Language：English   Publisher：宇宙航空研究開発機構宇宙科学研究所 (JAXA)(ISAS)  

レポート番号: ASTRO-2023-C019(031)/In-person

CiNii Research

Language：English   Publisher：宇宙航空研究開発機構宇宙科学研究所 (JAXA)(ISAS)  

レポート番号: ASTRO-2023-C001(010)/In-person

CiNii Research

Language：English   Publisher：宇宙航空研究開発機構宇宙科学研究所 (JAXA)(ISAS)  

レポート番号: ASTRO-2023-A014(018)/In-person

CiNii Research

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

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

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

Language：Japanese   Publisher：The Japan Joint Automatic Control Conference  

DOI： 10.11511/jacc.66.0_536

CiNii Research

Language：Japanese   Publisher：The Japan Joint Automatic Control Conference  

DOI： 10.11511/jacc.66.0_526

CiNii Research

Publisher：Lecture Notes in Electrical Engineering  

For deep space landing missions, spacecraft are required to identify their expected landing sites autonomously because of the extremely long time delay caused by the distance between the spacecraft and Earth. This identification process is desirable to finish within several seconds by onboard computers with limited calculation performance. Moreover, autonomous identification based on natural features of landing sites are highly recommended in future missions, although some artificial target markers have been used for navigation and control to the landing site in some previous missions. To make fast but reliable identification of landing sites for the automatic task, this research utilizes a deep learning processing for images taken in different light-conditions and altitudes. First, a semantic segmentation model for rocks in terrain images is developed. For robust identification, some improvements are introduced in the semantic segmentation process. Then, to identify the same place in images taken at different altitudes, a comparison algorithm based on triangular shapes is applied. Thus after training, the semantic segmentation model can detect the same place from several images in a relatively short computational time.

DOI： 10.1007/978-981-19-2635-8_66

Web of Science

Scopus

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

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

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

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

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

DOI： 10.2514/1.G005916

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

DOI： 10.3934/math.2022390

Language：Japanese   Publisher：The Japan Joint Automatic Control Conference  

DOI： 10.11511/jacc.65.0_1450

CiNii Research

Publisher：33rd Congress of the International Council of the Aeronautical Sciences, ICAS 2022  

In this paper, we propose an altitude estimation method by using Wide-Field-Integration (WFI) of optic flow. WFI of optic flow is a motion estimation method inspired by a biological research on the visual processing system of flying insects' compound eyes. In the standard estimation process, altitude information of a flying vehicle is necessary to estimate the motion variables of flying vehicles. The method proposed in this paper can estimate the altitude as well as the motion variables by using two patterns of optic flow acquired by two onboard cameras. The proposed method is useful for reducing onboard instrument mass and thus suitable for small autonomous vehicles. This paper discusses the mechanism of the altitude estimation and explores the suitable positions of the WFI of optic flow sensors in simulations.

Scopus

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

DOI： 10.2514/1.G006035

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

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

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

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

DOI： 10.2514/1.G005046

Repository Public URL： http://hdl.handle.net/2324/4479056

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

DOI： 10.1002/asjc.2363

Repository Public URL： http://hdl.handle.net/2324/4479598

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

DOI： 10.2322/tjsass.63.243

Repository Public URL： http://hdl.handle.net/2324/4476143

Language：English  

This paper proposes a new numerical method for finding libration point orbits in the vicinity of collinear libration points in the circular restricted three-body problem. The main advantage of this method is that it requires neither an initial guess nor complex algebraic manipulations for finding both quasi-periodic and periodic orbits. The proposed method consists of two steps: center manifold design and differential correction. The first step provides a quasi-periodic orbit parametrized by a single parameter vector. In the second step, the parameter vector in the first step is used to obtain an exact periodic orbit. This method is applied to find periodic and quasi-periodic orbits in the Sun-Earth restricted three-body problem around the L₁ and L₂ libration points.

DOI： 10.1016/j.actaastro.2019.02.029

Repository Public URL： http://hdl.handle.net/2324/4784354

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

This paper proposes a new approach of using differentials in aerodynamic drag in combination with thrusters to control satellite formation flying in low Earth orbits. Parameterized output regulation theory for formation-flying missions with combined control action is developed based on the Schweighart–Sedwick relative dynamics equations. The theory is implemented to precisely track the different trajectories of reference relative motion and eliminates the effects of the J₂ perturbations. The parametric Lyapunov algebraic equation is proposed to ensure the stability of the linear relative model subject to saturated inputs. The main goal of this study is to approve the viability of using the differentials in aerodynamic drag to precisely control different formation-flying missions. Numerical simulations using a high-fidelity relative dynamics model and a high-precision orbit propagator are implemented to validate and analyze the performance of the proposed control algorithm in comparison with the linear quadratic regulator algorithm based on actual satellite models.

DOI： 10.2514/1.G004219

Repository Public URL： http://hdl.handle.net/2324/4479055

Language：English  

This study applies a sliding mode control (SMC) strategy for a robust controller of a quad-rotor vehicle. First, a controller combined with a nested control loop and an SMC is introduced, because a quad-rotor vehicle has only four control inputs although the vehicle has six degrees of freedom. The control performance for the feedback gains in the nested loop is investigated in numerical simulations. Subsequently, the effects of practical system limitations (control cycle and rotor dynamics) on the control performance are examined. Finally, the robust performance of the SMC strategy on a quad-rotor vehicle is discussed.

DOI： 10.20965/jrm.2018.p0397

Language：English  

DOI： 10.2514/1.G003334

Repository Public URL： http://hdl.handle.net/2324/4479054

Language：English  

The swing-by maneuver is known as a method to change the velocity of a spacecraft by using the gravity force of the celestial body. The powered swing-by has been studied to enhance the velocity change during the swing-by maneuver. This paper studies another way of the powered swing-by using tether cutting, which does not require additional propellant consumption, and shows that the proposed powered swing-by can increase the effect of the swing-by as same as using impulsive thrust. Moreover, it is discussed whether the system has possibility to realize both the powered swing-by of a mother satellite and the planetary capture of a subsatellite simultaneously.

DOI： 10.1016/j.actaastro.2017.11.002

Language：English  

Station-keeping and formation flying along a libration point orbit in the circular restricted three-body problem are considered. In order to deal with the relative motion with respect to a reference trajectory, this paper extends our previous study which derives the station-keeping controller based on the output regulation theory. First the reference orbit of the chief satellite is represented as the output of an autonomous system called exosystem, assuming the reference orbit is given by a truncated Fourier series. For the formation flying, the relative trajectory of the deputy satellite with respect to the chief satellite is also represented by the output of an exosystem. Then the reference signal to be asymptotically tracked for the formation flying is obtained by the superposition of the two exosystems. The proposed controllers are applied and verified for the station-keeping and formation flying along a periodic orbit of the Sun-Earth L2 point.

DOI： 10.1016/j.actaastro.2018.02.004

Repository Public URL： http://hdl.handle.net/2324/4784017

Language：English  

A new station-keeping strategy based on nonlinear output regulation theory has been proposed for the periodic orbits in the circular restricted three-body problem. A Fourier series approximation was employed to generate the desired orbits, and then the output regulation theory for nonlinear systems was applied. The output regulation problem was solved in an analytical form. The proposed controller has been applied and verified in numerical simulations for the halo orbit of the sun–Earth L2 point. In the case that the reference orbit is the natural periodic orbit, the controller can approximate the optimal controller in the sense of minimizing the station-keeping cost. As an application of the proposed method, formation flying using circular orbits around the chief satellite has been worked out.

