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USAAVSCOM TR-85-D-4

AD-A162 097ADVANCED TECHNOLOGY HELICOPTER LANDING GEARPRELIMINARY DESIGN INVESTIGATION

D. Lowry

SIKORSKY AIRCRAFT DIVISIONUnited Technologies CorporationStratford, 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 DIRECTORATEUS 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 landinggear 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 wereconducted for the crashworthy designs to determine the energy absorbing component weights forvarious sink speeds and aircraft attitudes. These results are used with other landing gear designparameters to establish recommended crashworthy design requirements.

The results of this program represent a significant advance in the understanding of the parameterswhich influence crashworthy landing gear and energy absorbing structure weight. These findingswill be integrated with the parallel contract and past efforts in landing gear weight sensitivity todevelop 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, Safetyand Survivability Technical Area, assisted in this effort.

DISCLAIMERS

"The findings in 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.%"'aa.-. . . . . ... ., - .. .- . .- , .-. . •

UnclassifiedSECURITY CLASSIFICATION OF THIS PAGE

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Unclassified2a SECURITY CLASSIFICATION AUTHORITY 3 DiSTRIBUTIONIAJAILABIUTY OF REPORT

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"4 PERFORMING 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 TechnologyStratford, Connecticut 06601 Activitv (AVSCOM)

Fort Eustis, Virginia 23604-55778& 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 UNITELEMENT NO. NO. 1 L162- NO. ACCESSION NO.

62209A 209AH76 04 03311 TITLE (Include Security Classification)

Advanced Technology Helicopter Landing Gear Preliminary Design Investigation

12 PERSONAL AUTHOR(S)

F .D. Lowry13a. TYPE OF REPORT 13b TIME COVERED 14 DATE OF REPORT (Year. Month, Day) S. PAGE COUNT

Final I FROM 1983 TO 1984 IOctober 1985 I 12216 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 mainlanding 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 landing gear loads developed by a computer program KRASH. KRASH was used to obtain the design loads atvarious sink rates, and pitch and roll attitudes. The design loads were used to size the landing gear structurein 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. Therecommended 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 Unclassified22a 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 PAGEAll other editions are obsolete Unclassified', Ucasife

a-. -A

PREFACE

This Advanced Technology Helicopter Landing Gear Preliminary DesignInvestigation was conducted under contract DAAK51-83-C-0040 with theAviation Applied Technology Directorate, U.S. Army Aviation Researchand Technology Activity (AVSCOM), Fort Eustis, Virginia.

The work was performed under the general direction of Mr. G. Downerof the Aviation Applied Technology Directorate. Sikcrsky Aircraftprincipal 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 GearDesign.

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±iii

TABLE OF CONTENTS

PAGE

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

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

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

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

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

BASELINE LANDING GEARS .......... ............. 5Design Loading Requirements ..... ......... 5

Ground Handling ... ............. . 5Obstruction, Symmetric and

Asymetric Reserve Energy .... ....... 6Taxiing ............. ................. 6Supplemental Design Requirements . ... 7Vertical Impact Design Conditions

Envelope Crashworthy Gears ... ...... 8Additional Crashworthy Design

Requirements .......... ............. 8Pitch, Roll and Velocities .... ....... 8

Noncrashworthy, Retractable Main Gear . ... 9Crashworthy, Retractable Main Gear ... ..... 12Crashworthy, Fixed Main Gear ... ........ .. 15

"Tail Landing Gear ....... .............. .. 19Drag Beam - Main Gears .... ........... .. 21I LOADS ANALYSIS ................................ 22Design Mass Properties .... ........... .. 22Main Rotor Parameters ..... ............ .. 22Landing Gear Parameters .... ........... ... 22

Main Gear ....... ................ ... 22Tail Gear ....... ................ ... 22

Normal Operating Conditions ... ......... ... 27Static Design (Landing) ... ......... ... 27Static Design (Ground Handling) ..... ... 27Xjrmal Landing RequirementsIs ....... .... 27

Crashworthy Design Requirements ......... ... 28

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

v

TABLE OF CONTENTS (Cont'd)

PAGE

KRASH Program Results ..... ............ .. 37Time of Zero Vertical Velocity at

Aircraft C.G ...... .............. ... 37Landing Gear Strut Maximum Stroke

and Time of Occurrences ... ........ .. 39Ground Load ......... ................. ... 41Inertial Load Factors ..... ............ .. 47Seat Stroking ....... ................ ... 51Fuselage Crushing ...... ............. .. 54

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

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

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

RECOMMENDED DESIGN CRITERIA .... ........... ... 94Attitudes and Velocities .... .......... .. 94Detail Design Criteria for the Main

Landing Gear ............................ 96

DESIGN UPDATE ......... .................. ... 97Shock Strut Structural Analysis ......... ... 97Optimized Shock Struts - Baseline and

Updated Designs ....... .............. .100Weight Substantiation ..................... .. 104Aerodynamic Drag - Updated Design ......... .. 108Fabrication Methods and Processes ........ .. 108Reliability ................................ * 108Maintainability ....... ............... .. 108

CONCLUSIONS ........... ................... ... 109

SRECOMMENDATIONS .ii ........ 10REOMEDTIN.............................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 MainLanding Gear ....... ............ .i.i.. 11

5 Retractable, Kneeling, CrashworthyMain Landing Gear ... ........... .. 13

%-4

-. 6 Shock Strut, Crashworthy Retractable

Main Landing Gear ... ........... .. 14

7 Fixed, Kneeling, Crashworthy MainLanding Gear .... .............. ... 16

8 Shock Strut, Fixed, Crashworthy MainLanding 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 Long±tudinal Acceleration Time Historyon 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 MainGear Strut ...... ............... ... 64

30 Tail Gear Structure ... .......... .. 56

31 Tail Gear Drag Strut ... .......... .. 67

32 Composite Main Gear Drag Beam Comparedto 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 BeamWeight .......... ............... .. 81

38 Weight of Affected Structure as aFunction of Pitch and Roll Angles. . . . 85

39 Landing Gear System Weight as aFunction 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 Functionof Pitch and Roll Angle ... ......... ... 89

42 Weight and Cost Trends for LandingGear System ....... ............... ... 90

*° - F

ix

5. .,

t5,.

LIST OF TABLES

Table

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

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

3 Aircraft Properties and Parameters . . .. 24

4 Normal Landing Gear Design Conditionsand Ground Loads ..... ............. ... 29

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

6 Design Conditions for Crash Impacts . 32

7 Time of Zero Vertical Velocity atAircraft 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 ofSafety - Baseline Crashworthy Designs 99

x

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

LIST OF TABLES

Table Page

22 Baseline Shock Strut Weighc ReductionFor 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 WeightSummary ............ ................. .. 107

A-I Weight Sensitivity - Shock Struts andFittings - Fixed Gear - (42 FPS) ....... .. 114

A-2 Weight Sensitivity - Shock Struts andFittings - Fixed Gear - (36 FPS) ....... .. 115

A-3 Weight Sensitivity - Shock Struts andFittings - Fixed Gear - (30 FPS) ....... .. 116

A-4 Weight Sensitivity - Shock Struts andFittings - Retractable Gear - (42 FPS) 117



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 meetinghelicopter crashworthiness design goals of reduced crashinjuries, fatalities, and material losses. However, theability to absorb large amounts of energy in the landing gearsis not achieved without paying strict attention to the energyabsorption requirements in the design effort. Furthermore, ex-perience has shown that this capability cannot be realizedwithout some increase in gear weight and complexity, which inturn has an adverse effect on the aircraft performance andcost. These effects will, of course, vary with the level ofcrash energy which must be absorbed and with the severity ofthe crash attitude requirements.

The Army's most recent production helicopter, the UH-60 BLACKHAWK, 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, alanding gear has been designed to meet most of the crash re-quirements of MIL-STD-1290. However, these aircraft systemsare 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 ofthe Army's increased emphasis on increased forward flightperformance, in order to evaluate the applicability and effectof crashworthiness design requirements.

This report documents the results of a three-phase effort of apreliminary design investigation for a retractable and non-retractable landing gear system designed to crashworthy re-quirements.

1

-------------------------

BASELINE HELICOPTER

Sikorsky's Advanced Blade Concept (ABCt m ) LHX baseline configur-ation developed for the Advanced Technology Helicopter LandingGear Preliminary Design Study was derived from the utility"variant air vehicle provided in Reference 1. However, theLanding Gear Study aircraft was configured more in accordancewith the LHX requirements outlined in RFQ DAAK51-83-Q-0061,dated July 1, 1983.

"The principal physical characteristics of the ABC LHX baselineconfiguration are shown in Figure 1, the Baseline LHX GeneralArrangement drawing. The accommodations for six troops and acrew of two determined the cabin and cockpit size, and there-fore, the body length.

The takeoff gross weight of this utility helicopter, summarizedin Table 1, is 10,000 pounds. The counter-rotating main rotorsand the auxiliary propulsor are driven by twin advanced tech-nology engines (ATE's). The engines are installed behind themain gearbox enclosed by easy opening cowlings to allow foraccess, inspection, and/or maintenance.

A shroudeC pusher propellei is used to provide efficient andquiet cruise thrust. The shrouded configuration will protectground personnel and minimize damage to the propeller whileoperating on, or close to, the ground.

The Sikorsky ACAP diamond-shaped cross section was sele2ted forthe baseline body shape to provide a reduced body radAr returnat a minimal weight penalty. In addition, the shape of theaircraft is used to control the deflection of the structure.The body shape was iterated on as the landing gear designprogresses 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, fuelsystem, tub structure, and retention of high mass items. Thefuel system and the high mass items (engines transmission androtor head) are located Lehind the occupied area of he cabinand cockpit. Locating the fuel system and high mass itemsbehind 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 itemsbecomes the major airframe structure.

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

TABLE 1. WEIGHT SUMMARY

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

Total Weight Empty 7489Useful Load 2511

Gross Weight 10,000

4

BASELINE LANDING GEARS

The landing gear configurations described herein are compatibleto the LHX-utility class helicopter in the 6000- to 10,000-poundgross weight range. This version is configured for two crewmembers and six troops located forward of the heavy mass itemssuch as rotor head and transmission. The aircraft is con-figured with a tail wheel landing gear arrangement as shown inFigure 1.

To comply with the crashworthy requirements and to meet th, airtransportability requirement, the main landing gear is Locatedoutboard of the troop cabin area with the shock strut posi-tioned aft of the cabin door frame. The mounting point for theshock strut upper stage is provided by the aft cabin bulkheadwhich also supports the heavy mass items such as transmissionand rotor head. To simplify the kneeling and/or retractiongeometry, a trailing arm (or "drag beam") concept is utilized,thus allowing landing gear movement in a vertical planethroughout its stroke without obstructing the cabin door orinfringing upon the space occupied by per-onnel. In the eventof a crash impact this arrangement allows the gear to strokewithout penetrating the fuel cells or occupied areas. AMCP706-202 was used in determining the required angles defined inFigure 1.

3• All shock strut configurations are air-oil types designed inF 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 accordancewith MIL-G-5514. For the baseline concepts, the variouscomponents comprising the shock struts and mounting or support-ing structure are manufactured from heat treated, high strengthsteels (4340, 300M) or aluminum alloys (7075, 7175).

The tire sizes were selected based upon the landing gear staticreactions taken at the helicopter design gross weight and tomeet the requirement for a CBR of 2.5, equivalent to the H-60series helicopter. The tires meet the requirements ofMIL-T-5041, whereas wheels and brakes comply with requirementsof MIL-W-5013. All main and tail landing gears describedherein use the same wheels, tires, and brakes.

