iviel - dtic · landings, the particular case of a noseup landing with forward speeds, the...

S UARL-R-8-4 IVIEL0

'~EVALUATION OF AN ENERGY DISTRIBUTION SYSTEMFOR HELICOPTER LANDING GEARS DURING HARD LANDINGS

A. H. Logan, C. A. Waldon, E. FourtHughes Helicopters

SDivision of Summa CorporationO.. Culver City, CA 90230D D C

ICDI

'~November 1978

Final Report for Period March 1977 - September 1978

Approved for public release;distribution unlimited.

Prepared forAPPLIED TECHNOLOGY LABORATORYU. S. ARMY RESEARCH AND TECHNOLOGY LABORATORIES (AVRADCOM)Fort Eustis, Va. 23604

79 0301 055

BestAvai~lable

Copy

APPLIED TECHNOLOGY LABORATORY POSITION STATEM..T

The findings of a laboratory tested landing gear energy distributionsystem having crashworthiness capabilities are presented in this report.A hydraulic system of conventional oleo dampers, accumulators, equalizerspetc., with interconnection of each landing gear strut, is used to minimizeboth pitch and roll moments that occur during hard landings. The re-sulting attenuation and redistribution of the landing impact energyenhances the pilot's control of the aircraft and reduces the possibilityof aircraft damage and personnel injury. Both ground shake tests anddrop tests were performed to evaluate the viability of the concept. Acost savings of greater than 2:1 is indicated from analysis by incor-porating the interconnected landing gear system in new production air-craft.

Mr. Villiam T. Alexander of the Aeronautical Technology Division servedas project engineer for this effort.

DISCLAIMERS

Th findings In this report awe not to be construed as an official Department of the Army position unless sodesiated by other authorized documents.

When Government drawingp, specifications, or other dts are used for amy purpose other then In connectionWith a definitely rlt Government procurement operation, the United States Government thereby inae noresponsbllity nor any obligation whatsoever; end the fact that the Govermnt may hews foemulted, furnished.or In any mwy supplied the sid draw1ings, specifications, or other dam Is not to be rgwld by ImplinetOon orotherwso a in any manner licosing the holder or any other person or onprtion, or onvying ay r orpermission, to manufacture, use, or sell any patented Inmention that may in any way be reled hreto.

Trade names cited in this report do not constitute on official endorsement or approvel of the use of suchcornmercial hardwre or softswe

DISIPOS1TION INSTRUCTIONS

Destroy this report when no longer needd. Do not return it to the origimator.

L.|

UnclassnifiedLSECU RIT AIASSI FICA TIOM OF THIS PAGE flb.n DaAee ______EADINSRUCTION

J -USARTL-TR-78-44

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UnclassifiedSECURITY CLASSIFICATION OF THIS PAGSShb D04 MAuamr4

- ,to move from the flight position toward the fully extended position. Whenthese motions have been accomplished, the skids remain on the groundsurface throughout the landing, greatly reducing the pitching moments.

The landing gear was drop tested to demonstrate the effects of sink rate,

gross weight, center-of-gravity (CG) location, touchdown attitude (bothpitch and yaw), ground resonance, system damping, and spring rate. Thetesting included drop velocities up to 19. 5 feet per second and simulated

forward and lateral speed landings. For purpose of design and development,the OH-6A helicopter was used as the baseline aircraft, and the landing gearwas designed to require minimum modification to the OH-6A. Although thegear was designed for the OH-6A, the basic design princi developedalso apply to wheel-type landing gear.

The results of the testing showed that the interconnected landing gearreduces the nosedown pitching velocities and angles during autororationlandings, the particular case of a noseup landing with forward speeds,the interconn cted landing gear reduced the pitching velocities 60 percentas compared t the basic OH-6A. As compared to MIL-STD-1290, theinterconnected anding gear increased the OH-6A fuselage ground contactvelocities to 19 5 feet per second. In addition, analysis showed that if the

fuselage suppo t structure were strengthened, an OH-6A equipped with theinterconnected landing gear can absorb approximately a 33. 7-foot-per-second impact without serious crew injury. The experimental landing gearwas also inter ected in roll and demonstrated a 40-percent reduction inroll velocitie . A cost analysis indicated that incorporation of the inter-connected landing gear in new production aircraft would result in a returnon investment greater than 2:1.

UnclassifiedSECURITY CL AI PlC ATION O

F THIS PA6.l.mb Data mIMrOM.1

__p

PREFACE

This report was prepared by Hughes Helicopters, Division of SummaCorporation, under Contract DAAJ02-77-C-0019, funded by the AppliedTechnology Laboratory, U. S. Army Research and Technology Laboratories(AVRADCOM), Fort Eustis, Virginia. The ATL technical monitor for thiscontract was Mr. William T. Alexander. The Hughes Helicopters projectmanager was Mr. Andrew H. Logan. Mr. C. A. Waldon conducted thedrop test of the landing gear, and Mr. E. Fourt developed the cost/benefitsanalysis. The authors would like to acknowledge Mr. R. A. Wagner forhis support and many helpful suggestions during the program.

~AC.4,

NTiS

. , - -. ,

3

I- . 77

TABLE OF CONTENTS

Page

PREFACE ........ o.......... o...... . . .. . . 3

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

LIST OF TABLES . ............. ... * ... .... .. 11

INTRODUCTION ..... . .............. . ....... .... 13

INTERCONNECTED LANDING GEAR DESIGN ............. 17

Pitch Interconnection ......................... 17Roll Interconnection ........... ..... ......... 18System Installation ..... ... ... .. . ... ..... 18

System Weight ......... ..... . . ..... . 22

TEST EQUIPMENT AND INSTRUMENTATION ............ 24

Test Equipment . o .... . ..... .. .. ..... . 24

Instrumentation ...... . .... .... .... .. ..... .... 24

&&!,STS ... ....... ...... o........o......... ... 31

Shake Test ..... ... ........ . . .. . ... ...... . 31Drop Test ............... ....... 4 . ... .. 31

RESULTS AND DISCUSSION ........................ 32

Shake Test . . 0 . . ....... . ... .. . .... . 0. 32Drop Test .0....... ......... . .00*0.......... ..0 33

Life Cycle Cost ....... . 0 0 .. . . ... . .. . .. .... 43

CONCLUSIONS ................. ... . ... ... .. ... 49

RECOMMENDATIONS ..... ........... s o o . o50

REFERENCES ... .......... ................. . 51

5_________________________________

TABLE OF CONTENTS (CONT)

Page

APPENDIX A - INTERCONNECTED OH-6A LANDING GEAR DROPTESTS OF AN ENERGY DISTRIBUTION SYSTEMFOR HELICOPTER LANDING GEARS DURINGHARD LANDINGS .................... 52

APPENDIX B - COST ANALYSIS CALCULATIONS ......... 84

6

LIST OF ILLUSTRATIONS

Figure Page

1 The interconnected landing gear increases blade/tail boom separation and eliminates contact .......... 15

2 Diagram of pitch interconnected landing gear for theOH-6A ........ ................................ 17

3 Modified damper/sleeve assembly ................... 19

4 Pressure equalizer ........ ......................... 20

5 Hydraulic schematic of interconnected landing gearinstallation ....... ............................. 21

6 Typical test setup to simulate a forward speed landingwith pitchup attitude ............................. 25

7 Test setup to simulate a level autorotational landingwith lateral velocity ............................. 26

8 Strain gage locations .............................. 27

9 Schematic of modified damper/sleeve assemblyinstrumentation ....... .......................... 29

10 Velocity pickup operation .......................... 30

11 Effect of interconnected landing gear on pitch angleand velocity for a 6. 55-foot-per-second verticaldrop with simulated 30-knot forward speed ........ ... 34

12 Comparison of theoretical and experimental pitchvelocities for the basic OH-6A . ................. 35

13 Comparison of pitch angles and velocities for theinterconnected and basic OH-6A landing gear dur-ing a level autorotational landing ................... 36

7

LIST OF ILLUSTRATIONS (CONT)

Figure Page

14 Comparison of experimental pitch velocities forextended length standard gear and interconnectedgear during a level autorotational landing withforward CG .......... ......................... . 37

15 Effect of landing gear interconnection on roll angleand velocity for a 6. 55-foot-per-second verticaldrop . . .............. ............................... 39

16 Interconnected landing gear after a 19.2-foot-per-second vertical impact . . . ......................... 41

17 Limits of human tolerance to vertical deceleration(derived from Reference 6) . . . .................. 44

18 Cost savings of interconnected landing gear -

retrofit aircraft ..... . ................ . . . .. . 46

A-1 View of modified oleo damper interconnectinstallation . . o.... 0.... . . .. .... ... . 64

A-Z View of modified oleo damper interconnectinstallation ...... ... .... ..... ..... . 64

A-3 Landing gear drop test fixture used for theinterconnected landing gear tests . .... ........ 65

A-4 Plot of load versus frequency at input 0.25 inchP-P. 90 percent lift . . . .. .. .. . . . . . . . 0 66

A-5 Plot of load versus frequency at input I inchP-P, 90 percent lift . . .. .. . . ... .. .... 66

A-6 Plot of load versus frequency at input 2 inchesP-P, 90 percent lift . . . .. 0 * .. . . . .. 0. . . . .. . 67

A-7 Lissajous of 1-inch P-P at 1. 38 Hz, *77 pounds .... 67

A-8 Lissajous of 2-inch P-P at 2. 15 Hz, *175 pounds . . o 68

8

.- i


Figure Page

A-9 Plot of load versus frequency at input 0. 25 inchP-P, 0 percent lift ...................... 68

A-10 Plot of load versus frequency at input 2 inchesP-P, 0 percent lift ...................... 69

A-11 View of deflection transducers used to measureinterconnect and damper motions . ............ 69

A-12 View of aft right strut failure, Test Condition 12 . . • 70

A-13 View looking forward at failed landing gear,Test Condition 12 ........ ....................... 70

A-14 View of failed lug on front right strut .......... . . . 71

A-15 Record 9, Baseline Design Condition,Test Condition 1 . . . . . ..... ........... 72

