《中国空间科学技术》第26卷（2006年）总目次

第1期椭圆轨道编队的构形变化控制方法………………………………………………于萍,张洪华(1)航天器交会飞行设计方法研究……………………………………朱仁璋,汤溢,李颐黎等(9)基于积分域匹配滤波的系统误差检测……………………………段晓君,朱炬波,王光新(17)轨道方程计算中A     

编队飞行队形设计一般化方法

首先 ,在合理的假设条件下 ,基于运动学关系推导了环绕卫星相对参考卫星的相对运动方程 ;然后 ,给出了编队飞行队形设计一般化方法的具体计算公式 ,利用此方法 ,能在参考轨道为圆或椭圆轨道的情况下 ,直接求出编队中所有卫星的轨道参数 ;最后 ,给出了两个编队队形设计的算例     

星载InSAR辅星编队优化设计与分析

以“主星—辅星编队”InSAR系统干涉测高为应用背景,研究辅星编队的优化设计问题。首先,定义优化指标———“有效观测时间比”来描述全轨道周期观测性能,以共绕飞轨道、相位均匀分布的卫星编队(构形类似于CartWheel)作为研究的重点。选取描述编队卫星相对运动的振幅比和相位差作为寻优变量,通过仿真实例用搜索法得出了寻优变量与优化指标的三维关系图,并给出了优化构形的参数区域,得到了最优的编队卫星轨道根数。进一步研究了在主要摄动因素影响下卫星编队构形保持与轨道保持的燃料预算方法,并计算了上述优化构形一年的燃料预算。最后得出有益的结论,为全InSAR系统优化设计提供依据。     

基于交会模式的月球大椭圆轨道编队飞行控制

两颗微卫星进入环月大椭圆轨道后,在地面测控支持下,通过执行若干次轨道机动,最终实现从相距上千或上万km至相距1～10 km范围变化的环月轨道编队飞行。针对月球大椭圆轨道,基于多脉冲交会控制模式,设计了交会点满足编队飞行状态的轨道控制策略,采用线性制导方法迭代计算精确轨道控制参数;设计了顺序优化的5脉冲控制策略,对轨道平面、拱线、形状和相位等轨道全要素进行控制,通过远距离接近、中距离调整和近距离捕获的渐进式分段控制,在月球大椭圆轨道差异较大条件下,相对运动轨迹渐进稳定,最终实现近距离编队。     

椭圆轨道非线性相对运动模型的周期解与应用

在长期的航天器编队飞行中传统的基于线性相对运动模型设计编队保持轨道的方法会引起较多的燃料消耗.首先采用摄动法解析地求得了考虑二阶非线性项时椭圆轨道相对运动模型的周期性条件和周期解;然后以此周期解为参考轨道设计了基于Lyapunov稳定的PD保持控制律.仿真结果表明:基于椭圆轨道非线性相对运动模型的周期解较基于椭圆轨道线性相对运动模型的周期解,精度明显提高;以前者为参考轨道的保持控制与以后者为参考轨道的保持控制相比,燃耗明显降低.     

Walker星座摄动分析与保持控制策略

针对运行在中高轨道上的全球Walker星座的构形长期保持控制问题,通过分析星座卫星的轨道摄动及构形相对漂移,研究了星座构形长期保持的控制方法及控制策略.对于星座构形的保持,研究了星座构形摄动补偿法,提出了星座构形数值微分修正法.最后给出了全球星座构形长期保持控制策略并进行了仿真验证,仿真分析表明,星座构形数值微分修正法比构形摄动补偿法具有更高的控制计算精度.     

分布式SAR超分辨成像的卫星编队构形优化设计方法

编队构形设计是分布式卫星系统总体设计的关键问题,提出了一种编队构形优化设计方法。针对主星带辅星群体制分布式合成孔径雷达(SAR)超分辨率任务,从优化系统效能的角度出发,将满足超分辨性能的时间比率和分辨率改善因子作为编队构形优化设计的目标函数,以基线矢量为中间变量,建立了目标函数与编队卫星轨道根数的关系,并采用遗传算法求解。对不同卫星数目的编队系统进行了优化设计仿真,并与干涉车轮和钟摆编队构形方案进行了分析比较,数值结果验证了该优化设计方法的有效性和正确性。该设计方法还可改变优化目标函数,以适应用户的任务要求。     

近圆参考轨道卫星编队的安全性

针对近圆参考轨道卫星,利用相对轨道要素对编队构形进行了统一描述,提出了编队卫星在垂直于轨道平面方向相对运动最小距离的计算公式,分析了J2摄动对卫星编队安全性的影响。仿真结果表明环绕星在轨道平面内和垂直于轨道平面内的相位差对编队安全性的影响较大,给出的提高编队卫星安全性的措施可为卫星编队的实际工程应用提供参考。     

