, (2) German Aerospace Center (DLR), Porz-Wahnheide, Linder Hoehe, Koeln Germany,

Transcription

1 TRAJECTORY AND SYSTEM DESIGN OF AN ELECTRICALLY PROPELLED ORBITER FOR A MARS SAMPLE RETURN MISSION Uwe Derz (1), Wolfgang Seboldt (2) (1) EADS Astrium Space Transportation, Airbus Allee 1, Bremen Germany, , uwe.derz@astrium.eads.net (2) German Aerospace Center (DLR), Porz-Wahnheide, Linder Hoehe, Koeln Germany, , wolfgang.seboldt@dlr.de ABSTRACT The Mars Sample Return (MSR) mission is a flagship mission within ESA s Aurora programme and envisioned to take place in the timeframe of Corresponding studies developed a mission architecture consisting of two elements, an orbiter and a lander, both utilizing chemical propulsion and a heavy launcher like Ariane 5. The lander transports an ascent vehicle to the surface of Mars. The orbiter performs a separate transfer to Mars, conducts a rendezvous in Mars orbit with the sample container, delivered by the ascent vehicle, and returns the samples back to Earth in a small entry capsule. The focus of the present paper is on a mission and system analysis of a respective orbiter with low thrust electric propulsion. For each mission phase, i.e. Earth escape, Earth-Mars transfer, Mars capture, Mars escape and Mars-Earth transfer, the most important factors are investigated. The system analysis evaluates their influence on the launch mass, total mission time and time in Mars orbit. It has been found that a low thrust electrically propelled orbiter, launched by a Soyuz-Fregat into GTO, can conduct its mission within days. At the end as a supplement also results derived by the authors concerning a complete MSR mission with only one heavy launch vehicle (e.g. Proton) and the lander transported piggyback by the electrically propelled orbiter are included. 1 INTRODUCTION Since the 60ties of the last century orbiting and landing space probes are sent to Mars to search for extraterrestrial life and the Martian planetary evolution, allowing conclusions on the development of our own planet as well as the history of the solar system. Unfortunately, the scientific instruments, transported to Mars, have to fulfill enormous requirements. They should be as accurate as possible and should consume only small masses and volumes. In addition they have to withstand the high mechanical loads during launch, atmospheric entry and landing on Mars. Due to the large signal time delay between Earth and Mars, the instruments have to operate highly autonomously. A further problem is that in case new instruments become available, they should be sent to all interesting, already visited landing sites by new missions. This would result in high costs and is therefore not a practical solution. Hence, the Mars sample return mission became a flagship mission within ESA s Aurora programme. 2 MISSION OUTLINE AND BASELINE SCENARIO Previous studies [7][11][9][1] developed a mission architecture consisting of two elements, an orbiter and a lander, each utilizing chemical propulsion and a heavy launcher like Ariane 5. The lander transports sampling equipment (e.g. a rover) and a Mars ascent vehicle with a sample container to the Martian surface. After completion of surface operations, the samples are transferred to the sample container, which is launched by the ascent vehicle into Mars orbit. The orbiter, performing a separate impulsive chemical transfer to Mars, conducts a rendezvous in Mars orbit with the sample container and returns the samples back to Earth in a small Earth entry capsule.

