川口 淳一郎

The article “Orbit insertion strategy of Hayabusa2’s rover with large release uncertainty around the asteroid Ryugu” written by Yusuke Oki, Kent Yoshikawa, Hiroshi Takeuchi et al., was originally published electronically on the publisher’s internet portal (currently SpringerLink) on 05 November 2020 without open access. After publication in Volume 4, Issue 4, page 309–329, the author(s) decided to opt for Open Choice and to make the article an open access publication. Therefore, the copyright of the article has been changed to © The Author(s) 2020 and the article is forthwith distributed under the terms of the Creative Commons Attribution 4.0 International License (http://creativecommons.org/licenses/by/4.0/), which permits use, duplication, adaptation, distribution and reproduction in any medium or format, as long as you give appropriate credit to the original author(s) and the source, provide a link to the Creative Commons license, and indicate if changes were made.

This paper describes the orbit design of the deployable payload Rover 2 of MINERVA-II, installed on the Hayabusa2 spacecraft. Because Rover 2 did not have surface exploration capabilities, the operation team decided to experiment with a new strategy for its deployment to the surface. The rover was ejected at a high altitude and made a semi-hard landing on the surface of the asteroid Ryugu after several orbits. Based on the orbital analysis around Ryugu, the expected collision speed was tolerable for the rover to function post-impact. Because the rover could not control its position, its motion was entirely governed by the initial conditions. Thus, the largest challenge was to insert the rover into a stable orbit (despite its large release uncertainty), and avoid its escape from Ryugu due to an environment strongly perturbed by solar radiation pressure and gravitational irregularities. This study investigates the solution space of the orbit around Ryugu and evaluates the orbit’s robustness by utilizing Monte Carlo simulations to determine the orbit insertion policy. Upon analyzing the flight data of the rover operation, we verified that the rover orbited Ryugu for more than one period and established the possibility of a novel method for estimating the gravity of an asteroid.

Copyright © 2020 by the International Astronautical Federation (IAF). All rights reserved. Transformable spacecraft under development is an innovative system that consists of several structural components, such as panels, connected together by internal force actuators. The spacecraft can change its structure drastically by driving installed actuators and achieve the following four features simultaneously. The first feature is "attitude change by internal force using non-holonomic characteristic of the system". It is possible to orient the spacecraft to an arbitrary direction by repeating the deployment of the panel in an appropriate order by the internal force actuator. The second feature is that "change of the structure enables the multiple functions by switching modes". Two telescopes will be installed for scientific missions utilizing the features of the transformable spacecraft and used to realize two different observation modes. One is a mode in which each telescope is oriented to different directions to perform wide-field observation (single telescope mode). The other is a mode in which two telescopes are pointed in the same direction. This mode enables the spacecraft to work as an interferometer (interferometer mode). The third feature is "orbit control and orbit keeping by controlling the solar radiation pressure on the spacecraft with the use of change of spacecraft structure". Since the spacecraft can change its structure by the internal force actuator, the orbit control and orbit keeping are achieved without fuel consumption. By utilizing this feature, the spacecraft will be injected into an artificial halo orbit around Sun-Earth Lagrangian point L2, and the technology demonstration of the transformable spacecraft and the observation mission will be performed in the orbit. The fourth feature is "passive cooling of observation equipment by use of panels as sunlight shield". In the observation mode, observation in the infrared region is performed and sufficient cooling is required. Appropriate arrangement of panels enables shielding of sunlight, and then the passive cooling of the observation equipment is realized. As a result, disturbance due to refrigerator is eliminated, which contributes to precise observation in addition to the contribution by non-holonomic attitude control without disturbance. This paper shows the analysis and experimental results for feasibility studies and conceptual designs of above four features. Furthermore, development status of the system and each subsystem to realize the spacecraft are introduced.

Copyright © 2020 by the International Astronautical Federation (IAF). All rights reserved. The removal of space debris from an orbit is one of the definitive solutions to the increasing Earth-orbiting satellites issue. To achieve active removal of space debris, accurate estimation of the state of motion of the target is crucial. In addition to that, a safe approach is preferable. Many strategies have been proposed for estimation of motion of a non-cooperative target, but there are still problems with regards to conducting precise, robust and cost-effective estimation in real-time. This study proposes a new strategy to solve this issue - a real-time and practical method based on optical navigation around small bodies. A computer simulation results show the efficacy of the proposed method and verifies that it can be a viable option for use in the capture of non-cooperative space debris targets.