DESIGN LOADING REQUIREMENTS

Each landing gear system is designed to meet the followingrequirements:

Ground Handling

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

5

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b. Jacking - Jacking requirements shall be in accordance withMIL-A-8862, and the basic design gross weight shall apply.

c. Mooring - Mooring requirements shall be in accordance withMIL-A-8862 except that a 70-kt horizontal wind shall apply.

d. Transport - Transport requirements shall be in accordancewith AMCP 706-201 requiring a limit vertical load factor of2.67 at the basic design gross weight.

Obstruction, Symmetric and Asymmetric Reserve Energy

"a. Obstruction - Obstruction landing requirements shall be inaccordance with MIL-S-8698 at the basic design gross weight.

b. Symmetric and Asymmetric - Symmetric and asymmetriclanding conditions for tail wheeled gear configurations shallbe in accordance with AMCP 706-201 except that landing surfacesshall be flat and firm. Limit landing conditions shall also bein accordance with AMCP 706-201 at the basic design grossweight.

Taxiing

a. Two-Point Braked Roll - The requirements of MIL-A-8862shall apply for the two-point braked roll except that thevertical load factor at the center of gravity (CG) shall be 1.2for all gross weights.

b. Three-point Braked Roll - The requirements of MIL-A-8862shall apply to the three-point braked roll of a helicopter withnose wheel landin9 gear except that the vertical load factorat the CG shall be 1.2 for all grcss weights.

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

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

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

6

f. Pivoting - The pivoting requirements of MIL-A-8862 shallapply.

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

h. Special Tail Gear Conditions - The special tail gearconditions 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 speedsfrom zero up to 50 knots at limit sink speed on level groundand from zero un to 40 knots at reserve energy sink speed onlevel ground at th basic design gross weight.

b. Consideration siall be given to both crew-initiated andautomatic retractable wheeled landing gear during transitionto/from NOE flight.

c. Design fatigue loads shall be considered.

d. A landing gear shall have kneeling provisions for trans-portability.

7

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Vertical Impact Design Conditions Envelope Crashworthy Gears

The crashworthy landing gear systems are designed to sink speedand altitudes given in Figure 2.

40 40

200 20

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42 IT/SIC / • . . . . __\ 36 FT/S1C

Figure 2. Crashworthy Design Envelope

Additional Crashworthy Design Requirements

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

2. A retracted gear shall provide some vertical crash impactenergy attenuation.

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

Pitch, Roll and Velocities

1. No fuselage-ground contact shall occur, and there shall beminimal or no effect upon dynamic components during a crashimpact with the landing gear extended. For vertical velocitiesof 15 and 20 fps, all combinations of the following pitch androll 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 notbe transmitted to the crew with the gear retracted, assumingenergy absorbing subfloor structure and stroking seats. Forvertical velocities of 25, 30, and 35 fps, all combinations ofthe following pitch and roll attitudes shall be considered:


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

3. Injurious loadings, as defined in MIL-STD-1290, shall notbe transmitted to the crew with the gear extended, assumingenergy absorbing subfloor structure and stroking seats. Forvertical velocities of 30, 36, and 42 fps, all combinations ofthe following pitch and roll attitudes shall be considered:


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 anoncrashworthy, retractable landing gear with kneeling capa-bilities to meet the air transportability requirements. Theupper structural member, or retraction brace, and hydraulicretraction/extension actuator are mounted to the aft cabinbulkhead as shown in Figure 3. In addition to retracting andextending the landing gear, the hydraulic ;actuator also con-trois the kneeling function of the gear. Uplocks and downlockswithin the hydraulic actuator locks the gear in its respectiveposition during the retraction/extension cycle. The lowerstage is an air-oil shock strut, as shown in Figure 4, designedto withstand the normal landing loads. The strut consists of apiston inside a cylinder housing, both components manufacturedfrom 7075-T73 aluminum alloy. Internally, the strut assemblycontains an aluminum alloy metering pin and a floating piston

9

X.

which separates the air from the hydraulic fluid. This assem-bly acts as a shock strut capable of withstanding normallanding loads at 10 feet per second (fps) sink speeds with areserve energy capability to 12.25 fps. Total wheel strokefrom extended to compressed position is 12.00 inches, whichresults in a shock strut stroke of 11.90 inches per data fromthe KRASH computer program. The retraction brace connects thehydraulic actuator and the shock strut assembly to the aftcabin bulkhead. This member, as depicted in Figure 3, is madefrom 7075-T73 aluminum alloy. Mounting hardware is comprisedof standard self-lubricating bushings and bearings, and is

attached with high strength bolts.

A hydraulic system consisting of appropriate pump, valves, andreservoir with hydraulic fittings and hose assemblies isincorporated into the helicopter for providing power to thehydraulic actuator. An electrically operated system activatesthe hydraulic cycling of the landing gear and also includes therespective limit switches to indicate uplock and downlockpositions.

RETRACTED POSITION

HYDRAULIC

1A ACTUATOR

I \ ~RETRACTION IBRACE

il _\ -. -

all ~ 1Kf~ / FUEL CELL

' IPIVOT" ) POINT

300

SHOCK STRUTWL

43.1STATIC GROUND LINE , . / WL 35

iSTA259.25

Figure 3. Retractable, Kneeling, NoncrashworthyMain Landing Gear

10

"N• -• Y,"• • - ' "• -"- - . '-% ' + • % " - ' ' - _ .".°- " ' ' 'r -. '• '•._. ." -. " ' - ",• •". "• "

HYDRAULIC SERVICE/BLEED PORT

AIROIL

AIR VALVE FLOATING PISTON METERING PIN

REBOUND DAMPER VALVE

Figure 4. Shock Strut, Noncrashworthy, Main Landing Gear

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CRASHWORTHY, RETRACTABLE MAIN GEAR

The main landing gear configuration depicted in Figure 5 is aciashworthy, retractable landing gear designed to normaloperating requirements and the crashworthy requirements for thevertical impact design conditions. This concept is a trailingarm design with a universal-mounted, two-stage, air-oil strutdesigned to absorb 65 percent of the energy at a sink speed of42 fps at a pitch and roll of zero degrees. In addition tomeeting the crash impact criteria, the strut acts a6 a hy-draulic recraction and extension actuator which also controlsthe kneeling action required for air transportability. In theevent of crash impacts, the strut strokes in a vertical planeas the drag beam pivots about the fuselage attachment point,thus preventing the landing gear from protruding into the fuelcell or occupied areas. Also in a crash condition, the dragbeam does not obstruct the cabin door opening, thereby allowingrapid evacuation of personnel. With the landing gear retract-ed, some vertical crash impact energy absorption is providedthrough the compression of the tire, wheel deformation, andcontrolled failure of the upper housing attachment fitting onthe rear cabin bulkhead.

Both upper and lower shock strut stages are air-oil typesdesigned to the requirements previously specified but utilizinga variable orifice assembly in lieu of a metering pin, as shownin Figure 6. With a metering pin, gear retraction would behindered because of the interference between this pin and thefloating piston. The lower stage is designed to withstandlanding loads up to 12.25 fps reserve energy sink speeds within14.50 inches total strut stroke. This assembly contains apiston inside a housing of which the upper portion is machinedas the piston of the upper stage. Both components are madefrom 4340 heat treated steel. The upper housing which mountsto the airframe fitting is made from 7075-T73 aluminum alloy.Separating the air from the hydraulic oil in both stages is analuminum alloy floating piston. Each stage contains an orificeassembly, split bearings, a lower housing bearing, and aluminumalloy retraction/kneeling piston. To retract, or kneel, thelanding gear, hydraulic oil pressure is added to each respec-tive housing port. At the same time,ports in the upper portionof the housings are opened to allow hydraulic fluid displacedduring the retraction cycle to be bled into a reservoir. Asthe hydraulic fluid is pumped into the strut, both stagesretract, resulting in gear retraction or aircraft kneeling. Anexternal uplock system shuts the retraction system off when thegear is fully retracted. Reversing the cycle extends the gear.

12

Full strut travel is 27.00 inches which corresponds to an axlevertical travel of 29.00 inches, the travel required to fullyretract or kneel the main landing gear. Mounting hardwareconsists of self-lubricating bushings and spherical bearingwith high strength bolts.

BL SHOCK STRUTBL ST•A

' I34.25 275.25

ti WL.108.25

•-.• +

4 33.60F REF 270

" ~UEL• •,NK

" -E .7.87--9RETRACTED/KNEELEDKN E ELE..DD - ]•"-

POSITION B COMPRESSED

BL BL , -" STA SL10 24 200 26.42' 4_ 1.W.

- - - 3 EXTEND E

BL DRAG BEAM STA

0 6L STA 300.041.25 259.25

Figure 5. Retractable, Kneeling CrashworthyMain Landing Gear

13

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The main landing gear, in conjunction with the tail landinggear, has an energy absorption capability to prevent thefuselage from contacting the ground at a 20 fps sink speed.Results from the KRASH Computer program indicate that each maingear absorbs approximately 44% of the total energy for thissink speed rate. To meet this criteron, the upper and lowervariable orifice assemblies are designed in such a way as toallow the shock strut to absorb energy at crash impacts result-ing from 20 fps to 42 fps sink speeds and still react the loadsdue to normal landings up to 12.25 fps reserve energy sinkspeed.

The hydraulic system for controlling landing gear retraction,extension, and kneeling cycles is part of the helicopteroverall hydraulic system and contains the appropriate shuttlevalves, check valves, pumps, and accummulator or reservoi-,with fittings and hose assemblies. An electrically operatedsystem activates the hydraulic cycling of the landing gear andalso includes the respective limit switches to indicate uplockand downlock positions.

To comply with the Supplementa2 Design Requirements specified,it is recommended that the landing gear be automaticallyretractable with pilot override. The aircraft's flight speedindicator and radar altimeter will provide signal indicationsfor automatic retraction and extension, so that below (to bedetermined) knots and below (to be determined) feet from theground, the gear will always be in the extended position unlessthe pilot chooses to override the indicators. Automaticextension/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 ofpage 6 and the crashworthy requirements of page 8. This con-figuration is also a trailing arm design with the shock strutmounted to the lower end of the arm and to the upper fitting atthe rear cabin bulkhead. This arrangement allows the strut tostroke in a fore and aft vertical plane as the drag beam pivotsabout the fuselage attachment point in the event of crashimpacts, thus preventing the gear from puncturing the fuel cellor protruding into the occupied area.

15

V

V F .- -i[

,[• CRUSHABLE SPLIT SLEEVEFOR CRASH ENERGYABSORPTION, REMOVEABLE!iI\ FOR KNEELING/

-aIFUEL CELL

DRAG BEAM "- AIR-OIL SHOCK•,• .• WL STRUT

43.1 ;TA

STATIC GROUND LINE WL 35 300

K

_ /STA259.25

Figure 7. Fixed, Kneeling, Crashworthy MainLanding Gear

16

.4,Z

AIR VALVE -LOWER STAGE HYDRAULIC PORT

I-A-

AIROI SRU -CRUSHABLE SPLIT = SLEEVE HYDRAULIC 7PORTAIR-OIL STRUT

WITH VARIABLE FOR CRASH ENERGY FOR KNEELING

ORIFICE ASSEMBLY ABSORPTION, REMOVEABLEFOR KNEELING

Figure 8. Shock Strut, Fixed, Crashworthy MainLanding 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 sinkspeeds up to 42 fps crash conditions. The lower stage con-sists of a 4340 steel piston inside a housing made from7075-T73 aluminum alloy. The upper segment of this housing ismachined to become the piston for the upper stage. Separatingthe air from the hydraulic fluid is an aluminum alloy floatingpiston. The variable orifice assembly is designed to allowthis lower stage to absorb energy up to 20 fps sink speedswithin a 14.50-inch stroke. The upper stage contains an airand oil chamber which is bled off into a reservoir when kneel-ing the gear for air transportability of the helicopter. Theupper outer housing is also made from 7075 aluminum alloy andattaches to the airframe bulkhead.