A-16 Record 12, Damping Variation

Test Condition 2 ........... ............ 73

A-17 Record 15, Spring Variation, Test Condition 3 . . . .. 74

A-18 Record 16, Simulated Forward Speed, Condition 4 . 75

A-19 Record 17, Forward CG, Test Condition 5 ....... 76

A-20 Record 18, Simulated Lateral Speed, Test Condition 6 77

A-21 Record 20, Level Drop, Test Condition 7 .. ........ 78

; 9


Figure Pg

A-.22 Record 22, Roll, Test Condition 8 ..... ... 79

A-23 Record 26, Overload, Test Condition 9.............. 80

A-24 Record 21, Reserve Energy, Test Condition 10 . . . 81

A-25 Record 25, Nose Down, Test Condition 11. .. .. ...... 82

A-26 Record 27, Maximum Drop, Test Condition 12 .83

.4 10

LIST OF TABLES

Table Pa-e

1 System Weight ... ........................ 22

2 Cost Savings of Interconnected Landing

Gear for Retrofit (1972 Dollars) .................... 47

3 Cost Savings of Interconnected Landing

Gear for Forward Production(1972 Dollars) ......... .......................... 47

A-i Ground Resonance Test ................... 60

A-2 Landing Gear Drop Test Conditions ..... ............ 61

A-3 Drop Test Ballast Summary for the Model 369AInterconnected Landing Gear Configuration 62

A-4 Summary of Drop Test Data ....... ................. 63

-11

- *.- ~

INTRODUCTION

Blade/tail boom strikes occur with an excessive frequency during emergencyautorotations. Many of these strikes have resulted in substantial damageto the helicopter and in fatalities and injuries to personnel. In addition,

current Army-size limitations require more compact helicopter designswhich bring the tail boom and main rotor closer together, increasing thepossibility of blade-boom contact.

The sequence of events which results in blade/tail boom strikes in emer-gency (or practice) autorotations predominately follow this pattern:

a. Ground contact is made with the helicopter in a noseup attitude.

b. The vertical reaction loads act to give a nosedown moment on the

helicopter. This nosedown moment is increased due to drag loadsif forward speed is present at contact.

c. This nosedown moment causes angular acceleration and nosedownangular velocity (nosedown angular velocity is also tail boom-upangular velocity).

d. Pilot reaction to nosedown velocity is to pull the cyclic stick back.This brings the rotor blades down in the rear while the tail boomis coming up. This combination aggravates main rotor blade andtail boom interference.

It should be evident that whatever reduces the nosedown pitching momentwill reduce the tendency toward boom chops. This fact is widely recognized,and pilots are trained to level the helicopter prior to ground contact for thesole reason of reducing the nosedown moment.

Unfortunately, this maneuver requires considerable judgment and finessein handling both the cyclic and collective controls. Additionally, the act ofleveling the helicopter prior to touchdown reduces the angle of attack of the

rotor; and, hence, reduces the lift on the rotor at the wrong time in the

maneuver.

Recognizing these autorotation problems, a preliminary design study was

conducted to define a landing gear concept which reduces the nosedown

pitching moments by providing an interconnection between the front and rear

...

.13

1landing gears. Through the interconnection, as the rear landing gearmoves from the flight position toward the full compressed position underlanding impact, the front gear is impelled to move from the flight positionto a more extended position. When these motions have been accomplished,the skids (or front and rear wheels) are on the ground surface, and thevertical reactions inherent in absorbing the autorotational landing do notproduce a pitching moment.

The analysis showed that interconnection of the front and rear supports ofa skid-type landing gear significantly reduces the maximum nosedown pitch-ing angles and velocities that occur during autorotational landings. Thisresults in a more controllable autorotational landing and increased blade/tail boom separation (Figure 1). The increase in separation is larger in apure vertical landing than in a landing with forward speed. Although theincrease in blade/tail boom separation is smaller, the contribution of theinterconnected landing gear during forward speed autorotation landings issignificant because it eliminates blade/tail boom contact in a landing wherecontact has been recorded.

The lateral interconnection of the landing gear produces the same increasein helicopter controllability during autorotation with roll as does the foreand aft interconnection in pitch.

To verify the predicted benefits, an experimental study was conducted.Utilizing the preliminary design findings, I a full-scale skid-type landinggear was designed and fabricated to be capable of alleviating the landingloads and moments associated with both normal and hard landings. Thelanding gear incorporated both pitch and roll hydraulic interconnect sys-tems which distribute and attenuate the landing impact energy betweenlanding gears to minimize both rolling and pitching moments. The landinggear was designed by Hughes Helicopters and fabricated by the WesternDevelopment Center of MOOG, Inc.

The landing gear was drop tested to demonstrate the effects of sink rate,gross weight, CG location, touchdown attitude (both pitch and roll), groundresonance, system damping and spring rate. The testing included dropvelocities of 6. 5, 8. 2, and 19. 5 feet per second; design (2550 pounds) andoverload (2880 pounds) gross weights; simulated forward and lateral speed Jlandings; maximum fore and aft CG locations; and noseup, level, and

1. LOGAN, A. H., "Analytical Investigation of an Improved HelicopterLanding Gear Concept, " USAAMRDL-TR-76-19, August 19764U. S. Army Air Mobility Research and Development Laboratory,Ft. Eustis, Va., AD A029372

14

6 A~z - 6.56 FEET PIER SECOND

-10 DEGREES (NOSE UP)GROSS WEIGHT - 2550 POUNDS

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nosedown, and roll landing attitudes. For purpose of design anddevelopment, the OH-6A helicopter was used as the baseline aircraft, andthe landing gear was designed to require minimum modification to theOH-6A. An existing OH-6A landing gear drop test fixture was used forthe tests. Although the gear was designed for the OH-6A, the basic designprinciples developed also apply to wheel-type landing gear. The differenceis that, in wheel-type gear, the interconnected front and rear supports areattached to independent wheels and not a skid tube common to all supports.

The results of the testing were compared to the basic OH-6A helicoptertest data.and landing gear performance improvements were determined. Acost/benefits study was then conducted to determine the impact of the inter-connected landing gear on the OH-6A cost, reliability, and maintainabilitycharacteristics.

16

INTE~R ONNZC LNIGGETRDSG

bth test configuration landi' gear is a skid-.type landing gear incorporatin

bot Ptch an rol ydr ulC interconnect system s, The 01 -6A is the

baseline aicatand

the tes landin' ea

rndiaircrto tht Cng4 Tear Is designed to r equire nnijr

rrocjic to o h Q . A h CO mnPlete design dev l p ment 'of th e nt r.

Co ne t ys em spresented n e f erenc and a description of the e ite rtconfiguratio

is prsen ~ f1PITCH~ INTE;RCONNECIONwn scins eThe Pitch interconnected

landing gear is essentially the 0 1 -- 6A landing gear

with mnod fication as shown sche mnatically in Fig re 2 Th bai 1 -6

a u s e l a g e P i v o t P o i n t s , d r a g b r a c e s , a n k ir e i u n h a g T h e r o n tan eroe apers (369H~92

131) now fit intoneslvswhcarattached to the cross tubes. a h n w se v ro i e n a dto a

1. 74 inches Of travel both c a ndwn froee te nr vetra adition a h l e

damnper Upe attahmen

posfr teneta

tio toattchmnt oits are relocated frorn the basic Oti~ 10ac~rnnodtethis additional interconnect travel. Poion. eah oleve

MODIFIED OLEOOAMPER/SLEEVE

PRESSUREINTERCONNECTIO

CRtOSS TUBES EXTENDED

J L E F T -H A N O S DE S H o pMIGHT-HANO SIDE OWNE

Figure 2. Diagrann Of Pitch interconnected ln AEA ETHN le

ladng gear for the OH..6A

17

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two annular chambers filled with hydraulic fluid are formed by sealing thesleeve to the basic oleo damper (Figure 3). Matching hydraulic chamberson the front and rear damper/sleeve assemblies are connected through alow constraint interconnection system. The low constraint interconnectionsystem is comprised of a pressure equalizer (Figure 4), and two surgereservoirs. The pressure equalizer is a spring/piston assembly whichconnects the lower hydraulic chambers on the front and rear damper/sleeveassembly. In a nose-high landing when the aft skid experiences high forceand the forward skid little force, the aft skid hydraulic chamber iscompressed, forcing hydraulic fluid out of the lower aft chamber, into thepressure equalizer. This creates unequal pressure in the pressureequalizer, moving the piston forward and forcing hydraulic fluid into thelower forward chamber. This causes the forward skid damper to extenddown while the aft skid damper compresses upward.

The interconnect system spring rate and damping characteristics werechanged by modifications to the pressure equalizer. With reference toFigure 4, the spring rate was changed by installing different springs in thepressure equalizer. Two spring rates were tested: 17 pounds per inchand 11 pounds per inch, giving system spring rates of 34 and 22 pounds perinch, respectively. The damping was changed by installing a differentdiameter orifice in the piston face. Two orifice diameters were tested,0. 128 and 0. 059 inch.

ROLL INTERCONNECTION

The roll interconnection of the landing gear is accomplished in a mannersimilar to the pitch interconnection shown in Figure 2. For roll inter-connection, two additional pressure equalizers are needed: one intercon-necting the front modified damper/sleeve assemblies and one between therear damper/sleeve assemblies.

SYSTEM INSTALLATION

The complete pitch and roll interconnection system installation is shownin Figure 5. There are four pressure equalizers, four surge accumulators,and one reservoir. The surge accumulators are used to limit line pressure:and are placed on each side of the pressure equalizers and are connectedby a dual flow valve simulation using check and relief valves in parallel.The relief valve allows flow into the accumulators when the line pressureexceeds 1200 psi. The check valve allows flow out of the accumulatorwhen the line pressure is below 100 psi. The relief valve pressure is setso that when the landing gear is on the ground, the damper/sleeve assembl

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would be in a neutral position, providing approximately 1. 75 inchesadditional damper travel up and down. The reservoir is connected tothe upper hydraulic chambers of the damper/sleeve assemblies, ispressurized to 100 psi, and is installed to prevent cavitation in theupper chambers.

Each pressure equalizer was connected to the system by valves so thatthe removal of the pressure equalizers did not require replumbing theentire system. Hydraulic fluid was fed into the system by two valves,one for the upper hydraulic chambers and one for the lower hydraulicchambers.

