AAS 14-366

PRELIMINARY MISSION DESIGN FOR A CREWED EARTH-MARSFLYBY MISSION USING SOLAR ELECTRIC PROPULSION (SEP)

Stijn De Smet∗, Jeffrey S. Parker†, Jonathan F.C. Herman‡and Ron Noomen§

This paper discusses the preliminary design of a crewed mission to fly by Mars andreturn within 501 days using solar electric propulsion (SEP). The research demon-strates that new launch windows can be opened that would have been impossibleto achieve using conventional chemical propulsion with a reasonable payload andpresent launch vehicles. Furthermore, this paper will also investigate to what ex-tent applying SEP can minimize the launch mass or the re-entry velocity. SEPsystems are considered that use anything from 10 to 25 kW of power.

INTRODUCTION

A human visit to Mars is a challenging endeavor for many reasons. Considering the inherentrisk of such a mission, it is not unlikely that the first missions will attempt to arrive on free-returntrajectories, giving margin for the crews to return safely to Earth in the case of a critical systemfailure. Besides governmental interest for a visit to Mars, a number of private companies havesimilar plans. Based on the promising discussion of free-return trajectories to Mars discussed inReference 1, Inspiration Mars wants to perform a crewed Mars flyby mission launching as soon as2018.2 At the moment of writing, Inspiration Mars only considers the use of chemical propulsion.Taking into account that this mission relies on a very specific planetary alignment to achieve afeasible final payload mass, the next launch window would be in 2031.2 This paper shows thatsolar electric propulsion (SEP) could be used to open up more launch windows in 2018, 2019 and2021. Furthermore, SEP could be used to reduce requirements on other systems: it is demonstratedthat the return velocity can be decreased upto 1.5 km/s and that the required upper stage / launchvehicle performance can be lowered using realistic assumptions on SEP capabilities by the end ofthe decade.

METHOD

Low-thrust trajectory optimization can be done using a multitude of different methods and toolssuch as SEPTOP/VARITOP,3 Sims-Flanagan/MALTO,4 Mystic5 etc, each with their own level ofcomplexity and accuracy.6 Being a preliminary, proof-of-concept study, the speed with which thedeveloped software can scan the search space is considered more important than the level of fidelityof the method. Therefore, the usage of a fast, low-fidelity low-thrust optimization procedure is

∗MSc Student, Delft University of Technology, Delft, The Netherlands.†Assistant Professor, Colorado Center for Astrodynamics Research, University of Colorado, Boulder, CO 80309.‡Graduate Research Assistant, Colorado Center for Astrodynamics Research, University of Colorado, Boulder, CO 80309.§Assistant Professor, Delft University of Technology, Delft, The Netherlands.

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preferred for this study. The selected optimization procedure is known in literature as the Sims-Flanagan method.4 The most well-known implementation of this method is JPL’s MALTO tool. Aself-written version of the Sims-Flanagan method has been developed, and verified using existinglow-thrust tools.

Trajectory representation

The Sims-Flanagan method discretizes the thrust profile using multiple impulsive manoeuvres asan approximation of a continuous thrust profile. The trajectory is cut up into different legs, whichare bounded by control nodes that allow for a constrained discontinuous state. Such a control nodecan have any physical meaning such as a rendez-vous or a flyby of a celestial body, a probe beingreleased in deep space, etc. In this study, the control nodes represent the spacecraft’s encounterswith planetary bodies. The control nodes bounding the first leg represent the launch from Earth andthe arrival at Mars. Between the first and second leg of the trajectory, a Martian flyby is modeled asan instantaneous change in the hyperbolic excess velocity. The control nodes bounding the secondleg represent the departure from Mars and the arrival at Earth. This has been visualized in Figure 1.

