Comparison of Four Space Propulsion Methods for Reducing Transfer Times of Manned Mars Mission

0.5 MB·English

by  A. G. C. Guerra

Comparison of Four Space Propulsion Methods for Reducing Transfer Times of Manned
Mars Mission
Andre´ G.C.Guerraa,1,OrfeuBertolamib,PauloJ.S.Gilc,∗
aInstitutoSuperiorTe´cnico,UniversidadedeLisboa,Av.RoviscoPais,1049-001Lisboa,Portugal
bDFA,FaculdadedeCieˆncias,UniversidadedoPorto,RuadoCampoAlegre687,4169-007Porto,Portugal
cCCTAE,IDMEC,InstitutoSuperiorTe´cnico,UniversidadedeLisboa,Av.RoviscoPais,1049-001Lisboa,Portugal
5
Abstract
1
0WeassessthepossibilityofreducingthetraveltimeofamannedmissiontoMarsbyexaminingfourdifferentpropulsionmethods,
2
andkeepingthemassatdepartureunder2,500tonnes,forafixedarchitecture.Weevaluatedrepresentativesystemsofthreedifferent

v
state of the art technologies (chemical, nuclear thermal, and electric), and one advance technology, the “Pure Electro-Magnetic
o
Thrust” (PEMT) concept (proposed by Rubbia). A mission architecture mostly based on the Design Reference Architecture 5.0
N
is assumed in order to estimate the mass budget, that influences the performance of the propulsion system. Pareto curves of the

6durationofthemissionandtimeofflightversusmassofmissionaredrawn. Weconcludethattheionenginetechnology,combined
2withtheclassicalchemicalengine, yieldstheshortestmissiontimesforthisarchitecturewiththelowestmass, andthatchemical

propulsionaloneisthebesttominimisetraveltime. TheresultsobtainedusingthePEMTsuggestthatitcouldbeamoresuitable
]
hsolutionforfartherdestinationsthanMars.
p
Keywords: MannedMission,Mars,Propulsion
-
p
o
p
1. Introduction of the mission. Assessment of other mission architectures to
.
s determine their impacts on mission durations was outside the
c Interplanetary space travel takes a long time with current
scopeofthisstudy.
itechnology. Forexample,ittakesmanymonthstoreachMars,
s Historically,twomainconceptsformannedmissionstoMars
yand a few years for a return mission [1, 2]. In the case of
havereceivedsignificantattention: MarsDirectandtheDesign
h
mannedmissions,suchlongtraveltimesrequirehugelifesup-
p ReferenceMission.
port systems (food, air, etc.) capable of enduring the harsh-
[ Mars Direct was a mission developed by Robert Zubrin in
ness of the space environment for long periods. These long
1990[1].Thetotalmasswassettobebelow1,000t,whichwas
2traveltimesalsopresentothercomplications,e.g. healthprob-
v regardedasafeasibletechnologicallimitbytheauthor,andthe
lemsandanincreasedprobabilityofsolarstormsduringtravel,
7 technology restricted to what was available at the time. The
whichincreasestherisktotheastronautslifeandtothesuccess
5 proposalcomprisedtwoSaturnVtypelaunchers,onebearinga
4ofthemission.Thesedifficultiescanbeeventuallyminimisedif
45tunmannedcargomoduleandtheothera25thumanhabi-
6thepropulsionsystemusedispowerfulenoughtosignificantly
tat module, with a crew of four astronauts. The modules used
0diminishboththetraveltimeandthewaitingtimetoreturn.
. the chemical upper stage of the rockets to launch to a 180day
2 With the objective of evaluating the possibility of cutting
transfer to Mars. Upon arrival the modules were assumed to
0short the total travel time of an interplanetary manned space-
executeanaerocapturemanoeuvreforcaptureintoMarsorbit.
5craft,wecomparetheperformanceofdifferenttypesofpropul-
1 For the return the crew used fuel produced from Mars’ atmo-
sion systems. We focus the analysis on a manned mission to
: sphere,andagaina180daytransferbacktoEarth. Intotal,the
vMarsasitisthenextnaturalstepforhumanexplorationofthe
missionwouldlast910days,withanestimatedcostofUS$25
i
Xsolarsystem. Asinglemissionarchitectureisusedasthebase-
billionindollarsof1990.
line, while different propulsion systems, in type and size, are
r Theotherconcept,dubbed“DesignReferenceMission”(DRM)
atestedtoassesstheimpactonthedurationofthemission. The
wasdevelopedin1992-1993,andwasbasedonMarsDirect[2].
mission architecture was selected in order to minimize mass
It has been revised several times, with the last version desig-
(whichmeanscost)butwithoutusingoptionssuchasinsiture-
nated the Mars Design Reference Architecture (DRA) 5.0 [3].
sourcemanagementoraerobrakingthatcouldincreasetherisk
Itwasassumedtocompriseacrewofsixastronauts,andtrans-
fers of about 180days (for the crew). Together with a waiting
∗Correspondingauthor
time of about 500days, the total mission would take around
Emailaddresses:[email protected](Andre´G.C.Guerra),
[email protected](OrfeuBertolami),[email protected] 900days [3]. Moreover, a maximum of twelve Saturn V type
(PauloJ.S.Gil) launchers(theAresV),bearingadescent/ascentvehicle(DAV),
1Presentaddress:DFA,FaculdadedeCieˆncias,UniversidadedoPorto,Rua asurfacehabitat(SHAB)andamulti-moduleMarstransferve-
doCampoAlegre687,4169-007Porto,Portugal.
PreprintsubmittedtoActaAstronautica November30,2015
hicle (MTV), would be needed. Each module would use nu- 2. SelectedPropulsionTechnologies
clear thermal rockets as propulsion system and would have a
ManyconceptsformannedmissionstoMarsresorttochem-
massrangingbetween60tonneand80tonne.
ical propulsion or nuclear thermal systems, but do not include
Bothproposalsforeseeanextendedmissionduration. Itis
themostadvancedtypesofpropulsionsystems,includingnew
well known that in space humans are subjected to many haz-
ards, such as loss of bone/muscle mass and a substantial in- chemical engines [8, 9, 10, 11, 12]. We have selected four
examples of state of the art or advanced propulsion technolo-
crease in the probability of developing cancer [4, 5]. There-
gies, and evaluate their potential to reduce the duration of hu-
fore, a 180day transfer (plus waiting time and a similar time
man Mars missions (similar to the DRA 5.0 architecture), re-
to return) is a significant risk, most particularly in view of the
spectingareasonablecriteriaforthetotalmassofthemission
unpredictabilityofthesolarcycle.
(≤2,500t).