DOI： 10.2514/1.G002845

Repository Public URL： http://hdl.handle.net/2324/4784365

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

The hyperbolic instability of collinear libration point orbits dominates the distribution of orbit uncertainty, which is closely related with the unstable manifold of the local orbit. Exploiting the distribution of orbit uncertainty in an unstable orbital environment, an integrated orbit determination and maintenance approach is proposed. X-ray pulsar navigation embedded in the Kalman filtering architecture is adopted to determine the orbit of a spacecraft flying along the Earth-Moon L₂ halo orbit. To keep the spacecraft in the vicinity of the nominal trajectory, the newly proposed orbit maintenance approach based on the covariance matrix is proposed to design the orbit maintenance maneuver. The numerical simulations show that the proposed approach has a good performance in orbit determination and maintenance.

DOI： 10.1016/j.actaastro.2018.11.041

Repository Public URL： http://hdl.handle.net/2324/4784018

Language：English  

This study introduces a new robust nonlinear control scheme based on the theory of nonsingular terminal sliding mode control (NTSMC). Since conventional NTSMC utilizes a discontinuous switching function, a significant flaw called chattering can occur. The main purpose of this study is to design a new switching function based upon Lyapunov stability in order to alleviate this drawback over time. There are many approaches to mitigate the chattering drawback in SMC such as utilizing a smooth approximation of the switching element, or employing higher order sliding mode control (HOSMC) strategy. However, the use of a continuous approximation affects the system's performance and a finite reaching time to the sliding manifold, and in HOSMC the estimation of high-order derivatives of states is usually difficult and it still exhibits chattering in the presence of parasitic dynamics. In this study by employing a new sliding manifold including a time function, the chattering is attenuated as well as keeping the robustness. Finally, a second-order nonlinear dynamical system subject to disturbance is simulated to highlight the validity and applicability of the proposed method.

DOI： 10.1002/asjc.1477

Repository Public URL： http://hdl.handle.net/2324/4479597

Language：English  

DOI： 10.5923/j.jmea.20170702.04

Language：English  

Language：English  

This paper investigates the combination of optimal feedback control with the dynamical structure of the three-body problem. The results provide new insights for the design of continuous low-thrust spacecraft trajectories. Specifically, the attracting set of an equilibrium point or a periodic orbit (represented as a fixed point) under optimal control with quadratic cost is obtained. The analysis reveals the relation between the attractive set and original dynamics. In particular, it is found that the largest dimensions of the set are found along the stable manifold and the least extent is along the left eigenvector of the unstable manifold. The asymptotic behavior of the structure of the attractive set when time tends to infinity is analytically revealed. The results generalize the use of manifolds for transfers to equilibrium points and periodic orbits in astrodynamic problems. The result is theoretical and developed for a linearized system, but it can be extended to nonlinear systems in the future.

DOI： 10.2514/1.G000524

Repository Public URL： http://hdl.handle.net/2324/4784356

Language：English  

Language：English  

This paper presents the trajectory design and analysis of the near-Earth asteroid (NEA) deflection mission enabled by a kinetic impact of an intermediate asteroid. The sequential transfer trajectory is designed by solving two Lambert's problems that yield a chain collision between a spacecraft and an intermediate asteroid, followed by a collision between the intermediate asteroid and an NEA. The characteristics of the low-cost trajectories are then identified with respect to the optimal collision point. We show that the feasibility of the mission depends on the existence of an intermediate asteroid with small minimum orbit interception distance (MOID) with the NEA. Moreover the selection strategy of an intermediate asteroid that makes the mission feasible is discussed. We show that several asteroids exist that allow a 10 ton spacecraft with limited ΔV to be launched and impact the NEA 99442 Apophis allowing a deflection given several years warning time.

DOI： 10.1016/j.asr.2015.08.016

Language：English  

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

DOI： DliD2rhdIa

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

DOI： AUJ6UxEFLM

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

Language：English  

A new control strategy based on output regulation theory is proposed for formation flying along a halo orbit near a libration point in the circular-restricted three-body problem. The reference orbit is a periodic or shifted perturbation of the halo orbit generated by an exosystem, and modification of linear output regulation theory is employed. First, the station keeping problem is considered, and the total velocity change required for the halo orbit maintenance is found to be negligible. The reasonable size of a perturbation can be determined numerically, and hence the formation flying proposed in this paper is useful for the design of a reference orbit along a halo orbit. The proposed relative orbit and control strategy are independent of halo orbits and can be applied to formation flying along an arbitrary path.

DOI： 10.2514/1.G000463

Language：Japanese  

Language：English  

Language：English  

Language：English  

The in-plane motion of the Hill-Clohessy-Wiltshire (HCW) equations is controllable with along-track input but not controllable with crosstrack input. However, in the controllable subspace of the latter, the drift velocity is zero and the free motion is periodic. Thus, formation reconfiguration for the in-plane motion is possible by cross-track input. In the case of the continuous-time feedback controllers, the total velocity change of the along-track controller is larger than that of the two-input controller, but the difference is less than 8% for the settling time larger than five periods. In the case of pulse feedback controllers, the sampling time used for discretization is one-half of the period for the two-input controller and one-quarter of the period for all other single-input controller. The two-input and along-track feedback controllers are robust and effective for the nonlinear relative dynamics.

DOI： 10.2514/1.57197

Language：English  

Language：English  

The motion of a charged satellite subjected to the Earth's magnetic field is considered. The Lorentz force, which acts on a charged particle when it is moving through a magnetic field, provides a new concept of propellantless electromagnetic propulsion. A dynamic model of a charged satellite, including the effect of the Lorentz force in the vicinity of a circular or an elliptic orbit, is derived and its application to formation flying is considered. Based on Hill-Clohessy-Wiltshire equations and Tschauner-Hempel equations, analytical approximations for the relative motion in Earth orbit are obtained. The analysis based on the linearized equations shows the controllability of the system by stepwise charge control. The sequential quadratic programming method is applied to solve the orbital transfer problem of the original nonlinear equations in which the analytical solutions cannot be obtained.Astrategy to reduce the charge amount using sequential quadratic programming is also developed.

DOI： 10.2514/1.57060

Language：English  

An active formation flying for the Tschauner-Hempel equations (TH) is considered, in which the desired relative orbit of the follower is generated by an exosystem. This allows for flexibility of the shape and period of the reference orbit. The regulator differential equation is solved algebraically, exploiting the special structure of the Tschauner-Hempel equations. Hence, control implementation is straight forward. In the case of active formation, the total velocity change increases with penalty on the input. The settling time is independent of the frequency of the reference orbit. For the Tschauner-Hempel equations, the total velocity change for transition increases with frequency of the reference orbit, but the settling time remains almost constant, as in the circular case. For the eccentricity up to 0.3, the total velocity change for maintenance remains at the same level as in the circular case. Using these relations, feedback controls for active formation with given specifications and constraints can be designed effectively.

