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k-57.inddTrans. JSASS Aerospace Tech. Japan Vol. 12, No. ists29,
pp. Pk_43-Pk_50, 2014
Original Paper
Pk_43
Mission Analysis of Sample Return from Jovian Trojan Asteroid by
Solar Power Sail
By Jun MATSUMOTO1), Ryu FUNASE1), Osamu MORI2), Yoji SHIRASAWA2),
Go ONO1), Taku HAMASAKI1), Naohiro HAYASHI1), Toshihiro CHUJO1),
Norizumi MOTOOKA1) and Keita TANAKA1)
1)Department of Aeronautics and Astronautics, The University of
Tokyo, Tokyo, Japan
2)Japan Aerospace Exploration Agency, Sagamihara, Japan
(Received June 10th, 2013)
Japan Aerospace Exploration Agency (JAXA) are planning an outer
solar system exploration and sample return
mission from a Jovian Trojan asteroid using a 3000 m2 solar power
sail. The difficulty of this mission is a severe restriction on the
weight; only 300 kg is allocated for sampling and returning to
Earth. In this weight, fuel for trajectory and attitude control,
sampling mechanism, re-entry capsule, and other systems required to
return to the Earth are included. In this paper, a preliminary
analysis of this sample return mission is conducted. Three
scenarios for sampling are proposed; sampling with a 3000 m2 solar
power sail, with a detachable small solar power sail using electric
propulsion systems and with a small probe using chemical propulsion
systems. The mission analysis shows that the most feasible
configuration is to conduct the sampling with the 3000 m2 solar
power sail using an extension mast.
Key Words: Jovian Trojan Asteroid, Sample Return, Solar Power Sail,
Thin-film Solar Cell, Electric Propulsion
Nomenclature
t : time tf : final time r : position
V : velocity change Is : moment of inertia around spin axis It :
moment of inertia around normal axis
to the spin axis
1. Introduction Japan Aerospace Exploration Agency (JAXA) are
planning an outer solar system exploration and a sample return
mission using a solar power sail spacecraft with a sail area of
3000 m2 (Fig.1). It is a single-spin spacecraft. The
solar-power-sail membrane is fully covered with thin-film solar
cells to gain sufficient electric power to drive ion engines at the
outer solar system. The sail area is determined to be 3000 m2 in
order to satisfy the electric power supply requirement1). Since
this spacecraft has the sail and the ion engines, it is considered
to be a succession of the world’s first asteroid sample return
mission HAYABUSA2) and the world’s first solar power sail IKAROS3).
Main features of this mission are - The world’s first sample return
mission from a Jovian Trojan asteroid. - The world’s highest
performance ion engines. - The world’s first hybrid propulsion with
solar photon acceleration and electric propulsion. - The world’s
first observation of the infrared background radiation from the
outer solar system.
Fig. 1. 3000m2 solar power sail.
JAXA plan to initiate this project in a few years and expect a
launch in 2022. The target celestial body of this mission is
farthest from the Earth and mission period of 20 years is very long
in comparison to other sample return missions2,4-6). In this paper,
a preliminary mission analysis of the sample return from the Jovian
Trojan asteroid is conducted. This sample return mission is
challenging because the weight budget for the sample return is
restricted. A system analysis of the 3000 m2 solar power sail1)
shows that the weight allocated for sampling and returning to Earth
is only 300 kg. Within this weight budget, a system for the sample
return has to be designed. In this paper, two configurations for
sample return are mainly discussed; sampling with a 3000 m2 solar
power sail, or with a detachable small solar power sail using
electric propulsion systems. In the former scenario, fuel for
trajectory and attitude control, a sampling mechanism and
re-entry
Copyright© 2014 by the Japan Society for Aeronautical and Space
Sciences and ISTS. All rights reserved.
Trans. JSASS Aerospace Tech. Japan Vol. 12, No. ists29 (2014)
Pk_44
Table 1. Weight budget of the solar power sail. Wet Weight 1550
kg
Sail and Sail Expansion Mechanism 500 kg Bus Components 350
kg
Body 150 kg Science Components 50 kg
RCS Fuel (Outward only) 140 kg Ion Engine Fuel (Outward only) 60
kg
Weight for Sample Return 300 kg
Table 2. Configurations of the sample return spacecraft.
