Publisher：Advances in Space Research  

This paper proposes a method to estimate two sets of orbital parameters (inclination, right ascension of the ascending node, and nodal precession rate) of the collided objects from in situ measurements of sub-millimeter-size debris, using the first-ever accidental collision between US Iridium 33 and Russian Cosmos 2251 as an example. This collision released a large number of fragmentation debris and have had a tremendous impact on the environment around the Earth. Objects larger than 10 cm are cataloged and tracked, but it is difficult to observe sub-millimeter-size debris from the ground and predict their orbits accurately. Thus, this fact leads to insufficiently modelled sub-millimeter-size debris, despite their potential to space mission termination. Therefore, in situ debris measurements are essential to protect space systems against sub-millimeter-size debris. Previous studies assumed that an in situ debris measurements satellite records the position at the time of impact with sub-millimeter-size debris, and then demonstrated that it is possible to estimate the orbital parameters from in situ debris measurements when a single object is broken up. However, when two objects collide, measurements come from both objects, so the previous method cannot be simply applied. Besides, the inclination of the broken-up objects to which the previous method can be applied is constrained by the inclination of the measurements satellite. To solve this problem, this paper proposes using azimuth at the time of impact and applying the Lomb-Scargle Periodogram, a kind of frequency analysis. Finally, this paper demonstrates that it is possible to identify accidental collisions in space through in situ measurements.

DOI： 10.1016/j.asr.2025.03.075

Scopus

Language：English   Publisher：Japan Aerospace Exploration Agency (JAXA)  

In order to control the increase in space debris and to actively promote its removal, the activities of space companies are indispensable, and the incentive design for them needs to be properly implemented as a social system. However, under the fiat currency system, it has to rely on a design based on taxation and a system of subsidies, and there have been difficulties in obtaining a realistic solution. The authors have proposed a solution using a new digital currency (ADR currency) in the past, but there is another difficulty in that the concept is difficult to spread due to differences with the current monetary system. Therefore, in this study, a text-based game simulating the corporate activities of space companies in a society where legal tender and ADR currency coexist was developed as a serious game (a game whose main purpose is not entertainment) using generative AI. The method is simple and could be useful for simulating decision-making on various other issues related to space exploration.

DOI： 10.20637/0002001398

CiNii Research

Authorship：Lead author,　Corresponding author  

researchmap

Publishing type：Research paper (scientific journal)   Publisher：Acta Astronautica  

Laser ablation technology enables contactless debris removal, providing the advantages of safety and cost efficiency. To deorbit space debris using laser ablation, the attitude motion of the debris must be controlled in advance. Although a previous study derived an attitude control procedure, it empirically determines control gains and is easily affected by assumed preconditions. To solve this problem, this paper proposes an improved attitude control law using laser ablation. Previous methods use feedback control based on the angular velocity error and error quaternions to obtain a reference torque. By analogy with a magnetic attitude control law, this study designs a reference torque based on the angular momentum errors and presents theoretical condition for designing the control gains. Numerical examples using the proposed control law are conducted to control the target attitude from a random initial rotation of 1 rpm to an arbitrary attitude, which verify the effectiveness of the proposed method. Furthermore, the position, direction, and magnitude uncertainties of laser irradiation are introduced to numerically examine their effects on the control accuracy under environmental disturbances.

DOI： 10.1016/j.actaastro.2024.06.023

Scopus

researchmap

Publishing type：Research paper (scientific journal)   Publisher：Elsevier BV  

This paper proposes the formation keeping control law for deorbiting debris by a laser ablation. Laser ablation is vital technology for contactless active debris removal, where a chaser satellite with a laser system irradiates laser pulses to a target object to generate the ablation force for deorbiting. The deorbiting force decelerates the target, and the chaser must maintain its relative position and continue irradiating. In other words, both the chaser and the target are supposed to be deorbited simultaneously, where both have accelerations. Although conventional formation flying missions assume that only a chaser maneuvers, the formation flying in this paper considers that both a chaser and a target have accelerations. Thus, this paper derives the relative equations of motion between the chaser and the target in powered flight and their analytical solution using relative orbital elements. A control law based on the analytical solution is proposed, which determines the timings and directions of the laser ablation and the electrical thrust so that the formation periodically returns to a desired formation. Numerical simulations first examine the control law in two cases with different maneuver timings. Then, a Monte Carlo simulation is performed to verify the effectiveness of the control law for a variety of desired formations.

DOI： 10.1016/j.asr.2024.05.029

Web of Science

Scopus

researchmap

Language：English   Publisher：American Institute of Aeronautics and Astronautics  

Language：English   Publishing type：Research paper (scientific journal)  

Publishing type：Research paper (scientific journal)   Publisher：Springer Science and Business Media LLC  

This paper answers how multiple satellites are seen from the ground. This question is inspired by space-advertising, a public exhibition in the night sky using a dot matrix of satellites that are bright enough to be seen by the naked eye. Thus, it is important for space advertisement that the specific dot matrix is seen. Moreover, the stability of the dot matrix during a visible span is very valuable. To stabilize the dot matrix, this study formulates an apparent position of a dot from a representative dot seen from the ground. The formulation, linear functions of a set of relative orbital elements, reveals the appearance of the dot matrix. The proposed relative variable in the formulation drives the instability of the dot matrix, thereby revealing an initial stable configuration of deputies from a chief. The arbitrary dot matrix designed using the configuration is stable even at low elevations without orbital control during the visible span. (Figure presented.)

