight versus applying the weight to other
aircraft attributes.

94
 :3 LO C> ,f) W) kn C) I ff-
r4 LO m - r-4 a) Q) 0 a) 0 0)
mLO I I >- >- >- Z -
0) LA)

a_ q 00 Z: miJ 0-

MLO r-4 CY 0 0 0 0 0)

) V) fl 0 I CY)I CY
C.. 0D 0 0 LA ULA a) 0 ci 0 0)

L/)

C: )L OC) w~ a) a) 0 U) 0a) i~)~

a) cv) >- >- 4

L) 0

WdicoC (a U) cC) 00) CO

14 rI t C'4 mv Cl >- CD X: r
C,)

M x C~j v-q 0D 00 C fU c C
0 LA a) 0) Q
00~ -f 0 0 0 a) C- III I(vM
U- Z -

V)0l- C LA LA CD mv a) cu 0 di 0 0)-
LO 0I ? r_ - Z -
'A w
Z -X LO 0: U)
EC)
M: C0 to C0 mv m 0 0 0 a) 0 IIJ

C\J
.. J... J U) (U) W) U)
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I-

v-I 0o
C) - LA LA di) 0 0 d) 0 I- Ir

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anv- Ul) V) (A)
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mU 0) - d) C(p0) +-j (U 41) d41
a) - u
- 4--) C 0F) - c mU d) C U) 4-P(n U) S- CD- Lii
(U C c 0) 0)-- (V VI) 0 (a CL 0 wdiO0( z
U" <' C
c~ _%z V) u) S- 0CZu .- C "-4- LIi
V)f NUV) r- C .- v0 CL. u U S.- .(UMUO
Ca). *-- CLU - 0 S- 3. 0 - 4J*- L~L-
L4- S...4- 4--) 3 C -) di) 0 a) U) >-<'
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% 95
DETAIL DESIGN CRITERIA FOR THE MAIN LANDING GEAR

The wheel should be assumed to remain attached during the crash

sequence. A review of the structural analysis of the UH-60
main gear showed the drag strut to be designed by normal ground
conditions (Reference 4). The loads were applied at the ground
line with the gear two-thirds extended. For the crash condi-
tion it was assumed that the wheel had fractured and the loads
were applied where the axle attached to the drag strut. The
crash drop test conducted on the UH-60 main gear resulted in
half the wheel remaining on the axle (Reference 5) after the
wheel rim contacted the test rig platform. The tire should be
assumed destroyed.
The drag strut should be designed as a structure in plasic
bending during the crash sequence. However, the current KRASH
program cannot handle nor..linear behavior of materials. A 20-
millisecond time period for peak loads does not appear reason-
able based on oscillograph data obtained during the crash drop
test of the UH-60 main gear. A 20-millisecond time frame
results in vertical reactions far less than expected. Using
peak loads at 1 millisecond resulted in a vertical reaction
slightly greater than expected. A 3- to 5-millisecond time
frame appears reasonable.
The KRASH program appears to provide the proper shock strut
loads.

96
 DESIGN UPDATE

The baseline fixed and retracted crashworthy main landing gears

were sized based on an axial load, in the shock strut, of 50,390
"pounds, and the obstruction loads of Table 4 applied to the
drag beam. The preliminary structural analysis for the major
components of the shock strut with margins of safety less than
15 percent is discussed below.

SHOCK STRUT STRUCTURAL ANALYSIS

Shock Strut Assembly: (Crashworthy, Kneeling & Retractable

Gear) Reference Figure 26.

Cylinder Bore 4.00 Dia. A= 12.566

Piston O.D. 3.25 Dia. A = 8.296
4.270
Upper Piston:

D/t = 37.36
A = .8645 Pc = 50,390 lb
I = 1.0822 Press = 50,390= 6074 psi
p = 1.12 8.296

L1/p = .65.77/1.12 = 59

Section E-E Fc allow, = 81,000

Fc = 50,390/.8645 = 58288 Rc = 58,288= .720

"81,000

Fthoop = 6074(3.076) =107,377 Rtn = 107,377 .413

2(.087) 260,000

MS1 1 = +.01
MS = _.-72z+.4l3,(-_.7Z).413 -

Shock Strut Assembly: (Crashworthy, Kneeling, Fixed) Reference

Figure. 27.