Separating the two stages is a crushable split sleeve for crashenergy attenuation and is removable for kneeling the landinggear. The sleeve has a tapered section designed to crushprogressively as the load increases caused by sink speeds inthe range of 20 to 42 fps. Materials for this tube could beeither steel, aluminum, or composites. To kneel the landinggear requires this sleeve to be removed and the hydraulic fluidin both the upper and lower stages to be bled off until thedesired height is obtained. Reservicing the strut will extendthe upper stage, allowing reassembly of the sleeve. Allmounting hardware is comprised of standard high strengthfasteners used with self-lubricating, spherical bearings andbushings.

The stroking sequence of the shock strut during crash condi-tions is such that as the sleeve crushes the upper stagecompresses, which forces the hydraulic fluid out the serviceport to a reservoir. Blowout plugs could be designed into theinner tube in such a way as to allow the hydraulic fluid toenter the internal air chamber as the upper stage strokes, thuscontaining the fluid within the strut.

TAIL LANDING GEAR

The tail landing gear sho~m in Figure 9 is designed to thenormal operating requirements and to the crashworthy require-ments of a 20-fps sink speed. At 20 fps sink speeds, the shockstrut of the tail gear is capable of absorbing 12% of the totalenergy of the aircraft within 12 inches of strut travel, whichprevents the fuselage from contacting the ground. KRASHprogram data indicates that the total strut travel at 36 fpsand 42 fps crash impact conditions is 24 inches, resulting inthe deformation of the tail gear and tailcone structure. Thecombination of the two deformations results in 12% energyabsorption capability distributed over a larger area.

The location for the tail landing gear was chosen in order tominimize the strut stroke when preventing fuselage contact withthe ground at 20 fps sink speed conditions and to provideadequate rollover characteristics. This location also allowsthe landing gear to be retracted forward and to become com-pletely enclosed within the tailcone, thus maintaining smoothairflow through the ducted fan. From this retracted position,the combination of gear weight and airflow would provide apositive release and extension to a down and locked positionupon actuation of the emergency uplock release system. Ahydraulic system powers the actuator with uplocks and downlockswhich controls the retraction, extension, and kneeling of thetail landing gear. An electrical indicating switch system isalso incorporated to indicate gear position.

The air-oil shock strut is a cantilever-mounted type designedto absorb energy at sink speeds up to 20 fps. The strutassembly consists of a steel piston and fork insidtý an aluminumalloy trunnion that mounts to the airframe structure. Separat-ing the air from the hydraulic fluid is an aluminum alloyfloating piston. The assembly also contains a variable orificeassembly. Mounted to the piston and fork is the axle, wheel,and tire. The drag brace assembly connects the trunnion to theairframe structure and pivots at the hydraulic actuator mount-ing point for retraction and kneeling. Both sections of thedrag brake assembly are made from 7075 aluminum alloy. Stan-dard fasteners and self-lubricating bearings and bushings areused throughout.

For a nonretractable tail landing gear, the configuration andcomponents are the same as those fur the retractable conceptexcept for the removal of the hydraulic and electrical systems.

19

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

7 Aircraft

Nr

Hydraulic Actuator,• (Locked, Kneeling)

"-----" --- iBrace

•___ - •.• ... /Shock

Locked___

retracted

Kneeled PositionWL 56.9"'".,

Comn-pressed Position.+ _--ýW +.%41.9

Static Ground Line - WL 35 •15°0•--- Extended

Sta. 450

Figure 9. Baseline Tail Gear, Retractable

"20

DRAG BEAM -MAIN GEARSThe trailing arm (or drag beam) geometry as shown in Figure 10

is used for all concepts of the main landing gear. In additionto providing an attachment point for the lower stage shockstrut, it also mounts the axle, wheel, tire, and brakes. Ajack pad conforming to MIL-STD-809 and configured in accordancewith 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 crashworthydesigns is 300M steel with the section thicknesses being thesame for the normal and crash conditions. To mount the dragbeam to the airframe attachment fittings requires the insertionof the drag beam into the fitting at BL 10 and positioned 90degrees to the ground line. A bayonet fitting on the end ofthe beam locks into the airframe fitting when the beam isrotated 70 degrees to a static position. The drag beam isdesigned such that it can react the bending, torsional, andside loads imposed upon it during normal and crash impactlandings.

BL 10

REF

BEARING JOURNALS

5 BL24F REF

BRAKE MOUNTINGFLANGE• _ SHOCK STRUT

ZBAYONET FITTING AXLE JACK PAEND 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 employedto develop design mass properties. The CG (center of gravity)for the conf iguration is at fuselage station 288. 0 and thecorresponding moments and products of inertia are shown inTables 2 and 3.

MAIN ROTOR PARAMETERS

The aircraft assumed for this study is a modification of theSikorsky LHX air vehicle. The ABC main rotor is used and itsparameters are shown in Table 3.

LANDING GEAR PARAMETERS

Main Gear

The three baseline design configurations for the selectedaircraft landing gear are as follows:

1) A retractable gear designed to normal design operatingconditions (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 isacceptable in meeting the,-e requirements.

3) A nonretractable gear designed to crashworthy requirementsas stated in paragraph 2 above.

The load-deflection curve for the main gear tires is shown inFigure 11.

Tail Gear

The baseline tail gear design is a typical, retractable landinggear. Tail landing gear is designed to 20 fps landings.Figure 12 shows the load-deflection curve for the tail geartire.

22

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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-sec2 38207.

IY - lb-in-sec2 173924.

Iz - lb-in-sec2 174969.

Ixz - lb-in-sec2 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-

z0

w-JLL 1.5-w0w

1.0-

.5•

00 2000 4000 6000 8000 10000 12000

LOAD LB

Figure 11. Main Gear Tire Load Vs. Tire Deflection

25

CD

1 %

00

E-4

'44

NI N0103-130 3di

26o

NORMAL OPERATING CONDITIONS

Static Design (Landing)

Landing loads are developed for two categories of normallanding conditions. Limit conditions are done at a sink speedof 10 fps and a forward velocity of 50 knots. Reserve energyconditions are done at a sink speed of 12.25 fps and a forwardvelocity of 40 knots. Both limit and reserve energy conditionsare done at the basic design gross weight range using two-thirds rotor lift during landing. The landing gear is designedaccording to requirements set forth in MIL-S-8698, as well asadditional design criteria prescribed in specification ANC-2.A landing gear vertical extension position of two-thirds fromfully compressed serves as the application point for determin-ing maximum vertical reaction while a nine-tenths extendedposition is used to evaluate spin up and spring back. Normallanding requirements and the calculated ground loads to meetthe requirements are presented in Table 4.

Static Design (Ground Handling)

Loads data for braking, turning, pivoting, taxiing, towing, andjacking conditions are developed in compliance with MIL-A-8862.In every case, the gear static position serves as the loadapplication point. Ground handling conditions are done at thebasic design gross weight with the rotor lift equal to zero. Alist of the ground operations requirements and resulting loadsare presented in Table 5.

Normal Landing Requirements

For normal landings up to and including the reserve energycondition, the energy absorption system depends largely on thelanding gear stroking and tire deflection to attenuate theimpact loads without permanent structural deformation ormalfunction of the landing gear. In this case, the airframe isidealized as a rigid body and emphasis is placed on modelingthe landing gear characteristics. The actual KRASH model usedin this study is illustrated in Figures 14 and 15. An indepthdescription of the mechanics of the KRASH program is providedon page 33. For normal landing conditions, the KRASH programis used to evaluate gear stroke and efficiency. Associatedvalues of stroke and efficiency are thcn used as inputs to aSikorsky developed program of energy equations in order todevelop specification required landing and ground handlingconditions. 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 42fps using one-G rotor lift at basic design gross weight. Forvertical velocities up to 20 fps, the landing gear has beendesigned to attenuate the total impact energy while disallowingany fuselage ground contact or yielding of the airframe struc-ture. This criteron has been met for simultaneous fuselageangular alignments of ±10 degrees roll and +15 to -5 degreespitch. The landing gear and mounting system, however, mayexperience permanent deformation. For vertical velocitiesabove 20 fps and up to 42 fps, yielding of airframe structureis acceptable. Envelopes of sink speed vs. fuselage angularalignment with the ground, for which the crashworthy landinggear system has been evaluated, are depicted in Figure 13.

Further design requirements for the crashworthy gear system areas follows:

- Any gear upon crash impact shall not protrude intooccupied and/or fuel cell areas.

- A retracted gear shall provide some crash impactenergy attenuation.

- At vertical sink speeds of 20 fps to 42 fps, thelanding gear system shall provide for a limited 15%reduction in cabin space.

:-4

7.4

28

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LOL

E-4.

r- N

E-4 rX4 H0 W4 >.

UzH0 0 .0 000

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w "I C 4 -q HH HO)-

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0 10No

0 -,zN

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C, 0 0 0'~D~

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0- .(ýý

41r o a) O Co D 0

00

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-4

V 0- '-4 CC 3 )

;14 > 4 )

V~ ~ O.)i E0 H

C3

The light experimental helicopter (LHX) meets the objective ofproviding a high level of protection for its occupants insevere crash impacts with advanced landing gear systems andfuselage 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 verticalvelocity with 25 fps longitudinal velocity. The fuselagestructure and landing gear has been demonstrated to protect theoccupants and their living space in vertical crash impacts whenroll angles are limited to 10 degrees or if the impact is at 36fps with a 20-degree roll angle. The landing gear also meets

* crashworthiness requirements of decelerating the aircraft atnormal gross weight from an impact velocity of 20 fps onto alevel, rigid surface without allowing the fuselage to contactthe ground.

Crashworthy seats provided for both crew and passengers are allequipped with improved restraint systems. The crew seats meetthe crashworthiness requirements of USARTL-TR-79-22A "AircraftCrash Survival Design Guide", and the troop seats shall bedesigned to meet the above requirements also.

The crash impact conditions investigated using the KRASHcomputer program are impacts on level rigiu jround. The basicrequirement of vertical impact design conditions envelope is todemonstrate the capability of the aircraft to withstand verti-cal impacts of 42 fps without either a reduction of cockpit orcabin height of more than 15 percent or causing the occupantsto experience injurious decelerative loading. The envelopes ofpitch 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 geardesigns are designated by an asterisk (*). For all theseimpact conditions 25 fps longitudinal velocity has also beenincluded.

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 PITCHDESIGNATION 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

V0"42 FPS

ROLL ATT. DEG-20 -15 -10 -5 5 10 .15 20

"Figure 13. Pitch and Roll Attitudes in Crash Impacts

32

-- ..- - -5

KRASH ANALYSIS

KRASH PROGRAM INPUT

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

KRASH MODEL MASS PROPERTIES

The crash impact evaluation was conducted with the helicopterat its basic structural design gross weight of 10,000 poundsand a representative center of gravity at Fuselage Station288.6 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 (WaterLine) are presented in the Appendix.

In order to more accurately represent the structural members ofthe airframe and crushing of fuselage, use has been made ofnode points. These are points in space, each related to aspecific mass, and which move with the mass in a fixed relationto 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 19node points are presented in the Appendix.

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

The material properties included in the beam data are codedwith material code (MC) numbers. The properties associatedwith 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 TI L' • 17¶~ I 9.I ~ '.. i ~ ~ ~ ..I•,- • ,•i • *.. • ' .-ixi •Z<•I• •i• • :i• •A~ l R • I K 1 -'2 W2 . I

KRASH MODEL SPRING PROPERTITS

The KRASH program requires that those portions of the aircraftthat 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 thelanding gear or fuselage underside that they simulate. Thesprings 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 ofcrushable structure is such that a region is reached where theconfined crushing is very significant and the stiffness increasessubstantially.

