ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
harness,andairtubingmustbemeasured foreachhingelinebeforeindividualflightdamper characterization canbeperformed. Release &Spacecraft Interface Awingassembly (minus theyoke)attaches tothespacecraft at4points.Toallowforthedifference in thermal expansion coefficient between thealuminum spacecraft andtheGrCEarray,twosupport points, showninFigure4,useaball&socketanda'V'groove inaflexuremounting system toreactshearbut allowexpansion alongtheaxisofthe'V'groove. Thetworelease pointsofthemounting system interface between panelsthrough flatsthattakenoshear.Asillustrated inFigure5,aholethrough eachpanelat eachflatallowsan18.4-cm longBeCubolttotiethespacecraft tooneoftwoHi-Shear 9321lowshock seriesseparation nutsattached totheoutboard panel.Theboltisstrain-gaged topermitanaccurate measurement oftheboltpreload, anditspringloadedtopartially retractafterrelease. Asmallgaptothe spacecraft isprovided ateachrelease boltinterface toobtaina'pop'release usingstowed strainenergy frombending ofthepanelassembly. Figure4_.__..__.._-;_ .........:._.:.::.-; _:5:'!,'_> Figure5 Mechanical Ground Support Equipment TheAPLmechanical systems grouphasaverystrongphilosophy datingtothe1960'serathat"youfly whatyoutest".Thedesignteamwascertainfromthebeginning thatduetothesizeofthisarrayandthe limitedintegration timeavailable, thatwewouldnotbeabletodeployandtestthearrayonthespacecraft. Toreduce risk,theteaminsisted earlyonthatahighfidelity, full-scale engineering modelbefabricated andtestedpriortospacecraft levelCDR.However, toperform theengineering modelandflighttesting, unique deployment/G-negation ground support equipment (GSE)wouldhavetobedesigned. Itwas obvious thattheavailable manpower wouldnotallowustodevelop theGSEfromscratch andmaintain thetightGSEdevelopment schedule. OurNASAsponsor, Goddard Spaceflight Center (GSFC), offered designdrawings oftheirairbearing deployment system usedontheTRIMM andXTEmissions. Aftera quickreviewofthedrawings, someadditional designchanges weremadeandwesetouttofabricate the system. Inaddition, GSFCwasveryfamiliar withtheDEBManufacturing damper design chosen for TIMED andgraciously offered theirdamper characterization thermal chamber forouruse.Thehelp provided byGSFCjumpstartedourGSEdesigneffortandwasinstrumental inmaintaining schedule. Deployment TableSystem Thedeployment tablesystem showninFigure6consists ofaseriesofaluminum honeycomb panels mounted onanaluminum support frame.Thesetableswerejoinedtogether toformasmooth airbearing 220
compatible surfaceflattowithin0.38mmoverall.Thesetableswereapproximately 1.2mx2.4m(LxW) x0.6minheight. Figure6 Figure7 Theframeconsisted ofwelded tubular aluminum 6061alloyrectangular sections 10x15cmwitha0.64 cmwall.Thewelded framewassupported offthecleanroomfloorwithafour-point machine-tool typeof leveling foot.Thealuminum honeycomb panel(deployment surface) interfaced totheframevia10 adjustment screws. Theconcept wasthatbyemploying theadjustment screws, localleveling ofthe honeycomb surface couldbeachieved regardless ofthehoneycomb panelflatness. Thisuniquefeature proved quitevaluable lateronwhenwereceived "lessthanoptimum" honeycomb panels. Theschedule identified theneedtoquickly manufacture thetablesystem. Apanelvendorwaslocated viatheThomas Register whocoulddeliver withinaweek.Thevendor specialized inflooring systems forNavyshipsand wasnotconnected withtheaerospace industry. Thepanelsprocured were2.5-cm thickwith0.38-mm thickfacesheetsandcost$340each.Theentiretabledesignanddrafting process tookonemanweek withprocurement ofallcomponents aboutthreeweeks. Finding thepropersupport pointsfortheairbearing systemwasprobably themostchallenging aspectof thedeployment GSEsetup, Stability duringdeployment hadtobetraded withthecrowding ofthe supports whenstowed. Threedifferent airbearing designs werefabricated, mono-pods (1airpuck),bipodsandtri-pods, insufficient quantities tosupport multiple combinations. Thefinalcombination and location ofthesupports wasdetermined empirically onthefull-scale engineering modelandcarried over ontotheflightarray,Figure7presents thebJ-podversion oftheaJrbearing, andFigures 13and14show theairbearing arrangement chosen alongthelengthofthearray. Structural Dynamics, StressandKinematic Analysis Kinematic analysis wasperformed usingthe2Dversion 4.01ofWorking Modelsoftware byKnowledge Revolution, Inc.Choice ofdamper capacity, springsizeandpreload depended squarely ontheresultsof thedesignanalysis performed usingWorking Model. 221
