Mechanical Design of the LMDE.pdf - Heroicrelics

MECHANICALDESIGNDFTHE LUNARMODULEDESCENTENGINEBY .JACK M . CHERNETRW.n¥"M$ C_HTTP'I/SPACEAHOL.IC.COMHTTP://HERCICRELlCS.CRG

MECHANICAL DESIGN OF THELUNAR MODULE DESCENT ENGINEby Jack M. CherneManager, Engineering Design DepartmentPower Systems DivisionTRW SystemsRedondo Beach, California, U.S.A.IntroductionTo land astronauts on the moon, the Lunar Module must descend froma lunar orbit to a hovering position above the surface of the moon, select asite and descend to a soft landing. TRW Systems has translated these propulsionrequirements into the Lunar Module Descent Engine (LMDE).The LMDE is a pressure-fed, bipropellant, variable thrust, gimballing,chemical rocket engine with a maximum thrust of 9850 lbs, throttleabledown to approximately 1000 lbs . The throttling is obtained by means ofdual , variable area, cavitating venturi flow control valves mechanicallylinked to a variable area injector. Ignition is hypergolic in the combustionchamber since the propellants are nitrogen tetroxide (N 0 ) and aZ 450-50 mixture of hydrazine (N H ) and unsymmetrical dimethylhydrazine (UDMH).Z 4Figure 1.Lunar Module DescentEngineThe time available for developingand qualifying the LMDE formanned flight was short. To assurethat unforeseen problems did notdevelop during the latter part of thedevelopment and qualification phase,it was important to plan an adequatecomponent test program early in theengine's development. This programincluded tests of details and subassembliesto locate weaknesses in thedesign. Upon correction of thesedeficiencies, the next level of assemblywas tested. Then, step by stepassurance was developed that the completeengine performed as required.Details of the mechanical design ofthe LMDE (Figure 1) are describedlater in the paper. All aspects ofthe design, including discussion ofthe critical environments and thedevelopment problems anticipated, arepresented along with the componenttests used to find the solutions tothese problems.- 1 -

Engine DescriptionCOMffNSAT1NGSPItiNGFigure 2.The LMDE is a compact packageweighing 394 lbs and having overalldimensions of 90-1/2 inches by 59 inchesdiameter. Interface with the LunarModule vehicle is made at two trunnionsthrough the gimbal frame, providingcapability of gimballing 6° in bothlateral directions (Figure 2). Propellantsare supplied to the LMDEthrough two stainless steel propellantinlet lines with a simple bolted flangeinterface joint. The inlet linesaccommodate the 6° of gimbal motion bymeans of a pair of flexible bellowswelded in the lines at their intersectionwith the gimbal axes. Propellantsflow from the lines throughthe cavitating venturi flow controlvalves to the quad-redundant ball shutoffvalves. From the shutoff valve,the fuel enters the manifold of thehead end. The major percentage ofthe fuel is injected axially into thecombustion chamber through the annularorifice of the variable area concentricinjector, while the remainder is distributedto the 36 ports of the barriercooling ring where it is injected alongthe wall of the chamber. The oxidizerenters the center of the injectorassembly, flows down and is directedinto the combustion chamber radiallythrough the variable area ports in ahorizontal disc composed of 36 jetswhich impinge on and ignite with thesheet of fuel.Figure 3.I flOW"'''LYE.!l!!!Thrust Control AssemblyThrust control is accomplishedby adjusting the propellant flow ratesand varying the area of the injectionports to maintain near uniform velocities.Electrical control signals fromthe Lunar Module are received by theelectromechanical throttle actuatorwhich converts the signal to a linearposition of the actuator screw jack(Figure 3). Attached to the top ofthe screw jack is a cross beam. Theright side of the cross beam is connecteddirectly to the oxidizer flowcontrol valve through a flexural element.The left side of the cross beam- 2 -

