Launch Opportunities and Preliminary Orbit Design for Next Mars Exploration Program

Home , Areostationary orbit, Mars Exploration Program, Melos

Trans. JSASS Aerospace Tech. Japan Vol. 10, No. ists28, pp. Tk_19-Tk_25, 2012 Topics

By Naoko Ogawa1), Michihiro Matsumoto2), Nobuaki Ishii2), Yuichi Tsuda1,2), Yasuhiro Kawakatsu1,2), Jun’ichiro Kawaguchi1,2), Takeshi Imamura2), Ayako Matsuoka2), Takashi Kubota1,2) and Takehiko Satoh1,2)

1)JAXA Space Exploration Center, Japan Aerospace Exploration Agency, Sagamihara, Japan 2)Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency, Sagamihara, Japan

(Received June 27th, 2011)

Since 2008, a new plan for next Mars exploration program has been proposed and discussed by scientists and engineers in Japan. This exploration program is named MELOS, or Mars Exploration with Lander and Orbiter Synergy, a long-awaited program in the planetary science community in Japan after unsuccessful end of Nozomi Mars orbiter and ongoing challenge of Akatsuki Venus orbiter. The goal of the whole program is to understand Mars as a system, by elucidation of Martian climate, atmospheric escape, internal structure, surface environment and interaction by them. A series of missions has been planned, and several spacecraft including orbiters and landers are under discussion to be launched in early 2020’s. In this paper, we investigate launch opportunities during early 2020’s and estimate the payload mass in each case. Feasible interplanetary transfer trajectories from Earth to Mars are proposed. Preliminary design of insertion sequence into the Mars orbit and some orbit candidates derived from mission requirements are also shown together with numerical simulation results.

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

1. Introduction improved by Akatsuki at Venus2) and by MELOS1 at Mars. In this proposal, a meteorology orbiter will be inserted into the Since 2008, scientists and engineers in Japan have discussed Mars orbit. Another is to study escaping atmosphere that is the next Mars exploration program named MELOS, an acronym thought to be a key process for today’s tenuous atmosphere for “Mars Exploration with Lander-Orbiter Synergy”1). As its of Mars. Researchers propose “2-orbiter” configuration for name indicates, this is a programmatic series of several ambi- escaping atmosphere so that in-situ measurements and global tious missions composed of landers and orbiters. Combined views will be acquired simultaneously. We describe possible and networked exploration by multiple spacecraft is one of no- orbit sequences for the meteorology orbiter configuration and table features of the MELOS series. A working group for the 2-orbiter configuration in Sections 4. and 5., respectively. MELOS has been established in 2008, and more than 100 researchers have joined discussing the details of the first mission 3. Launch Opportunities toward the launch in early 2020’s. This paper describes the preliminary mission analysis and Table 1 shows launch opportunities to Mars from 2019 and orbit design for the MELOS mission series. Possible mission 2024. In one case, for example, the spacecraft will depart Earth plans to realize required configuration by a single launch and in July 2020 and arrive to Mars in February 2021, after about simple simulation results are reported. a half rotation around Sun. Vinf means the hyperbolic excess speed from the outgoing hyperbola from Earth or the incoming 2. Overview hyperbola to Mars, which corresponds to the velocity difference between the planet and transfer orbit. It is noteworthy that In this section, overview and scientific basis of the MELOS launch opportunities around early 2020’s require substantially series are introduced. high velocity to escape from Earth and to approach Mars. It The first mission called MELOS1 is planned to be launched means that the total mass which can be delivered to the Mars around early 2020’s, which will be an “orbiter primary” mission orbit is not so large. Some transfer orbits require more than one with one or more orbiters and a small lander as a precursor rotations around Sun. The table also includes the approximate to demonstrate entry, descent and landing (EDL). A larger estimation of the maximum dry mass that can be delivered by MELOS2 mission with a well-equipped lander will follow and an H-IIA 202 or 204 vehicle into the Mars transfer orbit (MTO) enhance our understanding about Mars. or into the Mars orbit, where the final orbit is assumed to be There are two proposals for the orbiter part of MELOS1. One 300 km 10 R and R is the Mars radius. Note that the dry × M M is to complement the comparative meteorology of terrestrial mass includes lander’s propellant for descent. Among the H- planets. Our knowledge of Earth meteorology will be greatly IIA series, 202 is the basic type with two solid rocket boosters

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Table 1. Launch opportunities to Mars during early 2020’s.