DOI： 10.2514/1.57703

Language：Japanese  

Language：English  

Language：English  

Graphical generation of periodic orbits of Tschauner-Hempel (TH) equations is presented. The general solution of the relative dynamics based on the orbital elements is obtained, and the relative motion invariant manifold, where the solution lies, is determined. The region where all periodic solutions lie is determined by two ellipses, whose size depends on the eccentricity of the leader orbit. Size and phase parameters for periodic orbits are introduced for the TH equations, and a simple way to draw a periodic orbit is proposed for a given pair of size and phase parameters and a given initial true anomaly of the elliptic orbit. Results show that the design of a relative orbit with size and position specifications is as simple as in the Hill-Clohessy-Wiltshire (HCW) case.

DOI： 10.2514/1.56326

Language：English  

In this paper, we investigate the possibility of the use of the Lorentz force, which acts on charged satellite when it is moving through the magnetic field, as a means of satellite attitude control. We first derive the equations of attitude motion of charged satellite and then investigate the stability of the motion. Finally we propose an attitude control method using the Lorentz force. Our method requires moderate charge level for future Lorentz-augmented satellite.

DOI： 10.1016/j.actaastro.2011.07.019

Language：English  

This paperisconcerned withafuel-optimal relative orbit transfer problem for Hill-Clohessy-Wiltshire equations. The input is of the pulse type and the performance index is the total velocity change required for a transfer. For the in-plane motion, the existence of optimal three-pulse strategies is shown and the minimum total velocity change is explicitly given under the condition that the difference of orbit size is relatively larger than that of drift velocity. For the out-of-plane motion, the existence of optimal single-pulse strategies is shown. The orbit transfer problem covers formation reconfiguration problem as a special case. Midpoints of the optimal pulses coincide with those points at which the velocity along the radial direction is zero. Using this information and null controllability with vanishing energy of the Hill-Clohessy-Wiltshire equations, suboptimal feedback controllers are designed. Simulation results with optimal open-loop strategies and feedback controls are given for two formation reconfiguration problems.

DOI： 10.2514/1.51230

Language：English  

Language：English  

The purpose of this paper is to investigate the possibility of on asteroid (NEA) survey mission enabled by advanced solar sailing technology. The study is focused not on the solar sail spacecraft itself but on its orbital dynamics to realize the missions. A novel NEA flyby survey mission with a lightweight solar sail spacecraft to increase the accessibility to NEA flybys located in the vicinity of the Earth's orbit is proposed. A numerical study suggests that our approach increases the opportunities in proximity to NEAs, which have eccentric and inclined orbits.

DOI： 10.1007/BF03321532

Language：English  

Language：English  

A study was conducted to demonstrate the formulation of the new Lambert Algorithm using the Hamilton-Jacobi-Bellman Equation (HJB). The two-point boundary-value problem (TPBVP) of the Hamiltonian system was treated as an optimal control problem where the Lagrangian function played a role as a performance index. The approach demonstrated in the study was based on the expansion of the value function in the Chebyshev series with unknown coefficients, considering the computational advantages of the use of Chebyshev polynomials. The differential expressions that arose in the HJB equation were expanded in Chebyshev series with the unknown coefficients. The new algorithm had the potential to provide a solution to the TPVBP using the spectral information about the gravitation potential function and was applicable to the problem under a higher-order perturbed potential function without any modification.

DOI： 10.2514/1.46751

Language：English  

This paper is concerned with formation acquisition and reconfiguration problems with an eccentric reference orbit. As a first step, the characterization problem is considered for all initial conditions that constitute periodic solutions of the nonlinear equations of relative motion. Under the condition that the inertial orbits of the leader and a follower are coplanar, initial conditions of all periodic relative orbits are generated in terms of parameters of their orbits and their initial positions. Then the inertial orbit of the follower is rotated successively around the two axes of its perifocal reference system, and the initial conditions of all noncoplanar relative orbits are derived. Based on these periodic relative orbits, formation acquisition and reconfiguration problems by a feedback controller are formulated. The main performance index of a feedback controller is the total velocity change during the operation. Using the property of null controllability with vanishing energy of the Tschauner-Hempel equations, suboptimal controllers are designed via the differential Riccati equation of the linear regulator theory of periodic systems.

DOI： 10.2514/1.46024

Language：English  

A new formulation of formation flying in an elliptic orbit with restricted control interval is discussed. The main performance index is the L₁ norm of the control input generated by an employed feedback control, which is proportional to the fuel consumption. The null controllability with vanishing energy (NCVE) property enables to design feedback controllers through the linear quadratic regulator (LQR) theory, which generate control inputs with arbitrarily small L₂ norms by imposing a large penalty on control. The L₁ norm of the control input decreases monotonically as the penalty on control increases, and suboptimal controllers are designed. The velocity of the satellite near the apogee is smaller, a relatively small velocity change is required for orbit control and it is assumed that t = 0 is the perigee passage time of the leader, and restrict the control input to the interval.

DOI： 10.2514/1.46139

Language：English  

Language：English  

Language：English  

Language：English  

In this paper, leader-follower formation and reconfiguration problems based on the periodic orbits of the nonlinear relative dynamics along a circular orbit are considered. First, initial conditions of coplanar and noncoplanar relative orbits are characterized by the initial true anomaly, mean motion, semimajor axis, eccentricity, and inclination angle of the follower's inertial orbit. Based on the property of null controllability with vanishing energy of the Hill-Clohessy-Wiltshire equations, L₁ suboptimal feedback controllers are designed via the linear quadratic regulator theory. Simulation results for three examples are given. A comparison with the reconfiguration problem based on the periodic orbits of the Hill-Clohessy-Wiltshire equations shows that replacement by nonlinear periodic orbits does not increase the L ₁-norm of the feedback control.

DOI： 10.2514/1.41438

Repository Public URL： http://hdl.handle.net/2324/4476142

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Repository Public URL： http://hdl.handle.net/2324/4479053

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Event date： 2023.6

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Event date： 2023.6

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Event date： 2023.6

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Event date： 2023.1

Language：English   Presentation type：Oral presentation (invited, special)  

Country：India  

Event date： 2022.10

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Event date： 2022.9

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Event date： 2022.9

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Event date： 2022.9

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Event date： 2022.8

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Event date： 2022.2

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Event date： 2022.2

Language：English   Presentation type：Oral presentation (general)  

Country：Japan  

Event date： 2022.2

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Country：Japan  

Event date： 2021.10

Language：English  

Country：Japan  

Event date： 2021.10

Language：English  

Country：Japan  

Event date： 2021.9

Language：English  

Country：Japan  

Event date： 2021.8

Language：English  

Country：Japan  

Event date： 2021.8

Language：English   Presentation type：Oral presentation (general)  

Country：Other  

Event date： 2021.6

Language：Japanese   Presentation type：Oral presentation (general)  

Country：Japan  

Event date： 2021.2

Language：Japanese  

Country：Other  

Event date： 2021.2

Language：Japanese  

Country：Other  

Event date： 2021.2

Language：Japanese  

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Event date： 2020.10

Language：Japanese  

Country：Other  

Event date： 2020.10

Language：Japanese  

Country：Other  

Repository Public URL： http://hdl.handle.net/2324/4479733

Event date： 2020.10

Language：Japanese  

Country：Other  

Event date： 2020.10

Language：Japanese  

Country：Other  

Event date： 2020.9

Language：Japanese  

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Event date： 2020.8

Language：Japanese  

Country：Other  

Event date： 2020.1

Language：Japanese  

Country：United States  

Event date： 2020.1

Language：Japanese  

Country：Japan  

Event date： 2019.10

Language：English  

Venue：Washington   Country：United States  

Many of the equations of motion appearing in the aerospace field are nonlinear, and the problem of input optimization under this equation of motion is important. There are two methods for solving the nonlinear optimal control problem: direct method and indirect method. In the direct method, the equation of motion is discretized, and the problem is solved as a nonlinear programming problem with motion equations as constraints. In the direct method, it is possible to solve the problem by adding various constraints, but the solution becomes complicated and it is difficult to guarantee the convergence to the optimal solution. In this study, we consider a set of unknowns that satisfy constraints as Riemannian manifolds, and treat the problem as an unconstrained optimization problem on Riemannian manifolds. A simplified rocket trajectory optimization problem illustrates the proposed method.