SYSTEM WEIGHT

The experimental prototype interconnect system was designed to be simpleand low in cost while retaining the dynamic characteristics of a more com-plex flightworthy installation. Consequently, no effort was made to designsmall, lightweight components. The weights of the prototype interconnectsystem's major elements are presented in Table 1. Comparisons to theweights of a basic OH-6A and a production version of the interconnectsystem are also presented. Only the major elements of the system

TABLE 1. SYSTEM WEIGHT

InterconnectLanding Gear

Basic OH-6A weights, lb

Item Weights, lb Production* Prototype

Dampers (4) 6.4 17.2 56

Surge Accumulators (4) 0 4.0 20.9

Reservoir (1) 0 1. 5 7.0

Pressure Equalizers (4) 0 10.8 15.0

+27. 1 +92.5

*From Reference I

22

are shown due to the difficulty of obtaining an accurate weight of theprototype plumbing. As shown in Table 1, the prototype interconnectsystem developed for this test weighed 92. 5 pounds more than a basicOH-6A landing gear. In production, however, where more efficient andcompatible components would be designed, the system elements wouldadd no more than 27. 1 pounds to the aircraft basic weight. As an exampleof where component weight could be reduced by more efficient design, thefour surge accumulators and one reservoir used in the prototype systemwere off-the-shelf items. These items were made of cast iron, had morecapacity than needed, and were developed for industrial applications. Thedampers and pressure equalizers were designed similarly. There was nomachining of excess material to reduce weight and a complex analysiswas not conducted to determine minimum wall thicknesses. A productionversion would incorporate both additional machining and detailed analysisresulting in reduced weight.

23

TEST EQUIPMENT AND INSTRUMENTATION

A complete description of the test equipment and instrumentation ispresented in the test report (Appendix A). This section contains asummary of the test equipment and instrumentation.

TEST EQUIPMENT

The experimental landing gear consisted of the interconnected landinggear system, described previously, mounted on an OH-6A extended lengthlanding gear. The experimental landing gear was mounted on the testfixture as shown schematically in Figure A-3.

With the landing gear installed, the test fixture simulates the helicopterweights, moments of inertia, and CG locations by relocating ballastattached to the fixture at designated weight par- positions. Simulated rotorlift is applied through the CG of the drop test fixture by the use of testlinkage connected to a special air cylinder -tank absorbing systemmounted on the test gantry as shown in Figure A-3. The downward dropvelocity is controlled by changing the free drop height. The required dropheight is obtained by a second hoisting mechanism (other than the simu-lated rotor lift hoist) located between the test fixture and the overheadelectric hoist on the gantry as shown in Figure A-3. This mechanism,which supports the test specimen until drop time, provides for remote airactuation of the safety pin and release hook.

The landing attitude, forward speed, and latera] speed are simulated by theorientation and comp3sition of the impact surfaces. A typical test setupfor a forward speed landing with a pitchup attitude is shown in Figure 6.Forward speed is simulated through the use of an inclined landing platformwith a rusty steel surface (coefficient of friction equals 0. 5), as shown inFigure 6. Reaction components are normal and parallel to the surface aswith drag induced by forward velcity. Landing attitude is measured rela-tive to the landing platform with a pitchup attitude relative to the inclinedsurface being shown in Figure 6. To simulate a landing with lateral speed,a plywood platform and side ramp (Figure 7) were used to produce a verticalright skid reaction and a left skid outboard reaction equal to approximately50 percent of the vertical reaction.

INSTRUMENTATION

The instrumentation was installed to substantiate the landing gear designand functional behavior. The data collected included the axial and verticalforces for forward and aft struts and drag braces. All four oleos were

24

Figure 6. Typical test setup to simulate a forward speed landingwith pitchup attitude.

instrumented for position and loads. All four modified oleo sleeves werealso instrumented for position to determine interconnect movement. Pitchand roll attitudes and rates as well as lateral and vertical accelerationswere measured. Also, lift, contact velocity, and interconnect system pres-sure were measured. The exact parameters to be measured are identi-fied by an "X" in Table 2 of the Engineering Test Request found in Appen-dix A, and the frequency response is listed next to each parameter.

The strut and drag brace forces were recorded by 1200 strain gagebridges that were applied at the locations shown in Figure 8. The left-hand landing gear was fully instrumented. In addition, the right forwarddrag brace was instrumented for axial and vertical loads to measuredifferences with respect to the left side caused by the 6-degree offset ofthe left upper forward oleo attach position relative to the right upper for-ward oleo attach position. The right and left rear oleo attach positionsare identical.

* All four oleo loads were recorded by load cells installed between the oleos

and the upper attachment fitting on the test fixture.

25b __ I

ROTOR LIFT SIMULATIONENERGY ABSORBING DEVICE

TEST FIXTUREDRAWING NO. 389-9302FIGURE A-3

VIEW LOOKING FORWARD

H 2660 HDROP HEIGHT

-777M rLEVELSURFACE

Figure 7. Test setup to simulate a level autorotational landing withlateral velocity.

26

P_____________________________________ _________77. 11

TYPICAL STRAIN GAGELOCATIONS ON STRUTS

+ Vb

+At +At

+V b

+At

+H bAXIAL BRIDGE

TYPICAL STRAIN GAGELOCATIONS ON BRACES 8.0 IN.

+At

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369H6006 +At +H b BENDING BRIDGESBRACE T Vb *'J (TYPICAL)(TYP(CALY

+At CD I

AXIAL BRIDGE IN.102I.36916001-1

VERTICAL STRUT ASSEMBLYBENDINGBRIDGE(TYPICAL)

369H921 11 FOOT ASSEMBLY 36H92114 SKIDAStiEMBLY

CALIBRATION LOADS ON LANDINGGEAR AT ATTACHMENT FITTINGS+A t AXIAL TENSION

+Vb VERTICAL LOAD (BENDING)

+Hb HORIZONTAL LOAD (BENDING)

: FORWARD

Figure 8. Strain gage locations.

27

The interconnect system is designed to move the landing gear relative to

the fuselage. This relative motion allows the landing gear to be on theground, reacting the landing force, without requiring corresponding motionof the fuselage. The motion of the landing gear relative to the fuselage

was measured directly at all four damper/sleeve assemblies. The relativelanding gear motion is comprised of two movements: the basic oleo pistonstroke and the interconnect system movement. These two motions weremeasured by linear position transducers installed as shown schematicallyin Figure 9 and on the test fixture in Figures A-I and A-2. The basicoleo piston stroke, aI, was measured between the oleo upper attachmentpoint and the inner barrel of the damper/sleeve assembly. The intercon-

nect system movement, A., was measured between the inner and outerbarrels of the oleo/damper assembly. By comparing the& 1 and A2 motionsof the four damper/sleeve assemblies, the motion of the landing gear rela-tive to the fuselage was determined.

In addition, the complete interconnect system performance can be deter-mined by combining the A 2 motion measurement with the line pressuremeasurements. The pressure measurements monitor the action of the

accumulators. The surge accumulators were controlled by a pressuresensitive valve which opened when the line pressure exceeded the valvepressure setting, thus stabilizing the line pressure near the valve pressuresetting. If the surge accumulators are closed, a compressive A 2 motionin the rear oleo/damper assembly is transmitted through the pressureequalizer and results in a A2 extension in the forward oleo/damper assem-bly. If the accumulators are open, the compressive A 2 motion resultsin fluid flowing into the accumulators and does not result in A2 extensionin the front oleo/damper.

The interconnect system pressure was measured by two sensors. one oneach side of the pitch interconnect pressure equalizers on the left-handgear as shown in Figure 5. These pressure sensors monitored theoperation of the dual flow valves and surge accumulators.

Both lateral and vertical accelerations were measured by standard linear

accelerometers (55 mV/G approximately) installed at the CG of the testfixture.

Pitch and roll rate gyros were mounted near the test fixture CG to

measure pitch and roll attitudes and rates.

The drop velocity was measured by the device depicted in Figure 10.

Simulated lift was measured by a load transducer situated near the test

fixture CG as shown in Figure A-3.

28

.t-.<2 -

COMPRESSED

LINEAR TRANSDUCERS

A1 . OLEO PISTON STROKE

- INTERCONNECT SYSTEM2.2MOVEMENT

J NOTE: TRANSDUCERS DRAWN OUT OF~PLANE FOR CLARITY

a. SIDE VIEW

'14 b. TOPVIEW

Figure 9. Schematic of modified damper/ sleeveassembly instrumentation.

29

WORE ATTACHED TO DROPFIXTURE NEAR CG

0.010-INCH-DIAMETER tMUSIC WIRE2-3 WRAPS AROUND I (2) ELECTROMAGNETIC

C S C 2PULLE PICKUPS

5 s o WIRES TO 50-CHANNELRECORDER

'I WIRES TO DATAFLASH RECORDER

80-TOOTH GEAR

I14-INCH-DIAMETER SUNGEE 8.00-INCH CIRCUMFERENCECORD STRETCHED 2 x L PULLEY

Figure 10. Velocity pickup operation.

A 50-channel visicorder with associated support equipment was used torecord the output of the instrumentation during the test drops. The tracewas made at 40 inches per second, allowing a resolution of at least0. 01 second. For each drop condition, the landing action was accomplishedwithin one-half second.

The vertical impact velocity was recorded as shown in Figure 10. As thelanding gear specimen fixture fell, the preloaded bungee cord reeled inthe wire which rotated the pulley-gear combination. Since the pulley cir-cumference equaled 8. 00 inches and the gear has 80 teeth, the magneticpickup(s) indicated one blip on the oscillograph record as 0. 1 inch of dropheight. The frequency of blips per unit of time (i. e., 1/200 second from200 cycle AC signal) indicated a velocity at any instant of the drop. Twoelectromagnetic pickups were used for permanent and quick-look recordsof the contact velocities measured. The exact instant of contact wasindicated by the accelerometers or the damper response.

Photographic coverage was made of all drop test conditions. The photo-graphic coverage consisted of two 16mm motion picture cameras set upto record landing gear motion in two orthogonal planes. One camera waspositioned in front of the test setup to record motion in the roll plane whilethe other camera was positioned to the right side of the test setup to recordthe motion in the pitch plane. Photographic coverage included a minimumof 100 feet of 400 frames per second coverage of each drop test condition.

