Astro Digital Ignis Orbital Debris Assessment Report (ODAR

Astro Digital Ignis ODAR – Version 1.0

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Astro Digital Ignis Orbital Debris Assessment Report (ODAR)

ASTRO-DIGITAL-IGNIS-ODAR-1.0

This report is presented as compliance with NASA-STD-8719.14, APPENDIX A. Report Version: 1.0, 11/12/2015

Astro Digital US, Inc. 3171 Jay St, Santa Clara, CA 95054 Document Data is Not Restricted. This document contains no proprietary, ITAR, or export controlled information. DAS Software Version Used In Analysis: v2.0.2

DocuSign Envelope ID: B28A276D-0687-4495-ACD8-FBAEEBFDF591


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Astro Digital Ignis Orbital Debris Assessment Report ASTRO-DIGITAL-IGNIS-ODAR-1.0

APPROVAL:

Chris Biddy CEO, Astro Digital

______________________________

JanA.KingCTO,AstroDigital

_______________________________________

PatrickShannonProgramManagerIgnisProgram

________________________________________


4/19/2019

4/19/2019

4/19/2019


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RevisionRecordRevision: Date: AffectedPages: Changes: Author(s):

1.0 4/11/2019 All–Initial DASSoftwareResultsOrbitLifetimeAnalysis

B.Cooper

TableofContents

Self-assessmentandOSMAassessmentoftheODARusingtheformatinAppendixA.2ofNASA-STD-8719.14:.............................................................................................................3Comments..............................................................................................................................................4AssessmentReportFormat:...........................................................................................................5MomentusFervorDescription:....................................................................................................5ODARSection1:ProgramManagementandMissionOverview.......................................5ODARSection2:SpacecraftDescription......................................................................................6ODARSection3:AssessmentofSpacecraftDebrisReleasedduringNormalOperations................................................................................................................................................10ODARSection4:AssessmentofSpacecraftIntentionalBreakupsandPotentialforExplosions.……………………………………………………………………………………………………….11ODARSection5:AssessmentofSpacecraftPotentialforOn-OrbitCollisions............16ODARSection6:AssessmentofSpacecraftPost-missionDisposalPlansandProcedures...............................................................................................................................................17ODARSection7:AssessmentofSpacecraftReentryHazards...........................................19ODARSection8:AssessmentforTetherMissions...................................................................20RawDAS2.0.2Output………………………………………………………………………………………21AppendixA:Acronyms.......................................................................................................................30



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Self-assessmentoftheODARusingtheformatinAppendixA.2ofNASA-STD-8719.14:AselfassessmentisprovidedbelowinaccordancewiththeassessmentformatprovidedinAppendixA.2ofNASA-STD-8719.14.

Note1:Theprimarypayloadsforalllaunchmissionsbelongtootherorganizations.ThisisnotaprimarymissionofAstroDigital.AllotherportionsofthelaunchcompositearenottheresponsibilityofAstroDigitalandtheIgnisProgramisnottheleadlaunchorganization.AssessmentReportFormat:ODARTechnicalSectionsFormatRequirements:AstroDigitalUS,Inc.isaUScompany.ThisODARfollowstheformatinNASA-STD-8719.14,AppendixA.1andincludesthecontentindicatedasaminimum,ineachofsections2through8belowfortheIgnissatellite.Sections9through14applytothelaunchvehicleODARandarenotcoveredhere.IgnisSpaceMissionProgram:ODARSection1:ProgramManagementandMissionOverviewProgramManager:PatrickShannonMissionManager:BrianCooperSeniorManagement:ChrisBiddyForeigngovernmentorspaceagencyparticipation:None.SummaryofNASA’sresponsibilityunderthegoverningagreement(s):N/A



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Scheduleofupcomingmissionmilestones:

