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A Survey of Missions using VASIMRfor
Flexible Space Exploration
April 2010
Prepared by:Andrew V. Ilin, Leonard D. Cassady, Tim W. Glover, Mark D. Carter, andFranklin R. Chang DazAd Astra Rocket Company141 W. Bay Area BlvdWebster, TX 77598
In fulfillment of Task Number 004 of NASA PO Number NNJ10HB38P
Document Number: JSC-65825
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VASIMR
For Flexible Space ExplorationTableofContents
Section1. FlexibleMissionStrategies .......................................................................................................1
Section2. MissionStudyAnalysisTools .................................................................................................... 2
2.1 Copernicus.....................................................................................................................................2
2.2 AdAstra3DTraj .................................................................................................................. ............. 2
2.3 Chebytop ....................................................................................................................................... 2
2.4 OptiMars ........................................................................................................................................ 2
2.5 HOT ................................................................................................................................................ 2
Section3. TypicalParameterAssumptions............................................................................................... 3
Section4. PrepositioningtotheEdgeofEarthsSphereofInfluence ...................................................... 3
4.1 DepartingLEOwithInclinationChange......................................................................................... 3
4.2 LunarTug .......................................................................................................................................4
Section5. RoboticMissionsBeyondEarthsSphereofInfluence.............................................................7
5.1 CargoDeliverytoMars .................................................................................................................. 7
5.2 MarsSampleReturn ...................................................................................................................... 8
5.3 EnhancingSolarPoweredCapabilitiestoreachJupiter .............................................................. 12
Section6. FastHumanMissionstoMarswithVariableSpecificImpulse...............................................16
6.1 HumanMissiontoMarsusing12MW ........................................................................................ 16
6.2 HumanMissionScenariostoMarsusing200MW......................................................................19
6.2.1 TechnologyRequirementEstimatesfromOptiMars..............................................................19
6.2.2 OptimizedResultsfromCopernicus.......................................................................................21
Section7. SummaryandFutureWork ....................................................................................................23
Section8. References..............................................................................................................................23
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FIGURESFigure1: ExampleEarthdepartureandplanechangetrajectory,inthiscasea200kWVASIMR
propelledspacecraftweighing4000kgatLEO.3
Figure2: VASIMRTransferfromLEOtoLLO..4
Figure3: CopernicusSimulationsofVASIMRTransferfromLEOtoLLO.5
Figure4: LunarTugmissiontrajectoryintheEarthMoonRotatingVisualizationframe..6
Figure5: OptiMarsSimulationofheliocentrictransferfromEarthSOItoMarsorbit...7
Figure6: VariableIspprofileforOptiMarsheliocentrictransfersimulation.8
Figure7: ChangingMinimalIspandPayloadMassasafunctionofTripTimeforCargoMissiontoMars.8
Figure8: VASIMRMSRScenario:EarthtoMars9
Figure9: VASIMRMSRScenario: MarstoEarth..9
Figure10: PowerScanforMSRMission. Foreachpowervalue,onewaytriptimes(andpropellant
mass)wereminimized............................................................................................................................10
Figure11: FromLEOtoMars:250kW.11
Figure12: FromMarstoLEO:250kW.12
Figure13: VASIMRCatapult:StartingPointforMassRelations.13
Figure14: JupiterCatapultmissiontrajectoryforIsp=5,000s.....14
Figure15: HumanpilotedheliocentrictransferwithIspvstimeplot(right)..17
Figure16: CTVarrivalatMarsandsubsequentcapture131dayslater(left)and7dayspiralmaneuver
intolowMarsorbit(right).18
Figure17: CTVspiraldeparturefromMarsandreturnheliocentrictrajectory.18
Figure18: OptiMarssimulationof31dayheliocentrictransferfromEarthtoMarsorbits....20
Figure19: Profileofvariablespecificimpulsefor31dayheliocentrictransferbetweenEarthandMars
orbitssecondstageofaveryfasthumanmissiontoMars...20
Figure20: Copernicussimulationof31dayheliocentrictransferfromEarthSOItoMars.............21
Figure21: Copernicussimulationofvariablespecificimpulseprofilefor31dayheliocentrictransfer
fromEarthSOItoMars...22
Figure22: ParametricstudyofhumanmissiontoMarstriptimedepartingfromL1withminimalIsp=
3,000sec.22
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NOMENCLATURE
A ThrustVector
an Normal(azimuthal)componentof
accelerationinOptimars
ar Radialcomponentofacceleration
AARC AdAstraRocketCompany
CDV CargoDeliveryVehicle
CTV CrewTransferVehicle
DRM DesignReferenceMission
EELV EvolvedExpendableLaunchVehicle
EEV EarthEntryVehicle
ERV EarthReturnVehicle
GNC Guidance,NavigationandControl
GUI GraphicUserInterface
HOT HybridOptimizationTechnique
IMLEO InitialMassatLEO[kg]
Isp SpecificImpulse
LEO LowEarthOrbit
LLO LowLunarOrbit
LMO LowMarsOrbit
ML MarsLander
MP PropellantMass[kg]
MPL PayloadMass[kg]
MPT PropellantTankMass[kg]
MSA SolarArrayMass[kg]
MSR MarsSampleReturn
MT MassofThruster[kg]
OTV OrbitalTransferVehicle
P Power[kW]
R RadiusVector
SOI SphereofInfluence
V VelocityVector
VASIMR VariableSpecificImpulse
MagnetoplasmaRocket
VX200 VASIMRlabexperimentat200kW
TotalSpecificMass[kg/kW]
SA SpecificMassofSolarArrays[kg/kW]
T SpecificMassofThruster[kg/kW]
PowerEfficiency
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Section1. FlexibleMissionStrategiesSpace
exploration
can
greatly
benefit
from
the
high
power
electric
propulsion
capabilities
the
Variable
Specific Impulse Magnetoplasma Rocket (VASIMR) provides. When combined with chemical rocket
technologies in a flexible architecture, the VASIMR allows new and dramatically improved mission
scenariostobeconsidered. Employingexistingstateoftheartsolarcelltechnology,VASIMRisableto
achieve dramatic propellant mass savings to move payloads near Earth and preposition payloads for
assemblynear themoon, theedgeofEarthsgravitationalsphereof influence,andbeyond. Robotic
prepositioningofassetsatkey locations inspaceallowscostand risk tobe reduced for later transits
between staging locations. The possibility of multimegawatt power levels also allows VASIMR
technologytosignificantlyreducethetraveltimeand improveabortoptionsforhuman interplanetary
missions between staging locations near the Moon and Mars. Power levels ranging from currently
availablesolartechnologiestothoserequiringthefuturedevelopmentofnuclearpoweredsystemsare
considered.