Sampling
& Returning Spacecraft
Sampling Strategy
(A-1) Extension Mast (A-2) 3000 m2 sail Electric Propulsion
Impactor (A-3) Tiny Probes (B-1) Extension Mast (B-2) Small sail
Electric Propulsion Impactor (B-3) Tiny Probes (C-1) Small probe
Chemical Propulsion Sampler Horn
capsule have to be designed within 300 kg. In the latter scenario,
whole system of the detachable small solar power sail including
fuel has to be designed within 300kg.
Moreover, characteristics of Jovian Trojan asteroids make it
difficult to realize this sample return mission. - The Jovian
Trojan asteroids are far from the Earth (>5 AU) so a
communication delay is large. The sampling has to be conducted
autonomously by the spacecraft. - The Jovian Trojan asteroids are
large (the diameter of these asteroids is of the order of 101 km –
102 km) and the gravity of these asteroids is strong. Fuel
consumption would be, therefore, significant in a touch-down
phase.
In this paper, mission sequences including new sampling strategies
and touch-down strategies are proposed to realize this sample
return mission. The purpose of this analysis is to determine the
minimum requirements for the mission. A system design which is
constructed by considering only an outward journey is, therefore,
used as preconditions and iteration processes to construct the
whole system design are not included in this analysis. All V of the
trajectory control mentioned in this paper do not include outward
trajectory control V. 2. Configurations for the Sample Return
2.1. Outward phase A planned scenario is as follow1);
- Phase 1: Electric Delta-V Earth Gravity Assist (EDVEGA) - Phase
2: Transfer from the Earth to Jupiter - Phase 3: Rendezvous with a
Jovian Trojan asteroid The value of C3 for the launch is assumed to
be 28 km2/s2
required to conduct 2-rev EDVEGA (Phase 1). After Jupiter swing-by
(Phase 2), the ion engines are driven to make V required to
rendezvous with a Jovian Trojan asteroid (Phase 3). Table 1 shows
the weight budget of the spacecraft.
Fig. 2. Scenario (A).
Fig. 3. Scenario (B).
Fig. 4. Scenario (C).
2.2. Mission scenarios for the sample return In this paper, seven
mission scenarios shown in Table 2 are discussed.
In scenario (A), the 3000 m2 solar power sail approaches the
asteroid and returns to the Earth with electric propulsion (Fig.
2.)
In scenario (B), a small solar power sail is separated from the
3000 m2 solar power sail near the asteroid. This small solar power
sail approaches the asteroid, and returns to the Earth with
electric propulsion by itself (Fig. 3). This small probe is
equipped with a small solar-power-sail membrane in order to gain
sufficient electric power to drive the ion engines at the outer
solar system. In scenario (C), a small probe approaches the
asteroid under the support of the 3000 m2 solar power sail. After
obtaining samples, this small probe produces V and returns to the
Earth by itself (Fig. 4). This scenario is not, however, feasible
because of the amount of fuel required for the trajectory control
maneuver. A preliminary trajectory analysis shows that the average
required V is 5.5 km/s. If the weight of this
J. MATSUMOTO et al.: Mission Analysis of Sample Return from Jovian
Trojan Asteroid by Solar Power Sail
Pk_45
Fig. 5. Extension mast.
Fig. 6. Impactor. small probe is 300 kg, the weight budget
excluding the fuel is only about 30 kg. Within this weight budget,
it is difficult to design the spacecraft with a re-entry
capsule.
Scenarios including re-docking after the touch-down can be
considered. However, it is difficult to realize docking around a
Jovian Trojan asteroid in this mission because - The 3000 m2 solar
power sail is spinning. - Communication delay is large so that the
docking has to be conducted autonomously. - There is a severe
weight restriction and it is difficult to make a light-weight
docking mechanism.
Scenario (A) and scenario (B) are, therefore, discussed mainly in
this paper. 2.3. Sampling strategy In scenario (A) and scenario
(B), it cannot get close to the asteroid because there is a risk
that the huge solar-power-sail membrane touches the surface of the
asteroid. A sampling strategy, therefore, has to be considered by
which the sampling is conducted at a certain distance. In this
paper, three sampling strategies are proposed; an extension mast,
an impactor, and tiny probes. - Extension Mast
A long extension mast such as a convex tape or a coilable mast is
equipped (Fig. 5). At first, the solar power sail extends the
convex tape or the coilable mast at a high altitude. Subsequently,
the touch-down is conducted with the mast extended. There are,
however, some difficulties in this strategy because the spacecraft
has the thin-film solar sail; (1) The solar sail is flexible so the
spin rate has to be high to keep its shape flat. (2) In order not
to touch the surface of the asteroid, the extension mast has to be
very long.