A typical load-deflection curve is shown in Figure 16; it repre-sents a combined load-deflection of three different stiffnesscharacteristics in series.

SF2 BOTTOMING

2j2 I LB/I5I I j

0.

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 hasbeen used for all the spiings to represent deflated landinggear tire and fuselage crushing.

In the KRASH model, the main landing gear simulation includesthe air-oil upper and lower oleo struts and the articulation ofthe gear that occurs during gear stroking. The tail gear hasalso been modeled as air-oil strut up to 20 fps verticalvelocity impacts. Beyond 20 fps it has been modeled as avertical spring.

The oleo input data given in the Appendix are used in theKRASH model.

The pilot seat and the 50th percentile occupant have beenmodeled in the KRASH analysis. The stroking seat has beenrepresented by a nonlinear member which deflects elastically0.75 inch then strokes at 14.5 'g'. The occupant and seat wereincluded in the analysis to develop the seat stroking distanceand then determine whether the seat had exceeded a maximumstroke of 12 inches.

LIFT FORCES

Masses in the KRASH model have lift loads applied to them equalto 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 impactavailable from a single KRASH computer program run results in600-800 pages of printouts. Selected data from nine impactconditions studied for the design of the landing gear aresummarized below.

TimE of Zero Vertical Velocity at Aircraft C.G.

The variation of beam ±-ads, inertial forces and fuselagecrushing is of significance until the vertical velocity at theaircraft center of gravity has dropped to zero. The time atwhich this occurs for each of the conditions studied is in-cluded in Table 7. Any loads developed by the KRASH programbeyond the time of zero velocity is the result of the crushedstructure still behaving as elastic structure. This behaviorthen causes excessive rebounding and additional loads.

"37

C 00 00 0

04

.4m 00 O LI N%

u o Lfl 00 m

CmNC1 H H H-

E-4

H 'LI) 00 N-

H- N LI) OLf LAIliN HI 0%

H 1 NLA AL 00 k~ N H- N-

C14

r-4H

-*.r -r-4 - ..

$i 4-) HM 1)iUo > iý4~ r.aHr4-) a)

4J -H> 7 U9-4 0 -P

o Q) 4-.)U> H-

38

Landing Gear Strut Maximum Stroke and Time of Occurrences

The main landing gear for the advanced technology helicopterconsists of a two-stage air-oil oleo system. The lower stageis capable of stroking 14.50 inches while the upper stagestrokes 12.5 inches. For normal landirg and vertical. impactsup 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 andlower stage stroke with one set of orifice openirgs. Forvertical crash impacts up to 42 fps both stages stroke dgainwith added orifice areas. The maximum strut stroke and thetime 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 notconsidered. The time at maximum stroking given in Table 8corresponds to the time of zero vertical velocity at theaircraft 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 11.5Velocity Vert. 42 fps L.H. Upper 0.09 11.5

Long. 25 fps R.H. Lower 0.1 14.59L.H. Lower 0.1 14.59

00 Roll, 00 Pitch Tail Gear .1 23.34

LHX03 R.H. Upper .17 8.3Velocity Vert. 42 fps L.H. Upper .16 11.1

Long. 25 fps R.H. Lower .16 14.01L.H. Lower .16 14.54




5' Roll, 150 Pitch Tail Gear .1 26.2



100 Roll, -50 Pitch Tail Gear .14 15.16

LHX17 R.H. Upper .17 3.26Velocity Vert. -CG fps L.H. Upper .17 9.32

Long. 2S fps R.H. Lower .19 9.79L.H. Lower .14 14.05




100 Roll, -100 Pitch Tail Gear .19 15.5




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 .98Velocity Vert. 20 fps L.H. Upper .42 6.25

R.H. Lower .44 9.12L.H. Lower .384 13.28



R.H. Lower .18 12.59L.H. Lower .18 12.59


GROUND LOAD

The landing gear structure has been designed for the dynamicload obtained from the KRASH analysis. Typical time histories ofground loads are shown in Figures 16, 17 and 18. As the geararticulates, the ground friction causes drag and side loads onthe gear. The ground friction coefficient of 0.34 has beenused in the analysis. Ground loads for the landing gear systemare 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 contactedthe ground and the helicopter is pitched forward. At 5 milli-seconds, the main gear contacts the ground. Between 50 and 60milliseconds an aft drag load is developed as the main geararticulates vertically and aft. As the assembly is moving aft,friction between the tire and ground develops which produces aforward acting drag load. The peak drag load is developed atapproximately 70 to 80 milliseconds after the tail gear hascontacted the ground. The next peak drag load occurs atapproximately 179 milliseconds, which is 10 milliseconds afterthe helicopter's vertical velocity at the 6.6 is zero as shownin Tables 7 and 8. Load developed by the KRASH analysis arenot considered after the C.G. velocity is zero.

41

- - - - -

Figure 18 shows the main gear side load at the ground as afunction of time. The helicopter is assumed rolled to theright. At approximately 60 milliseconds after the tail gearhas contacted the ground, a side load on the right main gear isdeveloped to the left. The aircraft is now rolling to the leftand a large side load is developed in the opposite direction at90 milliseconds. Lateral rebounding then develops until thehelicopter's vertical velocity is at zero.

Figure 19 shows the vertical load of the main right gear at theground as a function of time. Again, 50 milliseconds aftertail gear contacts, the main gear begins to develop maximumvertical loads. At approximately 80 milliseconds, the peakload is obtained. The helicopter is rolling to the left andthe left hand gear contacts the ground. Again load data beyond160 milliseconds is not considered.

Peak ground loads for the landing gear system are shown inTable 9. The peak loads, drag, side and vertical, are over a10 to 20 millisecond time period since the time response toimpact loads of all the materials in the gear vary. The loadsof KRASH condition LHX03, for example, were obtained fromFigures 17, 18, and 19 at 10 milliseconds for the verticalload. A similar method was used for the other KRASH conditionsof Table 9.

42

TABLE 9. GROUNE LOADS DATA

MAIN GEAR

KRASH Drag Load Side Load v7ertical LoadConditions (+ Aft) Lb (+ Left) Lb (+ Up) Lb

LHXIIVelocity Vert. 42 fps -16,000.0 -5,000.C 55,000.0

Long. 25 fps00 Roll, 00 Pitch

LHX03Velocity Vert. 42 fps -20,000.0 -15,000.0 60,000.0



Long. 25 fps50 Roll, 15' Pitch

LHX13Velocity Vert. 42 fps 21,000.0 2,000.0 62,000.0

Long. 25 fps10' Roll, -5' Pitch



LHX18Velocity Vert. 36 fps 11,000.0 1,500.0 32,000.0

Long. 25 fps100 Roll, -100 Pitch



TAIL GEAR

LHX30Velocity Vert. 20 fps -6,580.0 -1,390.0 20,015.0100 Roll, 150 Pitch

LHX31Velocity Vert. 20 fps00 Roll, 00 Pitch 5,880.0 0.0 17,300.0

43

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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 ofthe engine mass. Structures to retain these masses are de-signed to the inertia load factors obtained from the KRASHanalysis. Since all these high mass items are located awayfrom the living space, in a severe crash impact it is veryunlikely that these masses will penetrate into the livingspace. The maximum vertical load factors experienced by someof the representative masses are surmarized in Table 10.Typical time histories of accelerations of high mass item 3 areshown in Figures 20 and 21.

Figure 20 is a plot of the vertical acceleration at high massitem 3 as a function of time from the iime of initial groundcontact at a level attitude and 42 fps vertical, 25 fps forwardvelocity (Cond. LHX 1). As the vertical ground load on thegears are developing, vertical accelerations on the high massitems are also being developed. The peaks in the accelerationplot are caused by the structure flexing.

Figure 21 is a plot of the force and aft accelerations at highmass item 3. This is due to the forward velocity at groundcontact which is driving the wheel aft, then friction loadsdevelop. This motion of the wheel causes a reversal of thefore and aft accelerations.

Table 10 is a summary of vertical load factors at the cockpitcabin, fuel for all areas, and the transmission for each crashcondition specified.

47

H 0 0 0 0 N N- m N- m0 c%D lf) 00 LO N 14 N m) N Id Nm N H N

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CONDITION LHX I

0-10-

-20x0-3

0 20 40 60 80 100 120 140

TIME - MS

Figure 20. Vertical Acceleration Time Historyon High Mass Items

49

CONDITION LHX I15-

y~-. 10-

i~i--

0

x .

-10-

,,\•, 15

0 20 40 60 80 100 120 140

TIME- - S

[71 Figure 23. Longitudinal Acceleration Time Historyon High Mass Items

. "2- 5

k 5.0

"SEAT STROKING

The vertical stroking capability of the cockpit seat incluledin the KRASH model is simulated by beams that yield at aconstant load after a small initial elastic deflection. The

seat cushion has been represented as a soft elastic member, anda 50th percentile man with a Dynamic Response Index (DRI) spinehas been used as an occupant. The seat is capable of stroking12 inches at 14.5g. The seat, required in the 42 fps impact,strokes 9.98 inches, which is consistent with the cockpitcapability. To find the level of protection provided to the

occupant, a comparison between KRASH floor pulse and seatdesign pulse has been done. In all cases the velocity changeis less than 50 fps. A typical floor pulse with superimposedseat design pulse is shown in Figure 22. Table 11 summarizesthe cockpit seat stroking data. The 9.98 inches of seatstroking 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

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53

-,oL

FUSELAGE CRUSHING

The amount of crushing of the fuselage underfloor structure ispredicted ly the KRASH computer program in terms of deflectionsof 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 areused 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, verticalsine 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 12springs. The locations of these springs in the airframe areshown in Figure 15. The amount of fuselage crushing predictedin two crash impact conditions is illustrated in Figure 23. At42 fps level impact fuselage, crushing is less than 3-3/4inches. When the crash zone cxceeds 3-3/4 inches, the graphiteupper structure will crush locally to a point beyond whichthere will be local structural failure. In a 36-fps, 20-degreeroll,and 10-degree pitch condition, before the crush line goesabove the floor, there will be secondary structural failure oflateral beams, "parallelo-gramming" of the fuselage, and theremainingi structure will withstand the crash loads. Thus 85%of the li.ving space shall be maintained in a severe rollimpact.

Figure 25 presents the results of crushing tests conducted onbeam members with webs of Kevlar epoxy or aluminum. FromFigure 25 a 4-ply Kevlar sine wave web is shown to have crush-ing capabilities (load deflection) which are better thanaluminum 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 thestudy for each of the three baseline landing gear systems. Theanalysis was performed during the development of the KRASHmodel to provide model input data and establish baselineweights.

STRUCTURAL LOADS (MAIN GEARS)

The noncrashworthy main landing gea_ was analyzed for the MainGear Obstruction Ground Loads shown in Table 4. The shockstrut 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 thenoncrashworthy drag beam with a crashworthy drag beam.

The shock struts and drag beams were sized based on handcalculators. Loads were calculated for the crashworthy compo-nents assuming the critical design condition was at 42 fpsvertical, 25 fps longitudinal, and 10 degrees roll and pitch.The shock strut load developed was 50,390 pounds. Figures 28and 29 present the results of those calculations. Maximumbending 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 preliminarycrashworthy landing gear designs. The program developed designdata such as ground loads, shock strut loads, shears, andmoments in the drag beam, load factors, aircraft attitudes andgear geometry during a crash sequence. Table 13 presents themain gear ground loads, also given in Table 9, and the result-ing shock strut loads. It can be shown from Table 13 that theoriginal hand-calculated strut load of 50,390 pounds was withinthe range of loads obtained by the KRASH analysis. Maximumload was 55,000 pounds, minimum of 47,900 pounds from the KRASHanalysis. The loads from the KRASH analysis substantiated theoriginal baseline design of the main gear.