InSection2of thisreport,wedescribe thevariousstrengths, limitations,andassumptionsofmission
softwaretoolsusedforthisstudy. ThecapabilitiesenabledbyVASIMRtechnologyarethenexamined
usingapiecewiseapproachbuiltonthreemaneuvers:transferfromlowEarthorbittostaginglocations
innearlunarorbit,transferfromnearlunarorbittomoredistantobjects(includingrobotictransitsto
Marsortheouterplanets),andfinallyhumanmissionstoMars.
InSection 3, we give the parameters andassumptions typicallyused for these studies, unless stated
otherwiseinthespecificstudy.
InSection4wedescribemissioncapabilitiesneartheEarthandtheMoonusingaVASIMRwithrealistic
mass and performance values based on results from the VX200, a VASIMR laboratory experimentoperatingat200kW,and theVASIMR flightdesign,VF200,operatingat200kW.ThesenearEarth
missionsincludecosteffectivecargotransferfromLowEarthOrbit(LEO)toLowLunarOrbit(LLO)and
cargoprepositioningneartheedgeofEarthsgravitationalsphereofinfluence(SOI).
InSection5,weconsiderorbit transfers for solarpowered roboticmissions that start froma staging
areanearEarthsSOI todelivercargo toMars, returna sample fromMars,andcatapult payloads to
more distant destinations, particularly Jupiter. These studies to more distant objects help identify
possibleprepositioningstrategiesinsupportofmorecomplexmissions.
InSection6,wediscusstechnologydevelopmentsneededtosupporthighpowerfasthumanmissions
beginningfromLEOorprepositionedstagingareasnearEarthsSOIandendinginorbitaroundMars.
InSection7,wegiveabriefsummaryandsuggestscenariosthatwarrantmoredetailedstudyalongwith
basictechnologyrequirementsandfutureneedsforthesemissions.
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Section2. MissionStudyAnalysisToolsAdAstraRocketCompanyemploysseveralsoftwaretoolsforsimulatingtheVASIMRmissions. When
consideringelectricpropulsionsystems,thefundamentalequationsofmotionmustbeexaminedwith
care taken to avoid implicit assumptions commonly applied for chemical propulsion systems. For
example, the power and propellant mass flow are somewhat independent of one another, and thespecific impulsecanbechangedduringamaneuver.For thesemissionstudies,relativelysimple tools
areusedfirsttoidentifymissionssuitableforprogressivelyhigherfidelityanalysis. Then,moredetailed
surveysareperformedwithanalytictoolsandvariouscodesincludingAdAstra3DTraj,CHEBYTOP[2], and
OptiMars[4]. Wherewarranted,stillhigherfidelityanalysiscanthenbeperformedusingHOT[7],[8]and/or
Copernicus[1]. Inthissectionwegiveabriefdescriptionofthesemissionanalysistools.
2.1 CopernicusCopernicus[1] is a generalized spacecraft trajectory design and optimization system developed by the
UniversityofTexasatAustin. ThissoftwareisreleasedtoNASAcentersandaffiliates. Itissuppliedwith
a complex GUI (Graphic User Interface), and includes variable Isp (Specific Impulse) capability.
Copernicusisannbodytoolwithahighdegreeofflexibility.Theusercanmodelanumberofdifferent
missions, with varying gravitational bodies, objective functions, optimization variables, constraint
options,andlevelsoffidelity. Additionally,itcanmodelmultiplespacecraft,aswellasoptimizeforboth
constant and variable specific impulse trajectories. Copernicus employs multiple shooting and direct
integrationmethodsfortargetingandstatepropagation.
2.2 AdAstra3DTrajThe Fortran code AdAstra3DTraj was written in Ad Astra Rocket Company (AARC) for a direct 3D
trajectorysimulation. Itemploysvariousnumbersofgravitationalbodiesandcustommadenavigation
strategies, including variable specific impulse. It also allows for simple parametric scans and limited
optimization.
2.3 Chebytop
CHEBYTOP[2]isananalysistoolthatoptimizesonewaytrajectoriesbetweenplanetarybodies. Itisused
as a preliminary design tool for missions using electric propulsion with constant specific impulse.
CHEBYTOPusesChebychevpolynomialstorepresentstatevariables,whicharethendifferentiatedand
integrated in closed form to solve a variablethrust trajectory. This solution can then be used to
approximate a constant thrust trajectory. While it is considered a lowfidelity program, it is highly
valuedforitsabilitytorapidlyassesslargetradespaces.ItiswrittenonVisualBasicmacroswithExcel
GUIandgraphics.
2.4 OptiMarsOptiMars[4]
is a variable Isp Earth Mars transfer optimizer of very low fidelity. The software was
developed in 2000 2002 at University of Maryland. It assumes polynomial expression for the
acceleration. Thepositionsof theplanetsarenotconsidered,so transfer isoptimizedbetweenEarth
andMarsorbits.
2.5 HOTTheHOT(HybridOptimizationTechnique)software[7],
[8]isaFortrancodewrittenatNASAJSC. Itusesa
numerical optimization method based on the calculus of variations for minimizing a performance
function describing a mission trajectory with variable specific impulse. HOT simulates interplanetary
maneuversbyintegratingequationsofmotionandequationsforLagrangemultipliers.
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Section3. TypicalParameterAssumptionsThefollowingparametersandassumptionsaretypicallyusedthroughoutthispaperunlessspecifically
statedotherwise. ThemassbudgetcanbepresentedasM0=MPL+MP+MPT+MSA+MT,whereMPLis
payloadmass,MP ispropellantmass,MPT ispropellant tankmass (assumed tobe0.1MP),MSA= SA
max(P) isthemassofsolararrays ,andMT= TP isthemassofVASIMRthrusters,whereSA isthemasstopowerratioforthesolararraysinkg/kW,andsimilarly,T isthemasstopowerratioforthe
thrusterpackageincludingallpowerhandlingequipment.
FormissionsinsidetheEarthsgravitationalsphereof influence(SOI),weconsiderVASIMRpower(P)
levelsrangingfrom100500kW. Thenominalparametersforthesemissionsareaspecificimpulse,Isp,
of5,000swithatotalmasstopowerratio, = SA+ T,of10kg/kW.
Forroboticorcargointerplanetarymissions,weconsiderVASIMRpowerlevelsrangingfrom1 5MW.
Thenominalparameters for thesemissionsareaspecific impulse, Isp,of4,000or5,000switha total
powerefficiency,,of60%,andamasstopowerratio, (total),of4kg/kW.
Forhumaninterplanetarymissions,weconsiderVASIMRpowerlevelsrangingfrom10 200MW. The
nominalparametersforthesemissionsareavariablespecificimpulse,Isp,from3,000to30,000switha
total power efficiency,, of 60%, and a masstopower ratio, (total), less than 4 kg/kW. A more
accurate VASIMR model, considering the power efficiency to be a function of specific impulse and
power,isbeyondthescopeofthesestudies.