TIMED Solar Array Engineering Model Trial1 &Trial2 vsWorking Model Predictions Angular Position vsTime TI-T2=tl 29O =¢ 0._ 1_3 20 3O 40 50 6O _ -200 Elap_d Time, (_c)fl-¥_e _,T1.blr_ge1-2 Tl-blr_e2.3 TI-_noe3-4 oT2-¥oke ,:;T2-1_gel.2 T2.I-mge2-3 •T2-Hnge3-4 _&ldlnge 2-3 Figure8 _zi::l Figure9 Lockup loadswerederived fromthevelocities uncovered fromthedeployment analysis. Theengineering modeldeployment testresultswereusedtoverifytherigidbodykinematic model. Figure8givesthe deployment anglevs.timecurves generated foreachjointforboththeengineering modeltestand analysis model.Agreement isquitegoodgiventheaccuracy thatjointfrictionwasknownduringgroundbaseddeployment. Thecurves with*andsquare iconsalsoillustrate theabilityoftheanalysis to diagnose GSEproblems, inthiscaseanuneven floorsection thatslowed thedeployment ofhinges1-2 and3-4.Thediamond iconcurvesshowaslowerdeployment duringtestthanpredicted byanalysis for theyokedampers. Further investigation uncovered thattheyokedampers, whichwereloaded witha moreviscous silicone fluidthanthepaneldampers, required amoresensitive adjustment procedure than 222
ahandadjustment couldaccomplish. Forflight,theyokedamperviscousdampingcoefficient was adjustedusingtheGSFCthermalchamberforconfirmation. StructuralanddynamicanalysiswasperformedusingMSCNASTRAN finiteelementsoftwareforboth thestowedanddeployedconfigurations. Quasi-static loadcaseswererunconsistingof+15G'sinthrust (X),+15G'snormaltothepanels(Z),and+7.5G'sin(Y),alongwithacasethatenforced a2.1-mm preload (gap)attherelease bolts.Strength margins ofsafetywereallpositive forthefivestowed load casesrun.Figure9showsatypicalstressdistribution pattern ofthestowed arrayunderquasi-static loading. AsshowninFigure10,thewingdeployed firstbending natural frequency is0.26Hz.Whenstowed, the wingfirstnatural frequency is39Hz.Fixedbasesinetestingofeacharraystackreturned afundamental frequency of37.9Hz,verifying thedynamic FEanalysis. Figure10 Engineering ModelTestProgram Theengineering modeltestprogram wasthemostsignificant effortundertaken inthedevelopment ofthe arraysystem. Theplangoinginwastocomplete allfullscaleengineering modeltesting priorto spacecraft CDR.Theengineering modelwasdesigned tobefullscaleandofahighfidelitywiththe following requirements imposed: A.Panels/substrates: Physical sizeandmassidentical toflight.Insertsincluded forhingesandrelease mechanisms B.WiringHarness: Allwiringservice loopsweretobeflightlikeinsizeandcomposition. C.HingesandRelease Mechanisms: Flight-like aswell. D.Viscous Dampers: Flightdesignwithout fullqualification testing. ThesingleDEBManufacturing, Inc.Model1025viscous damper isflexmounted totheshaftbetween panels andprovides velocity control overtheentire1800oftravel.Extensive testingwasperformed to characterize typicaldamper performance undervarious temperature andloadconditions. Thetestsused acustom-built thermal chamber borrowed fromGSFC. 223
FlightArrayTestProgram Toaccurately predictthebehavior ofaparticular damper undertemperature andvarious needle valve settings, thenettorque thateachdamper seesmustbeaccurately measured. Therefore, thespring torqueanddragtorques frompinfriction, harness, andground support airtubingmustbemeasured for eachflighthingelinebefore individual flightdamper characterization canbeperformed. Sincethe dampers' damping coefficient variesgreatlywithtemperature, forflight,redundant heaters aremounted oneachdamper tocontrolthehousing temperature withintherangeof+15°Cto+20°C. Duetoaverylatedelivery ofthefullypopulated flightsubstrates fromthevendor, wewerelimitedtoonly