• ShipmentofSpacecraft:Q32019• FirstLaunch:Q32019

MissionOverview: TheIgnisspacecraftisatechnologydemonstrationspacecraftbuilttothe6UCubeSatstandard.ItincludestheApolloConstellationEngine(ACE),aHallThrusterwhich ionizesandexpels ahighdensityproprietarypropellant.ThespacecraftwillbelaunchedaboardaLauncherOnerocketbuiltbyVirginOrbit.Thespacecraftbus istheCorvus-6design.Thecommonsatellitebususesreactionwheels, magnetic torque coils, star trackers, magnetometers, sun sensors, andgyroscopestoenableprecision3-axispointingwithouttheuseofpropellant.LaunchVehiclesandLaunchSites:LauncherOneandCosmicGirl,GuamInternationalAirport,Guam,UnitedStates.ProposedInitialLaunchDate:Q32019MissionDuration:Theanticipatedlifetimeofthespacecraftis≤3yearsinLEO.Launchanddeploymentprofile,includingallparking,transfer,andoperationalorbits with apogee, perigee, and inclination: The selected launch vehicle willtransportmultiplemissionpayloadstoorbit.TheIgnisspacecraftwillbedeployedintoamediuminclinationlowEarthorbit.Oncethefinalstagehasburnedout,thesatellitepayloadswillbeautomaticallydispensed.TheIgnisspacecraftwilldeployaUHFantennaandsolarpaneloncereleasedfromtheLauncherOneupperstage.Thespacecraftwilldecaynaturallyfromoperationalorbitswithinthefollowingorbitalparameters:NominalOrbitalAltitude:500kmEccentricity:0.0000to0.0033Inclination:44.5°to45.5°TheIgnispropulsionsystemisconsideredtobeexperimental,andassuchisnotbeingreliedonasaviabledeorbitmethod.Thespacecraftwillbelaunchedintoanorbitthatwillresultinanaturalorbitaldecayinlessthan5years.ODARSection2:SpacecraftDescription:



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Physicaldescriptionofthespacecraft:The Ignis spacecraft uses the standard Corvus-6 bus, which is based on the 6UCubeSatformfactor.Basicphysicaldimensionsare32cmx21cmx11cmwithamassof no more than 12 kg. The superstructure is comprised of 6 outer panels withmultiplesubsystemsservingasinternalsupportingstructuresbridgingbetweenthepanels. There are L rails along each of the 32 cm edgeswhich accommodate thedeploymentofthesatellitefromthedeployer.Thebuselectronicsprovideadditionalinternalstabilitytothestructure.TheACEthrusternozzleislocatedonthe-Xfaceofthe spacecraft with the thrust axis passing through the center of mass of thespacecraft.Thespacecraftbusincludesaspring-loadedUHFantennaanddeployablesolarpanelwhicharedeployedafterjettisonfromthedeployerbyaburnwirecontrolledbyasoftware timer via the flight computer. Power is locked away fromall spacecraftplatformandpayloadcomponentsbymeansofredundantseriesseparationswitches.Theseswitchescannotbeactivateduntilthespacecraftseparatesfromthedeployerstructure.ThespacecraftisdepictedinFigure1.

Figure1:IgnisSpacecraft

Totalsatellitemassatlaunch,includingallpropellantsandfluids:Ignis:12.0kg+0.0kg/-1.0kg



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Drymassofsatellitesatlaunch:Ignis:10.9kg+0.0kg/-1.0kgDescription of all propulsion systems (cold gas, mono-propellant, bi-propellant,electric,nuclear): TheACEpropulsionsystemionizesahighdensityproprietary inert propellant (see Proprietary informationExhibit for details). Theionized propellant is expelled out of the thruster using the Hall effect at an Ispexceedingtraditionalchemicalpropulsionsystems.TheexpectedthrustandIspwillvarywith power input levels. Thrustwill not exceed 25mNmaximum. The Ignispropulsionsystemincludes1.1kgofpropellant.Thepropulsionsystemasdesignedhasapproximately12,000Nstotalimpulseavailable.The propulsion system is pressurized using inert nitrogen. The beginning-of-lifepressureis15psi,orapproximatelyatmosphericpressure.Thetankismanufacturedfromsteeltopreventanylifetimecyclingissuesfromcausinganunintendedruptureandreleaseofpropellants.Identification, includingmass and pressure, of all fluids (liquids and gases)plannedtobeonboardandadescriptionofthefluidloadingplanorstrategies,excludingfluidsinsealedheatpipes:Up to 1.1 kilograms of propriety propellant and less than 10 grams of gaseousnitrogen pressurant gas. These fluids will be loaded prior to integration of thespacecraft into the standard CubeSat deployer. There will be approximately zerogaugepressureat the timeofpropellant loading through launch.Theatmosphericpressurepresentwillresultinthetankpressurizingtoapproximately15psiaduringlaunch.Thepressurevesselhasbeentestedto3xtheproofpressureandisqualifiedfortransportationbytheDOT,asitisnotapressurevesselwhileintheatmosphere.FluidsinPressurizedBatteries:NoneThe Corvus-6 satellite bus design uses four unpressurized standard Lithium-Ionbatterycellsineachspacecraft.Theenergycapacityofeachcellis17.5W-hrs.Thetotalcapacityenergycapacityperspacecraftis70W-hrs.There is an additional payload battery system which uses nine unpressurizedstandardLithium-ionbatterieswiredinseriestoprovidea36voltoutputtotheACEthruster.Eachcellhasacapacityof17.5W-hrsforatotalbatterycapacityof157W-hrs.Descriptionofattitudecontrolsystemandindicationofthenormalattitudeofthespacecraftwithrespect to thevelocityvector: The Ignisspacecraftwillbeinitiallycontrolledbymagnetictorquecoilsembeddedinthefixedsolarpanelsofthespacecraft.Thesewillbeusedtodetumblethespacecrafttoalowenoughratesuchthatthereactionwheelscantakeoverandprovideprecision3-axisattitudecontrol.Therearefourprimaryattitudemodes.



