研究者業績
基本情報
- 所属
- 国立研究開発法人宇宙航空研究開発機構 宇宙科学研究所 教授東京大学 大学院工学系研究科 航空宇宙工学専攻 准教授
- 学位
- 博士（工学）(2007年3月 東京大学)
- 研究者番号
- 70509819
- J-GLOBAL ID
- 200901051354637504
- researchmap会員ID
- 5000090886
研究キーワード
11研究分野
1受賞
4論文
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Journal of Guidance, Control, and Dynamics 47(9) 1905-1919 2024年9月In autonomous space systems, the reliability of navigation systems is essential. Observability in autonomous orbit determination techniques depends on the spacecraft’s orbital motion, making the design of autonomous navigation systems and orbital maneuvers a coupled process. This study develops a stable and efficient algorithm based on differential dynamic programming to design maneuver sequences that improve navigation performance. Our approach incorporates the Fisher information matrix into a cost function to quantify state observability and facilitates its convergence using a semi-analytic gradient and Hessian derived under impulsive maneuvers. Two numerical examples show the validity and effectiveness of our algorithm. The results indicate the stability and efficiency in determining maneuver sequences and the improvement of state estimation accuracy along an optimized trajectory. It is also applicable to other observability-aware optimal control problems because the algorithm is independent of specific systems.
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Space Science Reviews 220(1) 2024年1月24日Abstract Here we describe the novel, multi-point Comet Interceptor mission. It is dedicated to the exploration of a little-processed long-period comet, possibly entering the inner Solar System for the first time, or to encounter an interstellar object originating at another star. The objectives of the mission are to address the following questions: What are the surface composition, shape, morphology, and structure of the target object? What is the composition of the gas and dust in the coma, its connection to the nucleus, and the nature of its interaction with the solar wind? The mission was proposed to the European Space Agency in 2018, and formally adopted by the agency in June 2022, for launch in 2029 together with the Ariel mission. Comet Interceptor will take advantage of the opportunity presented by ESA’s F-Class call for fast, flexible, low-cost missions to which it was proposed. The call required a launch to a halo orbit around the Sun-Earth L2 point. The mission can take advantage of this placement to wait for the discovery of a suitable comet reachable with its minimum $\varDelta $V capability of $600\text{ ms}^{-1}$. Comet Interceptor will be unique in encountering and studying, at a nominal closest approach distance of 1000 km, a comet that represents a near-pristine sample of material from the formation of the Solar System. It will also add a capability that no previous cometary mission has had, which is to deploy two sub-probes – B1, provided by the Japanese space agency, JAXA, and B2 – that will follow different trajectories through the coma. While the main probe passes at a nominal 1000 km distance, probes B1 and B2 will follow different chords through the coma at distances of 850 km and 400 km, respectively. The result will be unique, simultaneous, spatially resolved information of the 3-dimensional properties of the target comet and its interaction with the space environment. We present the mission’s science background leading to these objectives, as well as an overview of the scientific instruments, mission design, and schedule.
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 67(5) 274-284 2024年
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Journal of Astronomical Telescopes, Instruments, and Systems 9(03) 2023年9月12日
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Journal of Evolving Space Activities 1 n/a 2023年The low reliability of nano-satellites has become a problem, with the cause often being defects in the design and manufacturing process. It is required to take sufficient measures to account for those defects through ground tests to reduce on-orbit failures. The model-based fault diagnosis method can prevent oversight of fault candidates and compensate for a lack of knowledge in human-based diagnosis. Since the developing time of nano-satellites is limited, fault diagnosis needs to be carried out efficiently also while being careful not to cause secondary faults. This paper presents a novel method that can find the appropriate command to verify candidates of fault location in a satellite. The proposed method can consider the side effect of a command and the efficiency of narrowing them down by using the information transmission path model. In the case study, the proposed method is applied to an earth-orbiting 6U CubeSat, and this case study reveals that 33% of the failures that occur in ground tests can apply to the proposed method. In addition, the proposed method can identify the fault location of some of these failures, although there are several limitations to implementing the proposed method in the actual satellite development.