Launch Arrival Departure Vinf Arrival Vinf Dry mass in MTO [ton] Dry mass in Mars orbit [ton] [km/s] [km/s] H-IIA 202 H-IIA 204 H-IIA 202 H-IIA 204 Nov. 2019 Feb. 2022 3.00 2.70 2.0 3.3 1.4 2.4 Jul. 2020 Feb. 2021 3.70 2.62 1.7 2.9 1.2 2.1 Nov. 2021 Jan. 2024 3.01 2.98 2.0 3.3 1.3 2.2 Sep. 2022 Aug. 2023 3.83 2.64 1.6 2.8 1.1 2.0 Dec. 2023 Jan. 2026 3.12 3.31 2.0 3.1 1.3 2.0 Oct. 2024 Aug. 2025 3.36 2.45 1.9 3.1 1.4 2.3

(SRB), while 204 has an enhanced launch capability by adding and declination of the outgoing asymptote are 12.5 degrees and two more SRBs. 26.2 degrees, respectively. A small burn of TCM-1 (Trajectory Correction Maneuver 1) is performed 15 days after the launch. 4. Preliminary Orbit Design for Meteorology Orbiter After the six-month cruise, TCM-2 is planned 15 days before arriving Mars. The spacecraft approaches Mars with a B- Preliminary orbital elements for the meteorology orbiter are parameter of 7,751 km and a phase angle of 63.32 degrees. shown in Table 2. The orbit is elliptic in order to observe the Finally on 16th February 2021, the OME (Orbit Maneuver whole Mars globe from its apoapsis. In this paper we set its Engine) of the spacecraft burns with an 898-m/s delta-V at a inclination to be 63.4 degrees so that the argument of periapsis 300-km altitude of Mars to inject itself into the orbit as shown be kept constant, but this value is tentative. in Fig. 3.

Table 2. Preliminary orbital elements of the meteorology orbiter. 16 Feb.. 2021 Elements Values Periapsis altitude [km] 300 Apoapsis altitude [km] 30,564 (9 RM) Inclination [deg] 63.4 (TBD)

4.1. Transfer to Mars In this paper, we assume the launch year to be 2020. A possible mission sequence from the launch to Mars orbit insertion (MOI) can be designed as follows. Simulation was performed by using an aerospace mission analysis software STK 9 and its trajectory design module Astrogator (Analytical Graphics, Inc.). Note that the sequence is not optimized. Table 3 shows the overall sequence. 25 Jul. 2020

Table 3. A possible mission sequence for the meteorology orbiter. Dates Events Notes

25 Jul. 2020 Launch from Tane- 40.8-min coast Fig. 1. Mars transfer orbit for the meteorology orbiter. 21:41:40 UTC gashima 25 Jul. 2020 Insertion into Mars 3.81 km/s 22:29:16 UTC Transfer Orbit 4.2. Mars orbit Figure 4 and 5 show transition of orbital elements for the 9 Aug. 2020 TCM-1 (optional) 62.4 m/s meteorology orbiter after MOI, computed by STK/Astrogator 22:29:16 UTC considering Mars gravity field coefficients up to 80th degrees 1 Feb. 2021 TCM-2 (optional) 1.1 m/s in Goddard Mars Model 2B3), the spherical solar radiation 06:26:11 UTC pressure model4), and a simple exponential Mars air drag model, 16 Feb. 2021 MOI at 300 km peri- 898 m/s where we set the solar radiation pressure coefficient C to be 04:20:46 UTC apsis r 1, the atmospheric drag coefficient Cd to be 2.2, the spacecraft mass to be 500 kg, its area to be 20 m2, the reference air density A heliocentric view of the Mars transfer orbit for the to be 2 107 kg/km3 and the scale altitude to be 11.1 km, as meteorology orbiter is shown in Fig. 1. On 25th July 2020, the × preliminary values. It is indicated that the orbit is stable and spacecraft will be launched from Tanegashima Space Center, deviation of orbital elements is sufficiently small for at least one Japan. After 41-minute coasting at the 300-km altitude, the martian year. craft is injected into the outgoing transfer orbit toward Mars by a 3.81-km/s delta-V, as shown in Fig. 2. Right ascension

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63.90

63.85 63.45 Escaping Orbit 63.80

63.75

63.70 63.40

Launch 63.65 Feb 15 Mar 1 Mar 15 Apr 1

63.60

63.55

Coasting 63.50

63.45

63.40 Apr Jul Oct Jan 2022 Apr Jul Oct Jan 2023 2021 (UTCG)

Inclination (deg) Kick Burn Fig. 5. Transition of the inclination of the meteorology orbiter.