is connected to the fuel flow control valve through the mixture ratio trimlinkage, establishing the desired motion of the fuel flow control valve relativeto the motion of the oxidizer flow control valve. This produces therequired mixture ratio of the propellants. This linkage permits adjustmentof the mixture ratio during acceptance tests of the engine. Also attachedto the cross arm by means of two flexures is a walking beam, pivoted off themanifold assembly, which translates motion from the throttle actuator assemblyto the injector assembly. Motion of the movable sleeve in the injectorsimultaneously changes the gap in the fuel orifice and the size of theports in the oxidizer outlets.The manifold assembly which provides the mounting points for thethrottle mechanism is the distribution network for the propellants and theend closure of the combustion chamber. A welded titanium shell which boltsto the manifold forms the structural element of the combustion chamber andthe throat of the engine. The phenolic-fiberglass ablative liner providesthe thermal barrier between the burning propellants and the structural case.Attachments are provided in the throat area for the gimbal assembly. Thechamber and ablative liner extend to the 16:1 area ratio point where aflanged joint provides for attachment of the nozzle extension.The nozzle extension is a radiation-cooled, bell-shaped cone whichprovides for expansion of the gases from the 16:1 area ratio at the attachmentto the combustion chamber to the 47.5:1 area ratio exit.The total firing time requirement for the LM Descent Engine is1000 seconds. This is composed of 90 seconds of acceptance testing plus910 seconds of duty cycle. A nominal duty cycle is shown in Figure 4.Temperature of the titanium combustion chamber case does not exceed 800°Fduring duty cycle firing.Thermal requirements for the Lunar Module require that the externaltemperature around the engine above the nozzle extension remain below400°F during the duty cycle firing. This is accomplished by the use of acomposite stainless steel and fiberglass insulating blanket around thecombustion chamber.12,OOOr-= __ ~'-""----"-------""----,,,,,,,---,ORBIT INJECTION I10,0008,000 .THRUSTL66,0004,0002..000'(I~;' :=;:­!~ I ·1-1 ~"'M_ jt-- I ---+--1 ~I i IDESCENT ENGINE STARTSAND VEHICLE STABlUZATIONI+---' LANDi°O~~~-'~N\-~-~~~~~~-~~~~~~Figure 4.- 3 -

Engine GimbalThe gimbal assembly is a rectangular frame composed of four beamswith a modified I section machined out of 7075-T73 aluminum alloy. Theflanges of the beams are tapered in thickness and planform for minimumweight. A trunnion attaches to the center of each beam through a sphericalFabroid bearing. One pair of trunnions mounts to the engine thrust chamberand the other pair mounts to the Lunar Module support truss (Figure 5) .Figure 5.Thirteen load conditions were defined as applicable to the gimbalframe. These included pre-launch shock, launch and boost vibration, enginestart, lunar descent and lunar landing. Preliminary structural analysisindicated that seven of these were critical. Thermal analyses provided amaximum temperature in the gimbal beams of 200°F at the end of the mission .Ultimate static and sinusoidal load factors of 1.5 were used. An ultimateload factor of 2.25 was applied to random vibration environments.Prior to the tests of the complete engine in the gimbal frame fordynamic environment, several component tests were conducted to ensure thatthese components would survive their load and life requirements.After the selection of the configuration of the gimbal, two testprograms were initiated and completed successfully before the total systemtests were started. These were a static structural test to destruction ofone side of the gimbal frame and a series of life cycle tests on individualgimbal bearings for shock, cyclic rotation and vibration.Since the aluminum gimbal frame was designed for m1n1mum weight ,the static test of the single side beam was used to substantiate the stressanalysis of the tapered flange members. Subjecting an individual Fabroidbearing to the cyclic rotations experienced during gimballing of the engine- 4 -