2 Table 1: Mission characteristics for the orbiter element of the chemical reference scenario [9] Launcher Ariane 5 ECA Earth-Mars transfer 641 d including an Earth swing by Mars orbit injection chemical propulsive and aerobraking from an elliptic 24 h orbit Time in Mars orbit 533 d including aerobraking time of about 100 d Mars-Earth transfer 277 d Total mission time about 1460 d The focus of the present paper is on the orbiter element with a mission and system analysis, considering low thrust propulsion to minimize the launch mass. As reference, the conventional chemical orbiter according to [9] is chosen. Major characteristics of this reference configuration are summarized in Tab. 1 for a launch around SPACECRAFT DESIGN Tab. 2 summarises the assumed model of the various spacecraft subsystems. The spacecraft bus includes the avionics, the thermal control system, the part of the power system supporting avionics and thermal control and the corresponding carrying structure. It could be based on Mars Express heritage with an estimated mass of 450 kg. As main propulsion system a cluster of RIT 22 ion thrusters developed by the University Giessen and EADS Astrium is chosen. RIT stands for Radio frequency Ion Thruster. The RIT uses a radio-frequency generator to ionize propellant atoms (Xenon), resulting in an electrical eddy-field inside the ionizer vessel. The charged ions are extracted due to their positive charge and accelerated by the electric field between plasma holder grid and accelerator grid. This has been demonstrated under space conditions by the RIT 22 s precursor engine RIT 10 during the Eureca and Artemis missions. Spacecraft bus Table 2: Spacecraft model 450 kg including attitude control, guidance and navigation, communication, thermal, power and structure for spacecraft bus only Earth return capsule 200 kg including 1 kg samples (upscaled from [9]) Main Propulsion Radio-Frequency Ion Thruster (RIT) 22 [12] Versions low I sp medium I sp Beam voltage [kv] Specific impuls [s] PSCU power consumption [kw] Thrust force (nominal) [mn] Thrust to power ratio [mn/kw] Propellant flow per power [mg/kws] Engine mass [kg] Main engine numbers [-] Power (for propulsion) Structure Triple-junction solar cell (GaInP 2 /GaAs/Ge) Efficiency: 29% at Earth Specific mass: 10 kg/kw Tank and feeding system: 10% of propellant mass Supporting elements: see eq. 3

3 The spacecraft should be provided with power by two solar arrays. Most power is required for the electrical engines. To reduce the mass and to simplify the spacecraft design, it is assumed that the engines are not operated during eclipse phases. Hence, batteries are only required for the spacecraft bus during the eclipse phases and are contained in its mass. The mass of the solar panels and power distribution for propulsion purpose is considered separately and included in the power subsystem mass budget. The electrical power, provided by the solar array is calculated using eq. 1 ( ) P P el,s = A s η s (1) A with: ( P el,s : P ( A) P A) R : Electric power provided by the solar array Solar flux at the distance R from the sun : Solar flux at 1 AU = 1370 W E m 2 A s : Area of the solar array (always assumed to be perpendicular to the sun) η s : Conversion efficiency of the solar array Unfortunately, the conversion efficiency depends on the solar cell temperature and thereby on the solar constant. At large distances from the Sun, the conversion efficiency of relatively cold cells is higher than at Earth distance. To allow the determination of the electrical power, using a constant efficiency, it is helpful to assume for the solar flux in eq. 1 [5]: with R ( ) ( ) P P = A R A E R E : Orbit radius of the Earth R: Actual distance of the spacecraft to the Sun ( RE R ) 1.6 (2) The carrying structure of the remaining spacecraft (without bus and return capsule) is calculated as following: with: m tank : m prop : n eng : m eng : m power : m structure = 0.15 (m tank + m prop + n eng m eng + m power ) (3) Propellant tank mass (10% of propellant mass) Propellant mass Number of electrical engines Mass of one electrical engine Power system mass, including power converter, management and distribution Due to the fact, that only the mass of the samples, the sample container, the Earth entry capsule and the spacecraft bus is known at the beginning of the mission analysis, the mission phases are examined in reverse order. For each mission phase, the required propellant and the corresponding tank/structure masses are determined iteratively and the masses are fitted together w.r.t. the mission phase order. For simplification, it is assumed that the propellant is distributed to several tanks. After each mission phase an empty tank and its corresponding carrying structure is jettisoned.

4 4 MISSION ANALYSIS The mission analysis of the Mars sample return orbiter is divided into following phases: Earth Escape Mars capture and orbit injection Mars to Earth transfer Earth to Mars transfer Mars escape The low thrust trajectories are optimized concerning minimal flight time using the DLR low thrust optimizer InTrance [3][4]. As mentioned earlier, the analysis considers the use of RIT 22 engines with specific impulses of 3704 s (low = lo) and 4763 s (medium = me). Furthermore, two different power systems, introduced in section 4.1, are examined. To simplify the documentation, the spacecraft configurations are abbreviated in the following way: (Engine)-(power system)-(outbound transfer support) Example: me-stpo-booster me: RIT 22 electric engine with an I sp of 4763 s stpo: Standard power system, providing sufficient power to operate all spacecraft engines at 1 AU distance to the Sun booster: The spacecraft uses additional engines for the outbound transfer 4.1 Interplanetary Inbound Transfer (Mars to Earth) To enable the calculation of the inbound trajectories, Earth rendezvous conditions concerning distance and relative velocity (w.r.t. Earth s sphere of influence SOI) have to be defined. It is assumed that the spacecraft jettisons the Earth return capsule during flight within Earth s SOI. Hence, a maximal distance of 10 6 km is assumed. In reality this distance could be drastically reduced by implementing small correction maneuvers at a far distance to Earth. The relative velocity determines the atmospheric entry velocity which is limited by the used thermal protection system and the allowable deceleration loads. The fastest manmade Earth entry vehicle, the Stardust sample return capsule, entered Earth s atmosphere with about 12.6 km [10]. s This results in a maximal relative velocity of about 5.8 km, which can be accepted assuming s today s technology. Figure 1: Inbound low thrust transfer trajectories with low and high relative velocities at Earth, escape dates and flight time: May 31, 2022, 325 d (left); June 15, 2022, 235 d (right)