This paper investigates coupled orbit-attitude dynamics around asteroids subject to solar radiation pressure and gravity irregularities. The solutions of sun-synchronous orbits with sun-tracking attitude motion are analytically derived, and their stability is evaluated by applying linearization and averaging. To validate the analytical solutions, numerical simulations are performed based on nonlinear coupled orbit-attitude equations of motion. In addition, the nonlinear stability of such coupled motion is analyzed using finite-time Lyapunov exponents. It is demonstrated that the sun-synchronous orbit-attitude coupled motions exhibit long-term stability under certain conditions, and thus, these motions are promising options for asteroid missions.

© 2018 IAA In the exploration of small bodies, the orbiting operation is very significant in terms of fuel consumption and scientific observation of small bodies. Although the force field around a small body is strongly perturbed, there is a family of stable orbits, called terminator orbits, that exist. However, although a terminator orbit is stable, it also has disadvantages. The orbital plane of a terminator orbit must always face the Sun and lacks flexibility in orbit design. Moreover, since the orbital plane lies in the night side because of the solar radiation pressure (SRP) shifting the equilibrium point, optical observation of the small body is extremely restricted. Therefore, the present study focuses on the perturbed terminator orbit, i.e., the quasi-terminator orbit (QTO), which is a stable orbit that does not suffer from impact with the surface or escape. The present study attempts to reveal the solution space and the usability of the QTO. We succeeded in analytically deriving the existence range of the long-term QTO and verified that the numerical and analytical solutions coincide well. As a result, the calculation time required for solving the existence range of the long-term stable QTO can be greatly shortened, and it becomes possible to quickly apply our findings to other missions by accounting for the gravity field and SRP of the mission-specific small body.

© 2018 Elsevier Inc. Seismic shaking of small bodies has an important role in revealing information about internal structure and contributing to our understanding of microgravity geology. Since many small asteroids are likely to be rubble piles, it is important to understand their dynamics, which is largely different from that of monolithic asteroids. We introduce a new normal mode analysis method as an approach to seismic study based on a discrete element method (DEM), where a rubble pile is modeled as a group of elastic spheres bound together with gravitational force, with elastic repulsion forces following Hertzian contact theory and Mindlin's theory. Normal mode analysis is formulated by differentiating the interaction between particles around an equilibrium state. Our results show that normal modes and eigenfrequencies are independent of the size of the particles if their physical properties and geometry are identical. Furthermore, we apply a scaling law, which shows that the shape of normal modes does not change even if the size of a rubble pile is enlarged or contracted, but the eigenfrequency varies in inverse proportion with the scale to the power of two-thirds. We also compare normal modes between a rubble pile and a monolith and show that the shapes of normal modes are similar to each other, but the eigenfrequency is smaller in the rubble pile. Finally, we describe dynamic simulations to compare results from nonlinear DEM and results from linearization of the normal mode. It is found that normal mode analysis gives a good approximation for the frequency distribution when the kinetic energy is sufficiently small, and that the presence of tangential force between particles suppresses vibration motion.

Copyright © 2019 by the International Astronautical Federation (IAF). All rights reserved. In this study, we modify multi-particle method, a calculation method for predicting membrane behaviour, for evaluating collision between membranes and the change in membrane compression and bending stiffness due to devices mounted on membrane. This modification is verified its validity by the consistency with the results of the experiments conducted in the previous study. Then, we investigate the changes in the deployment behaviour when the temperature and thickness of solar cells change.

Copyright © 2019 by the International Astronautical Federation (IAF). All rights reserved. The membrane dynamics of spinning solar sails have a special relevance when considering attitude control of the spacecraft. So as to model the system accurately, the bending stiffness of the membrane has been included in the numerical approach, as it is believed to have strong effects on the behavior of the sail. First, this study shows that the influence of the bending moment on the attitude of the sail during its spin axis reorientation should not be neglected. Given the difficulty of measuring the actual bending stiffness of the membrane, finding a control system capable to perform the same regardless of its value is necessary. Therefore, this paper presents a new control system and its corresponding logic to lower the influence of the bending parameter on this attitude maneuver performance. Finally, a frequency analysis on the vibrations arising in the membrane when considering different bending stiffness values is done. This last analysis shows the shift in the natural frequencies obtained, remarking the importance of the bending stiffness when considering the dynamics of a mast-free sail.

