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Wang, Yi; Chen, Xiaoqian; Ran, Dechao; Zhao, Yong; Chen, Yang; Bai, Yuzhu
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In this paper, an adaptive artificial potential function (AAPF) method is developed for spacecraft formation reconfiguration with multiobstacle avoidance under navigation and control uncertainties. Furthermore, an improved Linear Quadratic Regular (ILQR) is proposed to track the reference trajectory and a Lyapunovbased method is employed to demonstrate the stability of the overall closedloop system. Compared with the traditional APF method and the equalcollisionprobability surface (ECPS) method, the AAPF method not only retains the advantages of APF method and ECPS method, such as low computational complexity, simple analytical control law and easy analytical validation progress, but also proposes a new APF to solve multiobstacle avoidance problem considering the influence of the uncertainties. Moreover, the ILQR controller obtains high control accuracy to enhance the safe performance of the spacecraft formation reconfiguration. Finally, the effectiveness of the proposed AAPF method and the ILQR controller are verified by numerical simulations.
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By
Li, Kebo; Su, Wenshan; Chen, Lei
5 Citations
The performance of the threedimensional differential geometric guidance law with proportional navigation formation against a target maneuvering arbitrarily with timevarying normal acceleration is thoroughly analyzed using the Lyapunovlike approach. The validation of this guidance law is firstly proved, and then the performance issues such as capturability, heading error control efficiency, line of sight rate convergence, and commanded acceleration requirement are analyzed, under the condition that the missile is initially flying toward the target with a speed advantage. It is proved that an intercept can occur and the line of sight rate and missile commanded acceleration can be limited in certain ranges, if the initial heading error is small and the navigation gain is sufficiently large. The nonlinear relative dynamics between the missile and the target is taken into full account, and the analysis process is simple and intuitive, due to the use of a convenient line of sight rotating coordinate system. Finally, the new theoretical findings are validated by numerical simulations.
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Liu, Wei; Zhang, Yongkang; Li, Zongfeng; Dong, Wenbo
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1 Citations
The Microgravity Active vibration Isolation System (MAIS), which was onboard China’s first cargospacecraft Tianzhou1 launched on April 20, 2017, aims to provide highlevel microgravity at an order of 10^{5}–10^{6}g for specific scientific experiments. MAIS is mainly composed of a stator and a floater, and payloads are mounted on the floater. Sensing relative motion with respect to the stator fixed on the spacecraft, the floater is isolated from vibration on the stator via control forces and torques generated by electromagnetic actuators. This isolation results in a highlevel microgravity environment. Before MAIS was launched into space, its control performance had been simulated on computers and tested by airbearing platform levitation and aircraft parabolic flight. This article first presents an overview of the MAIS’s hardware system, particularly system structure, measurement sensors, and control actuators. Its system dynamics, state estimation, and control laws are then discussed, followed by the results of computer simulation and engineering tests, including the test of the sixdegreeoffreedom motion by aircraft parabolic flight. Simulation and test results verify the accuracy of the control strategy design, effectiveness of the control algorithms, and performance of the entire control system, paving the way for operation of MAIS in space. This article also presents the steps recommended for the control performance simulation and tests of MAISlike devices. These devices are expected to be used on China’s Space Station for various scientific experiments that require a highlevel microgravity environment.
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Zhao, Shuge; Zhang, Jingrui; Xiang, Kaiheng; Qi, Rui
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5 Citations
A method is proposed to select the target sequence for a J_{2}perturbed multiple debris rendezvous mission aimed at removing dozens of debris from several thousand debris candidates running on sunsynchronous orbits (SSO). The solving methodology proceeds in two steps: Firstly, the variance of the right ascension of ascending node (RAAN) of the debris group is used for narrowing down the potential debris candidate; secondly, the debris of the candidate group that has closest RAAN to the current debris is chosen as the next debris. The low thrust nearminimumfuel trajectories of each rendezvous leg are obtained by the indirect optimization method. The proposed approach is demonstrated for the problem of the 8th China Trajectory Optimization Competition (CTOC). The radar cross section (RCS) of the debris is also considered in the first step since the primary performance index of the competition is to maximize the total RCS of the debris visited. The results show that the proposed approach achieves better performance within a competition period. Of the many rendezvous sequences found, the best one submitted for the competition obtained a total RCS of 184 by accomplishing rendezvous with 70 debris within a transfer duration of one year.
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By
Yang, Chihang; Zhang, Hao
Laser Interferometer Space Antenna (LISA) is a project to detect and measure gravitational waves. The project has three spacecraft flying in a formation of near equilateral triangle in a heliocentric orbit trailing Earth. Many sources of perturbations cause the configuration to deviate from the nominal. This paper studies the formation design problem for a LISAlike mission by considering ephemerisbased dynamics. This type of mission is wellknown for addressing several strict mission requirements under the realistic dynamics. The problem is formulated as optimizing multiple mission performance indices. It is observed that some indices are correlated with each other, whereas some indices have different sensitivities with respect to the semimajor axis. Therefore, the problem is transformed into a twostep cascade singleobjective optimization, in which the optimal solution of the first optimization problem is fed to the second optimization as initial value. In addition, the major perturbing celestial bodies are picked up to make a simplified but accurate enough dynamics to speed up the optimization. Numerical examples verify the analysis and show the effectiveness of the optimization procedure. The influences of mission lifetime and spatial scales on the optimal solutions are also presented.
