Adaptive Quasi Fixed-Time Orbit Control Around Asteroid with Performance Guarantees

This paper investigates a novel quasi fixed-time orbit tracking control method for spacecraft around an asteroid in the presence of uncertain dynamics and unknown uncertainties.To quantitatively characterize the transient and steady-state responses of orbit tracking error system,a continuous performance function is devised via using a quartic polynomial.Then,integrating backstepping control technique and barrier Lyapunov function leads to a quasi fixed-time convergent orbit tracking controller without using any fractional state information and symbolic functions.Finally,two groups of illustrative examples are employed to test the effectiveness of the proposed orbit control method.

Impulsive orbit control for spacecraft around asteroid

An impulse feedback control law to change the mean orbit elements of spacecraft around asteroid is presented. First, the mean orbit elements are transferred to the osculating orbit elements at the burning time. Then, the feedback control law based on Gauss’s perturbation equations of motion is given. And the impulse control for targeting from the higher circulation orbit to the specified periapsis is developed. Finally, the numerical simulation is performed and the simulation results show that the presented impulse control law is effective.

A Status Graph Based Control Allocation Algorithm of Digital Micro-Thruster Array for Micro/Nano-Satellites Orbit Control Application

Digital micro-thruster arrays can be used for special missions of micro/nano-satellites with the requirements of high precision and small impulse.This paper presents a novel control allocation algorithm for the digital micro-thruster array,namely status graph based control allocation(SGBCA)algorithm,which aims at finding the optimal micro thrusters combination scheme to realize the sequential control synthesis for micro/nano-satellite during real-time orbit control tasks.A mathematical model is set up for the control allocation of this multivariate over-actuated system.Through dividing thrusters into disjoint segments by offline calculation and combining segments dynamically online to provide a sequence of the required impulse for the micro/nano-satellite,the time complexity of the control allocation algorithm decreases significantly.All levels of impulse can be generated by the digital micro thruster arrays and the service life of the arrays can be extended using the segment converting strategy proposed in this paper.The simulation indicates that the algorithm can satisfy the requirements of real-time orbit control for micro/nano-satellites.

Chang’E-2 satellite asymmetric-descent orbit control technology

To accomplish high-resolution imaging of the preselected landing area,it was necessary for the Chang’E-2 mission to perform orbital maneuvering on the far side of the moon to meet the conditional height requirement of the imaging area.Engine shut-down would be executed invisibly on the back side of the moon if the descent maneuver mode opposite to the target perilune or the fuel optimal maneuver mode was used.To ensure the satellite safety,the project collectivety required that the engine shutdown should be designed to be executed in the domestic segmental arcs and meet the requirement of satellite emergency treatment simultaneously.Accordingly,the asymmetric-descent orbit control technology was adopted by offsetting the ma-neuver point,which obtained the orbit control parameters of finite-thrust mode with an iteration algorithm and modified the results with target perilune drift estimation.The Chang’E-2 satellite declined to the target of 100 km×15 km orbit successfully on 26 October 2010,and has been flying for 32 circles in the experimental orbit to accomplish the preselected landing area imaging.This paper describes the mechanism and realization method of the asymmetric-descent orbit control technology and evaluates the maneuver effect with the actual mission data.

Controlof Orbitand Controlof Chaosina Class of Dynamic System

The problem of control of orbit for the dynamic system x ¨+x(1-x)(x-a)=0 is discussed. Any unbounded orbit of the dynamic system can be controlled to become a bounded periodic orbit by adding a periodic step excitation to the system. By using a nonlinear feedback control law presented in this paper the chaos of the dynamic system with excitation and damping is stabilized. This method is more effectual than the linear feedback control.

Adaptive control for autonomous rendezvous of spacecraft on elliptical orbit

A strategy for spacecraft autonomous rendezvous on an elliptical orbit in situation of no orbit information is developed. Lawden equation is used to describe relative motion of two spacecraft. Then an adaptive gain factor is introduced, and an adaptive control law for auton- omous rendezvous on the elliptical orbit is designed using Lyapunov approach. The relative motion is proved to be ultimately bounded under this control law, and the final relative position error can achieve the expected magnitude. Simulation results indicate that the adaptive control law can realize autonomous rendezvous on the elliptical orbit with relative state information only.

Orbit and attitude control of spacecraft formation flying

Formation flying is a novel concept of distributing the functionality of large spacecraft among several smaller, less expensive, cooperative satellites. Some applications require that a controllable satellite keeps relative position and attitude to observe a specific surface of another satellite among the cluster. Specially, the target space vehicle is malfunctioning. The present paper focuses on the problem that how to control a chaser satellite to fly around an out-of-work target satellite closely in earth orbit and to track a specific surface. Relative attitude and first approximate relative orbital dynamics equations are presented. Control strategy is derived based on feedback linearization and Lyapunov theory of stability. Further, considering the uncertainty of inertia, an adaptive control method is developed to obtain the correct inertial ratio. The numerical simulation is given to verify the validity of proposed control scheme.