5 Fig. 1 shows low thrust inbound trajectories with low (1 km) and high (5 km ) relative velocities. s s The corresponding thrust profiles are displayed in fig. 2. The power system is designed to enable all engines to have full thrust at Earth s distance to the Sun. It is further assumed, that initially the spacecraft has escaped the Martian SOI with v inf = 0. Figure 2: Inbound trajectory thrust profile for low and high relative velocities at Earth It can be seen, that in case of a low relative velocity the spacecraft has to match its orbital velocity and thereby its orbit to the one of Earth. This requires continuous thrusting during the complete transfer. Although the magnitude of thrust varies over the time, two "high" thrust phases can be identified. In contrast to that, a transfer, allowing a high relative velocity, performs thrusting only in the first half of the transfer. Hence, this phase must be timed in a way that the spacecraft passes by the Earth within the maximal distance during the unpowered flight. Because the maximum power available for the electrical engines (and thus thrust) increases on the way towards Earth, there is a slight thrust increase visible during the inbound transfer. Figure 3: Flight times and initial masses after Mars escape of low thrust inbound transfers Fig. 3 left displays the flight times for a medium I sp of 4763 s in dependence of the number of engines and the relative velocity. The flight time decreases with increasing number of engines. It has to be pointed out that this decrease is more pronounced for low engine numbers than higher ones. Considering also that the spacecraft mass rises for increasing engine numbers (fig. 3 right), engine numbers above 5-6 appear less desirable. As mentioned earlier, the flight time can also be reduced by increasing the relative velocity at Earth. Regarding relative velocities above 3 km, this effect is relatively small in comparison to the engine number effect (all thrust s profiles are similar to the red solid line in fig. 2). Relative velocities below 3 km are not s

6 considered because they lead to much longer flight times. In principle a similar behaviour can be observed for a low I sp of 3704 s. But due to higher propulsion and power system masses, the overall system masses of medium I sp spacecrafts are higher than those of low I sp s for same engine numbers. Nevertheless, the higher thrust of the medium I sp spacecrafts enables a faster transfer. Figure 4: Flight times of low thrust inbound transfers with low and medium specific impulses and different power systems Figure 5: Inbound trajectory thrust profile for different power systems From analyses done so far it can be concluded that most of the interesting trajectories require all thrusting maneuvers in the initial phase of the return transfer. During this maneuver, the electrical engines cannot provide their full thrust, because the power system can only support full engine thrust at Earth distance. By increasing the power system output for a constant number of engines, each engine is used more efficiently. Hence, the power per engine is varied for both considered I sp s and a relative velocity of 5 km at Earth. The resulting transfer times s are displayed in fig. 4. All flight times show a decreasing behavior until a minimum at a power per engine of 6.75 and 10.4 kw, respectively. Notably, both values are 1.68-times the maximal engine power consumption. Using equation 2, it can be shown, that the power system is able to operate all engines up to a distance of 1.38 AU. This is also the largest distance from the Sun of the transfer trajectory. Hence, the engines can operate at maximal thrust during the complete powered flight phase. In fig. 5 the thrust profiles of 4 engine spacecrafts with and without such an enhanced power system are compared. If the power per engine is increased by more than a factor of 1.68, a further thrust increase is not possible, while the spacecraft dry mass increases, resulting in a flight time extension. In the following, the flight time optimal