Event date： 2019.10

Language：English  

Venue：Washington   Country：United States  

This paper investigated the manifold-based robust stationkeeping approach for the halo orbit near the Earth-Moon L2 point in the ephemeris model. The well-known Hamiltonian structure-preserved (HSP) method is utilized to design the controller for stationkeeping. However, conventional HSP method is proposed under the ideal condition, where the equations of motion are constructed in the circular restricted three-body problem, and the navigational uncertainty is not considered as well. Actually, these uncertainties always exist and have significant influence on the performance of stationkeeping. To make the results more realistic, the dynamics in the ephemeris model is used to describe the motion of a spacecraft. A filtering structure-based navigation is incorporated in the stationkeeping, where the navigational uncertainty is introduced. In addition, the navigation filter is reformed to compensate for the control uncertainty. Simulation is performed to observe the influence of the control gain and the magnitude of the uncertainty on the stationkeeping performance.

Event date： 2019.10

Language：English  

Venue：Washington   Country：United States  

This paper considers manifold-to-manifold transfers in the circular-restricted three-body problem enabled by low-thrust acceleration where an initial and target states lie on invariant manifolds associated to libration point orbits with different Jacobi constant. The basic idea is to utilize a family of stable and center manifolds that lie arbitrarily close to the target invariant manifold to reduce the cost of transfer. The linear quadratic regulator is used to design feedback control to transfer to the target manifold. Time invariant and time periodic controlleres are derived based on the linearized motion around the equilibrium point and periodic orbit respectively. The results show that the feedback controller can shape the linearized motion around manifold to be that around the equilibrium point or a periodic orbit. As a demonstration, transfer trajectories are designed to target the unstable manifold associated with an unstable Lyapunov orbit in the Earth-Moon system.

Event date： 2019.9

Language：Japanese  

Country：Japan  

Event date： 2019.6

Language：Japanese  

Country：Japan  

Event date： 2019.6

Language：Japanese  

Country：Japan  

Event date： 2019.6

Language：Japanese  

Country：Japan  

Event date： 2019.6

Language：Japanese  

Country：Japan  

Event date： 2019.1

Language：English  

Venue：Maui   Country：United States  

This paper generalizes the invariant manifolds of unstable libration point orbits through the application of continuous thrust. Considering Jacobi constant of the end of invariant manifolds, an artificial periodic orbit around a libration point realizes heteroclinic connections between itself and an unforced periodic orbit with same Jacobi constant of the end of invariant manifolds. Heteroclinic connections between libration point orbits are constructed by detecting intersections of states of manifolds on the Poincaré map. We reveal low-energy spacecraft can transfer to some periodic orbits with different Jacobi constant. In addition, this paper defines New Jacobi constant of low-thrust spacecraft. By utilizing new Jacobi constant, we illustrate zero-velocity curves of low-thrust spacecraft and reveal that there is a crack of zero-velocity curves and spacecraft can pass the crack.

Event date： 2019.1

Language：English  

Venue：Maui   Country：United States  

This paper proposes the controllability analysis of relative motion using different configurations for hybrid control actions of space environmental forces. It tends to present a minimal configuration for the missions exploiting these forces. It illustrates the constraints in each space environmental force and analyzes the integration between these forces to achieve full controllability of the satellite formation flight for near-circular low earth orbits with different orbit configurations. The paper implements a Kalman decomposition approach to decompose the system into controllable and uncontrollable subspaces for linear time-invariant control actions. Numerical simulations are investigated to study the controllability analysis for nonlinear models and substantiate the success of the controllability study for all hybrid control actions with different orbit configurations.

Event date： 2019.1

Language：English  

Venue：Maui   Country：United States  

This paper proposes a new hybrid control action of differentials aerodynamic drag and thrusters to control satellite formation flying in low Earth orbits. Parameterized output regulation algorithm for formation flying missions is developed based on the Schweighart-Sedwick relative dynamics equations. It is implemented to precisely track the different trajectories of reference relative motion and eliminates the effects of the J₂ perturbations. Parametric Lyapunov algebraic equation is derived to ensure the stability of the linear relative model subject to saturated inputs. The main goal of this study is to approve the viability of using the differentials in aerodynamic drag to precisely control different formation flying missions. Numerical simulations using a high fidelity relative dynamics model are implemented to analyze the performance of the proposed control algorithm in comparison with the linear quadratic regulator algorithm for actual satellite parameters. The paper exploits a high-precision orbit propagator model to verify the robustness of the control algorithm.

Event date： 2019.1

Language：English  

Venue：Maui   Country：United States  

A structure of the optimal trajectory for minimizing fuel consumption in an unstable dynamical environment such as the three-body problem is not well studied. Recently, it has been found that a sparse solution structure appears in the optimal control of a dynamical system. The concept of sparsity explains the property that the minimum fuel trajectory corresponds to the trajectory which minimizes the total thrusting time. In this paper, we propose a numerical method to obtain the minimum fuel sparse optimal trajectory in the unstable dynamical system. As an example, proposed methods are applied to the transfer in the Sun-Earth system.

Event date： 2018.10

Language：English  

Venue：Bremen   Country：Germany  

In trajectory design of a spacecraft, reducing the fuel consumption is important factor. In order to find the minimum fuel trajectory, this paper considers the L
¹
norm as the performance index. This paper discuss design parameters characterizing the L
¹
optimal trajectory, such as sampling time, flight time and magnitude of thrust. Also, this paper compare with the conventional method from the viewpoint of fuel consumption. In addition, this paper consider the rendezvous problem considering safety.

Event date： 2018.10

Language：English  

Venue：Bremen   Country：Germany  

This paper presents novel methods to provide libration point orbits (LPOs) and invariant manifolds (IMs) associated with libration point orbits in the elliptic restricted three-body problem (ERTBP). First, we introduce an extended autonomous system to describe the ERTBP. By introducing an autonomous system which generates a periodic function and integrating it with the ERTBP which is non-autonomous, we can obtain the extended autonomous system. Then, the center manifold design method which can uniquely provide an LPO in the circular restricted three-body problem is applied to the ERTBP. Next, we present a design method for IMs in the ERTBP by using the local stability characteristics obtained by the state transition matrix along LPOs of the extended system. The proposed design methods are applied and verified for the Earth-Moon L
₁
ERTBP. We show various LPOs in the ERTBP by changing design parameters, and IMs associated with a libration point and an LPO.