Fig. 7. Tiny probes.
Fig. 8. Touch-down sequence.
- Impactor The spacecraft throws an impactor down which explodes
on
the surface of the asteroid (Fig. 6). This explosion blows up some
dust for the spacecraft to catch. In order to catch the dust, the
spacecraft may not be able to avoid aftereffects of the explosion.
There is, therefore, a possibility of damaging the spacecraft. One
of the purposes of this sample return mission is to search for ice
and organic material in the asteroid. The explosion degrades this
scientific objective so this idea is not permitted in this mission.
- Tiny probes
The spacecraft throws many tiny probes down. These tiny probes
obtain samples on the surface of the asteroid. Subsequently, they
jump up using a force of a spring and the spacecraft catches at
least one of them (Fig. 7). The tiny probes have to jump up
accurately to the spacecraft with unknown surface conditions of the
asteroid so the success probability of this strategy may be,
therefore, very small. It is impossible to explore the surface
condition in the design phase of the tiny probes. It is, therefore,
difficult to increase the success probability of the
sampling.
It is, therefore, suggested that the most feasible candidate is the
extension mast strategy because it has no critical defect. 2.4.
Touch-down strategy In the sample return mission, the spacecraft
must approach the surface of the asteroid. In the HAYABUSA mission,
the spacecraft approached the asteroid slowly not to crash into the
asteroid7). Since the target asteroid Itokawa is relatively small,
the required fuel which was consumed to cancel the gravity of the
asteroid was small for the touch-down. The size of Jovian Trojan
asteroids which have been discovered is, however, relatively large.
If the spacecraft approaches the asteroid
Trans. JSASS Aerospace Tech. Japan Vol. 12, No. ists29 (2014)
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Fig. 9. Arrangement of ion engines (spin stabilized control
strategy).
Fig. 10. Throttling control strategy. slowly, the amount of
required fuel increases exponentially. The spacecraft, therefore,
has to approach the asteroid rapidly in this mission.
In this paper, a new touch-down strategy is proposed. This strategy
is located between a touch-down strategy on a microgravity asteroid
and a landing strategy onto the Moon. The touch-down strategy is
designed as follows (Fig. 8); -Step1: The spacecraft falls freely
from a starting point of the touch-down phase. -Step2: The
spacecraft decelerates using chemical thrusters in order that the
velocity is decreased to 1.5 m/s at a specific altitude from the
surface of the asteroid. -Step3: The spacecraft moves at a constant
velocity by canceling the gravity using RCS and the touch-down is
conducted. -Step4: The thrusters are operated at full power, and
the spacecraft escapes from the asteroid. The velocity value of 1.5
m/s is designed by referring to the moon landing strategy.
Moreover, in this paper, the fuel for the attitude control in the
touch-down phase is estimated as 15 % of the fuel for the position
control. 2.5. Attitude control strategy of the solar power sail
with electric propulsion system
As observed in IKAROS mission8), since solar radiation pressure
torque affected on the deformed sail was remarkably large, much RCS
fuel was required to control the attitude of the solar power sail.
In this section, therefore, new attitude control strategies for the
solar power sail with the electric propulsion system are introduced
to reduce the fuel for attitude control.
At first, the arrangements of the ion engines are explained. Fig. 9
shows an arrangement of the ion engines when the spin stabilized
attitude control strategy is applied. These ion engines have a
1-axis gimbal (parallel direction to the sheet
Fig. 11. Arrangement of ion engines (3-axis stabilized control
strategy).
Table 3. Computational conditions: scenario (A). Model Sun-centered
two-body problem
Planet orbits (coplanar) Earth: Circular orbit (r = 1 AU) Asteroid:
Circular orbit (r = 5.2 AU)
Boundary Conditions t = 0: Position and velocity are the same as
the asteroid
t = tf: Position is the same as Earth Constraints (1) Sun angle
< 45 deg
(2) Infinity velocity with respect to Earth < 9 km/s (Re-entry
velocity < 15 km/s)
Initial Mass 1300 kg Max Thrust Magnitude 150 mN
Isp 6000 s Computational strategy DCNLP ( V is optimized)
Table 4. Results of the trajectory design: scenario (A). Flight
Time [year] 8.26
Infinity Velocity [km/s] 8.61 V [km/s] 9.33
Fuel Mass [kg] 191 and the sail). By the gimbal control, the spin
rate can be controlled. By throttling control synchronized with the
spin (Fig. 10), it is possible to re-orient the spin axis.