Attitude dynamics and control of spinning solar sails are investigated considering the flexibility of sail membranes. Attitude maneuver of solar sails is, in many cases, performed using thrusters. In most studies, the attitude motion is analyzed assuming that the spacecraft is a rigid disk. However, the sail membrane deforms during attitude maneuver due to flexibility. This may cause coupled vibration between the spacecraft main body and sail membrane. This study presents an analysis model of spinning solar sail attitude dynamics considering sail deformation, based on modal analysis.

Copyright © 2018 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. Liquid crystal devices (LCDs) are proposedas fuel-free attitude control devices for spacecraft. The LCDs consist of multiple thin layers, including polymer-dispersed liquid crystal (PDLC) and a base film that has a microstructure with aluminum deposited onto it. When a voltage isapplied to anLCD, the crystalsinthe PDLC align with the electric field, and incident light passes through the PDLC layer and is reflected by the aluminum coating, whereas it is dispersed by randomly arranged crystals when no voltage is applied. This means that arrangements of LCDs can control both the magnitude and direction of reflection electrically, and they can achieve three-degree-of-freedom attitude control using solar radiation pressure. Because light diffracted by the layers causes optical interference, the PDLC has a polarization effect, and the microstructure shape has manufacturing errors, reflection from LCDs is a complex phenomenon and it must be understood to realize the desired performance." The first sentence all means the reason for the second sentence.

© 2017 The Author(s). Due to its important role in the sorting of particles on microgravity bodies by size, Brazil nut effect (BNE) is a major subject of study for understanding the evolution of planetesimals. Recent studies have revealed that the mechanism for the BNE on microgravity bodies is the percolation of particles or void-filling, rather than granular convection. This study also considers the mechanism for the BNE under 'less-convective' conditions and introduces three categories of behaviour for particles that mainly depend on the dimensionless acceleration of vibration Γ (ratio of maximum acceleration to gravitational acceleration), using a simplified analytical model. The conditions for Γ proposed by the model for each category are verified by both numerical simulations and laboratory experiments. 'Less-convective' conditions are realized by reducing the friction force between particles and the wall. We found three distinct behaviours of the particles when Γ > 1: the (i) 'slow BNE', (ii) 'fast BNE', and (iii) 'fluid motion' (the reverse BNE may be induced), and the thresholds for Γ correspond well with those proposed by the simplemodel. We also applied this categorization to low-gravity environments and found that the categorization scales with gravity level. These results imply that laboratory experiments can provide knowledge of granular mobility on the surface of microgravity bodies.

This paper proposes the control method which achieves fuel-free station-keeping around Sun-Earth L2 point (SEL2). This station-keeping is driven with only solar radiation pressure (SRP), and therefore the orbit is maintained without fuel consumption by controlling attitude of a spacecraft toward the sun. Furthermore, this attitude control is also achieved without fuel by using the technique of non-holonomic turn. The target orbit is the artificial orbit around SEL2, the size of which is much smaller than typical halo orbits, and thus provides more stationary thermal condition to spacecrafts.

As an innovative spacecraft system, a transformable spacecraft is being studied by the authors. One of the most innovative points of the spacecraft is that it is capable of non-holonomic turns for attitude maneuver. By the combination of the turns, the spacecraft can change its attitude largely without using fuel or external forces/torques. In this paper, we first show that it is possible to achieve any kinds of attitude with slight error in principle using a simple model, although it may not be optimal. After that, a method for optimal motion planning using genetic algorithm is proposed.

This study investigates coupled orbit-attitude dynamics around asteroids subject to solar radiation pressure and gravity irregularities. The solutions of Sun synchronous orbits with Sun-tracking attitude motion are analytically derived by applying linearization and averaging. To verify the validity of the analytical solutions, numerical simulations are performed based on non-linear coupled orbit attitude equations of motion. In addition, the stability of such coupled motion is analyzed using finite-time Lyapunov exponents. It is demonstrated that the Sun synchronous orbit-attitude coupled motions exhibit long-term stability under certain conditions, and thus, these motions are useful and feasible options for asteroid missions.

© 2018 Univelt Inc. All rights reserved. This study investigates coupled orbit-attitude dynamics around asteroids subject to solar radiation pressure and gravity irregularities. The solutions of Sun-synchronous orbits with Sun-tracking attitude motion are analytically derived by applying linearization and averaging. To verify the validity of the analytical solutions, numerical simulations are performed based on non-linear coupled orbit-attitude equations of motion. In addition, the stability of such coupled motion is analyzed using finite-time Lyapunov exponents. It is demonstrated that the Sun-synchronous orbit-attitude coupled motions exhibit long-term stability under certain conditions, and thus, these motions are useful and feasible options for asteroid missions.