Section4. PrepositioningtotheEdgeofEarthsSphereofInfluenceA VASIMR powered transfer from LEO to the Earths Sphere of Influence (SOI) is different than a
chemicallypropelledEarthdeparturebecauseofthelowthrustnatureofanelectricpropulsionsystem.
The low,butsteadythrustonthespacecraft leadstoaspiralorbitwith increasingradiustoreachthe
SOI.
4.1 DepartingLEOwithInclinationChange
Figure 1 shows an example of a
lowthrust Earth departure spiral
(inred)followedbya51.6degree
plane change (in green) using a
200 kW VASIMR system
propellinga4000kgIMLEO(Initial
Mass at LEO) payload. The
AdAstra3DTraj code was used for
the direct simulation without full
optimization. Theplanechange is
executedwhentheorbitalvelocity
is the lowest, at the end of the
spiral, to minimize the propellant
requiredforthemaneuver.
Figure1: ExampleEarthdepartureandplanechangetrajectory,inthiscasea
200kWVASIMRpropelledspacecraftweighing4000kgatLEO.
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ThemainadvantageofthistypeofEarthdeparturestage isthefuel,andhence,masssavingsforthe
propulsionsystem. Intheexamplescenario,lessthan600kgofargonpropellantisusedtoputa4,000
kgspacecraftonanEarthescape trajectory. Thecontinuousoperationof theVASIMR thrusteralso
allowsanoptimizedplanechangemaneuveratanyaltitude. This inturnallowsformoreflexibility in
the initial launch locationand results ina less severepayload reduction if the spacecraft is launched
from a high latitude location on Earth. The slow spiraling nature of the VASIMR for trajectories
departingtheEarthalsoenablesmissionoperatorstoperformvitalspacecraftsystemandhealthchecks
forweeksduringtheinitialspiral. Ifaproblemisdiscovered,ahighvaluespacecraftcouldbeputintoa
parkingorbitforlaterevaluation,docking,return,orrepair.
4.2 LunarTugLarge quantities of equipment and material must be transported from the Earth to the Moon for
ambitiousprogramsoflunarexploration. Iftransittimesofseveralmonthsareacceptableforcargo,a
reusableVASIMRtugallowsroughlytwiceasmuchpayloadtobedeliveredtotheMoonasachemical
system would require for the same initial mass in low Earth orbit (IMLEO). This implies cutting the
numberofheavylift launchesandtheirassociatedcostinhalfandamortizingthecostofthetugover
themanyyearsthatcargoisdeliveredbetweenLEOanddestinationsneartheMoon.
Fi ure2: VASIMRTransferfromLEOtoLLO
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Thefollowingmissionassumptionswereusedforthesimulation: inputpowerprovidedbysolarpanel
cellsP=500kW,VASIMRpowerefficiency=60%,Isp=5,000sec,IMLEO=25,200kg(includingboth
OTVandCDV).
Figure2demonstratesapossibleLunarTugarchitecture. ItwasassumedthatanEvolvedExpendable
Launch Vehicle (EELV) delivers a large Cargo Delivery Vehicle (CDV) into 500 km orbit, followed by
rendezvouswithaVASIMRpoweredOrbitalTransferVehicle(OTV)orTug. TheOTVtransfersCDV
betweena500kmlowEarthorbit(LEO)anda100kmlowlunarorbit(LLO)andreturnstoLEO. Forlow
thrustspiraltrajectorieswithoutaplanechange,theV isthedifferencebetweenthe initialandfinal
circularorbitalvelocities[10],whichis8km/sforthismission,includingboththespiralsaroundtheEarth
and the Moon. From the rocket equation, given in Figure 2, the VASIMR could deliver the 14 mT
payloadwithin6months,using3.8mTofpropellant. Forcomparison,achemicalpropulsionsystem
withspecific impulsearound450scanonlydeliver5.7mTpayloadforthesameinitialmassplaced in
LEO.
Figure3showsthetrajectory,calculatedbytheCopernicuscode,intheEarthframeofreference. Thetriptimeis178days,thepropellantusedis3,840kg,andbothnumbersareveryclosetotheestimates.
Figure4showsthesamemissiontrajectoryintherotatingEarthMoonframeofreference,inorderto
show both spiraling from LEO anddespiraling to LLO. For this case, the inclination changewas not
required. Copernicususesfourconsecutivesegmentsforthetrajectory;eachofthemmatchedattheir
connectingendpointswithrespecttoposition,velocity,massandtime. Thefirstandlongestsegmentis
spiralingfromLEOwith the thrustvectordirectedalong thevelocityvector. Thedurationofthefirst
Figure3: CopernicusSimulationsofVASIMRTransferfromLEOtoLLO
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segment isoneof the optimizedvariablesand calculated to be 140days and it requires 3,020 kgof
propellant. TheendofthesecondsegmentistargetedtothepointontheboundaryoftheMoonsSOI
(relative to theEarth) closest to Earth. It lasts 3daysand requires75 kgof propellant. Copernicus
optimizes thedirectionof the thrustvector in the secondand thirdphases,aswellas the transition
timesbetweenallphases.Thethirdsegment,capturingthepayload inorbitaroundtheMoon, lasts5
days and requires 105 kg of propellant. A fourth, optional segment performs despiralingmaneuver
withthrustdirectedopposedtothevelocityinpreparationforlandingontheMoonbychemicalmeans.
CopernicuscalculatesthefourthandthirdsegmentsbackwardfromLLOandconnectstheendsofthe
second and third segments iteratively. This despiraling maneuver requires 30 days and 640 kg of
propellant.
Figure4: LunarTugmissiontrajectoryintheEarthMoonRotatingVisualizationframe
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Section5. RoboticMissionsBeyondEarthsSphereofInfluence5.1 CargoDeliverytoMarsAparametricstudyofcargodelivery toMarswith2MWofpoweranda totalspecificmassof =4kg/kW(power+VASIMR)wasperformedusingtheOptiMarscodewiththevariableIspoptimizer,for
anIMLEO=20mT. Theparametervariedwastheminimumallowedspecificimpulse.Thefirstsegment
ofthemissionisthespiralingfromLEOwithaninitialaltitudeof1,000kmtotheEarthsSOI. ForIsp=
4,000sec,thespiralingtakes27daysandrequires3.7mTofpropellant. Figure5demonstratesthe80
dayheliocentrictransferfromEarthtoMarsorbit. Thearrivalspeedwasconstrainedtobe6km/sec,
assuming that aerocapture will be used for payload delivery to the Mars surface. The time varying
profileforthespecificimpulsedeterminedbyoptimizationisshowninFigure6. Amaximumtechnology
limitof30,000swasusedforthespecificimpulseoftheVASIMRengine. Thecoastingtimeisabout10
days. FortheoptimizedprofileofthespecificimpulsewithaminimumIspof4,200sec,theheliocentric
transferrequires3mTofpropellant. So,fromIMLEO=20mT,thetotalmassbudgetcanbepresented
as8mTforpowerandpropulsionplus7mTofpropellantand5mTofpayload. Thetotaldurationof
themissionisabout3.5months.