7 power configuration (called enhanced power system) is examined in parallel to the standard power system. The corresponding abbreviations are used: lo-enpo: RIT 22 with I sp = 3704 s, power system output: 6.75 kw/engine me-enpo: RIT 22 with I sp = 4763 s, power system output: 10.4 kw/engine 4.2 Mars Escape Figure 6: Low thrust Mars escape trajectory (from 250 km orbit) using 4 RIT 22-me-stpo (Martian SOI marked copper) Beginning from Mars orbit, the spacecraft has to escape Mars gravity before it can perform the interplanetary transfer described in the previous section. The interplanetary transfer calculation assumed that the spacecraft is orbiting initially the Sun at Mars distance with Mars orbit velocity. This means, that the relative velocity between Mars and spacecraft is zero. Therefore, the escape trajectories are calculated until the spacecraft reaches body escape and patched to the interplanetary trajectory. An example of an escape trajectory is displayed in fig. 6. A 250 km circular orbit as initial condition is chosen as baseline, but 1000 km and km circular orbits are addressed later on, too. Figure 7: Low thrust Mars escape flight times (from 250 km orbit) and initial masses in dependence of the number of main engines

8 As it can be seen from fig. 7, Mars escape trajectories show similar characteristics as the interplanetary inbound trajectories. Increasing engine numbers lead to shorter flight times, and this flight time decrease is more pronounced for lower engine numbers. Furthermore, the utilization of an enhanced power system can reduce the flight time significantly. This results from the fact that vehicles with standard power systems have to throttle their engines due to power shortage at this distance. It is also obvious that an engine number increase leads to higher propulsion and power system masses and thereby raises the total spacecraft mass. A further important factor of low thrust Mars escape trajectories is the initial orbit altitude. According to [8], the required V to spiral out from a circular orbit can be estimated using eq. 4: V el,esc = µ µ µ r i = (4) r i with V el : Electric V -demand r i : Radius of the initial orbit µ: Gravitational factor of the centre body r f : Radius of the final orbit (= ) r f Those theoretical V s are calculated for circular orbit altitudes between 250 and km and are displayed in fig. 8. The results show that V decreases with increasing orbit altitudes. Compared to real trajectory calculations with an example configuration considered in this study (4 RIT 22 me-stpo, marked with blue dots), the theoretical V is higher. This can be explained by the fact that real low thrust spacecrafts do not move on near circular orbits at the outer part of the Martian SOI. Therefore the assumption for the derivation of equation 4 gets violated. Taking also into account that the flight time until body escape is reduced, a higher Mars orbit seems desirable for low thrust spacecrafts. Since arrival and departure dates for minimal energy low thrust interplanetary transfers are approximately fixed, a reduction in Mars escape and in the almost symmetric Mars capture time leads to longer stay times in Mars orbit. During this time in final orbit, the orbiter can provide data relay for the lander. In addition, a later and shorter Mars escape would give the pre deployed rover on the surface more time to rendezvous with the ascent vehicle. On the other hand, it has to be noted that a higher Mars orbit increases the V requirements of the ascent vehicle for launch and orbit injection in order to perform rendezvous with the orbiter. Figure 8: V s of low thrust Mars escape from circular orbits in dependence on the altitude

9 4.3 Mars Capture Figure 9: Low thrust Mars capture trajectory (4 me-stpo to 250 km orbit) To achieve a successful low thrust Mars orbit insertion, the spacecraft, approaching from an interplanetary trajectory, has to adjust its orbit velocity to that of Mars in a way that it can be captured within the Martian SOI. Due to the low thrust of electrical engines the spacecraft can only compensate a limited relative velocity at the border of the Martian SOI. This relative velocity has to be achieved by the interplanetary transfer trajectory. Regarding the interplanetary trajectory the relative velocity should be maximized because this results in less restrictive rendezvous conditions, simplifying the interplanetary trajectory optimization (see also section 4.1). The corresponding capture trajectory calculations are performed backwards in time and aborted when the spacecraft reaches the border of the SOI. Figure 10: Low thrust Mars capture flight times (to 250 km orbit) and initial masses in dependence on the number of engines As it can be observed from fig. 9 and 10, the trajectories and thus flight times are similar to those for Mars escape. This can be explained by the fact that a capture trajectory is a reverse escape trajectory with "inverse" propellant mass flow. From fig. 11 it can also be seen that spacecrafts with enhanced power achieve the highest relative velocities, because they use the available engines more efficiently and thereby reach higher accelerations. Also, considering the shorter escape times of these spacecraft configurations, an enhanced power system seems to be desirable for operations around Mars.