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• Asunpointingmodethatisoptimizedforsolarpowergenerationfromthe

satellite.Thespacecraft’slargefixedpanelswillbeorientedtowardsthesun.Thismodewillmakeuseofmagnetometers,sunsensors,reactionwheels,andmagnetictorquerstoorientthespacecraftcorrectly.

• Avectortrackingmode,whichwillallowthethrustaxistobepointedinanydirectionininertialspace.Thismodewillmakeuseofreactionwheels,gyroscope,andstartrackertoorientthespacecraft.

• AnACSidlemode,whichwillallowthespacecrafttospinupinapredictablemannerwhilefiringthethrusterforlongdurationswithoutthereactionwheelsprovidinganytorque.

• Adetumblemodethatthespacecraftwillenterafterdeployingfromthelaunchvehicleorafterspinningupduringalongdurationthrusterfiring.

Description of any range safety or other pyrotechnic devices: None. Thespacecraft deploy its antenna and solar panel using a burnwire system. Systempower is locked off during launch by two series and two parallel deploymentswitches,buttheCubesatdeployerpreventsanyformofprematuredeployment,inanycase. Theantennaandpanelspringconstantsareverylowandcanbeheldinplacebyhand.Description of the electrical generation and storage system: Standard COTSLithium-Ion battery cells are charged before payload integration and provideelectricalenergyduringtheeclipseportionofthesatellite’sorbit.Thebusbatteriesareoperatedinan“all-parallel”arrangementthatresultsinincreasedsafetythankstonaturalvoltagebalancingbetweencells.Thepayloadbatteryusesacommercialbatterymanagementsystemtoensureallcellsinthestringareatthesamestateofcharge, preventing overcharging events. A series of Triple Junction Solar Cellsgenerateamaximumon-orbitpowerofapproximately36wattsattheend-of-lifeofthemission(5years forcalculationpurposes). Typicalbusoperationsconsume8wattsofpoweronaverage.Thethrustercanconsumeupto300wattsinshortbursts.Thecharge/dischargecycleismanagedbyapowermanagementsystemoverseenbytheFlightComputer.Identificationofanyothersourcesofstoredenergynotnotedabove:None.Identificationofanyradioactivematerialsonboard:None.ODARSection3:AssessmentofSpacecraftDebrisReleasedduringNormalOperations:Identificationofanyobject(>1mm)expectedtobereleasedfromthespacecraftanytimeafterlaunch,includingobjectdimensions,mass,andmaterial:None.



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Rationale/necessityforreleaseofeachobject:N/A.Timeofreleaseofeachobject,relativetolaunchtime:N/A.Releasevelocityofeachobjectwithrespecttospacecraft:N/A.Expectedorbitalparameters(apogee,perigee,andinclination)ofeachobjectafterrelease:N/A.Calculatedorbitallifetimeofeachobject,includingtimespentinLowEarthOrbit(LEO):N/A.AssessmentofspacecraftcompliancewithRequirements4.3-1and4.3-2(perDASv2.0.2)4.3-1,MissionRelatedDebrisPassingThroughLEO:COMPLIANT4.3-2,MissionRelatedDebrisPassingNearGEO:COMPLIANTODARSection4:AssessmentofSpacecraftIntentionalBreakupsandPotentialforExplosions.Potentialcausesofspacecraftbreakupduringdeploymentandmissionoperations:Therearethreepotentialscenariosthatcouldpotentiallyleadtoabreakupofthesatellite.Inorderofcredibility:

1) Ruptureofapropellanttank(N2pressurantgasandpropellant);and2) Lithium-ionbatterycellfailure.