59

TABLE 13 GROUND LOADS AND STRUT LOADS - MAIN GEAR

KRASH Drag Load Side Load Vertical StrutConditions (+ Aft - (+ Left - Load Load

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

Velocity Vert. 42 fps -16,000.0 -5,000.0 55,000.0 47,900Long. 25 fps

0' Roll, 0' Pitch

.LHX03

-' Velocity Vert. 42 fps -20,000.0 -15,000.0 60,000.0 48,900Long. 25 fps

100 Roll, 100 Pitch

LHX12Velocity Vert. 42 fps -22,000.0 -10,000.0 70,000.0 54,000

-- Long. 25 fpsS5 Roll, 150 Pitch

LIA13Velocity Vert. 42 fps 21,000.0 2,000.0 62,000.0 55,000

Long. 25 fps100 Roll, -5° Pitch

Average Strut Load 51,450

- -1

60

2008320083

(LBS i(LBS)

095

Aý 9416 A- 1222

I= 1173 1- 25638

3250 30= 1116 419 144b

D/i= 342 D/t- 441

10 do 7075-T73

F SECTION A-A7075-T73

SECTION B-8

"Figure 26. Shock Strut - Noncrashworthy, Kneeling,Retractable Main Landing Gear

K4

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

SECTION D-DSECTION C-C 087

D/t- 37 36A- 8645

P10822

SECTION E-E

Figure 28. Shock Strut Assembly, Crashworthy, Retractable,Kneeling Main Landing Gear

63

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64

TAIL GEAR

The tail gear was analyzed for a level 20-fps sink speed andnormal landing conditions. The trunnion cylinder and dragbrace were analyzed for the 20-fps crash condition. The axleand piston/fork were analyzed for the normal loads. Figures 30and 31 present a summary of the tail gear structural analysis.

BASELINE LANDING GEAR WEIGHT

Table 14 summarizes the weights of the components for eachbaseline landing gear. The weights for wheels, tires andbrakes were obtained from catalogs. Weights for the shockstruts, axles, drag beam, fittings, hardware and oil werecalculated from drawings. Table 15 compares the baselinelanding gear weight with the landing gear systems of otherSikorsky helicopter mcdels.

The percentage of gross weight for the landing gear system ofthe noncrashworthy desagn compares well with the percentage fora noncrashworthy commercial S-76A design and a noncrashworthydesign for the HH-3E military helicopter. The crashworthydesigns of this study appear reasonable when compared to theUH-60A. The crashworthy design for the ACAP (S-75A) appears to"be out of line. The larger weight for the crashworthy S-75A isin the shock strut and drag beam which is the result of thegeometry for a main gear aft of the c.g. and attached to thenarrower transition section.

The baseline crashworthy drag strut, shown in Figure 29, wasK:: calculated to weigh 76.6 pounds. This weight appeared too

heavy when compared to the drag strut of the UH-60A shown inTable 15. A review of the detailed structural analysis of theUH-60A drag strut showed that the strut had been designed for anormal ground obstruction condition. It was then assumed thatthe wheel, axle, and tire w#ould be destroyed during a crashonto a rigid surface. The crash loads were then applied to the

A+ strut at the location of the base of the axle. The result wasthat 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 landinggears.

65

. - -"

RETRACTING CYLINDER

ATTACH PINS (1"250 x 095)

S• • P3.5"N-FNG

8 A 47 7-T36 DIA

b SECION --B •255

DRAG BRACE 410

475 I A A.

PISTON-FORK

A

80 ,A S LUGS

100R=130

t-50

SECTION B-8 125 AXLE (1 492 x 0701 9" LONG4340 HI T 200.000

-170

L 40'90O35

a SECTION A-AE-E

100O325 SECTION D--

100 f 30020 0

C C

250SECTION C-C12

-r---------SECTION E-EPISTON -FORK26

300 M H T 280000Fl

SECTION F-F

Figure 30. Tail Gear Structure

66

'- 36

R LOWER LUGL:

R= ~R .625 RETACIN CYLNDE

t= .5.6

4J4.

">".:,•,•'•::' /, •UPPER LUGS:

• • ""D-- .5625 RETRACTING CYLINDER

.•.;::.t7- .44-•''>AIRFRAME (REF.)

S• •[--.180 (TYP')

I 7075-T73

,150

-2.0 1-

SECTION G-G (TYP.)

Figure 31. Tail Gear Drag Strut

67

COMPOSITE LAN•DING 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 beamb. Oleo extension fitting (noncrashworthy gear)c. Upper fitting, main gear/fuselaged. Tail gear drag brace

A preliminary analysis of the main gear drag beams for thethree 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 oleoextension fitting compared to an aluminum fitting.

A composite upper fitting, which attaches the top of the maingear to the airframe, was estimated to result in a 60-percent

weight saving compared to a steel fitting. The weight estimatewas based on the bearing strength density ratio (FBRU/P) for125,000 psi heat treat steel and ±450 graphite/epoxy.

A preliminary analysis of a composite tail gear drag braceresulted in a 47% weight savings compar-d to an aluminum brace.

The lower weight savings for a composite main gear drag beamwas the result of providing torsional stiffness to the beam toprevent excessive "scuffing" of the main gear tires.

A graphite/epoxy main gear wheel was considered but was notfeasible due to the heating of the wheel during braking. Ametal matrix main gear wheel appears feasible, but it is notclear that the technology will have matured sufficiently to becaptured by an accelerated LHX acquisition strategy.

Although composite materials appear to reduce the weight oflanding gear structures, design problems must be overcome. The

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 beconsidered. Possible fiber breakage could cause leakage of thesystem due to the high pressures in a cylinder. Repeated highimpact loads on epoxy matrix materials may result in matrixsplitting.

68

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WEIGHT SENSITIVITY ANALYSIS'.4A preliminary weight sensitivity analysis was conducted toevaluate the effect of various sink rates on the components ofa crashworthy system. The system included crashworthy crew/troop seats, crushable fuselage understructure, and the Wainlanding gears. The weights of seats systems, post-crash fireprevention, emergency egress, airframe crashworthiness and maingears were investigated by Sikorsky under a CrashworthinessDesign Parameter Sensitivity Analysis study conducted for theApplied Technology Laboratory under contract DAAK 51-79-C-0043.The results of that study were reported in Reference 2. A moredetailed analysis was conducted on the main landing gears toevaluate the effect of pitch and roll attitudes.

Ci'ASHWrWHY SYSTEMS

At present, creshworthy seats are designed independent of thelanding gear. This is ,lone to provide crew protection shouldthe heli-opter, for example, crash into the water. Under the

Reference 2 study, weiqnts for crew/troop seats were dex elopedfor various levels of zrashes up to 42 fps sink rates. Table16 summarizes the results of that study.

The study of crew/troop seat weights showed very small weightincreases when going from 20 fps capabilities to 42 fps capa-bilities. For the crew seat, the weight was 39 pounds at20 fps 'nd 12.5 pounds at 42 fps, a change of 3.5 pounds. -Forthe troop seat, the change was 7.73 pounds. At 20 fps, thetroop seat was 11 pounds. For the sink rates above 20 fps to42 fps, the weight remained unchanged at 18.7 pounds. Itappears from the study that very small changes in the weight ofa crew/troop seat is obtained due to large increases in sinkrates. Tae weight chanjes are considered insignificant for thislanding aear study.

Reference 3 concludes that a properly restrained human canwithstand the lateral and longitudinal loads Fssociated with;urvivable cras1' puls(s. Therefore, no load attenuation isnecessary in the.e directions. Based upon that conclusion, thecrew seat can be expected to perform in the crash attitudesspecified for tCe design of the larning gears. Also, since theseat is designed indepepdent of the landing gear, the seat canlimit the load on the oocupants, with gear retracted, as shownin 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, 000pounds for each crash condition. The maximum seat strokeduring the crash conditions is 12 inches. The duration of theseat loads are not injurious to the pilot. The peak loadfactor is approximately 15 g's for a 200-pound occupant.

.,

TABLE 16. CREW SEAT/TROOP SEAT SYSTEM WEIGHT

Crew Seat SystemCrash Level I II III IV

S-76 SH-3D Lamps

Seat Structure 19.8 lb 30.0 lb 33.5 lbCushions 7.0 lb 4.0 lb 3.9 lbRestraint System 3.2 lb 5.2 lb 5.1 lb

Total 30.0 lb 39.2 lb 42.5 lb

Troop Seat SystemCrash Level I II III, IV

CH-53 RH-53 UH-60A- (Troop) (Crew Chief) (Troop)

Seat 4 8 lb 10.) lb 15.55 lbRestraint 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

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A crushable fuselage under structure is not requited for sinkrates at 20 fps or less based on the requirement of this study.Beyond 20 fps, the amount of crushing structure depends on thedepth of the structure below the floor and the sink rate. Forthis study, the depth of the structure below the floor isapproximately 7 inches. Three and three-quarter inches areused to provide crushing structure. Four inches are requiredfor primary structure.

The weight of airframe crashworthiness was evaluated in Refer-ence 2. Table 17 summarizes the results. For this study, theweight of the crushable fuselage under structure is consideredpart of earth plowing and longitudinal impact structure ofTable 17. It is estimated that 70 percent of the earth plowingand impact structural weight can be attributed to crushablestructure. The aircraft weight of this study and that ofReference 2 are comparable.

The KRASH program evaluated the effects on the airframe forlevel landings, gear retracted. Figure 35 shows the result ofthose conditions Figure 35 indicates that the airframe verticalload 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. TheMIL-1290 (AV) requirement is an ultimate vertical load factor of20 G's (average). Also from Figure 34 it is apparent that thereis insufficient depth of crushable structure for level 25, 30,and 36 fps gear retracted crashes. It is noted that the loadson the pilots seat limit the pilot to approximately 15 G's fora 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 35Retention

2) High Strength/ 0 10 15 15Equipment Attach.

3) Earth Plowing 0 5 10 15Capability

4) Longitudinal 0 5 10 15Impacts

TOTAL WEIGHT 0 30 70 85

75

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Vertical load factors for the high mass items (rotor head,transmission, etc.) were developed by the KRASH program for thecrash condition with the gears retracted. Figure 35 presentsthose load factors for a level condition. The 35 foot-per-second sink rate results in a load factor (67 Gs) which is fargreater than MIL-STD-1290 (AV) requirements of 40 Gs peak, 20Gs average Sink rates greater than 25 feet per second, gearretracted, would result in the high mass items breaking looseduring the crash sequence. Load-limiting devices or heavierstructure would be required for the higher sink rate with the"gear retracted. This requirement is beyond the scope of thisstudy.

A study of crashworthiness effects with the gear retracted mustinclude the fuselage structure for attachment of seats, highmass items, fuel cells, backup structure for the landing gear,and retention of cargo.

Reference 2 also evaluated the weight trends for crashworthyfuel systems and main gearboxes. The results are shown in"Figure 36.

12F

77

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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 varioussink rates up to 42 fps.

A more detailed weight sensitivity analysis was conducted forthis study to evaluate the effect of various sink rates atvarious angles of pitch and roll on the landing gear shockstrut, drag strut, and fuselage fittings. The other componentsof the landing gear system are not designed by crash loadconditions. The KRASH program was run to develop main gearcrash loads at 42, 36 and 30 feet per second sink rates atcombinations of pitch angles and roll angles given on pages 8and 9. A pitch angle of -7.5 degrees was also used to deter-mine where the peak loads in the shock strut would occur. Theloads on the shock strut were obtained from KRASH for a sinkrate of 42 fps and the helicopter pitch and roll angles shownin Table A-1. The loads are plotted in Figure 37 and shown inTable A-1. A peak load was obtained at a nose down pitch angleof 5 degrees and a roll of 15 degrees. Shock strut loads wereobtained for sink rates of 36 and 30 fps. Again,maximum leadswere at a nose down pitch angle of 5 degrees and a roll of 15degrees. The weight of the shock strut is determined based onthe thickness of the structural load carrying components in thestrut. The thickness required to resist the loads imposed arebased on the strength allowables for the materials or onmachine shop limits.