Figure5: OptiMarsSimulationofheliotransferfromEarthSOItoMarsorbit
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TheresultsoftheparametricstudyareshownintheFigure7forthesameIMLEOinallcasesandvarying
minimumIsp. IftheminimumIspis
increasedfrom4,200secto6,200
sec,thetotaltriptimegoesupby
14 days but the amount
propellantrequiredgoesdownby
1.7 mT, so more payload can be
delivered.
5.2 MarsSampleReturnThe International Mars
Architecture for the Return of
Samples (iMARS) Working Group
has conducted extensive studies
on Mars Sample Return (MSR)
missions
using
chemical
propulsiontechnology [11]. Inthis
section,weinvestigatetheuseof
VASIMR technology to
accomplishaMSRmission that is
a modified version of the one
proposed by the iMARS Working
Group.
This analysis is based on
reasonablevalues for thespecific
massforpowerlevelsintherange
between 100 and 500 kW. This
calculation was performed using
Chebytopassuminganinitialmass
in LEO, IMLEO = 19,000 kg,
including 4,880 kg of payload
fixed, leaving 14,120 kg for the
OTVandpropellant. Powerreceivedfromsolarpanelwasassumedtodependondistancetothesunas
~1/R2. Figure8illustratesthevariousstagesofthemission:
1)
An
Atlas
V
552
puts
19,000
kg
into
LEO
at
an
altitude
of
500
km.
The
spacecraft
includes
an
OTV
with
aVASIMRthruster,a4,000kgMarsLander(ML),andan880kgEarthEntryVehicle(EEV). Themasses
forMLandtheEEVaretakenfromtheMSRmissiondevelopedbytheiMARSWorkingGroup.
2)ThespacecraftspiralsupfromLEOtotheEarthsSOI.
3)AheliocentrictransfertoMarsisperformed.
Figure7: ChangingMinimalIspandPayloadMassasafunctionofTripTimefor
CargoMissiontoMars
Figure6: VariableIspprofileforOptiMarsheliocentrictransfersimulation.The
variablesarandanareradialthenormalcomponentsofacceleration,respectively,
relativetotheSun.
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4)The4,000kgLanderisejectedjustbeforethespacecraftbeginsitsMarsorbitinsertion.TheLander
followsadirectdescentpath,justasitwouldinthechemicalmission.
5)TheOTVspiralsdowntoalowMartianOrbit(LMO)withanaltitudeof500km.
3. Heliocentric transfer to Mars
4. Lander separatesto make directlanding
5. OTV spirals downto 500 km LMO
Earth SOI
925,000 km radius Mars SOI
577,000 km radius
2. OTV spirals upfrom LEO toEarths SOI
1. Atlas V 552 puts 19,000 kg into500 km LEO; OTV carrying 4000kg Mars Lander and 880 kgEarth Entry Vehicle
Figure8: VASIMRMSRScenario:EarthtoMars
Figure9showsthe30kgsamplereturnwiththefollowingphases:
67)Thesamplecontainer(30kg)isreturnedtoLMO,whereitrendezvouswiththeOTV.
8) ThecombinedpackagespiralsuptoreachtheMarsSOI.
9) AheliocentrictransferreturnstheOTVandsamplebacktotheEarth.
1011)The910kgpackage,includingtheEEVtolandthesampleontheEarthisejectedfromtheOTVasitbegins itsEarthOrbit Insertion,sothatthesamplecontainer landsontheEarthandtheOTVspirals
downtoLowEarthOrbit.
Earth SOI Mars SOI6. Mars Ascent Vehicle launches samplecontainer to 500 km Mars orbit
7. OTV finds and captures sample container
8. OTV spirals up to Mars SOI
9. Heliocentric transfer to Earth
10. Earth Entry Vehicle separatesfrom OTV for direct landing
11. OTV spirals down to LEO from Earth SOI
Figure9: VASIMRMSRScenario: MarstoEarth
Figure10summarizeskeyresultsforthistypeofmissionwithdifferentpowerlevels.
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Figure10: PowerScanforMSRMission. Foreachpowervalue,onewaytriptimes(andpropellantmass)wereminimized.
Thepowerat1AUwasvariedbetween100kWand500kW.Theonewaytriptimewasminimizedfor
eachpower level. Theresultsoftheparametricscanaredemonstrated inFigure10. Theparametric
scanusedthefollowingvaluesofinputparameters:
Totalpowerefficiency=60%
Specificimpulse Isp=5,000sec,
IMLEO=19,000kg,
LEOaltitudeRLEO=500km,
MarsaltitudeRLMO=500km,
MinimalstaytimeonLMOis4months.
LanderMassML=4mT
PropellantTankmassMPT=0.1MP
InitialmassforLMOLEOsegment=finalmassofLEOLMOsegmentMLMPT
Insummary,aMarssamplereturnmissionispossibleatanypowerlevelbetween100kWand500kW
(at1AU). AsFigure10shows,theroundtriptimedecreaseswith increasingpowerwithasubstantial
decrease from6.1yearsat200kW to3.7yearsat250kWanda relatively slowdecrease forpower
levelsbetween250kWand500kW(downto3.1years). Thereasonforthebigdecrease intriptime
just less than250kW is that theLMOstay time isnotmonotonic functionofpower. For lowpower
levels(100200kW)theLMOstaytimeincreasesfrom0.4yearto2.1years,butfor250kWofpower,
thestaytimecanbereducedto140daysduetotheeffectoftherelativepositionofEarthandMars.
Theconclusion is that theoptimal levelofVASIMRpower for theMSRmission is250kW (at1AU),
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requiringaround triptimeof3.7years, including140days in lowMartianorbit. Themissioncanbe
performedforspecifictotalpower,
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-1.5
-1
-0.5
0
0.5
1
1.5
-1.5 -1 -0.5 0 0.5 1
Mass balance [kg]:Mass at LMO 9220Propellant(LMO-LEO) 2740Tank 280EEV with Mars sample 910
_________________________Mass at LEO 5390Total Alpha 21
Decelerating: 122 days
Coasting: 87 days
Accelerating: 61 days
Helio Transfer Departure: 1 July 2022
Helio Transfer Arrival: 28 Mar 2023
Power 250 kW (at Earth orbit)
stage Trip time PropellantSpiraling 87 days 493 kgHeliotrans 270 days 1361 kgDespiral 82 days 886 kgTotal 439 days 2740 kg
Departure 5 Apr 2022 (LMO)Arrival 18 Jun 2023 (LEO)
Figure12: FromMarstoLEO:250kW
5.3 EnhancingSolarPoweredCapabilitiestoreachJupiterInthissection,weexploretheconceptofareusableVASIMRspacecrafttocatapult5,000kgrobotic
payloadstoJupiterusingaHohmannliketransfer. Theintentofthisstudyistoidentifytheimportant
parameters for ejecting the payload and returning the VASIMR system back to Earth for additional
payloads inreasonableperiodsof time. Theoperationalparametersof interestare thepower level,
propellantmass,payloadreleasepoint,anddistanceofclosestapproachtotheSuntogainadditional
solarpower. Optimizationbasedonvariablespecific impulse isneeded to fullyexplore thisconcept.
