Two-dimensional launch window method to search for launch opportunities of interplanetary missions

Accurate and highly efficient approaches to obtain mission opportunities are still the goals of mission planners of interplanetary explorations. The search for launch opportunities not only determines the specified launch window of the mission but also presents the performance requirements for the interplanetary probe and its launch vehicle. An effective method, namely the two-dimensional launch window method, is developed from a completely new perspective to determine all the launch opportunities of the mission in this research. For a fixed launch time, the method to determine all the time windows in the dimension of Time-of-Flight (TOF) is firstly proposed and these time windows represent all the launch opportunities for the given launch time. And then, the two-dimensional launch window method is proposed, which computes the time windows in both the launch time and TOF dimensions to achieve all launch opportunities of the mission. Numerical examples are provided to demonstrate the accuracy and high efficiency of the method. Compared with the widely-used pock-chop plot method, the proposed method reduces the computational time by two orders of magnitude for the same search precision, and thus is especially suitable for the cases involving rapid, high-precision, and/or large-scale searches for mission opportunities.     

载人登月着陆器奔月窗口搜索方法

对环月轨道共面交会的载人登月任务中,着陆器(LM)奔月零窗口与轨道参数精确快速设计方法进行了研究。任务采用人货分离奔月模式,着陆器于载人飞船到达环月轨道前抵达环月共面交会轨道,着陆器近月点一次共面减速完成近月制动。提出一种三层快速精确奔月窗口搜索方法:第一层采用地心二体轨道理论解析计算月窗口及奔月轨道参数初值,作为正确性基本参考;第二层采用改进的双二体解析动力学模型求解月窗口内奔月轨道参数变化规律;第三层采用高精度轨道动力学模型和SQP_Snopt优化求解奔月零窗口及轨道参数精确解。仿真结果表明,本文提出的三层逐级奔月窗口搜索方法能快速精确求解载人登月任务中着陆器奔月窗口及精确轨道参数,也揭示了影响着陆器奔月窗口的主次因素和规律,为中国未来载人登月工程提供参考。     

交会对接任务碰撞规避策略制定研究简

针对交会对接任务目标飞行器与追踪器轨道运行特性,综合考虑规避策略计算方法与工程实际相结合的问题,提出高度规避、时间规避以及与正常轨控相结合的碰撞规避策略计算方法等三种空间目标碰撞规避策略计算方法.高度规避计算方法采用了Lambert飞行原理,用简化二体开普勒模型取代高精度轨道预报方法,迭代求解规避机动速度增量,实现了通过约束过交点与目标径向距离差得到速度增量的最优解;时间规避计算方法通过轨道周期与速度增量的关系,实现了通过约束过交点与目标的时间差得到速度增量的最优解;与正常轨控相结合的碰撞规避策略计算方法,在正常控制考虑冗余控制量的基础上,对控制策略的控制开始时间或沿迹方向的速度增量进行较小的修正,使两者通过碰撞点的时刻或径向距离错开,达到碰撞规避的目的,该方法不仅可以节省燃料、而且对任务的影响较小.通过对三种空间目标碰撞规避策略计算方法仿真分析结果表明,完全适用于交会对接任务,可为我国载人航天任务飞行安全提供技术保障.     

基于STK的小卫星轨道交会设计研究

利用球面三角形的几何关系进行了基于STK(卫星工具包)的小卫星轨道交会的规划和设计。采用一种较为方便的轨道交会设计方法,并使用功能强大的专业STK软件来仿真和演示,取得了良好的仿真效果,能够较好地满足航天器轨道交会设计的要求,对于其它类型的轨道规划有一定的借鉴作用,为各种航天器的仿真研究提供了一种新的方法。     

考虑潜在威胁区的航天器最优规避机动策略

随着一系列轨道转移飞行器的工程化实施,航天器可能面临的非合作交会威胁日趋严重。针对该问题,根据交会机动的特点定义了新的规避机动指标——潜在威胁区,相较于传统的相对距离和碰撞概率等规避指标,潜在威胁区更适合航天器在面对非合作交会追踪器时进行规避机动,能够有效提升航天器的抗交会能力。首先,建立追踪器多脉冲最优交会模型,以此为基础给出潜在威胁区的定义与计算方法;然后,以潜在威胁区弧长为优化目标,建立了目标器最优规避模型,采用遗传算法进行目标优化;最后,根据所建立的双层优化模型进行数值仿真,以初始相距100km为初始条件进行仿真并计算得到了使潜在威胁区为零所需规避脉冲值,验证了文中模型的正确性,结果显示所定义的潜在威胁区弧长随着规避脉冲的增大呈严格的单调递减关系。研究为在轨航天器在面对非合作交会时提供了有效的规避策略,提升了航天器的空间生存能力。     