Event date： 2018.10

Language：English  

Venue：Bremen   Country：Germany  

In this study, a concept of optimality is introduced into Port-Hamilton systems for attitude control of spacecraft. Port-Hamiltonian systems make it possible to design asymptotically stable controllers in a uniform procedure for a wide variety of physical systems. By introducing optimality to Port-Hamilton system, conventional error systems of Port-Hamilton systems are expanded to minimize a quadratic form of evaluation function. In this study, as an optimal control method of Port-Hamiltonian system, Hamilton-Jacobi-Bellman (HJB) equation is considered. Since solving HJB equation is not easy, the equation is simplified through generalized canonical transformation, which is a unique conversion of Port-Hamiltonian system. Furthermore, by considering the minimum evaluation function as a Hamiltonian transformed through the generalized canonical transformation, the analytical solution of the HJB equation can be derived. This method can be used for a time-varying error system and applied for tracking control of spacecraft to specified trajectories. The optimality of the control input obtained from the proposed procedure is verified in numerical simulations.

Event date： 2018.5

Language：English  

Venue：Changsha   Country：China  

This paper explores the influence of unstable dynamics on orbit determination. A halo orbit about the Earth-Moon L2 is considered to be the nominal orbit and the dynamics model is constructed based on the Earth-Moon circular restricted three body problem. X-ray pulsar navigation method is introduced for orbit determination that is performed by Kalman filter. Simulation results show that unstable manifold of halo orbit dominates the growth of orbit-determination error and the corresponding uncertainty. However, when the orbit-determination measurements are incorporated in, the results change and are greatly influenced by the characteristic of the X-ray pulse measurements.

Event date： 2017.9

Language：English  

Venue：Adelaide   Country：Australia  

In this paper, we extend the flow function method for finding resonant orbits. The relation between flow function, resonance number and order of the resonant orbit are revealed. The resonant orbits design in the Sun-Earth CRTBP are demonstrated and the results show that the resonant orbits can be found globally and systematically by the flow function method. The flow function method can also be applied to obtain higher order resonant orbits. As an application, we propose a transfer trajectory to the L₄/L₅ SPOs by using various resonant orbits. Our results reveals that two types of insertion is possible for SPO insertion with small ΔV.

Event date： 2017.9

Language：English  

Venue：Adelaide   Country：Australia  

This paper proposes a novel optic flow-based navigation system for planetary rovers. Optic flow is a vector field of relative velocities between the camera mounted on a rover and environments, and it is often utilized for motion estimations or relative distance estimations. However, only either one of them can be estimated in general optic flow processing. The method proposed in this paper enables both of the estimations by utilizing an image segmentation technique in computer vision and applying Wide-Field-Integration (WFI) of optic flow for robust estimations. Thus, both obstacle avoidance and avoidance of getting stuck, which are key technologies for planetary exploration rovers, are realized in the method. Furthermore, since the estimation is accomplished in linear processing, a real-time estimation is possible by onboard computers. The effectiveness of the proposed method is examined through numerical simulations, considering sensor noises and the shape of obstacles.

Event date： 2017.9

Language：English  

Venue：Adelaide   Country：Australia  

This paper presents a novel method to realize station-keeping and control attitude motion simultaneously on a libration point orbit based on the nonlinear output regulation. First, the coupled orbit-attitude motion of a rigid spacecraft subject to three-body dynamics is formulated using the quaternion. Next, the reference orbit and attitude are represented as the output of an autonomous system called exosystem, assuming the reference orbit and attitude are given by a truncated Fourier series. Then the controller is derived by solving the output regulation problem. The derived controller has the explicit form and can achieve the asymptotically tracking for any reference orbit and attitude represented in the Fourier series. The proposed controllers are applied and verified for the Sun-pointing attitude control and station-keeping along Sun-Earth L₂ Halo orbits.

Event date： 2017.9

Language：English  

Venue：Adelaide   Country：Australia  

This paper proposes a new method of optimal trajectory design for formation flying along an elliptical orbit. Under linearized assumptions and a quadratic performance index, the optimal cost is quadratic in the initial state. Attractive sets of optimal control are defined as contours of the optimal cost based on linear quadratic regulator theory. It describes a set of all initial states to reach a desired state by a given cost. By solving the optimal control problem for TH equations, the optimal cost is obtained as a time-periodic function. This paper develops the attractive set for a time-periodic system and procedure to draw the attractive set is shown. The advantage of using attractive sets for optimal trajectory design is that it can determine the optimal initial state immediately. The various shape of the periodic orbit and the attractive sets are demonstrated by weight parameters of optimal control theory.

Event date： 2017.9

Language：English  

Venue：Kanazawa   Country：Japan  

In previous research, a controller design procedure via generalized canonical transformations has been proposed to keep the passivity feature of Port-controlled Hamiltonian systems. The procedure has a general form and is applicable to spacecraft's attitude motion described with quaternion parameters. This paper investigates the roles of several design parameters in the generalized canonical transformations. Special attention is placed on the relation between the shape of Lyapunov functions and the convergence speed of the state variables. Furthermore, from the analysis utilizing the linearized form of the Port-controlled Hamiltonian system, the guideline to decide the preferable ratio between two parameters in the design procedure is proposed.

Event date： 2017.6

Language：English  

Venue：Boulder   Country：United States  

Event date： 2017.6

Language：English   Presentation type：Oral presentation (general)  

Venue：Matsuyama   Country：Japan  

Event date： 2017.6

Language：English   Presentation type：Oral presentation (general)  

Venue：Matsuyama   Country：Japan  

Event date： 2016.9

Language：English  

Venue：Guadalajara   Country：Mexico  

The swing-by maneuver is known as a method to change the velocity of a spacecraft by using the gravity force of the celestial body. The powered swing-by has been proposed and researched to enhance the velocity change during the swing-by maneuver, e.g. Prado (1996). The research reports that applying an impulse maneuver at periapsis maximizes the additional effect to the swing-by. However, such impulsive force requires additional propellant. On the other hand, Williams et al. (2003) researched to use a tether cutting maneuver for a planetary capture technique. This current paper studies another way of the powered swing-by using tether cutting, which does not require additional propellant consumption. A Tethered-satellite is composed of a mother satellite, a subsatellite and a tether connecting two satellites. In a swing-by trajectory, a gravity gradient force varies according to the position of the tethered-satellite, and consequently its attitude motion is induced; the tethered-satellite starts to liberate and rotate. Cutting the tether during the tethered-satellites' rotation can add the rotational energy into the orbital energy. In this research, since the gravity gradient effect on orbital motion is small, the orbit can be considered a hyperbolic orbit. Assuming that the tether length is constant, Eq. (1) describes the equation of the attitude motion of a tethered-satellite within the SOI (sphere of influence) of the secondary body. θ = 1/1+ e cos α {2e (θ' + 1) sin α - 3/2 sin 2θ} where e is an orbit eccentricity, α is a true anomaly, θ is an attitude angle and l is tether length. The prime means the derivative with α. The mother satellite and subsatellite can obtain not only the velocity change but the position change by the tether cutting; they are denoted as Δv and Δr, and described in Eqs. (2) and (3), respectively. Δv = -l₁(α + θ) [-sin (α + θ) cos (α + θ)] Δr = -l₁ [cos (α + θ) sin (α + θ)] where l₁ is a distance between the center of gravity of the tethered-satellite and the mother satellite. Since α and θ are functions of time, Δv and Δr are also functions of time. This means that changing the tether cutting point can maximize the velocity change in this proposed powered swing-by maneuver. Furthermore, the optimum cutting point depends on the attitude and angular velocity when the tethered satellite enters the SOI. We propose a systematic design procedure to obtain the desired velocity change by optimizing the cutting point, the initial attitude and the initial angular velocity of the tethered satellites.

