Mission Analysis of Sample Return from Jovian Trojan Asteroid by Solar Power Sail

Transcription

1 Trans. JSASS Aerospace Tech. Japan Vol. 12, No. ists29, pp. Pk_43-Pk_50, 2014 Original Paper Mission Analysis of Sample Return from Jovian Trojan Asteroid by Solar Power Sail By Jun MATSUMOTO 1), Ryu FUNASE 1), Osamu MORI 2), Yoji SHIRASAWA 2), Go ONO 1), Taku HAMASAKI 1), Naohiro HAYASHI 1), Toshihiro CHUJO 1), Norizumi MOTOOKA 1) and Keita TANAKA 1) 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 m 2 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 m 2 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 m 2 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 t f : final time r : position V : velocity change I s : moment of inertia around spin axis I t : 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 m 2 (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 m 2 in order to satisfy the electric power supply requirement 1). 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 HAYABUSA 2) and the world s first solar power sail IKAROS 3). 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 m 2 solar power sail. JAXA plan to initiate this project in a few years and expect a launch in 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 missions 2,4-6). 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 m 2 solar power sail 1) 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 m 2 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. Pk_43

2 Trans. JSASS Aerospace Tech. Japan Vol. 12, No. ists29 (2014) 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 Fig. 2. Scenario (A). Table 2. Configurations of the sample return spacecraft. Sampling & Returning Spacecraft Propulsion in Returning Phase Sampling Strategy (A-1) Extension Mast (A-2) 3000 m 2 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 10 1 km 10 2 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 follow 1) ; - 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 km 2 /s 2 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. 3. Scenario (B). Fig. 4. Scenario (C) Mission scenarios for the sample return In this paper, seven mission scenarios shown in Table 2 are discussed. In scenario (A), the 3000 m 2 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 m 2 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 m 2 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 Pk_44

3 J. MATSUMOTO et al.: Mission Analysis of Sample Return from Jovian Trojan Asteroid by Solar Power Sail Fig. 5. Extension mast. Fig. 7. Tiny probes. 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 m 2 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 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. 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 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 asteroid 7). 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 Pk_45

4 Trans. JSASS Aerospace Tech. Japan Vol. 12, No. ists29 (2014) Fig. 9. Arrangement of ion engines (spin stabilized control strategy). Fig. 11. Arrangement of ion engines (3-axis 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 Attitude control strategy of the solar power sail with electric propulsion system As observed in IKAROS mission 8), 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 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 = t f : 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 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 torque 8) 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 required 8). 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. Pk_46

5 J. MATSUMOTO et al.: Mission Analysis of Sample Return from Jovian Trojan Asteroid by Solar Power Sail Table 6. Specifications of the target asteroid of scenario (A) and (B). Target Asteroid Eurymedon Diameter 18.3 km Density 2.0 g/cm 3 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 210 s Fig. 12. Return trajectory (J2000EC): scenario (A). Table 5. Specifications of the 3000 m 2 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 I s = 1223 kgm 2, I t = 1178 kgm 2 Sail Mass 213 kg Shape Disk (S = 3000 m 2, deflection: 1 deg, torsion: 0.4 deg) Optical Characteristics Thin-film solar cells Moment of Inertia I s = kgm 2, I t = kgm 2 Arm Length RCS 210 s 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 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 mission 9). 3. Mission Analysis In this chapter, a detailed mission analysis is conducted and the weight budgets of the two scenarios are shown. Table 8. Weight Restriction Fuel for Trajectory Control Fuel for Attitude Control Fuel for Touch-Down Capsule Sampling Mechanism Weight budget result of scenario (A). 300 kg 191 kg 4 kg 44 kg 30 kg 31 kg Table 9. Computational conditions: scenario (B). Model Planet orbits (coplanar) Boundary Conditions Constraints Initial Mass Required Thrust Magnitude Computational strategy Sun-centered two-body problem Earth: Circular orbit (r = 1 AU) Asteroid: Circular orbit (r = 5.2 AU) t = 0: Position and velocity are the same as the asteroid t = t f : Position is the same as Earth (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) 300 kg 16.8 mn (Spin stabilized attitude Control) 12.6 mn (3-axis stabilized attitude Control) 2900 s 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 Scenario (A): Sample return by 3000m 2 solar power sail In this section, the mission analysis of scenario (A) is conducted. - Return trajectory In this configuration, the 3000 m 2 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 estimated 10). The computational condition is summarized in Table 5. In this configuration, the required fuel is 4.02 kg. Pk_47