during lunar descent, in addition to the shock and vibration environment oflaunch and boost and engine firing, revealed the need to make some changesto the bearing design. Those changes were accomplished before tests of afull gimbal frame was scheduled to start.When a complete gimbal assembly was available for further testing,an engine assembly still could not be used; therefore, it was planned touse an engine mass simulator. This unit simulated the mass properties inthe Y and Z axes but not in the X axis (the centerline of the engine thrust) .The gimbal was first subjected to the environments of salt fog,sand and dust and a limit load test to several of the critical loads. Abreakaway torque test on the bearings while subjected to a thrust load of10 , 500 lbs was used as one criteria for acceptance of the gimbal after eachtest. Alignment of the thrust axis was another criteria.Shock tests of the gimbal with the engine mass simulator wereconducted about all three axes . A shock impulse of 15 g peak with a6 millisecond rise time was applied three times along each axis. Duringthe shock test the gimbal was rigidly attached to the shock test equipment.No unexpected responses were noted.Vibration tests of the gimbal frame were next performed. Forthis test the gimbal trunnions were mounted to a structure which simulatedthe influence coefficients of the Lunar Module support structure. SeeFigures 6 and 7. During the first sinusoidal vibration sweep along theY axis an unexpected torsional mode was observed at 18 cps. Strain gageswere installed on the gimbal frame to measure bending stresses due to thetorsion and the survey repeated . As a result of the data obtained, itbecame apparent that it was necessary to correctly simulate the moment ofinertia of the engine mass simulator about the X axis. The first simulationof I was 34 , 000 lb-in 2 while the calculated engine inertia wasxx66,000 lb-in 2 . The change in inertia was made and the vibration test completed.These tests pointed out a minor structural weakness which wouldhave delayed the qualification tests if not discovered at this time. Thebolt which attached the bearing to the gimbal frame was restrained againstrotation by means of two cotter pins. These pins failed in shear permittingthe bolt to rotate during vibration. Retention of the bolt waschanged to the heat treated, machined lock ring shown in Figure 8.This specimen was then subjected to mechanical cycling of ± 6 0for 500 cycles about each of the rotational axes while subjected to10,500 lbs of thrust load.Upon completion of the mechanical cycling test, a static test toultimate loads was performed for the most critical load conditions. Thenload was applied until failure for the most critical condition with thrustapplied at 6 0 in both directions from center . Failure occurred at39,800 lbs which is 1.92 times the ultimate load taking account for gimballedangle, thermal degradation and other structural loads. SeeFigure 9.- 5 -

Figure 6.Engine Mass SimulatorMounted in GimbalFrame on VibrationTest FixtureFigure 7.View of Simulationof Lunar ModuleSupport StructureFigure 8 .Redesigned Bearing BoltRetention on Gimbal Frame- 6 -

Figure 9.Gimbal Fr ame Static Test SpecimenStress levels in the gimbal frame which were higher than anticipatedduring analysis due to the unexpected torsional mode aroused concernover the fatigue str ength of the assembly. The subject test specimen wasexposed to eleven times as many sinusoidal vibration cycles in the frequencyrange of 5 to 100 cps as is required for qualification. In addition, theexposure to random vibr ation was more than six times that required forqualification. Since the specimen withstood almost twice the ultimateload after experiencing all of this dynamic environment , confidence in itsability to perform during a lunar mission was established.As a result of the knowledge gained during the early componenttests and the gimbal frame tests with the engine mass simulator inOctober 1965, the gimbal on the design verification test of the completeengine in April 1966 and on the vibration test of the qualification enginein October 1966 performed without incident .Combustion ChamberThe combustion chamber consists of an ablative-lined titaniumalloy case to the 16 : 1 area ratio . Fabrication of the 6Al-4V alloytitanium case is accomplished by machining the chamber portion and the exitcone portion from forgings and welding them into one unit at the throatcenterline. Thickness of the shell is a uniform 0 . 035 inch except at theupper end where the head end bolts on, at the weld joint and at the lowerflange where the nozzle extension attaches. On either side of the throata pair of flanges are provided integral with the case for attachment ofZ-shaped aluminum rings . These rings form the structural members forattachment of the gimbal trunnions . See Figure 10.Since the titanium chamber is one piece, the ablative liner isfabricated in two segments and installed from either end . A metal lockingelement positively locks both halves of the liner together as they are installed.This lock is redundant since the shape of the nozzle extension- 7 -