基于弹道逃逸和小推力捕获的地月转移轨道设计

针对地月空间货运任务和环月轨道空间设施建设任务,提出一种弹道逃逸和小推力捕获相结合的新型地月轨道转移模式,并建立了一整套该类型轨道设计方法。首先,在三体模型假设下分别建立地心弹道逃逸轨道和月心小推力捕获轨道的二维极坐标动力学模型。对于弹道逃逸轨道,将地心旋转系对准角和地月转移加速速度增量作为控制变量,提出初值估计解析公式,并应用序列二次规划算法进行快速求解。对于小推力捕获轨道,以月心距为参考量设置与弹道逃逸轨道的拼接点约束,提出能量匹配方法预估飞行时间,采用最优螺旋轨道的初始伴随状态解析式预估近月点伴随变量初值。基于混合法和轨道逆推思想,采用人工免疫算法进行小推力捕获轨道求解。仿真结果表明,基于弹道逃逸和小推力捕获的地月轨道转移方式大幅降低了近月制动燃料消耗,能快速穿越地球辐射带,且飞行时间适中;同时,提出的轨道设计方法能快速搜索到基于弹道逃逸和小推力捕获的地月转移轨道,验证了该方法的有效性。     

Tangent-impulse transfer from elliptic orbit to an excess velocity vector

The two-body orbital transfer problem from an elliptic parking orbit to an excess veloc-ity vector with the tangent impulse is studied. The direction of the impulse is constrained to be aligned with the velocity vector, then speed changes are enough to nullify the relative velocity. First, if one tangent impulse is used, the transfer orbit is obtained by solving a single-variable function about the true anomaly of the initial orbit. For the initial circular orbit, the closed-form solution is derived. For the initial elliptic orbit, the discontinuous point is solved, then the initial true anomaly is obtained by a numerical iterative approach; moreover, an alternative method is proposed to avoid the singularity. There is only one solution for one-tangent-impulse escape trajectory. Then, based on the one-tangent-impulse solution, the minimum-energy multi-tangent-impulse escape trajectory is obtained by a numerical optimization algorithm, e.g., the genetic method. Finally, several examples are provided to validate the proposed method. The numerical results show that the minimum-energy multi-tangent-impulse escape trajectory is the same as the one-tangent-impulse trajectory.     

轨道拦截预警算法研究

针对空间轨道拦截作战模式,提出了一种拦截预警的计算方法.推导了求解轨道交点位置的表达式,然后结合拦截卫星的末端威胁距离给出了判断拦截是否发生的地心距判据和时间差判据,最后通过仿真算例验证了交点表达式以及判据的正确性,并讨论了拦截卫星的威胁距离与时间差门限的关系.     

Robust reliable control for autonomous spacecraft rendezvous with limited-thrust

This paper investigates the problem of robust reliable control for the spacecraft rendezvous with limited-thrust. Based on the Clohessy–Wiltshire (C–W) equations and by considering the uncertainties and the possible failures, the dynamic model for spacecraft rendezvous is proposed, and the orbital transfer control problem is transformed into a stabilization problem. Then, by a Lyapunov approach, the existence conditions for admissible controllers are formulated in the form of linear matrix inequalities (LMIs), and the controller design is cast into a convex feasibility problem subject to LMI constraints. With the obtained controllers, the rendezvous can be accomplished with the limited-thrust in spite of the possible thruster failures. The effectiveness of the proposed approach is illustrated by simulation examples.     

基于SAS算法的起飞一发失效应急路径规划方法

为解决起飞一发失效应急程序（EOSID）手动设计的不足,提出一种基于SRTM数据的稀疏A*搜索（SAS）算法的EOSID路径规划方法。首先采用航天飞机雷达地形测绘使命（SRTM）的网格地形数据,结合起飞一发失效相关规章,考虑爬升梯度与保护区限制确定可行搜索空间;然后基于可行搜索空间运用稀疏A*搜索算法搜索应急离场路径,在传统A*算法寻找扩展节点时加入起飞性能约束条件,同时利用地形高程数据进行地形和威胁回避,生成一条三维应急离场航迹;最后利用三次样条曲线对规划的应急离场航迹进行平滑处理。实验结果表明该方法能自动搜索出有效的EOSID三维航迹。     

