椭圆轨道非合作目标交会接近策略与控制

文章研究了追踪航天器与失控旋转非合作目标航天器在椭圆轨道中的交会接近策略与控制。在接近策略方面,首先,根据目标航天器大致结构设定一个安全的停泊点,使追踪航天器交会至停泊点;其次,通过在停泊点对旋转目标航天器姿态的观测,分析和预测其运动并确定合适的抓捕点位置,设计安全的接近轨迹,使追踪航天器沿着该轨迹接近至理想的抓捕实施点位置。在控制方面,考虑实际系统中的不确定性,只利用两航天器之间相对位置的测量信息,设计基于特征模型的自适应控制方法实现交会接近。最后通过数学仿真模拟整个交会接近过程,验证了文中所提出的接近策略和控制方法。     

空间交会最终平移段控制策略

提出空间交会最终平移段的控制方法,选取追踪航天器的相对位置与姿态角作为控制变量,同步控制追踪航天器的质心运动与姿态运动,修正制导机动执行偏差的影响,使目标航天器保持在光学导航视场范围之内,且满足对接操作对追踪航天器状态的要求。     

论空间交会最终平移段制导设计

文章阐述了空间交会最终平移段的制导设计方法。在最终平移段 ,追踪航天器沿视线方向作受迫运动 ,逼近目标航天器对接部位 ,追踪航天器相对速度的变化一般考虑指数型与等速型两种模式 ,采用分段制导策略。由追踪航天器相对运动轨迹及速度变化率确定机动加速度 ,在工程上采用多次有限常推力机动方式或多次冲量机动方式     

航天器交会飞行设计方法研究

针对半自主飞行追踪星,阐述航天器交会总体设计方法。根据对接点的地理位置范围、共面轨道倾角以及目标星轨道周期与追踪星入轨点地理位置,确定交会飞行时间和两星初始相位差范围。考虑最小轨道机动动力要求与飞行轨迹安全性等因素,并兼顾地面测控条件,设计追踪星远程导引段与相对导航段的轨道机动与飞行轨迹,特别是选择与比较不同的初始轨道、调相轨道与漂移轨道以及保持点停泊时间与最终逼近段飞行时间等交会飞行要素,调整飞行时间、相位差与对接点位置,确定最佳交会飞行方案,完成空间交会任务。     

非合作目标的自主接近控制律研究

主要研究了针对非合作航天器的自主交会和拦截两种接近模式的控制律。根据非线性相对运动方程,采用李雅普诺夫(Lyapunov)方法设计了自主交会的控制律;并对该控制律进行了改进,使其适用于拦截模式,能以一定的相对速度接近目标。同时针对非合作目标存在机动的情况,采用Lyapunov最小-最大方法设计了自主接近的控制律。仿真结果证明了基于Lyapunov方法的自主接近控制律的有效性。     

人控遥操作交会对接设计与验证

人控操作交会对接是指操作人员在远端通过遥操作方式控制追踪航天器进行交会对接,主要用于无人航天器自动交会对接系统故障条件下交会对接控制.简要介绍遥操作交会对接的基本概念,给出实现人控遥操作交会对接三种方案,设计一种易于工程实现的遥操作交会对接系统,分析关键技术并给出了解决方法;根据设计的系统方案,进行遥操作交会对接试验,试验结果表明设计方案合理可行.     

基于线性协方差方法的交会对接误差分析

将线性协方差分析方法和蒙特卡罗仿真相结合,按交会任务和飞行特征把交会过程分为变轨飞行、自由飞行和中途速度修正三种特征段,研究了状态误差的传播规律和交会过程中各种误差对交会对接精度的影响。在变轨飞行段,分析了追踪航天器的姿态误差、控制系统性能状态估计误差,以及目标航天器轨道摄动对状态误差传播的影响。在自由飞行段,分析了追踪航天器估计状态误差的先验值和测轨误差对状态误差传播的影响。在中途速度修正段,分析了追踪航天器姿态误差和控制系统性能误差对状态误差传播的影响。仿真结果表明,误差分析方法设计合理,可以指导交会对接的轨道设计工作,能对已经设计好的交会策略进行误差分析和设计验证。     

