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Automated spacecraft docking is a technology that has long been pursued. Deep space explorers and small spacecraft can carry fewer resources for docking, such as navigation sensors or latching structures, than can their larger near-Earth counterparts. The concept of the probe–cone docking mechanism is an effective solution to this problem. In this approach, a probe attached to the chaser satellite is guided automatically to the connection part of the target satellite by a conical structure. It is important to have a shock attenuation mechanism at the docking interface to prevent the chaser from being bounced away from the target. In the present paper, an automated docking mechanism that uses a flexible and deployable boom as the probe is proposed, and results of an analysis of the multi-body system dynamics are presented. Although analytical investigations into docking dynamics have been reported, the dynamics depend on many interdependent design parameters, the interaction of which is yet to be investigated. The present work involved a numerical analysis of the effect of each design parameter on the satellite behavior. An energy-based index that can predict the success or failure of docking was also developed in this study. In addition, a design scheme for the parameters is presented based on the results of the analysis in which the optimal combination of the design parameters is determined by searching the solution space.

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.

Accepted: 2019-07-03, 資料番号: SA1190135000

The solar power sail is an original Japanese concept in which electric power is generated by thin-film solar cells attached on the solar sail membrane. Japan Aerospace Exploration Agency (JAXA) successfully demonstrated the world’s first solar power sail technology through IKAROS (Interplanetary Kite-craft Accelerated by Radiation of the Sun) mission in 2010. IKAROS demonstrated photon propulsion and power generation using thin-film solar cells during its interplanetary cruise. Scaled up, solar power sails can generate enough power to drive high specific impulse ion thrusters in the outer planetary region. With this concept, we propose a landing or sample return mission to directly explore a Jupiter Trojan asteroid using solar power sail-craft OKEANOS (Oversize Kite-craft for Exploration and AstroNautics in the Outer Solar System). After rendezvousing with a Trojan asteroid, a lander separates from OKEANOS to collect samples, and perform in-situ analyses in three proposed mission sequences, including sending samples back to Earth. This paper proposes a system design for OKEANOS and includes analyses of the latest mission.

The present paper proposes a control method to excite spinning solar sail membranes for three-dimensional use. Using optical property switching, the input is given as the change in magnitude of the solar radiation pressure. The resonance point of this system varies with the vibration state due to its nonlinearity and the change in equilibrium state. To deal with this, a state feedback control law that automatically tracks the resonance point is developed in the present study. The proposed method enables decentralized control of the actuators on the sail, each of which determines the control input independently using only the information of vibration state. The proposed method is validated using numerical simulations. The results show that the nonlinear system behaves differently from the linear system, and the vibration grows using the decentralized control regardless of resonance point variation.

Accepted: 2018-01-07, 資料番号: SA1180389000

Accepted: 2018-12-05, 資料番号: SA1180272000

Accepted: 2017-06-22, 資料番号: SA1180299000

A novel approach for shape control of membrane structures is presented to realize their use in three-dimensional and variable configurations. The shape control is accomplished by exciting a spinnin...

This paper provides the thruster control approach for micro-satellites to realize autonomous orbit and attitude control. The Space Robotics Laboratory (SRL) of Tohoku University and ALE Co. Ltd. are currently developing a micro-satellite called “ALE-2” to generate artificial meteors. This satellite is required to keep the sun-synchronous orbit at the altitude of 375-400 km due to operational safety and mission constraints. ALE-2 is equipped with the High Density Cold Gas Jet System (HDCGJ) developed by Patchedconics, LLC for orbit and attitude control. Although a variety of thruster control methods have been proposed, this research focuses on a design of a reliable, uncomplicated and robust algorithm using the low-thrust, 4 nozzle propulsion system. A precomputed thruster selection table is utilized to reduce the processing power and to optimize the thrust force and torque distribution. The evaluation result from the full-software closed-loop simulation demonstrated that the proposed approach was valid for the orbit control of ALE-2 while high attitude control stability could be maintained in the long term.

