Preliminary Mission Analysis and Orbit Design for Next Mars Exploration

Trans. JSASS Aerospace Tech. Japan Vol. 8, No. ists27, pp. Tk_7-Tk_12, 2010 Topics

By Naoko OGAWA 1), Mutsuko Y. MORIMOTO1), Yuichi TSUDA1,2), Tetsuya YAMADA1,2), Kazuhisa FUJITA1,3), Tomohiro YAMAGUCHI4), Yasuhiro KAWAKATSU1,2), Takashi KUBOTA1,2) and Jun’ichiro KAWAGUCHI1,2)

1)JAXA Space Exploration Center, Japan Aerospace Exploration Agency, Sagamihara, Japan 2)Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency, Sagamihara, Japan 3)Aerospace Research and Development Directorate, Japan Aerospace Exploration Agency, Tokyo, Japan 4)The Graduate University for Advanced Studies, Sagamihara, Japan

(Received July 21st, 2009)

Japan has launched many interplanetary spacecraft for exploration of solar system bodies including Mars. Now we are planning the next Mars mission in the late 2010’s. This paper describes the preliminary mission analysis and orbit design for this plan. The combined exploration by several spacecraft requires complicated and careful consideration, different from those for single-probe missions. Mission plans to realize required conﬁguration by a single launch and simple simulation results are reported.

Key Words: Mars, Mission Analysis, Orbit Design, MELOS

1. Introduction 2.1.1. Meteorological orbiter for martian climate One orbiter of the two, called hereafter as the meteorological In 2008, Japan Aerospace Exploration Agency (JAXA) orbiter, aims understanding of the interaction between the has established a novel working group toward a novel Mars atmosphere and subsurface ice, and atmospheric dynamics. exploration program named MELOS, an acronym for “Mars Global, high-resolution and continuous mapping of water vapor, Exploration with Lander-Orbiter Synergy”1). As its name clouds, dusts and atmospheric temperature will be performed indicates, this is an ambitious mission composed of several with imaging cameras from the apoapsis of its highly elliptic landers and orbiters, schematically illustrated in Fig. 1. orbit. At lower altitude, a sub-millimeter sounder will also be Combined and networked exploration by multiple spacecraft is used for three-dimensional mapping of water vapor and winds. one of notable features of the MELOS mission, compared to Knowledge of Martian climate with the meteorological several missions to be launched in 2010’s 2–5). A working group orbiter will allow us to establish comparative meteorology for the MELOS mission has been established in 2008, and more of terrestrial planets such as Earth, Venus and Mars by than 100 researchers have joined discussing the details of the incorporating Earth climate data and those from Akatsuki, mission toward the launch in late 2010’s. JAXA’s Venus Climate Orbiter launched in May 20106). Both This paper describes the preliminary mission analysis and scientific data and engineering designs of PLANET-C will be orbit design for the MELOS mission. The combined inherited to the meteorological orbiter. exploration by several spacecraft requires complicated and 2.1.2. Atmospheric escape orbiter for escaping atmo- careful consideration, different from those for single-probe sphere missions. Mission plans to realize required configuration by a The other orbiter, here we call it the atmospheric escape single launch and simple simulation results are reported. orbiter, performs in-situ observations of escaping atmosphere in a low-Mars orbit, which was one of science objectives in 2. Mission Overview Nozomi – Japan’s past Mars orbiter mission7). It is now almost certain that Mars once had duration of warm In this section, overview and scientific basis of the MELOS and wet climate. The aim of the observation of atmospheric mission are introduced. escape is to obtain a clue of how and why the atmosphere 2.1. Scientific objectives and climate of Mars have evolved with time. Our target This mission is expected to consist of two orbiters and several is to elucidate non-thermal escape processes, in particular, landers, as implied by its name, MELOS, an acronym for solar wind-induced escape processes, which are pointed out to “Mars Exploration with Lander-Orbiter Synergy”. Two orbiters involve substantial uncertainties by previous measurements and and landers will perform cooperative and combined exploration theoretical studies. of Martian atmosphere, climate, interior structure and surface Elucidation of atmospheric escape will be strongly supported environment for understanding and elucidation of the evolution also by the meteorological orbiter in two ways. First, simul- of Mars environment. The MELOS mission challenges the taneous observation will be performed by the meteorological following three science objectives. orbiter to grasp global structures of escaping ions, and by the atmospheric escape orbiter to investigate the escape processes through in-situ measurements. Second, it will monitor solar

