vasimr for flexible space exploration

8/9/2019 VASIMR for Flexible Space Exploration

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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.

vasimr for flexible space exploration

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