Scheme design of the CHANG’E-5T1 extended mission

Flight schemes for the CHANG’E-5T1 extended mission are investigated in this paper. In the flight scheme and trajectory design, the remaining propellant of the CHANG’E-5T1 mission is utilized. The CHANG’E-5T1 mission is firstly introduced with feasible flight goals derived based on the terminal trajectory and satellite status. The flight schemes are designed to include a lunar return and the libration points in the Sun-Earth/Moon and Earth-Moon systems, with an emphasis on the Earth-Moon triangle libration point thus far unexplored. Secondly, three schemes are proposed for the CHANG’E-5T1 extended mission with different flight goals. The direct libration point orbit transfer and injection method is adopted to solve the issue in the transfer trajectory design. Furthermore, an innovative concept is proposed to transfer from the Earth-Moon collinear libration point to the triangle point using the Sun-Earth/Moon libration point. Finally, the merits and drawbacks of the three schemes are discussed in terms of flight time, control energy and frequency, flight distance, and goal value. As a result, the scheme including a lunar return and the Earth-Moon L2 libration point is selected for the CHANG’E-5T1 extended mission. A flight to the Earth-Moon libration point is achieved, replicating the achievement of the ARTEMIS mission.     

Earth-moon Trajectory Optimization Using Solar Electric Propulsion

The optimization of the Earth-moon trajectory using solar electric propulsion is presented. A feasible method is proposed to optimize the transfer trajectory starting from a low Earth circular orbit (500 km altitude) to a low lunar circular orbit (200 km altitude). Due to the use of low-thrust solar electric propulsion, the entire transfer trajectory consists of hundreds or even thousands of orbital revolutions around the Earth and the moon. The Earth-orbit ascending (from low Earth orbit to high Earth orbit) and lunar descending (from high lunar orbit to low lunar orbit) trajectories in the presence of J2 perturbations and shadowing effect are computed by an analytic orbital averaging technique. A direct/indirect method is used to optimize the control steering for the trans-lunar trajectory segment, a segment from a high Earth orbit to a high lunar orbit, with a fixed thrust-coast-thrust engine sequence. For the trans-lunar trajectory segment, the equations of motion are expressed in the inertial coordinates about the Earth and the moon using a set of nonsingular equinoctial elements inclusive of the gravitational forces of the sun, the Earth, and the moon. By way of the analytic orbital averaging technique and the direct/indirect method, the Earth-moon transfer problem is converted to a parameter optimization problem, and the entire transfer trajectory is formulated and optimized in the form of a single nonlinear optimization problem with a small number of variables and constraints. Finally, an example of an Earth-moon transfer trajectory using solar electric propulsion is demonstrated.     

基于空闲时间窗和多Agent的A-SMGCS航空器滑行路由规划

先进场面活动引导与控制系统(A-SMGCS)中的航空器滑行路由规划是一个典型NP难题。为解决航空器滑行路由规划的优化性和计算量之间的矛盾,提出一种基于空闲时间窗的路由规划方法,并利用多Agent系统(MAS)进行算法求解。首先,建立滑行资源图以对场面滑行区进行建模。其次,按照航班计划为航空器设置滑行优先级,并按优先级顺序依次规划路由,后规划的路由不破坏已有路由,即利用滑行路段的空闲时间窗进行规划。每次只需为一架航空器规划滑行路由,降低了问题的求解难度;通过搜索空闲时间窗获得路由使场面交通均衡分布,保证了路由规划的整体优化性。分析了空闲时间窗特性,指出空闲时间窗的可达性条件和避免同步资源交换冲突的条件。最后,设计MAS,把建立、维护和搜索空闲时间窗图的复杂集中式求解过程简化为通过路由管理Agent,航空器Agent和资源节点Agent相互协作实现对场面路由规划问题的分布式求解。仿真结果表明,设计的MAS能够快速找到空闲时间窗中的最优解;与固定预选滑行路径算法相比,航空器的平均滑行时间显著减少,最多可以节省19.6%的滑行时间。     

基于混合粒子群算法的上升段交会弹道快速优化设计

基于梯度搜索的高效性和粒子群搜索的随机性,提出了一种混合粒子群算法,并应用该算法研究了运载火箭上升段交会弹道快速优化设计问题.以运载火箭与目标飞行器在交会时刻的距离最小为目标函数,设计了运载火箭飞行程序,建立了运载火箭上升段交会弹道优化模型,同时分别采用混合粒子群算法、遗传算法和粒子群算法进行求解.仿真结果表明:基于本文算法对运载火箭上升段交会弹道进行优化设计,平均交会位置误差为4.137m,较遗传算法减少了17.940m,平均优化耗时488.922s,较粒子群算法缩短了2342.125s.混合粒子群算法搜索速度较快,收敛精度较高,可用于运载火箭上升段交会弹道的快速优化设计.      