DOI： 10.1007/s42064-023-0180-6

Web of Science

Scopus

researchmap

Other Link： https://link.springer.com/article/10.1007/s42064-023-0180-6/fulltext.html

Language：English   Publishing type：Research paper (scientific journal)  

Publishing type：Research paper (scientific journal)   Publisher：Elsevier BV  

This study introduces a new method to estimate the orbital parameters (inclination, nodal precession rate, and right ascension of the ascending node) of the broken-up object using in-situ debris measurements. To reduce collision probability with tiny debris generated from breakups of satellites, it is required to identify the origin of breakups and predict the orbit of debris. In general, small debris cannot be tracked from the ground, thus an in-situ measurement is an effective method to detect sub-millimeter-sized debris. Previous studies developed a method to identify the location of breakup origin from an in-situ measurement, however, it can be applied to an ideal situation where the detection occurs at the intersection of the orbit planes. This study introduces a new method to estimate the orbital parameters by fitting the model of geocentric declination to the measurement data. This method can be applied to practical situations that cannot be validated in previous studies. This paper also adopts iteratively reweighted least squares to improve estimation accuracy and mitigate initial value dependencies.

DOI： 10.1016/j.actaastro.2023.12.038

Web of Science

Scopus

researchmap

Language：English   Publishing type：Research paper (scientific journal)  

This paper solves how multiple satellites are seen from the ground. This question is derived by a space-
advertising, a public mission in the night sky with a dot matrix of satellites which are bright enough to
see in the naked eye. Thus, it is important for space advertisement that the dot matrix is seen as expected.
Moreover, the stable dot matrix during a visible span is more valuable. This study relatively formulates
an apparent position of a dot from a representative dot seen from the ground to stabilize the dot matrix.
The formulation is linear functions of a set of relative orbital elements and reveals the appearance of
the dot matrix. A proposed relative variable in the formulation drives the instability of the dot matrix,
discovering an initial stable configuration of deputies from a chief. The arbitrary dot matrix designed by
the configuration is stable even at low elevations without orbital control during the visible span.

Repository Public URL： https://hdl.handle.net/2324/7238296

Publishing type：Research paper (scientific journal)   Publisher：Journal of Space Safety Engineering  

This paper proposes less fuel strategies for space debris removal. To mitigate the risk of space debris cost-efficiently, multi-rendezvous missions are under development. On the other hand, multi-rendezvous missions often require changing orbital planes of removal satellites, which requires a huge amount of ΔV. Therefore, this study focuses on exploiting the J2 perturbation force as an auxiliary force and aims to establish maneuver rules that minimize ΔV consumption while maximizing the benefit of the J2 perturbation. The J2 perturbation equation is explored analytically, which clarifies whether the change in the semi-major axis or the inclination dominates the efficiency of the exploitation. A straightforward criterion is extracted which determines the efficient maneuver based on the initial inclination of the satellite.

DOI： 10.1016/j.jsse.2024.08.002

Web of Science

Scopus

researchmap

Language：English   Publishing type：Research paper (scientific journal)   Publisher：THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES  

<p>This paper proposes an attitude estimation technique that focuses on rapid change in light intensity, called glint. The states of space objects, such as shape, attitude, and surface properties, are essential information for performing active debris removal and monitoring the operational states of the satellite. Light curve inversion is a cost-effective measure to inversely estimate dynamic states. However, the light curve inversion suffers from divergence when the initial estimate is far from the true state. The glint occurs when a specific geometrical constraint is satisfied, and this glint constraint can be exploited to correct the attitude estimate in a Kalman filter. Thus, assuming a target object as a flat plate, this paper examines whether glint can be used in attitude estimation to improve estimation accuracy. Numerical simulations verify that the attitude estimation filter successfully converges, even when the initial estimate contains a significant error.</p>

DOI： 10.2322/tastj.22.59

CiNii Research

researchmap

Publisher：AIAA SciTech Forum and Exposition, 2024  

This paper proposes a close approach analysis method considering object attitude change. Implementing a detailed close approach analysis is important for space situational awareness. Conventional conjunction assessment methods assume the shape of the object as a encompassing sphere, which excludes considering the attitude change. However, most of the spacecraft are not spherical, and conventional methods result in overestimation. The proposed method in this paper consists of two steps: 1) calculation of the time of close approach with two objects as point masses and 2) interpolation of the first and last states in the time duration when the close distance is less than a given threshold. Dual quaternions are employed to express the translational and rotational motions, which ensures the uniqueness of the interpolation. Numerical simulations are performed to verify the proposed method and compare it with a conventional method.

DOI： 10.2514/6.2024-1632

Scopus

Publishing type：Research paper (scientific journal)   Publisher：Journal of Space Safety Engineering  

This study proposes a relative orbit control law for laser debris removal missions considering the uncertainties of laser ablation and atmospheric drag. A removal spacecraft irradiates laser pulses to a target debris to generate the ablation force for deorbiting. The deorbiting force lowers the target altitude, and the removal spacecraft must follow it to maintain its relative position for continuous laser irradiation. The difficulty stems from uncertainties of the magnitude of laser ablation and external disturbances such as atmospheric drag. To tackle this problem, this study derives an adaptive control method using the Gaussian process regression to cancel the uncertainties with a nonparametric regression model. Numerical simulations verify the proposed control law under the uncertainties of laser ablation and atmospheric drag. The proposed control law can contribute to the realization of a safer and more secure mission not only for laser debris removal missions, but also for other on-orbit services.

DOI： 10.1016/j.jsse.2024.04.007

Web of Science

Scopus

researchmap

Web of Science

Language：English   Publishing type：Research paper (scientific journal)   Publisher：Journal of Guidance, Control, and Dynamics  

This paper considers high-fidelity solar radiation pressure that is modeled with a detailed reflection model. The Cook-Torrance model is used to represent the isotropic reflection. Using the spherical Gaussian distribution, the analytical approximation method of the solar radiation pressure is derived.