Figure 37 also presents the bending moment and weight of thedrag strut based on the KRASH program. Also shown in Figure 37is the baseline moment used to design the baseline drag strut.The KRASH program appears to develop the proper loads in theshock strut but overestimates the bending moment of the dragstrut. The baseline drag strut was sized based on plasticbending allowables for steel. The bending stiffn'ss of thebaseline 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 notanalyze a plastic hinge.

80

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

BASELINEP50 P 7.5r

" 0 - FPS P=150-1 30

60 E=30 20P1 50 -

50 -

4cT 20 15-2

0-40

0 10<ROLL =150 42 FPS

-5 0 5 10 is

PITCH ANGLE "- DEG ROLL ANGLE ~- DEG

AXIAL LOAD ON SHOCK STRUT VS. PITCH AND ROLL ANGLE

FROM KRASH

60 -E =30 X i06 Psi so51 50 •-40

WEGH - BASELINEJ' •I• WEIGHT IZ 40 --

LBS. N;OI-/z 30 3-- Z

E 3 X 10 6 PSI DESIGNED0 20 BY

20 NONCRASHWORTHY

z CONDITIONS1a 10" 10mu I

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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 fuselagefittings. Table 18 summarizes the weight of the components inthe retractable and fixed shock strut and those componentsaffected by load.

The shock strut is an axially loaded structure; therefore, theweight of the affected components is assumed, in this study, tobe proportional to the shock strut lcads of Figure 37 and thebaseline 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 theshock strut load.

Figure 39 summarizes the fixed landing gear system weight as afunction of pitch angle.

A review of the landing gear system weight shown in Figure 39and Table 13 was performed to determine the factors that causethe changes in weight. From the earlier study (Reference 2),the weight of the landing gear was expressed as

^/DwG\8WLG= 62 .86 DG K L(100/0)

where DWG is the helicopter design gross weight and KLG thelanding gear weight factor. For various sink rates, the changein the weight of the landing gear is primarily dependent on theenergy that it must absorb; the weight coefficient is expressedby

b V2 LDK =a+LG 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 gearcomponents for wheels, tires, brakes and hardware. The con-stant "all was found to be 0.6861. The second expression giresthe 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. Forexample,

Noncrashworthy weight 244.9+73.3-26.6-22.2 = 269.4pounds

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 = 249.1pounds

The weight of the energy absorbing system for this studybecomes

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

This agrees closely with the weight of the shock struts shownin Table 13 for the crashworthy landing gears.

It appears that the weight of the energy-absorbing system inthe landinq gear is affected more by velocity than by attitude.The maximum load in the shock strut would occur when the shockstrut is perpendicular to the ground.

83

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U) 86

LANDING GEAR SYSTEM COST

The cost of the landing gear system is based on actual costdata of the UH-60A Lot I, fixed landing gear. Table 19 pre-sents the cost of that landing gear. Figure 40 presents thedata of Table 19 for the cost as a function of weight.

Figure 41 presents the cost of the fixed landing gear system at42 fps.

Based on the weight and costs shown in Figures 39 and 41;trending data was developed for the weight and cost as afunction of sink rate. Tile trend data is shown in Figure 42.Also shown in Figure 42 is a comparison of this study and theresults of a 11CRASHWORTHY DESIGN PARAMETER SENSITIVITYANALYSIS", conducted in 1980, for the SCOUT helicopter (Refer-ence 2). The SCOUT landing gear system is based on a designgross weight of 8683 pounds. Increasing the landing gearsystem weight by the ratio of the LHX design gross weight(10,000 pounds) results in a landing gear system weight veryclose to this study.

The SCOUT landing gear system cost shown in Figure 42 isupdated to 1984 dollars assuming an annual inflation rate of 5percent. The SCOUT cost is increased further by the ratio ofdesign gross weights. The result is a cost which is comparableto this study.

87

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

Cum. Avg. Cost ForWeight 1000 Units $/Lb @ T1000(lb (Thousands) (Dollars)

271.2 36.70 $135.3296.2 39.63 134.8321.2 42.53 132.4346.2 45.40 Cost = .2793 131.1 $/lb 288.67371.21 48.24 (WGT .8707 130.0 (WG) - .1348396.2 51.05 r = 1.00 128.8 r = - .006421.2 53.85 127.8446.2 56.62 126.9471.2 59.37 126.0

NOTES:

UH-60A Calibration Point (based on UH-60A Lot I ActualCost Data; 42 fps, fixed gear).

2 Costs represent Labor, Material and Factory Overhead, butno General OH or Fee ( 1.3 x cost, if desired).

3 FY '84 $, Constant

Methodology used: RCA Price System using UH-60A ActualCost Data.

88

70 -

U)•" COST =.2793 (wgt)**.8707I

00

U. so - . . .U./0

z 0 40- - - - ----------

0

30

00300 400 500 600

WEIGHT - LB

Figure 40. Landing Gear Parametric Cost Study

5 4 t--- --- ~~52

- -

• " 48 -

-46

4 o - I 100S42

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

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AERODYNAMIC DRAG

An aerodynamic drag estimate was made for the retractable andfixed/extended landing gear configurations. The power loss dueto drag is 7.2%/ft2. For the extended gears (main and tail)"the power loss is 26.6%. Retracted, the loss is 0.7%. Thedata in terms of square feet of drag area is as follows:

Landing Gear Type

Retracted Fixed/Extended

Main Gear 0.1 ft 2 2.3 ft2(both sides)

Tail Gear 1.4 ft 2

Total 0.1 ft2 3.7 ft2

A Drag +3.6 ft 2

(fixed-retracted)

The drag estimate ' for the components of the fixed/extendedmain gear are as follows:

Aerodynamic Drag, main gear components

Shock strut = .17 ft 2 /side (15%)Drag beam = .16 ft 2 /side (14%)Tire = .28 ft 2 /side (24%)Brake and axle = .15 ft 2 /side (13%)

Component total = .76 ft 2 /side (66%)Interference dragbetween shock strutand drag beam = .39 fc 2 /side (34%)

Total drag 1.15 ft 2/side (2.3 ft 2 maingear system)

The percentages of drag of the components in the extended tailgear are similar. The drag estimate for a "clean aircraft" is11-12 ft2.

The drag estimates have been based on the following assump-tions:

Aircraft weight 8000 to 10,000 lbForward velocity 200 knotsPower 1200 HP

91

The following is noted.

"" The main gear tires, brakes, and axles are the same forall gear concepts, therefore the drag values will remainunchanged.

* The outside diameter of the main gear drag beam is notchanged for each concept, therefore the aerodynamic dragvalues will remain unchanged.

* The outside diameter of the lower piston of the shockstrut for each main gear concept is not changed, thereforeany changes to the wall thickness of cylinders will have anegligible effect on the overall drag of the main gear.

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

92

RECOMMENDED DESIGN CRITERIA

Based upon the sensitivity analysis the following designcriteria 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 thelanding gear stroked at pitch angles between 20 degrees nosedown and 15 degrees nose up, and roll angles ranged from ze..odegrees to 15 degrees. Sink rates ranged between 18 fps and90 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. Ata nose down pitch angle of 20 degrees (V = 20 fps, H = 0), thecrew seats stroked; however, the troop seats did not. The resultwas 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 helicopterstudied, the criteria for attitudes and sink rates to developlanding gear loads should be as follows:

* The peak design load in the shock strut occurred at a nosedown pitch angle of 5 degrees and a roll of 15 degrees forall crash sink rates studied. It is therefore recommendedthat the criteria of 5 degrees nose down to 15 degreesnose up be used for crashworthy designs.

• The bending movement in the drag beam, assuming elasticbending, occurred at a pitch angle of 15 degrees androlled 15 degrees. It is recommended that the criteriafor a drag beam include plastic bending for roll angles of0 to 15 degrees.

Based on the loads developed for 42 fps sink rates com-bined with a longitudinal velocity of 25 fps, the criteriaof 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.

9

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% 95

DETAIL DESIGN CRITERIA FOR THE MAIN LANDING GEAR

The wheel should be assumed to remain attached during the crashsequence. A review of the structural analysis of the UH-60main gear showed the drag strut to be designed by normal groundconditions (Reference 4). The loads were applied at the groundline with the gear two-thirds extended. For the crash condi-tion it was assumed that the wheel had fractured and the loadswere applied where the axle attached to the drag strut. Thecrash drop test conducted on the UH-60 main gear resulted inhalf the wheel remaining on the axle (Reference 5) after thewheel rim contacted the test rig platform. The tire should beassumed destroyed.

The drag strut should be designed as a structure in plasicbending during the crash sequence. However, the current KRASHprogram 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 droptest of the UH-60 main gear. A 20-millisecond time frameresults in vertical reactions far less than expected. Usingpeak loads at 1 millisecond resulted in a vertical reactionslightly greater than expected. A 3- to 5-millisecond timeframe appears reasonable.

The KRASH program appears to provide the proper shock strutloads.

96

DESIGN UPDATE

The baseline fixed and retracted crashworthy main landing gearswere sized based on an axial load, in the shock strut, of 50,390"pounds, and the obstruction loads of Table 4 applied to thedrag beam. The preliminary structural analysis for the majorcomponents of the shock strut with margins of safety less than15 percent is discussed below.

SHOCK STRUT STRUCTURAL ANALYSIS

Shock Strut Assembly: (Crashworthy, Kneeling & RetractableGear) Reference Figure 26.

Cylinder Bore 4.00 Dia. A= 12.566Piston O.D. 3.25 Dia. A = 8.296

4.270

Upper Piston:

D/t = 37.36A = .8645 Pc = 50,390 lbI = 1.0822 Press = 50,390= 6074 psip = 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 .4132(.087) 260,000

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

Shock Strut Assembly: (Crashworthy, Kneeling, Fixed) ReferenceFigure. 27.

Upper Cylinder-Pc = 50,390 lb

A = 2.9166 Combine axial load withI = 6.5119 system pressure loadingp 1.494 of 4500 psi (ULT).B/t 20.2

97

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

Section J-J L/p = 65.77 = 441.494

Fc - 50390 -17277 Rc 17,277 .4022.9166 43,000

Ftn = 4500(4.0) = 40909 Rtn 40,909 .6712(.22) 61,C00

1MS + 4_.402'+.671-(_.402).671 -1 = +.07

Table 21 summarizes the margins of safety calculated for themajor components of the two baseline crashworthy shock struts.The large margin of safety for the lower cylinder of theretractable gear is the result of combining the cylinder withthe steel upper piston. Combining the two components into aone-piece machined structure caused the cylinder to be ofminimum wall thickness. The minimum thickness (0.083 inch wall)is required for the cylinder to be concentric. The minimummargin of safety for the drag beam designed to the obstructionloads was 4 percent. The maximum was 9 percent.

98

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OPTIMIZED SHOCK STRUTS - BASELINE AND UPDATED DESIGNS

The following analysis optimizes the baseline shock struts andthe shock struts required for the peak load at 42 feet persecond -5 deg pifrh and 15 deg roll.

