ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Adedicatedsynchronization system[3]guarantees thatthedeployingpanelsstaywithinacertain deployment trajectory.Thetorquerodsofthissystemarelocatedonthebackofthedeployedpanels.In stowedpositiontheyfitwithinthepanelspacing.Thedeployment phasesofatypicalTeledesictypearray aregiveninFigure2. i_ _[owed:ond_ion :'2 H,_d;downrelease <J iiii_ :_i::;¸_¸¸¸¸_:¸ Lii!ilii,'i_,,_iiiI_I;_..... i L 4 PDM mrY_or_ _3' ...... i!iI _iii;;¸¸¸_:....compteter_ .... i:ideployment _UbWlmg_ 110  Ilki: i i__ _]J i'T_ _iii¸ i Figure2."Curwin ©''deployment (typical foraTeledesic design case) 254
Detailed ArrayDescription Thesolararrayconsists ofthefollowing majorelements: 1.Primary Deployment Mechanism (PDM) 2.Wingsolarpanelstructure withPhotoVoltaicAssembly (PVA) 3.WingDeployment mechanisms 4.Holddownandrelease mechanism Thefollowing sections provide somedetailsofthewingconcept design. Emphasis hasbeenputonthe designofthewinghardware, asitcontains ahighdegree ofdesigncommonality anddesignmodularity withbothprograms, beingtheitems2,3and4oftheabovelist. Primary Deployment Mechanism (PDM) Thepurpose oftheprimary deployment mechanism istocreatesufficient clearance between therotating wingandthebaseline spacecraft structure. Inaddition, itensures minimisation ofshadow linesandalso provides adequate clearance fromtheexhaust plumes oftheS/Cthrustmotors. Theprimary deployment motionisquitedifferent perconcept. Still,modular elements suchasthelimbsandtheindividual hinges areforeseen. Atypicaldesignconcept isshowninFigure3.ThePDMalsoneedstoberestrained during launchby2oftheAdvanced RigidArrays(ARA)-type holddowns, asitisconsidered arelative heavy structure mechanism. • <_:_r_..:.... _ PDMarm  I I'-_'":__.... IIntermediate hingeI  Upperh,nge ._.<,_iiiii;-:-.,i.-;:: _I/ I "" -> ,_.".--;:_-:;" 04 Pressure plate  ..... ._".",z Damper assyM/RMholddown _ _i. '_:i.. •"-. Snubber frame Figure3.Solararrayinstowed configuration withPDM;(Celestri typical) 255
TheWinqPanelPackaqe: Thewinqsolarpanelpacka.qes structure (2-off) Eachwingpackage consists of: Fourteen (14)thinsemi-rigid CFRPpanels equipped withthinfilmtypesolarcells electrical interpanelpowertransfer harness Pressure plate Snubber-frame interfacing withtheS/Cincluding thejackcurving mechanism 2x6primary ARA-type hold-down points 2secondary modified ARA-type hold-down points Thinsemi-riqid CFRPsolarpanelswithPVA Thewingsconsist ofanumber ofrelatively thinsandwich panels, measuring 4.3mx0.50m.The sandwich isAluminum honeycomb withCFRPfacesheetsonbothsides.Thebaseline panelthickness is 4mm.Thephoto-voltaic assembly isbasedontheapplication oflargethinfilmcellssuchasamorphous SiliconorthenovelCopper Indium(Di)Sulphide (CIS)typecellsarenowemerging. Thefrontsideofthe panels areequipped withsolarcells,whilethebackside isusedfortheelectrical wiringandthe synchronization system. Thedesignofthecellsubstrate isaimedtoaccommodate twocelltypes,i.e.,the 'conventional' butheavier andmoreexpensive mono-crystalline SiliconorGaAstypecellsaswellasthe muchlighterandcheaper thinfilm,e.g.,amorphous Siliconcells. ThinfilmSolarCellTechnoloqy Theconcept tobeselected should allowtheuseofnovelthinfilmsolarcelltechnology. The Amorphous Siliconthinfilm(a-Si)cellisresistant tohighradiation environment [2].Anadditional reason forimplementing thiscelltypeisthelowassociated costandlowmass.Typically, the manufacturing process allowsforbulkproduction asrequired forconstellations. Panelandinter-panel electrical harness Therearsideofthepanelisequipped withaflexible electrical harness inordertominimize the resistance torque duringdeployment, especially undertheextreme temperature