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2022 IEEE Aerospace Conference (AERO) 152 1-11 2022年3月5日
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The Journal of the Astronautical Sciences 2022年2月4日
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SPACE TELESCOPES AND INSTRUMENTATION 2022: ULTRAVIOLET TO GAMMA RAY 12181 2022年GEO-X (GEOspace X-ray imager) is a small satellite mission aiming at visualization of the Earth's magnetosphere by X-rays and revealing dynamical couplings between solar wind and magnetosphere. In-situ spacecraft have revealed various phenomena in the magnetosphere. In recent years, X-ray astronomy satellite observations discovered soft X-ray emission originated from the magnetosphere. We therefore develop GEO-X by integrating innovative technologies of the wide FOV X-ray instrument and the microsatellite technology for deep space exploration. GEO-X is a 50 kg class microsatellite carrying a novel compact X-ray imaging spectrometer payload. The microsatellite having a large delta v (>700 m/s) to increase an altitude at 40-60 R-E from relatively low-altitude (e.g., Geo Transfer Orbit) piggyback launch is necessary. We thus combine a 18U Cubesat with the hybrid kick motor composed of liquid N2O and polyethylene. We also develop a wide FOV (5x5 deg) and a good spatial resolution (10 arcmin) X-ray (0.3-2 keV) imager. We utilize a micromachined X-ray telescope, and a CMOS detector system with an optical blocking filter. We aim to launch the satellite around the 25th solar maximum.
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JOURNAL OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 69(5) 179-186 2021年
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Journal of Guidance, Control, and Dynamics 43(4) 645-655 2020年3月 査読有り最終著者
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Acta Astronautica 171 208-214 2020年2月 査読有り
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Proceedings of the International Astronautical Congress, IAC 2020- 2020年Comet Interceptor, to be launched in 2028, is a recently-selected ESA/JAXA mission aimed to perform the first fly-by of a pristine long-period comet. As part of the mission's Phases A/B (feasibility analysis and preliminary design), this work preliminarily analyzes the attitude performance of JAXA's contribution to the mission, SC B1: one of the two small-spacecraft piggybacked along with ESA's main spacecraft. This study primarily focuses on the characterization of the dust environment, modelization of dust particle impacts, and analysis of attitude performance through the dust environment. Monte Carlo simulations are implemented to assess the performance of the attitude control system within a highly-active dust environment, for various fly-by altitudes and a worst-case 70-km/s fly-by speed. Results evidence the need for a wide-angle camera, and conclude that image acquisition shall tolerate angular velocities of several degrees per second. Solutions for improved attitude performance are also evaluated (i.e., upgrading of reaction wheels and angular momentum pre-loading), and a larger reaction wheels aligned with the camera line of sight is shown to provide largest performance improvements. Such strategies, nevertheless, still result in pointing errors and angular velocities that are only compatible with a short-exposure, wide-angle science camera.
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Proceedings of the International Astronautical Congress, IAC 2020-October 2020年Micro/nano-satellite technology development first started as either an educational or research tool primarily at universities and has spread rapidly across the world and found many practical applications. Recent technology advancement has enabled micro/nano satellites to become one of the platforms for deep space science and exploration missions. In order to provide an opportunity for students and researchers worldwide to propose a mission of deep space science and exploration using a micro/nano satellite, the 7th Mission Idea Contest (MIC) will focus on these missions. This is because the technological field of LEO satellites is already well-established, and so we consider that creation of a deep space mission will give the young generation more motivation towards the “Frontier.” MIC has provided aerospace engineers, college students, consultants, scientists, and anybody interested in space with opportunities to present their creative ideas and gain attention internationally. The 7th Mission Idea Contest (MIC7) is to open a door to new opportunities for proposing a deep space mission using a micro/nano satellite. There are several examples and emerging opportunities for micro/nano satellites in these types of missions. In this paper, we briefly introduce the past Mission Idea Contests and present examples of deep space missions using micro/nano satellites for MIC7, and follow this with a discussion on why it is important for UNISEC-Global to organize a “Mission Idea Contest for Deep Space Science and Exploration .