Othermissionconceptsweredevelopedafterthosepresented
The selected propulsion systems exhibit high specific im-
above,forexamplethe“AustereHumanMissionstoMars”[6].
pulse(I ),highthrust(F )andhighthrust-to-weightratio(F /w)
These are variants of the DRM, with updated technology, but sp T T
withoutconsiderabledifferencesinthemissionduration. Addi- among the numerous technologies available in the literature.
tionally, many concepts include insufficiently tested technolo- We have considered propulsion technologies currently being
testedorwithflightprovencapabilities.Inaddition,onepropul-
gies,suchasinsituresourceutilisationoraerocapture,increas-
sion system was examined that is an exception to this latter
ingtherisks,cost,orthedevelopmenttimeofthemission.
premise, the PEMT, presented by Carlo Rubbia [13]. For a
Therearetwokeyfactorsworthconsideringwhentryingto
broaderdiscussionofpropulsionsystems,includingaputative
develop an architecture for human missions to Mars. Firstly,
gravitycontrol,thereaderisreferredtoreferences[10,14,15,
onlythemannedcomponentofthesystemhasastrongrequire-
16,17].
menttominimisetraveltime;accessoriestobeusedlatercanbe
Thefourpropulsionsystemsconsideredarethe:
sent in advance and meet the astronauts afterwards, thus min-
imising cost. Secondly, mass is one of the drivers of the cost 1. CommonExtensibleCryogenicEngine(CECE)–repre-
andperformanceofthesystem. Consequently,aspacecraftdi- sentingtheclassicalchemicalpropulsion;
videdinmodules,servingspecificfunctionsatspecificphases, 2. NuclearEngineforRocketVehicleApplication(NERVA)
and discarded as soon as they have fulfilled their function can II–assumedtoberepresentativeofnuclearthermalpropul-
be used to minimise the total mass, following the successful siontechnology;
lunarorbitrendezvousapproachusedintheApolloproject[7]. 3. RadioFrequencyIonTechnology(RIT)XT–assumedto
Therefore,ourapproachistouseabaselinemissionarchi- berepresentativeofmodernelectricpropulsiontechnol-
tecturethatfulfilsallthemajorrequirementsofamannedmis- ogy;
siontoMars, andminimisesthemasslaunchedfromEarthby 4. PureElectro-MagneticThrust(PEMT)–anadvancedcon-
separating the manned part from the cargo component. Af- ceptpropulsionconcept.
terwards, we select four propulsion technologies for evalua-
tion,anddeterminetheminimumtransfertimepossibleforthe 2.1. ClassicalChemicalEngines
mannedcomponent,asafunctionofthesizeoftheengines(re- Inthisbriefdiscussionwedonotconsiderstorable,mono-
specting known engine size constraints). Selected propulsion propellant, or solid systems [9], but only cryogenic systems,
technologiesinclude:classicalchemicalengines(asithasbeen due to their higher thrust (when compared to the aforemen-
the workhorse of space exploration), nuclear thermal engines tionedsystems)andI ≈400s[10].
sp
(which in the 1940s was believed to be the future technology ChemicalenginespresenthighthrustandrelativelylowI ,
sp
forexploringthesolarsystem),modernelectricalengines(asit and accelerating continuously with a chemical engine would
is currently considered the best flight proven system for many rapidlyleadtoanimpracticalmassbudget. Theyaretherefore
applicationsduetoitshighspecificimpulse),andamorecon- usedasimpulsiveengines,i.e.onlyactiveforsmallintervalsof
jecturalconceptofthemanythathavebeenproposedasthenext time(andthespacecraftisinfreefallformostofitstrajectory).
revolutionary engine. We selected as the revolutionary engine Nevertheless, chemical engines can be crucial in escape and
the “Pure Electro-Magnetic Thrust” (PEMT) concept because capturemanoeuvres,becauseoftheirhighthrust. Furthermore,
itpromised,intheory,tobehighlyeffectiveduetoitsfullcon- impulsiveenginescanminimisethesocalledgravitylossesfor
versionofmassintoenergywithmomentumusableforthrust. the same total ∆v available, by being used at once in a more
We are interested in finding the minimum order of magnitude favourablelocation[18]—usuallydeeperinthegravitywell—
ofthemissionduration,withinfeasiblebounds,andcomparing and with the additional advantage of possibly dumping empty
themeritsofthedifferentpropulsionsystems.Wehavenotopti- fueltankssooner,maximisingperformance.
misedthemissionforeachpropulsionsystem. Wehaveinstead The most important limitation of the chemical propulsion
consideredonlytransfertrajectoriesthatproviderepresentative is that it is limited to the available chemical energy and ther-
performanceforeachsystem. modynamicconditionsofthepropellants[10]. Themainpoint
2
of developing other propulsion methods is to overcome these higherbutthedifferencetoNERVAIIisnottoolargeandthis
limitations. engineshowedoverheatingproblemsduringthetestsconducted
ForourMarsmission,weonlyconsideroperationsinspace, atthetime.
taking for granted that some launch vehicle takes the system OptionssuchastheORION,anuclearpulsepropulsionsys-
intoorbit. Anenginewiththeabilitytoberestartedisalsore- temdevelopedinthe1950/1960s,wasnotconsideredduetoits
quired, to cope with the various mission phases. Examples of lowF /w(betweenoneandsix),andtheneedtoblastnuclear
T
modern chemical engines fulfilling these criteria are the Vinci material[26].NoticethateventhoughtheORIONproposalhas
engine,thatisbeingdesignedfortheupperstageofAriane5[19], agreater F /wthanelectricpropulsion(discussedbelow), the
T
the RL10B-2, the latest version of the RL 10 engine and used I ofthelatterismuchlargerthantheformer.
sp
in the Delta IV launch vehicle [20], and the Common Exten-
sible Cryogenic Engine (CECE) of Pratt & Whitney Rocket- 2.3. ModernElectricEngines
dyne[21],alsoanevolutionoftheRL10.Ofcourse,forthcom- Electric propulsion overcomes the limitations of chemical
ing technological developments involving, for instance, zero engines by separating the energy source from the propellant
boiloffforcryogenicpropulsionfluidsmightbeconsidered,but material, and by not using thermodynamic mechanisms to ac-
asmostoftentheserepresentimprovementsratherthanbreak- celerateparticles. Commonsourcesofenergyaresolar,nuclear
throughstheywillprobablynotaffectconsiderablyourassump- powergeneration,andradioisotopethermalgenerators(RTG)[9].