4.3. Option: insertion into areostationary orbit (ASO) There are many suitable orbits for effective observation of the Martian climate. One potential candidate is the aerostationary Fig. 2. Launch, coasting and injection of the meteorology orbiter into orbit (ASO). Like a geostationary orbit, ASO keeps its nadir the Mars transfer orbit. on the fixed surface point about 17,000-km below within the Mars equatorial plane by synchronizing its rotation with Mars revolution. While ASO is useful for observation of climates on fixed points, it is relatively difficult to transfer to ASO from an interplanetary hyperbola, because its radius is large and the spacecraft cannot benefit from Mars gravity sufficiently for deceleration. In this paper we would like to discuss insertion Inserted Orbit into ASO as an optional study. In transfers from a hyperbola into a circular orbit, a 2- or 3-impulse transfer is sometimes more efficient than direct 5, 6) insertion by a single impulse . Assuming the arrival Vinf to be 2.62 km/s for example, a 2-impulse transfer requires less delta- V when Vinf is larger than the escape velocity Vesc for the ASO, which is 2,046 m/s, as shown in both Fig. 6 and Ref. 5. Fig. 7 and Ref. 5 also indicate that a 3-impulse transfer requires much less delta-V when the intermediate apoapsis is larger than the Incoming Orbit Insertion ASO radius. Burn 1000

0.88

0.98 0.86

0.84

0.92

0.9

0.94

900 0.96 Fig. 3. Mars orbit insertion of the meteorology orbiter.

800

30800 30800 365 308 360 700 30750 0.86 0.88 30700 0.98 306 355 0.84

0.92

0.9

1 350 0.94

30700 304 0.96 30600 600 345 302 30650 340

Feb 15 Mar 1 Mar 15 Apr 1 335 500 Periapsis altitude [km] 30600 0.82 330

30550 325 400 0.86 0.88

0.98 0.84 320

0.92

0.9

1 30500 0.94

315 0.96 300 310 30450 1 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9 305 Vinf / Vesc

30400 300 Apr Jul Oct Jan 2022 Apr Jul Oct Jan 2023 2021 (UTCG) Fig. 6. Ratios of total delta-Vs for 2-impulse transfers to 1-impulse

Apogee Altitude (km) Perigee Altitude (km) transfers. The ratio less than 1 is preferable.

Fig. 4. Transition of the apoapsis and periapsis altitude of the meteorology orbiter. Let us estimate delta-Vs in a particular case. In a 1-impulse transfer, 1,877 m/s is required. If we choose a 2-impulse

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1000 1.3 1.4 5.1. Transfer to Mars

0.9 1.1

1.2 In this paper, we assume the launch year to be 2022. A 1.5 900 possible mission sequence from the launch to the MOI can be designed as follows. Table 7 shows the overall sequence. 800 A heliocentric view of the Mars transfer orbit for 2-orbiter

700 constellation is shown in Fig. 9. On 2nd September 2022, the 1.3 1.4

0.9 1.1

1.2 spacecraft will be launched from Tanegashima Space Center, 1.5 600 Japan. After 28-minute coasting at a 300-km altitude, the craft is injected in the outgoing transfer orbit toward Mars by a 500 3.85-km/s delta-V, as shown in Fig. 10. Right ascension and Periapsis Altitude [km] declination of the outgoing asymptote are 80.1 degrees and 4.45 1.3 400 1.4