DOI： 10.2514/1.G005961

Web of Science

Scopus

researchmap

Repository Public URL： https://hdl.handle.net/2324/6791121

Language：English   Publisher：Elsevier  

An active debris removal method using a laser is a promising technology for its advantage in contactless operations. This paper deals with the attitude control of an uncooperative target by a laser, which is an important phase before deorbiting. The difficulty of attitude control by the laser stems from the torque directional constraint because the laser generates thrust along the normal vector of the irradiated face irrespective of the irradiating direction. Thus, the control torque along the normal vector cannot be generated, which makes the attitude control with the laser torque challenging. To tackle this problem, this paper first designs a reference controller that assumes arbitrary control torques are available. Then, a method for determining the irradiating point is proposed so that the difference between the reference torque and the actual one is minimized. Although the proposed controller does not guarantee theoretical convergence to the desired attitude, the effectiveness of the proposed controller is numerically verified for a box-type object. Furthermore, the robustness to the uncertainties of thrust magnitude and direction is also examined by Monte Carlo simulations.

CiNii Research

Authorship：Corresponding author   Language：English   Publishing type：Research paper (scientific journal)   Publisher：Elsevier {BV}  

An active debris removal method using a laser is a promising technology for its advantage in contactless operations. This paper deals with the attitude control of an uncooperative target by a laser, which is an important phase before deorbiting. The difficulty of attitude control by the laser stems from the torque directional constraint because the laser generates thrust along the normal vector of the irradiated face irrespective of the irradiating direction. Thus, the control torque along the normal vector cannot be generated, which makes the attitude control with the laser torque challenging. To tackle this problem, this paper first designs a reference controller that assumes arbitrary control torques are available. Then, a method for determining the irradiating point is proposed so that the difference between the reference torque and the actual one is minimized. Although the proposed controller does not guarantee theoretical convergence to the desired attitude, the effectiveness of the proposed controller is numerically verified for a box-type object. Furthermore, the robustness to the uncertainties of thrust magnitude and direction is also examined by Monte Carlo simulations.

DOI： 10.1016/j.actaastro.2022.04.024

Web of Science

Scopus

researchmap

Language：English  

This paper briefly introduces a new approach to estimate some orbital parameters of on-orbit satellite fragmentations (specifically, the direction of angular momentum at a specific time and the time change in direction of angular momentum) from in-situ debris measurements. This approach, as in previous studies, adopts a constraint equation derived from the fact that a piece of debris detected shares the geocentric position vector with an in-situ debris measurement satellite. However, unlike previous studies, this approach does not adopt a constraint equation that can be applied to the rate of change in right ascension of the ascending node of a broken-up object. Instead, this approach determines the inclination of a broken-up object from the maximum or minimum geocentric declination at the time of detection. Then, this approach finds out a candidate for the rate of change in right ascension of the ascending node of a broken-up object by assuming a circular orbit with a radius of the geocentric distance at the time of detection. Finally, using the constraint equation adopted, this approach estimates the right ascension of the ascending node at the time of breakup and calculates a correction for the rate of change in right ascension of the ascending node. This paper also verifies that this new approach works effectively under ideal conditions where all detections are assumed to be at the line of intersection of the two orbital planes of a broken-up object and an in-situ debris measurement satellite.

DOI： 10.3389/frspt.2022.867236

Web of Science

researchmap

Language：Others   Publishing type：Research paper (scientific journal)   Publisher：Journal of Space Safety Engineering  

End-of-life support of satellites is necessary to improve post-mission-disposal compliance rates for maintaining space environment. Deorbit mission with low thrust, e.g. a laser, induces a low-level deceleration on the target object that gradually lowers the target altitude. Since such a low-thrust trajectory is time-consuming, the risk of collision greatly influences the mission success rate. In this context, this paper assesses the collision risk during deorbit trajectories with low thrust. Furthermore, parametric studies for the relationship between the re-entry time and the risk of collision are performed.

DOI： 10.1016/j.jsse.2021.11.002

Web of Science

Scopus

researchmap

Language：Others   Publishing type：Research paper (scientific journal)   Publisher：Journal of Space Safety Engineering  

This paper presents an overview of space solar power satellites for the Moon and Mars mission and simultaneously demonstrates the compression of traditional power generation methods for the orbiter, lander, and habitat on Mars and the Moon. Interplanetary missions are where the space engineers work on the satellites, conceptual design of space habitat, and exploration system. The state of the art in those missions relies on radioisotope thermoelectric generators or solar panels attached with batteries to store power, both of which are plagued by limitations. For instance, when a spacecraft moves away from the Sun, the energy collection efficiency on the solar panel is reduced, and at temperatures below -100°Celsius, individual solar cells degrade unpredictably, deteriorating the performance of solar arrays. Furthermore, these power generation unit satellites carry a pack of batteries to store energy whose total system accounts for more than 10-25 % of the mass of the satellites. In an interplanetary mission, power generation and management are essential for research and investigation on the surface. For Mars, sandstorms affect the collection of energy at the attached solar panel to the rover. Moreover, the rovers must investigate at the far side of the Moon, where sunlight is unavailable for few consecutive days. This challenge can be suitably overcome by employing space solar power satellites, which can be used for wireless power transmission, independent of its location. Such techniques have possible applications towards power transmission for unmanned aerial vehicles for faster mapping purposes. As such, the dependence of those aerial vehicles towards fixed energy storage becomes alleviated. In hindsight, space solar power satellite serves as a potential for an improved energy transmission source than the traditional method for interplanetary rovers and habitat.