10 Figure 11: Relative velocities at the border of the Martian SOI before low thrust capture 4.4 Interplanetary Outbound Transfer (Earth to Mars) For outbound trajectory calculations it is assumed that the spacecraft is accelerated to v inf = 0 km/s by the launcher or reaches escape velocity by means of its own propulsion system. Examining near Earth operations of enhanced power spacecrafts during interplanetary transfer and Earth escape (see section 4.5) leads to the conclusion that these spacecrafts can provide only lower accelerations in comparison to the vehicles with a standard power system. This results from the fact that the enhanced power system is oversized for operations at Earth distance leading to disadvantageous mass to thrust ratios. For compensation spacecraft, configurations with additional engines are considered. These engines would be used as long as the power system can support them. After that the engines are detached at the border of the Martian SOI equivalent to a staged propulsion system. According to the naming of launcher strap-on propulsion systems, the additional electric engines are called (electric) "booster" (lo-enpobooster, me-enpo-booster). Even though those configurations have more operating engines at the beginning of the mission, they are classified by the "main engine" number at the end of mission. Details can be seen in tab. 3. It has to be considered that the number of booster engines is limited by the original power system, which is optimized for the complete mission. Table 3: Engine numbers of "booster" spacecraft configurations Configuration number Number of main engines Booster engines Number of engines at start of mission Fig. 12 shows a typical low thrust outbound trajectory. Interestingly, it looks like a conventional Hohmann-transfer although the relative velocity at Mars of 548 m/s (compare fig. 9) is remarkably smaller than that of a Hohmann-transfer (v inf 2.7 km/s). For all configurations the achieved relative velocities are displayed in fig. 11. Regarding transfer times, the left part of fig. 13 demonstrates that small engine numbers of about 2-3 lead to excessive flight times. Further on, it can be seen that the newly introduced booster configurations achieve a flight time benefit, but also lead to higher spacecraft masses due to larger dry masses and thereby higher propellant masses. The displayed spacecraft masses (fig. 13 right) can be seen as launch masses, in case a launcher accelerates the spacecraft to escape velocity. Regarding European launchers, the only option would be an Ariane 5, which would be oversized for such a payload. Hence, in the following a low thrust Earth escape from GTO is examined.

11 Figure 12: Low thrust interplanetary outbound trajectory (escape date: May 29, 2020; flight time: 287 d) Figure 13: Low thrust interplanetary outbound flight times and initial masses in dependence on the number of main engines (v inf = 0 km/s at Earth, masses correspond to a 250 km Mars target orbit) 4.5 Earth Escape In principle, Earth escape can be achieved in two different ways: One possibility is to launch the spacecraft directly into an escape trajectory with v inf 0 km/s. This is surely the most time saving option. The second option is a low thrust escape from an Earth orbit. It requires longer flight times, but due to the high specific impulse, the escape can be reached more efficiently and thereby reduce the launcher requirements. Compared to a low thrust escape from a low Earth orbit, the electrically provided V, the total flight time until escape and the stay time within the van Allen belts are smaller for an escape from a Geostationary Transfer Orbit (GTO). There is the additional advantage that launch vehicles all over the world inject payloads frequently to GTO. Hence, a lot of launch opportunities are commercially available. In case the spacecraft mass is below 4-5 t, a shared launch with a GEO communication satellite is possible, resulting in reduced launch costs compared to a dedicated launch to escape velocity. In addition, a shared GTO launch and the following low thrust operations were demonstrated by SMART 1. Hence, this study examines only a low thrust escape from GTO. Please note, plane change maneuvers are probably necessary to escape from a standard GTO. This maneuver depends on the GTO s orientation (thereby the launch site) as well as the launch and escape time. Because no specific GTO is chosen in this study, plane change maneuvers are neglected, too.