Event date： 2016.9

Language：English  

Venue：Guadalajara   Country：Mexico  

Invariant manifolds such as stable and unstable manifolds are often used in trajectory designs in the circular restricted three-body problem (CRTBP). Focusing on center manifolds, this paper proposes a simple but powerful trajectory design method for a periodic orbit based on the center manifold theorem. In general, there exist three types of invariant manifolds around equilibrium points: stable, unstable and center manifolds. The trajectories on center manifolds are bounded, whereas those on stable or unstable manifolds exponentially approach or diverge from the libration point as time passes. Hence, periodic orbits can be considered as systems lying on center manifolds. A general system around an equilibrium point is given by (Equation presented) where f, g and h are nonlinear terms and all real-parts of eigenvalues of matrices C, P and Q are zero, negative and positive, respectively. According to the center manifold theorem, in the neighborhood of the equilibrium point, the system (1) can be represented by (Equation presented) The first ordinary differential equation is called reduced order system. The solution to Eq. (2) represents solutions restricted on center manifolds. Based on Eq. (2), we propose a novel method to obtain periodic orbit. It includes two steps. The first step computes approximate solutions for the system shown as Eq. (2) in the CRTBP. This step requires iteration by using a successive approximation method proposed by Suzuki et al., which requires initial values of the reduced order system, ξ ∈ R
⁴
. Note that since in the first step quasi-periodic orbits are obtained, a further modification of ξ is necessary to obtain pure periodic orbits. The second step corrects ξ to form a periodic orbit. The required variation Δξ are computed by using the state transition matrix. The remarkable feature of the proposed method is that once we fix three components of ξ, the method provides the initial state for a periodic orbit uniquely by successive iteration mentioned above. Note that two of them become zero from symmetries of periodic orbits and the rest can be set arbitrarily. Contrary to the proposed method, conventional methods such as the differential correction method and the Lindstedt-Poincaré method require good initial guesses and complex algebraic manipulations. The proposed method is applicable to any equilibrium point, although as an example the proposed method is applied to the Sun-Earth L
₂
point and verified in numerical simulation. In conclusion, we reveal that the proposed method is a very powerful tool and can significantly reduce efforts to obtain periodic orbits.

Event date： 2016.9

Language：English  

Venue：Long Beach   Country：United States  

Trajectory design combining low-thrust and the three-body problem is a challenging area for study. This paper investigates the combination of optimal feedback control with the dynamical structure of the three-body problem. For the nonlinear system, the attractive set of an equilibrium point or a periodic orbit under optimal control is described by the Hamilton-Jacobi-Bellman partial differential equation. We obtain the solution to the Hamilton-Jacobi-Bellman equation by solving the truncation problem successively.

Event date： 2016.2

Language：English  

Venue：Napa   Country：United States  

This paper investigates the combination of optimal feedback control with the dynamical structure of the three-body problem. The results provide new insights for the design of continuous low-thrust spacecraft trajectories. Specifically, we solve for the attracting set of an equilibrium point under optimal control with quadratic cost. The analysis reveals the relation between the attractive set and original dynamics. In particular, we find that the asymptotic form of the attractive set to an equilibrium point or a fixed point under optimal control is completely defined by its left unstable eigenvectors and a term inversely proportional to its unstable eigenvalue.

Event date： 2015.10

Language：English  

Venue：Jerusalem   Country：Israel  

This paper studies a control law to stabilize the orbital motion in the vicinity of an unstable equilibrium point and periodic orbits in the circular-restricted three-body problem (CR3BP). Since the libration point orbits especially around collinear points are highly unstable, time-delay in sensor and/or acutor might cause instability. In this paper, we derive a continuous sliding mode controller (Cont. SMC) to achieve robust fuel-efficient periodic motion based upon a new Lyapunov function to smooth the control input. Then, the total velocity change is investigated to reach and maintain a halo orbit in the Sun-Earth CR3BP. Simulation results show that the proposed Cont. SMC has good performance and robustness in the presence of unmodeled time-delay.

Event date： 2015.8

Language：English  

Venue：Vail   Country：United States  

This paper studies a control law to stabilize the orbital motion in the vicinity of an unstable equilibrium points and periodic orbits in the circular-restricted three-body problem. Utilizing the eigenstructure of the system, the fuel efficient formation flying controller via linear quadratic regulator (LQR) is developed. Then the chattering attenuation sliding mode controller (CASMC) is designed and analyzed for the in-plane motion of the circular circular-restricted three-body problem. Simulation studies are conducted for the Sun-Earth L₂ point and a halo orbit around it. The total velocity change required to reach the halo orbit as well as to maintain the halo orbit is calculated. Simulation results show that the chattering attenuation sliding mode controller has good performance and robustness in the presence of unmodeled nonlinearity along the halo orbit with relatively small fuel consumption.

Event date： 2014.9 - 2014.10

Language：English  

Venue：Toronto   Country：Canada  

Space Situational Awareness (SSA) has been recognized to be important for the safe space activities including these problems. As low-thrust propulsion technology becomes increasingly popular, SSA for low-thrust spacecraft may become an area of increasing interest. More frequently use of low-thrust propulsion to place satellites in orbit create more opportunities for collisions and radio frequency interference as these spacecraft travel slowly through altitude ranges. The purpose of this paper is to develop a method for estimation of the mean orbital elements for low-thrust spacecraft. To overcome the instability of the estimation problem with low-thrust acceleration, we estimate the mean elements instead of osculating elements. By use of the averaging technique, Hudson and Scheeres (2009) proposed an analytical model of secular variations of orbital elements under thrust acceleration. The resulting averaged equation has a nice property in which only finite number of Fourier coefficients of the thrust acceleration appear because of the orthogonality of the trigonometric function. Based on the nonlinear state equation representation for the extended state variable which include not only orbital elements but also unknown Fourier coefficients, mean orbital elements and thrust history are estimated from observation data of mean orbital elements. Moreover, the mapping from mean to osculating elements is derived which can be replaced by a measurement equation. Proposed method is demonstrated through numerical simulations

Event date： 2014.9 - 2014.10

Language：English  

Venue：Toronto   Country：Canada  

This study discusses the condition of Wide-Field-Integration (WFI) of optic flow to make it as a realistic guidance and navigation system for space probes. WFI of optic flow is a bio-inspired navigation system mimicking visual processing system of flying insects using their compound eyes. However, in a real system, photoreceptors on a spherical surface used in WFI of optic flow are replaced by pixels on flat image surfaces of cameras, and the camera's field of view is limited. This paper examines the effect of the realistic restrictions for WFI of optic flows on the estimation performance in numerical simulations. The estimation accuracy is evaluated for several parameters: the field of view and the resolution of a camera, which is used instead of photoreceptors. Effects of space probe motion and sensor noises are also evaluated. It is shown that a camera with a wider field of view shows better accuracy, when the total number of pixels (or photoreceptors) is same.