However, two calculations at fixed specific impulse within the range of VASIMR capabilities, one
assuming 5,000 s and another assuming 4,000 s, demonstrate the advantages of variable specific
impulseandindicatethedirectionforpossiblefuturestudies.
ThemissionisbasedontheassumptionthatthecatapultspacecraftanditspayloadbeginattheEarths
sphere of influence (SOI), coasting in the Earths orbit about the Sun. The VASIMR engines power
ratingmustmatchthepeakpoweravailablewhenthespacecraftisclosesttotheSun. Thesolararrayisassumedtobeaplanararrayratherthanaconcentrator,tosimplifyitsoperationneartheSunwherea
concentratormightotherwiseoverheat thephotovoltaiccells. For this study,nogravityassist from
othercelestialbodiesareconsidered, inordertoallowforflexible launchwindowsdependingonlyon
Jupiter.
TheMESSENGERspacecraft[6]providesastartingpointforestimatingthemassofvarioussubsystemsof
theVASIMRCatapultwhichwillmakerepeatedflightsfromlowEarthorbittotheinnersolarsystem.
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MESSENGERisexpectedtohavealifetimeofapproximately10yearsinthistypeofenvironment. The
VASIMR Catapult will have similar requirements for communications and Guidance, Navigation and
Control(GNC). ThemassbudgetoftheCatapultspacecraftisshowninFigure13.
dockingpropulsion
(arcjet)
Argon mass depends on trajectory expect several thousand kg
Assume 4,000 k
launched to
rendezvousand dock withejector in LEO
Argonpropellant
payload
ejector
solar array
VASIMR engines
avionics
thermal
communications
GNC
structure
harness
T = 5 kg/kW; thrusterpower = power at closest approach to Sun
A = 18 kg/kW; P = power@ 1 AU; range 200 kW to 1 MW
100 kg
10% of sum of other components of the ejector (dominated by
VASIMR engines and solar array)
400 k
30 kg
35 kg
100 kg
based on MESSENGER spacecraft total:700 kg (this quantity is fixed)
35 k
Figure13: VASIMRCatapult:StartingPointforMassRelations
The mission trajectory was studied using the AdAstra3DTraj code and considered only the Suns
gravitationalfield. ThemissionbeginswithadecelerationphasefromtheEarthSOIwithaninitialmass
there (IMSOI) of 22mT where the solar panels provide power, PEarth = 500 kW, with a propulsion
efficiencyoftheVASIMRof=60%. Twooptionsforconstantspecificimpulsewereconsidered,one
with Isp=4,000sandanotherwith Isp=5,000sec. Several iterationswereperformed to findout the
optimal distance for switching from deceleration to acceleration, and then a return of the Catapult
spacecrafttoEarthsSOI.
During the acceleration phase, the orbital elements of the vehicles trajectory are continuously
changing. In particular, the aphelion distance grows larger. As soon as the instantaneous aphelion
matches
the
semi
major
axis
of
Jupiters
orbit,
the
robotic
payload
should
be
released
on
an
orbit
that
willcoasttoJupiter. Afterthepayloadisreleased,theCatapultsthrustvectoristhenchangedtobegin
itsrendezvouswithEarth.
Whentheenergyneededforejectionisreached,thepayloadwithamassofMPL=4mTandtheempty
propellanttankwithanassumedmassofMT=0.1MPisreleasedfromtheVASIMRCatapult. Whenthe
payload is released, the Catapult immediately directs its thrust opposite to its velocity vector to
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decelerate until the instantaneous aphelion distance is reduced to approximately 1 AU, thereby
initiatingitsrendezvouswiththeEarthforreuse.
AnoptimizedJupitercatapulttrajectoryisfirstconsidered,asshowninFigure14,forIsp=5,000sec.
Belowarethemissiondetails,whichwerecalculatedwiththeAdAstra3DTrajsimulations.
Figure14: JupiterCatapultmissiontrajectoryforIsp=5000sec
Phase1, shown in red, is adeceleration phase to begina close solar passwith the thrustvector, A,
directedperpendiculartotheradiusvector,R. Observationsofseveralsimulations indicatedthatthis
wasthebeststrategytogetclosetotheSunusingaminimalamountofpropellant. Theendtimefor
thefirstphasewasalsooptimizedbytrialanderrortoaccomplishthemissionwithminimalpropellant.
Phase1endsatadistancefromtheSunofRend=0.66AUafterTend=132.7daysusingapropellantmass
of3.6mT.
Phase2,shown ingreen, isanaccelerationphasenear theSunwithhighsolarpowerand the thrust
vectorparalleltothespacecraftvelocity. TheclosestapproachtotheSunisRmin=0.452AUduringthis
phase. Phase2endswhenthespecificenergyreachesthelevelneededforthepayloadtogettoJupiter
atRend =0.581AU, Tend=181.8days, requiring an additional 3.5 mT of propellant, so that the total
propellantusageinphases1and2is7.1mT.
Phase3a,showninred,involvesonlythereleasedofthe4mTpayload,whichcoaststherestoftheway
toJupiterat5.2AUafter1,036daysoftotaltimefromdepartingEarthssphereofinfluence.
Phase3b,alsoshowninred,describesthedecelerationphaseofthe0.8mTVASIMRCatapultsoitcan
begin itsreturntonearEarth. Thethrustwasdirectednearlyopposite to thevelocityvectorwithan
anglewhichwasoptimizediteratively. Theoptimalvalueangle(A,V)=165o,minimizespropellantusefor
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thereturntotheEarthsSOI. ThephaseendswhenthespecificenergyreachesthatofanEarthorbitat
distanceR=1.179AU,after231.4days,usinganadditional1.5mTofpropellant. Theremainingphases
are then performed with thrust directed orthogonally to velocity, so the specific energy does not
change.
Phase3c,showninlightblue,isamaneuvertoreturnthecatapulttotheEarthwiththethrustvector,A,
orthogonaltothevelocityvector,V,andtowardtheSun,endingat1.452AU,afteratotalof279.2days
usinganadditional0.6mTorpropellant.
Phase3d,showningray,isasimplecoastingphaseto1.478AUafteratotalof359.2days.