0.9

1.1 degrees, respectively. A small burn of TCM-1 is performed 1.2 1.5 15 days after the launch. After the ten-month cruise, TCM- 300 1 2 3 4 5 6 7 8 2 is planned 30 days before arriving Mars. The spacecraft Apoapsis / radius_ASO approaches Mars with a B-parameter of 7,663 km and a phase Fig. 7. Ratios of total delta-Vs for 3-impulse transfers to 2-impulse angle of 58.57 degrees. Finally on 13th August 2023, the OME transfers. The ratio less than 1 is preferable. of the spacecraft burns with a 1.01-km/s delta-V at a 300-km altitude of Mars to inject itself into the initial orbit with the 30- RM apoapsis as shown in Fig. 11. transfer with an intermediate orbit of 300 km 17,000 km, × 5.2. Orbit transformation for orthogonal constellation then the total delta-V is 1,697 m/s. And in the case of 3-impulse After insertion into the initial orbit, several maneuvers are transfer including intermediate orbits of 300 km 50 R and × M required to establish orthogonal constellation. In order to save 17,000 km 50 R , the delta-V is only 1,334 km/s. As Table × M fuel consumption, natural perturbation elements such as the J2 4 implies, a 3-impulse transfer is the best strategy to save fuel term of Mars gravity ﬁeld coefﬁcients or the air drag force are consumption. utilized for orbit transformation. The whole sequence is illustrated in Fig. 12. First, Table 4. Total delta-Vs in multi-impulse transfer orbits to ASO. two spacecraft are separated, and the remote-sensing orbiter Transfer Total delta-V descents to the orbit with 8 RM apoapsis by aerobraking. Next, 1 impulse 1,877 m/s the in-situ orbiter changes its inclination to 84.4 degrees at 2 impulses 1,697 m/s the apoapsis and enters into an intermediate orbit. Then, 3 impulses 1,334 m/s difference of J2 perturbation between the two orbits causes relative precession: the precession rate of the right ascension of ascending node is 0.19 deg/day in the remote-sensing − orbiter, while that in the in-situ orbiter is 0.005 deg/day, 5. Preliminary Orbit Design for 2-Orbiter Constellation − where the Mars J2 is 1.964 10 3 here1. The two orbital × − As previously mentioned in Section 2., another proposal planes will become almost orthogonal after 483 days. Finally with 2-orbiter constellation aiming elucidation of atmospheric the in-situ orbiter will descent to the required orbit, which escape is under discussion. It is composed of a remote-sensing will complete the orbit transformation and constellation with orbiter and an in-situ orbiter. precession synchronization of two orbital planes. Maneuver The in-situ orbiter has a low altitude orbit, while the remote- sequence is also shown in Table 8. sensing orbiter has a highly elliptical orbit. As illustrated in Fig. 8, it is assumed that the remote-sensing orbiter looks down upon the in-situ orbiter’s orbital plane from the apoapsis in order to perform simultaneous observation of atmosphere7); i.e., the In-situ Orbiter two orbits should be preferably orthogonal. More exactly, the in-situ orbiter’s normal vector and the remote-sensing orbiter’s eccentricity vector should be parallel, and it should be kept during the mission phase. How to establish and keep such constellation have been discussed in previous papers8). Preliminary orbital elements for the remote-sensing and in- situ orbiters in 2-orbiter constellation are shown in Tables 5 and 6, respectively. Inclination values for two orbiters are Remote-Sensing Orbiter chosen carefully so that the two orbits are orthogonal and its Fig. 8. Constellation of two orbiters around Mars. constellation is kept as long as possible8). These inclination values allow two orbital planes to share the common precession 1 9) rate of -0.19 deg/day, helping the maintenance of constellation. In this section, we used the J2 value on Rika-Nenpyo , for simplicity. The 3) 3 J2 value derived from GMM-2B , or normalized C20 is 1.956 10− , which is consistent with the value we used here within the scope of this× paper.

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Table 5. Preliminary orbital elements of the remote-sensing orbiter in 2-orbiter constellation. Elements Values Escaping Orbit Periapsis altitude [km] 300 Apoapsis altitude [km] 23,772 (7 RM)

Inclination [deg] 63.4 Launch

Table 6. Preliminary orbital elements of the in-situ orbiter in 2-orbiter constellation. Elements Values Coasting Periapsis altitude [km] 300 Kick Burn Apoapsis altitude [km] 7,000 Inclination [deg] 84.4