DOI： 10.1016/j.jsse.2021.10.008

Web of Science

Scopus

researchmap

Publisher：Proceedings of the International Astronautical Congress, IAC  

Space-advertisement is getting closer with advances in satellite formation flying, which displays an artificial constellation in the night sky. The formation shape composed of pixels of satellites equipping light sources is observed from the ground and different from formation in relative orbits. For realizing the advertisement shape of high quality, there are two requirements, placing the pixels of the advertisement shape properly and keeping the shape during a visible span. It is not enough to satisfy the requirements by the usual coordinate systems and equations. Considering a topocentric coordinate system, the new transforming equations are derived so that an arbitrary advertisement shape is designed at the target time. Also, stable orbits that keep the formation shape without control are found and an example of stable orbits is presented.

Scopus

Language：English   Publisher：THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES  

<p>This paper aims to assess the risk of mission termination for a large constellation of satellites in a low-Earth orbit. Many large constellations will be deployed to provide broadband network services using thousands of satellites. There is concern that such large constellations will have a serious impact on the long-term sustainability of outer space activities due to the rapid increase in population. First, therefore, the authors conducted an assessment under nominal activities (referred to as “business-as-usual”) on the basis of a prediction by ESA’s MASTER-2009 and NASA standard breakup model 2001 revision. The assessment found that nearly one catastrophic collision may happen in a large constellation, generating more than two million fragments as small as 1 mm in size. Second, the authors conducted a further assessment assuming a hypothetical collision of a satellite in a large constellation using the NASA standard breakup model and a spherical finite element model adopted in ESA’s MASTER-2009. In consequence, another catastrophic collision may happen to a large constellation, generating approximately a half-million fragments as small as 1 mm in size. Therefore, such catastrophic collisions and resulting secondary collisions should be prevented for large constellations.</p>

DOI： 10.2322/tastj.20.10

CiNii Research

researchmap

Publisher：Proceedings of the International Astronautical Congress, IAC  

The 21st century is marked by the unprecedented progress of human knowledge in space, quantum mechanics, and machine learning. This progress has in turn lead to a significant increase in satellite launches, with various organizations developing mega-constellations in space to provide new business opportunities. While current predictions estimate that space in Low Earth Orbit will be overcrowded in the future, and sustainability of space will become a critical question as the estimated traffic of 89,000 satellites will be launched by the year 2035. To allow sustainable development of space, especially in Low Earth Orbit, there is a need to develop accurate traffic management and active debris removal mission planning. Lately, several organizations proposed and completed space demonstration experiments for debris removal. Current their designs are part of an ongoing project to develop contact-based debris removal technology. However, this technology poses several problems for satellite removal. We propose a complementary method based on a laser system. The laser is capable to operate from a safe distance by stopping the continuous rotation (detumbling) of space objects. In this paper, we present a case study using the ENVISAT satellite. ENVISAT is an 8,211 kg satellite, launched by European Space Agency. Since 2012, communication has stopped, and it is no longer operational. It is predicted to self-deorbit in the next 150 years. However, the orbital simulation of existing objects found two cataloged space debris which will pass within 200 m., resulting in a close approach every year. The elevated mass of the satellite and the number of close approaches, therefore, pose a considerable threat, leading to the production of debris. SKY Perfect JSAT Corporation is working jointly with Japanese institutions to collaboratively develop laser satellites to stop detumbling and the careful removal of debris. The demonstration mission will be launched by the year 2026-2027. With this case study, a laser-based contactless method can be practiced for removing mega debris like ENVISAT from the Low Earth Orbit.

Scopus

Publisher：Proceedings of the International Astronautical Congress, IAC  

Laser ablation is a vital technology for contactless active debris removal, where a service satellite with a laser system irradiates laser pulses to a target satellite to generate the ablation force for deorbiting. The deorbiting force decelerates the target, and the service satellite needs to maintain its relative position and keep irradiating. In other words, both the service satellite and the target are supposed to be deorbited simultaneously, where both satellites have accelerations. The difficulty of simultaneous deorbiting stems from the relative motion between the service satellite and the target in powered flight because conventional formation flying missions assume that only a service satellite maneuvers. This paper derives the relative equations of motion between the service satellite and the target in powered flight. A control law for the simultaneous deorbit is proposed, which determines the timing and direction of the laser ablation and the electrical thrust so that the formation periodically returns to the desired formation. Numerical simulations are performed for two test cases to verify the control law under the uncertainties of thrust magnitude and orbital perturbation.

Scopus

Publisher：Proceedings of the International Astronautical Congress, IAC  

Collisions and explosions of satellites generate a large amount of sub-millimeter-size debris, which can cause fatal damage to a spacecraft. However, such tiny debris cannot be tracked from the ground. Therefore, an in-situ debris measurement satellite, which can detect impacts with sub-millimeter-size debris, has been proposed. Based on this concept, previous studies proposed the method to estimate some orbital parameters of the broken-up object. In those studies, it is assumed that the measurement satellite detects impacts at the intersection of the orbital planes of the measurement satellite and the broken-up object. However, simulated measurement data includes fragments contrary to this assumption. Thus, this study introduces a new approach to estimating the orbital parameters of the broken-up object from simulated measurement data. In this study, the inclination, the right ascension of the ascending node, and the nodal precession rate are estimated from the history of geocentric declination using the iteratively reweighted nonlinear least square method.