Phase3e,showninmagenta,isamaneuverturningawayfromtheSunwiththrustvector,A,orthogonal
tothevelocityvector,V,oppositetheSuntorendezvousat1AUafteratotal509.5days,usingan
additional2.4mTofpropellant. Thecumulativepropellantusageis11.5mT.
Toillustratetheeffectofspecificimpulseonthetechnologyrequirementsforthespecificmass,thecatapultmissionwasrepeatedforalowerspecificimpulsewithIsp=4000susingthesamemaneuvers.
TheresultsfromtheAdAstra3DTrajsimulationsat4000sare:
Phase1endsatRend=0.9AU,after80.7days,using2.9mTofpropellant.
Phase2endsatRend=0.729AU,withaclosestapproachtotheSunat0.657AU,afteratotalof157.4
days,usinganadditional4.6mTofpropellant.
Phase3adeliversthepayloadtoRend=5.2AU,afteratotalof1,057days.
Phase3bwithangle(A,V)=150oendsatRend=1.156AU,afteratotalof210.8days,using2.1mTof
propellant.
Phase3cendsatRend=1.27AU,after240.3days,using0.7mTofpropellant.
Phase3dcoaststoanendingradiusatRend=1.121AU,after363.3days.
Phase3eendsatRend=1AU,afteratotalof406.8days,requiringanadditional1.3mTofpropellant,
suchthattheentiremissionuses11.6mTofpropellant.
Oneadvantageof lowering the specific impulse from5,000s to4,000 s isa relaxation in technology
requirementsfrom =6.8kg/kWfor5,000secondsto =8.4kg/kWfor4,000seconds. However,the
efficiencyoftheVASIMRsystemmaybemorechallengingforlowerspecificimpulse. Inaddition,the
massofthepropellanttoreturntoEarth issignificant,somissionswithoutareturntoearthcanhave
larger,ormultiplepayloads,offeringoptionsforfuturestudy.
Theprimaryresult from these twostudiesatdifferent fixedspecific impulse is that the initialphases,
neededtoreleasethepayloadon itswaytoJupiter,canbeperformedwith7.1mTusing5,000s,but
require7.5 mTat4,000s. Alternatively, the returnof thecatapult spacecraft to theEarth for future
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reusecanbeperformedwith4.1mTofpropellantusingaspecificimpulseof4,000swhereasthereturn
phasesrequire4.5mTusingaspecificimpulseof5,000s. Clearly,varyingthespecificimpulsebetween
theoutboundandreturnmaneuverscanresultissignificantsavings,warrantingfurtherstudy.
Section6. FastHumanMissionstoMarswithVariableSpecificImpulseIn thissectionweassume thatveryhighpowersourcesareavailable,suchasmightbepossiblewith
advancednuclearpowertechnology. Withamplepowerforinterplanetarymissions,theVASIMRcan
significantlyreduce launchmassand trip timescomparedwithchemical thruster technology. At the
200 megaWatt power level,human missions to Mars in less than39 days become conceivable with
advancedVASIMRandpowertechnologies. TriptimestoMarsonthistimescalewerementionedasa
laudablegoal forNASAby itsadministrator,CharlesBolden[9], in July2009,because theydramatically
improveacrewssurvivabilityandsafety inthe interplanetaryspaceenvironment. Inthissection,we
presentconceptualstudiesshowingthedependenceofthetriptimeonthe initialdepartureposition,initialmassatthedepartureposition,massofthepayload,variablespecific impulse,andthemassto
powerratioforthepowerandpropulsionsystems. Thereturnmissionisnotpresentedinthisreport.
6.1 HumanMissiontoMarsusing12MWThevariablespecificimpulsemadepossiblebyVASIMRtechnologiescanbefullyappreciatedformulti
megawattinterplanetarymissions. ThesestudieswereconductedusingtheHOT[7],[8]
softwarepackage
anddemonstratethepossibilityof1218MWmissionsarrivingatMarswithin34months,abouthalf
the time estimated in the Design Reference Mission (DRM) 3.0[12], which utilizes Nuclear Thermal
RocketsfortransferringtoMars.Themissionanalysisassumedtheexactsamepayloadmassdelivered
totheatmosphereofMars(60.8mT)atthesamevelocityforaerocapture(6.8km/s)asintheDRM3.0
for the crewed portion mission with VASIMR engines. The crew time in space could be reduced by
aboutonemonthbyrendezvousingthecrewwiththeVASIMRpoweredspacecraftinhighearthorbit
afterthemainvehiclehadpassedthroughtheVanAllenbelts.
ForthismissioncargomustbepredeployedbothinMartianorbitandonthesurface. IntheDRM,one
rockettransfersthecargolanderandanothertransfersanearthreturnvehiclewithalivinghabitatfor
thecrewtoMars. Thetotalmassofthosepayloadsis91.4mTasoutlinedintheDRM3.0(includingan
aeroshell for thecargo landersizedappropriately forentryvelocity). TheVASIMRmissionproposes
combiningbothpayloadsplusthereturnpropellantononevehicletosavepropellantontheoutbound
journey. AVASIMRsystemwithonethirdthepower,4MW,comparedtothe12MWofthecrewed
vehicle, transfers the payload more slowly, requiring a 154 day spiral plus a 288 day heliocentric
transfer,butusesonethird less initialmasswithan IMLEOof202mT,comparedwith282mTforthe
DRM.ThereturnhabitisstoredinorbitatMars,asintheDRM3.0,waitingfortheMarsascentvehicle
andthecrewforthereturntrip.
ThecrewedspacecraftdepartsLEOonMay6,2018.Theship initialmass inLEO is188mT.12MWof
electricpower isassumedforthedurationofthetripwithatotalmasstopowerratioof=4kg/kW,
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which includes thenuclearpowersupply,VASIMR,andallothervehiclemassexceptpropellantand
propellant tanks. Therefore, the Crew TransferVehicle (CTV) mass is 48mT plus the propellantand
propellanttankmass.Themassbreakdownforthe60.8mTMarsLander(ML)isassumedtobe31.0mT
forthehabitat,13.5mTforanaeroshell,and16.3mTforadescentsystem.TheCTV,matedtotheML
(MarsLander),willtransportthecrewtoMars. ThespeedrelativetotheEarthis2.5km/sattheEarths
gravitationalsphereofinfluence. Figure15showsthepilotedmissiontrajectoryandIspprofile.
The ML will separate from the CTV at Mars arrival. The lander is designed to approach Mars with a
relativevelocityof6.8km/sandexecuteadirectdescenttothesurface. TheLanderdescentmaneuver
is identicaltothatoutlined intheDRM.TheCTVwill initiallyexecuteaflybyofMars,closeenoughto
dropofftheML.TheCTVwillcontinuepastMars inanarcingtrajectorytobecapturedbytheplanet
approximately fourmonths later.Theapproach toMarsand the elliptical spiral to low Martianorbit
(LMO)areshowninFigure16.