Baseline design load Peak design loadP = 50,390 lb P = 57,000 lbcB C

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=18.9Compression stress Compression stress

fc=50,390/.718 = 70,181 psi fc=57,000/.796 = 71,600 psiFc=172,500 psi Rc=.406 Fc=174,000 psi Rc=.411

Hoop tension Hoop tensionft= 6074x3.106 = 131,000 psi fz= 6870x3.09 = 132,676 psi

2 x .072 2 x .080

Ft= 180,000 Rt = .727

1M.S. = 1 _--40 T?1 M.S. = 0.004-.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 internalpressure. The cylinder is designed by an internal pressure anda resulting tension load on the cross section of the cylinder.The minimum wall thickness for machining a steel cylinder isbased on a diameter/thickness ratio, D/t, of 50.

Lower cylinder, section D-DO.D. = 4.16 in., t = .083 in., A = 1.064 in. 2 D/t = 50Pressure area = 4.27 in. 2 4340 STEEL 260,000 H.T.

100

Baseline pressure = 6074 psi Peak pressure = 6870 psiP = 6074x4.27 = 25,936 lb PT = 6870x4.27 = 29,334 lbTinsion stress

ft = 25,936/1.064 = 24,376 psi ft = 29,334/1.064 = 27,570 psiFt = 260,000 Rt = .094 Ft = 260,000 Rt = .106

Hoop tensionftn = 6074x4.00 = 146,361 psi ftn = 6870x4.00 = 165,542 psi

2 x .083 2 x .083Rtn = .56 R .63

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

Upper piston - section E-EO.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 lbPissure 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=33.10I=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 = .733F Hoop tension Hoop tension

"f = 6074(3.076) = 107,377 psi ftn= 6870(3.054) = 107,045th 2 x .087 2 x .098

Rth = .413 Rtn =.4121 -

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 along column. No pressure is applied inside the piston.

101

. . . ..

- N

B = 50 390 lb P = 57,000 lbSPCB ' Pet=.137 in., A=1.447 in2, D/t=25.5 t=.156 in., A=I.638 in 2 , D/t=22.4I=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 psir 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 uppercylinders (F-F and J-J) of both struts are of 7075-773 aluminumand 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 loadPressure = 6074 psi Pressure = 6863 psiP= 25,936 lb PT = 33,717 lb

"t=.190 in, A=2.50 in 2 t=.210 in, A=2.77 in 2

Tension stressft = 25,936/2.50 = 10,374 psi f = 33,717/2.77 = 12,172 psiF = 66,000 psi = .157 =R = .184

HooT tensionfth = 6074x4 = 63,940 psi f = 6863x4 = 65,362 psi

2x.190 2x.21FT = 61,000 psi RTH = 1.04 R = 1.07

""' MS 1 MSMSV 1 04l'+. i57'-.157(1.04) S 1 4• 07T+. 184,-.-184(1. 07V

= 0.03 = 0.02

The sleeve is designed to crush due to loads that develop fromsink rates greater than 20 fps or loads exceeding the maximum"load for the noncrashworthy gear shock strut. No change isrequired. The weight changes for the crashwortny baselineshock struts are shown in Table 22.

102

W (d a) 0

I-CO E- 4Val) 1- 10-1

0" UL) 0 Mc:0c- ý-D (n (4 01 r~- LA

0 0 r-4 Cf4cl ý

IX4

0

H (

-4> ý r-I U)i

u, ) E4 - %~

Hr t0ir-4 04 CZ - 0 Or" C:

mr 0 ) *- - OD co -)- r-. ON ., .- %

t ~ ~ ~ E- >4rh-

U) E-4 a

0) U))) U

Oat

000

a) U) '

030

WEIGHT SLBSTANTIATION

The weights of major components in a shock strut assembly areproportional to the square of the velocities at a fixed grossweight. The lower piston of the noncrashworthy shock strut,for example, is 2.8 pounds (Ref. Tablc 18). The weight of theretractable and fixed piston is 11.3 pounds (Ref. Tables 18).Therefore, adjusting for the margins of safety of 0,

11.3-.9 = 4.1 and 422/202 = 4.42.8-.29

The weight of the components in a shock strut assembly at a"fixed sink rate and gross weight are directly proportional tothe axial load in the shock strut as a result of pitch and rollangles. The weight is a result of changes in the wall thick-ness of pistons and cylinders as shown in the structuralanalysis. End caps, floating pistons, seals, lugs and bearingsare changed. Table 23 compares the wall thickness of thebaseline shock strut components adjusted for zero margins ofsafety and the updated shock strut. The original assumptionthat the weight of the shock strut is proportional to the load,used to develop the curves of Figure 38, is substantiated by thedata of Table 23.

TABLE 23. SHOCK STRUT COMPONENTS, WALL THICKNESS

BASELINE UPDATED RATIO(M.S.=0) (M.S.=0) UPDATED/BASELINE

Retracted/FixedLower Pistons t=.072 in t=.086 in 1.11Upper Pistons t=.087 in t=.098 in 1.13Upper Cylinders t=.190 in t=.210 in 1.11Lower 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 zeromargins of safety. Tables 25 and 26 present the ..eights ofthe shock strut components for the updated landing gears. Theweight changes of Tables 24, 25 and 26 are negligible for thispreliminary study when compared to the system weight of 380pounds.

104

ElE-j LO0 N 0C'4c 0 0 0 0'.3 c; 0- L o c

IH

* 4ii 4

En ) J0 ' '

Q) H4 U) 0J Hn (y 4OLOAD'l 4

U U

EH

U) a

)) 4J

A 0 4-)

(a (d 0 kD0 Hor- 0 0 0 0 0H4IcYr) 0S-4 4 H4-i) U E

UlUU) 4-)

000004a)a* i

1 ) 0 04-) 4-) 4-) 4J-~--P 4iJ aMW LO (n

En L -rq-H 1 >l--f>1H -H -r- P4134EnU)t (d '- 4-)

H- aL L ) 1) %D0 a)ý4 1 Mtpt~lb) ) ) 0 14-4

Q) u 0)) 4-J 4J 0 0 - -1- '0 d) (aVo 00u U -H 4 r4-rA-4Cfl) > LA C)IU)

*-r-4 4) 0 ýlýlý 4( 0 f a a > (L) > 4-4r- 4) (d(34-3 4- 00 4-) 4-H M) I 14 0

)iU) >1 al w )H0 0 a-rU ,0 a) 9 ý-

P~4a

04 4 - 400000 0- 4-) -H- 4-)0) 0: 0l w

105

0 4)

H4-) H 4-)r

"I PE- I'l -0 I A HLn HM~ HO1

-H LO iN r, Hf a% C'i m OO 1 0 0

+ + + + + + +

t-4 d

C) 0 Gl00 0 0 OL0H- 4..) -r4 00 tpQ a% I H CD - Lf) H

(4 0 r. ~ 0 e.JNH N

U H4-) (4-) 4- -) .4-ip4- 4-)H

0C) 3

0

4.' U)

U)

N CY 0' N1 0rI 0H- 0 00 ,Hf C0 r-4H H-

Hl 4i' 4-i 4-) 4-i - ) i - 4I..-) 4-)

e.- i - U)

1 -44

H-q-H W 0 ~0 0 0 ~0 04) U) HU J rI 4- H 4J HU- 4) Hr

Ul ý PQ En rA E r-IUl -i .- r-qaO E~a E-a4 U 4 ) P4 u) Na a

w 0 0 P4 04 '- 0 0~ 040

106

4- :VO -)HN 0O

1.C 4C ý 4'

a)0 ) Ln ko Jncn N kolv 0 0

UH z

r44

El) 4)4-1

L) Q) . . I . 11I111I1 II0 44 tac 0 00 C4 H

4-4 H H-

U)m

r-q

4-)

H (d. ul '0 0 'DCD0lI:: 4a( ) I I I I I . . .i It) 0 itPL ( 4-4 O000HH q44 H- LI'

MrC 4-1ý-44-)a1) 0

'00

A3 0 00)WUQ ) 0 04-P4-J4- 4-) H

4-) 4-) nU) En tLOrUED w -Hi-H- 1 >1( >1

-H r-4-HP41, U) U) M 4-)

>4M l M a ) '0 4)

a) u )M44- W r-I ko a) as1: 1:: L) 0 - a) > LA) u U)

4)I 0 ý-4 I4 -Hý 00rI 0 -

0) -irý4-) M M 4J 4-) 0 0 -) 4-J ClJ ) 1 ý4 CEll>1 U 1) a)()a() H Q) a) > r_ - a)

()V () 0 a) ()a) k) () 9) (va) 4) u ::

04:30rH HiP- 00040 0 4 0 -H 4-) -1 -P rZ)~~U Q: 0) w D3C 0 w

107

Aerodynamic Drag - Updated Design

There are no changes in the aerodynamic drag characteristics ofthe updated landing gear designs since the outside diameter ofthe exposed components remain unchanged.

Fabrication Methods And Processes

Cylinders and pistons are machined from forged aluminum orsteel as requirec. The wheels are of forged aluminum. Thedrag beam would be fabricated from two steel iorgings, welded,and machined to obtain a tapered thickness as required. Thecost reriains basically unchanged.

RELIABILITY

The landing gear reliability benefits from being designed tostringent crash survivability and ground flotation require-ments. The gears are designed for sink speeds up to 10 feetper second with a reserve energy capability of 12.25 feet persecond. The main gears and the tail gear are capable ofpreventing the fuselage from contacting the ground at a 20-foot-per-second sink rate. Both main gears are the articulatingtype which, because of the drag beam, enables the gear to climbover ground obstacles without becoming overloaded. The longwheel base is ideal for rough terrain operation and providesprotection for the fan shroud during flared and tail downlandings. The articulated gear eliminates oleo bending loadsto provide longer bearing life and reduced friction. The tireswere selected to satisfy the most adverse center of gravity andwill, 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 fornormal 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 landinggears permits easy replacement of wheels or brakes. Since the

gears are similar to the UH-60A the maintenance man-hour perflight 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 bemade:

1) The vertical impact design condition envelope at 42 feetper second sink rate can be extended from ±10 degrees ofroll to ±15 degrees of roll.

2) The maximum load in the shock strut occurs when thehelicopter is at a nose down pitch angle of 5 degrees atany sink rate above 20 feet per second.

3) The maximum weight increase is 77 pounds for a retractablecrashworthy landing gear, and 48.5 pounds for a fixedcrashworthy gear when compared to a retractable non-crashworthy gear. The major weight increase is in theshock strut of both crashworthy gears.

4) The aerodynamic drag of the fixed crashworthy main gearrepresents a power loss of 26.6 percent compared to 0.7percent 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 whenF the helicopter crashes with the gear retracted. However,

the load factor on the high mass items is excessive forsink rates beyond 25 feet per second.

6) The weight increase of a shock strut is more affected bysink 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 thebeam is allowed to bend plastically during a crash se-quence.

8) Composite materials such as graphite/epoxy cannot be usedin the main wheels due to high temperatures generatedduring braking. Broken fibers in a shock strut assemblymay cause loss of pressure during normal or crash land-ings.

109

RECOMMENDATIONS

The following recommendations are made for further research toimprove 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 bodygroup and the fuel system are the largest weight driversfor a 10,000-pound class helicopter.

2. Conduct design studies to allow the cockpit/cabin sectionof the airframe to pivot about a low point in the struc-ture as a section of the upper aft cabin is collapsing dueto a nose down crash, gears retracted.

3. Evaluate the effect of repeated high impact loads oncomposites with epoxy matrix.

110

REFERENCES

1. Meir, W. H.,"Application Of The Advancing Blade Concept ToThe LHX Concept Developmenti, Sikorsky Engineering ReportSER-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 IncreasedEnergy-Absorbing Seat Effectiveness, USAAVRADCOM TR82-D-42, Applied Technology Laboratory, U.S. Army Researchand 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, QualificationTest Report" Sikorsky Engineering Report, SER-70174,November 1978.

ill

APPENDIXWEIGHT SENSIVITY AND KRASH ANALYSES

Tables A-I through A-6 of this appendix present the weightsensitivity analysis for the crashworthy landing gear systemsat three sink rates and sixteen pitch/roll conditions.