Scopus

Language：English   Publisher：Space Solar Power Systems Society  

<p>Wireless Power Transmission (WPT) technology using a satellite-to-satellite system represents a valuable and convenient technology for transferring power wirelessly among Space Solar Power Satellites (SSPS) to Satellite and potential upcoming interplanetary missions. This direct transmission offers a possible solution to deliver continuous, convenient, and unlimited energy supply to satellites to help replace traditional power storage and reduce the weight and ultimately the costs of launching satellites. Satellite industries traditionally use photovoltaic cells and nuclear generators to satisfy the needed electricity by spacecraft. Current power generation and effective management systems occupy up to 10-25% of the satellite's mass. The concept of laser-based WPT from Energy Satellite (E-Sat) can overcome substantial problems. This consistent idea can be adopted for spacecraft by developing a constellation of E-Sat called Energy Orbit (E-Orbit) to supply sufficient power to spacecraft within range. It will increase the impressive performance and operational lifetime. In addition, creating 1600 E-Sat’s constellations to fulfill the power demand in Low Earth Orbit. The overall efficiency variation depends on the selection of Laser, transmitter, transmission distance, and photovoltaic cells, the same as increasing the maximum transmission efficiency of information in a wireless communication network. Consequently, in general, and adequate guidelines of the satellite-to-satellite power transmission system design in practice. The development and demonstration of this technology can help fulfill Space Solar Power Satellite’s idea to transfer gigawatts of renewable energy to Earth.</p>

DOI： 10.24662/sspss.7.0_12

CiNii Research

Publisher：Proceedings of the International Astronautical Congress, IAC  

Solar radiation pressure (SRP) is a major disturbance for a spacecraft operated in deep space missions, affecting its orbit and attitude, and high accuracy of the SRP calculation is required. The accuracy of SRP calculation is affected by the shape and surface characteristics of the satellite. However, existing computational models do not sufficiently take into account the exact reflective properties of the satellite surface. A detailed reflection model and considering the optical characteristics yield a different SRP and its torque from the conventional ones, resulting in a different attitude motion. In this study, an SRP calculation method that considers the exact reflective properties is applied to investigate the attitude dynamics.

Scopus

Language：Others   Publishing type：Research paper (scientific journal)  

This study aims to develop an effective strategy for geostationary satellites to avoid collisions with space debris satisfying various operational constraints. Recently, the number of space objects is continuously increasing to threaten operational satellites in the geostationary region. Studies on collision avoidance have been advanced for a long time, but not sufficiently considered orbit maintenance after maneuvering. Not to shorten satellite lifetime by collision avoidance, it is necessary to consider orbit maintenance after maneuvering. Besides, orbit maintenance regularly conducted differs between chemical thruster and electric thruster, so that two different avoidance strategies may be required. This study adopts a Multi-objective Genetic Algorithm to find out an optimal strategy for each thruster under various constraints on the operation. This paper also demonstrates that the optimal collision avoidance maneuvering can achieve both collision avoidance and orbit maintenance with less fuel than regular orbit control.

DOI： 10.1016/j.jsse.2021.08.004

Language：English   Publishing type：Research paper (scientific journal)  

Language：English   Publishing type：Research paper (scientific journal)  

Language：English  

DOI： 10.2514/1.G005877

Language：English   Publishing type：Research paper (scientific journal)  

Language：English   Publishing type：Research paper (scientific journal)  

Light curves are formulated using bidirectional reflectance distribution functions (BRDFs), which describe the ratio of light intensity between irradiance and radiance. The simplest BRDF is parameterized with Lambertian diffusion and perfect specular reflection. This model has been widely used in research on light curves as well as solar radiation pressure (SRP) modeling. SRP and its torque are one of main disturbances that contribute to spacecraft translational and rotational motion, which are also formulated with BRDFs.

DOI： 10.2514/1.G004966

Repository Public URL： http://hdl.handle.net/2324/4481550

Language：Others   Publishing type：Research paper (scientific journal)  

© 2018 COSPAR Even sub-millimeter-size debris could cause a fatal damage on a spacecraft. Such tiny debris cannot be followed up or tracked from the ground. Therefore, Kyushu University has initiated IDEA the project for In-situ Debris Environmental Awareness, which conducts in-situ measurements of sub-millimeter-size debris. One of the objectives is to estimate the location of on-orbit satellite fragmentations from in-situ measurements. The previous studies revealed that it is important to find out the right nodal precession rate to estimate the orbital parameters of a broken-up object properly. Therefore, this study derives a constraint equation that applies to the nodal precession rate of the broken-up object. This study also establishes an effective procedure to estimate properly the orbital parameters of a broken-up object with the constraint equation.

DOI： 10.1016/j.asr.2018.07.034

Language：English   Publishing type：Research paper (scientific journal)  

An optimal formation reconfiguration method under the constraints of a satellite attitude with respect to an inertial frame is addressed. Both the satellite position and attitude are controlled by only two body-fixed thrusters for an in-plane maneuver. To tackle the underactuated control problem, an attitude controller for tracking reference accelerations is firstly derived on the basis of Lyapunov approach. This controller allows us to consider the attitude constraints as input directional constraints because the satellite attitude is controlled so that the thrust direction is coincide with the force direction required for the orbit transfer. Secondly, a formation reconfiguration method based on the Fourier series is used as the reference inputs, and boundary conditions that make the resulting input trajectory an ellipse are shown. Such elliptic input trajectory changes the input direction monotonically, which enables bounding it around an desired direction. The proposed underactuated-controller achieves a reconfiguration maneuver while keeping the satellite attitude within a range from a specified direction, and thus is useful when several thrusters of a satellite fail due to malfunctions. Finally, numerical simulation results validate the effectiveness of the proposed relocation method by comparing energy consumptions and bounded satellite attitude angles.