Figure 10.Combustion Chamber Assemblyis such that the ablative liner is retained in the exit cone during transportationand launch and boost. During engine firing, thrust loads forcethe exit cone liner against the case.The titanium head end assembly attaches to the combustion chamberwith thirty-six A-286 steel 1/4 inch bolts. This joint also permits thetransfer of moments around the corner insuring a continuous pressure vessel.In order to keep the maximum operating temperatures of the titaniumcase in the vicinity of 800°F, the ablative liner was designed as acomposite material providing the maximum heat sink at minimum weight. Theselected configuration consists of a high density, erosion-resistant silicacloth/phenolic material surrounded by a lightweight needle-felted silicamat/phenolic insulation.Temperature of the area around the combustion chamber is maintainedbelow 400°F with the aid of a heat shield attached to the outsideof the titanium case. The heat shield consists of two layers of0.0015 inch stainless steel foil separated by fiberglass wool.Structurally, the ablative liner is not required to be loadcarrying in the pressure vessel. The maximum internal pressure is 116 psiwith an ultimate load factor of 3.0. Figure 11 presents the variation ofchamber pressure at full thrust and shell temperature with distance alongthe combustion chamber used for design analysis.To analyze the chamber shell and head end assembly for the structuralloads applied, it was necessary to utilize a TRW digital computerprogram developed to perform static and dynamic analysis of axisymmetricthin shells of revolution subjected to arbitrary loads. Static and dynamicresponse of complex multi-segment shell assemblies are calculated. Loads,- 8 -

including temperature effects, of arbitrary meridional and circumferentialvariation, are computed - the latter by expansion in a Fourier series ofharmonics. Orthotropic material properties, variable stiffness in themeridional direction, discrete circumferential stiffeners, and all physicallypossible boundary conditions are included.~,ov~TEMPfRATURE AND PRESSURE120I:~I:2000V,/'00MIIU8TI0N~ V~ --'IIfO'T"" 7'0I1AMI!fft f'ftES8URf ~~ ~~8I£Ll. TBIf'fIIAllJIIf.\\ ' /f-,V-I~lOOil!J900M"/OOiti"- 1=4 1Z 18 20 Z4 Z8 3Z 36 '10 44ClIWI!BI LB«mI - INCHSFigure 11.The solution is based on Fltigge's shell equations, which arereduced to four second-order partial differential equations in terms ofthe components of displacement and the meridional moment.Substantiation of the structural integrity of the combustionchamber and head end assembly was accomplished in several steps at variousstages of the program. Early in the program a room temperature bursttest of an 0.050 inch thick chamber was used to check on basic materialproperties of the fabricated unit. This did not develop any knowledge ofthe complex stress distribution of the joint between the cylindrical caseand the head end assembly since a heavy weight closure on the chamber wasused. After the final selection of the ablative liner configuration wasmade and sufficient test data was accumulated to be sure of the temperaturedistribution in the titanium shell, it was decided that an appreciableamount of weight could be saved by reducing the case to 0.035 inch thick.Final proof of the system was obtained by means of an elevatedtemperature burst test of a chamber and head end assembly, which weresealed at the throat and the interior volume of the unit reduced by fillingwith a metal plug. The exterior of the case was heated to 800°F withquartz lamps while pressure was applied to the interior of the assemblywith argon gas. The chamber withstood 427 psig, at which time straingages indicated that failure of the wall was imminent. The design requirementof 378 psig (three times maximum chamber pressure of 116 psig)was substantially exceeded.- 9 -