Tables A-7 through A-13 present the input data for the KRASHcomputer analysis. This data is provided as reference dataonly.

113

0 1.0

0) o -Ci* nHo m',-vqta )-co r r

.- 4 c ; C '.t6J40 ý 1 1-j4 tt , q Of.WMI Wr - V rL

I _ C)cnm CY mmen C) n CV Y y

LfI

L-'U

-r X- Im

%w i-i CJC c 14m - Dq

LLL

LA-

I ~4Ji0 u .- ( . .

11- u +~ a + 1 +4 + 1 4 -

m01

V)

4C.4

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4 W ON -itO n r4r NL)

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<- C3 4-0)(i L

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cn 0)R44 4- .' -U

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tCVj

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a) .

. c o *,-W r-v- t L0o e ooo( Ol)Ll1 :4a ~ ca3 a uiE 1r~~ rc~~ vr)lr

Al ) rý- ý 6a k i- ;6 t i ýr

'. 4 J 4 JU. 4 C2 ý 119 k9 019 10 10 190ý10 10 919 k91

to-.-I cn CY~V) M V) ci MvC~ enC) ( m m Y

(a cJ C~j cliCj C.C~j~~j JcJ'i C1 % - N Cj 04 0

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ol l- Ur gal UY gag ggal Waggg D% r 4-z

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en CL %- . ) mC< L-~ m 41 0

LO .- U) * (l *IA 01.n 0

115r

en CA)ML nr- DC 0Ln H( 0L 90q l 0

0 r)

41 4.) n 4.. ,-,4C = -4G) C'LfCA4D% 00wUw w-40wU Wt4.Nw

P- a,

'. 4 J4 .)=-- WWI m~0O~~.C. D tt

N) WO). C -c *r De 4L m c ~ nC 0L

1 '4J-. j (ACACACV '.40 )C-: "; )Ca; d ;')ý4 " CVC' C'C6))6C 3C4J 4 ---'r- ivCVIA V)( -A r4 H H( Cki CA)4AHCA%. NNCJ N N CICI(. CJCJ

0 : W

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uI - n 0Ne (L) ) N- ; 0 ;* ;L;ý

Z - 3 d e ObJ H,. M CV4bC- iDC')r-0 ~i CV) 4

V4)0VI --

u- II 44J

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V)e nr CD wý 00 00cv0 CDr- 00 4t 0 H CD Cl " C C>

U- Cl mC CV .M M C 1 c0CjC4(jC4+

>- CiL- 4..0i.

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4J .J (a 4-,4.Z~~~~( 4) 00 0 0 0 0 0 400 i C

LJJ Lo 0 L aLoC L0aL a.04. 0Ll ý ~Lnl. - 4-U)J 0-4 c)f r-40 r-4~ r-4-I0 0N04 H0 r- 4

(. N 0

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a)- min (DL ) iL LA-

U. H Wu U

C. 4. r

CO r 4N CV) C"L -C C n

7*4 C Ci * 116

r x - - - ~ - .*.- - .

ff 0-OJ 3

cm4 d-C 1 r- n

0- V

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0l )0)LL 4*- : t t t t t0 Vt-Wt I WWW

LL) 3.0w D 0) - -~f0 ~ - 0 f0~t C~r4

CN . - Dq 4 0) -4C% P- N C04 D tD 4 m O)-LI)o cr ,m t 4 6-.00.4 'O.-4 ; .~a 06 4 4 DN6

CZ + r+++ -

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4- +4

LA.) +~tf

c)1. CDC Dt Mm C " t : l= Dc za

I-.

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4- -0 U 4-4- (1) 1 1 11< c 0

C1) r- C ') tV) Q. .

I -~ -17

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Ln 3t.0

LC; v:U L; 4 o r-r-OT-- a;oN6Ola4- t P, 0N0H (n Lnr-Ln F% nwl wt~lH qO~t %

I U) mYCCl en M CV) m CV) en m M M e CV) CY mY)

r-.4

4J 4J4.U = - toW to to to .tto %D w to to W( to WtD Wt~~~c N m 0 .0 . * . . .. . . .

'4- -J r~-r-r- iR-.tr qtrr-r. rr-rW- LU H, C,

Ln C" m0H1-I--

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

w. im to 4-J czO1 011lt ~ t0 r,40'1- cm '4. =. - HC.HH N 'H tlC J

-. ~~4- 4- Udi11 11 g I g l

AC C .C

a) c Z cOl- to--- (

(L)~~I 0)i ) 1 L L-

0 0C

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N . rL or *L DtD c)- nL V nlW4.1 nme mC)e V n f r v

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L) CV)

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W ý .0 C

I.--

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(A 40 .3c -gm0) H R , mm44- 01C I0C I1'. ~ .

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a) UW -

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cm C;C

o Ln (m COC n( O C n DL lC n4

U)

IL LL r 4- CO-

4)0 cm0U c-

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Cn 0 mU)i t

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 z1 Z.234C0D+03 2.000000.02 0.0 S.85000.01 l1.000C00e02 1.000000+02 1.000000*02 1

L.. L8170 0003 Z.79000,0o __..0.0 -. . ';.85000DO02 1 _.00:3+02. 1.OO0OoDOZ 1.0000+0,02 23 2.753003:113 3.000000.02 0.0 1.800000D02 1.000000o02 1.00000oo 02 1.000C00002 34 1.930000.03 3.200000402 0.0 5.850000.01 1.0000.0002 1.000000+02 1.000000D02 45 9.000000+02 4.500000+02 0.0 8.8500c0o01 1.000000+02 1.OCOo0,o02 1.000000*02 S

-.ft P-500000+01-.2.5575•0.2 -3.425000+01 - 5.fl3coo30l -1.0coo00002 1.000003tO2 -- 1.00ý0002 -- 67 2.500000+111 2.S575CDO02 -3.405000+01 5.413000+01 1.0000CO02 1.00000.D02 1.03300D#02 78 5.000000+01 2.495000+02 3.425000.01 3.7050C0+01 1.0000co+02 1.000000+02 1.000000+02 aI 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 1112 7.300000.01 2.0000CD002 0.0 9.340000.01 1.130000.01 1.0CoCo.01 6.530000+00 1213 7.300000.01, ?.000000+02 0.0 9.340000+01 1.1300C3.01 1.0&0000+01 6.5:0000*00 13l. .. 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-_POIN•T 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+02Z 2 2.261700+C2 3.4,50-10+01 5.78700D.012 3 2.740000+02 -3.4^5000*01 1.057500.012 4 2.261700.02 -3.425000+01 5.737000+01S8 2.502500+02 3.425000+01 3.11C0000019 1 2.502500.02 -3.425000+01 3.ilOco.o011 1 2.000000+02 2.400009+01 6.050000+011 2 2.000000+02 -2.400000+01 6.050000+01 _

2 5 2.790000.02 2.40300C0+01 6.050300+012 6 2.790000.02 -.. 4O03C00C1 6.0500CD+014 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+012 7 2.790000+02 4.6C00VD+01 6.050000+012 8 2.790000402 -4.6C000+01 6.0500C0+012 9 2.2b1700+02 2.7C000001 5.73/00+012 10 2.261700D02 -2.700000+01 5.787000401

120

Q

.4

TABLE A-9. BEAM DATA

6~DAMPIPNG TBAN 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-. t1 t 0 0 2.000D#01 4.5000402 1.0000+03 2.0000.02 1.000C0.¢ 1.002CO00 0.0 7.9000.01 5.00ý3-02 s 0 a 0 0 1 1 22 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 33 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.ý33.0Ofc S.00.0-02 5 0 0 0- 0 ..'. 4t 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 12 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 36 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 8 0 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 0z 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 2 910 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 1 C

.- 11-0-.0 2.00•D-03 5.0 v0-0^ 5.01^0-02 1.002D-01 1..C0ctC0 1. C 30210. 0.0 7.430.0 5 C220-CZ 4 3 0 c 0 a1210 ! 111 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 32 011 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 OFELASTICITY RIGIDITY

MC MATERIAL (PSI) (PSI)

2 6150H Steel 30.0E6 I1.0E6

4 2024-T3 10.5E6 4.0E6Alu minum

5 6061-T3 10.0E6 3.8E6Aluminum

8 Zero Torsion 1.0E6 0.0Material

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

12

-'C.

-'C. . •

'Cm.

i-,.~%, 2

-- _C.'

TABLE A-12. SPRING DATA

FREE FRICTION BOTTOMINGSPRING LENGTH COEFFICIENT MFIN1G

M•K lLBtAR•ZKM)-t'J(IKII) .... KE(IZI)-

14 3 0 9.750000#00 3.400000-01 4.000000+0415 3 0 9.750000+00 3.400000-01 4.000000.045 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 7.500000.042 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+034 3 3 7.500000+00 3.400000-01 7.000003+03S4 3 3 7.5O0000#00 3.400CO0-01 7.000000+03

- .J..._3-3 .-.7.500000*00 .3.400000-01- 7.000000+031 3 4 7.500000400 3.400000-01 7.000000.03"2 3 7 7.500000.00 3.4000C0-01 7.000000.032 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 4.5000004005.030000+00 1.00000+01 2.000000.01 2.400000.011.000000-01 2.750000300 2.760000.00 3.750000.00

-1.000000-00 1.7500CO0oo 1Z.7600000.00 i3.750000oo01.000000+00 1.0000C0g0o 1.010000.00 1.020000.00S1.010000+00 J.()(00cO#*O0 1.010000#00 1.0Z0000+*001.000000-01 2.7500C0000 2.760000+00 3.750000+00

--. :0000OD-01,2.7500CD00 - Z.760000+00 _ 3.750000+00S1.O000C-01 2.7500C0+00 2.760C00000 3.750000+001.0000GD-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- -2.937110,001.000000.04 2.00000C04 1.0000C0-01 2.937110.C01.950000.04 1.9500C0004 1.000000-01 1.9n6qo+0,03.64CO00004 3.64&000.04 2.000000-02 5.809120.013.64ý000.04 ... 3.648C0D004 .....2.OOCOCO-02 -. 5.809120*.17.500000.04 7.5C0000+04 2.0000C0-02 Z.3754704017.500000.04 7.500000.04 2.00c000-02 2.373470+014.032000+04 4.03M0004 Z.OCCOCO-OZ 5.507810+014.03:0C0.04 .-. 4.0.ZO0.04- 2.000003-02 5.507310+311.824000.04 1.824000#04 2.000000-02 3.817970+011.824000.04 1.824000+04 2.000000-02 3.817970+C11.368000.04 1.368000.04 2.000040-02 3.557350D011.368000#04 -.. 1.363000+04 -... 2.0=0000-02-3.55733•,01G-1.438000+04 1.488000.04 2.000000-02 3.345960.011.488000.04 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

_J I J " N MLEO FAO FAA . EXPOLE YMAX -

k 2 6-1 0 1.3270+01'4-.980 •h03 1.2,200.02 1.3CC.00 1.1000.016 2 7 3 0 1.3270401 '.96C0003 1.2200+02 1.30'D000 1.l000+017 6 8 0 0 1.5030+01 7.3740.02 1.:.004ZC 1.3003+00 1.4,500*018 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 5.50Co0006 2 7 3 0 2.0000•00 3.0000.00 2.0000#05 2.000.03 5.5000.007 6 8 0 0 5.0000-01 3.0000+00 2.0000&05 2.000005 5.S00j000

_0-7.-9 - .90 J.o.500-.o.1 A..oCo.go0.z.oooo-os.ZOOOo.s 5.5000o0o.0

1244770-85

ad-a162 097 - dtic · 2011-05-14 · usaavscom tr-85-d-4 ad-a162 097 advanced technology helicopter...

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