tionsandconclusions. Thethrustproducedbycurrenttechnologyisverysmall,when
We selected as representative of the chemical propulsion comparedtochemicalengines.However,theycanachievemuch
system the CECE engine. Although the thrust of the Vinci larger I , allowing the engine to run for longer periods with
sp
engine (180kN) is higher than the others (≈ 110kN), it has lessfuel. Theelectricengineistreatedasanon-impulsiveen-
a much higher mass (almost the double of the lighter). Ad- ginesinceitisworkingmostoftheflight. Inthesesystemswe
ditionally, they all have similar I . Consequently, the CECE have to take into account not only the mass of the engine but
sp
hasamuchhigherthrust-to-weightratio, which, togetherwith alsothemassoftheassociatedenergysystem,e.g.solarpanels
theabilitytorestartmanymoretimes(50, insteadof5or15), andthepowersupplyandcontrolunit(PSCU)[27].
makesitthebestchoiceforsuchacomplexmission. Relevant The selected power source is the solar photovoltaic, since
characteristics of one CECE engine can be found in Table 1, nuclearpowersystemsrepresentalargeincreaseoftheengine
whereinallcasesthemassofpropellantsandtheirdepositsis mass, andtheRTGtechnologycanonlyachievespecificpow-
notincludedsincetheycanbediscardedandwillbeaccounted ers of 5W/kg (and are under development) [9]. The size and
separately. massoftherequiredsolararraycanbeestimatedforthepower
leveloftheengineused(asindicatedinTable1)[28],andmust
2.2. NuclearThermalEngines be taken into account on the mass budget of the mission. We
Nuclearthermalengineshavebeendevelopedsincethe1940s, consideredanefficiencyof29.5%forthesolarcells,adensity
andwereevenconsideredfortheupperstageontheNovarocket of ρsp = 0.84kg/m2, with degradation rate of 0.4% per year,
A
(forthelunardirectlaunchmission)[10]. Theirworkingprin- and the worst case scenario of a three-year mission [29, 28].
cipleissimilartochemicalengines,withhighthrustandI ≈ For the PSCU a direct-drive concept was used, with a corre-
sp
800s, and therefore should be treated as impulsive. A single spondingmassgivenbyascalingparameter(MPSCU =0.35W+
propellant,usuallyhydrogen,isheatedbythenuclearcoreand 1.9kg)[27]. Thesevaluesgiveadensityfortheassociateden-
isexpelledthroughanozzlewhileexpanding.Thecore,usually ergy system of about 7kg/kW, well within what was used in
anuraniumderivative(likedioxideorcarbide)orplutonium,re- otherstudies[17].
leasesheatduetothenuclearreaction,providingenergytothe Several electric propulsion technologies are in use such as
gas expansion, and resulting in an I approximately the dou- the arcjet, Hall effect, and gridded ion thrusters [9]. Modern
sp
bleofthechemicalengines. Theheatreleasedislimitedbythe examples of electric engines are NASA’s Evolutionary Xenon
meltingpointofthematerials[9]. Thruster(NEXT),anevolutionofthealreadytestedNASASo-
Oneoftheengineswithhighestpowertobedevelopedwithin lar Technology Application Readiness (NSTAR) used in the
the Nuclear Engine for Rocket Vehicle Application (NERVA) DeepSpace1mission[30]andDawn[31],theradiofrequency
programwastheNERVAII,whichhadthegoalofachievinga ionthrusterRIT-XT,whichworksbygeneratingionsusinghigh
higher I ,andthrust,withalowerweightthanpreviousmod- frequencyelectromagneticfieldsandisverysimilartotheRIT-
sp
els[22]. Itwasintendedtoserveasthepropulsionsystemfor 22engine(indesignandthrust-to-weightratio)buthashigher
manned interplanetary missions with masses close to 1,000t. I [23, 32], and the PPS 1350-G, which is a plasma thruster
sp
NERVAIIproducedtherequiredpowerwithauraniuminven- withflightprovencapability(SMART-1mission)[33].
toryof360kg,andhad2mindiameter[22]. Temperaturesof To select the electric propulsion system, including all ele-
thehydrogenfuelcouldreach2,755K[22].Oneoftherequire- mentsrequiredsuchaspowerprocessingelectronicsandpower
ments of the program was an endurance of over 600min [24]. source, we considered not only the specific impulse I but
sp
ItsmainfeaturesareshowninTable1. also the thrust-to-weight ratio, because the latter also affects
Anotherengine,withinterestingfeatures,developedwithin the transfer time [34], which is our main concern. Moreover,
ProjectRoverprogram,wasthePhoebusIIengine,whichfea- thepowerrequirementsoftheenginealsoinfluencetheengine
turedF /w=38andI =790s[25]. Thesevaluesarealittle selection, since it has an impact on the size and mass of the
T sp
3
Table1:Maincharacteristicsofselectedenginesforasystemconsistingofoneengine(datafrom[21,22,
23]).Themassofpropellantsandtheirdepositsisnotincluded
Engine Power[W] I [s] Thrust[N] Mass[kg] F /w[N/kg]
sp T
CECE - 465 0.11×106 256 435
NERVAII 5.0×109 785 1.0×106 34×103a 30
RIT-XT 3,260 4,600 0.12 32b 3.7×10−3
PEMT 6.1×109 30.6×103 20 32×103c 0.64×10−3
aThereactor’smassis11×103kg.
bIncludesmassoftherequiredsolarpanelsandpowerprocessingelectronics.
cIncludesmassoftheradiatorandreflector.
associated energy system. Consequently, the selection of the thepotentialforsignificantlyreduceflighttimesforhumanmis-
electricengineprovedtobemoredifficultthantheothercases. sionstoMars. Thisengineusesthemomentumofemittedpho-
WeevaluatedboththePPS1350-GandtheRIT-XTengines tons,insteadofexpellingaworkingfluidtocreatethrust[13].
insimulatedtransferstoMars.Differencesinperformancewere The thermal energy produced in a nuclear reactor is used
foundtobesmall,withanapparentadvantagetothelatter,and toheataradiator,whichemitselectromagneticradiation(pho-
weendupdecidingfortheRIT-XTengineastherepresentative tons).Theradiatorisinfrontofareflectingsurfacetodirectthe
engine of the electric propulsion. Its main characteristics are radiationthatproducesthrust. AWinstoncone(Fig.1),anon-
shown in Table 1 (for a single RIT-XT engine, including the focusing reflecting conical structure, can be used to collimate
energysystem). theradiation,resultinginatotalthrustof F = W/c,whereW
T
Using the RIT-XT engine as baseline, we also briefly dis- isthepowerandcthespeedoflight[13].