Copyright © 2018 by the International Astronautical Federation (IAF). All rights reserved. The solar power sail mission OKEANOS, planned to be launched in 2026, aims for sample return from a Jovian Trojan asteroid. In this mission, propagation delay (100min) is longer than that of Hayabusa2 (60min). Moreover, it is difficult to get sufficient information to make a success of the image-based navigation used in the previous mission. Another GNC method is required under the condition. This paper proposes an image-based on-board GNC method.

© 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. As fuel-free attitude control devices for spacecraft, we introduce Liquid Crystal Devices (LCDs). LCDs consist of multiple thin layers (∼ 150 µm thick in total) including Polymer Dispersed Liquid Crystal (PDLC) and a base film that has microstructure with aluminum deposited on it. By applying voltage, reflectivity of the surface of LCDs changes as the PDLC is directed to the electric field, and applying Solar Radiation Pressure (SRP) changes accordingly. In addition, the microstructure changes the angle of reflection only when voltage is applied such that incident light is reflected on the aluminum on it. These mean that LCDs can control both the magnitude and direction of reflection electrically, and three-degree-of-freedom attitude control can be realized by multiple LCDs mounted on a single plane controlling SRP. As light is diffracted on the microstructure and optical interference occurs, and PDLC has polarization effect, reflection on LCDs is a complex phenomenon and it must be understood for the design to realize desired performance. In this paper, we introduce basic reflection principles and a manufacturing method, and confirm them by making prototypes and measuring the distribution of luminous intensity. The performance of prototypes of LCDs is estimated based on the measurement. The results indicate validity of the reflection principles and the usefulness of LCDs considering the future view as attitude control devices.

© 2018 Univelt Inc. All rights reserved. Attitude dynamics and control of spinning solar sails are investigated considering the flexibility of sail membranes. Attitude maneuver of solar sails is, in many cases, performed using thrusters. In most studies, the attitude motion is analyzed assuming that the spacecraft is a rigid disk. However, the sail membrane deforms during attitude maneuver due to flexibility. This may cause coupled vibration between the spacecraft main body and sail membrane. This study presents an analysis model of spinning solar sail attitude dynamics considering sail deformation, based on modal analysis.

Copyright © 2018 by the International Astronautical Federation (IAF). All rights reserved. As an innovative spacecraft system, a transformable spacecraft is proposed, which consists of multiple bodies connected with each other. They are equipped with actuators that move them relatively within a certain range of angle. The shape of the bodies is arbitrary, and the simplest is panel shape, for example. Supposing a transformable spacecraft consisting of a number of panels, they can be folded to be compact as a whole, unfolded to configure a large plane, and recomposed to configure various kinds of three-dimensional shape. The system enables a single spacecraft to have multiple functions by transforming the shape. A distinctive characteristic of a transformable spacecraft is its capability of performing nonholonomic attitude control. By transforming to another shape and transforming back to the original shape in a different path, the attitude is changed even if the shape is unaltered. This nonholonomic control is realized only by internal torque, and does not require any fuel consumption. We introduce the concept of a transformable spacecraft and its applications to missions, utilizing the nonholonomic control. For example, combining the function of variable-shape structure with the nonholonomic control, a multifunction space telescope can be realized orbiting around the Sun-Earth Lagrange point.

© 2017 IAA The motion of a spacecraft in proximity to a small body is significantly perturbed due to its irregular gravity field and solar radiation pressure. In such a strongly perturbed environment, the coupling effect of the orbital and attitude motions exerts a large influence that cannot be neglected. However, natural orbit-attitude coupled dynamics around small bodies that are stationary in both orbital and attitude motions have yet to be observed. The present study therefore investigates natural coupled motion that involves both a Sun-synchronous orbit and Sun-tracking attitude motion. This orbit-attitude coupled motion enables a spacecraft to maintain its orbital geometry and attitude state with respect to the Sun without requiring active control. Therefore, the proposed method can reduce the use of an orbit and attitude control system. This paper first presents analytical conditions to achieve Sun-synchronous orbits and Sun-tracking attitude motion. These analytical solutions are then numerically propagated based on non-linear coupled orbit-attitude equations of motion. Consequently, the possibility of implementing Sun-synchronous orbits with Sun-tracking attitude motion is demonstrated.