Fig. 11 shows an arrangement of the ion engines when the 3-axis
stabilized attitude control strategy is applied. These ion engines
have a 2-axis gimbal. By the gimbal control, a control torque
vector in a plane which is normal to the thrust direction can be
generated.
Disturbance torque such as solar radiation pressure torque8) or
swirl torque can be canceled as follow; - Spin stabilized attitude
control strategy When the ion Engines are on;
Windmill torque can be canceled by the 1-axis gimbal control and
other disturbance torque can be canceled by the spin synchronized
throttling control strategy. When the ion engines are off;
The spin axis direction converges to an equilibrium point
automatically so the control is not required8). The spin rate
control is conducted by RCS. - 3-axis stabilized attitude control
strategy When the ion engines are on;
The disturbance torque which is normal to the thrust direction can
be canceled by the 2-axis gimbal control. The disturbance torque
which is parallel to the thrust direction cannot be canceled by
using ion engines. It is, therefore, necessary to use RCS to cancel
them.
J. MATSUMOTO et al.: Mission Analysis of Sample Return from Jovian
Trojan Asteroid by Solar Power Sail
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Fig. 12. Return trajectory (J2000EC): scenario (A).
Table 5. Specifications of the 3000 m2 solar power sail: scenario
(A). Main Body (located in the center of the sail)
Mass 1087 kg Shape Cylinder
(radius: 1.5 m, height: 2.5 m) Moment of Inertia Is = 1223 kgm2, It
= 1178 kgm2
Sail Mass 213 kg Shape Disk
(S = 3000 m2, deflection: 1 deg, torsion: 0.4 deg)
Optical Characteristics Thin-film solar cells Moment of Inertia Is
= 1.64×105 kgm2,
It = 8.20×104 kgm2 RCS
Isp 210 s Arm Length 1.5 m
When the ion engines are off;
All perturbation torque have to be canceled with RCS. The
differences between these two strategies are the fuel required for
the trajectory control and the attitude control. The 3-axis
stabilized solar power sail can effectively drive ion engines to
change the trajectory; on the other hand, spin-stabilized solar
power sail can effectively cancel the solar radiation pressure
torque and some other disturbance torque applied on the solar power
sail by using ion engines. 2.6. Re-entry capsule
In the sample return mission from the Jovian Trojan asteroid, the
infinity velocity relative to the Earth is very high. Therefore
high-performance re-entry capsule has to be developed. In this
analysis, as an example, the weight of the re-entry capsule is set
to be 30 kg. This is based on the Marco-Polo sample return
mission9).
3. Mission Analysis
In this chapter, a detailed mission analysis is conducted and the
weight budgets of the two scenarios are shown.
Table 6. Specifications of the target asteroid of scenario (A) and
(B). Target Asteroid Eurymedon
Diameter 18.3 km Density 2.0 g/cm3
Table 7. Computational conditions of scenario (A).
Starting Point of the Touch-Down 500 km (Altitude) Touch-Down Point
50 m (Altitude)
Thrust of Reaction Control System 80 N Isp 210 s
Table 8. Weight budget result of scenario (A).
Weight Restriction 300 kg Fuel for Trajectory Control 191 kg Fuel
for Attitude Control 4 kg
Fuel for Touch-Down 44 kg Capsule 30 kg
Sampling Mechanism 31 kg
Table 9. Computational conditions: scenario (B). Model Sun-centered
two-body problem
Planet orbits (coplanar) Earth: Circular orbit (r = 1 AU) Asteroid:
Circular orbit (r = 5.2 AU)
Boundary Conditions t = 0: Position and velocity are the same as
the asteroid
t = tf: Position is the same as Earth Constraints (1) Sun angle
< 45 deg
(2) Flight time < 6 years (3) Infinity velocity
with respect to Earth < 9 km/s (Re-entry velocity < 15
km/s)
Initial Mass 300 kg Required
Thrust Magnitude 16.8 mN (Spin stabilized attitude Control)
12.6 mN (3-axis stabilized attitude Control) Isp 2900 s
Computational strategy DCNLP ( V is optimized)
In return trajectory analyses, two type trajectories, with and
without Jupiter swing-by, are calculated. These is, however, no
difference in these trajectories from a perspective of mission
period. The trajectories without Jupiter swing-by are, therefore,
shown in this paper. 3.1. Scenario (A): Sample return by 3000m2
solar power sail In this section, the mission analysis of scenario
(A) is conducted. - Return trajectory
In this configuration, the 3000 m2 solar power sail returns to
Earth by using ion engines. Table 3 shows the computational
conditions. Constraint (1) in Table 3 means that the solar power
sail must be oriented to the Sun in order to gain sufficient
electric power. An example trajectory is shown in Fig. 12. The
required V is 9.33 km/s and the required fuel is 191 kg (Table 4).