Figure 1: Structure of the Sims-Flanagan formulation on a generic trajectory (adapted from J.Sims4)

Each of those legs is discretized into segments. The thrust on a segment is represented by animpulsive manoeuvre at the midpoint of that segment, of which two have been depicted in Figure 1.These impulsive manoeuvres influence the forward and backwards propagation of the leg startingrespectively at the initial control node and the arrival control node of the leg. These two propagations

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meet each other in the middle of the leg at a so called matchpoint. The heliocentric coordinates andvelocities, the hyperbolic excess velocities and the spacecraft mass at the control nodes are usedas the starting point for these propagations. These initial conditions are then propagated usingthe control parameters on each segment, which are the magnitude and direction of the manoeuvrealong with the specific impulse. The propagation between the impulsive manoeuvres is done usinga two body model and an RK7(8)13M integrator, also known as DOPRI8.7 These forward andbackward propagations have to be consistent in heliocentric coordinates, velocities and mass at thematchpoints in order to ensure a continuous trajectory at the matchpoints. This propagation structurehas been visualized in Figure 15 in the Appendix.

Optimization variables

In this study, the launch date, flyby date and arrival date are kept constant for the optimization of asingle trajectory. These dates can then be translated into the control nodes’ heliocentric coordinatesand velocities using Meeus’ polynomials,8 which remain fixed throughout the optimization of thistrajectory. The available power and the dry mass of the spacecraft are also kept constant. This drymass is the payload of the mission, based on the Inspiration Mars system.9 As such, the final massat Earth return, which can be found using the following equation, is fixed:

MEarthreturn = Mdry +MSEP

= Mdry + P0 · kP0 (1)

where MSEP is the mass of the SEP system, including the power supply (solar panels) as well asthe propulsion system itself. kP0 is the power to mass ratio of the SEP system and is assumed to be30 kg per kW.10 P0 is the the available power for the SEP subsystem at a heliocentric distance of 1AU.

The parameter being minimized is the mass at Earth departure: the launch mass. The parameterswhich can be changed to minimize the launch mass are: the mass at Mars arrival, the mass atMars departure (which has to be equal to the final mass at Mars arrival), the departure hyperbolicexcess velocity at the Earth’s launch, the incoming and outgoing velocity vector at the Mars flyby,the incoming hyperbolic excess velocity at arrival at Earth and the magnitude and direction of themanoeuvres. The specific impulse remains constant and has been set to 2000 s.

Constraints

In order for a trajectory to be feasible, it must meet the imposed constraints. The constraints on atrajectory for the Sims-Flanagan method depend on the type of control nodes on the legs. However,some constraints are present for all types of control nodes. These will be explained first, followedby an explanation of constraints specific for this study.

Matchpoint constraints. In each Sims-Flanagan problem, constraints must be imposed such thatthe heliocentric coordinates and velocities and the spacecraft mass from the forward and backwardspropagation at the matchpoint are equal. These will be called the matchpoint constraints and are ofthe form

− ε < αforward − αbackwards < ε (2)

where α represents the heliocentric coordinates, velocities and the mass at the matchpoint of theforward and backwards propagation and ε represents a small value.

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Thrust constraints. The magnitude of the manoeuvre applied at the midpoint of each segment islimited. It must be ensured that the magnitude of this manoeuvre does not exceed the maximum thatthe spacecraft can provide during the duration of that specific segment with a certain power level.Note that due to the usage of solar electric propulsion, the available power on each segment dependson the heliocentric distance. The thrust constraints can be found by combining Equations (3) and(4) into Equation (5):

Pjet = ηjet · P =1

2TIspg0 (3)

where Pjet is the power of the exhaust jet, P is the available power, ηjet is the power conversionefficiency assumed to be 60%11 and T the thrust

T =∆Vimi

DT(4)

with DT the duration of the segment, ∆Vi the size of the manoeuvre and mi the mass of the systembefore the manoeuvre.

Pjet < ηjetP0AU2

R2i

∆VimiIspg02DT

< ηjetP0AU2

R2i

∆VimiR2i

P0<

2ηjetAU2DT

Ispg0(5)

where P0 is again the available power at a heliocentric distance of 1 AU and Ri is the heliocentricdistance.