cussthepossiblegainsinperformanceifsomefuturetechnolo- Thepoweremittedisrelatedtotheareaofthesurfaceofthe
gieswouldenhancethespecificimpulseorthethrust-to-weight radiator(S)anditstemperature(T)throughStefan-Boltzmann’s
ratio(seesection4.3). Weconsideredaspecificimpulseanda law, and the power produced by the nuclear engine. It is pos-
thrust-to-weightratiomorethantwotimesandthreetimesthe sibletoensurethatareasonablesizedradiatoryieldssay,20N,
correspondingvalueoftheRIT-XTengine,respectively. While foraradiatortemperatureofabout3,300K(closetotheboiling
the considered specific impulse increase would correspond to pointofacoolant).Amongthematerialsthatcanwithstandthis
adirecttechnologyenhancementintheengine,theincreasein radiator high temperature without melting, carbon nanotubes
the thrust-to-weight ratio was based in foreseen developments (as suggested by Rubbia), are the lightest [13]. The cone re-
intheenergysystem. flector must then envelop the radiator, and should have high
Anotherelectricalpropulsiontechnologythathasbeendis- reflectivityforthewavelengththeradiatorisemitting. Forour
cussedformannedMarsmissionsisthemagnetoplasmadynamic radiatortemperature,thereflectorconeshouldbecapableofre-
thruster(MPD)[17,35]. Theexpectedtotalthrustgeneratedby flecting visible and infrared radiation to be effective. Current
theseenginesisordersofmagnitudelargerthanthepreviously technologyforsolarsailsusecompositebooms(onthesupport
discussed electrical engines, even though the thrust-to-weight structure)andAluminisedMylarsails(orcarbonfibresailsub-
ratio and specific impulse is of the same order. However, the strate), with densities of 10g/m2 (including the support struc-
MPD that is claimed to have better results [17], explains that ture)[11].Combiningthedensitiesofthematerialsandsizesof
theenginecouldworkonlyforabout1000hoursbeforedegra- the structures, the radiator and reflector mass can be extracted
dationoccurs[35],whichisconsiderablylessthantheexpected (usingtheideasof[13]).
triptimes. Furthermore,the F /wand I valuesfortheMPD As power source we have selected a NERVA-like reactor,
T sp
arewellwithinwhatwetestforpossiblefuturetechnologiesin resized and improved, as this is one of the discussed nuclear
section4.3.Consequently,wedecidednottoexplicitlyconsider reactors with a power level closer to our intended figure. It
MPDthrustersascontinuousimpulseenginesinourstudy,al- is assumed a “NERVA 2000” reactor which can produce 22%
thoughwiththeexpectedlevelsofthrustoftheseenginesthey moreenergywithanincreaseinmassof26%,togetanet20N
couldbeconsideredtobeinthefrontierbetweenimpulsiveand engine. Theneedforanuclearreactorisjustifiedbythe50km2
non-impulsiveengines,andfuturestudiesshouldbeperformed of solar panels, weighting 44kilotonnes, required to produce
(not only to increase the endurance time, but also to explore the same power using the already mentioned current technol-
whatcanbeachievedwiththisoption). ogy[11,27].
In the PEMT concept no mass is expelled. However, it is
2.4. NuclearPureElectromagneticThrust possibletodetermineits I , tocompareitwiththepropellant
sp
Propulsion technologies based on new concepts are con- mass spent by chemical thrusters [13]. If we assume that the
stantly being proposed and tested. For instance, Electrody- fraction of mass transformed into energy through nuclear fis-
namic Tethers, MagSails, Plasma Sails, and Solar Sails [11]. sion is ξ = 10−3 [13], and it is ejected at the speed of light c,
WeincludedtheNuclearPureElectromagneticThrust(PEMT) thentheeffectiveexhaustspeedisvex = ξc. Thus,thespecific
conceptinourevaluationtoseeifsuchanadvancedconcepthad impulse is given by Isp = 30,600s. The main features of the
4
Figure1:SchemeofRubbia’sengineconcept[13]
engineareshowninTable1. fractional burnup β, defined as the ratio of the number of fis-
The PEMT engine is, nevertheless, still a theoretical con- sionsforaspecifiedmassoffueltothetotalnumberofheavy
ceptandpresentmanypotentialdifficultiestobeimplemented atoms [37]. Furthermore, if all fuel atoms where fissioned,
in practice: control of nuclear reactions in space (in free fall β = 1, this would lead to 950GWd/t for the uranium isotope
cooledonlybyradiation),insulationoftheengineandtherest 235 (235U) [37]. Consequently, the specific burnup is sb =
ofthespacecraft,transferofenergyfromthereactortothera- 950βGWd/t. Throughout our work we have used β equal to
diator,etc. Therequiredhightemperatureisanotherchallenge 4%[37,38].
since the radiator material must withstand it without melting. Todeterminethemassofuranium,M ,requiredtoproduce
U
Somesolutionssuchasdiscussedin[13]aresomewhatdifficult theneededenergy,wedividethetotalburnupobtained,bythe
toimplement. Weneverthelessdecidedtoincludethistechnol- specificburnup, i.e. weuse M [t] = B[GWd]/sb[GWd/t] =
U
ogy,anduseitsoptimisticcharacteristicsatfacevalue,toassess W ×t /950β. If the nuclear reactor is loaded with ura-
reactor on
if such an engine proposal would offer some real advantages nium dioxide (UO ), enriched to almost weapon’s level (80%
2
overotherwellknownoptions. of235U),usingM = M ×AUO2/AU (whereAXistheatomic
UO2 U w w w
mass of element X), we can compute the needed mass of ura-
2.5. FuelCalculation niumdioxideM [39,40,37].
UO2
For engines that can be treated as impulsive, we consider
theusualapproximationofinstantaneouschangeinvelocity∆v, 3. MissiontoMars
obtaining the propellant mass (M ) through the Tsiolkovsky
P
rocket equation. The complete mission trajectory requires N The mass and time required for a mission to Mars is de-
manoeuvres that will be determined from the last to the first, pendent on its architecture. We consider a mission similar to
since the fuel mass depends on the initial mass for each seg- theDRA5.0concept,thatincludesamannedspacecraftwitha
ment. Thefiniteburnlossesaretakenintoaccountbydefining crewoffourastronautsandanunmanned,orcargo,spacecraft.
alossfactor[36]. The manned spacecraft is comprised of a human habitat
For continuous thrust systems the spent fuel is determined module(whichhousesthecrewduringtransittoandfromMars,
ateachinstantbym˙ =−F /I g ,andtakenintoaccountinthe andincludesalllifesupportsystemsforthemission),apropul-
T sp 0
numericalintegrationoftheequationsofmotion. sion system and a transport capsule. The first two might be
In the case of PEMT, the fuel required is the nuclear ma- assembledinorbit. Thetransportcapsulecarriesthecrewfrom
terialforthereactor, whichisnotexpelled(ityieldsphotons). lowEarthorbit(LEO)tothemainspacecraftonEarth,andbe-
Therefore,weneedtocomputetheamountofnuclearmaterial tweenthemainspacecraftandloworbitonMars(uponarrival).