Upper Cylinder-
Pc = 50,390 lb

A= 2.9166 Combine axial load with

I = 6.5119 system pressure loading
p 1.494 of 4500 psi (ULT).
B/t 20.2

97

- .---.. -. '!-- - -.-

Section J-J L/p = 65.77 = 44
1.494

Fc - 50390 -17277 Rc 17,277 .402

2.9166 43,000

Ftn = 4500(4.0) = 40909 Rtn 40,909 .671

2(.22) 61,C00
1
MS + 4_.402'+.671-(_.402).671 -1 = +.07

Table 21 summarizes the margins of safety calculated for the

major components of the two baseline crashworthy shock struts.
The large margin of safety for the lower cylinder of the
retractable gear is the result of combining the cylinder with
the steel upper piston. Combining the two components into a
one-piece machined structure caused the cylinder to be of
minimum wall thickness. The minimum thickness (0.083 inch wall)
is required for the cylinder to be concentric. The minimum
margin of safety for the drag beam designed to the obstruction
loads was 4 percent. The maximum was 9 percent.

98
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17, 99
OPTIMIZED SHOCK STRUTS - BASELINE AND UPDATED DESIGNS

The following analysis optimizes the baseline shock struts and

the shock struts required for the peak load at 42 feet per
second -5 deg pifrh and 15 deg roll.

Baseline design load Peak design load

P cB = 50,390 lb P C = 57,000 lb

Pressure = 6074 psi Pressure = 6870 psi

Lower Piston

Sections C-C and G-G (Figures 26 and 27)

O-D = 3.250 4340 STEEL, 180,000 psi HEAT TREAT

I=.908 in 4 ., p=l.12, Lp=20.0 I=1.078 in 4 , p=l.16, L/p=18.9

Compression stress Compression stress
fc=50,390/.718 = 70,181 psi fc=57,000/.796 = 71,600 psi
Fc=172,500 psi Rc=.406 Fc=174,000 psi Rc=.411
Hoop tension Hoop tension
ft= 6074x3.106 = 131,000 psi fz= 6870x3.09 = 132,676 psi
2 x .072 2 x .080
Ft= 180,000 Rt = .727

M.S. = 11 _--40 T?1 M.S. = 0.00

4-.406z+.72Th-(-.406).727 1 MS .0
=0.00

Lower cylinders/upper piston - retractable strut,

sections D-D and E-E (Reference Figure 26).

The lower cylinder/upper piston is a one-piece steel component.

The piston is designed by the strut axial load and an internal
pressure. The cylinder is designed by an internal pressure and
a resulting tension load on the cross section of the cylinder.
The minimum wall thickness for machining a steel cylinder is
based on a diameter/thickness ratio, D/t, of 50.

Lower cylinder, section D-D

O.D. = 4.16 in., t = .083 in., A = 1.064 in. 2 D/t = 50
Pressure area = 4.27 in. 2 4340 STEEL 260,000 H.T.

100
 Baseline pressure = 6074 psi Peak pressure = 6870 psi
P = 6074x4.27 = 25,936 lb PT = 6870x4.27 = 29,334 lb
Tinsion stress
ft = 25,936/1.064 = 24,376 psi ft = 29,334/1.064 = 27,570 psi
Ft = 260,000 Rt = .094 Ft = 260,000 Rt = .106
Hoop tension
ftn = 6074x4.00 = 146,361 psi ftn = 6870x4.00 = 165,542 psi
2 x .083 2 x .083
Rtn = .56 R .63

M.S. is high M.S. is high

Upper piston - section E-E

O.D. = 3.250 in., 4340 STEEL 260,000 H.T.
Column length L = 65.77 in.

Baseline load, Peak load,

P = 50,390 lb P = 57,000 lb
Pissure 6074 psi PFssure = 687C psi
t=.087 in.,A=.865 in, 2 D/t=37.36 t=.098 in. A=.960 in 2 , D/t=33.10
I=1.0822 in 4 , p=1.1 2 , L/p =59 I=1.205 p=l.12 L/p=59
. Compression stress Compression stress
-" fc=50,390/.865 = 50,280 psi fc=57,000/.960 = 59,375 psi
1 Fc=81000 psi Rc = .720 Rc = .733
F Hoop tension Hoop tension
"f = 6074(3.076) = 107,377 psi ftn= 6870(3.054) = 107,045
th 2 x .087 2 x .098
Rth = .413 Rtn =.412
1 -
M.S. = 72L+.413(_.72).413 - M.S. = +0.00

= +0.01

Lower cylindler/upper piston - fixed strut,

sections H-H and I-I (Reference Figure 27)

The lower cylinder/upper piston is a one-piece aluminum com-

ponent. The piston is designed by the strut axial load as a
long column. No pressure is applied inside the piston.