twoweekstoassemble, deployment test,vibrate, deployment testagainandintegrate eachwingintothe spacecraft forspacecraft- levelenvironmental testing! Thelessons learned ontheengineering model proved invaluable andmorethanpaidfortheinvestment duringthisdifficult period. Thefabrication tooling, dataacquisition system andthedeployment system performed flawlessly. Sinevibration testingof eachwingwentsmoothly withnogapping observed attherelease bolts,verifying thepreload inthestrain boltwascorrect. Post-vibration deployment testing, showninFigures 13and14,notednochanges inthe damper settings orindeployment behavior. Theyokewasintegrated firsttoallowforafull180°ofrotation oftheSADA, verifying proper harness looprouting andtiedown. Integration ofeachwingwiththe spacecraft wentwellwithout anymodification ofinterface components required. Thefollowing testmatrix detailstheentiretestprogram implemented fortheengineering modelandflightsolararrayhardware. Table2 t........ TIMED S_acecraft  ...........i.................................................... Mo_i_i.................... T Oncethearrayswereintegrated, thespacecraft underwent sinevibration testing atJHU/APL, and acoustics, massproperties, andself-induced shocktesting atGSFC. Aspartoftheshocktest,each wing'sseparation nutpyrotechnics werefired,andeachtethered wingwasallowed toswingafew degrees fromthespacecraft. This'pop&catch'testdemonstrates thatsufficient clearances exist between thearrayandlastminuteadditions duringtheintegration process suchasthermal blankets and harness. Conclusions andLessons Learned 1. Aswehavelearned onotherJHU/APL mechanism designs aswell,anengineering modelisan incredible resource, provided itsimulates allimportant aspects oftheflightdesign suchasfunction, behavior, lubrication andsurface finish,GSEhandling fixtures, tooling, etc.Wewereabletodevelop the actualflightassembly andhandling processes andprocedures ontheengineering model allowing the flightarraystobeassembled, deployment tested, environmentally tested,anddelivered tothespacecraft intwoweeks! 225
, . . .Figure13 Figure 14 Encourage outside reviewers. Fresheyesprovided valuable design suggestions duringthearray EDR.TheGSFCreviewers remained helpful contacts, sharing experiences onsimilardesigns and providing insightintodesignintent(notalways reflected) ontheairbearing GSEdrawings thatwere provided byGSFC. Bewillingtotossoutareally"bad"ideaevenlateinthegamewhererisktotheflighthardware is possible. Unfortunately theoneassembly wecouldnottestoutwasthewingintegration fixture. Itwas anoffsetCG-compensating fixtureusedsolelyforflightintegration onthespacecraft, andthuscould notbeaccurately testedusingtheengineering modelarray.Whenwetriedtouseitforflight integration, thefixturewastoomassive tosafelyhandlethemuchlighterweightflightwing.Wethen substituted 4menforonefixture,greatlyreducing therisk. Establish goodpersonal relationships withpeersinotherorganizations. Wedeveloped good relationships withsuppliers andGSFCemployees thusallowing usto"cutthrough theredtape"when quickresponses wereneeded toproblems. Design andimplementation ofdeployment GSEcanbeasmuchofaneffortastheflightarraydesign. Assuch,cost,schedule andmanpower resources mustbeallocated insufficient quantities orthetest program canbejeopardized. 226
.s; , / Validation ofTSX-5SolarArrayDeployment byCorrelated Simulation andTest BrianW.Gore" Abstract Aninnovative spacecraft hingemechanism hasbeenanalyzed andflightqualified byauniquemethod that joinsvarious DADSsimulations andcomplementary labtesting, sinceTheTSX-5solararraysubsystem is incapable ofbeingdeployed intheEarth's gravitational field.Results fromsub-scale testing and corresponding validated models exposed detrimental deployment effects inafull-scale simulation that wouldnothaveotherwise beendetected usingstandard aerospace qualification techniques. Those deficiencies werefixed,andthespacecraft solararraysrecently deployed successfully on-orbit. The methodology usedandthesubsequent resultsarepresented andexplained indetail. Introduction