conditions. A singleuninterrupted flex-print isdesignbaseline toeliminate the'conventional' inter-panel power transfer harness andconnectors. Thiswillresultinalowermass,compact flatstowage, lowcost andareliable performance. Thisflex-print designalsoreduces themultiple ohmiclossesatthe 'conventional' inter-panel connector junctions. Redundancy requires thatthereare2 uninterrupted flex-print harness. Snubber frame: Thepurpose ofthesnubber frameistoobtainastiffinterface towards thespacecraft structure. Asthe thinpanels havearelative highflexibility, itisrequired tohaveastiffframesuchtomeetthestowed resonance frequency requirement offl>35Hz.TheCFRPframehasinterface pointstowards the12 holddownpoints.Itincludes the"jackmechanism". (see:deployment mechanisms section) Pressure Plate(optional) Toclosethebox,apressure plateofCFRPisforeseen. Thetotalpanelpackage iskepttighttogether by means oftheearliermentioned snubber frameandthepressure plate.Thepressure plateinterfaces with thelastlifepanel.Thepressure platecanbedeleted incasethespacecraft actsasasupporting structure. Thismeansacloseinteraction withtheS/Cbuilder. Assuch,25mmofstowage heightcould besaved. 256
Deployment Mechanisms ofWings Thefollowing deployment mechanisms areforeseen forthewings: 1.Secondary deployment system, i.e.,panelhinges 2.Panelcurving system withtension wires 3.Thehold-down element basedonexisting andprovenARAdesign. 4.Deployment damper forspeedcontrol Secondary Deployment System Thepanelsareequipped withthree(3)hingesalongtheirfoldlines.Theactualdeployment ofthesolar panelsstartsafterthelastandcentralconnecting element hasbeencut.Springmotors actuate thesolar panels fold-out. Ateveryhingelinethereisadeployment springintheformofatorsion rod,The deployment ofthepanelsissynchronized bya"soft"synchronization system [3].Between twoadjacent hinge-lines thereisatorquetubeinstead ofapulleyandcablesystem. Thesynchronization tubeismade ofstainless steel,butincasethebending stiffness hastobeless,carbon fiberisanoption. Figure4 showssomedetails. Thepanelhingesonlyhaveastopbutnolock.Thismeansthatthereishardlyany deployment shock.Still,atleastonedamper isrequired tobettercontrolthedeployment trajectory [4].A smalleddy-current orviscous damper isforeseen unwinding acablepulleysystem. Hingebacklash is inherently eliminated whenthehingelineiscurved. Thepanelhinges arepre-loaded viathecurving mechanism andare"forced" toonesideofthehingeedge.  Powertranfer '.,/_ harness._._ /_'/ Synchronization :=: tubes \ // ETens'onw'resI I"'noe''o'O"oeI Figure4.Design ofthesecondary deployment mechanism (incl.thesynchronization "tubes") Duringdeployment, thepanelhinge/fold linesarestiff,because thepanelsareindifferent planes. As soonasthepanelshavelinedupattheendofthesecondary deployment phase, thepanelscanbe curved toformacurvedwing.Inaddition totheprimary curving system, thecurvature ofthepanels is 257
initiatedandmaintained byanumberofspringloadedriggingwiresorstraps,whichrundiagonallybehind eachpaneltoallfourcorners. Whenthepanelsareflat,thewiresarestretchedforabout25mm,butassoonasthepanelsareallowed tocurve,thewiresshorten25mmandpullthepanelsinadefinedcurvature.Adiagonallayoutofthe wiresisselectedtoincreasethetorsionstiffnessofthedeployedwing.Existing,spacequalifiedandlowmassKevlarcablesareforeseentobeusedaspanelriggingwires.Thetensionatstartisabout100N percable.Atthefinalposition,oncethepanelsarecurved,theremainderofthepre-tension isabout60N Theprimary rootcurvinq system (i.e.,"lack"mechanism) Attherootofthewing,embedded inthesnubber frame,several provisions arerequired toallowthewing tochange fromaflatsurface toacurvedone.Thefunction ofthis"jack"mechanism