第 3 回 月火星着陸探査シンポジウム

火星エアロキャプチャミッション
2018/3/15
Kazuhisa Fujita, Ryuta Hatakenaka, Kentaro Nishi, Toshinori Ikenaga,Toshiyuki Suzuki, Misuzu Haruki, Takashi Ozawa, Satoshi Nomura
Japan Aerospace Exploration Agency
藤田和央, 畠中龍太, 西顕太郎, 池永敏憲,
鈴木俊之,春木美鈴, 小澤宇志, 野村哲史

宇宙航空研究開発機構
Tomohiro Usui
Tokyo Institute of Technology
臼井寛裕

Hirdy Miyamoto
Graduate School of the University of Tokyo宮本英昭

東京工業大学

東京大学大学院
Shingo Kameda
Rikkyo University
亀田真吾

立教大学

第 3 回 月火星着陸探査シンポジウム

Purpose, Goal, & Objectives of Mars Aerocapture Orbiter 2 Strategy of Mars ExplorationIn the situation that not only NASA and ESA have performed but also Asian countries areplanning Mars exploration in 2020s, if Japan launches a Mars exploration, its significance,purpose, and outcome will be strongly questioned. The key word may be novelty and originality,and bold challenges with high creativity will be needed for exploration engineering.
Meanwhile, to realize Mars surface mission that the JSPS supports, secure and step-by-stepacquisition of EDL technology is needed (security & reliability > novelty & originality).
Mars Aerocapture Obiter as a small-class mission must be a best solution for this question,
following MMX, leading to Mars surface missions, and performing high cost effectiveness.

 Mission Objectives and Expected Results
Mission Objectives
To insert a Mars orbiter by aerocapture

Expected Results
Partial demonstration of EDL technologies required
for future surface missions (technical continuity),
and the world’s first acquisition of AC technology.

To provide platform functions for surface International cooperation for existing platforms /missions (data relay, EDL support etc.)
provision for future Japan’s surface missions.
To provide platform functions for long- Realization of remote-sensing, Martian spaceterm scientific observation of Martian weather & climate exploration observation, etc.
environments
Acquisition of information required for landing point
selection in future surface missions.

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Casting in Exploration Roadmap (Proposed)
Small bodies
Comets2015
Procyon
HAYABUSA2

DESTINY

Jupiter
Trojan SR

S #4

M #2

Sample return

M #1

MMX

SLIM

Mars
exploration
with Rover
HTV-M

S #3

SELENE-R

HLEPP

Soft-landing to gravitational body
Image navigation

EarthS #6

MACOMarsMoon

2030

International SR

Deep space transportation by IE
Image navigation

2025

2020

Obstacle avoidance

HTV SRC
Aeroassist guidance

HTV-X

Light-weight TPS

ISS

Human lunarexploration
M #3

SPICA

International human Mars missions

Target3

第 3 回 月火星着陸探査シンポジウム

Orbit Planning (1/2)
 Mars Transfer Orbit

Launch @ 2026/10/31, Arrival at Mars on 2027/08/20 （0 rev. about sun）
Arrival C3 = 7.36 km2/s2 （𝑉𝑉∞ = 2.713 km/s）

 Launch Capability of Epsilon


Similar to SLIM （using KM）
Assessed as TMI from SLIM IForbit（250 x 19,000 km）
Launch epoch2026 window

Departure C3

9.18 km2/s2

TMI ΔV

1.61 km/s

TMI WET Mass（Isp =310 s）

323 kgArrival C3

Arrival 𝑉𝑉∞

Flight time

7.36 km2/s2
2.713 km/s
293 days4

第 3 回 月火星着陸探査シンポジウム

Orbit Planning (2/2)

5 Visibility & Communicability on ArrivalDistance from Earth is 2.2 AU (relatively long but communicable)
SEP = 38º (not bad). Conjunction occurs 2.5 month after MOI (late enough for initial operation).
ArrivalArrival

第 3 回 月火星着陸探査シンポジウム

Aerocapture Corridor Assessment (1/2)

6 Trans-Mars OrbitLaunch on 2026/10/31, Mars arrival on 2027/8/20（𝐿𝐿s = 148.5） as previously defined (0 rev.)
Arrival 𝑉𝑉∞ = 2.713 km/s，

 Mars Atmosphere ModelNominal, High-density, Low-density models are developed for arrival epoch by using MarsGRAM 2005 v13．

Nominal model is defined as a daily and global average, while high-density and low-density
models are defined as upper and lower limits with +5σ and –5σ dissipations, respectively.

 Mars Gravity ModelMartian surface is assumed as a spheroid.