基于马达代数的交会对接相对位姿测量算法

精确测量追踪航天器与目标航天器之间的相对位姿关系是成功完成航天器交会对接任务的关键。传统的位姿测量算法将旋转和平移分而视之,破坏了三维运动的统一性,同时增加了算法的复杂性和计算难度。针对这个问题,采用双目视觉测量方法,在马达代数框架内,以两个特征光点确定的特征直线为变换基元,统一描述并设计算法测量追踪与目标航天器之间的位姿关系,最终将两航天器之间的相对位姿解算问题简化为求解两个线性方程组。该方法在计算形式上更为简洁,且测量受特征光点的安装位置限制较小。仿真结果表明,算法具有较高的精度和稳定性,可以满足航天器交会对接任务的要求。     

航天器空间自主交会故障诊断

研究了空间自主交会中最终逼近段轨道控制的故障诊断问题.采用C-W方程描述圆轨道上的目标航天器与追踪航天器的相对运动关系,首先针对C-W方程设计了基于开环模型的未知输入鲁棒故障观测器,然后针对空间自主交会的闭环控制问题,设计了基于线性矩阵不等式(LMI)的H∞控制器.在有干扰条件下,进行鲁棒控制和PD控制两种控制器的自主交会对比仿真,仿真结果表明设计的控制器可以完成自主交会任务,但故障诊断器在不同的控制器下诊断的效果并不一样,最后进行了总结并探讨了这一阶段新的研究方向.     

目标飞行器交会对接轨道的设计和控制

根据目标飞行器轨道高度和追踪飞行器入轨轨道高度,给出了目标飞行器交会对接轨道初始相位的设计方法。针对目标飞行器交会对接轨道控制要求,建立了共面相位计算模型以及轨道相位、高度和圆化度的多目标参数求解模型。基于定轨误差、轨道控制误差和轨道预报误差的调相时间分析,制定了目标飞行器调相控制策略。仿真计算表明,实现的目标飞行器交会对接轨道满足要求,验证了调相控制量优化原则的正确性,并对标称共面与虚拟共面的共面时刻和共面相位进行了比较。所提出的计算模型、控制策略和分析方法适用于目标飞行器交会对接轨道设计和控制实施。     

非合作自主交会对接的动态障碍物躲避制导

首先,在视线坐标系下建立了系统相对运动状态方程,将人工势函数制导方法应用于航天器的非合作自主交会对接任务和动态障碍物躲避问题。其次,利用Lyapunov稳定性理论分析证明了在该制导方法控制下系统的稳定性,并且研究讨论了两种不同情形下的动态障碍物躲避效果,分析了人工势函数制导方法的应用能力。最后,用精确的数学模型进行了数值仿真,验证了制导方法应用于所研究问题的正确性和有效性。     

Optimal guidance and collision avoidance for docking with the rotating target spacecraft

The guidance and control strategy for spacecraft rendezvous and docking are of vital importance, especially for a chaser spacecraft docking with a rotating target spacecraft. Approach guidance for docking maneuver in planar is studied in this paper. Approach maneuver includes two processes: optimal energy approach and the following flying-around approach. Flying-around approach method is presented to maintain a fixed relative distance and attitude for chaser spacecraft docking with target spacecraft. Due to the disadvantage of energy consumption and initial velocity condition, optimal energy guidance is presented and can be used for providing an initial state of flying-around approach process. The analytical expression of optimal energy guidance is obtained based on the Pontryagin minimum principle which can be used in real time. A couple of solar panels on the target spacecraft are considered as obstacles during proximity maneuvers, so secure docking region is discussed. A two-phase optimal guidance method is adopted for collision avoidance with solar panels. Simulation demonstrates that the closed-loop optimal energy guidance satisfies the ending docking constraints, avoids collision with time-varying rotating target, and provides the initial velocity conditions of flying-around approach maneuver. Flying-around approach maneuver can maintain fixed relative position and attitude for docking.     