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Flight Model Development and Ground Demonstration of Water Resistojet Propulsion System for CubeSatsTRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 63(4) 141-150 2020年The University of Tokyo has proposed a water resistojet thruster with a high certainty of liquid-vapor separation and low power consumption. In this propulsion system, liquid water is periodically vaporized in a pulsating manner to generate thrust. A vaporization chamber with a labyrinth-shaped flow path catches droplets using their surface tension to separate the liquid and vapor, and the droplets vaporize under normal temperature to reduce the input power by reusing the heat from the surrounding components. In this study, we designed and fabricated a flight model of the proposed propulsion system for 6U CubeSat and evaluated the performance of this propulsion system, including the control method. The results confirm the concept of the proposed liquid-vapor separation method and its low power consumption. Moreover, we revealed the relationships between the vaporizing duty cycle, input power, and thrust.
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日本航空宇宙学会論文集 68(2) 89-95 2020年 査読有り最終著者<p>A Fault Detection, Isolation, and Recovery (FDIR) algorithm for attitude control systems is a key technology to increasing the reliability and survivability of spacecraft. Micro/nano interplanetary spacecraft, which are rapidly evolving in recent years, also require robust FDIR algorithms. However, the implementation of FDIR algorithms to these micro/nano spacecraft is difficult because of the limitations of their resources (power, mass, cost, and so on). This paper shows a strategy of how to construct a FDIR algorithm in the limited resources, taking examples from micro deep space probe PROCYON. The strategy focuses on function redundancies and multi-layer FDIR. These ideas are integrated to suit the situation of micro/nano interplanetary spacecraft and demonstrated in orbit by the PROCYON mission. The in-orbit results are discussed in detail to emphasize the effectiveness of the FDIR algorithm. </p>
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JOURNAL OF THE ASTRONAUTICAL SCIENCES 67(3) 950-976 2020年1月 査読有り最終著者This paper presents the trajectory design for EQUilibriUm Lunar-Earth point 6U Spacecraft (EQUULEUS), which aims to demonstrate orbit control capability of CubeSats in the cislunar space. The mission plans to observe the far side of the Moon from an Earth-Moon L2 (EML2) libration point orbit. The EQUULEUS trajectory design needs to react to uncertainties of mission design parameters such as the launch conditions, errors, and thrust levels. The main challenge is to quickly design science orbits at EML2 and low-energy transfers from the post-deployment trajectory to the science orbits within the CubeSat's limited propulsion capabilities. To overcome this challenge, we develop a systematic trajectory design approach that 1) designs over 13,000 EML2 quasi-halo orbits in a full-ephemeris model with a statistical stationkeeping cost evaluation, and 2) identifies families of low-energy transfers to the science orbits using lunar flybys and solar perturbations. The approach is successfully applied for the trajectory design of EQUULEUS.
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JOURNAL OF SPACECRAFT AND ROCKETS 56(5) 1400-1408 2019年9月 査読有りThis study proposes a micropropulsion system unifying ion thrusters and resistojet thrusters and assessing that propulsive capability. The remarkable features of the system are the usage of water propellant and unification of the two types of thrusters by the single propellant. Water has been regarded as an attractive propellant in the view points of safety, availability, handling ability, low molecular mass, and future procurement in space. A multimode propulsion system is an attractive solution for the increasing demand for nano-/microsatellite missions. The proposal is to use microwave discharge water ion thrusters, tolerant for oxidization by water, and low-temperature water resistojet thrusters, enabling reuse of the waste heat. As a result of the assessment, it was expected that the propulsion system would have 3U size (10 x 10 x 30 cm(3)) and 3.70 kg mass, which realize in total a 6U and 10 kg satellite with 3U and 6 kg satellite bus system. The ion thruster would provide the maximum Delta V of 630 m/s by 47 W system power and the resistojet thruster would have 3.80 mN thrust and 72 s specific impulse by 19.4 W. Additionally, reuse of the waste heat from ion-thruster power supplies would enable the simultaneous operations of the two thrusters even at 50 W, which is almost the same power as the single ion thruster operation.