Summaryof failuremodesandeffects analysesof all credible failuremodeswhichmayleadtoanaccidentalexplosion: Thein-orbitfailureofabatterycellprotection circuit could lead to a short circuit resulting inoverheatingandaveryremotepossibilityofbatterycellexplosion.Thebatterysafetysystemsdiscussedinthe FMEA (see requirement 4.4-1 below) describe the combined faults thatmustoccurforanyofseven(7)independent,mutuallyexclusivefailuremodestoleadtosuchanexplosion.Detailedplanforanydesignedspacecraftbreakup,includingexplosionsandintentionalcollisions:Therearenoplannedbreakups.ListofcomponentswhichshallbepassivatedatEndofMission(EOM)includingmethodofpassivationandamountwhichcannotbepassivated:Thirteen(13)LithiumIonBatteryCells.Rationaleforallitemswhicharerequiredtobepassivated,butcannotbeduetotheirdesign:None.



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AssessmentofspacecraftcompliancewithRequirements4.4-1through4.4-4:Requirement4.4-1:Limiting therisk tootherspacesystems fromaccidentalexplosions during deployment andmission operations while in orbit aboutEarthortheMoon:“Foreachspacecraftandlaunchvehicleorbitalstageemployedforamission,theprogramorprojectshalldemonstrate,viafailuremodeandeffectsanalyses or equivalent analyses, that the integrated probability of explosion for allcredible failure modes of each spacecraft and launch vehicle is less than 0.001(excludingsmallparticleimpacts)(Requirement56449).”Compliancestatement:RequiredProbability:0.001Expectedprobability,Ignis:0.0000SupportingRationaleandFMEAdetails:PressureTankExplosion:Effect:Aruptureofthepropellanttankwouldreleasegaseousnitrogenandupto1.1kgofliquidpropellant.Duetothelowpressure(15psia),thepenetratingenergyofany debris would be relatively low. The propellant tank is enclosed in the solidaluminumstructuralpanelsofthespacecraft.Thesealuminumwallswouldcontainanysoliddebriswithinthebodyofthespacecraft.Dropletsofpropellantwouldbeexpectedtoleakoutofventholesinthespacecraftbody,butwouldreenterquicklyduetotheirhighareatomassratio.Probability:Verylow.Astructuralfailureofthetankwouldneedtooccur,andthemechanismsbywhich these failuresoccurareverywellunderstood.Cubesatsaretypicallyvolume-limitedasopposedtomass-limited.Thismeansthatitisveryeasytoaddmass toagivenstructure toprotectagainst failure,andstructuralstrengthmarginscanbeveryhigh.ThisapproachisemployedinthedesignofthepressurevesselsfortheIgnisspacecraft.Additionally,thetankdesignwillbecertifiedbyUSDoT for commercial transporation.Whereas typical aerospace componentswouldhaveamarginofsafetyunder2,allstructuresontheCorvussatellitedesignshavestrengthtofailuremarginsof3orgreater.Batteryexplosion:



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Effect:Allfailuremodesbelowmightresultinbatteryexplosionwiththepossibilityoforbitaldebrisgeneration.However,intheunlikelyeventthatabatterycelldoesexplosively rupture, the small size, mass, and potential energy, of these smallbatteries is such that while the spacecraft could be expected to vent gases, mostdebrisfromthebatteryruptureshouldbecontainedwithinthespacecraftduetothelackofpenetrationenergytothemultipleenclosuressurroundingthebatteries.Probability:ExtremelyLow.Itisbelievedtobelessthan0.01%giventhatmultipleindependent(notcommonmode)faultsmustoccurforeachfailuremodetocausetheultimateeffect(explosion).Failuremode1:Internalshortcircuit.Mitigation1:Protoflightlevelsineburst,sineandrandomvibrationinthreeaxesofbothspacecraft,thermalvacuumcyclingofbothspacecraftandextensivefunctionaltestingfollowedbymaximumsystemrate-limitedchargeanddischargecycleswereperformed to prove that no internal short circuit sensitivity exists. Additionalenvironmentalandfunctionaltestingofthebatteriesatthepowersubsystemvendorfacilitieswerealsoconductedonthebatteriesatthecomponentlevel.Combinedfaultsrequiredforrealizedfailure:EnvironmentaltestingANDfunctionalcharge/dischargetestsmustbothbeineffectiveindiscoveryofthefailuremode.FailureMode2:Internalthermalriseduetohighloaddischargerate.Mitigation2:Batterycellsweretestedinlabforhighloaddischargeratesinavarietyofflight-likeconfigurationstodetermineifthefeasibilityofanout-of-controlthermalriseinthecell.Cellswerealsotestedinahot,thermalvacuumenvironment(5cyclesat 50°C, then to -20°C) in order to test theupper limit of the cells capability.Nofailureswereobservedoridentifiedviasatellitetelemetryorviaexternalmonitoringcircuitry.Combined faults required for realized failure: Spacecraft thermal design must beincorrectANDexternalover-currentdetectionanddisconnectfunctionmustfailtoenablethisfailuremode.FailureMode 3: Excessive discharge rate or short-circuit due to external devicefailure or terminal contact with conductors not at battery voltage levels (due toabrasionorinadequateproximityseparation).Mitigation3:Thisfailuremodeisnegatedby:a)qualificationtestedshortcircuitprotectiononeachexternalcircuit,b) design of batterypacks and insulators such that no contactwithnearbyboardtracesispossiblewithoutbeingcausedbysomeothermechanicalfailure,