30

7777

TESTS

The testing was conducted in two phases: shake test and drop test. Theshake test phase was conducted to explore the ground resonance charac-teristics of the interconnected landing gear. The drop test phase was con-ducted to explore the operation of the interconnected landing gear and itseffect on the aircraft landing characteristics. A complete description ofthe tests is presented in Appendix A. This section is a summary of thetests conducted.

SHAKE TEST

The shake test was conducted at one gross weight, 2550 pounds, for nolift and 90-percent lift conditions. The interconnected landing gear wasmounted on the drop test fixture which was, in turn, in ground contactthrough greased Teflon pads sitting on four steel plates. This simulateda friction-free contact. The shake test rig was oscillated through a rangeof frequencies (1 to 10 hertz) for a range of amplitudes (1 to 2 inches).The frequency range included the predictedl critical frequency of 1. 3 hertz.

DROP TEST

The drop test was conducted over a range of conditions selected to repre-sent the full range of operating conditions. The test conditions includedthe following:

a. Drop velocities of 6. 5, 8. 2 and 19. 5 feet per second. Thesevelocities represent the current OH-6A helicopter limit energydrop velocities, the reserve energy drop velocity, and theanalytically determined maximum allowable drop velocity foxthe new landing gear, respectively.

b. Design and overload gross weights of 2550 and 2800 pounds,respectively.

c. Simulated forward and lateral speeds to OH-6A limit conditions.

f. d. Maximum fore and aft CG locations.

e. Level ground contact.

f. ±10 degrees-pitch slope and ±10 degrees slope.

31

RESULTS AND DISCUSSION

The complete results of the shake test and drop test are presented inAppendix A, Structural Test Report. The results are presented in bothtabular form and as time histories of all data recorded during the droptest. In this section, salient results of the testing are presented anddiscussed.

SHAKE TEST

No ground resonance point could be identified over the amplitude andfrequency range tested. Possible ground resonance frequencies wereidentified in the range from 1. 38 to 2. 15 hertz. However, there was noconsistency as test conditions were changed.

The ground shake testing did reveal the problem of putting the orifice inthe pressure equalizer piston face. Prior to each test, the landing gearhad to be centered manually. No attempt was made to correct the problembecause it was felt that it would not affect the drop test results and becauseof schedule and budget constraints.

Due to the limited scope of the test program, the pressure equalizer wasdesigned to be simple and inexpensive, yet retain the basic features of amore sophisticated system. Consequently, system damping character-istics were achieved by a simple orifice in the piston face. This designworked well in the dynamic mode but in static situations and during veryslow movement, the orifice design compromised the interconnected landinggear concept. Since the fore and aft chambers were connected, duringstatic conditions there was no counteracting centering moment. Conse-quently, if a static moment was applied (such as a man standing offsetfrom the center of CG), one hydraulic chamber would eventually compressand the other extend without restoration to a neutral position when themoment was removed. In a more sophisticated production system, thischaracteristic would be eliminated by a more complex damping arrange-ment. One way of achieving damping, yet having static centering, is tohave a movement sensitive orifice in the piston face. At low piston speeds,the orifice would be closed and at high speeds it would be open. Anothermethod would be to have piston damping achieved by a system separatefrom the hydraulic interconnect system. Both of these concepts requirefurther design effort.

Another ground resonance shake test would be required for this refineddesign of the interconnected landing gear.

"-32

t7

DROP TEST

The primary drop test objective was to demonstrate that a hydraulicallyinterconnected landing gear would reduce pitching and rolling velocitiesduring autorotational landings. The data indicate that the interconnectedlanding gear does reduce pitching and rolling velocities resulting in morecontrollable autorotational landings.

The test results for both pitch and roll landing conditions were comparedto predicted behavior from Reference 1. A comparison between predictedand measured pitch angle and velocities is shown in Figure 11. Due toinstrumentation problems, the variation of the interconnected landing gearexperimental pitch angle is calculated by integrating the measured pitchrate trace. The test data were recorded during a 6.25-foot-per-seconddrop with the test rig angled 10 degrees noseup relative to a 26. 5-degreeinclined surface. This condition simulates a noseup autorotation landingwith forward speed. The test data indicate lower pitch rates than werepredicted for the interconnected landing gear. This occurred due to twofactors: First, based on other comparisons, the computer simulation isconservative in that it predicts slightly higher loads and rates than aremeasured. Second, the interconnect spring rate used in the test landinggear is lower than the spring rate used in the computer simulation whichpredicted the landing gear behavior. The lower spring rate was used inthe experimental gear due to fabrication difficulties and size limitationsassociated with using the simulation spring rate.

In addition to predicted interconnected landing gear behavior, the predictedbehavior for the basic OH-6A landing gear is presented in Figure 11. Acomparison shows that the interconnected landing gear reduces the nose-down pitch rate by approximately 60 percent when compare' to pitch ratespredicted for the basic OH-6A landing gear. The basic OH-6A landinggear has not been tested for these conditions. However, during develop-ment of the computer simulation, which predicted the basic OH-6A landinggear pitch rates, good correlation was demonstrated between predicted andexperimental behavior for other drop conditions. This is shown inFigure 12, which is taken from Reference 1. Consequently, it is felt thatthe experimental data for the OH-6A would be close to the predicted valuesif it was tested under the proper conditions. Consequently, the comparison

j of experimental and predicted pitching behavior results is a validdetermination of the benefits of the interconnected landing gear.

33

WITH SIMULATED 30-KNOTFORWARD SPEEDAVZ - 6.55 FEET PER SECOND

M1--10 DEGREES (NOSE UP)GROSS WEIGHT -2550 POUNDS

-101

TES

0Z,

1d 0

101.

U

40

z I--

0 0.10203O

TIME, SECOND@S

Figure 11. Effect of interconnected landing gear on pitchangle and velocity for a 6. 55-foot-per -secondvertical drop with simulated 30-knot forwardspeed.

34

*1 .7

REFERENCE 1, FIGURE A-7

0 BASIC OH-GA EXPERIMENT

U 40 OT-72 RECORD 3481280-POUND,O | 100-PERCENT LIFTLu (a 26.5 INCLINE PREDICTED ""--,-,--

>+8 12 NOSE UP [ '

FORWARDCG""" 120'

0 0.05 0.10 0.15 0.20 0.25 0.30 0.35

TIME AFTER IMPACT, SECONDS

Figure 12. Comparison of theoretical and experimentalpitch velocities for the basic OH-6A.

The benefits of the interconnected landing gear are also demonstratedduring a level autorotation landing. A comparison of experimental datafor the interconnected and basic OH-6A landing gear 2 is shown inFigure 13. The comparison shows that the standard gear pitches oversharply to approximately 5 degrees while the interconnected landing gearresulted in less than a 2-degree noseup attitude following the drop.

A further comparison to experimental landing gear data indicates thatfor a forward CG location the effect of interconnection is minimized. Acomparison is shown in Figure 14 for the interconnected landing gear, the

3 4extended length landing gear, and an improved OH-6A landing gear

2. MAGULA, A. W., "369A6000B Production Landing Gear Drop Tests2800 lb Gross Weight, using 369A6300 Dampers with 369A340-601Bladders and 369ASK 150 Double Acting Dampers, " Hughes ToolCompany - Aircraft Division Report 369-BT-3609.

3. MAGULA, A. W., "369H90006 Regular Production Extended LandingGear Drop Tests (2550 lb Gross Weight)" Hughes Tool Company -

Aircraft Division Report 369-BT-3033, 1969.

4. MAGULA, A. W., "Drop Tests of Improved Landing Gear for Model369A Helicopters, " Hughes Tool Company - Aircraft Division369-BT-3613, May 1971.

35r

,a , I i ,i . -l II lI . l I

-2-

INTERCONNECTED

0

CA,

0U

w-J_j

2

-N

SNDARD OH4J

z

0

LU

Cz

LUw.

01 DEGRE2S

TIMEH SECONDSE

RE9EC 36REOD 7

25,-0 .901 .403 .004

NOSE UP

-1

LU

GROS WEGHT %. 2550EPOUND

67 PERCENT LIFTSURFACE INCLINE -26.5 DEGREES PITCH0 DEGREES PITCH AND ROLL___ __

CG -STATION 97

NOEDOWN__________________

NOSE UP4

j2 E

06

zI0U _

S -6

W -10I EXTENDED LENGTH

STANDARD GEAR0. -12 -- - -

GROSS WEIGHT is 4 ONSIMPROVED OH4A

-14 -IMPACT VELOCITY - 6.5 FEET PER SECONDGERKIM026

I SURFACE INCLINE - 26.5 DEGREES PITCHI 0 DEGREES PITCH AND ROLL -$ CG -STATION 97

NOSE DOWN I I __

-0.04 0 0.04 0.06 0.12 0.16 0.20 0.24 0.26 0.32 0.36 0.40 0.44 0.46 0.52

SECONDS

Figure 14. Comparison of experimental pitch velocities for extendedlength standard gear and interconnected gear during alevel autorotational landing with forward CG.

I3

The geometry of the extended length landing gear is essentially theinterconnected landing gear without interconnection. The other differenceis that the interconnected gear uses four 369H92131 dampers, while theextended gear uses 369H6340 dampers in the front and 369H92131 dampersin the aft struts. As compared to the basic OH-6A landing gear, theimproved OH-6A landing gear (kit M30245) has a swivel joint at the aftcross tube-to-skid attachment and aft cross tubes with increased yieldingcapability. The comparison shows that for this drop condition, the per-formance is approximately equal for all three landing gears. The inter-connected landing gear, however, did have the lowest pitching rate,approximately 12 percent less than the improved OH-6A landinj gear.

In the roll mode, the drop test data indicate that increases in autorotationallanding controllability are shown for the interconnected landing gear. Asshown in Figure 15, for test condition 8 (6. 5 foot-per-second impactvelocity and 10-degree roll attitude), the roll interconnection reducesmaximum roll velocities by 40 percent as compared to calculatedmaximum values for the basic OH-6A.

The dynamics of the interconnected landing gear system also result inthe helicopter seeking a level attitude without overshooting, as shown inthe top part of Figure 15. Due to instrumentation malfunction, theexperimental roll angles are calculated using the measured roll velocities.