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Advances in the Astronautical Sciences 168 281-300 2019年Recent low-thrust space missions have highlighted the importance of designing trajectories that are robust against uncertainties. In its complete form, this process is formulated as a nonlinear constrained stochastic optimal control problem. This problem is among the most complex in control theory, and no practically applicable method to low-thrust trajectory optimization problems has been proposed to date. This paper presents a new algorithm to solve stochastic optimal control problems with nonlinear systems and constraints. The proposed algorithm uses the unscented transform to convert a stochastic optimal control problem into a deterministic problem, which is then solved by trajectory optimization methods such as differential dynamic programming. Two numerical examples, one of which applies the proposed method to low-thrust trajectory design, illustrate that it automatically introduces margins that improve robustness. Finally, Monte Carlo simulations are used to evaluate the robustness and optimality of the solution.
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Advances in the Astronautical Sciences 168 4015-4026 2019年Autonomous orbit determination method using active maneuvers and inter-satellite ranging between multiple spacecraft is applicable to general dynamics situations, but large uncertainty of information about maneuvers results in inaccurate orbit estimation. This paper proposes an accurate and robust estimation method using sequential filter, RTS smoother, and EM algorithm. Proposed method estimates not only states but also maneuver results. Results from simulations of Mars-Phobos system show that the proposed method improve orbits determination accuracy.
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Advances in the Astronautical Sciences 168 301-318 2019年Temporarily-captured orbiters (TCOs) are a new population of asteroids that are temporarily gravitationally bound around the Earth-Moon system. Because of its small geocentric distance and energy, short-term exploration with small Δv is expected possible. This study aims to construct low-energy transfers to 2006 RH120, one of the TCOs, from low-Earth orbit using an analogy with the Earth-Moon low-energy transfers. The initial guess was sought by back-propagating perturbed 2006 RH120’s trajectory, then it was optimized through the direct multiple shooting method. The result shows that various transfers are possible with rendezvous Δv below 100 m/s, and they form diverse family structures.
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Advances in the Astronautical Sciences 168 3511-3526 2019年Navigation and mutual communication are indispensable functions to manage a spacecraft formation flight. To achieve these functions, this paper presents a navigation and communication network architecture based on ultra-wide band (UWB) communication devices, which have been widely used for many applications like indoor-localization and navigation. By measuring distances between pairs of spacecraft, it is possible to estimate the relative positions and velocities of spacecraft in the formation. When it comes to a large-scale formation, a limited number of distances between two spacecraft can be measured by UWB devices at any one time due to physical limitations, such as available communications bandwidth. In this sense, it is necessary to select measured pairs of spacecraft to efficiently achieve high estimation accuracy. We propose an algorithm for the optimal selection of measure pairs of spacecraft to increase the estimation accuracy of the positions of spacecraft in the formation. The performance of the proposed method is evaluated in the simulation.
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Aerospace Technology Japan 17(1) 43-50 2019年1月 査読有り
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Aerospace Technology Japan 17(3) 315-320 2019年1月 査読有り
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Advances in the Astronautical Sciences 168 1331-1343 2019年 最終著者© 2019, Univelt Inc. All rights reserved. In order to deeply understand orbital disturbances, the flight data of the PROCYON, which is the 50kg-class interplanetary micro-spacecraft was analyzed. In the telemetry data, we found two unexpected behaviors of angular momentum in Z-axis as compared with the accurate solar radiation pressure model. In order to clarify the causes of the angular momentum anomalies, several small disturbances like thermal radiation pressure, deformation of the structure, and interplanetary magnetic field effect, which are usually ignored are discussed in this study. The thermal radiation and deformation of the structure can explain the over-large Z-axis anomaly. The interplanetary magnetic field effect is correlated with the sudden change of Z-axis torque anomaly in several cases, but the cause of the anomaly is not completely revealed yet.