12 Figure 14: Low thrust Earth escape flight times from GTO and initial masses for various spacecraft configurations (masses correspond to a 250 km Mars target orbit) Fig. 14 left shows the flight times and the initial masses for an escape from GTO. It can be seen that configurations with booster achieve the shortest flight times due to their high thrust. As mentioned earlier, spacecrafts with enhanced power systems require the longest flight times, because their accelerations are the lowest. The corresponding launch masses are displayed in fig. 14 right and compared to Soyuz-Fregat and Zenit launch performances. 5 SYSTEM ANALYSIS A major system analysis evaluation criterion is the launch mass of the spacecraft. Another important mission characteristic is the overall mission time, which drives costs especially for mission operations and ground support. Furthermore, the spacecraft hardware has to be qualified for a longer mission time. A further evaluation criterion is the time in Mars orbit. During that time the orbiter can support the surface segment by providing communication relay. Another aspect is the departure date at Mars. By shifting the departure to later days, surface time will be increased. This can be used for rendezvous with pre-deployed surface elements, resulting in lower required landing accuracies, or leave more time for scientific research. Figure 15: Mission time with a launch to v inf = 0 km/s (left) and to GTO (right) The mission analysis shows that the mission times are in case of launch to v inf = 0 km/s with about 1, 000 days in a similar range as for chemical missions (see fig. 15 left) [2]. Larger engine numbers, enhanced power systems and electrical booster engines can reduce the mission time, but with a maximal change in mission time of 10% the effect is quite low and would not justify the higher complexity of the considered spacecrafts.

13 In contrast to that, the spacecraft configuration has a significant impact on the total mission time in case of a launch to GTO (see fig. 15 right). Due to the high power system mass of enhanced power configurations with unused excess power at Earth distance, those configurations require more time to provide the Earth escape. This time can be reduced significantly by adding additional booster engines. Then they could be further operated during outbound transfer until they are detached at the border of the Martian SOI. It can be concluded that the implementation of a booster stage reduces the total mission time significantly, but the higher spacecraft complexity has to be considered. Fig. 15 right also shows that the standard power configurations require mission times between both extremes of enhanced power and booster configurations. But due to their shorter stay times in Mars orbit (fig. 16), they seem less desirable and are excluded from further considerations. Figure 16: Stay time in 250 km and km circular Mars orbit As mentioned earlier, electrical sample return orbiters have to spiral from the border of the SOI to the final orbit during Mars capture and escape from it again between more or less fixed Mars arrival and departure dates. Therefore, the stay time in Mars orbit depends predominantly on the required times for Mars capture and escape. To maximize the stay time in Mars orbit, the thrust of the spacecraft at Mars should be as high as possible. Alternatively, the Mars orbit altitude can be increased to extend the time in Mars orbit. The displayed stay times in orbit in fig. 16 left and right represent two possible extremes and should demonstrate the potential. But it has to be noted, that very high orbit altitudes (e.g km) increase the requirements on the Mars ascent vehicle drastically, such that they are only interesting in combination with propellant production on the Martian surface. Therefore, an orbit altitude of 1000 km is preferred. Booster engines cannot increase the stay time in Mars orbit significantly, but such configurations would allow to increase the payload mass to Mars in terms of a piggyback transferred Mars lander without an excessive transfer time increase. The evaluation of the most attractive spacecraft configuration will be mainly based on the launch mass and in particular on the choice of the launch vehicle. As can be seen from fig. 13 right, the launch masses of t into an Earth escape trajectory with v inf = 0 km/s range between the capabilities of Ariane 5 (7 t) and Soyuz Fregat (1.6 t). Since a shared Ariane 5 launch seems not feasible for direct injection into an escape trajectory and a dedicated one is oversized, a launch into GTO is preferred. As mentioned earlier the lo/me-stpo configurations have been excluded and the two remaining configurations (lo/me-enpo and lo/me-enpo-booster) represent extremes in terms of Earth escape (fig. 14 left) as well as total mission time (fig. 15 right) and differ in latter by