The present status and prospects in the research of orbital dynamics and control near small celestial bodies

Small celestial body exploration is of great significance to deep space activities. The dynamics and control of orbits around small celestial bodies is of top priority in the exploration research. It includes the modeling of dynamics environment and the orbital dynamics mechanism. This paper introduced state-ofthe-art researches, major challenges, and future trends in this field. Three topics are mainly discussed： the gravitational field modeling of irregular-shaped small celestial bodies, natural orbital dynamics and control, and controlled orbital dynamics. Finally, constructive suggestions are made for China’s future space exploration missions.

Spacecraft formation control strategy on Sun-Earth Lissajous orbit

To carry out the deep space exploration tasks near Sun-Earth Libration point L2, the CRTBP dynamic model was built up and the numerical conditional quasi-periodic orbit (Lissajous orbit) was computed near L2. Then, a formation controller was designed with linear matrix inequality to overcome the difficulty of parameter tuning. To meet the demands of formation accuracy and present thruster’s capability, a threshold scheme was adopted for formation control. Finally, some numerical simulations and analysis were completed to demonstrate the feasibility of the proposed control strategy.

Shift control method for the local time at descendi ngnode based on sun-synchronous orbit satellite

This article analyzes the shift factors of the descending node local time for sun-synchronous satellites and proposes a shift control method to keep the local time shift within an allowance range. It is found that the satellite orbit design and the orbit injection deviation are the causes for the initial shift velocity, whereas the atmospheric drag and the sun gravitational perturbation produce the shift acceleration. To deal with these shift factors, a shift control method is put forward, through such methods as orbit variation design, orbit altitude, and inclination keeping control. The simulation experiment and practical application have proved the effectiveness of this control method.

Orbit control strategy for Lagrange point orbits based on an analytical method

The periodic or quasi-periodic orbits around collinear Lagrange points present many properties that are advantageous for space missions. These Lagrange point orbits are exponentially unstable. On the basis of an analytical method, an orbit control strategy that is designed to eliminate the dominant unstable components of Lagrange point orbits is developed. The proposed strategy enables the derivation of the analytical expression of nonlinear control force. The control parameter of this strategy can be arbitrarily selected provided that the parameter is considerably lower than the negative eigenvalue of motion equations, and that the energy required keeps the same order of magnitude. The periodic or quasi-periodic orbit of controlled equations remains near the periodic or quasi-periodic orbit of uncontrolled equations.

Ground demonstration system based on in-orbit assembly oriented manipulator flexible force control

To eliminate the load weight limit of carrier rockets and reduce the burden on support structures,in-orbit assembly is a key technology to make design of scattering a large diameter telescope into submirror modules,which requires smooth operation of assembly robots,and flexible force control technology is necessary. A ground demonstration system is presented for in-orbit assembly focusing on flexible force control. A six-dimensional force/torque sensor and its data acquisition system are used to compensate for gravity. For translation and rotation,an algorithm for flexible control is proposed. A ground transportation demonstration verifies accuracy and smoothness of flexible force control,and the transportation and assembly task is completed automatically. The proposed system is suitable for the development of in-orbit assembly robots.

太阳成像仪镜筒舱内压力控制技术研究

文章阐述了太阳成像仪镜筒舱内压力控制技术研究,主要包括限流小孔机械尺寸理论分析、慢速泄压阀设计、镜筒压力控制模拟测试等内容。目前,所研制的慢速泄压阀已顺利交付使用,并随46.5 nm极紫外太阳成像仪在中科院空间新技术试验卫星上进行了搭载,实现了在轨顺利开机,验证了太阳成像仪镜筒舱内压力控制手段的正确性。

对地观测卫星编队协同容错姿轨耦合控制

面向对地观测场景下卫星编队的姿轨耦合特性和完全失效故障,提出了基于空间构型重构的协同容错控制方法。首先,根据对地观测的任务需求,分别建立了无故障情况下的卫星编队构型,以及故障发生后将故障卫星移出编队的构型重构方案,包括拓扑网络重构、卫星编号更新和编队构型重构算法;其次,根据上述构型,考虑外部干扰和姿态轨道耦合特性,设计了基于非奇异终端滑模控制的分布式控制律,使得编队在有限时间内稳定,即在故障发生前后卫星编队都可以实现对观测区域的协同覆盖;最后,分别通过数值仿真和实验验证表明了所提方法的有效性。

Design and simulation of fault diagnosis based on NUIO/LMI for satellite attitude control systems

This paper presents a scheme of fault diagnosis for flexible satellites during orbit maneuver. The main contribution of the paper is related to the design of the nonlinear input observer which can avoid false alarm arising from the disturbance from orbit control force. The effects of orbit control force on the fault diagnosis system for satellite attitude control systems, including the disturbing torque caused by the misalignments and the model uncertainty caused by the fuel consumed, are discussed, where standard Lu- enberger observer cannot work well. Then the nonlinear unknown input observer is proposed to decouple faults from disturbance, Besides, a linear matrix inequality approach is adopted to reduce the effect of nonlinear part and model uncertainties on the observer. The numerical and semi-physical simulation demonstrates the effectiveness of the proposed observer for the fault diagnosis system of the satellite during orbit maneuver.