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winds, which is crucial to understand present escape processes or fluxes as well as their dependencies on the external condi- Arrival 1.0e+08 Dec 2018 tions. 2.1.3. Landers for internal structure and surface environment 0.0e+00 MELOS surface landers will scope several scientific topics such as mass spectrometry, seismology, geochemistry, thermal Y [km] activity, crater chronology and atmospheric electricity. They Launch -1.0e+08 will carry several science packages for measurement of these May 2018 properties. It is noteworthy that we are planning networked exploration by several landers distributed in the wide area on -2.0e+08 Mars Mars. They can benefit from network science with ESA’s Mars Earth S/C NEXT mission. -1.0e+08 0.0e+00 1.0e+08 2.0e+08 3.0e+08 Accumulation for surface science and landing technologies X [km] through recent ambitious missions such as Hayabusa (JAXA’s Fig. 2. The transfer orbit from Earth to Mars in the inertial ecliptic asteroid sample return mission)8) and its follow-ons, and coordinate system (case 1, launched in 2018). SELENE-2 (JAXA’s lunar lander)9) will be of great use for Mars landers. Experience of reentry into the atmosphere and knowledge of aerodynamics cultivated through Hayabusa’s 2.0e+08 reentry will be also inherited.

3. Orbit Sequence 1.0e+08 3.1. Launch opportunities We are planning the launch of the all spacecraft in the late Launch 2010’s by a single H-IIA rocket, aiming data acquisition under May 2018 conditions of high solar activity expected to be maximized 0.0e+00 Y [km] Sun around 2022. There will be several launch windows toward Mars around 202210). In one case, for example, the spacecraft will depart Earth in May 2018 and arrive at Mars in December 2018, after a half revolution around the sun. The total delta-V -1.0e+08 for the departure and the arrival is 5.75 km/s. In other case, the Arrival departure from Earth will be in October 2017, and the arrival at Dec 2018 Mars Mars will be February in 2020. Though it requires one and a S/C -2.0e+08 half revolutions around the sun, the total delta-V is 5.59 km/s, 0.0e+00 1.0e+08 2.0e+08 3.0e+08 less than that for the former case. Figs. 2-5 show transfer orbits X [km] from Earth to Mars in the inertial ecliptic coordinate system Fig. 3. The transfer orbit from Earth to Mars in the Sun-Earth ﬁxed (J2000) and Sun-Earth ﬁxed ecliptic coordinate system. ecliptic coordinate system (case 1, launched in 2018). 3.2. Orbital design around mars After the cruising phase around the sun, the spacecraft is 3.0e+08 injected to the orbit around Mars. Two orbiters are then

2.0e+08

1.0e+08 Launch Mars Exploration Oct 2017 with Lander-Orbiter 0.0e+00 Synergy (MELOS) Y [km]

-1.0e+08

-2.0e+08 Arrival Feb 2020 Mars Earth S/C -3.0e+08 -3.0e+08-2.0e+08-1.0e+08 0.0e+00 1.0e+08 2.0e+08 3.0e+08 X [km]

Fig. 4. The transfer orbit from Earth to Mars in the inertial ecliptic Fig. 1. Concept of MELOS. coordinate system (case 2, launched in 2017).

Tk_8 N. OGAWA et al.: Preliminary Mission Analysis and Orbit Design for Next Mars Exploration

3.0e+08 Atmospheric Escape Orbiter

2.0e+08 Arrival Feb 2020

1.0e+08 Landers Meteorological Launch Orbiter Oct 2017 0.0e+00 Y [km] Sun

-1.0e+08 Fig. 6. Orthogonal constellation required for two orbiters.