101

. . . ..
- N

SPCB
B = 50 390 lb ' PPe= 57,000 lb
t=.137 in., A=1.447 in2, D/t=25.5 t=.156 in., A=I.638 in 2 , D/t=22.4
I=2.04 in 4 , p=l.19 L/p = 55.3 I=2.29 in 4 , p=l.18 L/p 55.6
F = 35,000 psi F = 34,750 psi
r callow c allow
fc = 50,390/1.447 = 3481.3 psi fc = 57,000/1.6 = 34,750 psi
M.S. 0.00 M.S. = 0.C

The lower cylinder (H-H) of the fixed shock strut and the upper
cylinders (F-F and J-J) of both struts are of 7075-773 aluminum
and are designed by the same axial load and internal pressures.

Lower cylinder, fixed (1H-H)

Upper cylinder, fixed (J-J)
"Upper cylinder, retracted (F-F)

I.D = 4.00 in. 7075-T73 Aluminum

Baseline load Peak load

Pressure = 6074 psi Pressure = 6863 psi
P= 25,936 lb PT = 33,717 lb

2 2
"t=.190 in, A=2.50 in t=.210 in, A=2.77 in
Tension stress
ft = 25,936/2.50 = 10,374 psi f = 33,717/2.77 = 12,172 psi
F = 66,000 psi = .157 = .184
=R
HooT tension
fth = 6074x4 = 63,940 psi f = 6863x4 = 65,362 psi
2x.190 2x.21
FT = 61,000 psi RTH = 1.04 R = 1.07

""' MS 1 MS
MSV 1 04l'+. i57'-.157(1.04) S 41 07T+. 184,-.-184(1. 07V

= 0.03 = 0.02

The sleeve is designed to crush due to loads that develop from

sink rates greater than 20 fps or loads exceeding the maximum
"load for the noncrashworthy gear shock strut. No change is
required. The weight changes for the crashwortny baseline
shock struts are shown in Table 22.

102
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030
WEIGHT SLBSTANTIATION
The weights of major components in a shock strut assembly are
proportional to the square of the velocities at a fixed gross
weight. The lower piston of the noncrashworthy shock strut,
for example, is 2.8 pounds (Ref. Tablc 18). The weight of the
retractable and fixed piston is 11.3 pounds (Ref. Tables 18).
Therefore, adjusting for the margins of safety of 0,

11.3-.9
2.8-.29 = 4.1 and 422/202 = 4.4

The weight of the components in a shock strut assembly at a

"fixed sink rate and gross weight are directly proportional to
the axial load in the shock strut as a result of pitch and roll
angles. The weight is a result of changes in the wall thick-
ness of pistons and cylinders as shown in the structural
analysis. End caps, floating pistons, seals, lugs and bearings
are changed. Table 23 compares the wall thickness of the
baseline shock strut components adjusted for zero margins of
safety and the updated shock strut. The original assumption
that the weight of the shock strut is proportional to the load,
used to develop the curves of Figure 38, is substantiated by the
data of Table 23.
TABLE 23. SHOCK STRUT COMPONENTS, WALL THICKNESS

BASELINE UPDATED RATIO

(M.S.=0) (M.S.=0) UPDATED/BASELINE
Retracted/Fixed
Lower Pistons t=.072 in t=.086 in 1.11
Upper Pistons t=.087 in t=.098 in 1.13
Upper Cylinders t=.190 in t=.210 in 1.11
Lower Cylinder t=.137 in t=.156 in 1.14
(Fixed)
Design Load 50,390 lb 57,000 lb 1.13

Table 24 presents the shock strut weights adjusted for zero

margins of safety. Tables 25 and 26 present the ..eights of
the shock strut components for the updated landing gears. The
weight changes of Tables 24, 25 and 26 are negligible for this
preliminary study when compared to the system weight of 380
pounds.