DOI： 10.1016/j.ast.2018.03.021

Repository Public URL： http://hdl.handle.net/2324/4481549

Language：English   Publishing type：Research paper (scientific journal)  

The fault tolerance of spacecraft actuators significantly affects the reliability of satellites and the likelihood of successful missions. To enhance the fault tolerance of the actuators, this study derives optimal fault-tolerant configurations of fixed thrusters that maximize the controllability of a fully-actuated or underactuated satellite. The proposed method optimizes thrust and torque directions generated by the thrusters. Thus a cost function in terms of the thruster locations and directions is defined as the summation of the generated control forces and torques with respect to the body-fixed frame. The optimal configuration is obtained by the successive use of an energy potential method that is motivated by Thomson's problem. Some numerical examples are provided that show the effectiveness of the proposed formulation and optimization method.

DOI： 10.1016/j.asr.2018.01.014

Language：English   Publishing type：Research paper (scientific journal)  

This paper presents the results of line-of-sight (LOS) attitude control using control moment gyros under a micro-gravity environment generated by parabolic flight. The W-Z parameters are used to describe the spacecraft attitude. In order to stabilize the current LOS to the target LOS, backstepping-based feedback control is considered using the W-Z parameters. Numerical simulations and experiments under a micro-gravity environment are carried out, and their results are compared in order to validate the proposed control methods.

DOI： 10.1016/j.actaastro.2017.11.020

Language：English   Publishing type：Research paper (scientific journal)  

The study examines the control algorithm of a three-dimensional attitude and position of a free-floating rigid body with three thruster forces in which the force directions are fixed with respect to the body. This study provides a theory to develop a control method of an underactuated satellite with the minimum thruster number. In the procedure, three switching controllers are used in conjunction with motion planning in the final angular-rate deceleration phase to individually control the six state variables to the target values. The switching controllers have a hierarchical structure by using invariant manifolds as switching surfaces. The state variables in higher class manifolds that include lower class ones are adjusted by repeatedly adding intentional disturbances while the lower class state variables are returned to the original values by using lower class invariant manifolds. This study describes methods to define the invariant manifolds and also the intentional disturbance for achieving the forementioned control strategy. Finally, the motion planning in the angular-rate deceleration phase from a remained single-axis rotation finalizes the six state values of the body to the target values. Numerical simulations verify the proposed method.

DOI： 10.1155/2018/2825681

Language：English   Publishing type：Research paper (scientific journal)  

Spacecraft Line of Sight Maneuver Control Using Skew-arrayed Two Single-Gimbal Control Moment Gyros

DOI： 10.2322/tastj.14.Pd_31

Language：English   Publishing type：Research paper (scientific journal)  

Underactuated control offers fault-tolerance for satellite systems, which not only enables the position and attitude control of a satellite with fewer thrusters, but also can reduce the number of thrusters equipped on the satellite even when considering the need for backups. Due to having fewer thrusters, the coupling effect between the translational motion and rotational motion of the satellite cannot be avoided, and the coupled motion must be considered in control procedures. This paper presents a global trajectory design procedure required for the position and attitude control of an underactuated satellite. The satellite has four thrusters with constant thrust magnitudes on one plane of the satellite body. Then, an analytical solution for coupled motion between the rotation and translation of the satellite is obtained using three-step maneuvers of attitude control. The trajectory design based on the analytical solution is shown for the control of translational and rotational motion in three dimensions. Finally, a numerical simulation is performed to verify the effectiveness of the proposed design procedure.

DOI： 10.2322/tjsass.59.107

Language：English   Publishing type：Research paper (scientific journal)  

an optimization method for fault-tolerant control moment gyros (CMG) configurations, which maximizes the angular momentum envelope of the CMGs even after a portion of CMGs has failed. To incorporate weighting according to the number of available CMGs, the optimization problem is formulated as a min?max problem and reduced to Thomson's problem under appropriate constraints. Using different weights enables the determination of optimal CMG configurations that take into account a maximization of the angular momentum that admits the possibility of CMG failure. Therefore, the resulting configuration is both optimal and fault tolerant, and maintains satellite attitude controllability in underactuated conditions. Furthermore, reliability-based weighting is shown as one possible criterion for determining the appropriate weights. By adding constraints to the azimuth angles of the gimbal axes, it is shown that the zero-momentum condition of CMG configurations can be satisfied using the proposed optimization method. The proposed method enables calculating various types of optimal CMG configurations without providing initial estimations.

DOI： 10.2514/1.G001249

Language：English   Publishing type：Research paper (scientific journal)  

Proximity operations of satellites such as formation flying and on-orbit servicing offer more advanced missions than missions achieved by a single satellite. In a practical situation of formation flying, thrust directions for keeping and controlling a relative orbit is limited, e.g., for astronomical observation and plume impingement avoidance. The aim of this paper is to provide an energy efficient control method for a formation reconfiguration under input directional constraints with respect to both an inertial and a leader-fixed frames. The proposed controller is designed consisting of two parts: (1) guaranteeing a formation reconfiguration to a desirable formation on the basis of an energy optimal controller and (2) satisfying the input directional constraints by superimposing additional inputs. The analytical form of the control input shows that the input direction forms an ellipse in the leader-fixed frame when a particular boundary condition is satisfied, which is exploited as a nominal controller to take into account the input directional constraints. Due to the singularity avoidance of the nominal controller, the additional inputs can be analytically obtained. The effect on the follower trajectory due to the additional inputs is compensated by setting a virtual target orbit, and thus the successful formation reconfiguration is still guaranteed. Some numerical simulation results verify the effectiveness of the proposed method and compare the energy efficiency. (C) 2015 COSPAR. Published by Elsevier Ltd. All rights reserved.