-2.0e+08 Mars Table 1. Examples of orbital elements of the two orbiters. S/C -3.0e+08 -3.0e+08-2.0e+08-1.0e+08 0.0e+00 1.0e+08 2.0e+08 3.0e+08 Meteorological Atmospheric X [km] Orbiter Escape Orbiter Fig. 5. The transfer orbit from Earth to Mars in the Sun-Earth fixed Apo. Alt. 6.9 Rm 7,000 km ecliptic coordinate system (case 2, launched in 2017). Peri. Alt. 300 km 300 km Incl. 8.67 deg 102.09 deg Period 16 hrs (3 revs in 2 5 hrs separated, and several maneuvers are performed to establish the days) constellation appropriate for the combined observation of Mars. RAAN Rate 0.4142 deg/day 3.2.1. Mission requirements for orbits around mars As mentioned above, simultaneous observation of the Martian atmosphere by two orbiters is one of the most important 3.2.2. Mars orbit insertion and initial orbit phase goals in the MELOS mission. The atmospheric escape orbiter After about a 7-month voyage at shortest along the Mars has to be in a low-altitude polar orbit for in-situ and local transition orbit, the spacecraft will be inserted into a Mars observation of the Martian atmosphere. The meteorological orbit by an orbit maneuver engine (OME) mounted on the orbiter needs to capture global images from the distance of atmospheric escape orbiter, with the initial periapsis altitude of several radii of Mars. Its orbital period is preferable to be in about 300 km. We adopted insertion of the two orbiters and synchronization with the daily rotation of Mars, if possible, in landers into a single initial orbit as a whole, instead of one-by- order to observe the transition of the Martian climate under high one insertion of already separated spacecraft. It has benefit such resolution and to capture whole global images in one or two that only a single OME is needed, and that concurrent insertion days. is avoided. How to establish different dedicated observation To comply with these requirements, several plans are now orbits for the two orbiters is, however, a difficult issue. under discussion. In one candidate plan, tentative values for the Several constellation plans and maneuver sequences are now apoapsis altitude and periapsis altitude of the meteorological under consideration. The following sequence is one example. orbiter are 6.9Rm and 300 km, respectively, where Rm is the First, the spacecraft is inserted into a highly elliptical orbit with radius of Mars. Those for the atmospheric escape orbiter are nearly 50 Rm apoapsis for example and a low inclination, which 7,000 km and 300 km, respectively. Note that these values allows us to change inclination of the orbiter with a relatively are just tentative, and desirable values are currently under small maneuver on the apoapsis, and to spend sufficient time for discussion in the science community. precise orbit determination such as DDOR (Delta Differential Moreover, the most essential requirement is the orthogonal One-way Range) prior to deorbit of landers. constellation of two orbiters, as shown in Figure 6. This How to create and maintain orthogonal constellation is constellation must be kept at least during the mission period another challenge. One solution is to utilize perturbation of two Martian years. Because orbital elements are largely of Mars gravity non-uniformity, especially J2 terms effecting 11) perturbed by various external forces such as the J2 term in Mars orbital planes to rotate . This will save chemical thrust, though gravity non-uniformity, orbital elements of the orbiters must be it requires considerably long time for 90-degree rotation of the designed carefully so as to keep orthogonality11). For example, orbital plane. in the case of the launch in 2018, the minimum inclination As for descent of the orbiters from the initial orbit to the of the initial insertion orbit is 8.67 degrees. If we adopt observation orbit, aerobraking will be very effective12). During this as the inclination of the meteorological orbiter, we can the braking, the spacecraft will be dipped into the atmosphere keep orthogonality by setting the inclination of the atmospheric with about 110-120 km altitude. It will also be of great help for escape orbiter 102.09 degrees. Table 1 shows tentative orbital observation of atmospheric escape, because such lower altitude parameters of the two orbiters in the observation phase. The is scientifically significant. period of the meteorological orbiter is set to be 16 hours so that Table 2 shows an example of possible maneuver sequences it synchronizes with the daily rotation of Mars every two sols. for the orbiters. Delta-Vs in (F) and (H) correspond to “walk- in” and “walk-out” maneuvers between the upper atmosphere