Event date： 2014.3

Language：English  

Venue：Rome   Country：Italy  

In this paper, leader-follower formation flying problems based on the periodic orbits of the Hill-Clohessy-Wiltshire equations and Tschauner-Hempel equations are considered. the control input is assumed to be impulsive. For a given final relative orbit, the admissible controls are feedback controls such that the follower tracks the final orbit asymptotically. The main performance index is the ΔV which is proportional to the fuel consumption. The number of impulses is fixed, but the impulse time is arbitrary and varied. The optimal location of impulse times is sought numerically. Simulation results for a reconfiguration problem are given.

Event date： 2014.1

Language：English  

Venue：Mexico   Country：United States  

In this paper, an approach to optimize deflection mission of Near Earth asteroids under uncertainty is proposed based on control theoretical framework. By using the nonlinear mapping from initial deviation to the final deviation, the performance index J which is expressed as a function of the distance of the Earth in the Earth's closest approach time is minimized for the worst case initial condition. We also formulate the robust orbital transfer problem of kinetic impactor spacecraft where the initial state estimation error exists.

Event date： 2013.9

Language：English  

Venue：Nagoya   Country：Japan  

The relative motion of a follower satellite with respect to the leader in a given circular orbit is described by autonomous nonlinear differential equations. The linearized equations around the null solution are known as Hill- Clohessy-Wiltshire (HCW) equations. In this paper, active formation flying for the HCW system with pulse (impulse) control input is considered, where the desired relative orbit of the follower is generated by an exosystem. This allows for flexibility of the shape and period of the reference orbit. The output regulation theory for discrete linear system is employed. To show the effectiveness of this approach, numerical examples are given.

Event date： 2013.9

Language：English  

Venue：Nagoya   Country：Japan  

In this note, we propose a new driving unit for a rover on small gravity asteroid MINERVA-II. MINERVA-II will be carried by Hayabusa-2 which will be carried by H-IIA rocket launched in 2014. In designing the rover, the most difficult point is lack of information of the target asteroid. To overcome this diffilulty, we propose a kind of passive type driving system. Concretely, we try to apply buckling phenomenon and use spring. As the result, very simple driving unit for unknown field will be developed.

Event date： 2013.8

Language：English  

Venue：Hilton Head Island, SC   Country：United States  

The linearized equations of relative motion of a follower satellite with respect to the leader in a given elliptic orbit are described by Tschauner-Hempel (TH) equations. In this paper, formation flying for the TH system with pulse control input is considered. First, the reconfiguration problem where the desired relative orbit of the follower is natural periodic orbit. Next, the active formation problem where the desired relative orbit is generated by an exosystem is considered. This allows for flexibility of the shape and period of the reference orbit. To realize such a formation flying, the output regulation theory for linear periodic systems is employed. To achieve asymptotic tracking, stabilizing feedback controls are designed by the periodic solution of discrete-time Riccati equation of the linear quadratic regulator (LQR) theory.

Event date： 2012.5

Language：English  

Venue：Kyoto   Country：Japan  

The trajectory design of the interplanetary mission "Demonstration and Experiment of Space Technology for INterplanetary voYage, DESTINY" is discussed. The trajectory optimization of low-thrust spacecraft pose a difficult design challenge. Moreover, DESTINY mission requires many constraints for the orbital design. In advance of optimizing the whole transfer mission, we investigated the basic theory to design which can take into account such constraints.

Event date： 2012.3

Language：English  

Venue：Kowloon  

We consider a plug-in control system via a concept of Implicit and Explicit Controls as the two stage controllers. We first show two conditions under which the plug-in control system can be optimized by a suitable design of explicit control in the 2nd stage, when an implicit control is given in the 1st stage. We then show another condition under which the resultant optimal cost can be minimized by a suitable design of the Is' stage implicit control, which is characterized as a feedback control that allocates all the closed-loop poles onto the imaginary axis. These two results clarify the system theoretic meaning of the implicit control from the viewpoint of inverse optimal control problem.

Event date： 2011.9

Language：English  

Venue：Tokyo   Country：Japan  

In this note, we consider a plug-in control system via a concept of Implicit Control and Explicit Control. We show that the plug-in control system can be optimized by a certain plug-in control law. And also we show that a property of integrity can be embedded in the optimal plug-in control system.

Event date： 2011.9

Language：English  

Venue：Tokyo   Country：Japan  

In this paper, we consider the motion control of spacecraft in the vicinity of the peanut-shaped asteroid. The implicit control concept is studied and orbital maintenance strategy to observe the peanut-shaped asteroid where the model of the gravity field is difficult to obtain is proposed. The implicit control law is characterized in terms of inverse problem of optimal control.

Event date： 2011.7 - 2011.8

Language：English  

Venue：Girdwood, AK   Country：United States  

In this paper, leader-follower formation flying problems based on the periodic orbits of the Hill-Clohessy-Wiltshire equations are considered. For a given final relative orbit, the admissible controls are feedback controls such that the follower tracks the final orbit asymptotically. The main performance index is the L₁-norm of the control input which is proportional to the fuel consumption. The L_∞-norm, which gives the magnitude of the input, is an additional performance index under the thruster level limitation. Feedback controls are designed via two algebraic Riccati equations with parameters, and their L₁- and L_∞- norms are given as functions of the parameters. Suboptimal controls are designed by choosing appropriate parameters. Simulation results for a reconfiguration problem are given to demonstrate the effectiveness of the design method.

Event date： 2010.9 - 2010.10

Language：English  

Venue：Prague   Country：Czech Republic  

This paper presents a procedure to design multiple rendezvous missions. We propose a new low-thrust trajectory optimization method to find the optimal trajectory in analytical form. The resulting control law is a feedback control, so that it is effective in overcoming sensitivities to small variations in initial orbit and thrust profile. This study determines the optimal approach to find an effective control law to sequential transfer problem. For comparison, optimal sequence using impulsive and continuous thrust method are investigated. Our approaches are illustrated through the trajectory design of the multiple flyby mission and resulting fuel costs are compared.

Event date： 2010.9 - 2010.10

Language：English  

Venue：Prague   Country：Czech Republic  

Orbiting artificial satellites naturally tend to accumulate electrostatic charge which may in some circumstances cause malfunctions in measuring instruments and other devices on the satellite body and ultimately lead to difficulties in satellite operation. In this paper, we investigate the possibility of the use of the Lorentz force, which acts on charged satellite when it is moving through the magnetic field, as a means of satellite attitude control. We first analyze the stability of the motion and then propose novel attitude control method using the Lorentz force.

Event date： 2010.9

Language：English  

Venue：Nara   Country：Japan  

In this paper, a formation problem of a spacecraft with a leader in an elliptic orbit is considered, where the follower is required to track asymptotically a given periodic relative orbit. The linearized equations of relative motion along an eccentric orbit are known as Tschauner-Hempel equations. The Tschauner-Hempel equations are transformed into simpler ones by the change of variables. For the transformed system, the active formation problem is formulated, and feedback controllers are designed by the regulator equations associated with the Hill-Clohessy-Wiltshire equations. Feedback controllers are also designed using a DRE associated with the TH equations, and the Li-norms of designed controllers are compared.

Event date： 2010.9

Language：English  

Venue：Nara   Country：Japan  

The aim of this paper is to investigate the dynamics of a restricted three-body problem governed by two types of forces: gravity force and Coulomb force. The electrostatic charging and hence the Coulomb force is introduced in order to change the relative position by positive control of the charge amount without using fuel expenditure.