DOI： 10.1016/j.asr.2015.07.037

Language：English   Publishing type：Research paper (scientific journal)  

This study proposes a new control method for under-actuated planar satellites having body-fixed thrusters that can generate unilateral and constant forces. The proposed method is applicable when the thrusters are not balanced, which means a rotational motion is inevitably induced whenever the thrusters are turned on. By applying the method, the position and attitude of any planar satellite can be controlled to arbitrary target states so long as a controllability condition is satisfied. The control logic is based on a switching control using plural manifolds. First, the satellite is transferred to an initial manifold, and then it moves to another manifold while decreasing its velocity and approaching the target states. The ON/OFF timing of the thrusters to transfer among the manifolds is explicitly specified in this paper. A numerical simulation validates the proposed control logic.

DOI： 10.1299/kikaic.79.225

Language：English   Publishing type：Research paper (scientific journal)  

This paper deals with the three-dimensional attitude control of an underactuated satellite equipped with thrusters whose force directions are fixed to the satellite. First, the necessary number of thrusters for the satellite's attitude control is discussed utilizing the Minkowski-Farkas theorem. Then, using the wz parameters for a satellite's attitude expression, this paper proposes a non holonomic attitude controller which is effective for any satellite regardless of its moment of inertia. Numerical simulation demonstrates the effectiveness of the proposed controller. Furthermore, the efficiency of the controller for different thruster positions is also discussed.

DOI： 10.20965/ijat.2011.p0892

Event date： 2024.10

Language：English   Presentation type：Oral presentation (general)  

Venue：Milan, Italy   Country：Italy  

Event date： 2024.9

Language：English   Presentation type：Poster presentation  

Event date： 2024.7

Language：English   Presentation type：Oral presentation (general)  

Venue：Busan, Korea   Country：Korea, Republic of  

Event date： 2024.7

Language：English   Presentation type：Oral presentation (general)  

Venue：Busan, Korea   Country：Korea, Republic of  

Event date： 2023.9

Language：English   Presentation type：Oral presentation (general)  

Venue：久留米   Country：Japan  

Event date： 2023.9

Language：English   Presentation type：Oral presentation (general)  

Venue：久留米   Country：Japan  

Event date： 2023.9

Language：Japanese   Presentation type：Oral presentation (general)  

Venue：富山   Country：Japan  

Event date： 2023.9

Language：Japanese   Presentation type：Oral presentation (general)  

Venue：富山   Country：Japan  

Event date： 2023.9

Language：Japanese   Presentation type：Oral presentation (general)  

Venue：富山   Country：Japan  

Event date： 2023.9

Language：Japanese   Presentation type：Oral presentation (general)  

Venue：富山   Country：Japan  

Event date： 2023.9

Language：Japanese   Presentation type：Oral presentation (general)  

Venue：富山   Country：Japan  

Event date： 2023.9

Language：Japanese   Presentation type：Oral presentation (general)  

Venue：富山   Country：Japan  

Event date： 2023.9

Language：English   Presentation type：Oral presentation (general)  

Venue：Sugarland, Texas   Country：Japan  

Event date： 2023.9

Language：English   Presentation type：Oral presentation (general)  

Venue：Sugarland, Texas   Country：Japan  

Event date： 2023.9

Language：English   Presentation type：Oral presentation (general)  

Venue：Orland, Florida   Country：United States  

Event date： 2023.6

Language：English   Presentation type：Oral presentation (general)  

Venue：Paris   Country：France  

Event date： 2023.6

Language：English   Presentation type：Oral presentation (general)  

Venue：Paris   Country：France  

Event date： 2023.6

Language：English   Presentation type：Oral presentation (general)  

Venue：Paris   Country：France  

Event date： 2023.6

Language：English   Presentation type：Oral presentation (general)  

Venue：Paris   Country：France  

Event date： 2023.6

Language：English   Presentation type：Oral presentation (general)  

Venue：Paris   Country：France  

Event date： 2023.6

Language：Japanese   Presentation type：Oral presentation (general)  

Venue：熊本   Country：Japan  

Event date： 2023.6

Language：Japanese   Presentation type：Oral presentation (general)  

Venue：熊本   Country：Japan  

Event date： 2023.6

Language：Japanese   Presentation type：Oral presentation (general)  

Venue：熊本   Country：Japan  

Event date： 2023.6

Language：Japanese   Presentation type：Oral presentation (general)  

Venue：熊本   Country：Japan  

Event date： 2023.6

Language：Japanese   Presentation type：Oral presentation (general)  

Venue：熊本   Country：Japan  

Event date： 2023.6

Language：Japanese   Presentation type：Oral presentation (general)  

Venue：熊本   Country：Japan  

Event date： 2023.6

Language：English   Presentation type：Oral presentation (general)  

Venue：Kurume, Fukuoka   Country：Japan  

Event date： 2023.6

Language：English   Presentation type：Oral presentation (general)  

Venue：Kurume, Fukuoka   Country：Japan  

Event date： 2023.3 - 2024.3

Language：Japanese   Presentation type：Oral presentation (general)  

Country：Japan  

Event date： 2022.6

Language：English   Presentation type：Oral presentation (general)  

Venue：online   Country：Japan  

Event date： 2022.6

Language：English  

Venue：online   Country：Japan  

Event date： 2022.6

Language：English   Presentation type：Oral presentation (general)  

Venue：online   Country：Japan  

Event date： 2022.6

Language：English   Presentation type：Oral presentation (general)  

Venue：online   Country：Japan  

Event date： 2022.6

Language：English   Presentation type：Oral presentation (general)  