A novel cooperative mid-course guidance scheme for multiple intercepting missiles

In the interception engagement,if the target movement information is not accurate enough for the mid-course guidance of intercepting missiles,the interception mission may fail as a result of large handover errors.This paper proposes a novel cooperative mid-course guidance scheme for multiple missiles to intercept a target under the condition of large detection errors.Under this scheme,the launch and interception moments are staggered for different missiles.The earlier launched missiles can obtain a relatively accurate detection to the target during their terminal guidance,based on which the latter missiles are permitted to eliminate the handover error in the mid-course guidance.A significant merit of this scheme is that the available resources are fully exploited and less missiles are needed to achieve the interception mission.To this end,first,the design of cooperative handover parameters is formulated as an optimization problem.Then,an algorithm based on Monte Carlo sampling and stochastic approximation is proposed to solve this optimization problem,and the convergence of the algorithm is proved as well.Finally,simulation experiments are carried out to validate the effectiveness of the proposed cooperative scheme and algorithm.     

Repair of Salyut 7

The Soviet Union's mission to rendezvous with and repair the Salyut 7 Space Station, which had gotten out of control, is described. This was the Soviets' first attempt to dock with an uncontrolled object, and some of their instruments were being used for the first time. Both the equipment and the procedures used are detailed     

基于正交试验设计的交会远距离导引误差分析

由于测量-规划-执行闭环过程的引入,交会远距离导引段的导航执行误差之间存在交互作用。对一个实际的远距离导引任务,采用正交试验设计方法安排数值试验,通过离差平方和评估误差之间的交互作用。结果表明,影响推进剂消耗的主要误差源之间存在显著的交互作用,而影响终端相对位置速度的主要误差源之间的交互作用很小或不明显。     

非合作目标交会的双层MPC全局轨迹规划控制

针对空间非合作目标近距离视线交会中的全局最优鲁棒轨迹规划与控制问题,提出了基于高斯伪谱方法（GPM）和线性时变模型预测控制（LTVMPC）的双层模型预测控制（MPC）算法。在轨迹规划方面,以视线坐标系下的相对轨道动力学为模型、能量最少和控制精度最优为性能指标构建最优控制问题,利用GPM精度高、收敛速度快的特点将最优控制问题转化为易于求解的全局非线性规划问题,在MPC框架下求解得到全局最优的标称轨迹,克服了传统的MPC不适用于全局大范围非线性规划的缺点;在轨迹跟踪控制方面,考虑预测时域内状态转移矩阵的时变特性,设计了LTVMPC算法对标称轨迹进行追踪,避免了存在不确定性时轨迹的重规划,从而降低在线计算量,保证算法在线自主实施,并且采用滚动优化的策略使算法对不确定性具有鲁棒性。由于规划层和控制层考虑的约束相同,因此规划的轨迹是可控、可达的。数字仿真表明,在燃料消耗和交会时间等方面,提出的方法均显著优于传统的MPC方法,相较于传统的MPC方法,新算法的交会时间减少50%左右,燃料消耗降低30%以上。     

自适应动态规划算法在飞行器追逃中的应用

针对飞行器追逃对抗的二人零和微分对策问题,提出基于数据的积分策略迭代自适应动态规划算法,以求解数学模型未知系统的控制律.该算法利用固定时段内有效的状态和输入信息,建立数据模型,并对其进行基于值函数和控制策略的算法迭代,在平面拦截系统完全未知的情况下得到追逃双方的近似最优策略.仿真结果表明,所得到的双方控制策略能在有限界内无限接近最优解,验证了所提出算法的有效性.     

Orbital reference frame estimation with power spectral density constraints for drag-free satellites

The drag-free satellites are widely used in the field of fundamental science as they enable the high-precision measurement in pure gravity fields. This paper investigates the estimation of local orbital reference frame (LORF) for drag-free satellites. An approach, taking account of the combi-nation of the minimum estimation error and power spectral density (PSD) constraint in frequency domain, is proposed. Firstly, the relationship between eigenvalues of estimator and transfer func-tion is built to analyze the suppression and amplification effect on input signals and obtain the eigenvalue range. Secondly, an optimization model for state estimator design with minimum estima-tion error in time domain and PSD constraint in frequency domain is established. It is solved by the sequential quadratic programming (SQP) algorithm. Finally, the orbital reference frame estimation of low-earth-orbit satellite is taken as an example, and the estimator of minimum variance with PSD constraint is designed and analyzed using the method proposed in this paper.