toloadthereactor. The unmanned cargo spacecraft consists of the propulsion
The thrust force of PEMT is given by the radiated power system,thepayloadforMarsoperations,includingthedescent
(W ) divided by the speed of light (F = W /c) [13]. We andascentvehicle,andthefuelrequiredfortheastronautstore-
rad T rad
assumethattheradiatedpowerisequaltothepowergenerated turntotheEarth,thatwillbetransferredtothemainspacecraft
by the reactor, i.e. W = W . Combining this with the whileinMarsorbit.
rad reactor
time of flight (ToF), which is equal to the operational time of
thereactor,t ,wecomputethetotalburnupofthereactor,B= 3.1. MissionArchitecture&MissionTimeline
on
Wreactor×ton,expressedinGW×day[13,37]. Thesamearchitectureisusedtocomparetheperformance
The specific burnup, sb, of a nuclear material is the total ofthedifferentpropulsionsystems. Consideringthattheobjec-
energy released per unit of mass of nuclear fuel, and is ex- tive is to minimise the travel time for the astronauts, the crew
pressed in MW×day/tonne [38]. This is proportional to the
5
and most of the cargo are sent separately. This allows send- 3.1.3. CaptureatMars
ing cargo through a slower and more economic way, reducing ThecapturemanoeuvreatMarsisperformedusingtheclas-
substantiallytheinitialmassofthemannedspacecraft. sicalsingleimpulsebrake,sincecontinuousthrustbrakewould
Themissiontimelineforthemannedphaseisdisplayedin requiretoomuchtimetoexecuteforitslevelofthrust[13,41,
Fig.2. 9]. We also considered that aerobraking would be too dan-
gerousforthemannedspacecraft[42], andalsowouldrequire
3.1.1. ParkingOrbit&Departure shieldingandasomehowcompactandstrongarchitecture,adding
We have considered that the mission starts at LEO, where tothemassandloosingpartofitsappealing(andpossiblynot
a launch system can deliver all the required modules for the compatible with a high area of solar panels in some options).
mission. Boththemain(manned)andthecargospacecraftcan Aerobraking,ifoneacceptstherisk,requiresawholedifferent
beassembledataninitialcircularparkingorbitatabout500km analysisbeyondthescopeofthiswork.
altitude,similartotheInternationalSpaceStation,andwhichis Thecaptureorbit,wherethemainspacecraftwillbeparked,
highenoughtoallowassemblywithoutdecay. hasaperiapsisaltitudeof300km,selectedbycomparisonwith
Afterassemblingandtesting,thespacecraftwouldraiseits other missions [43, 44, 45] and eccentricity 0.9. We selected
apogeetoanorbitwitheccentricity0.9,inaseriesofimpulsive this high energy orbit such that its low periapsis allows for an
manoeuvresatperigee, sotominimisefiniteburnlosseswhen effectivecapturemanoeuvreunderhighvelocity, butminimis-
compared to a direct interplanetary injection from LEO [36]. ingthefuelconsumptionduringthecapture. Themainspace-
Thissavesmassbydiscardingthefueltanksalreadyused. The craft is also used in the return trajectory and fuel is saved by
finalellipticalorbitisselectedtomaximiseitsenergy(butnot notloweringtheapoapsis. Thefuelforthereturntrajectoryis
tooclosetotheescapeenergytoavoidcomplicationwithorbital transported by the cargo spacecraft (sent previously and with
perturbationsandlongperiodsofrevolution).Atthesametime, thesamecaptureorbit),andcannowrefuelthemannedspace-
wekeeptheperigeelowtousetheeffectivenessoftheOberth craftforthereturn.
effectwhenescapingfromtheEarthinfluence. Wedidnotconsiderinsituresourceutilisation(ISRU),ex-
Ifcontinuousthrust(whenapplicable)wouldbeusedatde- cept for recycling of air and water (possibly included into the
parture, too much time would be required to escape, since the capsule),asourarchitectureisadequateforafastfirstmission,
continuousthrustconsideredisyetofrelativelysmallintensity. andISRUpresentsanewsetofchallenges.Apartfromtherisk,
Therefore, weonlyconsideredchemical(ornuclearthermalif use of ISRU in a large scale to grow food and obtain mission
applicable)propulsionforescapeandcapturemanoeuvres,and support resources in general only makes sense for prolonged
for the apogee raising initial manoeuvre. In this context, we stays, which is not the case we are considering. The only re-
consideredseveralescapevelocities(cid:126)v leadingtodifferentini- maininginterestinguseofISRUisforgeneratingreturnpropel-
∞
tialvelocitiesfortheinterplanetaryphase. lants. However,inthepresentcase,theywouldhavetoascend
into orbit (with a 0.9 eccentricity), demanding more complex
3.1.2. InterplanetaryTransfer meansoftransportationthanjustbringingastronautsback.This
The proximity of Mars implies that interplanetary transfer woulderodetheadvantageofISRU,astheachievedgaininpro-
solutionswithfly-bymanoeuvreswillincreasethetraveltimes pellantmasswouldhavetocompensateaconsiderableincrease
making these uninteresting for the main spacecraft. For sim- inthemassofthetransportvehicle,andtheincreaseinpropel-
plicity, we did not considered this option for the cargo mis- lantmassrequiredtotransportthisvehicletoMars. Therefore,
sionaswell.Wearemainlyinterestedincomparingtherelative itisnotcertainthatISRUwouldbringconsiderableadvantages
performance of the different propulsion systems and different, fortheconsideredtypeofmissionand,althoughitsanalysisis
betteroptions, forthecargospacecraftcanbeusedbyall, im- beyondthescopeofthiswork,itremainsanopenissue.
provingalltheconsideredalternativesequally. Wealsodidnot Meanwhilethecrew(usingthetransportcapsule)wouldget
consider deep (impulsive) space manoeuvres as these are usu- alltherequiredpayloadfromthethecargospacecraft,andde-
allyusefultosynchronisewithplanetsforfly-by,andtheyseem scendtoa300kmcircularorbit. Afterwards,thecrewchanges
nottoofferanyadvantageascomparedwithalarger∆vnearthe totheMarsoperationsvehicle,thatwilldescendtothesurface,
Earth(initialorfinalmanoeuvre),duetotheObertheffect. whilethetransportcapsulewaitsinthecircularorbit.