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c)observationofsuchothermechanicalfailuresbyprotoflightlevelenvironmentaltests(sineburst,randomvibration,thermalcycling,andthermal-vacuumtests).Combinedfaultsrequiredforrealizedfailure:Anexternalloadmustfail/short-circuitANDexternalover-currentdetectionanddisconnectfunctionmustalloccurtoenablethisfailuremode.FailureMode4:Inoperablevents.Mitigation 4: Battery venting is not inhibitedby thebatteryholderdesignor thespacecraftdesign.Thebatterycanventgasestotheexternalenvironment.Combinedeffectsrequiredforrealizedfailure:ThecellmanufacturerORthesatelliteintegratorfailstoinstallproperventing.FailureMode5:CrushingMitigation5:Thismodeisnegatedbyspacecraftdesign.Therearenomovingpartsintheproximityofthebatteries.Combinedfaultsrequiredforrealizedfailure:AcatastrophicfailuremustoccurinanexternalsystemANDthefailuremustcauseacollisionsufficienttocrushthebatteriesleadingtoaninternalshortcircuitANDthesatellitemustbeinanaturallysustainedorbitatthetimethecrushingoccurs.FailureMode6:Lowlevelcurrentleakageorshort-circuitthroughbatterypackcaseorduetomoisture-baseddegradationofinsulators.Mitigation6:Thesemodesarenegatedby:

a) batteryholder/casedesignmadeofnon-conductiveplastic,and

b) operationinvacuumsuchthatnomoisturecanaffectinsulators.

Combinedfaultsrequiredforrealizedfailure: Abrasionorpiercingfailureofcircuitboard coating orwire insulatorsAND dislocation of battery packsAND failure ofbatteryterminalinsulatorsANDfailuretodetectsuchfailuresinenvironmentaltestsmustoccurtoresultinthisfailuremode.FailureMode7:Excesstemperaturesduetoorbitalenvironmentandhighdischargecombined.Mitigation7:Thespacecraftthermaldesignwillnegatethispossibility.Thermalrisehasbeenanalyzed in combinationwith spaceenvironment temperatures showingthat batteries do not exceed normal allowable operating temperatures under a