Figure15: HumanpilotedheliocentrictransferwithIspvstimeplot(right)
The separation of the ML from the propulsion system at Mars arrival and its direct entry are
operationallyreasonablebasedontheDesignReferenceMission.Thedelayinachievingorbitalinsertion
of the interplanetarypropulsionmoduleatMars,results inconsiderable fueland timesavings.While
someriskisinvolvedinthisapproach,thecrewhasapossiblebackupbecausetheCargoVehicleinLMO
contains
the
Earth
Return
Vehicle
and
the
return
propellant,
as
well
as
a
fully
functional,
albeit
lower
power, VASIMR module. This configuration could be used in a contingency, should the prime
propulsionsystemfailtoachieveLMO.Suchanoptionwillresultinalongerreturntriptime.
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Figure16: CTVarrivalatMarsandsubsequentcapture131dayslater(left)and7dayspiralmaneuverintolowMarsorbit
(right)
UponarrivalatMars,theCTVpropulsionmodulewilldockwiththeEarthReturnVehicleandpropellant.
ThecargoengineandreactorpackagecanbereleasedafterdockingandleftinMarsorbit. Acommon
dockingmechanismwillallowtheEarthReturnVehicletomatewiththeCTVpropulsionmodule.
Figure17: CTVspiraldeparturefromMarsandreturnheliocentrictrajectory. Thespirallasts4daysandtheheliocentric
transferlasts85days.
AscentfromMarstotheReturnHabitatisaccomplishedwithachemicalascentcapsule,asintheDRM
3.0.ThiscapsulestaysattachedtotheCTVduringinterplanetaryflightandisusedfordirectdescentto
the Earth surface. Since the capsule is designed for a reentry at a relative velocity to Earth of 6.8
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km/sec,thevehicleistargetedtoapproachEarthwiththatvelocity.Thereturnmissionfollowsasimilar
strategy.Itconsistsofa4dayMarsspiralfollowedbyan85dayheliocentrictransfer,asshowninFigure
17.
6.2 HumanMission
Scenarios
to
Mars
using
200
MW
6.2.1 Technology Requirement Estimates from OptiMars
A200MWhumanmissiontoMarswasevaluatedusingtheOptiMarsprogramassuminganinitialmass
atLEOof600mT,withinitialaltitudeof1,000km,aVASIMRefficiencyof60%,andaspecificimpulse
of3,200sec. Forthispowerlevel,thespiralingfromLEOtoEarthsSOIrequires8days,and165mTof
propellant. Thisspiralsavesasignificantamountofpropellantoverachemicalmaneuvertonearthe
EarthsSOI,however,achemical rendezvousnearEarthsSOIbetweenastronautsand theVASIMR
spacecraftcouldbeconsideredtoreducethecrewsexposuretotheVanAllenbelts,furtherreducetrip
time,andalso to relax theminimumspecific impulse requirements for theVASIMR technology. No
chemical
based
rendezvous
is
considered
here,
but
future
study
may
be
useful.
In this evaluation, propellant usage during the second stage heliocentric transfer is optimized by
escapingtheEarthwithavelocitydirectedawayfromtheSunwithzerotangentialvelocityrelativeto
Earth at a speed of 5.5 km/s. The mass and velocity at the end of this spiral maneuver is used to
initialize the secondstageheliocentric transferbetweenEarthandMars,asshown inFigure18. The
optimal 31day heliocentric transfer requires 295 mT of propellant with variable specific impulse
followingtheprofileshowninFigure19between3,000sand30,000sec. A3dayperiodinthemiddle
oftheheliocentrictransferhadtheVASIMRthrustersrunningatmaximumspecificimpulsepriortothe
slowingof thecraft inpreparation fororbitalcapturebyMars. Thearrivalvelocityat theendof the
heliocentric transfer is 6.8 km/sec, which requires aerocapture to slow the lander for capture. The
arrivalmass toMars for thisscenario is140mT,and themassof thepayload issomewhatarbitrarily
pickedtobe20mT,themassbudgetforthepowersystemsandVASIMRthrustersis120mT,requiring
atotalalphaoflessthan0.6kg/kW. Currenttechnologiescannotachievespecificmassvaluesthislow,
although some theoretical studies predict power plant specific masses below 1 kg/kW. Thus, pre
positioningofcargoandcrewsupportfacilitiesinMartianorbitmayberequiredtoachievetheshortest
times for crewed missions. Nevertheless, the orbital mechanics identified by the OptiMars program
warrantsamore thoroughevaluationofmissions to identify the technologyneeded,asmeasuredby
specificmass,usingCopernicus.
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Figure18: OptiMarssimulationof31dayheliocentrictransferfromEarthtoMars. Themaximumspeedrelativetothesun
isroughly40km/spriortotheslowingmaneuvertotransferintotheorbitofMars.
Figure19: Profileofvariablespecificimpulsefor31dayheliocentrictransferbetweenEarthandMarsorbitssecondstage
ofaveryfasthumanmissiontoMars. Thehorizontalaxisindicatestherelativeheliocentrictransfertriptimestartingwith
departurefromtheEarthSOIandendingwitharrivaltotheMarsSOI
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6.2.2 Optimized Results from Copernicus
Theeffectson trip timeandpayload for200MWhumanscenarioscausedbyvarying the technology
requirements,measuredbytotalspecificmassandarrivalvelocity,arepresentedinthissection. Figure
20demonstratesanoptimized31dayheliocentrictransferforthemissiontoMarsusingtheCopernicus
program. TheoptimaldeparturedatefromEarthSOIinthenext20yearswascalculatedtobeJuly15,
2018. Thecorrespondingoptimizedvariablespecific impulseprofile isshown inFigure21. The final
mass after the heliocentric transfer is 145 mT, in good agreement with, and slightly higher than
estimatedbyOptiMars. Inthisstudy,werelaxthetotalspecificmassrequirementsbyparametrically
varyingthearrivalspeedatMarsSOIandthemissiontimewhilekeepingtheinitialmassatEarthsSOI
andthefinalpayloadmassatMarsfixed. Additionalimprovementsbeyondthescopeofthisstudycan
likelybeobtainedbychemicalrendezvousofahumancrewwithprepositionedresourcesat libration
pointsoftheEarthMoonsystem.
Figure20: Copernicussimulationof31dayheliocentrictransferfromtheEarthsSOItoMars. Theyellowlinesindicate
thrustvectors.
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0
3000
6000
9000
12000
15000
1800021000
24000
27000
30000
33000
0 5 10 15 20 25 30
time [days]
Isp
[sec]
Isp [sec]
Figure21: Copernicussimulationofvariablespecificimpulseprofilefor31dayheliocentrictransferfromEarthSOItoMars
Figure22shows the required totalspecificmass,, fordifferentmission timesandarrivalvelocities.