istoinitiatepanel curving oncethedeployment iscompleted. Thenumber andthedesignoftheseprovisions areafunction oftherequired curvedshapeofthewing afterdeployment. Thiscurving system issandwiched between thesnubber frameandthefirstpanel. Figure5showssomedetailsofthedesignconcept. Thefirstpanelisfixedtothesnubber frameatitstip andbasewithtwoshearwebs.Oneshearwebisequipped withonlyonehinge/fold lineandtheother shearwebwithtwohinge/fold lines.Inthisway,thepanelisallowed tohaveexcursions inthecenter,ina similarwayasasuspended leafspring. Theconcertina structure pushes thein-board panelinthecenterawayforabout200mm.Hence, the remainder ofthewingpanelsisurgedtofollow,because ofthepanelcurving system. Thehollowsideof thecurvedwingisnormally pointing awayfromthesun.Fromlaterbreadboard testingitturnedoutthat themechanism couldbeavoided, asthecurving motion wasinitiated bythestrainenergy induced from thetension cablestotheindividual panels. .I//'\\ ....,..,,.__i/ "---. ...+s_Figure5.Panelprimary rootcurving mechanism. 258
Thehold-down system Thehold-down system isdivided intotwosystems: -Theprimary hold-down system -Thesecondary hold-down system Theprimary system restrains the2stowed wingstacksandthePDMduringlaunch, whilethesecondary hold-down system functions inthefirstpartoftheoperational orbittokeepthestillstowed panelstiedto thesnubber frameandprevent premature secondary deployment. ThePDMboomandthetwowingstacksofconcertina-wise foldedpanels, areheldstowed bytwo(2) hold-down elements inaddition tothetwelve(2x6)pointsforeachwing,including therestraint elements. Theother'hardpoints'arenotequipped witharestraint element; theyactas'snubber' points.Incasethe number ofpointsonthespacecraft isnotsufficient tosupport thetwopanelstacks, aframe-like adapter canbeused.Thisframewillprovide therequired extrahardpointsforthesolarpanelstacksagainst the S/Csidewall. Thissnubber-frame itselfisfixedtothebrackets onthesidewall ofthesatellite. Ahold-down pointconsists ofabracket, ahardpointinthepanels, arestraint element, atensioning element andasetof'Thermal Knives'. Therestraint element willbebasedonexisting andprovenARAtypetechnology [5](Figure 6). Also,anewholddownmechanism isbeingdeveloped toachieve amorecompact stowage volume. Itis calledtheMultipurpose HolddownandRelease Mechanism (MHRM). Thecuttingactionisachieved bya thermal (redundant) knife.ThecabletypeisDyneema ®,whichhasalowermelting pointthantheKevlar cableusedintheARAtypeholddown. Also,thecableisnowconfigured inaclosedloop,allowing amore compact stowage volume. Thelocation oftheMHRM wouldbealongthelongedge,aseachwingpanel package isratherslender. Figure7showsthedesignconcept oftheMHRM, typicalforonepaneloption. Extension toalargerpanelstackwillbeimplemented inthefollowonstudy. Thermal Knife CableloopI \ IRestraint bracket I Figure6.ARArestraint basedonaThermal Knife[5]Figure7.MHRM: Multipurpose Holddown& Release Mechanism [7]. 259
Structural Performance Deployed WingAnalyses: Bothforthestowed aswellforthedeployed wing,structural analyses wasexecuted (FEM/NASTRAN). Forthedeployed wing,alsoanon-linear analysis wasdone,inwhichthesensitivity against thefollowing threeparameters wasinvestigated: _. Tension loadsinthestraps/tension wires p- Amount ofcurvature ofthesemi-rigid panels ;,- Stiffness ofthetension wiresorstraps Itwasfoundthatthestiffness ofthe(sub)wingcouldbetunedwiththeaboveparameters. Thelowest frequency wasatorsion mode(f_-0.278Hz),followed byabending mode(fl~0.577Hz).Increase in torsionrigiditycouldbeachieved byincreasing thestiffness ofthetension strapsincombination withthe panelcurvature. Also,thePDM-stiffness determines theoverall deployed wingbehavior greater than expected. Thiscanbetunedbythegeometry oftheCFRPtubeoftheprimary linkage. Y __'-"-_ _*....... =-._;:........,....................3.0;/ X Torsion mode:0.278Hz ":W:::;/-.