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Table 2. A possible maneuver sequence for the orbiters. Events Total Mass Meteorological Atmospheric Landers Notes Orbiter Mass Escape Mass Orbiter Mass (A) Before MOI 2,956 kg 345 kg 1,713 kg 448 kg (B) MOI 2,226 kg 983 kg 50Rm 300 km, i=8.67 deg × (888.9 m/s) (C) Orbiter Separation (D) Landers Separation Delta-V: TBD (E) Inclination Maneuver 2,180 kg 937 kg 8.67 deg 102 deg → (150.9 m/s) (F) Meteorological Orbiter 2,178 kg 343 kg 50Rm 8Rm → Aerobraking (16.8 m/s) (G) Orbital Plane Rotation >204 days (H) Atmospheric Escape 2,169 kg 928 kg 50Rm (TBD) 7,000 km → Orbiter Aerobraking (31.2 m/s) with 110-km altitude and upper nominal orbits for start and stop 120 Entry of aerobraking. 100 4. Entry, Descent and Landing of Landers 80 The landers are initially mounted on the meteorological orbiter just after MOI (Mars Orbit Insertion), and separated 60 Peak Heat from the orbiter. Compared to direct entries adopted in most Flux Altitude [km] Mars lander missions by other countries, the entry from the 40 Mars orbit has an advantage of more flexible selection of the 20 landing point and higher landing accuracy. It is also intended to Parachute Powered avoid concurrent critical operation for entry of the landers and Deploy Decsent 0 MOI of the orbiters. 0 2 4 6 8 10 12 Deorbit of the landers is performed near the apoapsis. Time [min] One candidate scenario under discussion includes the landers Fig. 8. Time sequence of the lander altitude during entry, descent and targeting maneuver (LTM) along with the meteorological landing. orbiter toward the Mars atmosphere interface point, followed by separation of the landers, the orbiter deflection maneuver (ODM) and a cleanup, in a similar manner to the separation based guidance will be of a great help for accurate landing on sequence of the Huygens probe from the Cassini spacecraft the aimed point14). 13). It leads to reduction of AOCS (Attitude and Orbit Control We are also planning some mobilities for landers or a part of Subsystem) components in the landers’ capsule, though it needs them after landing. Several ambitious plans including rovers, a solar array panel and other subsystems for survival over airplanes and airships are being discussed. several days. A case of deorbit from the nominal observation orbit of the meteorological orbiter is also discussed, where only 5. Tentative Mass Budget the landers will deorbit and reach the entry point after several hours. Table 3 shows a tentative mass budget for orbiters and After coasting, the spin-stabilized lander reaches the entry landers. As for the vehicle launch capability, 2,500 kg, 3,000 point. One example of the entry, descent and landing sequence kg and 3,500 kg were supposed provisionally, and one, two and three lander(s) corresponding to the launch capability were is illustrated in Fig. 7. In this case, deorbit from the 8Rm apoapsis of the meteorological orbiter is tentatively assumed. assumed respectively. Details are to be determined according to The lander enters the atmosphere at the velocity of 4.6 km/s the mission sequence and spacecraft configuration. and the flight path angle of 8.34 degrees. After experiences of the peak heat flux and the peak dynamic pressure, a parachute 6. Summary is deployed and the heat shield is jettisoned. Powered descent with thrusters is triggered by the altitude threshold and finally This paper described the preliminary mission analysis and soft landing is achieved. Figs. 8 and 9 also show the altitude orbit design for Japanese next Mars mission named MELOS. and velocity of the lander, respectively. During this sequence, The scientific objectives, outlines of the mission and several several guidance techniques such as guided entry or image- topics were briefly described. Note that these descriptions

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E−7.8h: Deorbit at Apo ΔV=20m/s E−0m0s: Entry h=120km, V=4.6km/s, γ =−8.34deg

E+2m32s: Peak Heat Flux (0.08MW/m2) h=48km, V=4.1km/s E+3m16s: Peak Dynamic Pressure (1.2kPa) h=39km, V=2.8km/s

E+6m00s: Parachute Deploy & Shield Jettison, Peak Accel (7.88G) h=10km, V=200m/s

E+11m00s: Powered Descent h=214.5m, V=24.2m/s

E+11m02s: Parachute Jettison h=166.1m, V=24.1m/s

E+11m47s: Touchdown V=0.95m/s

Fig. 7. Entry, descent and landing sequence of the landers.