6 Trans. JSASS Aerospace Tech. Japan Vol. 12, No. ists29 (2014) Fig. 13. Return trajectory (spin stabilized attitude control, J2000EC): scenario (B). Fig. 14. Return trajectory (3-axis 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] Flight Time [year] Infinity Velocity [km/s] V [km/s] Fuel Mass [kg] Table 11. Specifications of the thin-film solar cells: scenario (B). Weight 1800 W / kg at 1AU Efficiency 120 W / m 2 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. 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 I s = 76.7 kgm 2, I t = 54.8 kgm 2 Sail Mass 27.0 kg Shape Disk (S = 200 m 2, deflection: 1 deg, torsion: 0.4 deg) Optical Characteristics Thin-film solar cells Moment of Inertia I s = 859 kgm 2, I t = 429 kgm 2 Arm Length Swirl Torque Arm Length Ion Engines RCS 0.5 m 2.0 Nm 210 s 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 Pk_48

7 J. MATSUMOTO et al.: Mission Analysis of Sample Return from Jovian Trojan Asteroid by Solar Power Sail Fig. 15. Square type solar-power-sail membrane. Fig. 16. Hexagon type solar-power-sail membrane. 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 210 s Table 14. Weight restriction Ion Engines Solar-Power-Sail Membrane Boom Boom Expansion Mechanism Primary buttery Reaction Control System Re-entry Capsule Weight budget result of scenario (B). Spin stabilized Attitude Control 300 kg 48.9 kg 12.7 kg 3.5 kg 14.1 kg 1.7 kg 13.1 kg 30 kg 3-axis stabilized Attitude Control Fuel for ion engines 73.0 kg 62.6 kg Fuel for RCS kg 3.63 kg Fuel for Touch-Down 15.6 kg Communication System Power Supply System Date Handling Unit Attitude Orbit Control System Electric System Thermal Protection System Instruments for Touch-Down 9.6 kg 8.3 kg 5.0 kg 6.1 kg 10.0 kg 5.0 kg 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 solar system. In order to satisfy this requirement, the area of the solar-power-sail membrane becomes 200 m 2 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 strategy 11) 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 estimated 10). 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 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. Pk_49

8 Trans. JSASS Aerospace Tech. Japan Vol. 12, No. ists29 (2014) - 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 spacecrafts 2,9,12). 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 m 2 solar power sail with the extension mast. By the analysis conducted in this paper, the feasibility of this mission is shown. References 1) Funase, R., Mori, O., Shirasawa, Y. and Yano, H: Trajectory Design and System Feasibility Analysis for Jovian Trojan Asteroid Exploration Mission Using Solar Power Sail, Transactions of the Japan Society for Aeronautical and Space Sciences, Aerospace Technology Japan, 12, ists29 (2014), pp. Pd_85-Pd_90. 2) Kawaguchi, J. et al.: The MUSES-C mission for the sample and return its technology development status and readiness, Acta Astronautica, 52 (2003), pp ) Tsuda, Y. et al.: Achievement of IKAROS - Japanese deep space solar sail demonstration mission, Acta Astronautica, 82 (2013) pp ) Lauretta, D. S. et al.: An Overview of the OSIRIS-REx Asteroid Sample Return Mission, 43rd Lunar and Planetary Science Conference, ) Barucci, M. A. et al.: MARCO POLO-R near earth asteroid sample return mission, Exp Astron, 33 (2012), pp ) Marov, M. Y. et al.: Phobos-Grunt: Russian sample return mission, Advances in Space Research, 33 (2004), pp ) Shirakawa, K. et al.: Accurate Landmark Tracking for Navigating Hayabusa Prior to Final Descent, Advances in the Astronautical Sciences, 124 (2006), pp ) Tsuda, Y. et al.: Modeling of Attitude Dynamics for Ikaros Solar Sail Demonstrator, Proceedings of 21 st AAS/AIAA SPACE FLIGHT MECHANICS MEETING, ) M Class Internal Review Report, Marco Polo, Technical & Programmatic Report, , Retrieved ) Chujo, T. et al.: Study about Difference in Attitude Control Method of Solar Power Sail and Application to Mission, 23 rd Workshop on JAXA Astrodynamics and Flight Mechanics, 2013, A ) Natori, M. C. et al.: Deployable Membrane Structures with Rolled-up Booms and Their Deployment Characteristics, 54 th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials (SDM) Conf., ) Yamakawa, H. et al.: Current status of the BepiColombo/MMO spacecraft design, Advances in Space Research, 33 (2004), pp Pk_50