104
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107
Aerodynamic Drag - Updated Design

There are no changes in the aerodynamic drag characteristics of

the updated landing gear designs since the outside diameter of
the exposed components remain unchanged.
Fabrication Methods And Processes
Cylinders and pistons are machined from forged aluminum or
steel as requirec. The wheels are of forged aluminum. The
drag beam would be fabricated from two steel iorgings, welded,
and machined to obtain a tapered thickness as required. The
cost reriains basically unchanged.

RELIABILITY

The landing gear reliability benefits from being designed to

stringent crash survivability and ground flotation require-
ments. The gears are designed for sink speeds up to 10 feet
per second with a reserve energy capability of 12.25 feet per
second. The main gears and the tail gear are capable of
preventing the fuselage from contacting the ground at a 20-foot-
per-second sink rate. Both main gears are the articulating
type which, because of the drag beam, enables the gear to climb
over ground obstacles without becoming overloaded. The long
wheel base is ideal for rough terrain operation and provides
protection for the fan shroud during flared and tail down
landings. The articulated gear eliminates oleo bending loads
to provide longer bearing life and reduced friction. The tires
were selected to satisfy the most adverse center of gravity and
will, therefore, provide extended life under average condi-
tions.

The mission safety and system/total mean time between failures

(MTBF) for the landing gears is similar to the U-H-60A.

MAINTAINABILITY

The landing gears designed to crash loads is overstrength for

normal operations, thus maintenance is reduced. The shock-
absorbing wheel landing gears can take high contact loads.
They are not as apt to distant or spread, as skid gears do,
when high loads are applied. The simplicity of the landing
gears permits easy replacement of wheels or brakes. Since the
gears are similar to the UH-60A the maintenance man-hour per
flight hour is estimated to be 0.035.

108
 CONCLUSIONS

Based upon the study of crashworthy landing gears for a 10,000-

pound LHX Utility helicopter, the following conclusions can be
made:

1) The vertical impact design condition envelope at 42 feet

per second sink rate can be extended from 10 degrees of
roll to 15 degrees of roll.

2) The maximum load in the shock strut occurs when the

helicopter is at a nose down pitch angle of 5 degrees at
any sink rate above 20 feet per second.
3) The maximum weight increase is 77 pounds for a retractable
crashworthy landing gear, and 48.5 pounds for a fixed
crashworthy gear when compared to a retractable non-
crashworthy gear. The major weight increase is in the
shock strut of both crashworthy gears.
4) The aerodynamic drag of the fixed crashworthy main gear
represents a power loss of 26.6 percent compared to 0.7
percent for the retracted crashworthy gear.

5) Crew/troop seats designed independently of the landing

"gear, when combined with a crushable fuselage under
16 structure, can provide protection for the occupants when
F the helicopter crashes with the gear retracted. However,
the load factor on the high mass items is excessive for
sink rates beyond 25 feet per second.
6) The weight increase of a shock strut is more affected by
sink rate conditions than pitch and roll angles.

7) The drag beam is sized by notmal obstruction load cri-

teria. No increase of the drag beam is required if the
beam is allowed to bend plastically during a crash se-
quence.
8) Composite materials such as graphite/epoxy cannot be used
in the main wheels due to high temperatures generated
during braking. Broken fibers in a shock strut assembly
may cause loss of pressure during normal or crash land-
ings.

109
 RECOMMENDATIONS
The following recommendations are made for further research
to
improve crashworthiness of LHX helicopters.
1. Evaluate the effect of crash loads on the high mass items
and the fuel system. Studies have indicated that the body
group and the fuel system are the largest weight drivers
for a 10,000-pound class helicopter.
2. Conduct design studies to allow the cockpit/cabin section
of the airframe to pivot about a low point in the struc-
ture as a section of the upper aft cabin is collapsing due
to a nose down crash, gears retracted.

3. Evaluate the effect of repeated high impact loads on

composites with epoxy matrix.

110
 REFERENCES

1. Meir, W. H.,"Application Of The Advancing Blade Concept To

The LHX Concept Developmenti, Sikorsky Engineering Report
SER-510096, April 15, 1983.

2. Lowry, D. W., B.L. Carnell, G. Thomas and M.J. Rich,

"CRASHWORTHY DESIGN PARAMETER SENSITIVITY ANALYSIS" SER-
510076 (Unpublished), Contract DAAK51-79-C-0043, July 6,
1982.

3. Coltman, J.W., Design And Test Criteria For Increased

Energy-Absorbing Seat Effectiveness, USAAVRADCOM TR
82-D-42, Applied Technology Laboratory, U.S. Army Research
and Technology Laboratories (AVRADCOM) Fort Eustis,
Virginia, 23604, March 1983, AD A128015.