TheTri-Service Ex_periment-5 (TSX-5) spacecraft isaU.S.AirForceprogram thatprovides aplatform for twotechnology experiments tobelaunched onalowercost,higherrisk,andfasterschedule basis.The onlydeployable onboardisanon-articulated solararray,consisting oftwowings,eachcontaining six solarpanels(seeFigure1). Figure1.TSX-5Spacecraft (Courtesy ofOrbital Corp.) Sincecomplete mission failuremayresultifthearrayfailstodeploy, itisofparamount importance thatthe system deploys successfully. Inordertoensure thatsuccess, thestandard procedure istosubject sucha moving mechanical assembly torigorous testing inalaboratory environment sothatitskinematic and dynamic behavior arewell-understood andcomply withallrequirements. However, theinnovative solar arrayhingedesign usedbyOrbitalCorporation onTSX-5cannotsupport thegravitational weightofthe solararray.Thisdesignissatisfactory whenthereisnogravity, butitposesaparticularly complex problem forengineers whowishtotestitsbehavior ontheground. Moreover, thisproblem canbereadilysolvedinmostspacecraft deployment systems sincetheyusually consist ofasingle,oreventwoparallel, hingelines.Insuchcases,various methods ofgravityoff-loading aregenerally employed. Inthisuniquecase,eachofthewings(whichdeployindividually) possesses multiple andmoving hingelinesaboutwhichdeployment occurssimultaneously, thusrulingoutstandard off-loading fixtures. Assuch,engineers werefacedwiththreepossibilities: 1)deploythesystem inasimulated zero-gravity field°-theKC-135 aircraft canprovide this duringparabolic flightpaths TheAerospace Corporation, ElSegundo, CA Proceedings ofthe35thAerospace Mechanisms Symposium, AmesResearch Center,May9-11,2001 227
2)employagravityoff-loadingfixtureusingballoons,cables,etc. 3)useacombination ofanalyticalandsub-scaletestingtechniquestoevaluatethevalidityofthe deployment system Thefirstchoicewascompletely cost-prohibitive giventhelowbudgetoftheprogram.Thesecondchoice wasattemptedbutwithquestionable resultssincetheballoonshadasignificantaidingorhinderingeffect onindividualpanels.Hence,thelatterwaschosenasthecourseofaction.Thispapershalldiscussthe testingandanalyticalprocedureaswellasthelessonslearnedthroughouttheevolutionofthisevaluation process. Threeorganizations tookpartinthistesting: Orbital Corporation (TSX-5 primecontractor), Planetary Systems Corporation (TSX-5 solararrayexpert), andTheAerospace Corporation (customer technical representative. Background Inordertounderstand theissuesinvolved withthisdeployment system, itisfirstnecessary tounderstand thedesignofthehinge,sinceitservesasthestructural linkbetween panels, aswellasthemotiveforceto deploythem.Figure2depicts thehingeofinterest, whichwasinvented byWalterHolemans andpatented byOrbitalCorporation (Patent #US05715573). Figure2.Hinge,Stowed &Deployed Onecanseethatthelink(thediamond-shaped piecebetween thepanels) hasaplunger attached oneach sideviaarevolute joint.Ontheotherendofeachplunger, whichresides withintheinteriorvolume ofits respective panel,isaseatagainst whichthedeployment springs actuate. Whenthecompressed spring releases itsenergy, theplunger ispulledintothepanel,andthelinkrotates andslidesalongthepanel edgeandupintoitsseatedposition, alsoinsideeachpanel.Oneofthemostimportant features tonote aboutthishingeisthatithasvirtually nocapability tosupport shearloading whennotlatched, thusgravity offloading becomes amajornecessity. 1.I1<-q .... [.i Figure3.SolarArrayDeployment Sequence (Courtesy ofOrbital Corp.) 228