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variety ofmodeled cases, includingworst case orbital scenarios. Analysis showsthesetemperaturestobewellbelowtemperaturesofconcernforexplosions.Combinedfaultsrequiredforrealizedfailure:ThermalanalysisANDthermaldesignAND mission simulations in thermal-vacuum chamber testing AND over-currentmonitoringandcontrolmustallfailforthisfailuremodetooccur.Requirement 4.4-2: Design for passivation after completion of missionoperationswhileinorbitaboutEarthortheMoon:“Designofallspacecraftandlaunchvehicleorbitalstagesshall includetheabilitytodeplete all onboard sources of stored energy and disconnect all energy generationsources when they are no longer required for mission operations or post-missiondisposalorcontrol toa levelwhichcannotcauseanexplosionordeflagration largeenoughtoreleaseorbitaldebrisorbreakupthespacecraft(Requirement56450).”Compliancestatement:TheIgnissatelliteincludestheabilitytofullydisconnecttheLithiumIoncellsfromthechargingcurrentofthesolararrays.AtEnd-Of-Life,thisfeaturecanbeusedtocompletelypassivatethebatteriesbyremovingallenergyfromthem.Intheunlikelyeventthatabatterycelldoesexplosivelyrupture,thesmallsize,mass,andpotentialenergy,ofthesesmallbatteriesissuchthatwhilethespacecraftcould be expected to vent gases, the debris from the battery rupture should becontainedwithinthespacecraftduetothelackofpenetrationenergytothemultipleenclosuressurroundingthebatteries.The thruster propellant will not be actively vented due to the desire to preventsolidified propellant droplets from causing a debris hazard at survivable orbits.Instead,thethrusterwillbefireduntilallpropellantsandpressurantsareexpelled.Requirement4.4-3. Limiting the long-term risk to other space systems fromplannedbreakups: Compliance statement: This requirement is not applicable.Therearenoplannedbreakups.Requirement4.4-4:Limitingtheshort-termrisktootherspacesystemsfromplanned breakups: Compliance statement: This requirement is not applicable.Therearenoplannedbreakups.ODARSection5:AssessmentofSpacecraftPotentialforOn-OrbitCollisionsAssessmentofspacecraftcompliancewithRequirements4.5-1and4.5-2(perDASv2.0.2,andcalculationmethodsprovidedinNASA-STD-8719.14,section4.5.4):Requirement4.5-1.LimitingdebrisgeneratedbycollisionswithlargeobjectswhenoperatinginEarthorbit:



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“Foreachspacecraftandlaunchvehicleorbitalstage inorpassingthroughLEO,theprogramorprojectshalldemonstratethat,duringtheorbitallifetimeofeachspacecraftandorbitalstage,theprobabilityofaccidentalcollisionwithspaceobjectslargerthan10cmindiameterislessthan0.001(Requirement56506).”LargeObjectImpactandDebrisGenerationProbability:0.00001;COMPLIANT.Requirement4.5-2.LimitingdebrisgeneratedbycollisionswithsmallobjectswhenoperatinginEarthorlunarorbit:“Foreachspacecraft,theprogramorprojectshalldemonstratethat,duringthemissionof the spacecraft, the probability of accidental collision with orbital debris andmeteoroidssufficient topreventcompliancewith theapplicablepostmissiondisposalrequirementsislessthan0.01(Requirement56507).”SmallObjectImpactandDebrisGenerationProbability:0.0000;COMPLIANTIdentificationofallsystemsorcomponentsrequiredtoaccomplishanypost-missiondisposaloperation,includingpassivationandmaneuvering:NoneODARSection6:AssessmentofSpacecraftPost-missionDisposalPlansandProcedures6.1Descriptionofspacecraftdisposaloptionselected:Thesatellitewillde-orbitnaturallybyatmosphericre-entry.6.2 Plan for any spacecraftmaneuvers required to accomplish post-missiondisposal: No maneuvers are required to accomplish post-mission disposal. Theexperimental thrusterwill beused to attempt to lower theorbit, but this isnot arequirementtoachievethedescribeddisposalplan.6.3Calculationofarea-to-massratioafterpost-missiondisposal,ifthecontrolledreentryoptionisnotselected:SpacecraftMass:12.0kg(selectedasworstcasemass)Cross-sectionalArea:0.124m^2(averagetumbling)(CalculatedbyDAS2.1.1).Areatomassratio:0.0103m^2/kg6.4AssessmentofspacecraftcompliancewithRequirements4.6-1through4.6-5 (per DAS v 2.0.2 and NASA-STD-8719.14 section): Requirement 4.6-1.DisposalforspacestructurespassingthroughLEO:



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“Aspacecraftororbitalstagewithaperigeealtitudebelow2000kmshallbedisposedofbyoneofthreemethods:(Requirement56557)a.Atmosphericreentryoption:Leavethespacestructureinanorbitinwhichnaturalforceswillleadtoatmosphericreentrywithin25yearsafterthecompletionofmissionbut no more than 30 years after launch; or Maneuver the space structure into acontrolledde-orbittrajectoryassoonaspracticalaftercompletionofmission.b.Storageorbitoption:Maneuverthespacestructureintoanorbitwithperigeealtitudegreaterthan2000kmandapogeelessthanGEO-500km.c.Directretrieval:Retrievethespacestructureandremoveitfromorbitwithin10yearsaftercompletionofmission.”Analysis:The Ignis spacecraftwill follow a concept of operations to ensure a safe disposalwithin5yearsoftheendofthemission.Thespacecraftislaunchedintoacircular500kmaltitudeorbit.Evenifthethrusterisunabletoprovideanyimpulse,thespacecraftwillpassivelyreenterduetoatmosphericdragwithin3.0yearsofthelaunchdate.Anythrustmaneuverswillbeplannedsuchthattheorbitaltitudeisreducedbelow500km,therebyacceleratingthereentrytime.ThisanalysiswasperformedwiththeNASADebrisAssessmentSoftware2.1.1.Figure2showstheoutputdatafromthisanalysis.