© Copyright 2016 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. Delta-V assisted periodic orbits are introduced as a new type of periodic orbit around small bodies subject to strong solar radiation pressure. Delta-V assisted periodic orbits are made periodic by introducing a small deterministic delta-V within each period. This type of orbit has a simpler shape and provides higher flexibility than other periodic orbits and therefore enables missions with higher scientific value. The general theory of delta-V assisted periodic orbits is described, including the orbit design methodology, solution-space analysis, and stability analysis. This paper clarifies that delta-V assisted periodic orbits are useful and feasible options for small-body missions and exhibit unique dynamic characteristics.

Copyright © (2017) by International Astronautical Federation. All rights reserved. In the exploration mission of small bodies orbiting operation is very significant in terms of fuel saving and scientific observation of small bodies. Although the force field around a small body is strongly perturbed, there is a family of stable orbits which is called "Terminator Orbit". While the terminator orbit is stable, it also has disadvantages. The terminator orbit's orbital plane must always face the sun and lack flexibility in orbit design. Moreover, since the orbital plane lies in night side, optical observation of the small body is extremely restricted. Therefore, this study extends the terminator orbit concept to a group of stable orbits which includes the terminator orbit and does not suffer from impact with surface or escaping. This extended terminator is called "Quasi-Stable Terminator Orbit (QSTO)" in this study. This work aims to reveal the solution space and the usability of QSTO. Thus, we succeeded in analytically deriving the existence range of the long-term QSTO and verified that numerical and analytical solutions coincide well. As a result, the calculation time required for solving the existence range of the long-term stable QSTO can be greatly shortened, and it became possible to obtain the range when changing the strength of the small body's gravity and the SRP in accordance with various missions in a short time.

Copyright © (2017) by International Astronautical Federation. All rights reserved. The authors are currently developing a novel attitude control device for solar sails. The device is called Advanced Reflectivity-Control-Device (A-RCD) which can change the optical properties between diffuse reflection and obliquereflection. This device is mounted on the solar sail's membrane and can control the torque perpendicular to the membrane. The A-RCD is composed of microstructured film and polymer dispersed liquid crystal (PDLC), and these structures cause the strong diffraction when the light propagates in the device. Simulating such light propagation is indispensable to evaluate the performance by using computer calculations, which is inevitable if one tries to achieve the optimum design via parametric studies. Besides, the ability to simulate the light propagation in the A-RCD will give designers new insights into the A-RCD design. In this context, we will introduce the simulation method for the A-RCD in this paper.

The total electric power consumption of heaters in a spacecraft has to be controlled in order not to be in short. In conventional system, a server communicates with each heater and controls the total power. In this system, the time to communicate with heaters depends on the number of heaters and the time to develop a spacecraft is long. This paper proposes the autonomous distributed system which does not need a server. This system enables the control where the time to communicate does not depend on the number of heaters and the time to develop a spacecraft is short. This paper demonstrates that it is possible to control power consumption and temperature of heaters by both a simulation and an experiment. In the experiment, ZigBee, which is a wireless unit, is used to broadcast from the transmitter to the heaters.

Delta-V assisted periodic orbits (DVAPOs) are introduced as a new type of periodic orbit around small bodies subject to strong solar radiation pressure. DVAPOs are made periodic by introducing a small deterministic delta-V within each period. This type of orbit has a simpler shape and provides higher flexibility than other periodic orbits and therefore enables missions with higher scientific value. The general theory of DVAPOs is described, including the orbit design methodology, solution-space analysis, and stability analysis. This paper clarifies that DVAPOs are a useful and feasible option for small body missions and exhibit unique dynamic characteristics.

<p>The purpose of this work was to develop escape trajectories from distant retrograde orbits in the Sun&ndash;Earth circular restricted three-body problem. Previous studies have suggested installing a space port outside the gravitational sphere of influence of Earth for future deep space exploration. This paper proposes the placement of a space port on a Sun&ndash;Earth distant retrograde orbit (SE-DRO), which is stable over a long period. The characteristics and fuel consumption efficiency of escape trajectories from the space port on the SE-DRO using one or two impulses &Delta;V were investigated and compared with those obtained in previous studies that considered a space port located in the vicinity of the Sun&ndash;Earth L2 Lagrangian point. The connection of the escape trajectories from an SE-DRO with Earth gravity-assist (EGA) trajectories was then investigated based on the results of the analysis with the impulse &Delta;V. Finally, a series of trajectories from SE-DRO to EGA were shown to have a high efficiency from the perspective of both &Delta;V and the flight time.</p>