•01 .__, -__:._.. ._.__..Bending mode:0.577Hz Yy .Yy - __ × Figure8,Typical modeshapes ofafullydeployed wing(Ref:Teledesic) Massperformance: Apreliminary masstable(SeeTable2)isprovided showing thetypicalmasses involved (based onthe relative lowvalueofabout8%cellefficiency atEOLconditions). ThePDMandallmechanisms are included whichcontribute to-30%oftheoverallmass.Especially, thePDMwiththedoublegimbalSADM isaheavystructure, andisfullydrivenbytheconstellation requirements. Butincaseanormal V-yoke structure isused,andnoactivemotorized PDMhingesareused,considerable masssavingisforeseen. 260
Note,thatincasetransferismadetothemuchmoreefficientCIStypethinfilmcells(i.e.,withapredicted efficiencyof15%EOL),thepowertoweightratiowouldincreasetoabout70-80W/kgEOLforthe2 individualwingstructure(excluding thePDM,excluding 20%margin). Table2:Predicted massperformance ofatypicalsolararray(Ref2;Teledesic) Assembly Wing SolarArrayItem Unitmass No Mass typ.(kg) (kg) Solarpanelstructure CFRP 1.599 14 22.386 PVA(basedon8%a-Sicells,70W/m2EOL)0.540 14 7,560 Pressure Plate 5.626 1 5,626 Snubber Frame 3.511 1 3,511 Interpanelhinges 0.030 40 1,200 Interpanelsnubbers 0.003 285 0.855 Primary hold-down system (per2) 0.385 6 2,310 Snubber frameshearwebs 0.146 1 0,146 Synchro !panelcurving system 2.495 1 2,495 Wingdamper system 0.920 2 1840 Powertransfer harness 1.143 1 1,143 Telemetry deployment /wingstatus 0.120 1 0.120 Telemetry /sunsensing 0.500 1 0.500 49.692 2 99.384 Wings EOLWatts/kg(excl.PDMetc) Primary Deployment Arm PDMHolddownSystem PDMPowerTransfer Harness SADM/gimbalmechanism43W/kg 14.187 1 14.187 0.266 2 0.532 2.407 1 2.407 15.000 1 15.000 EOLWatts/kg(withPDMetc) 32W/kg131.51 Deployment characteristics: Topredict thedeployment behavior, aspecial purpose program thatisnormally usedforourstandard ARA-type solararrayswasupdated (SMX-Deploy). Also,thedamper effectwasincluded. Asensitivity studywasdonetoinvestigate criticalbehavior, Itappeared ratherquickly thatthecalculation method with SMXDeploy couldnotbeused.Anewdeployment program wasrequired thatcoulddealwiththehighly non-linear behavior. Thisanalysis activitywasinitiated withtheMulti-Body software codecalledADAMS, andisnowinprocess. Breadboarding aspartofD&D A50%scalemodelwasbuiltcomprised of15panels of2.2mx0.25m.Aspecial deployment rigwas manufactured tosimulate zerogravity. Thisairbearing deployment rigwasbasedonaprevious Fokker Spacein-house development [6].Thefollowing testswereperformed: Functional deployment andlifetesting Torque surplus measurements (indicative) _"Stiffness andbacklash (qualitative &quantitative) Themodelwasstowed anddeployed several times(>20x). Thewingdeployment wasinitiated bythe torsion springs incombination withthesynchronization tubes. Initialreservations concerning the deployment initiated weretakenaway.Duringthedeployment, somepanelsalready curved totheirfinal curvature whileothersstillremained flat.Thiswasnotexpected andcouldalsonotbepredicted asitwas considered tobeahighlynon-linear kinematics deployment behavior. Itcouldbeexpected however, that 261