Venue：online   Country：Japan  

Event date： 2019.12

Language：English  

Country：United States  

Event date： 2019.10

Language：English  

Venue：Washington   Country：United States  

This study aims to develop a new collision avoidance strategy for geostationary spacecraft. Recently, the continuous growth of space objects is increasing the collision risk among them. In the future, it will be necessary to operate the spacecraft while avoiding much space objects. However, there is a limit of fuel loading in a spacecraft, so it will be important to save fuel consuming each collision avoidance maneuvers not to shorten spacecraft lifetime rapidly. The proposed strategy considers the following points to save fuel and extend a spacecraft lifetime as much as possible. The first point is to avoid as many objects as possible at once maneuver so that the maximum collision probability is less than 1.0 × 10^%&. The second point is to minimize fuel consumed by collision avoidance maneuvers and the next orbit maintenance maneuver. The last point is to maximize the period to keep a spacecraft in the station-keeping slot. To realize this strategy, multi-objective genetic algorithm optimizes maneuver time and maneuver vector. In the simulation, to simplify, a spacecraft avoids two approaching objects. The Spacecraft conducts collision avoidance maneuver as the East-West control maneuver and the North-South control maneuver, separately. As a result, in the case of once maneuver, collision avoidance is succeeded but the station-keeping period was too short, especially in the case of an East-West control maneuver. Therefore, it needs to conduct many maneuvers before the closest approach.

Event date： 2017.9

Language：English  

Venue：Adelaide   Country：Australia  

Magnetorquers enable energy efficient attitude control of microsatellites in low Earth orbits. The challenging problem for the magnetic attitude control is the limitations on their component capability, and high-frequency Pulse- Width-Modulation (PWM) actuations are difficult to implement. Moreover since the magnetorquers generate control torques into a plane orthogonal to the geomagnetic field, three-axis attitude control using only the magnetorquers is challenging. In this context, this study deals with the magnetic attitude control of satellites using only magnetorquers that are driven with coarse PWM. The coarse PWM actuations in this study mean that the magnetorquers are successively actuated by on-state and then off-state in each actuation cycle of a few seconds. The difficulties of the coarse PWM actuations stem from the followings: 1) Conventional techniques for feedback controllers are hard to be applied because the magnitude of the magnetic moment is constant. 2) The dynamics of the satellite has nonlinear terms in control torques due to coarse PWM actuations. First, the dynamics of the satellite with on-off inputs are formulated as a discrete system in order to deal with the coarse PWM actuations of the magnetorquers. Then, to tackle with the nonlinearity of the control torques, averaging method for general case is used for reducing to the problem that is dynamic quantizer is applicable. As a preliminary study, dynamic quantizer is designed for satellites in sun-synchronous orbits and is verified for the coarse input system.

Event date： 2016.9

Language：Japanese  

researchmap

Event date： 2011.8

Language：English  

Venue：Portland, OR   Country：United States  

This paper discusses a position and attitude control problem of an underactuated satellite which uses on-off thruster mechanisms for control. Each thrusters has constant orientation relative to the satellite's body, and can generate only unilateral forces. The purpose of this paper is to clarify the control law for simultaneous position and attitude control for this type of satellite configuration. Considering the input constraints, we obtain the analytical solutions of the satellite's translational and rotational motion. Then the control procedure utilizing an invariant manifold is derived. Some numerical simulations are shown to verify the effectiveness of the proposed controller.

Event date： 2011.7 - 2011.8

Language：English  

Venue：Girdwood, AK   Country：United States  

This paper deals with the two-dimensional position and attitude control of a satellite by on-off thrusters which are fixed to the satellite's body. Since the thrusters generate constant and unilateral forces, the equations of motion form nonholonomic constraint and cannot be integrated. Furthermore, the satellite's mass changes due to the fuel consumption. Utilizing Fresnel integrals and the partial integrals, we derive the approximate analytic solution considering the satellite's mass change to achieve a specified satellite's position and attitude. The accuracy of the proposed approximate technique is verified by applying it to a condition considered in SLIM project of JAXA.

Language：English  

Language：Japanese  

Language：English  

Language：English  

Language：Japanese  

Language：Japanese  

Language：English  

Language：English  

Language：English  

Language：English  

Language：Japanese  

Language：Japanese  

Language：Japanese  

Language：Japanese  

Language：English  

Language：English  

Language：Japanese  

Language：Japanese  

Language：Japanese  

Language：Japanese  

Language：Japanese  

Language：Japanese  

Language：Japanese  

Language：Japanese  

Language：Japanese  

Language：Japanese  

Language：Japanese  

Language：Japanese  

Language：English  

Language：English  

Language：Japanese  

Language：Japanese  

Language：Japanese  

Language：English  

Language：English  

Language：English  

Language：English  

Language：English  

Country：Japan  

Language：English  

Language：English   Presentation type：Oral presentation (general)  

Venue：online   Country：Japan  

Language：Japanese   Presentation type：Oral presentation (general)  

Venue：online   Country：Japan  

Language：Japanese   Presentation type：Oral presentation (general)  

Venue：online   Country：Japan  

Language：Japanese  

Venue：オンライン   Country：Japan  

Language：Japanese  

Venue：オンライン   Country：Japan  

Language：Japanese  

Venue：調布，東京   Country：Japan  

Language：Japanese  

Venue：調布，東京   Country：Japan  

Language：Japanese  

Venue：調布，東京   Country：Japan  

Language：Japanese  

Venue：調布，東京   Country：Japan  

Language：Japanese  

Venue：調布，東京   Country：Japan  

Language：Japanese   Presentation type：Oral presentation (general)  

researchmap

Type：Peer review 

Number of peer-reviewed articles in foreign language journals：2

Type：Peer review 

Number of peer-reviewed articles in foreign language journals：5

Proceedings of International Conference Number of peer-reviewed papers：1

Type：Peer review 

Number of peer-reviewed articles in foreign language journals：4

Type：Peer review 

Number of peer-reviewed articles in foreign language journals：1