For the impulsive-type propulsion, chemical and nuclear
3.1.4. ReturnTrajectory
thermal, once the initial velocity is defined a coast trajectory
(Lambert arc) takes the spacecraft to the arrival planet, where Aftercompletionofthegroundoperations,thecrewascends
it is captured. For the continuous thrust propulsion, electric from the surface in the Mars operations vehicle (or part of it),
and PEMT, the direction and intensity of the thrust can make returning to the transport capsule at low Mars altitude. Af-
aconsiderabledifferenceandmustbedeterminedineachcase terrendezvousanddiscardoftheMarsoperationsvehicle, the
toobtainasuitablesolution. Thecontinuousthrustcanalsobe transportcapsuleraisesitsorbittoreturntothemainspacecraft
allowedtobrakebecauseitcouldimplyasmallerrequirement fortheMtE(MarstoEarth)injectionmanoeuvre.
ofpropellantforthecapture. The capture at Earth is similar to the one at Mars (the en-
Ourgoalwasnottocompletelyoptimisethetrajectory,but tirespacecraftiscapturedandpossiblyreused;also,propulsion
tocomparetherelativeperformanceofthepropulsionsystems. systems with nuclear material would have to be carefully de-
commissioned). Subsequently,thecrewentersthetransportve-
6
Launch Rendezvous
with Main Departure
Spacecraft
Parking Orbit Rendezvous
between
Crew and C3 Value
Command
Module
Assembly Interplanetary
Phase
Refuelling
C3 Direction
Apoapsis
Raising Ascending
Thrust
Direction
Launch Orbit Command
Module
Parking Brake
No Brake
Landing
Thrust
Direction
Departure
Low Circular
Orbit Capture
C3 value
Interplanetary Main Command
Phase Spacecraft Module
Main Command
C3 Direction Spacecraft Module
Rendezvous
Thrust with Cargo
Direction Low Circular
Orbit
Capture Orbit
Brake No brake
Landing
Capture
Thrust
Direction
Figure2:Sketchofthemannedmissiontimeline. DiamondsrepresentSingleinstantsoftime;rectangles,Missionphases;squares,Actions;ovalshapes,Variable
parameters(C3equalsthesquarerootofthedeparturevelocity)
hicle and returns to LEO where it will be transported back to injection.Continuouspropulsioncouldbeusedforthispurpose
Earth’ssurface. advantageously. As the cargo spacecraft mass has mostly the
A mission trade tree is displayed in Fig. 3, where we can sameimpactonthefourdiscussedpropulsionsystems, wese-
seewhichpropulsionsystemisusedineachmissionsegment. lected,withoutanyadverseeffectsontherelativeresults,anim-
Whenacoasttransferisselectednopropulsionsystemisused pulsive(purelychemical)transferapproximatedbyaHohmann
duringtheinterplanetaryphase. transfer,forsimplicity,andsinceafasttrajectoryisnotrequired
forthecargo(themaingoalistominimisetherequiredenergy
3.1.5. CargoMissionPhase savingmass).
Starting from the same assembly circular LEO, the cargo ThecargoistoarriveatMarsbeforethecrew,forthepay-
spacecraftraisestheapogeealtitude,similarlyandforthesame load to be ready to descend to Mars. Upon arrival on Mars
reasonsthanthemannedspacecraft,tothesamehighelliptical thecaptureofthecargospacecrafttoahighlyellipticalorbitis
orbitofeccentricity0.9,andperformsanEtM(EarthtoMars) similartothemannedspacecraft,sincethelatterwillhavetobe
7
Consequently, the dry mass of the manned spacecraft is
Initial Orbit MEsectahpoed InteTrrpalnasnfeerta ry MCaeptthuorde Option MDMraynned = MMfixaendned + MPS, where the tanks are not included
since they are discarded after each manoeuvre. Its total mass
Coast CECE 1 is MTMoatanlned = MDry + MPT + MTT. The dry mass of the cargo
(no thrust) spacecraftisMCargo = MCargo+M +MFP. Thefuelpayload
Dry fixed PS R
CECE RIT-XT CECE 2 entry(MFP)representsthepropellantsandtanksneededforthe
R
manned return trajectory. The total cargo mass is MCargo =
PEMT CECE 3 Total
Parking Orbit M +MT +MT.
Dry P T
Combining the total manned spacecraft mass for the Earth
Coast
NERVA II NERVA II 4
(no thrust) toMarstransfer,withthetotalcargomass(whichasmentioned
aboveincludesthereturnfuelforthemannedspacecraft),yields
thetotalmissionmass(M = MManned+MCargo).
Total Total Total
Figure3:Differentengineconfigurationsassumedforthemannedspacecraft
3.3. TrajectoryProblem&Solution
To determine the trajectory and time of flight we adopt a
refuelledbytheformertoreturntoEarth.
simplepatch-conicapproximation[48], withanumericalinte-
grationintheinterplanetaryphaseofthemission. Weconsider
3.2. MassBudget
Earth and Mars to be in the same plane and in circular orbits
Whilecrew,modules(habitatandcargo),andtransportcap-
aroundtheSun(withradiusequaltothetruesemimajoraxis).
sules remain constant, the type of propulsion and number of
Thismakestheproblemonlydependentontheheliocentrican-
enginesarevariablesoftheproblem. Propellantsandtanksare
gle between the planets and the characteristics of the mission,
alsodependentontheexecutiontimeofthemanoeuvres,since
andnotthespecificlaunchdate.
we considered different levels of finite burn losses [36]. The
The spacecraft escapes from, and is captured into, an el-
humanhabitatandcargomodules(includingallnecessarysys-
liptical orbit with instantaneous manoeuvres at the periapsis
temsandpayloads),andtransportcapsules,areanextrapolation
of the orbits, using chemical or nuclear thermal propulsion.
from Mars Direct [1] and DRA 5.0 [3] concepts using typical
The terminal velocity, (cid:126)v , of the escape hyperbola can make
∞
guidelines[46].
any angle with the velocity of the departure planet, within the
Themassbudgetforthemissionisthereforecomprisedof:
planet’sorbitalplane. Thearrivalvelocityangleisdetermined
by the interplanetary phase, a simple Lambert arc in the case
• Manned spacecraft (a fixed value including all systems
andcrew,exceptthepropulsionsystem): MManned =38t; ofimpulsive-typepropulsion,oratrajectorydeterminedbythe
fixed continuousthrustthatcanincludeabrakesegmenttoeasethe
• Cargospacecraft(againwithexceptionofthepropulsion capturemanoeuvre.
system): MCargo =42t; Arelativelysimpleapproachisusedtoobtainasolutionof
fixed
the continuous thrust trajectory. Continuous thrust has a high
• Marsoperationsvehicle(whichwillfollowwiththecargo
I , and the expense of propellant is not the main issue, so to
spacecraft)MMOV =18t; sp
minimisethetransfertimethepropulsionsystemworksatfull
power [49, 34]. Consequently, only the direction of the thrust
• PropulsionSystem,M :
PS
remainsas acontrol parameter. Aconstant angleof thethrust
– Impulsiveengines(chemicalornuclear); with the instantaneous velocity vector (when accelerating and
otherforbraking)isconsideredasbeingacompromisebetween
– ContinuousThrustEngines,including:
asimplesolutionandtheoptimisationprocedureusedin[34].