4. Kohler, K., "Structural Analysis of Main Landing Gear",

Sikorsky Engineering Report SER-70525, June 15, 1983.

5. Young, B., "Landing Gear, Main and Tail, Qualification

Test Report" Sikorsky Engineering Report, SER-70174,
November 1978.

ill
 APPENDIX
WEIGHT SENSIVITY AND KRASH ANALYSES

Tables A-I through A-6 of this appendix present the weight

sensitivity analysis for the crashworthy landing gear systems
at three sink rates and sixteen pitch/roll conditions.

Tables A-7 through A-13 present the input data for the KRASH
computer analysis. This data is provided as reference data
only.

113
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4J4J4.
U = - toW to to to .tto %D w to WtD Wt
N ~~~c
0 m .0 . .* . . . . .
to toW(

'4- -J r~-r-r- qtrr-r. iR-.tr rr-r

LU H, C,
W-

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4--

* .0 CDJCD(D(m N c%.) CJ r-oCi-D-DtC ClI.CD C I.O 0

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-t+ C.8 CD- r4CDj wLO D rl,o-4 CD n t- 0 n
V)m Cc) rA ujc H r-.~ F4. n o,- HHLHn-c
nc r
7d n -iti 45 ml0cl 1 "r lC l r--4-llr v
LnL

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44.)

4 - - IV) 4
V4 C,
V- 4- '0L
O U ,4 ~ , to o L t ~~~~~~~~- (1) to)H
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to 4-J
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czO1 011lt

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(

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- .4 -118
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N . rL or *L DtD c)- nL V nl
W4.1 nme mC)e V n f r v
to .
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< C. .U . .D
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CA 4-.-Gi -tw rI q c 1-tuq'. L0 -r -Wv
-W04. qtONCR 4

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0)-r_ q.r-r0).N-W n-1.NND
~ 0 ~ 0 m r-4 cofI- I0'*
W .0 C

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0)ww 6L; 40 l 6 64C i 0 S
(A 40 .3c -gm0) H R , mm4
4- 01C I0C I1'. ~ .
CJ
tv 0 f,4) .r)v-0 %
4 C - O~
L) r_ ) ,4- -4-,4 i -I-,I C t4

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C> C*C D( D D MC DC ) 4 0a(

Oj *D o0a D I D- D C n Oc 0L c Ds =

UW a) -

w4. 394.1

cm C;C
o Ln (m n( O C n DL lC n4 COC
U)
IL LL r 4- CO-

4)0 cm0U c-

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CL < '041
Cn
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119
 TABLE A-7. MASS PROPERTIES AND LOCATION

_ NUON!S...5 A5..CO_OIWIjUM.T.F...F. f. LAL.. .....flSS JIOntlTS1 0LIrnaTI2A.LB.IM-SECR I~.-.....___

I M X"Y Z. ix lY rz z
1 Z.234C0D+03 2.000000.02 0.0 S.85000.01 l1.000C00e02 1.000000+02 1.000000*02 1
L8170 0003
L.. Z.79000,0o __..0.0 -. . ';.85000DO02 1 _.00:3+02. 1.OO0OoDOZ 1.0000+0,02 2
3 2.753003:113 3.000000.02 0.0 1.800000D02 1.000000o02 1.00000oo 02 1.000C00002 3
4 1.930000.03 3.200000402 0.0 5.850000.01 1.0000.0002 1.000000+02 1.000000D02 4
5 9.000000+02 4.500000+02 0.0 8.8500c0o01 1.000000+02 1.OCOo0,o02 1.000000*02 S
-.ft P-500000+01-.2.55750.2 -3.425000+01 - 5.fl3coo30l -1.0coo00002 1.000003tO2 -- 1.000002 -- 6
7 2.500000+111 2.S575CDO02 -3.405000+01 5.413000+01 1.0000CO02 1.00000.D02 1.03300D#02 7
8 5.000000+01 2.495000+02 3.425000.01 3.7050C0+01 1.0000co+02 1.000000+02 1.000000+02 a
I 5.000000.01 2.495000.02 -3.425000.01 3.725000+01 1.000000*02 1.O000CO+02 1.000000.02 9
-20. -2.000000+(,1 Z.000000.02- .0.0 - 1.08500D+01 - 5.9500300 6.610000+00 - 3.150030+00 - 10 -
11 7.300000+01 2.000000.02 0.0 7.830000D01 1.395000.01 1.477000.01 1.683000+01 11
12 7.300000.01 2.0000CD002 0.0 9.340000.01 1.130000.01 1.0CoCo.01 6.530000+00 12
13 7.300000.01, ?.000000+02 0.0 9.340000+01 1.1300C3.01 1.0&0000+01 6.5:0000*00 13
l. .. 2.SOOO5.01 -. 50Z0C0002 -. 3.825000+01 -3.i10000+01 -. 5.0000C0303 5.00^003+03 - 5.3CC0003 -14
"15 Z.5000S0.01 2.50oScr.02 -3.805000.01 3.110000.01 5.0000C9#03 5.00000.+03 5.00030003 15