This trajectory designed is used for an estimation of the required
fuel for attitude control later. - Fuel for the attitude control
The fuel required for the attitude control is estimated10). The
computational condition is summarized in Table 5. In this
configuration, the required fuel is 4.02 kg.
Trans. JSASS Aerospace Tech. Japan Vol. 12, No. ists29 (2014)
Pk_48
Fig. 13. Return trajectory (spin stabilized attitude control,
J2000EC): scenario (B).
Table 10. Results of the trajectory design: scenario (B).
Spin stabilized Attitude control
3-axis stabilized attitude control
Max Thrust [mN] 16.8 12.6 Flight Time [year] 5.8 6.0
Infinity Velocity [km/s] 8.79 8.86 V [km/s] 7.93 6.65
Fuel Mass [kg] 73.0 62.5
Table 11. Specifications of the thin-film solar cells: scenario
(B). Weight 1800 W / kg at 1AU
Efficiency 120 W / m2 at 1AU
- Fuel for the touch-down
In this paper, asteroid Eurymedon is set to be a hypothetical
target asteroid. Table 6 shows the specification of this asteroid.
This asteroid is relatively small among all discovered Jovian
Trojan asteroid. In the touch-down phase, a considerable amount of
fuel is required to cancel the gravity of the asteroid if the size
of the asteroid is large. The target asteroid, therefore, has to be
small to realize this sample return mission within the restrict
weight budget. The computational conditions are summarized in Table
7. By a calculation based on the new touch-down strategy mentioned
in subsection 2.4, the required fuel is 14.6 kg for one touch-down.
In reality, rehearsal touch-down operations are required. In this
paper, the number of the rehearsals set to be two. Therefore the
total fuel is 43.8 kg. - Weight Budget
Weight Budget of scenario (A) is summarized in Table 8. Taking the
required weights from the 300 kg, the weight budget for the
sampling mechanism is 31 kg. In this scenario, it is necessary to
consider a manufacturing strategy of the sampling mechanism within
this weight budget.
Fig. 14. Return trajectory (3-axis stabilized attitude control,
J2000EC): scenario (B).
Table 12. Specifications of the small solar power sail: scenario
(B). Main Body (located in the center of the sail)
Mass 273 kg Shape Cylinder
(radius: 0.75 m, height: 0.85 m) Moment of Inertia Is = 76.7 kgm2,
It = 54.8 kgm2
Sail Mass 27.0 kg Shape Disk
(S = 200 m2, deflection: 1 deg, torsion: 0.4 deg)
Optical Characteristics Thin-film solar cells Moment of Inertia Is
= 859 kgm2, It = 429 kgm2
Ion Engines Arm Length 0.5 m Swirl Torque 2.0 Nm
RCS Isp 210 s
Arm Length 0.74 m 3.2. Scenario (B): sample return by small solar
power sail
In this section, the mission analysis of scenario (B) is conducted.
There are two possible systems in this configuration from a
perspective of the attitude control; a spin stabilized attitude
control strategy and a 3-axis stabilized attitude control strategy.
Therefore the comparison of these two control strategies is
discussed. - Return trajectory
In this section, return trajectory to the Earth is designed. Table
9 shows the computational conditions. Constraint (1) in the Table 9
means that the spacecraft has to be oriented to the Sun in order to
gain sufficient electric power. Constraint (2) is imposed from a
perspective of the required operation period. Preliminary analysis
shows that the spacecraft has to generate 16.8 mN (spin stabilized
attitude control strategy) or 12.6 mN (3-axis stabilized attitude
control strategy) around a Jovian
J. MATSUMOTO et al.: Mission Analysis of Sample Return from Jovian
Trojan Asteroid by Solar Power Sail
Pk_49
Table 13. Computational conditions of scenario (B).