∗ Electric propulsion case: solar panels includ-
This simple procedure does not assure a real optimal solution
ing structure, and power processing electron-
foragivenmassofthemission,butitshouldbeenoughtocom-
ics;
paretheperformanceofthepropulsionsystemsanddetermine
∗ Rubbia’s concept engine case: the radiator & anapproximatetimeofflight.
reflector; The duration of operations on Mars is determined by the
waitingtimet .OurgoalwastominimisethemannedToF,and
• Propellants,MT; w
P ultimatelythetotaldurationofthemannedpartofthemission
• Propellanttanks,MT; (tmission =tEtM+tMtE+tw),andforafirstexplorationmissionan
T
extendedtimeforoperationsisnotrequired. Thewaitingtime
Themassofthepropellanttankscanbecomputedfromthefuel is determined by the time of flight of the EtM and MtE trans-
mass(computedusingthealgorithmdescribedinsection2.5), fers, and depends critically on the heliocentric angle between
usingafittingfunction(fromseveralspecificationsofexisting theplanetsuponarrivalatMars. Asweshowonsection4,for
tanksasinputparameters[47]). Furthermore,theyaresizedin allreasonablevaluesoftheparametersoftheproblem,itisim-
ordertobeemptyanddiscardedaftereachlargemanoeuvre,as possibletoavoidaprolongedstayonMars.
possible.
8
3.4. ProblemParameters larger. Usually,reversingthethrustisonlyconsideredafterthe
Once the baseline mission is defined (including the mass middle of the trajectory [34, 50, 51], but after testing we de-
budget), and settled what is transported to Mars, we focus on cided to consider that the reverse could occur at any of nine
theobjectiveoftryingtominimisetheToFofthemannedseg- equallyspacedradialdistancesbetweenEarthandMars.
ment. For each propulsion system, the available thrust is in-
creasedbyvaryingthenumberofenginesemployed.
4. Results&Discussion
Asarguedin[1](theMarsDirectmission)aspacecraftwith
more than 1,000t should be avoided. However, since we aim 4.1. EffectsofAngleVariationandDirection
totestnewpropulsionsystemsandestablishhowfastwecould Tobetterperceivetheeffectthatθ andθ haveonthesys-
reachMarsandgetback(andsincewearenotcompletelyop- v FT
temmassandToF,weplottedtheevolutionoftimeversusmass,
timizingthecontinuousthrust,andforsimplicityuseonlyim-
foranengineconfigurationandasingleinitialvelocity(without
pulsive thrust on the cargo spacecraft), we stretch the limit to
brakesandforanonewaytrip).
2,500tforthewholemission(i.e.thesumofthemannedand
In Fig. 4 we can see how a variation on the direction of
unmannedspacecraft). Withoutamasslimit,themissioncould
theinitialvelocity(maintainingthesameabsolutevalue)result
growindefinitelywithsomeToFgains,butataprohibitivecost.
invariationsofmassoftheorderofmagnitudeofthehundred
We consider that the trajectory, and travel time, is defined
tonne,andtensofdays.Forsimplicityonlythe|θ |variationare
v
bythefollowingparameters:
annotatedinthefigure,andonlyforanEarthtoMarstransfer.
1. Type of propulsion (see Table 1), considering different Thevariationofθ doesnotproduceasstrongeffectasθ here,
FT v
numberofenginestoincreasetotalthrust; although for |θ | = 70◦ some dispersion of the results can be
v
2. Escapevelocityondeparture,(cid:126)v (moduleanddirection); seen(whichisrelatedwithθ ).
∞ FT
3. Thrustdirectionintheinterplanetaryphase,ifapplicable; Acloseranalysisoftheresultsrevealsthatthecauseforthe
4. Brakelocationduringtheinterplanetaryphase(thrustdi- mass variation is the velocity on arrival. Through the change
rectionafterbrakecanbedifferentthanbefore); of the initial angle we induce a variation on the interplanetary
ToFand,mostimportantly,adifferentspacecraftvelocityvector
Forthechemicalpropulsionweconsideracombinationof
whenarrivingatthetargetplanet,affectingthearrivalmanoeu-
twoCECEenginesastheunitpropulsionsystem(i.e.thevalues
vre (and the required propellent), and hence every manoeuvre
ofTable1multipliedbytwo).InthecaseoftheNERVAengine
beforethat.
theunitpropulsionsystemisoneengine. Moreenginescanbe
The fact that the thrust direction is not as important as the
added to increase the available thrust and thus avoid too long
initialangleisexplainedbytherelativelysmallcontinuousthrust
workingperiods(thatleadstohigherfiniteburnlosses).
consideredinthissituation. Higherthrustmagnitudesinducea
For the RIT-XT continuous propulsion engine we consid-
higherdispersionofresultsforthesameinitialvelocityangle.
ered 5, 10, 15, 25, 35, 45, 60, 75, 100, 150, and 250 engines,
AninterestingfactshownbyFig.4istheabsenceofhigher
because of its low thrust and mass. Nevertheless, we must re-
valuesofθ . Forsomevaluesoftheinitialvelocityandthrust
memberthatthereisanassociatedenergysystem(thatincreases v
magnitude, the spacecraft cannot reach the target planet when
the mass), and that more thrust does not necessarily means a
usinghighvaluesoftheinitialvelocityangle.Aclassicalexam-
betterperformance. InthecaseofthePEMTconcept,weonly
pleofthisistheHohmanntransfer,forwhichtheonlyallowed
consideredcombinationsof1,2,3,and4enginesduetoitshigh
anglefortheinitialvelocityiszero.
mass.