Copyright © 2016 by the International Astronautical Federation (IAF). All rights reserved. Rendezvous missions to small bodies, such as asteroids and comets, have been of interest in recent years. The motion of a spacecraft in the proximity of a small body is strongly perturbed mainly because of the irregular shape of the small body and the solar radiation pressure (SRP). In order to understand this unique environment, orbital dynamics and attitude dynamics around small bodies have been analyzed in many previous studies. In those analyses, the orbital motion and the attitude motion of a spacecraft were analyzed separately, and dynamic interaction between these motions has not been revealed. Moreover, the past research on attitude dynamics focused on the attitude of a spacecraft with respect to a small body, although the attitude with respect to the Sun is often the information of interest for solar power generation and thermal design. This research, therefore, investigates natural motion in orbit-attitude coupled system which involves both a Sun-synchronous orbit and Sun-tracking attitude motion. Sun-synchronous orbits, which are also called as frozen orbits, are periodic orbits in the Sun-centered rotating frame. On the other hand, Sun-tracking attitude motion is the motion that a spacecraft keeps tracking the Sun with small oscillation. This orbit-attitude motion enables spacecraft to maintain its orbital geometry and attitude with respect to the Sun, ideally with no active control. Thus, the proposed method can reduce the usage of orbit and attitude control system, such as thrusters and reaction wheels, which results in reducing the weight and prolonging the mission life time of a spacecraft. This paper consists of two parts. First, the orbital motion and the attitude motion are analyzed independently, which are modeled as Lagrange planetary equations and linearized Euler equations, respectively. As a result, analytical solutions of Sun-synchronous orbits and Sun-tracking attitude motion are successfully obtained. Next, orbital motion and attitude motion are propagated simultaneously by numerical integration based on orbit-attitude coupled equations of motion. It is then demonstrated that the analytical solutions can be valid approximation of the numerical solution. This research clarifies that Sun-synchronous orbits with Sun-tracking attitude motion are feasible for small-body missions and exhibit unique dynamic characteristics.

Delta-V assisted periodic orbits (DVAPOs) are introduced as a new type of periodic orbit around small bodies subject to strong solar radiation pressure. DVAPOs are made periodic by introducing a small deterministic delta-V within each period. This type of orbit has a simpler shape and provides higher flexibility than other periodic orbits and therefore enables missions with higher scientific value. The general theory of DVAPOs is described, including the orbit design methodology, solution-space analysis, and stability analysis. This paper clarifies that DVAPOs are a useful and feasible option for smallbody missions and exhibit unique dynamic characteristics.

© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. An asteroid explorer (Hayabusa, Hayabusa2, …) takes many images of the target asteroid, and the images are used for the navigation around the asteroid. Some of representative optical navigation methods detect feature points of the asteroid in the images such as rocks and craters. Hayabusa succeeded to touch down on the asteroid by use of the optical navigation method called GCP-NAV. However, these kinds of navigation methods need to construct the database of the feature points beforehand by the precise measurement of the asteroid ground. In this research, we constitute virtual feature points from the topographical information captured in the images, and propose an efficient, autonomous optical navigation method without large-scale measurement in advance.

Attitude control methods of a solar sail which maximize the effectiveness of ΔVEGA (ΔV Earth Gravity Assist) are proposed. An orbit of a solar sail is controlled by changing its attitude, and there are two types of method to control an orbit of a spinning solar sail. One is to control its attitude directly, and the other is to control its attitude indirectly using an attitude drift motion due to the solar radiation pressure by controlling its spin rate. These methods are applied to an orbit of ΔVEGA, and their efficiency is quantitatively evaluated by solving optimal control problems to maximize the relative velocity with respect to the Earth.

Hayabusa was the first asteroid sample return mission. It was launched in May 2003, and arrived at the target asteroid (25143) Itokawa in September 2005. The mission enabled us to see close up a very tiny asteroid in detail for the first time. Hayabusa observed Itokawa with its scientific instruments, and attempted to collect surface material. The mission experienced several serious problems, but successfully returned to Earth in June 2010. After retrieving the capsule, we found thousands of small grains that had been captured from the asteroid. We studied Itokawa in detail with both the remote sensing data and the returned samples, which revealed a great deal of new information to shed light on its origin. In this chapter, we review the Hayabusa mission and summarize the scientific results.