基于视线测量的卫星编队机动控制方法

采用视线测量的方法,建立一种编队卫星队形保持与机动的协同控制策略。编队中每一个卫星跟踪自己轨道前方邻近卫星,产生一个视线测量矢量,编队的第一个卫星根据高级控制层指令追踪期望轨道,产生链式编队,将编队卫星之间的视线距离作为反馈控制量来实现队形控制。通过推导J2相对摄动力的表达式,控制模型考虑了模型不确定性和摄动影响,采用滑模控制器,实现了基于视线测量的编队卫星链式跟踪协同控制。仿真算例结果表明,该方法在实现编队卫星队形保持与整体机动控制上具有可行性。     

日心悬浮轨道混合推进航天器编队构型保持研究

针对混合推进航天器编队日心悬浮轨道保持控制问题进行了研究.首先推导出在日心悬浮轨道附近的航天器编队相对运动方程，考虑到航天器间距离变化值较小且航天器间距离与航天器到太阳的距离的比值为小量，将其在悬浮轨道附近线性化.基于该线性化方程，设计了一种LQR编队控制方式，该控制方式可通过调节太阳帆的姿态及航天器间库仑力的大小对编队构型进行改变或保持，具有响应速度快和控制简单的特点.最后对控制律进行数值仿真，表明该控制方法能实现编队.     

Spacecraft formation and orbit control using differential attitude-dependent solar radiation pressure

This paper introduces a linear model for spacecraft formation dynamics subject to attitude-dependent solar radiation pressure (SRP) disturbance, with the SRP model accounting for both absorption and specular/diffuse reflection. Spacecraft attitude is represented in modified Rodriguez parameters (MRPs), which also parameterize the orientation of individual facets for a spacecraft with fixed geometry. Compared to earlier work, this model incorporates analytic approximation of the SRP-perturbed chief orbit behavior in a manner enabling its use in applications with infrequent guidance updates. Control examples are shown for single-plate representations of hypothetical spacecraft with generally realistic optical parameters. The results demonstrate the validity of the model and the feasibility of SRP-based formation and rendezvous control in orbits around small bodies and in high orbits around the Earth such as the GEO belt.     

基于有限Fourier级数的航天器转移轨迹设计

针对航天器小推力转移轨迹的初始设计问题，利用基于三阶Fourier级数的设计方法实现了航天器小推力的多圈转移。同时，基于有限Fourier级数的形状法，对具有多个约束条件的小推力多圈转移轨迹进行了优化设计。选取了共面同轴同偏心率的初始和末端轨道位置，对所提出的方法进行了仿真验证。结果表明：与改进逆五阶多项式形状法相比，所提出的方法虽然增加了转移时间，但当转移圈数为5圈、有限Fourier级数的项数为10时，可减少将近75%的转移速度增量，同时大大减小了所需的最大推力加速度的值。     

近地轨道集群航天器电磁编队飞行非线性反馈控制方法

针对近地轨道集群航天器电磁编队飞行的动力学和控制问题, 提出了一种非线性反馈控制方法. 基于电磁力模型和地磁场模型, 分析了地磁场对近地轨道电磁编队的影响; 建立了集群航天器电磁编队高精度相对轨道动力学模型; 基于Lyapunov稳定性理论设计了一种非线性反馈控制律, 利用该方法对两星电磁编队维持控制进行了仿真验证. 仿真结果表明, 地磁场引起的电磁干扰力可以忽略, 但是电磁干扰力矩的影响必须考虑; 近地轨道集群航天器电磁编队是可控的, 所设计的控制方法是可行的.      

Taurid complex meteoroids detected near aphelion with Ulysses

Between 3.4 and 4.0 AU the dust detection system aboard the Ulysses spacecraft showed an increase in detection rate for particles with masses greater than 5 × 10⁻¹³ g. The spacecraft meteoroid encounter geometry indicates highly eccentric orbits detected near aphelion. The outer limit of the enhanced flux is imposed as meteoroids on such orbits move outside the aperture of the dust detector. The inner edge of the enhanced flux would be consistent with the aphelion distance acquired by 50-200 μm particles evolving for 10-20 kyr under Poynting-Robertson drag from an Encke type orbit. We propose such meteoroids provide a source population from which collisional fragmentation produces particles in the mass range to which the Ulysses detector is sensitive. Daughter fragments produced away from the aphelia of the parent orbits, a 2.2 AU, e 0.85, enter hyperbolic orbits which are not evident in the Ulysses data. The spatial density of fragments from collisions very near aphelion drops off rapidly as they evolve inward under Poynting-Robertson drag while collisions closer to 3.4 AU leave the subsequent peak density outside that radius for a significant fraction of the fragment's subsequent lifetime. The rapid orbital evolution for these collision fragments implies a recent breakup and probably a large reservoir of parent meteoroids.     