A comparison between experimental and calculated rolling velocities forthe interconnected landing gear is also shown in Figure 15. The compar-ison shows that the experimental values are less than calculated primarilydue to the lower interconnected spring and damping rates used in the

experimental hardware. The reasons for this have been discussedpreviously during the discussion of landing gear pitching behavior.

Generally, the interconnected landing gear reacted dynamically as expected.The geometric action of the interconnected landing gear can be determinedby examining the interconnected displacements shown in Table A-4. Inthe longitudinal axis for nose-high landings, both the rear right and leftinterconnect chambers compressed and the forward right and left inter-connect chambers extended. The reverse was true for the nosedownlanding. In level landings, all four interconnect chambers compressed.The compressions were generally unequal due to the aft CG location andslight roll and pitch angles at contact.

In the lateral axis for left skid down landings (condition 8), both the leftfore and aft interconnect chambers compressed and the right fore and aftinterconnect chambers extended. The interconnect action was also

t 38

-10 NERNNC

-j0

4 : BASIC OH46A.j o (CALCULATED)

1i 0

0a

U 20__ _

AZ - 6.5 FETCPRNSCON

wi--0DGES RGTSI PGROS CALCULATEDPUND

LUINECONC

o 40 ___

so6

I 00 __ _ _ _ _ _ _ _

0 0.05 0.1 0.15 0.2 0.25

TIME, SECONDS

Figure 15. Effect of landing gear interconnection on roll angle and

velocity for a 6.55-foot-persecond vertical drop.

39

evident in the simulated lateral speed landing (condition 6). In this mode,as shown in Figure 7, the right skid contacts first on a horizontal platformwhile the left skid contacts an inclined surface. For this condition, theright fore and aft interconnect chambers compressed and the left fore andaft interconnect chambers extended.

A detailed comparison of the interconnect action indicates that theextension was generally unequal and less than the compression interconnectmovement. This is independent of landing attitude or test condition. Evenin level landings with compression of all four interconnect chambers, theaft interconnect chamber compressed more than the forward chamber dueto the aft fCG location. The inequality of interconnect compression andextension was due to a combination of hydraulic fluid leakage and airtrapped within the system. The sources of hydraulic fluid leakage werethe pressure relief valve which was shown to have a slow leak during theshake tests. Anotjher source of leakage may have been the seal betweenthe upper and lower interconnect chambers. During some drop tests, ahydraulic mist was observed being expelled from the modified oleodampers. In addition, the damping orifice in the face of the pressure

equalizer piston also contributed to the unequal compression and extension.This design deficiency and the corrections have been discussed earlier.

The landing gear did not fail until the final drop condition, which wasdesigned to evaluate the MIL-STD-1290 criterion of 20-foot-per-secondcontact velocity without fuselage impact. For this drop condition, thelanding gear was dropped from a skid height of 6 feet, impacting at19. 2 feet per second in a level attitude. The lift load was approximately75 percent of the desired level of 2550 pounds because the increase indrop energy excet.ded the lift simulation capability of the system.

The results of the MIL-STD-1290 evaluation are shown in Figure 16 andin Figure A-12, A-13, and A-14. Three landing gear cross tubes yieldedand the fourth fractured. Three of the four oleo dampers bottomed andthe right forward oleo attach lug fractured. (

The fractured cross tube was the right aft cross tube and it may have

fractured for reasons other than the forces experienced during this drop.The data indicate that the maximum right aft oleo load was the smallestof the four oleo loads. The oleo load is a qualitative indication of theloads in the individual cross tubes. This implies that the fractured crosstube may have been affected by previous testing and cause it to fracturebefore the other cross tubes. If the drop test was repeated with freshcross tubes, it is probable that all four cross tubes would have yielded butnot fractured.

40

1*

4112

-4

0

C)

U)

4-.

-4

I-.C)4--.*1

C)

-4

UC)

00

C)

C)

ii

The level of the oleo damper load is a strong indicator of the severity ofthe loads in the landing gear and its supporting fuselage structure. Thedata from Test Condition 12 indicate that the right front oleo maximumload was 10, 002 pounds and the left front oleo maximum load was 6202pounds. The present OH-6A is designed for a 6750-pound ultimate loadin the front oleo. This value reflects a minimum ultimate margin ofsafety of 6 percent. As can be seen by comparison, the interconnectoleo damper loads exceeded or approximately equaled the OH-6A fuselagestructural strength. In the case of the right front oleo, the oleo attach-ment lug was fractured during testing.

A comparison of front oleo loads indicates that the right front oleo was3252 pounds overloaded with reference to the design ultimate load. Thiswas caused by the fracture of the right aft cross tube and subsequenttilting of the helicopter. With reference to Figure A-26, the right frontoleo experiences a sharp 3000 pounds overload in the hundredth of asecond following the right aft cross tube fracture. The overload conditionis terminated by the fracture of the oleo attachment lug. Consequently,it appears that the landing gear structural elements such as the oleoattachment lug and the cross tube are the limiting elements.

It is difficult to predict the probable performance of the interconnectedlanding gear in satisfying the 42-foot-per-second vertical impactrequirements of MIL-STD-1290. Specifically, MIL-STD-1290 requiresthat the landing gear must be capable of decelerating the aircraft atnormal gross weights from 20-foot-per-second downward vertical velocitywithout allowing the fuselage to contact the ground. The aircraft structureexcept the rotor blades and the landing gear shall be flightworthy afterthis impact.

The interconnect landing gear was predicted to have an energy-absorbingcapability of 19. 5 feet per second based on ground contact and predictedloads. The results of Test Condition 12 indicate that the interconnectedlanding gear would provide that capability based on ground contact. How-ever, the loads exceeded the OH-6A fuselage design loads. The OH-6Alanding gear was designed to absorb a 12-foot-per-second impact and thesupporting structure designed accordingly. Consequently, the intercon-nected landing gear is limited to approximately 14 feet per second basedon fuselage structural limits. (The 14-foot-per-second value is derivedby interpolating between the maximum okeo loads measured at Test Condi-tion 12, AV2 = 19.2 feet-per-second, and Test Condition 7, AV? = 6.25feet-per-second. ) If the OH-6A structure is strengthened to accept thehigher loads, the interconnected landing gear raises the OH-6A maximumenergy absorption capability to 33. 7 feet-per-second. If the structure isnot strengthened, the capability is 30. 9 feet-per-second, which isapproximately the present design value for the OH-6A.

42

• - '_ _ _ _ _ _ _

I I Ia I I i . . ..-

The maximum energy absorption for the OH-6A with the interconnectedlanding gear is determined using data and an analysis outlined inReference 5. In brief, the analysis adds the increase in landing gearcapability in the following manner.

(aV)AV) 2 + (AV ) 2 (AV)2z OH-6A z OH-6A z nt. z OH-6A

with Int. L.G. L.G. L.G.

F 2 22= (30 fps) + (19.5 fps) - (12 fps)

(AVz)OH-6A = 33.7 fps

The load factors experienced during the final drop indicate that if thefuselage structure had been reinforced, a survivable landing with minoror no injury could have resulted. The peak load factor experienced at theCG was 5.49G. This peak value was a spike value superimposed on amean load factor of 4. 23G for approximately 0. 20 second duration. Usingthe data on the limits of human tolerance to vertical deceleration as definedin Reference 6, it is seen that the mean load factor and its duration arewithin the boundary of minor injury (Figure 17).

LIFE-CYCLE COSTS

A detailed cost estimate was conducted for the interconnected landinggear to determine the benefits for both retrofit and forward production(initial installation) in the OH-6A. This section presents a summary ofthe analysis used, assumptions, and results. The detailed calculationscan be found in Appendix B.

The analysis followed the procedures of a bottom-up approach ratherthan the technique of parametric relationships, such as changes in weightor piece part count. In support of the bottom-up approach, a document

i search and review was conducted to determine landing damper (or oleo)performance in the past in terms that would have bearing on operating

and Support costs. Failure modes, failure rates, remove and replacerates, and average time for maintenance action were included in theavailable information. This information along with an analysis of the

43

- i , ! a ,,m ! I .i i l-',

100 -' . . .-- V =65 FT SEC t

0% 1 1 -- xV-60FTSEC -

00u 30 MODERATE ,

LL l INJURY I ,EJECTION SEAT2 2

- IMINOR INJURY DESGN15 IW NO INJURY-J 10

II II

C -- I IINTERCONNECTED2 t LANDING GEARccIZ - 19. FPS o TSE ,

U I

I FORCE 3OFTSECN2

V INITIAL VELOCITY IN FEET PER SECOND

0.001 0.01 0.10 1.0

t - DURATION OF ACCELERATION, SECONDS

Figure 17. Limits of human tolerance tovertical deceleration (derivedfrom Reference 6).

components and functional design, enabled prediction of the MTBF forthe new configuration damper and an average MTBF for the other com-ponents of the interconnected landing gear system.

The cost of retrofit and forward production used in this analysis wasgenerated by the HH DTUPC (Design to Unit Production Cost) group andrepresents an average unit production cost. Learning curve correlationswere used with retrofit quantities of 100, .200, and 400 shipsets and withthe forward production of 100, 200, and 500 shipsets.

The assumptions used in the analysis are generally conservative in thatonly blade/tail boom strike benefits were included. Increases in auto-rotational landing controllability due to pitch and roll interconnectionwere not included due to uncertainty in quantifying the benefits. All theassumptions are presented below.

0 The increase in operating and support cost for the interconnectedlanding gear is due to an increase in unscheduled removals andreplacements.

44

............. 4

" The distribution of repair time for "On" aircraft and "Off" air-craft repairs is identical.

* The old-style damper replacement rate of 3. 4/1000 hours issuperseded by a projected new style damper replacement rateof 3. 635/1000 hours.

* The additional hydraulic elements of the interconnecting designwill have a combined replacement rate of 1. 280/1000 hours.

• The mean MMH/UMA (Maintenance Man-hour per UnscheduledMaintenance Action) for each of the hydraulic elements of thelanding gear is 3.5 hours.

* The inventory of OH-6As for retrofit consideration is 400 aircraft.