anyprematurecurvingofthepanelwouldimpose excessive friction forcesatthehinges. Inreality,this wasnotthecase.Onthecontrary, ithelped thedeployment motion bytheenergyreleased fromthe contracting tension cables. Figure9.Thescalemodel(50%)oftheCurwin infullydeployed state(onewing) Arestartofthedeployment wasalwayspossible, evenwhenacertainamount ofthepanelshadalready reached theircurved (=deployed) position. Itwasobserved thatthecurving ofthepanels nearthe deployed condition isnotapotential resistance source, butinstead stimulates thefurtherdeployment of theremaining panels. DuringthewingBBtestingitwasclearlydemonstrated thatsomeformofdamping wouldberequired tocontroltheadditional deployment energycoming fromthecurving panelsandalsoto controlthedeployment trajectory. Table3:Stiffness performance ofatypical 50%scalewing(Ref2;Teledesic) Outofplanestiffness Elaverage Inplanestiffness Elaverage Torsion stiffness GIpaverage1600Nm2 20000Nm2 740Nm2 Thestiffness asmeasured fulfilled therequirement. Anexpected knock-down of4wasobserved in bending andtorsion, however, stillmeeting theanticipated requirement offl>0.3Hzfortheflightdesign case.Lifetesting showed noobservable degradation; thetension wiresremained intactwithno degradation. Although nolargeshocks occurred, itwasconsidered mandatory toimplement a 262
deployment damperinordertolimittheoff-axis trajectory ofthesecondary deploying solararray.The relative softsynchronization requires sometypeofdamper tocoordinate thedeployment behavior. This wasimplemented intheupdate designsketches. Thedesignsketches andtheirinputswereusedfora complete PhaseC/Dcosting andtodetermine thefeasibility ofthe$1M/array.Unitdevelopments were identified forfurtherdetailing andcosting forphaseC/D.Moreeffortisrequired fortheAITstreamlining as thisdetermines thebulkcosting fortherecurring phase. SinglerailO-gdeployment rig1 Winginstowed configuration I Figure10.Curwin fullydeployed, shown fromrearsideshowing synchronization tubesand tension cables 263
Conclusions Thereservations concerning thedeployment feasibility andassociated deployment principles werefully covered bythe50%scalebreadboard model. Deployment wasfeasible despite firstreservations. Stiffening ofthedeployed arraybybending theultra-thin panelsnotonlyworked onpaper,butwasalso proveninextensive BB-testing. Summarizing: 1.Compact stowage ofthethinpanelstackcanbeachieved, stillmeeting thefrequency requirements forboththestowed configuration aswellasthefullydeployed configuration. 2.Themechanism complexity suchasHDRS andhinges arewellknown andtherefore arenot considered alargerisk(Fokker SpaceBVdesignheritage). 3.ThePDMisarelatively complicated andcostlymechanism duetotherequired deployment trajectory. Commonality forTeledesic andCelestri couldstillbeachievable. 4.Apower-to-weight ratioof35W/kgatEOLforMEOandalsounderarelative highradiation environment after10yearsseemsachievable. Thisisthecaseforthethinfilmcells,butevenmorefor theCIStypecells,whichofferbetterperformance. (i.e.,70to80W/kgatEOLconditions) 5.Implementation ofthinfilmtechnology onultrathinpanelswasselected asthemostfavorable cell technology. Thiscouldbeeithera-Sitypecells,newCIStypecellspresently underdevelopment oreven ultrathinGaAscellswhicharenowalsoemerging. Acknowledgements IwanttothanktheDutchNIVRforfunding thisstudy.Without theirhelp,a"jumpstart"wouldneverhave beenpossible. Theyrealized theimportance tobethereintime.Also,thephase-A andphase-B team, whichcreated atremendous outputinsuchashorttime.Thishelped thedesign teamtoestablish a baseline relatively quickly. Lastbutnotleast,thehardworking BB-team, which, withtheirenthusiasm, wereabletocreate,buildandtestthis50%scalemodelinjust3months. References 1.CurwinSolarPanelSystem: PatentNo:US6,091,016 andEuropa 97204099.2 2.PhaseB1datapackage toNIVRCONSA anda-Sidevelopments: CONSA-SER-SER-095 3.SolarPanelSynchronization System patentNoUS6,031,178 andEP98200951.6 4.BBtestresultsof1/2scaleCurwin©model:PhaseB1studynoteCONSA-FS-SER-094 5.Thermal Cutting Hold-down &Release Mechanism. Second European SpaceMechanisms &Tribology Symposium, Meersburg, Germany, Symposium BookESASP-231. 6.Fokker Space SolarArrayDeployment Rigs(A.P.Eggers andH.J.Cruijssen; 33r_Aerospace Mechanisms Symposium, Pasadena, California May1999;NASA/CP-1999-209259, 7.System forholding together andseparating partsofaconstruction: USpatent: 5,716,157 and EP:0716982 264