Study specific constraints. Besides the general constraints, there are also some case specific con-straints. At launch from Earth, constraints are active to ensure that the C3 and launch mass matchthe performance of the selected launch vehicle. The Martian flyby also adds specific constraints.The incoming and outgoing relative velocities must be equal in magnitude, as must the mass bebefore and after the flyby. Furthermore, a restriction on the pericenter altitude of the flyby has beenset to ensure that the spacecraft does not impact Mars or enters Mars’ atmosphere:12

rperiapse − (rMars + hsafety) ≥ 0

µMars

V +2

∞

[1

sin(δ2

) − 1

]− (rMars + hsafety) ≥ 0 (6)

where µMars is the standard gravitational parameter of Mars, V +∞ is the outgoing hyperbolic flyby

velocity vector, hsafety is a safety altitude of 200 km above Mars’ surface and δ is the deflectionangle defined by

δ = acos

[ −→V −∞ ·

−→V +∞

|V −∞ | ∗ |V +∞ |

](7)

with V −∞ the incoming hyperbolic flyby velocity vector.

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Finally, a constraint on the re-entry velocity of 14.2 km/s at 200 km altitude has been imposed.As mentioned before, this has been translated using the vis-viva equation to a constraint of 8.969km/s on the incoming hyperbolic excess velocity :

V 2re−entry

2− µEarthrre−entry

=V 2∞2

(8)

Optimization procedure

In order to find the optimal solution to the formulated problem, the Sparse Nonlinear OPTimizer(SNOPT)13 has been used. SNOPT uses the gradient of the constraints with respect to the statevariables to find feasible and optimal solutions. SNOPT has the option to calculate these gradientsusing finite differencing. However, to increase the speed of the optimization procedure, it was optedto find analytical expressions for all these derivatives. These are lengthy formulas and as such,are not included in this paper. The majority of these equations were presented in Reference14.Furthermore, SNOPT requires proper scaling of the problem. This discussion is again not includedin this paper, but can be found in Reference 14.

In simple cases, the optimization procedure does not require an initial guess of the state vector.However, since the optimization procedure does not optimize for launch, flyby and re-entry date,it was decided to first narrow down the search space by finding feasible, classical chemical thrusttrajectories using Copernicus.15 These trajectories were then used as an initial guess. This approachallows for a faster converging towards a solution and for a quick assessment of the search space.

Based on this initial guess trajectory generated with Copernicus,15 a two dimensional grid searchhas been performed where the launch date and flyby date vary. The re-entry date at Earth is thendetermined by the total mission duration length of 501 days. This grid search demonstrated whichcombinations of launch date and flyby date were feasible and as such determined several launchwindows for this mission. These launch windows will be further elaborated upon in the Resultssection.

Launch vehicle

The selected launch vehicle has a significant impact on the launch windows. In order to do a faircomparison with the chemical launch window identified for Inspiration Mars,9 the Space LaunchSystem vehicle will be used for this study. Throughout this paper, two different configurations of theSpace Launch System vehicle are considered. Both configurations use liquid oxygen and hydrogenand have a specific impulse of 462.5 s.16

The first configuration is the SLS/iCPS 1xRL10B2. This stands for Space Launch System InterimCryogenic Propulsion Stage which uses one RL10B2 engine. The SLS/iCPS 1xRL10B2 is alsoknown as the Block 1 configuration and is a derivative of the Delta-IV 5m upper stage presently inproduction. This configuration can lift 70.0 tons into LEO.16

The second configuration is the SLS/LUS 4xRL10C2. This stands for Space Launch SystemLarge Upper Stage which uses four RL10 C2 engines. This configuration can lift 93.1 tons intoLEO.16

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Important design parameters and assumptions

In order to obtain realistic results, several design parameters had to be obtained or assumed.These have been summarized in Table 1.

Table 1: Important design parameters and assumptions

SEP and method related parametersPower to mass ratio SEP system10 30 kg/kWJet efficiency11 60%SEP system duty cycle17 90%Specific impulse 2000 sNumber of segments entire mission 120 segments ≈ 4 days/segmentMission related parametersLaunch vehicles16 SLS/iCPS 1xRL10B2

SLS/LUS 4xRL10C2Total mission duration9 501 daysMaximum re-entry velocity9 14.2 km/s

Operational considerations

Being a crewed mission, the solution must be made robust and operationally achievable. There-fore, some operational considerations have been made.

