Mechanical Design of the LMDE.pdf - Heroicrelics

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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 -
Page 1 and 2: MECHANICALDESIGNDFTHE LUNARMODULEDE
Page 3 and 4: Engine DescriptionCOMffNSAT1NGSPIti
Page 5 and 6: Engine GimbalThe gimbal assembly is
Page 7 and 8: Figure 6.Engine Mass SimulatorMount
Page 9: Figure 10.Combustion Chamber Assemb
Page 13 and 14: Figure 15.Nozzle Extension in Crush
Page 15 and 16: (a)Cross BeamFlexure(b)Injector Dri
Page 17 and 18: Figure 21.Figure 22.Section of Two-
Page 19 and 20: out of the first two analyses on th
Page 21 and 22: Figure 26.Vibration Test of LMDEsys
Page 23: designers and builders of the Lunar

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