Table 3. A preliminary mass budget of MELOS mission (Unit: kg). Meteorological Orbiter Atmospheric Escape Orbiter 1 Lander 2 Landers 3 Landers AOCS 41.4 11.3 32.8 32.8 32.8 Power 80.2 63.8 59.2 59.2 59.2 Comm 22.0 22.0 18.8 18.8 18.8 DHU 6.9 6.9 6.9 6.9 6.9 EDL SS 0.0 0.0 80.0 80.0 80.0 Launch SS 0.0 2.2 0.0 0.0 0.0 Propulsion 37.7 108.4 40.9 40.9 40.9 Wire Harness 12.0 59.6 15.2 15.6 16.4 Structure 36.0 178.8 55.6 56.8 59.2 Thermal 6.0 29.8 7.6 7.8 8.2 SS Total 242.2 482.8 316.9 318.7 322.3 Payload 27.8 37.2 13.1 21.3 37.7 Dry 270.0 520.0 330.0 340.0 360.0 Fuel 30.0 970.0 50.0 50.0 50.0 Wet 300.0 1490.0 380.0 390.0 410.0 Margin (15%) 45.0 223.5 57.0 58.5 61.5 Total 345.0 1713.5 437.0 448.5 471.5 AOCS: Attitude and Orbit Control System. DHU: Data Handling Unit. EDL: Entry, Descent and Landing. SS: Subsystem.

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5000 SELENE Follow-On, in Proceedings of Joint Annual Meeting of LEAG-ICEUM-SRR, 2008. 4500 Entry 500 10) Ishii, N., Ogawa, N., Morimoto, M. Y. and Sato, T.: Launch Oppor- 4000 tunities for Mars Exploring Mission in 2015-2020, in Proceedings of 3500 The 27th International Symposium on Space Technology and Science Parachute Deploy (ISTS 2009), 2009. 3000 Peak Heat 11) Ogawa, N., Morimoto, M. Y., Kawakatsu, Y. and Kawaguchi, J.: 2500 Parachute Flux Constellation of Two Orbiters around Mars, in Proceedings of The Jettison 2000 27th International Symposium on Space Technology and Science (ISTS 0

Velocity [m/s] 6 8 10 2009), 2009. 1500 12) Fujita, K., Kubota, T., Ogawa, N., Morimoto, M. Y., Suzuki, 1000 T., Takayanagi, H., Yamada, T. and Kawaguchi, J.: Assessment of 500 Aeroassist Orbital Maneuver Technologies for Next Mars Exploration, in Proceedings of The 27th International Symposium on Space 0 Technology and Science (ISTS 2009), 2009. 0 2 4 6 8 10 12 Time [min] 13) Kazeminejad, B., Atkinson, D. H., Perez-Ay´ ucar,´ M., Lebreton, J.-P. and Sollazzo, C.: Huygens’ entry and descent through Titan’s Fig. 9. Time sequence of the lander velocity during entry, descent and atmosphere–Methodology and results of the trajectory reconstruction, landing. Planetary and Space Science, 55 (2007), pp.1845-1876. 14) Kubota, T., Okada, T., Miyamoto, H., Ogawa, N., Morimoto, M. Y., Fujita, K., Yamada, T., Mizuno, T., Kawakatsu, Y., Satoh, T. and Kawaguchi, J.: Preliminary Study on Lander System and are tentative as of 2009 and to be modified, confirmed and Scientific Investigation for Next Mars Exploration, in Proceedings of determined through ongoing discussion among researchers. The 27th International Symposium on Space Technology and Science (ISTS 2009), 2009. Acknowledgments

The authors are deeply grateful to Prof. N. Ishii in Japan Aerospace Exploration Agency and Prof. H. Yamakawa in Kyoto University, for enlightening discussions on orbital dynamics and mission design.

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