First of all, it was decided to put a higher level of restriction on the thrust level. Normally, onewould expect the highest allowable thrust level to be the maximum achievable thrust level on thatsegment. However, it has been decided to restrict it to 90 percent of what is actually achievable.Such a margin has been shown to effectively prevent negative consequences of missed thrust.17

Furthermore, this 90% does not only account for missed thrusts, it also accounts for tasks that mayinterfere with thrusting periods such as uploads & maintenance and communications & tracking.17

Furthermore, it is unwise to have manoeuvres near planets. The reason from an operationalpoint being to avoid additional tasks to be performed during the critical phases near the planetsduring the mission. From an astrodynamical point of view, it is important to avoid a dependencyon manoeuvres near planets. Manoeuvres near planets have a much larger effect than heliocentricmanoeuvres. If the thrusters fail during these crucial manoeuvres, the system might not be able torecover from this failure. Therefore, 3 coast arcs have been imposed on the trajectory. A 5 daycoast arc has been imposed upon leaving Earth to account for the early checkout phase, a coast arc2 weeks prior and 2 days after the Martian flyby for pre- and post-flyby operations and a final coastarc of 2 weeks prior to arrival at Earth for re-entry operations.

As mentioned above, there are also considerations with respect to the flyby altitude. The flybyaltitude has a lower bound of 200 km since it must not impact Mars, but also avoid the Martianatmosphere. Besides this lower bound, also an upper bound has been imposed. Considering that thepurpose of this mission is to be the first human visit to Mars, it is assumed to be desirable to makea close visit to Mars. Therefore, the upper bound has been set at an altitude of 2000 km. In practicehowever, the majority of the solutions does not approach this upper bound.

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Finally, the spacecraft has to be able to withstand the heat load, which is strongly related to there-entry velocity. Inspiration Mars sets a boundary at 14.2 km/s or better.9 This has been translatedto a hyperbolic excess velocity limit of 8.969 km/s.

The combination of all these margins make all of the presented solutions extremely robust, inorder to ensure the safe return of the crew.

RESULTS

Using Copernicus, 3 nominal chemical thrust scenarios in 2018, 2019 and 2021 have been identi-fied. These have been summarized in Table 2 including the magnitude of the Deep Space Maneuvers(DSMs), if applicable. Note that these scenarios have not been fully optimized yet. They were op-timized to the point of being a feasible starting point for the SEP design.

Table 2: Nominal scenarios identified using Copernicus

Launch date Launch C3 DSM 1 Flyby date DSM 2 Return V∞(mm-dd-yyyy) (km2/s2) (m/s) (mm-dd-yyyy) (m/s) (km/s)

01-05-2018 37.45 N.A. 08-20-2018 N.A. 8.8912-13-2019 58.37 647.31 09-02-2020 517.28 6.1012-10-2021 21.44 1650.41 09-18-2022 2227.85 5.17

Around each nominal scenario, a grid search will be performed using these nominal scenarios asan initial guess. An overview of all these cases has been given in Table 3. In the next subsections,the grid searches for these cases will be analyzed.

Table 3: Cases around which grid searches have been done

Case Scenario Power level Payload mass Launcher configuration(kW) (tons)

1 2018 10 19 SLS/LUS 4xRL10C22 2018 10 13.139 SLS/LUS 4xRL10C23 2018 10 13.139 SLS/iCPS 1xRL10B24 2018 25 15 SLS/iCPS 1xRL10B25 2019 10 13.139 SLS/LUS 4xRL10C26 2021 25 13.139 SLS/LUS 4xRL10C2

The results of these optimization runs have been verified using an independent propagation. Thispropagation was fed with the optimized output to verify that it corresponded to a feasible trajectory.