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Acta Astronautica 152 299-309 2018年11月 査読有り© 2018 IAA This study proposes a solar sailing method for angular momentum control of the interplanetary micro-spacecraft PROCYON (PRoximate Object Close flYby with Optical Navigation). The method presents a simple and facile practical application of control during deep space missions. The developed method is designed to prevent angular momentum saturation in that it controls the direction of the angular momentum by using solar radiation pressure (SRP). The SRP distribution of the spacecraft is modeled as a flat and optically homogeneous plate at a shallow sun angle. The method is obtained by only selecting a single inertially fixed attitude with a bias-momentum state. The results of the numerical analysis indicate that PROCYON's angular momentum is effectively controlled in the desired directions, enabling the spacecraft to survive for at least one month without momentum-desaturation operations by the reaction control system and for two years with very limited fuel usage of less than 10 g. The flight data of PROCYON also indicate that the modeling error of PROCYON's SRP distribution is sufficiently small at a small sun angle (<10°) of the order of 10−9 Nm in terms of its standard deviation and enables the direction of the angular momentum around the target to be maintained.
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JOURNAL OF ASTRONOMICAL TELESCOPES INSTRUMENTS AND SYSTEMS 4(4) 2018年10月 査読有りToward an era of x-ray astronomy, next-generation x-ray optics are indispensable. To meet a demand for telescopes lighter than the foil optics but with a better angular resolution <1 arcmin, we are developing micropore x-ray optics based on micromaching technologies. Using sidewalls of micropores through a thin silicon wafer, this type can be the lightest x-ray telescope ever achieved. Two Japanese missions, ORBIS and GEO-X, will carry this telescope. ORBIS is a small x-ray astronomy mission to monitor supermassive blackholes, while GEO-X is a small exploration mission of the Earth's magnetosphere. Both missions need an ultralight-weight (<1 kg) telescope with moderately good angular resolution (<10 arcmin) at an extremely short focal length (<30 cm). We plan to demonstrate this type of telescope in these two missions around 2020. (C) The Authors. Published by SPIE under a Creative Commons Attribution 3.0 Unported License.
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JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 41(2) 377-387 2018年2月 査読有りLow-thrust propulsion is a key technology for space exploration, and much work in astrodynamics has focused on the mathematical modeling and the optimization of low-thrust trajectories. Typically, a nominal trajectory is designed in a deterministic system. To account for model and execution errors, mission designers heuristically add margins, for example, by reducing the thrust and specific impulse or by computing penalties for specific failures. These conventional methods are time-consuming, done by hand by experts, and lead to conservative margins. This paper introduces a new method to compute nominal trajectories, taking into account disturbances. The method is based on stochastic differential dynamic programming, which has been used in the field of reinforcement learning but not yet in astrodynamics. A modified version of stochastic differential dynamic programming is proposed, where the stochastic dynamical system is modeled as the deterministic dynamical system with random state perturbations, the perturbed trajectories are corrected by linear feedback control policies, and the expected value is computed with the unscented transform method, which enables solving trajectory design problems. Finally, numerical examples are presented, where the solutions of the proposed method are more robust to errors and require fewer penalties than those computed with traditional approaches, when uncertainties are introduced.
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Flight Model Development and Ground Demonstration of Water Resistojet Propulsion System for CubeSatsTransactions of the Japan Society for Aeronautical and Space Sciences 16(5) 427-431 2018年 査読有り
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10th Symposium on Space Resource Utilization, 2017 2017年Recently, asteroid mining has been studied by various groups of people including researchers, governments, and private companies. One of the target resources is water which is thought to be contained in C-type asteroids because it can be used as propellant, drinking water, etc. A temporarily captured orbiter (TCO), a celestial body that is captured by the Earth’s gravity and temporarily revolves around the Earth, can be a good mining target because it is energetically close to the Earth. This paper proposes a simple yet novel system that directly uses solar energy to extract and collect water from TCOs.
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SICE Journal of Control, Measurement, and System Integration 10(3) 192-197 2017年 査読有り
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日本航空宇宙学会論文集 65(6) 219-226 2017年 査読有りEarth observation satellites can improve the flexibility of observation sites by having “maneuverability,” and low-thrust obtained by ion thruster will be a promising method for orbital change for micro-satellites. Designing low-thrust trajectories for these satellites is a multi-revolution and multi-objective (time/fuel-optimal) optimization problem, which usually requires high computational cost to solve numerically. This paper derives an analytical and approximate optimal orbit change strategy between two circular orbits with the same semi-major axis and different local time of ascending node, and proposes a graph-based method to optimize the multi-objective criteria. The optimal control problem results in a problem to search a switching point on the proposed graph, and mission designers can design an approximate switching point on this graph, by using two heuristic and reasonable assumptions that 1) the optimal thrust direction should be tangential to orbit and 2) the optimal thrust magnitude should be bang-bang control with an intermediate coast. Finally, numerical simulation with feedback control algorithm taking thrust margin demonstrates that the proposed method can be applicable in the presence of deterministic and stochastic fluctuation of aerodynamic disturbances.