TABLE A-8. NODE POINT LOCATIONS

_NODE-_POINT PATA.. . . . . . . . . .

'ASS N.P. NODE POINT COORDINATES F.S.,B.L.,I.L.

I I X's Y96 Z.6
..... 3
___3.000000402-. 0.0 - 5.850000+01..
2 1 2.740000+02 3.425000.01 1.057509+02
Z 2 2.261700+C2 3.4,50-10+01 5.78700D.01
2 3 2.740000+02 -3.4^5000*01 1.057500.01
2 4 2.261700.02 -3.425000+01 5.737000+01
S8 2.502500+02 3.425000+01 3.11C000001
9 1 2.502500.02 -3.425000+01 3.ilOco.o01
1 1 2.000000+02 2.400009+01 6.050000+01
1 2 2.000000+02 -2.400000+01 6.050000+01 _
2 5 2.790000.02 2.40300C0+01 6.050300+01
2 6 2.790000.02 -.. 4O03C00C1 6.0500CD+01
4 2 3.200000.02 2.4GO000+01 6.0530000.C1
--- 4 3_ 3.200000+02_ -2.403000D+01 6.050000+01_
1 3 2.0000^0+02 4.IOCOCD+O1 6.0.52000+01

1 4 2.000000+02 -4.COC0O001 6.05CO00+01

2 7 2.790000+02 4.6C00VD+01 6.050000+01
2 8 2.790000402 -4.6C000+01 6.0500C0+01
2 9 2.2b1700+02 2.7C000001 5.73/00+01
2 10 2.261700D02 -2.700000+01 5.787000401

120
Q

.4

TABLE A-9. BEAM DATA

6~DAMPIPNG T
BAN AREA MIOMENTS
OF INERTIA LENGTH RATIO L P-COVES BEA"