Nozzle ExtensionThe radiation cooled nozzle extension is attached to the combustionchamber case at the area ratio of 16:1 and extends to an exit arearatio of 47.5:1. Columbium alloy C-l03 was selected as the material forthe nozzle extension because of its high temperature structural properties .An aluminide coating provides oxidation resistance and high emissivity .In addition to being able to survive the aerodynamic loads offiring and the vibration environment of launch and boost, a major designcriteria for the nozzle extension is that it be capable of collapsingwithout upsetting the Lunar Module should it strike a protrusion on thesurface of the moon during landing.Figure 12.Aluminum Nozzle ExtensionCrushing Test SpecimenThe program ofdesigning a nozzle extensionwhich would collapse 28 inchesat an impact velocity of10 ft/sec without absorbingan excessive amount of energystarted in 1963 with sometests of aluminum right circularcones . Figure 12 isrepresentative of one of thespecimens . Variations ofconstant thickness, taper andcombinations of the two weretested. The test data ob ­tained provided assurance thatthe analytical methods werecorrect.Having confirmed themethod of analysis, a nozzleextension was designed withHaynes 25 material as a choicesince the temperatures werepredicted to be well below the2400°F melting point of thatmaterial . Subsequent changesto the Lunar Module reducedthe radiation cooling of thenozzle extension and thecorresponding temperature profileprediction exceeded2400°F. As a result Columbium alloy C-l03 was used on the new nozzle extensionsince that material, when coated with the aluminide coating, canwithstand temperatures up to 2700°F. The configuration consists of sheetwelded in segments, stepped down in thickness towards the exit plane. Atthe attachment plane to the combustion chamber, the thickness is 0 . 060 inchin order to have a stiff surface for the seal between components . Threeand one-half inches below that is a nine-inch segment 0.030 inch thick .The next l3-inch segment is 0.020 inch thick and the last 15 inches is0.010 inch thick. At the exit plane a titanium I-shaped ring is attached- 10 -

in order to provide stiffness during handling and launch and boost vibration.Figure 13 shows the temperatures measured during a high altitude test firing.2400f.----- 0.029V WALLDATA ARE MAXIMUM VALUES FROMTEST NO. HATS 060 • AEDC AD-lit -12TEMPERATURE.of200018001----- 0.017 ----iWALL1----0.010 --­WALL1600~OL. ____" ____ ~ __ ~ ____ ~ ____" ____ ~ ____ ~~ m 24 m n H 40AREA RATIOFigure 13.Thicknesses shown in thischart are actual measurementsof that specimen. One of thenozzle extensions which wasused on the qualification testengines is shown in Figure 14after the test was completed.Proof that the nozzleextension would collapse asrequired without impartingforces large enough to upsetthe lunar vehicle during a10 ft per sec descent to thelunar surface was undertakenwith the aid of the testsetup shown in Figure 15.The columbium alloynozzle extension was attachedto a servo controlled hydrauliccylinder which provided energyFigure 14. LMDE Nozzle Extension necessary to collapse thenozzle extension exit flangeagainst the lower platform which was instrumented for measuring load. Surroundingthe test specimen was a bank of quartz lamps which heated up thecolumbium to the temperature simulating the condition at lunar touchdown.The temperature gradient controlled during this test was 2l00°F at the- 11 -

Figure 15.Nozzle Extension in Crushing Test Fixture16:1 joint to l250°F at the exit plane . The 10 ft per sec impact velocitywas accomplished by compressing the nozzle extension 28 inches in less than1/2 second after several inches of free motion. Resultant loads andenergies recorded are shown in Figure 16 .10 15 20 25 30CRUS HIN G DISTANCE - INCHESFigure 16.- 12 -