Theseresultsshowthatheliocentrictransfertimesfordeliveringafixed20mTpayloadfromL1toMars
usingafixedinitialmassatEarthsgravitationalsphereofinfluenceof600mT,canbeintherangeof55
to60daysdependingonthearrivalvelocityatMarswithatotalpropulsionofapproximately2kg/kW.
A39daymissiontoMarswithanarrivalvelocity10km/scanbeachievedwithtotalspecificmassof1.2
kg/kW. Achievingtriptimesoflessthan90daysforthesemasstransfersrequiresthetotalpropulsion
technology,measuredby, to belessthanabout2.5kg/kW. Notethatslowmissionsdontneedsuch
highpowerlevels. Forexample,90daymissiontoMarscanbeaccomplishedwithpowerlevelscloseto
12MWasreportedinprevioussection. AdditionalmassstagingfordeparturenearL1with800mTmay
further
relax
the
technology
requirements
on
the
propulsion
system
to
around
4
kg/kW,
but
further
studyisrequired.
0.00
0.50
1.00
1.50
2.00
2.50
3.00
35 45 55 65 75 85
Trip time [days]
T
otalalpha[kg/kW]
V_arr=0km/s
V_arr=2km/s
V_arr=4km/s
V_arr=6km/s
V_arr=8km/s
V_arr=10km/s
Figure22: ParametricstudyofhumanmissiontoMarstriptimedepartingfromL1withminimalIsp=3,000sec,IML1=600
mT,P=200MW,=60%,MPL=20mT.
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Section7. SummaryandFutureWorkAsurveyofspaceexplorationmissionsenabledbyVASIMRtechnologieshasbeenpreformedusinga
wide range of simulation tools including nearEarth, deepspace robotic and human missions. This
surveyprovidesamaptoguidemorespecificanddetailedanalysisofscenariosenabledbyexploitinganddevelopingVASIMRtechnologies. Italsoprovidesguidanceforevaluatingthetradeoffs,risks,and
rewardsinherentindevelopingnewtechnologies.
A500kWLunarTugscenariobasedonaVASIMRdrivenspacecraftandexistingstateoftheartsolar
cell technologywitha totalpropulsionalphaof10kg/kW, isable toachieve significantmass savings
overallchemicalthrustertechnology. ForanIMLEOof25.2mT,theVASIMRtugcandeliver14mTof
payloadtolowlunarorbit(LLO)versus5.7mTdeliveredbytraditionalchemicalmeans. Thus,thecost
to preposition cargo and equipment near the Moon can be cut in half using technologies that are
availableinthenearterm. Similarly,a250kWMarsSampleReturnmissionusingaVASIMRisableto
accomplishthemissionin3.5years,aboutthesameasusingchemicalthrustertechnology,butrequiring
muchlesspropellantlaunchedintoLEO.
The major advantage of Variable Specific Impulse technology is demonstrated for high power deep
spaceroboticmissions. Ascenariobasedonpresentlyavailablesolarpowertechnologieswith500kW
ofpowerandalphaof6.4kg/kWcanlaunch4mTofpayloadbeyondtheorbitofJupiter. Besidesusing
muchlesspropellantthanchemicallypoweredmissions,theVASIMRCatapulttoJupiteroptionallyhas
the advantage of reusing the hardware to amortize the initial investment. This technique warrants
furtherstudytobetteridentifyitscapabilities,tradeoffs,andadvantages.
Perhapsthemostlaudablegoalsforthetechnologyarethatitshouldeventuallyenablehumanmissions
toMarsthataremuchfasterandsaferthancanbeachievedwithchemicalrockets. Tripstoothernear
Earthobjectsclearlywarrantsimilarstudies. Using12MWofpowerandatotalspecificmassforthe
entirepowerandpropulsionsystemofachallenging,butpresentlyrealizable4kg/kW,allowsascenario
withacrewedonewaymissiontimeofapproximately3months. Assumingadvancedtechnologiesthat
reducethetotalspecificmasstolessthan2kg/kW,triptimesoflessthan60dayswillbepossiblewith
200MWofelectricalpower. OnewaytripstoMarslastinglessthan39daysareevenconceivableusing
200MWofpower iftechnologicaladvancesallowthespecificmasstobereducedtonearorbelow1
kg/kW.
Section8. References[1] Ocampo,C.AnArchitectureforaGeneralizedTrajectoryDesignandOptimizationSystem,
ProceedingsoftheInternationalConferenceonLibrationPointsandMissions,Girona,SpainJune
2002
[2] Hahn,D.W.,Johnson,F.T.,andItzen,B.F.,FinalReportforCHEBYCHEVOptimizationProgram
(CHEBYTOP),TheBoeingCompanyReportNo.D21213081,NASACR73359(or69N34537),July
1969
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[3] Yam,C.H.andLonguski,J.,M.,OptimizationofLowThrustGravityAssistTrajectorieswitha
ReducedParameterizationoftheThrustVector,AAS/AIAAAstrodynamicsSpecialistConference
andExhibit,AASPaper05375,1995,LakeTahoe,CA
[4] K.
Karavasilis,
Earth
Mars
Trajectory
Calculations,
4
th
VASIMR
Workshop,
Houston,
TX,
March
28,
2001.
[5] BBCNews(2007).ConcernoverChina'smissiletest.RetrievedJanuary20,2007.
"SatellitesCollide,PutSpaceStationatRisk".TheWashingtonPost.February11,2009.
http://www.washingtonpost.com/wpdyn/content/article/2009/02/11/AR2009021103387.html
[6] Messengerfactsonweb: http://www.astronautix.com/craft/mesenger.htmor
http://en.wikipedia.org/wiki/MESSENGER
[7] Chang Daz F. R., Braden E., Johnson I., Hsu M. M., Yang T. F. (1995) Rapid Mars Transits With
Exhaust-Modulated Plasma Propulsion, NASA Technical Paper 3539, 10p
[8] Chang Daz F. R. (2000) The VASIMR Rocket, Scientific American, November 2000, 283 (5) 90-97.
[9] Grossman,L. (2009) Ionenginecouldonedaypower39day trips toMars,NewScientist,22 July
2009.
[10]Edelbaum,T.N.(1964)PropulsionRequirementsforControllableSatellites,ARSJournal,31:1079
1089.August1961.
[11]PreliminaryPlanningforanInternationalMarsSampleReturnMission, ReportoftheInternational
MarsArchitecturefortheReturnofSamples(iMARS)WorkingGroup,June1,2008.
[12]HumanExplorationofMars:TheReferenceMissionof theNASAMarsExplorationStudyTeam,
Hoffman,S.J.,Kaplan,D.I.NASA/SP6107ADD,NASAJSC,60p.,1998.