Accepted: 2018-06-20, 資料番号: SA1180198000

Accepted: 2018-01-19, 資料番号: SA1180311000

Physical characteristics: Original contains illustrations, Accepted: 2018-05-15, 資料番号: PA1920002000

Accepted: 2018-03-02, 資料番号: SA1180094000

Accepted: 2018-04-11, 資料番号: SA1180095000

Accepted: 2017-11-25, 資料番号: SA1170256000

Accepted: 2017-01-13, 資料番号: SA1160256000

資料番号: SA1170276000

Accepted: 2016-09-21, 資料番号: SA1160305000

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.

Accepted: 2016-06-22, 資料番号: SA1160096000

Release from a high-altitude balloon is a cost-effective means of achieving supersonic flight, but there have been only two prior applications of this method to aircraft flight testing. The Japanese Aerospace Exploration Agency’s D–SEND#2 flight experiment is the latest of this kind of test. Due to the one-shot nature of the experiment, it was not sufficient to guarantee the stability of the test vehicle, but it was also necessary to achieve the mission objectives on the first flight. Accordingly, we performed an inclusive pre-flight evaluation of D–SEND#2 including the flight controller and operating conditions. This approach was very effective at enhancing the probability of mission success under off-design flight conditions. The actual flight results presented in this paper demonstrate the effectiveness of this approach.

Microwave components, prototypes of 4×4 transmitting active integrated antenna array(AIAA) and 2×2 receiving array with direction of arrival function operating at the X-band were demonstrated in this paper. For the components of the satellite on-board tracking system, a wideband and high gain low noise amplifier (LNA) as well as the compact planar antenna were fabricated using thin substrates. High isolation between the transmitting signal and the receiving signal was achieved for the good satellite tracking system. Good agreement on measured antenna patterns with the theoretical ones was obtained.

資料番号: SA1140118000

– The Planetary Material Sample Curation Facility of JAXA (PMSCF/JAXA) was established in Sagamihara, Kanagawa, Japan, to curate planetary material samples returned from space in conditions of minimum terrestrial contaminants. The performances for the curation of Hayabusa-returned samples had been checked with a series of comprehensive tests and rehearsals. After the Hayabusa spacecraft had accomplished a round-trip flight to asteroid 25143 Itokawa and returned its reentry capsule to the Earth in June 2010, the reentry capsule was brought back to the PMSCF/JAXA and was put to a series of processes to extract recovered samples from Itokawa. The particles recovered from the sample catcher were analyzed by electron microscope, given their ID, grouped into four categories, and preserved in dimples on quartz slide glasses. Some fraction of them has been distributed for initial analyses at NASA, and will be distributed for international announcement of opportunity (AO), but a certain fraction of them will be preserved in vacuum for future analyses.

Accepted: 2012-03-27, 資料番号: SA1004104000

A study on the post-HINODE Solar Observation Mission has been started by members in the solar physics community. One candidate of the mission targets on the observation of the solar polar region from the orbit largely inclined with the ecliptic plane. In order to achieve this severe mission target, possible trajectory sequences are investigated considering the application of various trajectory manipulation techniques. Three major trajectory options are listed up to achieve the mission objective. The first option, "SEP option", is characterized by the usage of solar electric propulsion (SEP) combined with the earth gravity assists. The other options are characterized by the usage of planetary gravity assists, and SEP is not used in these options. They are named "Jupiter option" and "Venus option" respectively. The comparison among the trajectory options is discussed in the paper, not only from the aspect of the orbital mechanics, but also from the aspects of the spacecraft design and operation.

The orbit of a 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 is managed by Japan Aerospace Exploration Agency (JAXA). The IKAROS attitude, i.e., the direction of its spin-axis, is nominally controlled by the rhumb-line control method. By utilizing the solar radiation torque, however, we are able to change the direction of the spin-axis by only controlling its spin rate. With this spin rate control, we can also control indirectly the solar sail’s trajectory. The main objective of this study is to construct the orbit control strategy of the solar sail via the spin-rate control method. We evaluate this strategy in terms of its propellant consumption compared to the rhumb-line control method. Finally, we present the actual flight attitude data of IKAROS and the change of its trajectory.