Figure 17 shows the test specimen after the test. The tests performedearly in the program gave assurance that the analytical methods were correct ,therefore the risk taken by leaving the qualification crushing test tilllate in the program was small.Figure 17 .Columbium Nozzle ExtensionAfter Crushing TestThrottle Control LinkageWe have reviewed the structural design of the larger componentsof the LMDE. Now , we shall look at some of the small details which are ofmajor importance in the function of this engine. The linkage between thethrottle actuator, the flow control valves and the variable area injectordrive mechanism requires high precision. The ratio between the motion ofthe flow control valves and the injector is 5:1. Total motion of theinjector sleeve is 0.15 inch. A ten per cent increment in throttle settingis therefore a motion of only 0.003 inch. In order to maintain the positionrelationship between the injector sleeve and the flow control valves,looseness in the pivot bearings of the linkage had to be virtually eliminated.Another constraint on the design was that the joints of themechanism be capable of motion in a complete vacuum.Where possible, moving elements are connected by flexures. SeeFigure 3. Photographs of these flexures are shown in Figure 18. Tensionand compression loads are transmitted through these flexures while theypermit relative rotation of the connected members. In order to insureagainst a structural failure of these critical components, they were designedwith redundant load paths . Either side of the flexure is capableof full load and life should one side fail.- 13 -

(a)Cross BeamFlexure(b)Injector DriveFlexureFigure 18.ThrottleLinkageFlexuresIn those places where a flexural element was not practical, pivotbearings were used. The first concept was to use a non-metallic bearingfabricated of fiberglass and phenolic. Initial tests showed that the wearrate of the bearing and the compliance of the material exceeded the requirements.An all metal configuration was selected. The extremely simpledesign consists of a steel bushing with a steel pin - both coated withmolybdenum disulfide in an inorganic binder. This configuration also providesredundancy. The primary motion occurs between the pin and the innerdiameter of the bushing. Should this not function, the bushing wouldrotate in the bore of the link of the mechanism. To substantiate thereliability of the design, an assembly of the throttle linkage was subjectedto a cycling test in a vacuum chamber at 8 x 10- 10 Torr. Criteriafor success was completion of 10,000 cycles (33 mission duty cycles) withno significant change in load.Proof that the complete assembly met the requirements for ruggedness,rigidity and repeatability was obtained in the deflection test shownin Figure 19. In this test motions of the two flow control valve pintles,- 14 -

Figure 19.Deflection Test ofThrottle Linkagethrottle actuator and the injectorsleeve were observed and recordedunder the influence of variousfluid pressures and chamberpressures.Bellows AssembliesBoth the flow controlvalves and the injector driveassembly contain a moving elementused to control propellant flowrates while containing the propellantunder pressure . Bellows areused as dynamic seals, permittingdirect external actuation of internalmetering surfaces . The configurationselected to accomplishthis was a thin wall (0.006 to0.010 inch) stainless steel weldedbellows. Figure 20 shows thelocation of the bellows in theflow control valve. A similarapplication of bellows is used inthe injector drive . In 1963 atest program to verify the str engthand life of these bellows was instituted.Ten thousand cycles (33mission duty cycles) was selectedas a minimum life requirement . The original configuration did not survivethe tests. Fatigue cracks occurred in the single ply convolution at theweld joints. The final configuration (Figure 21) a two-ply design with0 . 004 inch convolutions successfully passed all of the tests which consistedPINTLEBEARINGINLETDIFFUSERFLOW SPLITTERVENTURI THROATFigure 20.LMDE Flow Control Valve- 15 -

