Launch Opportunities and Preliminary Orbit Design for Next Mars Exploration Program - J-Stage
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Trans. JSASS Aerospace Tech. Japan
Vol. 10, No. ists28, pp. Tk_19-Tk_25, 2012
Topics
Launch Opportunities and Preliminary Orbit Design
for Next Mars Exploration Program
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 re-
searchers 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 RM and RM is the Mars radius. Note that the dry
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
1
Copyright© 2012 by the Japan Society for Aeronautical and Space Sciences and ISTS. All rights reserved.
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Trans. JSASS Aerospace Tech. Japan Vol. 10, No. ists28 (2012)
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 Cr to be
04:20:46 UTC apsis
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
2
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N. OGAWA et al.: Launch Opportunities and Preliminary Orbit Design for Next Mars Exploration Program
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
insertion by a single impulse5, 6) . 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
0.94
0.9
0.96
1
900
Fig. 3. Mars orbit insertion of the meteorology orbiter.
800
Periapsis altitude [km]
30800 30800 365
0.86
308 360
0.88
700
0.98
0.84
30750
0.92
30700 355
0.94
0.9
306
0.96
1
30700 350
304
30600 600
345
302
30650 340
Feb 15 Mar 1 Mar 15 Apr 1 500
2
335
0.8
30600
330
0.86
0.88
0.98
0.84
30550 325
400
0.92
0.94
0.9
320
0.96
1
30500
315
300
30450
310 1 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9
305 Vinf / Vesc
30400 300
2021
Apr Jul Oct Jan 2022
(UTCG)
Apr Jul Oct Jan 2023
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
3
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Trans. JSASS Aerospace Tech. Japan Vol. 10, No. ists28 (2012)
1000 5.1. Transfer to Mars
0.9
1.3
1.4 1.5
In this paper, we assume the launch year to be 2022. A
1
1.1
1.2
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
Periapsis Altitude [km]
700 constellation is shown in Fig. 9. On 2nd September 2022, the
0.9
spacecraft will be launched from Tanegashima Space Center,
1.3
1.4 1.5
1
1.1
1.2
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
declination of the outgoing asymptote are 80.1 degrees and 4.45
400
degrees, respectively. A small burn of TCM-1 is performed
0.9
1.3
1.4 1.5
1
1.1
1.2
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 RM and
required to establish orthogonal constellation. In order to save
17,000 km × 50 RM , the delta-V is only 1,334 km/s. As Table
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
5. Preliminary Orbit Design for 2-Orbiter Constellation orbiter, while that in the in-situ orbiter is −0.005 deg/day,
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 In this section, we used the J2 value on Rika-Nenpyo9) , for simplicity. The
rate of -0.19 deg/day, helping the maintenance of constellation.
J2 value derived from GMM-2B3) , or normalized C20 is 1.956 × 10−3 , which is
consistent with the value we used here within the scope of this paper.
4
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N. OGAWA et al.: Launch Opportunities and Preliminary Orbit Design for Next Mars Exploration Program
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 RM → 7 RM
Inclination i = 63.4 deg → 84.4
Maneuver deg
Precession ∼ 483 days
Descent 30 RM → 7,000 km
Fig. 9. Mars transfer orbit for 2-orbiter constellation.
5
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Trans. JSASS Aerospace Tech. Japan Vol. 10, No. ists28 (2012)
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
the two orbiters are inserted into the initial Mars orbit, and then
In-Situ Orbiter
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.
However, SBI has two potential difficulties; each spacecraft
Precession of
needs its own OME, and almost simultaneous MOI of two
Remote-Sensing Orbiter 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
7. Summary
Establishment of
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.
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