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Transactions of the Japan Society for Aeronautical and Space Sciences 60(3) 181-191 2017年 査読有り最終著者This paper describes development strategies and on-orbit results of the attitude determination and control system (ADCS) for the world's first interplanetary micro-spacecraft, PROCYON, whose advanced mission objectives are optical navigation or an asteroid close flyby. Although earth-orbiting micro-satellites already have ADCSs for practical missions, these ADCSs cannot be used for interplanetary micro-spacecraft due to differences in the space environments of their orbits. To develop a new practical ADCS, four issues for practical interplanetary micro-spacecraft are discussed: initial Sun acquisition without magnetic components, angular momentum management using a new propulsion system, the robustness realized using a fault detection, isolation, and recovery (FDIR) system, and precise attitude control. These issues have not been demonstrated on orbit by interplanetary micro-spacecraft. In order to overcome these issues, the authors developed a reliable and precise ADCS, a FDIR system without magnetic components, and ground-based evaluation systems. The four issues were evaluated before launch using the developed ground-based evaluation systems. Furthermore, they were successfully demonstrated on orbit. The architectures and simulation and on-orbit results for the developed attitude control system are proposed in this paper.
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Geophysical Research Letters 44(23) 192-197 2017年 査読有り
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Advances in Space Research 58(4) 528-540 2016年8月 査読有り
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Initial Flight Operations of the Miniature Propulsion System Installed on Small Space Probe: PROCYONAerospace Technology Japan 14(ists30) Pb_13-Pb_22 2016年7月 査読有り
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AIAA/AAS Astrodynamics Specialist Conference, 2016 2016年Solar sail spacecraft is attractive for deep-space exploration since it can control the trajectory without any propellant. On the other hand, it has greater uncertainties in the dynamical system than the conventional spacecraft. Therefore, the solar sail trajectory should be designed with enough robustness against the disturbances. Similarly to the duty cycle in low-thrust trajectory design, a typical method designs a nominal trajectory with an underestimated acceleration model, and controls the solar sail spacecraft to follow the nominal trajectory by shedding the excess thrust in the actual operation. However, this method might be conservative or not robust enough to disturbances. To derive the robust trajectory, this paper introduces Stochastic Differential Dynamic Programming, which solves the quadratically expanded Bellman equation for the stochastic-optimal control problem. Numerical examples show that the proposed trajectories have more robustness to disturbances, and the expected value of performance becomes better.
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AIAA/AAS Astrodynamics Specialist Conference, 2016 2016年For a precise calculation of the solar radiation pressure force acting on a space vehicle, the effect of shadows casted by the spacecraft structure itself is a challenging issue. The calculation cost of the shadow effect for a high-fidelity geometrical model is too high to apply to a practical astrodynamics analysis. Thus, authors have proposed a pre-computed method for a fast and precise solar radiation calculation. In this method, the visibility function for each micro-surface is calculated beforehand and then expanded to a set of basis functions in an offline pre-computation stage. In the real-time calculation phase, the coefficients corresponding to the basis functions are used for fast and precise solar radiation computation. One of the key points of the proposed method is the selection of the basis functions, and this paper compares two major spherical basis functions for accurate and efficient shadow representation. Several calculation examples of the three-dimensional objects are demonstrated to evaluate the computational cost and accuracy of the proposed method. The examples show that the proposed method is capable of providing an efficient and accurate real-time solar radiation pressure computation with a capability of the optical property estimation.