Figure 21.Figure 22.Section of Two-PlyBellowsComponent Test ofInlet Lines , FlowControl Valve andShutoff Valveof cycles of surge pressure, andextension and compression cyclesunder static pressure, followedby burst tests in excess of tentimes the design pressure.Here again, early testingof a component revealed a designdeficiency which was correctedwithout affecting the engine developmentschedule. The final configuration(two-ply bellows) hadan additional advantage of havinga lower spring rate than the singleply bellows originally proposed.Testing of the bellowscontinued when they were combinedinto assemblies. A componentdesign verification test of theflow control valve included longterm exposure to propellants, shockand vibration tests representinglaunch aud boost and engine firing,and 33 mission duty cycles of pressurizationand actuation. Uponcompletion of these exposures, thecomponents were required to passtheir acceptance criteria.Figure 22 shows the arrangementfor testing the flow control valve,shutoff valve and inlet line asa subassembly.Engine TestsAs the next step in thetest program, an assembly of thehead end - composed of injectormanifold, injector drive, ducting,shutoff valves, flow control valves,throttle actuator, and throttlelinkage - was subjected to the completequalification level of vibrationenvironment (Figure 23). Inorder to establish the margin ofstrength built into the head endassembly, upon completion of acceptancetests following the qualificationtests, it was tested to 125% ofthe qual levels, then again to 150%in an attempt to find out whenfailure would occur. Finallyfailure did occur at 175% of- 16 -

Figure 23.Figure 24.Vibration Test ofHead End AssemblyLMDE Installed in HighAltitude Test Standqualification vibration tests ofthe complete engine.At this point, only thecombustion chamber and nozzle extensionhad not been subjected to thelaunch and boost vibration environment.They had however receivedmany thousands of seconds ofreactive firing during the developmenttests of the engine in theTRW Sea Level Vertical Engine TestStands (VETS) and High AltitudeTest Stand (HATS). As a lastmeasure before the start of thequalification series of tests, adevelopment engine was vibrationtested to the qualification environmentand then successfullyfired in HATS for a full missionduty cycle. The engine shown inFigure 24 installed in the HATSfacility is awaiting final checkoutprior to mission duty cyclefiring. The support structureto which the engine is attachedmeasures thrust and six componentsof thrust to determinethrust alignment.Analysis and Test for Dynamic LoadsA linear spring-massanalogy for the LM Descent Enginewas developed as a design tool tocompute the dynamic loads on theengine structure . Initially,this analysis consisted of asingle mass, five degree-offreedomsystem. Two componentmasses were added to the initialanalysis, increasing the number ofdegrees of freedom to eleven.Hand computation of the dynamicloads became very time consumingso the response equations were programedfor the digital computer.The first vibration testof the gimbal, described earlier,indicated that a strong enginetorsional mode existed (enginerotation about the thrust axis).The corresponding rotationalcoordinate,e, had been leftx- 17 -

out of the first two analyses on the assumption that the torsional mode wouldnot be strongly excited. Based on this experience, a third and final springmassmodel was developed. The primary purpose of this analysis was to includethe torsional coordinate. In addition, the gimbal was represented astwo support masses with a system of springs connecting them to the main enginemass. Each side of the gimbal is connected to the Lunar Module with a fourmembertruss . Modeling the gimbal in the above manner facilitated representationof the support trusses as a system of springs .The final model is basically a three mass, twelve degree-of-freedomsystem (Figure 25) . The motion of the main engine mass is described by sixcoordinates, three displacements and three rotations, in a rectangularcoordinate system with the origin at the center of the gimbal. The twogimbal/support mass motions are described by three displacement coordinateseach.'3FINAL SPRING - MASS MODEL OF ENGINETHRUST ENDSUPPORT MASS, '"3"SUPPORT MASS , m 2HEAO ENDKG,,' K ' K - GIMBAL SPRINGSCy GzK, - GIMBAL TORSIONAL SPRINGK,,2' K y2' 1