Onalargerlevel,Figure3showshowthehingesacttomovethesolararraysfromtheirstowed configuration totheirdeployedconfiguration. Twoadditionalpointsshouldbeidentifiedatthistime:1)duringdeployment, thehingesallowacertain amountofover-travel ofthepanelsbeforesettlinginthedeployedconfiguration, and2)thedeployed configuration ofthetwowingsisnotplanar(seeFigure1).Attheoutsetoftheprogram,itwasarguedthat awingdeployment staggeringmechanism shouldbeemployedsothatthetwowingswouldnothiteach other.Inaddition,itwassuggestedthatenergydissipatingsoftstopsbeaddedtopreventdamageshould eitherwinghitanypartofthespacecraftbottomcoreplate.Oncethetechnicalexpertsembracedthese ideas,severalmajorquestionsremained: 1.Howlongshouldthedeployment staggeringdevicewaittoreleasethesecondwing? 2.Does2n_wingdeployment causea"tail-wagging-the-dog" disturbance sufficienttounseatthe 1stdeployedwing? 3.Doesover-travelofeitherwinginduceexcessivebendingstressinthepanelsorspacecraft coreplate? 4.Willanyotherpanelover-travel(withinaparticularwing)causedetrimental contactofthe deployingsolararray? 5.Docoreplatesoftstopsnecessarily preventdetrimental over-travel ofthedeploying solar wing? Notethatthereareatotalof12panelsandhingelines,veryfewofwhichareparallel duringdeployment. Assuch,itisnearlyimpossible toeffectively offloadtheeffectsofgravityduringawingdeployment test. Thissamesubject wasmentioned inReference 1(butwithhighlyquestionable ground testresults), and thiseffortstrovetofindasolution. Itwasclearthatavalidated modelwouldbenecessary tosatisfactorily address alloftheseissues,thusthesoleremaining optiontoqualify thisdeployment scheme design wasacorrelated analysis andtest. Modeling andTesting Figure4illustrates theoverallprocedural flowthatwillbediscussed inthissection; notethethreedifferent boxstylesforeachstepoftesting, output, andmodeling, andthefactthatthisisahighlyintegrated methodology. Testandcharacterizationofsamplehinge (test) (output) I(model) Testof2-panel systerr Figure4.Integrated Correlation Methodology 229
Thefirststepwastoperformaseriesoftestsonasinglehingethatwasconnectedtotwosmallcoupons ofhoneycomb panelsubstrates. Thistestwasdonetocharacterize thespringproperties, friction,and over-travel tendencies ofthehingedesign.ThosepropertieswerethenusedtocorrelateaDADS (DynamicAnalysisandDesignSystem;LMS-CADSI, Coralville,IA)modelofthatsamesample hinge setup. Themodelcontained averyhighleveloffidelitybyincluding notonlythespring, massproperties ofthe panelcoupons, andairdragresistance forcesonthem,butalsothegeometric rolling/sliding contact behavior between thelinkandthehingehousing. Withthefrictioncoefficient between thelinkandthe panelhousing beingtheonlyrealunknown, itwasvariedwithinnormal expected ranges untiltheovertravelshowninthemodelwasverynearwhatwasfoundinthesample hingetest. Atthattime,theneedtoincorporate awingstaggering mechanism intothespacecraft design was beginning tobecome clear.Withthelightsample substrate, anover-travel angleof20-30degrees was observed intest.Asstatedabove,withthetwowingspossessing anon-planar deployed configuration, this levelofover-travel wouldcertainly produce animpact between thewingsduring asimultaneous deployment. Thisisahighlyundesirable condition giventhefragilenatureofthesolarcellsmounted tothe panels. Another conclusion drawnatthispointintheprocedure wasthatasimplified hingemodelwouldbe necessary inordertocreatetheeventual goal-acorrelated spacecraft model.Duetothehighlevelof calculation necessary duringthemodelsolving process (primarily duetotherolling/sliding contact elements inthehingemodel), itbecame evident thatitwouldbeunfeasible toduplicate thissetupinthe 3._22necessary locations inaspacecraft model. Therun-time forthissample hingemodelwas approximately 40minutes for1.25seconds ofsimulation ofonehinge.Thespacecraft modelwas expected totakeapproximately 60seconds andcontain 24-32hinges, andtherelationship isnon-linear whencombining suchsetsofelements inacomprehensive model. Hence, itwasdecided thatthefull spacecraft modelwouldneedtohavesimplified revolute joint/spring hinges, whilestillcapturing the correctbehavior oftheseinnovative devices. Thenextstepintheprocess wasprobably thekeytotheentireeffort-performing ameaningful sub-scale testtobeusedasaplatform fromwhichtheentirespacecraft modelwouldbeextrapolated. Figure5 illustrates thesetupusedforthis2-panel test.Agrounded-body spacecraft mockup wasused,alongwith twodummy panelswhichhadlargeholescutouttominimize theeffects ofairresistance. Thecore/root hingelinehadfourhingesacrossit,similartothecorresponding location ontheflightunit;theroot/outer hingelinehadtwohingesacrossit,simulating themostoutboard hingelinesonthespacecraft. Wiring harnesses similartothoseusedinflightwereroutedacrossbothhingelines. Core/Root HingelineRoot/Outer Hingeline RootPanel Potentiometer (tyIj,. OuterPanel ._ Figure5.2-Panel TestSetupSchematic Theentireassembly wassupported usingairbearings onagranitetable-recallthatthesehingeshave nocapability toprovide lateralstiffness (vertical direction here)untiltheyarelatched. Theairbearing 230