assembly iscaptured usingaseriesofretaining springs, whichsecure itinthepostdeployment position. Todeploy thesatellite theseparation boltsarecutwithpyrotechnically actuated boltcutters EJECTION SYSTEM SPRING  SATELLITE LOADED PUSHER PLATEMARMAN CLAMP \ N SEPARATION .... ., _ _ .... PLANE0° 180° PYROTECHNIC CU'Iq-E RHOUSING EJECTION SPRINGS (NOT SHOWN) PUSH ROD Figure 1.PUSH PLATE ASSEMBLY  LINEAR BEARING Integrated Satellite, PES,Marman BandAssembly Section INTERCONNECT LEAF SPRING BOLT CATCHER 270° \ 90° MARMAN CLAMP\ \ 180° SECTION B-B (ORIGINAL DESIGN)SECTION THROUGH SEPARATION PLANE/MARMAN CLAMP Figure 2.Original SSPPMarman BandAssembly 266
Figure 6.Linkage ETU-Unfolded (Sideview) Vibration testing Thetestarticlewassubjected torandom vibration testing totheGeneral Environmental Verification Specification forSTS&ELVPayloads, Subsystems &Components, GEVS-SE, protoflight levelsof 14.1Grr,sforoneminuteineachoftwoaxes.Thefirstwasthe"Z"orvertical axisandthesecond a "X/Y"composite orhorizontal axis.Ironically theonlyanomaly noticed duringvibration testing wasthe backing offofimproperly stakedsetscrews usedtoeliminate vibration intheclamped configuration of thelinkage. Apostvibration functional testconsisting ofdualpyrotechnic boltcutterfiringswas conducted toverifyresults. Thepostvibration functional mechanism demonstration wasobserved to actuate nominally andinacontrolled manner asinthefunctional tests. Conclusion Themarman bandinterconnect linkage mechanism permits theexisting SSPP ejection system marman clampassembly tomeettheprimary requirement ofeliminating satellite re-contact during deployment bychanneling thestoredenergy present inthesystem toactuate andcontrol assembly motion. Compatibility andadaptability requirements aresimplymetsincethelinkage isaone-to-one replacement fortheexisting assembly components. Itdemonstrates asignificant improvement tothe wellused,testedandanalyzed mechanism. Utilizing the3Dmodeling capability ofProEngineer TMfordesign development provided anefficient andaccurate method formeeting thedesign goals. Theuseofhigh-speed videorecording asatestandinvestigation toolhasproven tobeanimportant development forSSPPinmechanism evaluation anddocumentation. Sinceit'sinitialuseduringthe HHonflightanomaly investigation, high-speed filmandhigh-speed digitalvideohavebeenused extensively forthetestandevaluation ofmechanisms ontheHH,Triana, andSHELS projects. The videoplayback isalsoaneffective toolfordemonstrations andpresentations. Thefirstscheduled application andflightforthenewlinkage willbeduringthesummer of2001.The SimpleSAT experiment willbedeployed usingthePESaboard Discovery STS-105. 272
PULLDOWN STOPS ---_ _ MARMAN BANDINTERCONNECT LINKAGE ASSEMBLY / k S _"_LY)ISOMETRIC VIEW TOPVIEW MARMAN BANDASSEMBLY /PES-CLAMPED CONFIGURATION Figure7 --PULLDOWNSTOPS Y(X)i X(Y) ISOMETRIC VIEW TOPVIEW EXPANDED MARMAN BANDASSEMBLY -RESTING ONPULLDOWNSTOPS Figure8 273
References 1.NASAGSFC, "Hitchhiker Customer Accommodations &Requirements Specification", 740-SPECO08,1999. 2.NASAMSFC, "Design Criteria forControlling Stress Corrosion Cracking", MSFC-SPEC-522B, 1987. 3.Slebzak, G.R.,"Hitchhiker Marman BandInterconnect Linkage TestProcedure", SSPP, 870PROC-587, 2000. 4.NASA GSFC, "General Environmental Verification Specification forSTS&ELVPayloads, Subsystems, andComponents" GEVS-SE, 1996. 274