Table 1 presents the qualification sinusoidal and random vibrationtest levels applied to the attachment interface on the gimbal frame.The configuration of the final engine vibration test specimen is shown inFigure 26. It was necessary to mount the engine upside- down in order tokeep the mass of the supporting fixture low, since the mass of the fixtureplus the engine approached the maximum limits of the shaker . The supportingstructure is the same as used for the gimbal test shown in Figures 6and 7.TABLE 1VIBRATION LEVELS FOR QUALIFICATIONTEST (LAUNCH AND BOOST)SINUSOIDAL VIBRATIONFlightEnvironmentFrequencyInput LevelSweep RateLaunch andBoost5-1616-1000.2 inches d.a.2.5 g vector3 oct/min upand downRANDOM VIBRATIONPhaseFrequency Band(cps)Power SpectralDensityTime perTest AxisLaunch 10-23 12 db/octave rise23-80 0.025 g2/cpsand 80-100 12 db/~ctave rise100-1000 0 . 06 g /cpsBoost 1000-12000 12 db/octave fall1200-2000 0.025 g2/cps5 minutesof randomvibrationTable 2 presents a comparison of the calculated and measured frequenciesand loads. It is seen that the final mathematical model permitsaccurate estimation of the dynamic characteristics of the engine.The analytical model has been used primarily to determine theeffects of design changes . For example, following vibration of the gimbalwith a mass simulated engine during design verification testing, the stiffnessesof the support trusses that mount the engine in the Lunar Modulewere decreased to save weight. The mathematical model was used to determinethe effects of the softer mounting system on engine dynamic response. Othersimilar analyses include the effects of changing the gimbal actuator stiffness,effects of changing the engine c . g . location, and an evaluation of thevibration test fixture.In addition, the dynamic loads on the gimbal nozzle extension andthrottle control system have been computed for use in associated stressanalyses of these components. The loads computed for the throttle control- 19 -

Figure 26.Vibration Test of LMDEsystem, verified by the results of the engine mass simulator vibration testresults, allowed simplification of the structure mounting the throttlecontrol system on the engine.TABLE 2COMPARISON OF ANALYTICAL WITH TESTDYNAMIC RESPONSE OF ENGINEANALYSISMODALFREQUENCIES(cps)(All Axes)11.212.627.533.137.9QUAL LEVELTEST RESONANTFREQUENCIES(cps)AxisX Y ZMAXIMUM GIMBAL LOADSLOAD TEST ANALYSIS11.4 11.6Mx12.7 12.728170 in-lbs 31330 in-lbs28.4 29.2 1:'y 5580 lbs 5220 lbs30.8 34.1 34.3 Px 7105 lbs 7520 lbs39.0 38.7 Pz 6723 lbs 6275 lbs- 20 -

SummaryThus at the start of the engine qualification test series, everycomponent of the TRW Lunar Module Descent Engine had been subjected to tests -step by step - first as a unit, then in its next assembly, then as a majorsubassembly, and again as a complete engine. Simultaneous with all of themechanical and structural testing described, reactive testing was being performedin the TRW facilities at San Juan Capistrano and at the USAF ArnoldEngineering Development Center at Tullahoma, Tennessee, and at NASA's 1Mvehicle test facility at White Sands, New Mexico. Table 3 presents thedetail of this test experience to 1 July 1967.TABLE 3LMDE REACTIVE TEST SUMMARYNo. ofNewBuildsNo. ofStartsFiringDuration(seconds)Injector Tests1,80070,374Throttling Head EndAssemblies on WaterCooled CombustionChambers2894161,666Ablative EnginesSea Level (E = 2:1)21675,368High Altitude (E = 47 . 5:1)3524226,549Total Test Experience3,050163,957The successful completion of qualification of the TRW LMDE demonstratesthat, in order to develop light weight and reliable high performancerocket engines in a short time span, it is necessary to substantiate all ofthe component design concepts early in the program. A carefully designedcomponent development test program verifies the assumptions made duringdesign analysis and provides assurance that tests of the complete systemwill proceed without unexpected difficulty. The engine shown in Figure 28being prepared for shipment to Grumman Aircraft Engineering Corporation,- 21 -

designers and builders of the Lunar Module, is one of several being testedas part of the complete propulsion system and space vehicle in order toassure success for the first Lunar Module flight later this year.Figure 27.LMDE Being Prepared for Shipment to theGrumman Aircraft Engineering Corporation- 22 -