JL- 4- N -.A - IYYt - - Z2 - - JX -IQ z XLi - -- crOAO1- McI- I1 J J.1. L- i-.t

1 t 0 0 2.000D#01 4.5000402 1.0000+03 2.0000.02 1.000C0. 1.002CO00 0.0 7.9000.01 5.003-02 s 0 a 0 0 1 1 2
2 4 0 1 2.000#01 4.5000S02 2.0000#0. 2.0000402 1.0000900 1 000.01 0. 0 2.100o.01 5.0000-02 S 0 0 0 0 2 2 4 3
3 4 0 1 2.0000.01 3.5000.03 3.5000.03 2.40C0003 1.0=3,00 1.0003.00 0.0 1.21SD002' 5.003-02 S 0 0 0 0 3 3 4 I
.A4...Sfl0 Z.000..OD2J..2 1.0000#0z I..oc20.oa 1.0=0~3* 0.0 A.33.0Ofc S.00.0-02 5 0 0 0- 0 '... 4
t 6 1 0 1.1000.00 1.0000.02 1.-00C302 1.0000.00 1.OCO-00 1.OOC.00 0.0 5.4750.01 5.00C-02 0 1 1 0 0 5 2 6 1
2 7 3 0 1.1000.00 1.0000.02 1.0000.02 1 0000,00 1.0000.00 1 0.L0,00 0 0 5.4700.01 5.00CD-02 3 1 1 0 0 6 2 7 3
6 a 0 0 2.0300.00 1.0000.02 1.0c00.02 I CCCoI*2 1.00co00 1 02O0.00 0.0 1.r0cq001 5 00.0-U2 08 0 1 1 7 6 5 ^
-7-9-.O_0 2.00CO*00 1.0000-OZ 1.0=.02 1.0"O3D 1.0070,00 1.0020.00 0.0 1.20201 5.-0:0-O2 3 0 0 1 1 8 7 9 0
z 8 2 1 2.5303400 1.0000.01 1.0000.01 2.0227.01 1.00..04CO 1.00"'0C 0.1 3 6010.01 5.00C0-02 2 1 0 0 0 9 2 8 Z
9 4 1 2.50OD000 1.000^D01 1.0000.01 '.0000401 1.0=0.-00 I OLO+00 0.0 3.6010-01 5.0C00-02 2 1 0 0 0 10 2 9
10 0 0 4.3000-03 2.0000.00 2.0-0-00 4.00C0O00 1.0070CO 1 CjCo'*3 0 0 1 235J#01 5 0%0-), 4 0 0 0 0 11 1 1C
.- 11-0-.0 2.00D-03 5.0v0-0^ 5.01^0-02 1.002D-01 1..C0ctC0 1. C 30210.0.0 7.430.0 5 C220-CZ 4 3 0c 0a1210
! 1
11 12 0 0 2.00V0-02 7.5000-03 7.5000-03 1.5000-02 1.0027.00 1.00L0003 0.0 1.5100-01 5.0003-02 4 0 0 0 0 13 11 32 0
11 13 0 0 7.2150-03 4.8000-03 4.C000-03 9.6000-03 1.00-0000 1.0030000 0.0 1.5100+01 3.1110-0110 0 0 0 0 14 10 13 3
8 14 1 0 1.0D00.00 0 5000.01 2.5000+01 I.C3OD*00 1.000,00 1.C0O3D03 0.0 4.00C0+03 5.C220-02 8 3 0 1 1 15 8 14
. - 0.-3_...O..OOCOOo
2.5000.01 Z.5000S.0 I 00C0O00 1.000-00 1.0000.00 0.0 4.0000,00 5.0000-02 a 0 0 1 1 16- 9 15 1

TABLE A-10. MATERIAL PROPERTIES

MODULUS OF MODULUF OF
ELASTICITY RIGIDITY
MC MATERIAL (PSI) (PSI)

2 6150H Steel 30.0E6 I1.0E6

4 2024-T3 10.5E6 4.0E6

Aluminum
5 6061-T3 10.0E6 3.8E6
Aluminum
8 Zero Torsion 1.0E6 0.0
Material
10 DRI Spine 1.0E6 0.3E6
(DRI)

121
 TABLE A-Il. BEAM END FIXITY DATA

BEAM P-CODES

IJ I J M N IY IZ JY JZ

R.H. Upper Oleo Attachment 5 2 6 1 0 1 1 0 0

L.H. Upper Oleo Attachment 6 2 7 3 0 1 1 0 0
R.H. Lower Oleo Attachment 7 6 8 0 0 0 0 1 1
L.H. Lower Oleo Attachment 8 7 9 0 0 0 0 1 1
R.H. Drag Beam Attachment 9 2 8 2 1 1 0 0 0
L.H. Drag Beam Attachment 10 2 9 4 1 1 0 0 0
R.H. Wheel Axle 15 8 14 1 0 0 0 1 1
L.H. Wheel Axle 16 9 15 1 0 0 0 1 1

-'C.

-'C..

'Cm.

i-,.~%, 2

-- _C.'
12
 TABLE A-12. SPRING DATA

FREE FRICTION BOTTOMING

SPRING LENGTH COEFFICIENT MFIN1G
MK
lLBtARZKM)-t'J(IKII) .... KE(IZI)-

14 3 0 9.750000#00 3.400000-01 4.000000+04

15 3 0 9.750000+00 3.400000-01 4.000000.04
5 3 0 6.535000+01 3.400000-01 1.000000#04
. L..3.II 1--7..500000+00.
- .3.400000-OL. ...7.000000403-
- 1 3 2 7.500000.00 3.400000-01 7.000000.03
2 3 9 1.500000+00 3.400000-01 7.500000.04
2 3 10 1.500000.00 3.400000-01 7.500000.04
"2 3 6-7-500000.00 3.400C00-01 7.00=000+03-
2 3 6 7.500000.00 3.400000-01 7.000000+03
S4 4 33 33 7.500000+00
7.5O0000#00 3.400000-01
3.400CO0-01 7.000003+03
7.000000+03
- .J..._3-3 .-.7.500000*00 .3.400000-01- 7.000000+03
1 3 4 7.500000400 3.400000-01 7.000000.03
"2 3 7 7.500000.00 3.4000C0-01 7.000000.03
2 3 8 7.500000.00 3.400000-01 7.000000.03