* The utilization of OH-6As after retrofit varies from 8 through

30 flight-hours per month.

e The service life of the OH-6As after retrofit is projected to befrom 10 to 13 years.

* The maintenance float is 14. 2 percent of the year-end inventoryof aircraft. For flight utilization less than 20 hours/month themaintenance float is reduced proportionally.

* The mean retrofit and production rates for the OH-6A will be100 aircraft per year (8. 3 per month).

0 The service life of new production OH-6As will be 20 years.

* Tail boom chops of OH-6A aircraft of the current configurationoccur at a rate of once every 5600 flight-hours.

* Tail boom chop repair requires an expenditure of $30, 968 (1972dollars).

Is Downtime for retrofit causes a loss of 24 flight-hours.

a Replacement part supply utilizes 20 percent new parts and80 percent rebuilt parts.

45

0 Increase in maintenance man-hour requirement is consideredto be so small that it does not require additional numbers ofmaintenance personnel.

* The interconnected landing gear is effective in eliminating atleast 80 percent of tail boom chops.

* All monetary calculations are based on 1972 dollar values.

The results of the analysis are summarized in Figure 18 and Table 2for retrofit costs and in Table 3 for forward (initial) production. Theanalysis shows that the cost of the integrated landing gear is lower inforward production than in retrofit. This is because the expected lifeof the aircraft is longer (20 years versus 13 years), making greater costbenefits possible. The analysis shows that the cost of the interconnectedlanding gear adds $3, 000 (1972 dollars) to the cost of an OH-6A landinggear, but the reduction in tail boom chops and elimination of theassociated repair costs offset the initial cost of forward productionaircraft.

600K

400K

2000A/200K ";100 A/C

BENEFIT

0FLIG HT-HOURS PER MONTH_

24 2

-j-200K

-400K O

DEFICIT -400K

-600K

-1000K

Figure 18. Cost savings of interconnectedlanding gear - retrofit aircraft.

46

TABLE 2. COST SAVINGS OF INTERCONNECTED LANDING GEARFOR RETROFIT (1972 DOLLARS)

100 A/C 200 A/C 400 A/C

10 years 13 years 10 years 13 years 10 years 13 years

8 hours/ -392, 156 -298, 826 -584, 703 -423, 208 -1, 333, 295 -993, 077month

ZO hours/ -69,338 30,071 26,363 22,654month

30 hours/ 85,894 182,298 244,284 528,278month

TABLE 3. COST SAVINGS OF INTERCONNECTED LANDINGGEAR FOR FORWARD PRODUCTION (1972 DOLLARS)

Utilization 100 A/C 200 A/C 500 A/CRate. 20 years 20 years 20 years

Z0 hours/month $784,658 $1,568,965 $4,312,039

Investment 336,700 673,400 1,683,500

Return on 2: 1 2.3:1 2. 6:1Investment (RDI)

-us 47

The cost benefits of retrofitting the interconnected landing gear are

dependent on the number of flight-hours. As shown in Figure 18, the

cost reductions due to elimination of tail boom chops offset the cost of

the interconnected landing gear when flight-hours exceed approximately20 hours per month. The present OH-6A fleet are operated primarily in

the National Guard, and monthly flight-hours are difficult to accurately

estimate. Present estimates are approximately 8 hours a month, butthere are indications that this will rise to 20 hours a month due to agreater military reliance on the National Guard. The effect on retrofit

costs of a reduction in aircraft service life is shown in Table 1. When

service life is reduced to 10 years from the assumed 13 years, the costbenefits are reduced. Again, service life is difficult to estimate due to

uncertainty in National Guard usage.

The cost benefits of incorporating the interconnected landing gear in new

production are shown in Table 3. In new production aircraft, use of the

interconnected landing gear results in cost savings of up to $4, 312, 039

for a fleet of 500 aircraft. These savings represent a return on invest-ment of from 2:3 to 2.6:1 in constant 1972 dollars relative to the initial

cost of the interconnected landing gear.

48

CONCLUSIONS

1. The interconnected landing gear reduces the nosedown pitchingvelocities and angles during autorotation landings. In the particularcase of a noseup landing with forward speed, the interconnected land-ing gear reduced the pitching velocities approximately 60 percent ascompared to predicted values for the standard OH-6A landing gear.

2. The benefits of the interconnected landing gear are also found in theroll mode. In one case, the roll velocities were reduced 40 percentand attenuated over a longer time.

3. The landing gear was shaken over a wide range of frequencies andno resonance was identified.

4. As compared to MIL-STD-1290, the interconnected landing gearincreases the landing gear absorption capabilities as compared tothe standard OH-6A landing gear. An OH-6A equipped with the

interconnected landing gear could absorb approximately a 33. 7-foot-

per-second impact without serious injury to the crew if the fuselagesupport structure was strengthened.

5. A cost analysis indicates that the incorporation of the interconnectedlanding gear in new production aircraft would result in savings morethan twice the costs. Retrofitting the current OH-6A fleet with theinterconnect landing gear would save money if the flight hours peraircraft exceeds 20 hours per month.

p49

I~~~M *- -...

RECOMMENDATIONS

Based on the results of this effort, it is recommended that:

1. A wheel-type interconnected landing gear be designed and tested.

2. An interconnected landing gear be designed and tested for a Scout-type helicopter, such as the OH-58, with cross tube-skid landinggear.

3. A flight test version of the interconnected landing gear be designed,manufactured, and flown.

so

R EFERENC ES

1. LOGAN, A. H., "Analytical Investigation of an Improved HelicopterLanding Gear Concept, " USAAMRDL-TR-76-19, August 1976,U. S. Army Air Mobility Research and Development Laboratory,Fort Eustis, Virginia, AD A029372.

2. MAGULA, A. W., "369A6000B Production Landing Gear Drop Tests2800 lb Gross Weight, using 369A6300 Dampers with 369A340-601Bladders and 369ASK 150 Double Acting Dampers, " Hughes ToolCompany - Aircraft Division Report 369-BT-3609.

3. MAGULA, A. W., "369H90006 Regular Production Extended LandingGear Drop Tests (2550 lb Gross Weight)" Hughes Tool Company -

Aircraft Division Report 369-BT-3033, 1969.

4. MAGULA, A. W., "Drop Tests of Improved Landing Gear ior Model

369A Helicopters, " Hughes Tool Company - Aircraft Division369-BT-3613, May 1971.

5. GOODALL, R. E., "Advanced Technology Helicopter LandingGear," USAAMRDL-TR-77-27, October 1977, U. S. Army AirMobility Research and Development Laboratory, Fort Eustis,Virginia, AD A048891.

6. SMITH, H. G. and McDERMOTT, J. M., "Designing forCrashworthiness and Survivability, " American Helicopter Society24th Annual National Forum Proceedings, Washington, D. C.,Mpo- !968.

51

APPENDIX A

INTERCONNECTED OH-6A LANDING GEAR DROP TESTSOF AN ENERGY DISTRIBUTION SYSTEM FOR

HELICOPTER LANDING GEARSDURING HARD LANDINGS

1. 0 INTRODUCTION

This appendix presents the results of drop tests to determine the structuralintegrity and functional response for the interconnected landing gear. Alltesting was accomplished within the Hughes Helicopters Complex. Theground resonant portion of the tests was discontinued after failure to obtainany resonant data. The drop tests were performed from 20 January 1977to 27 February 1977 at Hughes Helicopters, Culver City, California.

2.0 TEST OBJECTIVE

The objective of these tests was to experimentally determine the structuralintegrity and functional response of the interconnected landing gear in thelanding modes established in Reference 1.

3.0 DESCRIPTION AND LOCATION OF HARDWARE

An existing OH-6A drop test fixture was modified by lowering the braceand cross tube attachment points 1. 72 inches. This allowed the instal-lation of the four modified oleo dampers. The interconnect system alsorequired four pressure equalizers, five surge accumulators, and four dualflow valves, which were installed per Drawings 369-ASK-2060 and 369-ASK-2058. Figures A-I and A-2 show this system installed on the testvehicle.

The ground resonant test was conducted for the baseline configurationper the conditions in Table A- 1. The interconnect spring constant wasapproximately 34 pounds per inch and the damping orifice diameter was0. 128 inch.

The drop tests were conducted according to the configurations defined inTable A-Z.

Changes to the interconnect system included spring and damping variations.The center of gravity (CG) for the vehicle was FS 104 for all tests exceptTest Condition 5, which was FS 97. All gross weights were 2550 poundsexcept Test Condition 9, which was 2800 pounds to simulate overload.

52

3. 1 Conformity

The simulated OH-6A test vehicle conformed to the test plan. The appli-

cable drawings for the test configuration are as follows:

OH-6A Drop Vehicle 369-9302 SH-2Shake Test Setup 999-0490Hydraulic Schematic, Damper Interconnect 369 ASK 2058

AssemblyInterconnected Landing Gear Installation 369 ASK 2060Displacement Transducer Installation 999-0489

4.0 METHOD OF TEST

4. 1 Shake Test

Input shake amplitudes, measured by a built-in linear variable differentialtransformer (LVDT), were used to control a hydraulic actuator installedbetween a test structure and the drop vehicle. A load cell in series withthe actuator allowed simultaneous load monitoring. The capacity of thishardware was ±1000 pounds and a ±3-inch stroke.

A spectral dynamics sweep oscillator was used to input the stroke fre-quency. An MTS servo controller was used to monitor input and feedbackfrom the actuator. Load and frequency were monitored on an X-Y plotter.An oscilloscope was used to obtain load-deflection Lissajous figures. Theinterconnect deflections, pressures, landing gear strain gages, acceler-ometers, lift load, and gyros were monitored on a 50-channel visicorder.Teflon rings were fastened to the skids at the aft and middle pad locations.The vehicle was placed on four greased steel plates. The lift load wasapplied through a torsion bar attached by a cable.