paddleshadthecapabilitytosupportvarioussteelweightssuchthatthemasspropertiesofthewingcould bevaried.Thereasonforthataccommodation wasanunresolved discussionregardingthedistribution of wingweight.Oneargumentwasthatthewingweightshouldbedividedequallyamongtheinboardand outboardpanels(properdistribution relativetothecore/roothinge).Theothersaidtheweightofonepanel shouldbelefttotheouterpanelandtheremainingfivepanelweightslumpedattheinboardpanel(proper distribution relativetotheroot/outerhinge).Thecompromise wastotestbothconfigurations andattempt tocorrelatebothwiththesamesetofsimulatedhingeparameters. Inaddition, apotentiometer measurement system wasusedtogather thetimehistory dataofthe deploying panels. Thisdatawasusedtorefineandcorrelate theensuing DADSmodel. Thatmodel included allofthecorrect massproperties andgeometric features ofthe2-panel testsetup,including contact/interference between thepanelsandthespacecraft andcontact/stiffness between thedeployed panels andthecoreplate.Thesimplified hinges nowconsisted ofbasicrevolute joints,drivenby Rotational Spring/Damper/Actuator (RSDA) elements. Torque-vs-angle datawasusedtofine-tune thenon-linear springcharacteristics ofthehingeforthe RSDA's, aswellasthefriction andcableharness resistance inputs.Thethirdenergy dissipation factor usedwasa"buckling factor." Whenthesehinges passthrough theirdeployed position, thereissome energy lossduring thattransition, verysimilar tothebehavior exhibited bycarpenter tapehinge deployments. Figure6illustrates thereversible stepfunction thatwasassumed forthisfactor,whichwas alsoassumed toactonlywithin+10degrees ofthedeployed position basedonobserved hingegeometry. Themagnitude oftheordinate (torque) wasessentially foundusingatrialanderrorapproach untilit resulted inaclosematchofthebehavior forthebothoftheaforementioned massproperties cases. i I::°m EO I-- Angle(rad) Figure6."Buckling Factor" Resistance StepFunction Thereweremanymoresensitivities inthesystem thaninitiallyassumed. Thedynamics andtimingofone paneldefinitely hadnon-linear effectsonthedeployment behavior oftheother.Assuch,extremely good correlation couldbeachieved foreitherofthechosen representative cases(Tests17and18),butthebest overall correlation isshown inFigures 7and8forbothrespective casessimultaneously. Thiswas achieved bysplitting thedifference forsomeoftheparameters sothattheerrorsinbothcaseswere minimized simultaneously. 231
180 150 9O Angle(rad) 6o 3o 30Core/Root HingeCorrelation (Green=Test 17;Red=Test 18)(Dash=Test; Solid-DADS) =- _-:"............. 1 r _........ ,........ .... :......" '/T__ct17I ' 'I .-. _ .... E I ....... "r........... -".._q"_-".,............................ T........ 1........  / _="-, I t I /_-_%. , , ,................. /'_.........-_r......... _-................. -_........ -_........ ,. r .,_. ,,.- , , Test18| '\i_:_:","" "'. ' ; ................... i................................................................... ,............................. :,"_:__::.=_*_"::_'__ ................ _:_-:_'_ ' ; ; i iI i i II _. 4 F, 14 111 1_ 14 Ih Time(sec) Figure7.Core/Root HingeTestandAnalysis Correlation 15D lO.D 9D Angle(rad) ftll 3D -3Oi i i i i''i I,I ,,I I I ILl 2 4 r_ I_ ]El 12 ]d 10 Time(sec) Figure8.Root/Outer HingeTestandAnalysis Correlation Onecanseethattheovershoots matchwellbetween thetestsandthemodel.Also,allofthesignificant timingofevents iswithin0.5-1.0 second orbetterbetween computer prediction andtest.Afterreviewof theseresults, themodelwasdeemed byallparties tohavesufficient correlation toproceed tothe extrapolation oftheentirespacecraft configuration. 232