船瀬 龍

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.

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 .

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.

<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>

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.

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.

© 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.

© 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.

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.

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.

Earth observation satellites can improve the flexibility of observation sites by having &ldquo;maneuverability,&rdquo; 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.

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.

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.

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 1.4^&ordm; 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.

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.

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.

A thermal-infrared (TIR) imager system is developed for HAYABUSA2, which is planned to be launched in 2014 and aims at sample-return from a C-class near-Earth asteroid 162173 (1999JU3) considered to contain organic or hydrated materials. The system consists of a TIR imager and digital electronics, which are used not only for the scientific investigation of physical properties of the asteroid surface, but also for the assessment of landing site selection and safe descent operation onto the asteroid surface with in situ measurement. TIR adopts an uncooled bolometer. Image operations such as multiple images summation, dark image subtraction, and the compensation of dead pixels are processed onboard. A processing module is connected to sensor interfaces through SpaceWire in order to provide deterministic processing time. Data compression is also provided to reduce the restriction of transmission time, which provides the equivalent compression ratio as JPEG2000 in 1/30 processing time in average. A high-speed data recorder is connected through SpaceWire in order to record TIR data in parallel with other sensor data. The modularity of SpaceWire enables us to use these as built devices for TIR and inherits the same design as the long-wavelength infrared imager developed for the Venus climate orbiter Akatsuki. (C) 2014 Society of Photo-Optical Instrumentation Engineers (SPIE)

Planetary protection has been recognized as one of the most important issues in sample return missions that may host certain living forms and biotic signatures in a returned sample. This paper proposes an initiative of sample capsule retrieval and onboard biosafety protocol in international waters for future biological and organic constituent missions to bring samples from possible habitable bodies in the solar system. We suggest the advantages of international waters being outside of national jurisdiction and active regions of human and traffic affairs on the condition that we accept the Outer Space Treaty. The scheme of onboard biological quarantine definitely reduces the potential risk of back-contamination of extraterrestrial materials to the Earth. (C) 2014 The Authors. Published by Elsevier Ltd. This is an open access article under the CC BY license (http://creativecommons.org/licenses/by/3.0/).

Copyright ©2014 by the International Astronautical Federation. All rights reserved. PROCYON (PRoximate Object Close flY by with Optical Navigation) is world's first mission aimed to demonstrate the technology of a micro spacecraft deep space exploration and proximity flyby to asteroids. The mission is developed by the University of Tokyo in collaboration with ISAS, JAXA. The spacecraft is scheduled to be launched as a secondary payload in late 2014 with Hayabusa 2 spacecraft. PROCYON will first target back to the Earth using its miniature ion engine; then it will transfer to the target asteroid using Earth gravity assist; finally it will use optical navigation to perform proximity flyby of the asteroid. Due to the very low thrust and limited propellant of the mission, it is therefore important to ensure that the mission objective and requirements can still be satisfied under different conditions and parameters. In this paper, we present the results of a broad sensitivity study of PROCYONs trajectory due to various launch dates and mission parameters.

An attitude model for a general spinning solar sail spacecraft under the influence of solar radiation pressure is presented. This model, called "Generalized Spinning Sail Model", can be applied to realistic sails with nonflat surfaces that have nonuniform optical properties. The unique behaviors predicted by the generalized spinning sail model are verified by actual operation of the Japanese spinning solar sail spacecraft IKAROS. It is shown how imperfections in the sail surface affect the attitude motion of spinning sails, and a compact mathematical model that can precisely reproduce the spin-averaged motion of the spinning sails is derived. The stability conditions and a reduced model that preserves the key characteristics of the generalized spinning sail model are also derived to reveal the unique properties of the attitude behavior of spinning sails.

This paper describes achievements of the IKAROS project, the world's first successful interplanetary solar power sail technology demonstration mission. It was developed by the Japan Aerospace Exploration Agency (JAXA) and was launched from Tanegashima Space Center on May 21, 2010. IKAROS successfully deployed a 20 m-span sail on June 9, 2010. Since then IKAROS has performed interplanetary solar-sailing taking advantage of an Earth-Venus leg of the interplanetary trajectory. We declared the completion of the nominal mission phase in the end of December 2010 when IKAROS successfully passed by Venus with the assist of solar sailing. This paper describes the overview of the IKAROS spacecraft system, how the world's first interplanetary solar sailer has been operated and what were achieved by the end of the nominal mission phase. (c) 2012 Elsevier Ltd. All rights reserved.