4. 2 Drop Test

Drop tests were configured per the test plan. The vehicle was hoisted abovethe landing surfaces by a gantry located at remote Test Site 2. Figure A-3shows a schematic view of a typical drop setup. The vehicle was releasedfrom an air-release hook mechanism, and free fall developed the requireddrop velocity. At a preset position, the lift beam begins to apply the liftload. An air cylinder, pressurized to a preset pressure, provides thesimulated rotor lift while allowing the test vehicle to continue its descent.A hydraulic actuator attached between the lift beam and load cell is usedto help dampen any oscillatory loadings that may occur during the suddenlift loading. The lift load cell attaches to the vehicle at the CG specifiedin the test plan.

j 53

:o .' -_ _ __'_ _......_ _l._-__ _

The drop velocity is measured by a magnetic pickup that senses therotating gear teeth. The gear is driven by a wire attached to the vehicle.This wire wraps around the gear's drive pulley and then attaches to abungee cord which maintains tension. Table A-3 presents the ballastand locations to obtain the required CG for each configuration.

5.0 DATA ACQUISITION AND REDUCTION

5. 1 Shake Test

Load versus frequency data was recorded on an X-Y plotter and load versusdeflection data on a;i oscilloscope. A Polaroid camera was used to obtaina permanent record of the load-deflection Lissajous figures. The requestedparameters pertinent to the interconnect system response were recordedon a 50-channel visicorder.

5. 2 Drop Test

All parameters listed in the test plan were recorded on a 50-channel visi-corder. High-speed photography was used on each drop condition. Thiswas accomplished using two cameras positioned to view the most criticalmotions of the gear. Generally, the cameras were positioned to view theside and front of the landing gear.

The drop velocity was calculated using 0. 01-second increments.

6.0 TEST RESULTS

6. 1 Shake Test

Input shake amplitudes of 0. 25, 1. 0, and 2. 0 inches peak-to-peak wererun for a lift condition of 90 percent. The gross weight of 2550 poundswas used for all shake tests. Frequency scans were run per Table A-Iand the load versus frequency plots are presented in Figures A-4, A-5,and A-6. It did not seem clear where any resonant points were for the0. 25-inch amplitude run. For the 1. 0-inch P-P run, the Lissajous at1. 38 Hz is shown in Figure A-7. It may be noted that the correspondingfrequency in Figure A-5 does not indicate a resonant point. For the2. 0-inch P-P run, the Lissajous at 2. 15 Hz is shown in Figure A-8. It

54

47 ; .. .

may also be noted that the corresponding frequency point in Figure A-6does not indicate a resonant point.

The lift load was removed and frequency scans were run for 0. 25-inch and2. 0-inch P-P amplitudes. The load versus frequency plots are presentedin Figures A-9 and A-10. Again, nautral frequency response was unde-finable, and when the input rod failed, testing was discontinued. It appearedthat the data presented may have been affected by hydraulic leakage past theseals and the pressure relief valve. No usable data were obtained on thevisicorder trace of the other requested instrumentation.

6. 2 Drop Tests

The drop test vehicle was configured to the parameters given in Table A-2for each drop. The vehicle was hoisted into position and ropes were tiedto the skids to prevent vehicle rotation. The interconnect system was setso that approximately 1. 9 inches of extension were showing on the aft oleosand 1. 6 inches of extension on the forward oleos. Figure A-11 shows thedeflection transducers used to measure interconnect and damper motions.It can be noted in the photograph that the interconnect is fully retracted dueto leakage in the pressure relief valve. When the weight of the vehicle wasremoved, the interconnects would extend toward their normal positions, butnot equally at each oleo due to differences in internal friction and the abil-ity of the oil to flow around the system. The drop test data are summarizedin Table A-4 and the complete visicorder traces of all data for all test

conditions are presented in Figures A-15 through A-26.

Test Conditions I through 5 used a 26. 5-degree ramp as the landingsurface. The ramp was covered with a steel plate and is the same type ofsurface as used for previous drop testing. The forward end of the landinggear was pitched up 10 degrees to the surface prior to the first drop. Thegross weight was 2550 pounds with the CG at Sta 104. Rotor lift was set at67 percent and application began approximately 1. 5 inches above groundcontact. As the landing gear made contact, the aft interconnects con-tracted and the forward interconnects extended. The amounts are given inTable A-4 along with the other recorded data. Oil mist was seen blowingoff the dampers during initial contact. This was attributed to oil filmbuildup on the inner piston or seal blowby. This, along with the leakagein the pressure relief valves, and trapped air in the system are possiblereasons why the deflection in the interconnect system does not add up(i. e., compression = extension). It was noted that the pitch and roll

attitude traces showed little or no motion even though pitch and roll rate

t 55

traces did show motion. This indicated that the gyros were probablynot functioning properly. Since no replacements were available, nochanges were made. A calibration trace, taken prior to each drop,indicated the proper response, but during each test little or no responsewas obtained.

For Test Condition 2, the pressure equalizers were removed and sent toMOOG, and the pressure equalizer piston was exchanged for one with asmaller orifice. The Test Condition I orifice was 0. 128 inch in diam-eter and Test Conditions 2 through 12 used an orifice of 0.0595 inchin diameter. The results of the drop test are presented in Table A-4.Similar results to Test Condition I were obtained except the pitch ratewas noticeably reduced by approximately 16 percent.

For Test Condition 3, the pressure equalizers were again removcd andsent to MOOG, and the springs were replaced with softer ones. The newspring constant was approximately 22 pounds per inch and remained sothroughout the rest of the tests.

The results for Test Condition 3 show similar values to Test Conditions1 and 2, except for pitch rate, which was about 4 percent greater thanCondition 2 (12 percent less than Condition 1).

Test Condition 4 was conducted with the skids parallel to the 26. 5-degreeslope. This condition simulated forward speed with an aft CG. The testresults presented in Table A-4 show a low positive pitch rate, 3. 58degees per second.

Test Condition 5 was 6onfigured similar to Test Condition 4, except theCG was forward at Sta, 97. The drop test results (see Table A-4) showedthat an initial pitch rate was - 12. 45 degrees per second, pitching down infront.

Lateral speed was simulated by dropping the test vehicle on a plywoodplatform with one side sloped at 26. 5 degrees. The left skid was placedover the slope and was parallel to the fore and aft surface. This drop,Test Condition 6, was completed and the results are presented inTable A-4. The right interconnects compressed significantly and the leftextended. The vehicle was noticeably more level than it normally wouldhave been without interconnects. The roll rate of -27. 31 degrees persecond seems high, but the final position of the vehicle indicates littlemotion, even though there are no previous roll drop data with which tocompare these data.

S56

- ,% --

Test Condition 7 was conducted on plywood sheets placed on level groundand the skids were parallel to the plywood. The drop test results arepresented in Table A-4 and do not indicate any excessive or reduced values.The pitch rate was noted to have increased from approximately 4 degreesper second (Report 369-BT-3033 Record 896) to approximately 15 degreesper second.

It was decided to conduct Test Condition 10 next, as the only changerequired was to elevate the vehicle enough for an increase in free fallvelocity to a maximum of 8. 02 feet per second. Also, the lift pressurewas increased so 100 percent lift would be obtained. A peak lift of 2503pounds was measured at contact, but it had reduced to 2115 pounds by 0.25second, giving the reported average of 2309 pounds. The results given inTable A-4 are noticeably similar to Test Condition 7. The pitch ratemade the only significant change, and it was in the positive direction to17. 50 degrees per second.

Test Condition 8 was conducted similar to Condition 7, except the rightskid was raised 10 degrees for a roll attitude. As the skids made contact,the left interconnects compressed and the right extended. The roll ratewas 57. 56 degrees per second maximum. The drop vehicle appeared toroll right until reaching the level position then it continued to descend atthis attitude with no further roll motion. The lateral acceleration reacheda maximum of 2.23 G with the ground load factor.

Test Condition 11 was the next drop to be run. The flat wood was used asa landing surface. The drop vehicle was pitched forward (nosedown) sothat the base of the skids was inclined -10 degrees to the landing surface.As expected, the forward interconnects compressed and the aft extended.Table A-4 gives the recorded amounts. The pitch rate, 45. 0 degrees persecond, was in the desirable direction. The trace indicated little or nonegative pitch motion. The positive pitch motion continued until the forwardtips of the skids had raised off the landing surface to approximately +10degrees. Due to instrument problems, this drop-was made three times.It was noted that after the first drop the front pads on the skids no longercontacted the landing surface when the vehicle was sitting at rest. Bothpads appeared to be approximately 1/8 to 1/4 inch above the surface.There did not appear to be any change after the second and third drops.

The overload configuration, Test Condition 9, required that the total dropweight be increased to 2800 pounds. The CG (FS 104), drop velocity(6. 5 feet per second), and rotor lift (67 percent) were the same as pre-vious tests. The results in Table A-4 indicate that the vehicle pitched

, " ,, "' III II --

mop

forward (nosedown) and rolled left. The motion picture data show theaft descent stopping and the front continuing downward, which gave theforward pitching data results.

The final test condition, number 12, is the maximum drop condition. Inorder to obtain the 19. 5-foot-per-second drop velocity, the vehicle wasraised 72 inches above the wood landing surface. This height resulted ina 19. 2-foot-per-second fall. The lift pressure was increased to obtain100 percent lift at impact. The lift came up very strong and overshot,which may have been caused by the hydraulic damping system. The liftimmediately dropped and oscillated with an average from 1943 to 1899pounds, which was well below the desired 2550 pounds. At impact theload measured 2734 pounds. It appears that the increase in energy wastoo much for the hydraulic damper on the lift bar to sustain.

As shown in Table A-4, three of the four oleo dampers bottomed; one ofthose, the left aft, was loaded to only 2989 pounds, which was less thanthe oleo proof and some other test condition loadings on the forwarddampers. The results in Test Condition 9 did not give any indication ofoleo damper problems. Failure of the right aft strut occurred approxi-mately 0. 17 second after impact. The pitch rate of -66 degrees per secondwas caused by the forward end continuing to descend just prior to failure.This may have been due to the dampers that had bottomed approximately0. 12 second after impact. The peak of negative pitch rate was reached at0. 24 second. The pitch rate immediately became positive and by roughly0. 32 second, it had peaked at 93 (estimated) degrees per second. Thiswas observed in the movies showing the flat descent stop as the aft endbottomed and the forward end continued to descend. At failure, the aftend again descended until contact with the ground was made with the stubstrut. Total fracture of the strut was incurred, as was failure to theright forward oleo attach lug. The failures are shown in Figures A- 12,A-13, and A-14. Figure A-12 shows the total fracture of the rear rightstrut, below the elbow. Other items of interest in the picture include thehydraulic hose broken away from both rear oleo interconnects (arrows 2and 3). These lines attach to the reservoir. Also, note a failure of thefixture at the forward center attach point of the cross tubes (arrow 4).Figure A-13 shows the same details as Figure A-12, but it also shows therear left damper almost fully compressed. The front right damper waseven more compressed, as can be seen in Figure A-14 (arrow 2). Thelug failure can also be seen (arrow 1) along with damage to the damperdeflection transducer.

f 58

The right and left forward brace vertical bending bridges appear to havefailed shortly after the strut failure. The values in Table A-4 are takenbefore or near the failure time.