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Figure2:IgnisOrbitHistory(500kmpassivedeorbit)

Requirement4.6-2.DisposalforspacestructuresnearGEO:Analysisisnotapplicable.Requirement4.6-3.DisposalforspacestructuresbetweenLEOandGEO:Analysisisnotapplicable.Requirement4.6-4.ReliabilityofPost-missionDisposalOperations:Analysisisnotapplicable.Thesatellitewillreenterpassivelywithoutpostmissiondisposaloperationswithintheallowabletimeframe.



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ODARSection7:AssessmentofSpacecraftReentryHazards:AssessmentofspacecraftcompliancewithRequirement4.7-1:Requirement4.7-1.Limittheriskofhumancasualty:“Thepotentialforhumancasualtyisassumedforanyobjectwithanimpactingkineticenergyinexcessof15joules:a)Foruncontrolledreentry,theriskofhumancasualtyfromsurvivingdebrisshallnotexceed0.0001(1:10,000)(Requirement56626).”SummaryAnalysisResults:DASv2.1.1reportsthatIgnisisCOMPLIANTwiththerequirement.ThemaximumvaluesreportedbytheDASsoftwareare:

• DemiseAltitude=0.0km• DebrisCasualtyArea=0.61m^2• ImpactKineticEnergy=12Joules• RiskofHumanCasualty=1:100,000,000

Thisisexpectedtorepresenttheabsolutemaximumcasualtyrisk,ascalculatedwithDAS'smodelingcapability.Requirements4.7-1b,and4.7-1c:Theserequirementsarenon-applicablerequirementsbecausethespacecraftdoesnotusecontrolledreentry.4.7-1,b):“Forcontrolledreentry,theselectedtrajectoryshallensurethatnosurvivingdebrisimpactwithakineticenergygreaterthan15joulesiscloserthan370kmfromforeignlandmasses,oriswithin50kmfromthecontinentalU.S.,territoriesoftheU.S.,andthepermanenticepackofAntarctica(Requirement56627).”NotapplicabletoIgnis.Thesatellitedoesnotusecontrolledreentry.4.7-1 c): “For controlled reentries, the product of the probability of failure of thereentry burn (from Requirement 4.6-4.b) and the risk of human casualty assuminguncontrolledreentryshallnotexceed0.0001(1:10,000)(Requirement56628).”Notapplicable.Thesatellitedoesnotusecontrolledreentry.ODARSection8:AssessmentforTetherMissionsNotapplicable.TherearenotethersusedontheIgnismission.ENDofODARforIgnis.



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TherawDASreportasfollowsforIgnis:===============EndofRequirement4.3-1===============04052019;17:00:33PM ProcessingRequirement4.3-2:ReturnStatus:Passed=====================NoProjectDataAvailable====================================EndofRequirement4.3-2===============04052019;17:00:37PM Requirement4.4-3:Compliant===============EndofRequirement4.4-3===============04052019;17:03:36PM ProcessingRequirement4.5-1: ReturnStatus:Passed==============RunData==============**INPUT** SpaceStructureName=Ignis SpaceStructureType=Payload PerigeeAltitude=500.000000(km) ApogeeAltitude=500.000000(km) Inclination=45.000000(deg) RAAN=0.000000(deg) ArgumentofPerigee=0.000000(deg) MeanAnomaly=0.000000(deg) FinalArea-To-MassRatio=0.010300(m^2/kg) StartYear=2020.000000(yr) InitialMass=12.000000(kg) FinalMass=12.000000(kg) Duration=1.000000(yr) Station-Kept=False Abandoned=True PMDPerigeeAltitude=-1.000000(km) PMDApogeeAltitude=-1.000000(km) PMDInclination=0.000000(deg) PMDRAAN=0.000000(deg) PMDArgumentofPerigee=0.000000(deg) PMDMeanAnomaly=0.000000(deg)**OUTPUT**