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By
Wang, Enmei; Wu, Shunan; Liu, Yufei; Wu, Zhigang; Liu, Xiangdong
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1 Citations
To deal with the vibration problem of the solar power satellite (SPS), the distributed vibration control approach is investigated in this paper. Taking the MultiRotary joints SPS as the research objective, the control unit (CU) and the location relationship matrix are firstly defined for distributed controller design according to the configuration of SPS. The dynamic model of each CU is therefore established based on the finite element method. The dynamic model of the whole SPS structure is then developed using the CU models, and is updated along with onorbit assembly. The distributed cooperative controller, using proportional and differential feedback and the interaction feedback among adjacent CUs, is proposed to suppress vibration. The closeloop distributed cooperative control system is then achieved by integrating all distributed controllers, and the asymptotic stability is proofed by the Lyapunov’s stability theorem. To verify the feasibility of the proposed control system, three numerical cases are finally presented. The results demonstrate that the distributed cooperative controllers can effectively suppress vibration during onorbit assembly and operation after assembly, and the closedloop system has good fault tolerance.
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By
Caruso, Andrea; Quarta, Alessandro A.; Mengali, Giovanni
3 Citations
This paper deals with the optimization of the transfer trajectory of a solar sailbased spacecraft between circular and coplanar heliocentric orbits. The problem is addressed using both a direct and an indirect approach, while an ideal and an optical force model are used to describe the propulsive acceleration of a flat solar sail. In the direct approach, the total flight time is partitioned into arcs of equal duration, within which the sail attitude is assumed to be constant with respect to an orbital reference frame, and a nonlinear programming solver is used to optimize the transfer trajectory. The aim of the paper is to compare the performance of the two (direct and indirect) approaches in term of optimal (minimum) flight time. In this context, the simulation results show that a direct transcription method using a small number of arcs is sufficient to obtain a good estimate of the global minimum flight time obtained through the classical calculus of variation.
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Bengtson, M.; Wilson, K.; Hughes, J.; Schaub, H.
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7 Citations
The number of operational satellites and debris objects in the valuable geosynchronous ring has increased steadily over time such that active debris removal missions are necessary to ensure longterm stability. These objects are very large and tumbling, making any mission scenarios requiring physical contact very challenging. In the last 10 years, the concept of using an electrostatic tractor has been investigated extensively. With the electrostatic tractor concept, active charge emission is employed to simultaneously charge the tug or services vehicle, while aiming the charge exhaust onto the passive space debris object to charge it as well. The resulting electrostatic force has been explored to actuate this debris object to a disposal orbit or to detumble the object, all without physical contact. This paper provides a survey of the related research and reviews the charging concepts, the associated electrostatic force and torque modeling, and the feedback control developments, as well as the charge sensing research.
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By
Quarta, Alessandro A.; Mengali, Giovanni; Niccolai, Lorenzo
Insitu measurements are necessary for a longterm analysis of the spatial structure of the geomagnetic tail. This type of mission requires the use of a propellantless propulsion system, such as a classical solar sail, to continuously rotate the design orbit apse line such that it remains parallel to the SunEarth direction. To reduce the mission costs, this paper suggests the employment of Sunpointing smart dusts, which are here investigated in terms of propulsive acceleration level necessary to guarantee a mission’s feasibility. A Sunpointing smart dust can be thought of as a millimeterscale solar sail, whose geometric configuration allows it to passively maintain an alignment with the Sunspacecraft line. The smart dust external surface is coated with an electrochromic reflective film in such a way that it may change, within some limits, its propulsive acceleration magnitude. A suitable control law is necessary for the smart dust to enable an artificial precession of its Earthcentred orbit, similar to what happens in the GeoSail mission. This paper analyzes the required control law using an optimal approach. In particular, the proposed mathematical model provides a set of approximate equations that allow a simple and effective tradeoff analysis between the propulsive requirements, in terms of the smart dust acceleration, and the characteristics of the design orbit.
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By
Tang, Gao; Hauser, Kris
Nonlinear optimal control problems are challenging to solve due to the prevalence of local minima that prevent convergence and/or optimality. This paper describes nearestneighbors optimal control (NNOC), a datadriven framework for nonlinear optimal control using indirect methods. It determines initial guesses for new problems with the help of precomputed solutions to similar problems, retrieved using knearest neighbors. A sensitivity analysis technique is introduced to linearly approximate the variation of solutions between new and precomputed problems based on their variation of parameters. Experiments show that NNOC can obtain the global optimal solution orders of magnitude faster than standard random restart methods, and sensitivity analysis can further reduce the solving time almost by half. Examples are shown on optimal control problems in vehicle control and agile satellite reorientation demonstrating that global optima can be determined with more than 99% reliability within time at the order of 10–100 milliseconds.
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