Mars gravity is assumed as a mass point located at the spheroid center.

 Aerocapture Procedure & Target OrbitBased on mission requirement from Martian Space Weather & Climate Exploration Mission, a300×5,000 km polar orbit synchronized with Mars rotation is selected as the target orbit.

After deceleration by atmosphere, periapsis raise maneuver is conducted to raise the periapsis
altitude to 300 km at the first apoapsis, then apoapsis altitude is adjusted to 5,000 km by
apoapsis adjust maneuver at the second periapsis.

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Assessment of Aerocapture Corridor (2/2)

7 Constraint Conditions on Post-Aerocapture Maneuver
1. The total post-aerocapture ΔV (the sum of periapsis raise and apoapsis adjust maneuver ΔV)
should be less than 100 m/s. This constraint is required for aerocapture to be superior to
propulsive orbit insertion with respect to mass budget.
2. The first apoapsis after atmospheric exit should be less than 30 Rm (the orbiter should besafely captured by Martian gravity). This is automatically satisfied when 1 is fulfilled.
3. Flight time from atmospheric exit to the first apoapsis should be greater than 30 minutes. This
period is required for attitude determination and attitude control prior to periapsis raisemaneuver.
4. Corridor width should be wider than 3º.
According to past studies, orbit
determination by using DSN realizes
flight path angle accuracy within 0.1º.

 Numerical Method

Parameters

UnitErrors (±3σ)

Position error (X)m12459.5

Position error (Y)m3058.6

Position error (Z)m

5482.5Velocity error (magnitude)

m/sec

8.356

Velocity error (flight path)

deg

0.06

Velocity error (flight azimuth)

deg

0.04



Flight dynamics code SDF28-012 is used in the mass-point mode (3-degrees-of-freedom).

Range of flight path angle acceptable for aerocapture is determined by changing the ballisticcoefficient from 10 to 1,000, and the lift-to-drag ratio (L/D) from –0.4 to +0.4.

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Mars Atmosphere Model MarsAtmosphere-20270820

8

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Aerocapture Corridor & Design Point Design Point

β = 85.7 kg/m2

|L/D| = 0.2



γ = – 13.28º

 Corridor Characteristics
Δ γ = 3.2º （> 3.0º）Post aerocapture ΔV <100.0 m/s(minimum = 46.3 m/s)

9

第 3 回 月火星着陸探査シンポジウム

Aerodynamic Shape Definition (1/2)
 Influential Parameters for Aerodynamic Shape

Dependence of L/D on angle of attack (linear)
Sensitivity to center of gravity (small)
Aerodynamic heat transfer rate (small)

Stability around trim angle (high)Accommodatability (high)Surface to volume ratio (low)10第 3 回 月火星着陸探査シンポジウム

Aerodynamic Shape Definition (2/2)
 Aerodynamic Shape Optimized for AerocaptureHalf cone angle of 60º is favorable rather than 70º (Viking-type aeroshell).

11

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Aerocapture Orbit Baseline
 TRJ-20270820-180110


Peak heating is about 0.32 MW/m2 at stagnation point. Total heat is 42 MJ/m2.
Peak dynamic pressure is approximately 1.3 kPa.
Baseline orbit data are given as TRJ-20270820-180110.dat

12

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Summary of Design Parameters
Parameters
𝑉𝑉∞ , km/s

13Specification

Remarks

2.713

Total mass (WET), kg
Aeroshell mass, kg
Others, kg

323
45+24254

Forebody + Aftbody

Aeroshell diameter, mm

2,000

Storable in Epsilon fairing

Aeroshell height, mm

1,200

Storable in Epsilon fairing

Atmospheric entry I/F altitude, km

125Ballistic coefficient (β), kg/m2

85.7

CD

1.20

L/D

0.2

Flight path angle, deg−13.28ºCorridor width (Δγ), deg

0.32

Periapsis altitude (target), km

300

Apoapsis altitude (target), km

5,000

Apoapsis altitude range (3σ), km

4,870 to 5,050

CDS = 3.77 m2
Trim angle of attack = −18º
> 0.3º (requirement)

Adjusted after MOI

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Communication Subsystem Design14 Orbiter-GroundAUbi t ra te (bps )TLMma rgi n (dB)
1.2s ta tus
ma rgi n (dB)ca rri er
s ta tus
bi t ra te (bps )TLM
ma rgi n (dB)
1.54s ta tus
ma rgi n (dB)ca rri er
s ta tus
bi t ra te (bps )TLM
ma rgi n (dB)
2.25
s ta tus
ma rgi n (dB)ca rri er
s ta tus