组合体航天器姿态的智能自适应控制方法

研究了交会对接后组合体航天器构型变化带来的姿态控制问题,对执行机构在控制量受限时的控制能力进行了分析,应用基于特征模型的智能自适应控制方法,设计了能适用于不同构型的姿态控制器,分别对组合体在直线构型和L构型对接情况下进行了数学仿真,仿真结果验证了智能自适应控制方法可行并且具有一定的优越性。     

基于鲁棒协同控制方法的多航天器交会问题

提出一种解决多航天器交会问题的协同控制算法。首先应用图论中邻接矩阵及拉普拉斯矩阵的定义及其相关性质,描述了多航天器之间的通信拓扑关系;其次对目标航天器轨道为椭圆形情况下的交会问题进行构建,并设计了相应的协同控制算法;最后利用李雅普诺夫函数证明该系统的稳定性,并且能够保证消耗的能量最优以及最大推力受限。仿真实验表明:提出的方法可以实现多航天器的协同交会,验证了该方法的有效性。     

Constant thrust fuel-optimal control for spacecraft rendezvous

In this paper, constant thrust rendezvous is studied and the optimal rendezvous time is calculated by using continuous genetic algorithm. Firstly, the relative position parameters of the target spacecraft are obtained by using the vision measurement and the target maneuver positions are calculated through the isochronous interpolation method. Then, the results of the calculation of constant thrust rendezvous is founded by processing with multivariate linear regression method. Next, a new switching control law is designed based on the thrust acceleration sequence and the on time of thrusters which can be computed by the time series analysis method. The perturbations and fuel consumptions are addressed during the computation of the on time of thrusters.     

Robust fault-tolerant saturated control for spacecraft proximity operations with actuator saturation and faults

In this paper, the motion control problem of autonomous spacecraft rendezvous and docking with a tumbling target in the presence of unknown model parameters, external disturbances, actuator saturation and faults is investigated. Firstly, a nonlinear six degree-of-freedom dynamics model is established to describe the relative motion of the chaser spacecraft with respect to the tumbling target. Subsequently, a robust fault-tolerant saturated control strategy with no precise knowledge of model parameters and external disturbances is proposed by combining the sliding mode control technique with an adaptive methodology. Then, within the Lyapunov framework, it is proved that the designed robust fault-tolerant controller can guarantee the relative position and attitude errors converge into small regions containing the origin. Finally, numerical simulations are performed to demonstrate the effectiveness and robustness of the proposed control strategy.     

Tube-based robust output feedback model predictive control for autonomous rendezvous and docking with a tumbling target

In this paper, a tube-based robust output feedback model predictive control method (TRMPC) is proposed for controlling chaser spacecraft docking with a tumbling target in near-circular orbit. The controller contains a simple, stable, Luenberger state estimator and a tube-based robust model predictive controller. Several practical challenges are also considered under dock-enabling conditions, such as the control saturation, velocity constraint, approach corridor constraint, and collision avoidance constraint. Meanwhile, uncertainties are carefully analyzed when designing the controller, including dynamics uncertainty, measurement error, and control deviation. The TRMPC ensures that all possible state trajectories with uncertainties lie in the minimum robust positively invariant set (mRPI, i.e., the so-called tube in this paper). The tube center is the solution of a nominal (without uncertainties) system. Another important contribution of this paper is to propose a technique where it is unnecessary to calculate the mRPI explicitly. Thereby, the ‘curse of dimensionality’ can be avoided for a six-dimensional system. To verify the feasibility of the proposed TRMPC strategy in the presence of uncertainties, two scenarios of autonomous rendezvous and docking (AR&D) are simulated. The simulation results show that the TRMPC method is more efficient in minimizing the uncertainties, fuel consumption, and computational cost, compared to the classic model predictive control (MPC) method.     

Two-level optimization approach for Mars orbital long-duration,large non-coplanar rendezvous phasing maneuvers

A relative dynamics equation-set based on orbital element differences with J₂ effects is derived, based on which a two-level approach is proposed to optimize the Mars orbital rendezvous phasing with a large difference in the initial ascending node. The up-level problem uses the revolution deviation between the target spacecraft and the chaser as the design variable, and employs a linear search to find the optimum. The low-level problem uses the maneuver revolutions, locations, and impulses as the design variables, and is solved using a hybrid genetic algorithm combined with sequential quadratic programming. To improve the solution accuracy, an iteration method is developed to satisfy the terminal constraints of the absolute numerical integration trajectory. Test cases involving Mars sample return missions with large initial node differences are presented, which show that the relative dynamics, two-level optimization model, and hybrid optimization algorithm are efficient and robust. Compared with previously published results, the total velocity increment has been further reduced by utilizing this proposed approach. It is found that a five-impulse plan requires the least quantity of propellant, and a propellant-optimal minimum rendezvous duration exists for this long-duration, large non-coplanar rendezvous problem.