Accepted: 2011-02-19, 資料番号: SA1003628000

Accepted: 2011-06-01, 資料番号: SA1002966000

The cathode-less electron cyclotron resonance ion engines, µ10, propelled the Hayabusa asteroid explorer, launched in May 2003, which is focused on demonstrating the technology necessary for a sample return from an asteroid, using electric propulsion, optical navigation, material sampling in a zero gravity field, and direct re-entry from a heliocentric orbit. It rendezvoused with the asteroid Itokawa after a two-year deep space flight using the ion engines. Though it succeeded in landing on the asteroid on November 2005, the spacecraft was seriously damaged. This delayed Earth return in 2010 from the original plan in 2007. Reconstruction on the operational scheme using thrust vector control of ion engines, Xe cold gas jets and solar pressure torque made Hayabusa leave for Earth in April 2007. Although most of the neutralizers were degraded and unable to be used in fall of 2009, a combination of an ion source and its neighboring neutralizer has kept the orbit maneuver to Earth including a series of final trajectory correction maneuvers. Finally, the spacecraft decayed in atmosphere and only the reentry capsule was retrieved from the Australian outback on June 14th, 2010. For the round trip space odyssey between Earth and the asteroid, the ion engines served the total accumulated operational time 39,637 hour·unit, the powered spaceflight in 25,590 hours, delta-V of 2.2 km/s, total impulse of 1 MN·s and 47 kg Xenon propellant consumption. 1 2

Hayabusa spacecraft performed approaching to and landing on the asteroid Itokawa from September to November in 2005 with sophisticated vision-based navigation and guidance. The spacecraft had some autonomous functions of station-keeping, descending, touching down and sampling. Since two of three reaction wheels for attitude control had malfunctions and the autonomous guidance did not work as designed, however, a ground-based landmark optical navigation system was developed. This paper presents the navigation, guidance and control system of Hayabusa spacecraft with some flight results.

We propose a novel strategy for precise landing of a space probe on an asteroid surface. In Hayabusa mission, a Japanese asteroid sample-return spacecraft launched in 2003, a small marker was dropped onto the ground and utilized as a landmark for landing. In this paper, we propose two scenarios to drop multiple markers one by one, with a view to more precise landing. Preliminary Monte Carlo simulation of guidance using multiple markers with assumption of some control errors are presented. It was indicated that more than one markers allow us to increase the landing accuracy compared to the conventional method using a single marker.

Accepted: 2009-01-12, 資料番号: SA1000721000

The instruments and actuators in the attitude control system of small spacecraft are restricted in weight and space, and need to be reduced in weight and size. The rhumb line control strategy is one of the most popular schemes in reorientation of spin-stabilized spacecraft, since it requires only a spin sun sensor and a single axis reaction control system. By being combined with active nutation control, rhumb line control can reorient the spin axis of the spacecraft to any direction. To verify the control strategy, we manufactured an attitude controller and demonstration experiments were conducted using a motion table at ISAS/JAXA. This paper reviews the configuration of the controller and the outlines of the experiments, and evaluates the control performance. The same type of controller was installed in the Solar Sail Subpayload Satellite (SSSAT) launched in September 2006. An attitude control experiment in orbit was not conducted because of trouble with the satellite, but new control logic for the SSSAT was implemented for the attitude controller. This paper also reviews the control logic of the SSSAT.

Reported in this paper are the results of a mission analysis conducted for an asteroid exploration mission. Following the results obtained from HAYABUSA, a Japanese asteroid explorer, the Japan Aerospace Exploration Agency has started studying the possibility of the next asteroid exploration mission. The mission studied gives priority to the “early” retrieval of a sample from an asteroid with a primitive composition. Therefore, the design of the spacecraft follows that of the HAYABUSA, basically as it is, and the spacecraft is planned to be launched early in the next decade. The objective of the mission analysis is to design a mission sequence which has a launch window early in the next decade, is feasible utilizing a HAYABUSA-type spacecraft, and whose target asteroid complies with the scientific objectives. The results include the selection of the target asteroid, the design of the nominal mission sequence, and the alternative sequence to overcome the drawbacks of the nominal sequence.