XHGA
φ0.3m
φ0.4m
1024
2048
2.3
1.8
OKOK
4.3
6.8
OKOK
512
1024
3.1
2.6
OKOK
6.9
4.6
OKOK
256
512
2.3
2.1
OKOK
10.2
10OK
OKXMGA(Bores i ght) (±12.5deg)
256
128
3.1
2.4
OKOK
3.9
4.2
OKOK
128
64
3.8
2.9
OKOK
5.7
4.7
OKOK
64
16
3.2
4.1
OKOK
56.1
OKOK

XLGA (bea con)
(Bores i ght) (±60deg)
N/A
6.2
OK

0.2
NG
N/A

4
OK-2
NG
N/A

0.7
NG-5.3NG

 Mars Surface – Orbiter

UHF Channel based on Martian data-relay standard
By using Electra-Lite manufactured by L3 CCE, which is a standard compliant product, a bit rate
higher than 64 kbps is possible.

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Block Diagram15

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System Configuration
Subsystem

Mass, kg

16Remarks

Aeroshell
Forebody
Aftbody69

COM
X-band (LG, to Earth)X-band (HG, to Earth)U-band (to Mars)

40

AOCS

14.4

DHS

10.6

RCS
Propellant
HW (4N RCS ×12)
Tank & tubing

59

TCS

5EPS

23

SAP = 1.5 m2, 1.2kWh/sol
300 Wh Li-ion battery

STR

38

Combined with aeroshell

INT

14

Scientific Instruments

30

Margin

20

Total

323

45
24
0.4
22.6
1732
522

Advanced NALT aeroshell isassumed.

Conceptual View of MACO
Redundant 2 system
Aftbody
aeroshell（deployed）

Combined system, 2-liquidForebody
aeroshellRCSSAP
Orbiter body
ANT（deployed）
Mars Space Weather &Climate Instruments etc.

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Comparison to Propulsive Orbit Insertion17

 Direct MOI by Chemical Propulsion

ΔV required for direct MOI into a 300×5,000 km orbit with 𝑉𝑉∞ = 2.713 km/s is 1.514 km/s.
Propellant mass ratio for propulsion system with Isp = 315 sec is 38.8 %. Assuming structuralcoefficient of 0.25 for the propulsion system, mass ratio of the propulsion system is 48.5 %.
Mass ratio of the aerocapture subsystem is estimated to be 42.4% in a conservative analysis,
even considering all the propulsion subsystem HW, which is expected to be more advantageousthan MOI by the propulsion system.
Note: this advantage is degraded when propulsive MOI is combined with aerobraking after MOI,
since aerocapture MOI cannot enjoy mass reduction by aerobraking.

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Technical Problems18

 Deployment TechnologyFor thermal control and power generation, the aftbody aeroshell has to be deployed during
interplanetary flight, closed before atmospheric entry, and redeployed after atmospheric exit.
Genesis is a good reference.

 Communication SubsystemA small-sized X-band communication system with high-gain antenna is needed.
A U-band communication system compliant with Mars communication standard is needed.
Currently, the conceptual study result by MEISEI ELECTRIC CO., LTD. is used.
A gimbal mechanism to control the direction of antenna may be needed.

 Attitude Orbit Control SubsystemA small-sized AOCS with RWs is needed, since three axis attitude control is preferred for datarelay and science missions.

 Aeroshell Subsystem
Further lightweight aeroshell than current BBM is desirable.

 Propulsion SubsystemCurrent RCS is applicable, but embedment of RCS in deployable aeroshell is difficult problem.

 Redundancy

HW redundancy is applied only to COM and DHS because of limited mass budget. Other type ofsmall-sized bus system with sufficient redundancy will be entertained in the future.

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Summary19

 Conceptual Study of Mars Aerocapture Orbiter (MACO) has been conducted.
 Based on an Epsilon-Enhanced Launcher Equivalent to SLIM (combined with a KM), MTI
Mass as Large as 323 kg is Expected for 2026 Launch Window.
 Within a Mass Budget of 323 kg, an Aerocapture Orbiter Including 30 kg of Scientific
Instruments (Mars Space Weather and Climate Mission) is Considered to be Possible.
 Major Key Technical Issues to be Solved Have Been Extracted.
 From Here on, in Cooperation with Mars Remote Sensing Mission, Mars Space Weatherand Space Climate Exploration Mission, and Mars Atmosphere Exploration Mission etc.,
More Detailed System Study will be Conducted to Create a Small Science Mission
Proposal to be submitted to AO in 2018 or later.