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CollisionProbability=0.000000 ReturnedErrorMessage:NormalProcessing DateRangeErrorMessage:NormalDateRange Status=Pass=============================EndofRequirement4.5-1===============04052019;17:09:05PM Requirement4.5-2:Compliant04052019;17:09:12PM ProcessingRequirement4.6 ReturnStatus:Passed==============ProjectData==============**INPUT** SpaceStructureName=Ignis SpaceStructureType=Payload PerigeeAltitude=500.000000(km) ApogeeAltitude=500.000000(km) Inclination=45.000000(deg) RAAN=0.000000(deg) ArgumentofPerigee=0.000000(deg) MeanAnomaly=0.000000(deg) Area-To-MassRatio=0.010300(m^2/kg) StartYear=2020.000000(yr) InitialMass=12.000000(kg) FinalMass=12.000000(kg) Duration=1.000000(yr) StationKept=False Abandoned=True PMDPerigeeAltitude=491.221054(km) PMDApogeeAltitude=501.948986(km) PMDInclination=44.996381(deg) PMDRAAN=178.806045(deg) PMDArgumentofPerigee=149.519091(deg) PMDMeanAnomaly=0.000000(deg)**OUTPUT** SuggestedPerigeeAltitude=491.221054(km) SuggestedApogeeAltitude=501.948986(km)



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ReturnedErrorMessage=PassesLEOreentryorbitcriteria. ReleasedYear=2023(yr) Requirement=61 ComplianceStatus=Pass=============================EndofRequirement4.6===============04052019;17:24:06PM *********ProcessingRequirement4.7-1 ReturnStatus:Passed***********INPUT****ItemNumber=1name=Ignisquantity=1parent=0materialID=5type=BoxAeroMass=12.000000ThermalMass=12.000000Diameter/Width=0.210000Length=0.320000Height=0.110000name=ThrusterStructurequantity=1parent=1materialID=65type=FlatPlateAeroMass=0.154000ThermalMass=0.154000Diameter/Width=0.140000Length=0.240000name=Tankquantity=1parent=1materialID=59type=CylinderAeroMass=0.400000ThermalMass=0.400000Diameter/Width=0.100000



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Length=0.040000name=ACComponentquantity=1parent=1materialID=-1type=BoxAeroMass=0.050000ThermalMass=0.050000Diameter/Width=0.050000Length=0.050000Height=0.050000name=ZCComponentquantity=1parent=1materialID=-2type=BoxAeroMass=0.015000ThermalMass=0.015000Diameter/Width=0.015000Length=0.015000Height=0.015000name=TANComponentquantity=1parent=1materialID=-3type=BoxAeroMass=0.010000ThermalMass=0.010000Diameter/Width=0.010000Length=0.010000Height=0.007000**************OUTPUT****ItemNumber=1name=IgnisDemiseAltitude=77.999115DebrisCasualtyArea=0.000000ImpactKineticEnergy=0.000000*************************************name=ThrusterStructureDemiseAltitude=0.000000



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DebrisCasualtyArea=0.613564ImpactKineticEnergy=11.517694*************************************name=TankDemiseAltitude=68.662743DebrisCasualtyArea=0.000000ImpactKineticEnergy=0.000000*************************************name=ACComponentDemiseAltitude=0.000000DebrisCasualtyArea=0.422500ImpactKineticEnergy=7.511644*************************************name=ZCComponentDemiseAltitude=0.000000DebrisCasualtyArea=0.378225ImpactKineticEnergy=7.542167*************************************name=TANComponentDemiseAltitude=0.000000DebrisCasualtyArea=0.371148ImpactKineticEnergy=9.459887*************************************===============EndofRequirement4.7-1===============



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AppendixA:AcronymsArgperi ArgumentofPerigeeCDR CriticalDesignReviewCm centimeterCOTS CommercialOff-The-Shelf(items)DAS DebrisAssessmentSoftwareEOM EndOfMissionFRR FlightReadinessReviewGEO GeosynchronousEarthOrbitITAR InternationalTrafficInArmsRegulationsKg kilogramKm kilometerLEO LowEarthOrbitLi-Ion LithiumIonm^2 Meterssquaredml millilitermm millimeterN/A NotApplicable.NET NotEarlierThanODAR OrbitalDebrisAssessmentReportOSMA OfficeofSafetyandMissionAssurancePDR PreliminaryDesignReviewPL PayloadISIPOD ISISCubeSatDeployerPSIa PoundsPerSquareInch,absoluteRAAN RightAscensionoftheAscendingNodeSMA SafetyandMissionAssuranceTi TitaniumYr year


Documents

Astro Digital Ignis Orbital Debris Assessment Report (ODAR