The high roll rate was caused by the strut failure and occurred afterfracture.

7.0 CONCLUSIONS

The interconnect landing gear system was able to perform responsivelyto reduce many of the undesirable characteristics involved with anautorotational-type landing. This functional capability was obvious inTest Conditions 6 and 8, where expected pitching did not occur. The flatsurface landings were all showing level descent and virtually no forwardpitching.

The structural integrity of the oleo damper was more than sufficient andthe hose failure incurred on the final drop did not affect the functioningof the system for that landing, as it was part of the upper system whichextends the interconnects to their middle position after lift-off.

8.0 RECOMMENDATIONS

Based on the testing, it is recommended that the following changes beincorporated in the design. First, better seals with wipers be installedin the charging and pressure relief system to eliminate leakage. Second,a self-centering ability be incorporated in the system. These changescould be accomplished by eliminating the orifice in the pressure equalizerpiston face, and by a motion sensitive orifice or a separate dampingsystem.

j59

TABLE A-I. GROUND RESONANCE TEST

P-P Input Frequency

GW CG Lift Amplitude Sweep Range(Ib) (Station) (%) (in.) (Hz)

2550 104.0 90 0.25 1.0 - 10.0 - 1.0

1.00 1.0 - 5.0 - 1.0

2.00 1.0 - 4.0 - 1.0

0 0.25 1.0- 10.0- 1.0

1.00 1.0 - 5.0 - 1.0

2.00 1.0 - 4.0 - 1.0

60

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H 0 0 0 0 0 0 0 0 0e

cy,.- : !n ! t t N N N1N N N9

p. 4 4

0.U'4 m 0 0 0 0 0- 0D 0, 0a

+6

TABLE A-3. DROP TEST BALLAST SUMMARY FOR THEMODEL 369A INTERCONNECTED LANDING GEAR

CONFIGURATION*

Moment of Inertia

Load Gross Weight Horizontal CG Vertical CG (slug-ft2 )Condition (Ib) FS (in.) WL (in.) Roll Pitch Yaw

1 2550 104.0 28.9 318 868 751

2 2550 97.0 29.9 318 872 755

3 2800 104.0 28.9 328 888 768

Ballast Weight (lb)Ballast Locations (in.) lad C it(ls

Ballast Load ConditionsBox H-Arm (FS) L-Arm (BL) V-Arm (WL) No. I No. 2 No. 3

A 35 0 26.8 260 400 260

B 165 0 26.8 444 301 444

C 79 0 63.8 133 128 133

D 121 0 63.8 83 88 83

E 100 -35 29.3 0 0 0

F 100 135 29.3 0 0 0

GL 84 - 8 18.3 0 0 50

GR 84 8 18.3 0 0 50

HL 116 - 8 18.3 0 0 75

HR 116 8 18..3 0 0 75

*Using drop test fixture, P/N 369-9302

62

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. ... . .t ... N.0.

! S 4 00. tz-N~

..* OONO~f.11 -TNNONO.

-- 7 - -

---- 4 0 0 . .-. C -- NN-

Figure A- I. View of modified oleo damper interconnect installation.

1 igutrec A -2. View\ of ii od ifi olvd d) iamnper inter to ntie t install ation.,

W4

AIR CYLINDER-TANK ASSEMBLYHTC-AD 999-0410

-T -r- ISTON ASSEMBLY

GANTRY

ENERGY ABSORPTION SYSTEM PROVIDESSIMULATION OF ROTOR LIFT AFTER 2-TON ELECTRIC HOISTREQUIRED FREE (H) FALL (SLIDE RODSWHICH STRADDLE HOIST, CATCH RELEASE HOOK MECHANISM (AIRSPECIMEN AT END OF DROP BY MEANS ACTUATE SOR ESTOF ATTACHED NUTS) _jACTUATED) SUPPORTS TEST

SPECIMEN AT SPECIFIED DROP

HEIGHT UNTIL RELEASE

PRESSURIZED HYDRAULIC STRUTAND RELIEF VALVE SYSTEM TOSNUB LIFT LOAD AT IMPACT

IF

LIFT BEAM BELIF

DROP FIXTUREHTD-AD 369-9302

SKIDS LEVEL 00

DROP SURFACE VELOCITY PICKUP/H ,HI IH ,H rift#

LEVEL DROP CONDITION

Figure A-3. Landing gear drop test fixture used for theinterconnected landing gear tests.

65

420

350

280

z0

. 210

0-J

140

70

0 1 2 3 4 5 6 7 8 9 10

FREQUENCY, Hz

Figure A-4. Plot of load versus frequency atinput 0. 25 inch P-P, 90 percent lift.

420-

3 INPUT ROD SHAKING

350-

280

z

210 -

o0 X

J140-

70- _

LISSAJOUS SHOwNIN FIGURE A-7

00 1 2 3 4 5 6 7 8 9 10FREQUENCY, Hz

Figure A-5. Plot of load versus frequency atinput 1 inch p.p, 90 percent lift.

66

a-.i

42 -- __ --

350

280

z

a210-

0

140to

\LISSAJOUS SHOWNIN FIGURE A-S

70

0 1 2 3 4 5 6 7 8 9 10FREQUENCY, Hz

Figure A-6. Plot of load versus frequency atinput 2 inches P-P, 90 percent lift.

Figure A-7. Lissajous of 1-inch P-P at 1. 38 Hz, *77 pounds.

67

Y-

Figure A-8. Lissajous of 2-inch P-P at 2. 15 Hz, +175 pounds.

420- - - - - - -

350

280 - __________'fl~ INUT ROD SHAKE

z

0

70 A

01,1

FREQUENCY, Hx

Figure A-9. Plot of load versus frequency at input.0. 25 inch P-P, 0 percent lift.

68

INPUT ROD FAILURE

280

210 _

0

140

0

01

0 1 2 3 4 5 6 7 8 9 1

FREQOi2NCY, Hz

Figure A- 10. Plot of load versus frequency at input2 inches P-P 0 percent lift.

Figure A-il1. View of deflection transducers used to measureinterconnect and damper motions.

69

Figure A-12. View of aft right strut failure, Test Condition 12.

Figure A-1 3. View looking forward at failed landing gear,Test Condition 12.

t 1 701

Figure A-14. View of failed lug on front right strut.

71

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APPENDIX B

COST ANALYSIS CALCULATIONS

LIST OF CALCULATIONS

Page

CALCULATION OF PHASED UTILIZATION

Retrofit of 100 aircraft flying 8 hours/month for 10 years.. 87



Forward Production of 100 aircraft flying 20 hrs/mon. forZ0 years .................................. 90

Forward Production of Z00 aircraft flying 20 hrs/mon. for20 years ..... ...... ............ ........... 91

Forward Production of 500 aircraft flying 20 hrs/mon. for20 years .................................. 92


Retrofit of ZOO aircraft flying 8 hours/month for 13 years.. 94


Retrofit of 100 aircraft flying 20 hours/month for 13 years. 96

Retrofit of 400 aircraft flying 20 hours/month for 13 years . 97





*Retrofit of 100 aircraft flying 30 hours/month for 10 years. 102

Maintenance Float Increased in Proportion to Flight Hour. Increasefrom 20 to 30 hours per month.

84

. w n n ,m mm m. n mm . mmm .0. -mm

LIST OF CALCULATIONS (CONT)

Page

TAILBOOM CHOPS -NUMBER OF OCCURRENCES

New procurement of A/C flying 20 hours/month for20 years .. . . . . . . . . . . . . . . .. 103

Retrofit 100, 200, 400 A/C flying 8 hours/month for10 years . * ... * . . ....... . * . . . . 0. . . ....... 104

Retrofit 100 A/C flying 20 hours/month for 10 years ..... 105

Retrofit 100 A/C flying 30 hours/month for 10 years ...... 106

Retrofit 100 A/C flying 30 hours/month; 10 years, Maint.Float 21.3% ...... ....... ................ . 107

Retrofit 100 and 400 A/C flying 30 hours/month for13 years ............................................ 108

Retrofit 100 and 400 A/C flying 20 hours/month for13 years ............. ........ .. ... ; ....... 109

Retrofit 100, 200, 400 A/C flying 8 hours/month for13 years ............................................ 110

INVESTMENT COST

Purchase of 100, 200, 500 new procurement A/C ....... .. 1

Retrofit at Hughes of 100, 200, 400 A/C ................ 112

SPARE PARTS COST AND OVERALL COST IMPACT OF CHANGE

Retrofit 100, 200, 400 A/C; flying 8 hrs/mon; 10 years;all new spares ...... ................................ 113

New Production 100, 200, 500 A/C; flying 20 hrs. mon.;Z0 years all new spares ....................... 114

Retrofit 100, 200, 400 A/C; flying 8 hours/mon; 10 years;new and rebuilt spares ...... ......................... 115

New Production 100, 200, 500 A/C; flying 20 hours/mon;20 years; new and rebuilt spares ...................... 116

85

LIST OF CALCULATIONS (CONT)

Page

Retrofit 100 A/C; flying 20 hours/month; 30 hours/month

10 years; new and rebuilt spares ................. 117

Retrofit 100 A/C; flying 8, 20, and 30 hours/month; 13 years;

new and rebuilt spares ...... ....................... 118

Retrofit 400 A/C; flying 8, 20, and 30 hours/month; 13 years;

new and rebuilt spares ........ ....................... 119

Retrofit ZOO/A/C; flying 8 hours/month; 13 years; new and

rebuilt spares ............................... . 120

4

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