2018

Using classical chemical propulsion, a launch window of 12 days has been identified betweenDecember 14, 2017 and January 4, 2018 for Inspiration Mars.9 This launch window takes intoaccount a fully margined payload mass of 19 tons.9

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Using low-thrust SEP, this launch window can be substantially increased for a payload mass of 19tons using the SLS/LUS 4xRL10C2 launch configuration. Remember that this payload mass doesnot include the mass of the propulsion subsystem. Figure 2 shows the grid search which has beenperformed around the nominal scenario. On the horizontal axis, the variation of the launch datew.r.t. the launch date of the nominal scenario has been plotted. On the vertical axis, the flyby datehas been plotted w.r.t. the nominal scenario launch date. The color scale represents the minimizedlaunch mass in tons for the combinations of launch and flyby dates. The largest dot represents thecombination of launch and flyby date resulting in the lowest minimized launch mass. As can beseen from Figure 2, this launch window would be increased to around 57 days between the 8th ofDecember, 2017 and the 3th of February, 2018, using only 10 kW.

Figure 2: Launch opportunities for Case 1: 19 tons, 10 kW and 4RL10 upper stage stage

The final mass at Earth arrival, which is the sum of the payload mass and the mass of the SEPsystem, is fixed for a certain grid search. For Case 1, the final mass is 19.3 tons of which 19 tonspayload mass and 300 kg SEP subsystem mass. Since this mass is fixed, the difference in launchmass in Figure 2 can be attributed to the difference in propellant mass. One can see from Figure2 that most data points have a launch mass quite close to this 19.3 tons, indicating that there arehardly any manoeuvres. This does not really came as a surprise, since the reference trajectory fromCopernicus in 2018 is a free return trajectory. An example of such a trajectory with hardly anymanoeuvres can be seen in Figure 3.

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Figure 3: Trajectory of a scenario with not many manoeuvres

Figure 4: Thrust profile of a scenario with not many manoeuvres

Some data points however are further away from this 19.3 tons. Those data points representtrajectories that require more manoeuvres. An example of such a trajectory has been plotted inFigure 5. In Figure 6, one can also observe the effect of the coast arc near the Martian flyby.Around 0.65 years, the coast arc is clearly visible.

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Figure 5: Trajectory of a scenario with many manoeuvres

Figure 6: Thrust profile of a scenario with many manoeuvres

The launch window could be much larger if it is assumed that the vehicle can be constructedmuch closer to the baseline mass than the fully margined mass. To illustrate the potential impactof this, Case 2 is identical to Case 1 but instead uses the baseline mass of 13139 kg.9 The launchwindow could then be increased to 124 days between the 10th of October, 2017 and the 11th ofFebruary, 2018, again using only 10 kW. This can be seen in Figure 7.

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Figure 7: Launch opportunities for Case 2: 13.139 tons, 10 kW and 4RL10 upper stage

The previous two cases require the more complex LUS 4xRL10C2 upper stage. It is not entirelysure that this upper stage will be finished by 2018. Therefore, it has been investigated if the baselineSLS/iCPS 1xRL10B2 upper stage configuration can be used for this mission. Using the baselinepayload mass of 13139 kg and using only 10 kW, the launch window for this configuration wouldbecome 44 days between the 25th of December, 2017 and the 7th of February, 2018, as can be seenin Figure 8.

Figure 8: Launch opportunities for Case 3: 13.139 tons, 10 kW and 1RL10 upper stage

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44 days is still a relative big launch window. Therefore, the payload mass has been graduallyincreased to identify what the largest mass would be with a sufficient launch window of at least 20days. For a payload mass of 13.5 tons, the launch window has decreased to 33 days between thefirst of January 2018 and the 3th of February 2018. For a payload mass of 13.75 tons, this decreasesto a 21 day launch window between the 16th of January, 2018 and the 6th of February, 2018. Fora payload mass of 14 tons, no launch window at all could be identified. This distribution betweenpayload mass and size of the launch window behaves clearly non-linear.