Event date： 2010.8

Language：English  

Venue：Taipei   Country：Taiwan, Province of China  

The optimal control problem of a spacecraft using continuous thrust where the terminal state and time interval are explicitly given is considered. We consider a new robust controller to the spacecraft rendezvous problem. Optimal rendezvous problem using continuous low-thrust is treated as optimal control problem and H_∞ control. Using the property that both problems can be treated as two-point boundary value problem (TPVBP) of Hamiltonian system, we formulate the linear quadratic control with worst case disturbance rejection and worst case initial condition related by generating functions. Numerical simulations are given to illustrate the theory.

Event date： 2010.8

Language：English  

Venue：Toronto, ON   Country：Canada  

The motion of a charged satellite subjected to the Earth's magnetic field is considered. Lorentz force, which acts on charged particle when it is moving through the magnetic field, provides a new concepts of propellant-less electromagnetic propulsion. We derive a dynamical model of a charged spacecraft including the effect of Lorentz force in the vicinity of circular and elliptic orbit and consider its application to formation flight. Based on Hill-Clohessy-Wiltshire equations and Tschauner-Hempel equations, analytical approximations for the relative motion in Earth orbit are obtained. The analysis based on linearized equations in the equatorial case show that the stability of the periodic solutions. Numerical simulations revealed that the periodic solutions which are useful for the formation are obtained for both circular and elliptic reference orbits.The control strategy for the propellantless rendezvous and reconfiguration are developed.

Event date： 2010.8

Language：English  

Venue：Toronto, ON   Country：Canada  

This paper is concerned with formation flying along a circular orbit. The control input is of pulse type and the performance index is the total velocity change required for formation acquisition. For the in-plane motion, optimal three-pulse strategies which use control acceleration only in the flight direction are given under the condition that the difference of orbit size is relatively larger than that of drift velocity. If the inputs are restricted to the radial direction, only formation reconfiguration is possible and optimal two-pulse strategies are given. For the out-of-plane motion, the existence of optimal single pulse strategies is shown. Formation flying with feedback control is also considered. For the in-plane motion, suboptimal feedback controllers are designed with the help of null controllability with vanishing energy of the Hill-Clohessy-Wiltshire equations. Feedback controls restricted to flight direction are also designed, and the total velocity changes are compared with those of full control. Simulation results for a formation reconfiguration problem are given both for optimal open loop strategies and feedback controls.

Event date： 2010.7

Language：English  

Venue：Montreal, QC   Country：Canada  

The purpose of this paper is to investigate the possibility of asteroid survey mission enabled by advanced solar sailing technology. The study is focused not on the solar sail spacecraft itself but on its orbital dynamics to realize the missions. A novel NEA flyby survey mission with a light-weight solar sail spacecraft to increase the accessibility to NEAs located in the vicinity of the Earth's orbit is proposed. Numerical study suggests that our approach increase the opportunities in proximity to NEAs which have eccentric and inclined orbits.

Event date： 2010.2

Language：English  

Venue：San Diego, CA   Country：United States  

In this paper, we consider the canonical transformations and its applications appearing in astrodynamic problems. First we address stabilization of relative motion via generalized canonical transformation and passivity-based control. Then we propose a method to solve Lambert's problem based on the Hamilton-Jacobi-Bellman (HJB) equation in optimal control theory. Using the generalized canonical transformation, we transform the performance index to positive- definite one and then solve the optimal control problem. We also apply our method to obtain solution to two-point boundary-value problem by the generating function. As an application of the generating functions approach, we consider the problem of multiple flyby mission with impulsive thrust.

Event date： 2009.8

Language：English  

Venue：Hong Kong   Country：China  

In this paper, a formation problem of a satellite with a leader in an eccentric orbit is considered, where the transition to a periodic relative orbit is required, and the control input is restricted to a subinterval of the period of the leader's orbit. As preliminaries, the same problem in the special case of the circular orbit is considered, and feedback controls based on a differential Riccati equation with a periodic control matrix are designed. The performance index is the L₁-norm of the feedback control, and suboptimal controls are obtained by virtue of the null controllability with vanishing energy of the controlled system. Then the original problem is transformed into a special form by the change of variables, and using the structure of the resulting system, feedback controls are designed based on those of the circular problem.

Event date： 2009.8

Language：English  

Venue：Pittsburgh, PA   Country：United States  

This paper considers halo orbit control for the Earth-moon elliptic restricted threebody problem. Expressing equations of motion with true anomaly, Lagrangian points are defined and a halo orbit control problem at the L ₂ point is discussed. By the change of control variables, constant feedback controllers are designed which maintain a halo orbit of the circular restricted problem. Considering equations of motion relative to the moon, and letting the mass of the moon go to zero, the equations of relative motion along an eccentric orbit are derived. Then formation and reconfiguration problems are formulated, and feedback controllers, based on the Hill-Clohessy-Wiltshire systems, are designed from the point of view of L₁-norm minimization.

Event date： 2009.2

Language：English  

Venue：Savannah, GA   Country：United States  

The optimal control problem of a spacecraft using impulsive and continuous thrust where the terminal state and time interval are explicitly given is considered. Using a recently developed technique based on Hamilton-Jacobi theory, we develop a method to approximate the solution of the Hamilton-Jacobi equation which can solve the two-point boundary-value problem. The proposed method is based on the successive approximation and Galerkin spectral method with Chebyshev polynomials. This approach is expected to derive the analytical solution of the optimal control problem in the large domain. Numerical simulation is given to illustrate the theory.

Event date： 2007.8

Language：English  

Venue：Saint Petersburg   Country：Russian Federation  

In this paper the output regulation problem for nonlinear systems described by multiple linear models with unknown parameters is considered. Based on the Lyapunov stability theory, an adaptive controller which stabilize the system is derived. Then sufficient conditions for the output regulation problem with full information to be solvable are established. Simulation results are given to illustrate the theory.

Event date： 2006.10

Language：English  

Venue：Busan   Country：Korea, Republic of  

In this paper, the output regulation problem for linear time-invariant systems with unknown parameters is considered. Based on Lyapunov stability theorem, an adaptive controller which achieves regulation is derived. Then, sufficient conditions for the existence of a controller that solves the output regulation problem with state feedback are established and simulation result are given.

Event date： 2004.8

Language：English  

Venue：Sapporo   Country：Japan  

A control system, which consists of several cooperative modules whose combination and structures change dynamically according to the state and environment, is discussed in this paper. We propose a method to design a control system by modular learning. Numerical simulations and flight experiment of an autonomous aero-robot demonstrate the effectiveness of the proposed method.

Type：Peer review 

Number of peer-reviewed articles in foreign language journals：9

Proceedings of International Conference Number of peer-reviewed papers：2

Type：Peer review 

Number of peer-reviewed articles in foreign language journals：10

Proceedings of International Conference Number of peer-reviewed papers：12

Type：Competition, symposium, etc. 

Number of participants：20

Type：Competition, symposium, etc. 

Type：Peer review 

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

Proceedings of International Conference Number of peer-reviewed papers：10

Type：Peer review 

Number of peer-reviewed articles in foreign language journals：6

Number of peer-reviewed articles in Japanese journals：2

Type：Competition, symposium, etc. 

Number of participants：60

Audience：Infants,　Schoolchildren,　Junior students,　High school students

Audience：Infants,　Schoolchildren,　Junior students,　High school students

Audience：Infants,　Schoolchildren,　Junior students,　High school students