Formation flying for magnetic sails around artificial equilibrium points

This paper investigates the problem of magnetic sail-based spacecraft formation control around the artificial equilibrium points (AEPs), which can eliminate the requirement of the propellant. The thrusts are achieved by utilizing the interaction between the solar wind and the artificial magnetosphere generated by superconducting current coil onboard. The circular restricted three-body problem (CRTBP) of magnetic sail is discussed including the allowed regions and linear stability of AEPs, the locations of collinear AEPs and the possibility of existence of periodic orbits around the collinear AEPs. Next, the dynamical models of magnetic sail formation around the collinear AEPs are established. A novel fast fixed-time nonsingular terminal sliding mode controller (FFNTSM) based on fixed-time disturbance observer (FTDO) is developed to account for external disturbances. Several numerical simulations are conducted to substantiate that spacecraft formation can be precisely controlled by the proposed propellantless propulsion method in the presence of external disturbances.     

Atmospheric reentry hemisphere prediction for prograde orbits using logical disjunction

A unique logic-based algorithm for atmospheric reentry hemisphere prediction is presented for spacecraft in low-eccentricity, prograde low Earth orbits at altitudes of 300 km and lower. Using two-line element (TLE) data for initial orbit conditions, coupled with coarse estimates for spacecraft aerodynamic characteristics, the algorithm relies on logical disjunction operations based on a dual analysis of histogram and two-weighted Gaussian probability density function (PDF) fits of predicted reentry latitude data. The algorithm requires the execution of a series of parametric simulations to determine the reentry hemisphere for variations in spacecraft aerodynamic coefficients and drag reference area. When implemented, the algorithm yields accurate hemisphere predictions on average 15 days from reentry as demonstrated by historical reentry cases from 1979 to 2018. All reentry cases were selected to demonstrate the algorithm’s ability to deliver accurate reentry hemisphere predictions for spacecraft with varying physical size and mass, and reentering during different periods of solar cycle activity.     

Connectivity preservation and collision avoidance control for spacecraft formation flying in the presence of multiple obstacles

This paper addresses connectivity preservation and collision avoidance problem of spacecraft formation flying with multiple obstacles and parametric uncertainties under a proximity graph. In the proximity graph, each spacecraft can only get the states of the neighbor spacecraft within its sensing region. Connectivity preservation of a graph means that the connectivity of the graph should be preserved at all times during spacecraft formation flying. We consider two cases: (i) the obstacles are static, and (ii) the obstacles are dynamic. In the first case, a distributed continuous control algorithm based on artificial potential function and equivalent certainty principle is proposed to account for the unknown parameters and the static obstacles. In the second case, a sliding surface combined with a distributed adaptive control algorithm is proposed to tackle the influence of the dynamic obstacles and the unknown parameters at the same time. With the distributed control algorithms, the desired formation configuration can be achieved while the connectivity of the graph is preserved and the collisions between the spacecraft and the obstacles are avoided. Numerical simulations are presented to illustrate the theoretical results.     

A guidance approach to satellite formation reconfiguration based on convex optimization and genetic algorithms

This paper presents a new approach for autonomous reconfiguration of distributed space systems, which ensures safe guidance of spacecraft formations towards the desired patterns while optimizing the total propellant consumption. The orbital transfer is reduced to the form of a convex optimization problem to guarantee rapid computation of control laws. Hence, tasks are iteratively assigned to the component platforms to detect the best reconfiguration strategy. The path-planning is entrusted to a reference satellite of the cluster, that coordinates the remaining ones by means of a procedure based on genetic algorithms. Two methods are proposed, depending on the organizational architecture of the spacecraft formation. In the first one, the maneuver is completely planned by the reference satellite, that determines final tasks and control actions for the whole cluster. As an alternative to such a fully-centralized approach, a distributed version of the algorithm is proposed: tasks are sorted by the reference satellite and transfer orbits are computed by exploiting the computational resources of the whole cluster. Whatever the considered framework, both the planners ensure a safe transition of the formation towards the target geometry. Simulation results show that, when relative distances are of the order of hundreds of meters, a mean delta-v per satellite of the order of 0.1 m/s is required to reconfigure LEO clusters within one orbital period.     

航天器相对运动与Hill解的适用性分析

研究了空间飞行器编队中最具基础性的问题之一 ,即相对运动的解析表达及 Hill方程的适用条件。通过建立相对运动的通解公式 ,针对不同性质的初值深入地分析了其相对运动轨迹的本质特征 ,并给出了 Hill方程的适用条件。此外 ,文中还给出了一个新的编队设计简化公式     

Improving magnetosphere in situ observations using solar sails

Past and current magnetosphere missions employ conventional spacecraft formations for in situ observations of the geomagnetic tail. Conventional spacecraft flying in inertially fixed Keplerian orbits are only aligned with the geomagnetic tail once per year, since the geomagnetic tail is always aligned with the Earth-Sun line, and therefore, rotates annually. Solar sails are able to artificially create sun-synchronous orbits such that the orbit apse line remains aligned with the geomagnetic tail line throughout the entire year. This continuous presence in the geomagnetic tail can significantly increase the science phase for magnetosphere missions. In this paper, the problem of solar sail formation design is explored using nonlinear programming to design optimal two-craft, triangle, and tetrahedron solar sail formations, in terms of formation quality and formation stability. The designed formations are directly compared to the formations used in NASA’s Magnetospheric Multi-Scale mission.