In the overall context of a solar sail mission in the vicinity of the Trojan asteroids swarm around the L4 Lagrange point, the purpose of this work is to find an optimal sequence of asteroids rendezvous that accommodates given mission constraints. A currently considered strategy to solve this problem will be presented and design choices will be outlined. A subset of the Trojan asteroids database is first extracted based on orbital elements considerations and a tree containing all the potential sequences is generated. A first set of pruning techniques is applied to the tree in order to quickly reduce the search space by several orders of magnitude. A global optimization method is then used: it combines a branch-and-bound approach and an evolutionary algorithm to find good sequence order, departure date, transfer and coasting durations. In order to enable a fast computation of potential transfer costs, the dynamics are linearized around L4 and the ΔVs are computed analytically. Preliminary results show a drastic reduction of the search space along with a reasonable accuracy on the cost prediction. This method has been used to analyze a tour scenario starting from 588 Achilles and including three rendezvous; the final result provides a list of sequences of potential interest. © 2013 2013 California Institute of Technology.

The Japan Aerospace Exploration Agency (JAXA) makes the world's first solar power sail craft IKAROS demonstration of photon propulsion and thin film solar power generation during its interplanetary cruise. The spacecraft deploys and spans a membrane of 20 meters in diameter using the spin centrifugal force. It also deploys thin film solar cells on the membrane, in order to evaluate its thermal control property and anti-radiation performance in the real operational field. The spacecraft weighs approximately 310kg, launched together with the agency's Venus Climate Orbiter, AKATSUKI on May 21, 2010. This paper presents the summary of development and operation of IKAROS.

This paper describes a method of evaluating sail quality utilizing in-flight attitude behavior of spinning solar sailer IKAROS. Since the successful deployment of the sail, IKAROS has received SRP which strongly affects both translational and rotational motion of the spacecraft. The authors have derived the "Generalized Spinning Sail Model (GSSM)" to reproduce observed unique attitude behavior of IKAROS. Following the previous work, this paper attempts to relate the GSSM with sail quality such as sail shape and flatness. An optical FEM model is constructed to evaluate the precise SRP effect on the spacecraft, and some candidates of deformed sail shape is reproduced which is consistent with the observed attitude motion. We also conclude by the in-flight attitude behavior that the surface roughness of the IKAROS sail is 0.33% at minimum.

Solar power sail is a deep space probe to be powered by hybrid propulsion of solar photon acceleration and ion engines to explore outer planetary region of the Solar System without relying on nuclear technology. The Japan Aerospace Exploration Agency (JAXA) launched the world's first deep space solar sail demonstration spacecraft "IKAROS" (Interplanetary Kite-craft Accelerated by Radiation Of the Sun) on May 21, 2010. IKAROS succeeded in deploying a 20m-span solar sail on June 9 and demonstrated several key technologies for solar sail utilizing the deep space flight environment. JAXA is currently studying an outer solar system exploration mission using the demonstrated solar power sail technology. The mission plans to fly for Jupiter, where the spacecraft drops a tiny Jovian probe and performs a swing-by for a Trojan asteroid. Current scenario consists of the rendezvous with one of the Trojan asteroids that are at the Lagrange points L4/L5 associated with Sun-Jupiter system. About as large as 50m sail should be deployed for this mission according to preliminary mission analysis and related research is intensively being carried out in JAXA. JAXA plans to initiate the project in a few years and looks at the launch around 2020. © 2012 by JAXA.

Solar power sail is a deep space probe to be powered by hybrid propulsion of solar photon acceleration and ion engines to explore outer planetary region of the Solar System without relying on nuclear technology. The Japan Aerospace Exploration Agency (JAXA) launched the world's first deep space solar sail demonstration spacecraft "IKAROS" (Interplanetary Kite-craft Accelerated by Radiation Of the Sun) on May 21, 2010. IKAROS succeeded in deploying a 20m-span solar sail on June 9 and demonstrated several key technologies for solar sail utilizing the deep space flight environment. JAXA is currently studying an outer solar system exploration mission using the demonstrated solar power sail technology. The mission plans to fly for Jupiter, where the spacecraft drops a tiny Jovian probe and performs a swing-by for a Trojan asteroid. Current scenario consists of the rendezvous with one of the Trojan asteroids that are at the Lagrange points L4/L5 associated with Sun-Jupiter system. About as large as 50m sail should be deployed for this mission according to preliminary mission analysis and related research is intensively being carried out in JAXA. JAXA plans to initiate the project in a few years and looks at the launch around 2020. ©2012 AIAA.

It is well known that the thrust force of the solar sail due to the solar radiation pressure is changed by the orientation of the sail with respect to the Sun direction. Therefore, the orbit of the solar sail can be controlled by changing the attitude of the spacecraft. In this study, we consider the spinning solar power sail IKAROS (Interplanetary Kite-craft Accelerated by Radiation Of the Sun), which succeeded to become the world's first flight solar sail in orbit. The IKAROS attitude, i.e. the spin-axis direction is nominally controlled by the rhumb-line control method. By utilizing the solar radiation pressure (SRP) torque, however, we are able to change the direction of the spin-axis only by controlling its spin rate. This is because the spin axis direction relates to the balance between the angular momentum of spinning and the SRP torque. Thus, we can control the solar sail's orbit by controlling the spin rate. The main objective in this study is to construct the orbit control strategy of the spinning solar sail via the spin rate control.