14 about 100 days. The reduced mission time and thereby the lower ground segment costs are "paid" by a more complex, tow-stage spacecraft. Especially, a separation of electrical engines has not been performed yet. A further important aspect, the stay time within the van Allen belts and thereby their effect on the spacecraft has to be examined. Due to the relative high thrust of the spacecrafts, the stay times of the remaining configurations range within days. Based on those results, a two stage vehicle seems not justified and the 4 engine, enhanced power spacecraft with an I sp of 4763 s is chosen as baseline. The major characteristics of such a 4 engine sample return orbiter (with 1000 and km orbit altitudes) are displayed in tab. 4. Although the enpo orbiter with 4 engines is preferred, for comparison the lo-enpo orbiter with 3 engines, requiring only a 20 kw solar array, is presented as a minimal configuration. Such a solar array could be based on today s GEO communication satellites. Furthermore, a combined orbiter/lander spacecraft, mentioned earlier, is displayed for information. Details can be found in [6]. It has to be noted that the additional piggyback lander spacecraft of 1600 kg can only deliver an ascent vehicle to the Martian surface and requires support by a pre-deployed rover. Table 4: MSR orbiter characteristic data 4 me-enpo 3 lo-enpo 4 me-enpo with lander No. of engines [-] 4 RIT 22 medium I sp 3 RIT 22 low I sp RIT 22 medium I sp Power at Earth [kw] Mars orbit altitude [km] Total mission time [d] Time in Mars Orbit [d] max. 283 Mass budget [kg] Mars lander N/A N/A N/A 1600 Cruise stage + Adapter N/A N/A N/A 160 Earth return capsule Spacecraft bus Power Propulsion Docking Tanks Structure Propellant Dry mass Wet mass Launcher Soyuz Fregat Soyuz Fregat Soyuz Fregat Proton M Breeze Launcher AR5 2. payload AR5 2. payload AR5 2. payload Atlas Launch orbit GTO GTO GTO v inf = 0 km/s 6 CONCLUSIONS The present paper outlines that a Mars Sample Return mission can benefit from an electrically propelled orbiter. Such an orbiter can be launched by a medium size launcher as the Soyuz- Fregat instead of an Ariane 5 ECA as foreseen for a chemical orbiter. Depending on the configuration, an orbiter with electrical engines can conduct its mission within days. As launch orbit, a geostationary transfer orbit seems an appropriate compromise between a

15 launch to v inf = 0 km/s (most time saving option) and a low thrust escape from a low Earth orbit (option with least requirements on launcher). To maximize the stay time in Mars orbit, the power system should provide sufficient power to operate all electrical engines at Mars. Hence, the power system output at Earth should be about 1.68 times larger than required for the electrical engines. The V for Mars orbit injection as well as the time for Mars capture and escape can be reduced by a Mars orbit altitude increase. Due to the limited thrust of electrical engines, a successful Mars capture requires lower relative velocities than in case of chemical propulsion. Hence, a low thrust propelled spacecraft should enter the Martian SOI with about 550 m/s. The investigations on Mars-Earth return trajectories show that the flight time for interplanetary transfer can be reduced significantly by allowing relative velocities at the border of Earth s SOI of more than 3 km/s. References [1] Preliminary Planning for an International Mars Sample Return Mission. International Mars Architecture for the Return of Samples (imars) Working Group, [2] L. Bessone and D. Vennemann et al. CDF Study Report, Human Missions to Mars, Overall Architecture Assessment. European Space Technology Centre, ESA, [3] B. Dachwald. Low-Thrust Trajectory Optimization and Interplanetary Mission Analysis Using Evolutionary Neurocontrol. Universitaet der Bundeswehr Muenchen, Fakultaet fuer Luft- und Raumfahrt, Doctoral thesis. [4] B. Dachwald. Optimization of very-low-thrust trajectories using evolutionary neurocontrol. Acta Astronautica, [5] B. Dachwald, W. Seboldt, H.W. Loeb, and K-H. Schartner. Main Belt Asteroid Sample Return Mission Using Solar Electric Propulsion. Acta Astronautica 63, pp , [6] U. Derz. Analysis of a Robotic Mars Sample Return Mission. Institute of Material Physics in Space, German Aerospace Center and Chair of Flight Dynamics, RWTH Aachen University, Diploma thesis. [7] M-C. Perkinson et al. Mars Sample Return Study, Executive Summary. EADS Astrium under contract by ESA, [8] E. Messerschmid and S. Fasoulas. Raumfahrtsysteme (in German). Springer, Berlin, Heidelberg, New York, [9] F. Mura. Mars Sample Return ENG-02 Mission Architecture Definition. Thales Alenia Space, [10] NASA. Planetary Mission Entry Vehicles - Quick Reference Guide V3.0, number SP , [11] M.A. Perino and F. Mura et al. MSR Phase A, Executive Summary. Thales Alenia under contract by ESA, [12] W. Seboldt, B. Dachwald, and J. Streppel et al. Lander Mission to Europa with Solar Electric Propulsion. In: proceedings of the 7th International Symposium on Launcher Technologies, Barcelona, Spain, BACK TO SESSION DETAILED CONTENTS BACK TO HIGHER LEVEL CONTENTS