Table 7. A possible mission sequence for 2-orbiter constellation. Date Event Note Fig. 10. Launch, coasting and injection of 2-orbiter constellation into 2 Sep. 2022 Launch from Tane- 27.6-min the Mars transfer orbit. 03:09:01 UTC gashima coast 2 Sep. 2022 Insertion into Mars 3.85 km/s 03:43:28 UTC transfer orbit 17 Sep. 2022 TCM-1 (optional) 25.9 m/s 03:43:28 UTC 14 Jul. 2023 TCM-2 (optional) 49.3 m/s 03:43:28 UTC 13 Aug. 2023 MOI into initial orbit 1.01 km/s 03:17:38 UTC at 300-km periapsis Inserted Orbit 25 Aug. 2023 Separation & Aerobrake or 15:42:19 UTC Remote-Sensing 213 m/s orbiter descent 27 Aug. 2023 Inclination maneuver 70 m/s 17:55:59 UTC for In-situ orbiter

26 Dec. 2024 In-situ orbiter de- Aerobrake or Insertion 22:59:25 UTC scent 594 m/s Burn 1 Jan. 2025 Completion of con- Incoming Orbit stellation

Fig. 11. Mars orbit insertion of 2-orbiter constellation.

Table 8. Changes in orbital parameters in each orbiter.

13 Aug. 2023 Events Remote-Sensing In-Situ Orbiter Orbiter

MOI 300 km 30 RM, i = 63.4 deg 2 Sep. 2022 × Separation Descent 30 R 7 R M → M Inclination i = 63.4 deg 84.4 → Maneuver deg Precession 483 days ∼ Descent 30 R 7,000 km M → Fig. 9. Mars transfer orbit for 2-orbiter constellation.

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6. Discussion on Simultaneous or Sequential Insertion 21 Aug. 2023 Initial Orbit In 2-orbiter constellation, two spacecraft are required to enter into two orbits with different altitude and inclination. It requires so many maneuvers and long waiting time before completion of constellation, as we saw above. The simplest way to realize such constellation is to launch two orbiters independently from Earth. Such scenario is not realistic, however, because launch opportunities, budgets, facilities and resources are limited in the planetary exploration program in Japan. Thus, we assumed that the two orbiters will be launched simultaneously. 3 Sep. 2023 After simultaneous launch, we can propose two scenarios for Separation & Descent of finally acquiring two different orbits around Mars: Insertion Remote-Sensing Orbiter before Separation (IBS) or Separation before Insertion (SBI). Inclination Change of IBS is a strategy that we adopted in the former section, where In-Situ Orbiter the two orbiters are inserted into the initial Mars orbit, and then separated and several orbit transfer maneuvers are performed to achieve orthogonal constellation. In SBI, on the other hand, the orbiters are separated before arriving Mars, and inserted independently and sequentially into the different orbits. An advantage of SBI is that maneuvers and waiting time after MOI are not necessary, which leads to increase of the mass 31 Aug. 2024 for more payloads and immediate start of the mission phase. Precession of However, SBI has two potential difficulties; each spacecraft Remote-Sensing Orbiter needs its own OME, and almost simultaneous MOI of two independent spacecraft will be a quite complicated and risky operation for the ground station. If we intend to shift MOI timing between two spacecraft, considerable amount of delta- V will be required in the cruising phase. As for the mass impact of an OME on each craft, we found that there is not so large difference compared to the fuel mass needed in IBS10). We need further quantitative trade-off assessment for operational impacts in the future works. 5 Dec. 2024 Establishment of 7. Summary Orthogonality This paper described the preliminary mission analysis and orbit design for Japanese next Mars exploration program named MELOS. The scientific objectives, outlines of the mission and several topics about orbits were briefly described. Note that these descriptions are just a tentative candidate, and they are to be modified, confirmed and determined through ongoing discussion among researchers.

Acknowledgments 6 Jan. 2025 Descent of In-Situ Orbiter The authors are deeply grateful to Dr. K. Fujita, Mr. Y. Saitoh and Ms. C. Hirose in Japan Aerospace Exploration Agency and Mr. T. Yamaguchi in The Graduate University for Advanced Studies, for enlightening discussions on aerodynamics, launch vehicle capability and orbital dynamics. The authors also greatly appreciate helpful advices and suggestions given by anonymous reviewers.

Fig. 12. Orbit maneuver sequence for establishment of orthogonal constellation in 2-orbiter constellation. 6

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