One could further increase the payload mass if the available power for the SEP system is higher.Considering that NASA states that the top technical challenge for in-space propulsion is the devel-opment of high-power electric propulsion system technologies to enable high ∆V missions withheavy payloads,18 it is not unreasonable to assume that power levels beyond 10kW will be achiev-able in the near future. As an example, a payload mass of 15 tons with 25 kW of available power isused. As can be seen in Figure 9, a launch window of 36 days between the 10th of January, 2018and the 15th of February, 2018 can be opened up. This payload mass can be increased upto 15250kg for a launch window of 22 days between the 16th of January, 2018 and the 7th of February, 2018.

Figure 9: Launch opportunities for Case 4: 15 tons, 25 kW and 1RL10 upper stage

A second set of mission designs have been examined with the goal of minimizing the re-entryvelocity at Earth, rather than the launch mass. This has been done for several cases, of which themost interesting cases will be highlighted. For Case 2, the result can be seen in Figure 10. Note thatinstead of the launch mass, the color bar now represents the re-entry velocity in km/s. Comparingthis Figure to Figure 7, one can see that the launch windows in both cases are almost identical.

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Figure 10: Launch opportunities for Case 2: 13.139 tons, 10 kW and 4RL10 upper stage for mini-mized re-entry velocity

Figure 11: Relation between required launch window length and design re-entry velocity for Case2: 13.139 tons, 10 kW and 1RL10 upper stage

Normally, a mission has to be designed around its most constrained parameters. For this launchwindow, this would mean that the mission should be designed around 14.2 km/s if one would like tobe able to utilize the entire launch window. However, this launch window is quite large. Therefore,if one would accept a smaller launch window around the minimal re-entry velocity, the maximumencountered re-entry velocity in this smaller launch window would be smaller. Obviously, the

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smaller the acceptable launch window is, the more the re-entry velocity can be reduced. This hasbeen visualized in Figures 11 and 12 for Case 2 and 4, in which one can see the departure datew.r.t. the nominal scenario launch date on the horizontal axis. On the vertical axis, the re-entryvelocity in km/s can be found. For each departure date, the minimal encountered re-entry date hasbeen plotted. For different sizes of the launch window, different limiting re-entry velocities can beidentified. For the different run cases, the results have been summarized in Table 4, showing thatthe re-entry velocity can be reduced to as little as 12.67 km/s.

Figure 12: Relation between required launch window length and design re-entry velocity for Case4: 15 tons, 25 kW and 1RL10 upper stage

Table 4: Results 2018 cases optimized w.r.t. the re-entry velocity Vre

Case Launch window Min Vre Max VreMax Vre 20 dayslaunch window

Max Vre 30 dayslaunch window

(days) (km/s) (km/s) (km/s) (km/s)1 57 13.33 14.19 13.37 13.422 124 12.63 14.18 12.66 12.673 44 13.55 14.19 13.62 13.724 36 13.89 14.19 13.95 14.02

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2019

First of all, it has been investigated if the baseline SLS/iCPS 1xRL10B2 upper stage configurationcan be used for this mission. However, the 1RL10 configuration could not provide enough energyto even launch the baseline payload mass of 13139 kg. In order to establish how high the payloadmass could be while still having a feasible launch window of at least 20 days, several cases havebeen run. Using 10 kW, the maximum payload mass would be 9700 kg for a launch window of 23days. Obviously, this mass is not even close to the required 13139 kg. Therefore, the same scenariohas been run with a higher power level. Using a power level of 25 kW, the maximum achievablepayload mass is 11830 kg with a 24 day window between the first of November, 2019 and the 25th

of November, 2019.

As this shows that the 1RL10 has great difficulty with the 2019 scenario, cases using the 4RL10configuration have been run. For 13139 kg and using 10 kW, the launch window becomes 108 daysbetween the 22nd of November, 2019 and the 9th of March, 2020 as can be seen in Figure 13.Obviously, there is quite some margin to increase the payload mass. Therefore, the payload masshas been gradually increased until a launch window of approximately 20 days was encountered.This was the case for 15750 kg, in which a 20 day launch window has been identified between the1st of February, 2022 and the 21st of February, 2022.