AnotheraspectseeninFig.4isthattherearecombinations
Fortheescapevelocity,westartedwiththeonecorrespond-
of angles which yield missions with higher system mass and
ing to the Hohmann transfer (HT) and searched for solutions
ToF (e.g. a mission with θ = 0◦ has a higher mass and ToF
withexponentiallyincreasingvalues(v =[vHT(1+0.02)i]2, i= v
∞ ∞ than one with |θ | = 20◦). As we aim to minimise both, these
0,1,2,3,...,200).Onthereturntransfer,wewouldinprinciple v
missionscanbediscarded. Therefore,foreachescapevelocity
obtainretrogradeinterplanetarytransferorbits,whenthevalue
ondeparturewecomputeallθ andθ casesandselectthebest
of(cid:126)v∞ is larger than the velocity of the planet. This case has using a Pareto efficiency critevria (i.eF.Tby ordering the cases in
noadvantagesrelativelytotheequivalentdirecttransfer,andit
ascendingorderoftimeandthenofmass,thepointsthatshow
willnotbeconsidered.
equalorhighertimeswithhighermassesarediscarded).
Regarding the angles of both the escape velocity with the
Resultsshowthatmostmissionsselected,asmentionabove,
velocityoftheplanet(θ ),andofthecontinuousthrustinthein-
v involve braking during the interplanetary phase (less than 2%
terplanetaryphase(ifapplicable)withthevelocityvector(θ ),
FT do not include a brake), being more probable for the brake to
weconsidervaluesfrom0◦ to−90◦ witha−5◦ step. Theneg-
startafterthemiddleofthetransfer,inaccordancetothelitera-
ative sign indicates that the velocity or thrust makes a retro-
ture[34].
grade angle with the velocity of the planet and the velocity of
the spacecraft, respectively. We only considered negative an-
4.2. ChemicalversusNuclearThermalpropulsion
glesastestsdemonstratedthat,allthingsequal,positiveangles
leadtoworstperformance,asexpected[34]. Inthearchitecturesweconsidered,animpulsivepropulsion
Brakingisdefinedasnegativethrustintheconsidereddirec- system is always required, as it is needed when performing
tion, witch is equivalent to a (positive) thrust at an angle 180◦
9
150
θ
140 v = 70 θ EMaarrths--ttoo--EMaartrhs ((VV∞∞ == 67,,510293 mm//ss))
v =
130 65
ay] vθ =
Time [d 111012000 60 vθ = 55 vθ = 50 vθ = 45 vθ = 40 vθ = 35 vθ = 30 vθ = 25 vθ = 20 vθ =vθ = 10vθ = 0vθ = 5
1
90 5
0 200 400 600 800 1000 1200
Mass [tonne]
Figure4: InterplanetarytimeofflightversusmannedtotalspacecraftmassfortenRIT-XTengines,withconstantinitialvelocityandmultiplevelocityandthrust
angles
coast interplanetary transfers, or by the continuous thrust sys- differencebetweentheminimumandmaximumvaluesisvery
tems for the escape and capture manoeuvres (as explained in small).
section3.1). Therefore,thefirsttestistounderstandwhichof We compare the selected RIT-XT engine with possible fu-
thesystems(chemicalornuclearthermal)haveabettermission turetechnologieswithincreasedspecificimpulse I orthrust-
sp
performanceforsimplecoasttrajectories(sincethecontinuous to-weightratio F /w,asexplainedinsection2.3. Forthepur-
T
thrustmaybeseenasanadditiontothesesolutions). poses of understanding the possible gains of such enhanced
ResultsshowthatthepurechemicalCECEpropulsionsys- electric propulsion technology, we just computed the Earth to
tem performs better than the nuclear thermal NERVA propul- Mars transfertime offlight of themanned spacecraft. In each
sion, for total mission masses lower than ∼ 1,250t. Between case we selected the best combination of electric engines (in
about 1,250t and about 1,600t both options are fairly simi- terms of number of engines), resulting in the shortest ToF for
lar. Whereas, fortotalmissionmasshigherthanabout1,750t theselectedmassofthespacecraft.Resultsforamannedspace-
theNERVAsystemyieldsslightlybetterresults(differencesin craft with 1,000t can be found in Table 2 and show that, al-
missionmassoflessthan22%ofthetotalmissionmass,forthe though the considered values of specific impulse and thrust-
samemissiontime). to-weight of the enhanced technology are considerably larger,
Up to the mission mass order of magnitude for which the theperformancechangeismarginal,andcanevenbedegraded.
NERVA was designed for, about 1,000t, the CECE smaller Thesameresultisobtainedforothermasseswithintheoverall
propulsion system mass outperforms the NERVA. Afterwards, rangeofinterest.
the∼ 34toftheNERVApropulsionsystemrepresentlessthan Our interpretation of these results is that continuous thrust
3.4%ofthemissionmass,anditshigher I andthruststartto isnotthebestchoiceofpropulsiontechnologyfortheproposed
sp
haveapositiveimpactonthetotalmissionmass. Theseresults goalofminimisingtheToFforthisarchitecture. Lowintensity
alsosuggestthateveniftheperformancecouldbeenhancedas continuousthrustrequirestimetobeeffective, buttodiminish
promisedbythePhoebusproposal[25],nosignificantimprove- thetraveltimeisexactlythegoal. Asthetechnologyimproves,
mentwouldbeachieved. by improving the specific impulse or the thrust-to-weight ra-
Considering these results, we only used chemical propul- tio, the travel time diminishes, making the propulsion less ef-
sion for departure and capture of the continuous thrust mis- fectiveandrequiringmorechemicalfuelforthecapture. This
sions. Additionally, chemical propulsion has the advantage of agreeswiththeexistingliterature: keepingthethrust-to-weight
beingsimpler,andmoreenvironmentallyfriendly,thannuclear ratioconstantandincreasingthespecificimpulseincreasesthe
propulsion[24,52]. interplanetarytransfertimewhiletheinterplanetarypropellant
massdecreases(i.e. changingthevalueofI hasapositiveef-
sp
4.3. ElectricEngines fectonthemass,butanegativeoneonthetransfertime)[34].
Forthecaseofelectricpropulsion,wetestedwhichconfig- Whereastheoppositeeffectforthetransfertimeandpropellant
uration(intermsofnumberofengines)yieldsthebestresults. mass is seen when the Isp is fixed and the thrust-to-weight is
Asthedifferentengineconfigurationsleadtosimilaroutcomes increased[34]. Thesurprisinglysmallincreaseinperformance
wedecidedtodeterminewhichonehasthelowestParetocurve suggests that large gains cannot be achieved even with future
on average (for mission masses lower than our defined limit), technological developments of the continuous propulsion. We
usingatrapezoidalintegration. FortheRIT-XTenginethecon- thereforefocusouranalysisontheRIT-XTenginewhencom-
figurationwithlowestvalueinvolves25engines(althoughthe paringwithothertypesofpropulsion.
10