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Advances in the Astronautical Sciences 158 239-258 2016年This paper proposes a robust-optimal trajectory design method for uncertain system to minimize the expected value of cost-to-go function in Dynamic Programming. The fundamental idea is introducing Stochastic Differential Dynamic Programming (SDDP), which solves stochastic-optimal control problem by the second-order expansion of Bellman's equation around reference trajectory. Most recent studies have focused on trajectory optimization assuming that the spacecraft can control the trajectory perfectly as planed however, the assumption is violated in realistic operations where uncertain events, such as navigation error or uncertainty on dynamical system, perturb the predetermined trajectory. Conventionally, experienced specialists empirically determine "margin" on optimal low-thrust trajectory by duty cycle or forced coast period. A proposed SDDP autonomously provides "margin" in optimization for future feedback as well. Numerical results by V-infinity leveraging problem show that SDDP has "margin" without duty cycle or coast period. Monte-Carlo simulation shows the SDDP solution has better performance than DDP considering uncertainty.
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Advances in the Astronautical Sciences 158 933-951 2016年This paper presents the time-optimal attitude control for spinning solar sail utilizing the controllability of the membrane reflectivity. Although solar sail is an ideal spacecraft due to its propellant-free acceleration by solar radiation pressure, conventional solar sails require fuel for its attitude control, which prevents completely propellant-free space exploration. It was demonstrated that fuel-free attitude control can be achieved by reflectivity control however, no effective control strategy has been proposed. The time-optimal control law proposed in this paper gives a solution for the problem. Besides, the time-optimal control law enables us to solve the orbital and attitude control simultaneously. As an example of orbital control, V-infinity leveraging problem is numerically solved, where fast calculation was realized by analytical formulations derived from the proposed control law. The result indicates the importance of incorporating transient thrust generated during attitude maneuvers into orbital design of reflectivity-controlled solar sail.
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Advances in the Astronautical Sciences 158 3053-3068 2016年The calculation of Solar Radiation Pressure (SRP) for complex shape spacecraft is complicated because the calculation cost increases along with the number of facets, and the error caused by the shadow effect tends to increase by the complexity of the spacecraft's structure. This research newly proposes a generalized semi-analytical SRP calculation method for such complex shape spacecraft. It was derived from the ideas of Generalized Sail Model, which is a technique for solar sails, and a computer graphics rendering method called Pre-computed Radiance Transfer. In order to verify the utilities of our method, comparison results with the ray-tracing method and the optical property estimation will be shown in this paper.
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Advances in the Astronautical Sciences 156 1617-1630 2016年Recently, reflectivity control device (RCD) is proposed as a fuel-free attitude control system for spinning sail spacecraft. In this research, an attitude control model for spinning sail spacecraft with reflectivity control capability was derived as an extension of Generalized spinning Sail Model (GSSM). It was found that attitude control capability is determined by three parameters, which depend only on geometric property and optical performance of RCD. The proposed model suggests that the attitude, or the spin axis direction of the sail, converges toward an equilibrium point, which can be controlled within some range determined by the three parameters by switching RCD. Finally, the fidelity of the model was evaluated using actual flight data of IKAROS during RCD operation.
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Proceedings of the International Astronautical Congress, IAC 2016年Copyright © 2016 by the authors. All rights reserved. The Exploration Mission-1 (EM1) is the first test flight of NASA's new Space Launch System. Scheduled for launch in 2018, EM1 will carry the Orion Multi-Purpose Crew Vehicle (MPCV) into a cislunar orbit, together with a secondary payload composed by 13 cubesat. Two of these cubesat are currently proposed by JAXA: EQUULEUS, a 6U Earth-Moon Lagrangian-Point orbiter (in collaboration with the University of Tokyo); and SLSLIM, a 6U Moon lander. This paper presents the mission analysis work for EQUULEUS, while a second paper presents the mission analysis work for SLSLIM. EQUULEUS mission objectives are demonstrating cubesat orbit control techniques within the Sun-Earth-Moon regions; understanding the Earth's radiation environment; characterizing the flux of impacting meteors at the far side of the Moon; and demonstrating future exploration scenarios with a deep-space port at the Lagrange points. Following MPCV disposal, EQUULEUS is separated by the upper stage towards a lunar flyby, which, if not corrected, would result in an Earth escape trajectory. For this reason, after one-day orbit determination a trajectory correction maneuver is performed by the onboard thrusters to pump up the flyby perilune and put the spacecraft into an Moon-return orbit. Exploiting Sun perturbations, multiple lunar flybys and small trajectory correction maneuvers, EQUULEUS will be finally placed into a libration orbit around the Earth-Moon L2 point. We present the trajectory design process and a few sample trajectories, with the current baseline and the launch window analysis. Several astrodynamics techniques are described, including the search for Lunar-return orbits in the Earth-Sun Circular Restricted Three-Body Problem (first introduced by Lantoine in [1], and further developed by Garcia [2] for EQUULEUS and other applications); and the design of Libration orbits and low-energy transfers in real ephemeris.