Starting Point of the Touch-Down 500 km (Altitude) Deceleration
Point 4 km (Altitude)
Thrust of Reaction Control System 8 N Isp 210 s
Table 14. Weight budget result of scenario (B). Spin
stabilized
Attitude Control 3-axis stabilized Attitude Control
Weight restriction 300 kg Ion Engines 48.9 kg
Solar-Power-Sail Membrane 12.7 kg Boom 3.5 kg
Boom Expansion Mechanism 14.1 kg Primary buttery 1.7 kg
Reaction Control System 13.1 kg Re-entry Capsule 30 kg
Fuel for ion engines 73.0 kg 62.6 kg Fuel for RCS 0.066 kg 3.63
kg
Fuel for Touch-Down 15.6 kg Communication System 9.6 kg Power
Supply System 8.3 kg Date Handling Unit 5.0 kg
Attitude Orbit Control System 6.1 kg Electric System 10.0 kg
Thermal Protection System 5.0 kg Instruments for Touch-Down 21.2 kg
Body & Sampling Mechanism 22.1 kg 29.0 kg
Trojan asteroid to satisfy these constraints. The spacecraft,
therefore, has to equip a solar-power-sail membrane which can
generate sufficient electric power to generate these thrust
magnitudes. Fig. 13, Fig. 14 and Table 10 show the results of the
calculated trajectory. These results show that the 3-axis
stabilized attitude control strategy is advantageous from a
perspective of the required V . - Solar-Power-Sail Membrane
As mentioned above, the spacecraft has to equip a solar-power-sail
membrane to drive the ion engines. According to the return
trajectory analysis, 546 W of power is required to be generated to
drive the ion engines at the outer
Fig. 16. Hexagon type solar-power-sail membrane.
solar system. In order to satisfy this requirement, the area of the
solar-power-sail membrane becomes 200 m2 assuming the specification
of the thin-film solar cells which is under development (Table 11).
- Sail Deployment System
In this paper, a boom deployment strategy is applied as a sail
deployment strategy. This is because it takes little time to deploy
the solar-power-sail membrane. Moreover, since the booms are rigid,
the solar power sail will be more like to tolerate the impact of
the touch-down. The large sail, however, interferes with the
touch-down maneuver. In this analysis, an intermittent deployment
strategy11) is, therefore, introduced. The deployment is stopped
halfway until the end of the touch-down phase, and the remained
deployment is conducted just before leaving for the Earth to gain
sufficient electric power to drive the ion engines. In the
touch-down phase, the required electric power is small so the
mission can be continued if the sail is not fully deployed. In
order to realize the intermittent deployment strategy simply, two
types of sail can be considered; square type (Fig. 15) and hexagon
type (Fig. 16). Weight estimation of the mast is conducted in each
form and the result shows that the mast weight of the square type
is 9.49 kg and the mast weight of the hexagon type is 11.5 kg.
These weights are required not to be buckled by the force of
thrusters. These results show that the square type is advantageous.
- Fuel for the attitude control
The fuel required for the attitude control is estimated10). The
computational conditions are summarized in Table 12. The required
fuel for the spin stabilized attitude control strategy and the
3-axis stabilized attitude control strategy are 0.066 kg and 3.63
kg respectively.
These results show that the spin control strategy is advantageous
from a perspective of fuel consumption. - Fuel for the
touch-down
Also in scenario (B), Eurymedon is set to be the hypothetical
target asteroid. The computational condition is summarized in Table
13. By a calculation based on the new touch-down strategy, the
necessary fuel is 5.2 kg for one touch-down. The total fuel
required is 15.6 kg including two rehearsals.
Trans. JSASS Aerospace Tech. Japan Vol. 12, No. ists29 (2014)
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- Weight Budget Weight Budget of scenario (B) is summarized in
Table 14.
The weights of bus components are based on the weight budgets of
several spacecrafts2,9,12). Taking the required weights from the
300 kg, the weight budget of the body and the sampling mechanism is
22.1 kg (spin stabilized attitude control) or 29.0 kg (3-axis
stabilized attitude control) in this scenario. The 3-axis
stabilized attitude control is advantageous from a perspective of
the weight budget. This advantage is due to the arrangement of the
ion engines.
In any case, the weight budget for the body and the sampling
mechanism is small. It is, therefore, difficult to make the
spacecraft in this configuration in reality.
4. Conclusion
In this paper, the analysis of a sample return mission from a
Jovian Trojan asteroid is conducted. Among seven configurations,
the most feasible one is (A-1) the sampling by the 3000 m2 solar
power sail with the extension mast. By the analysis conducted in
this paper, the feasibility of this mission is shown.
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