Design and control of multiple spacecraft formation flying in elliptical orbits

Spacecraft formation flying is an attractive new concept in international aeronautic fields because of its powerful functions and low cost. In this paper, the formation design and PD closed-loop control of spacecraft formation flying in elliptical orbits are discussed. Based on two-body relative dynamics, the true anomaly is applied as independent variable instead of the variable of time. Since the apogee is considered as the starting point, the six integrating constants are calculated. Therefore, the algebraic solution is obtained for the relative motion in elliptical orbits. Moreover, the formation design is presented and both circular formation and line formation are provided in terms of an algebraic solution. This paper also discusses the PD-closed loop control for precise formation control in elliptical orbits. In this part, the error-type state equation is put forward and the linear quadratic regulator （LQR） method is used to calculate PD parameters. Though the gain matrix calculated from LQR is time-variable because the error-type state equation is time variable, the PD parameters are also considered as constants because of their small changes in simulation. Finally, taking circular formation as an example, the initial orbital elements are achieved for three secondary spacecraft. And the numerical simulation is analyzed under PD formation control with initial errors and J2 perturbation. The simulation results demonstrate the validity of PD closed-loop control scheme.

利用径向力平衡飞行控制的航天器高精度轨道捕获方法

为实现对探测器轨道形状与高度的精准调整,提出一种径向力平衡飞行的航天器连续推力控制新方法。建立连续推力平衡飞行的动力学极坐标模型,并推导出特殊条件下的解析轨道解,进一步分析边值条件,给出连续推力的控制律。利用这一平衡飞行控制理论,构建轨道捕获的最优控制策略。考虑推力器的推力水平,通过一次或多次的控制过程,实现对轨道形状、轨道高度及轨道相位的综合调整。数值仿真表明:利用平衡飞行的轨道控制方法,配置微小推力器的空间引力波探测器可以实现高精度的轨道捕获;该方法具有控制过程可解析、计算量小、简便、实用等特点。

火星上升器入轨段自适应迭代制导方案设计

针对火星上升器入轨段的制导问题,设计了一种自适应迭代制导策略。为降低初始状态偏差和火星环境不确定性的影响,每个周期内在制导坐标系下迭代计算剩余飞行时间,求解出在上升器推力大小固定情况下满足目标点位置与速度矢量约束的最优控制角,进而实时修正飞行轨迹。仿真结果表明,相较于传统的开环制导方案,提出的方案显著地提升了制导精度,其中高度误差降低3个量级,最大速度误差降低至原开环制导的三分之一,入轨点轨道倾角和偏心率误差均满足基本工程需求,可作为未来火星上升器入轨段制导的一种可靠方案。

空间机动目标双臂协调柔顺操作控制方法

针对被操作目标运动状态变化导致双臂与目标间内力发生突变的问题,提出了一种空间机动目标双臂协调柔顺操作控制方法。首先,针对空间双臂协调柔顺操作任务,分别建立双臂空间机器人开链动力学模型和闭链动力学模型;其次,考虑目标机动产生的扰动,设计阻抗参数自主调整策略,在双臂末端与目标间作用力发生较大变化时,可以保证双臂协调操作任务的平稳性和安全性;最后,搭建双臂协调柔顺操作地面物理试验系统,并开展物理仿真试验验证,试验结果表明了本文所提方法的有效性。

分布式航天器编队约定时间姿轨耦合协同控制

针对具有参数不确定性、外部扰动和预设性能需求的航天器编队约定时间姿轨耦合控制问题,本文基于预设性能控制方法提出了一种低复杂度的约定时间编队姿轨耦合协同控制器,使得航天器在设定时间内形成编队,且编队误差满足预设的各种性能指标.首先,通过结合有限时间稳定概念引入一种约定时间性能函数,其系统稳定时间可以由使用者任意设定;然后,把约定时间性能函数与预设性能控制方法结合起来,提出了不依赖航天器质量和转动惯量等信息的约定时间编队协同控制器,保证了编队状态量的收敛性能和收敛时间,并使用李雅普诺夫理论证明了其稳定性.最后,通过仿真验证了该控制方案的有效性.