"DEFLECTION COORDINATES
SI(I"Km) SA(IKM) SBEIKM) SF(IKM)
....3.000000000-.3.5000.00,00 __ ot.000000o00o. -4.500000+00
3.000000+00 3.500C00.00 4.0000C0+00 4.500000400
5.030000+00 1.00000+01 2.000000.01 2.400000.01
1.000000-01 2.750000300 2.760000.00 3.750000.00
-1.000000-00 1.7500CO0oo 1Z.7600000.00 i3.750000oo0
S1.010000+00 1.000000+00 1.0000C0g0o
J.()(00cO#*O0 1.010000.00
1.010000#00 1.020000.00
1.0Z0000+*00
1.000000-01 2.7500C0000 2.760000+00 3.750000+00
--. :0000OD-01,2.7500CD00 - Z.760000+00 _ 3.750000+00
S1.O000C-01 2.7500C0+00 2.760C00000 3.750000+00
1.0000GD-01 2.750000.00 2.760000.00 3.750000.00
... 1.000000-01 2.75000+,00 2.760000+00 3.750000+00
.- 1.000000-01 .Z750000#00 2.760000.00 3.750000.00
1.000000-01 2.750000.00 2.760000+00 3.750000.00

SPRING AXIAL FORCfS

FSPOFI KKM) FSFOF( IKM1) CRIT.0AMP CDAMP( 1Kt)
.1.C0C00*04... -2.OCOO3D.04..- i.00oo
...- 0-01- -2.937110,00
1.000000.04 2.00000C04 1.0000C0-01 2.937110.C0
1.950000.04 1.9500C0004 1.000000-01 1.9n6qo+0,0
3.64CO00004 3.64&000.04 2.000000-02 5.809120.01
3.64000.04 ... 3.648C0D004 .....2.OOCOCO-02 -. 5.809120*.1
7.500000.04 7.5C0000+04 2.0000C0-02 Z.375470401
7.500000.04 7.500000.04 2.00c000-02 2.373470+01
4.032000+04 4.03M0004 Z.OCCOCO-OZ 5.507810+01
4.03:0C0.04 .-. 4.0.ZO0.04- 2.000003-02 5.507310+31
1.824000.04 1.824000#04 2.000000-02 3.817970+01
1.824000.04 1.824000+04 2.000000-02 3.817970+C1
1.368000.04 1.368000.04 2.000040-02 3.557350D01
1.368000#04 -.. 1.363000+04 -... 2.0=0000-02-3.55733,01G-
1.438000+04 1.488000.04 2.000000-02 3.345960.01
1.488000.04 1.488000+04 2.000000-02 3.345960.01

123
 TABLE A-13. OLEO STRUT INPUT DATA

OLEO STRU' BEAM DATA

SEA" AIR CURVE PARAJIETERS

_JI J " N MLEO FAO FAA . EXPOLE YMAX -

2 6-1
k 0 1.3270+01'4-.980 h03
1.2,200.02 1.3CC.00 1.1000.01
6 2 7 3 0 1.3270401 '.96C0003 1.2200+02 1.30'D000 1.l000+01
7 6 8 0 0 1.5030+01 7.3740.02 1.:.004ZC 1.3003+00 1.4,500*01
8 7 9 0 0 1.5030.01 7.374D+02 1.2Z00O2Z 1.3000+00 1.4500'01
BEA" DAMIPING CCIfSTANTS,COULOM3 FRICTION AtN LINEAR 'SPRINGS |EXTENSION&CO!IFRESSICO)

_ZJI_ J._H_B1_N OLEO .BROLEO.__. XEXT -... XKCCiP -. FCCUL --

5 a 6 1 0 1.0000o00 3.000.O0 Z.OO0o0o0 Z.coo0D05 5.50Co000
6 2 7 3 0 2.000000 3.0000.00 2.0000#05 2.000.03 5.5000.00
7 6 8 0 0 5.0000-01 3.0000+00 2.0000&05 2.000005 5.S00j000
_0-7.-9 - .90J.o.500-.o.1 A..oCo.go0.z.oooo-os.ZOOOo.s 5.5000o0o.0

124
4770-85