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日本航空宇宙学会論文集 64(2) 131-138 2016年 査読有りWe propose thrust vector management by correctly positioning the thruster on a spacecraft by thrust vector measurement to decrease unwanted torque of thrust vector misalignment. A ground test was performed to measure 2-dimensional ion current distribution of 10W-class miniature ion thruster by electrostatic probe. The thrust vector measurement test showed that the thrust vector inclining angle was 1.4º from the geometrically symmetric axis of the thruster. The thruster was positioned on the first interplanetary micro-spacecraft: PROCYON after redesigning thruster bracket. Thrust vector estimation in the initial on-orbit operation of 6.5 hours showed that thrust vector passes through within 5mm of the PROCYON's center of gravity.
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Proceedings of the International Astronautical Congress, IAC 7 5231-5239 2015年Copyright © 2015 by the American Institute Federation of Aeronautics and Astronautics. Inc. All rights reserved. PROCYON is the first deep-space micro-spacecraft; it was developed at low cost and short time (about one year) by the University of Tokyo and JAXA, and was launched on December 3rd, 2014 as a secondary payload of the H II A launch of Hayabusa2. The mission primary objective is the technology demonstration of a microspacecraft bus for deepspace exploration; the second objectives are several engineering and science experiments, including an asteroid flyby. This paper presents PROCYON high-fidelity, very-low-Thrust trajectory design and implementation, subject to mission and operation constraints. Contingency plans during the first months of operations are also discussed. All trajectories are optimized in high-fidelity model with jTOP, a mission design tool first presented in this paper. Following the ion engine failure of March 2015, it was found the nominal asteroid could not be targeted if the failure was not resolved by mid-April. A new approach to compute attainable sets for low-Thrust trajectories is also presented.
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Aerospace Technology Japan 12(ists29) Tr_1-Tr_5 2014年11月 査読有り
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Aerospace Technology Japan 12(ists29) Pk_43-Pk_50 2014年11月 査読有り
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Aerospace Technology Japan 12(ists29) Pd_85-Pd_90 2014年11月 査読有り筆頭著者
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Aerospace Technology Japan 12(ists29) Tk_19-Tk_24 2014年8月 査読有り
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Aerospace Technology Japan 12(ists29) Tk_7-Tk_11 2014年7月 査読有り
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 12(ists29) Pb_33-Pb_38 2014年5月 査読有りSolar sail is a spacecraft that has a large-scale membrane to utilize the solar radiation pressure for its thrust. Hence, maintaining the membrane structure during space flight is a critical issue to keep thrust performance of the spacecraft. In this paper, we focused on the electrostatic force due to spacecraft charging on the membrane as one of the possible factor to cause the deformation of the membrane structure. We had estimated the electrostatic force via charging simulation for the IKAROS spacecraft in solar wind plasma at 1.0 AU. We had also made a structural analysis for the deployed membrane of IKAROS with the electrostatic force. The structural analysis showed that the electrostatic force could hardly affect the membrane structure in this case.
MISC
279所属学協会
1共同研究・競争的資金等の研究課題
5-
日本学術振興会 科学研究費助成事業 基盤研究(S) 2016年5月 - 2021年3月
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日本学術振興会 科学研究費助成事業 若手研究(B) 2013年4月 - 2017年3月
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日本学術振興会 科学研究費助成事業 若手研究(B) 2011年 - 2012年
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日本学術振興会 科学研究費助成事業 若手研究(スタートアップ) 2008年 - 2009年
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日本学術振興会 科学研究費助成事業 特別研究員奨励費 2004年 - 2006年