Figure 13: Launch opportunities for Case 5: 13.139 tons, 10 kW and 4RL10 upper stage

2021

Much like the 2019 scenario, the 1RL10 configuration has great difficulty with the 2021 scenario.Therefore, cases using the 4RL10 configuration have been investigated. It was found to be impos-sible to launch the baseline payload mass of 13139 kg using 10 kW. Therefore, the power level hadto be increased to 25 kW to make this scenario feasible. In the end, a launch window of 33 days canbe observed in Figure 14 between the 20th of November, 2021 and the 23th of December, 2021.

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If the payload mass is increased to 13.3 tons, the window becomes 22 days between the 26th ofNovember, 2021 and the 18th of December, 2021.

Figure 14: Launch opportunities for Case 6: 13.139 tons, 25 kW and 4RL10 upper stage

FUTURE WORK

The presented results could be further improved. There may be a significant potential for im-provement when the mission duration is allowed to change. So far, it has always been kept fixed at501 days, but allowing variable flight time might create more opportunities. Therefore, the capabil-ity to optimize the epochs of important events should be added. Furthermore, higher power levelscould be investigated along with lower and higher specific impulses. These specific impulses couldalso be made variable along the trajectory to improve current results. Further improvement couldbe achieved by investigating the effect of different launch vehicles.

Besides improving the current results, additional studies could be performed such as investigatingadditional flight opportunities in the 2020s/2030s. Furthermore, different mission concepts could beinvestigated. For instance, an additional Venus flyby could be added to increase the scientific returnof the mission and to open up new launch windows. Another possible follow-up investigation couldbe the usage of SEP to adjust the flyby geometry. Another possible study is to impose more specificconstraints on thrusting. For example: thrusting only before the Mars encounter, to ensure the flybycan enter a free-return that requires no further thrusting.

CONCLUSIONS

In this paper, it has been shown that solar electric propulsion can be used to significantly improvecrewed flyby missions of Mars. Using a modest amount of solar electric propulsion, less perfor-mance is required by the upper stage of the launch vehicle, giving more margin for developmentof the system, and possibly allowing for the use of a smaller upper stage altogether. Alternatively,re-entry velocities can be reduced, which reduces the system requirements for the re-entry system.

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Allowing launch windows of 20 days, the reduction in re-entry velocity w.r.t. the launch massoptimization cases were between 0.8 and 1.5 km/s.

In order to give a quick overview of the capabilities of SEP, the combination of parameters thatlead to preliminary launch windows of 20 days have been listed in Table 5.

Table 5: Limiting cases: 20 days launch window

Scenario Power level Launcher configuration Payload mass(kW) (tons)

2018 10 SLS/LUS 4xRL10C2 >192018 10 SLS/iCPS 1xRL10B2 13.752018 25 SLS/iCPS 1xRL10B2 15.252019 10 SLS/iCPS 1xRL10B2 9.72019 25 SLS/iCPS 1xRL10B2 11.832019 10 SLS/LUS 4xRL10C2 15.752021 25 SLS/LUS 4xRL10C2 13.30

This study has demonstrated that SEP improves the launch period and mission performance of amission to fly by Mars and return within 501 days. Many parameters may still be adjusted, whichmay open up further mission concepts and improve the performance even more. This warrantsfurther study, that would undoubtedly create even further improvements on those discussed here.

ACKNOWLEDGMENT

We would like to thank Mike Loucks for providing relevant information about the InspirationMars mission and about the Space Launch System.

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APPENDIX: PROPAGATION STRUCTURE

Figure 15: Structure of the propagation

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[17] D. Oh, J. Snydert, D. Goebel, R. Hofer, D. Landau, and T. Randolph, “Solar Electric Propulsion forDiscovery Class Missions,” 33rd International Electric Propulsion Conference, October 2013.

[18] M. Meyer, L. Johnson, B. Palaszewksi, D. Goebel, H. White, and D. Coote, “In-Space PropulsionSystems Roadmap - Technology Area 2,” tech. rep., NASA, April 2012.

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