跨流区超低轨航天器快速气动力计算方法

超低轨飞行器常工作在过渡流区中,过渡流区中气动特性多变,建模困难,目前缺少较好的快速气动力计算方法。在面元法的基础上,提出了设计迎风单元筛选算法与logistics-log桥函数,对连续流区与自由流区结果进行加权,实现超低轨卫星的气动力估算。在方法验证中,对圆柱与钝头双楔体在宽努森数(Kn)范围下的气动力进行测算,发现结果与直接模拟蒙特卡洛法(DSMC)数据一致。对典型超低轨飞行器地球重力场和海洋环流探测卫星(GOCE)的气动力进行分析,发现GOCE随着姿态变化阻力变化明显,需要舵面或陀螺仪进行姿态控制;在不同高度下的阻力量级变化大,在超低轨运行时需要动力系统维持高度。     

DOA estimation for attitude determination on communication satellites

In order to determine an appropriate attitude of three-axis stabilized communication satellites, this paper describes a novel attitude determination method using direction of arrival （DOA） estimation of a ground signal source. It differs from optical measurement, magnetic field measurement, inertial measurement, and global positioning system （GPS） attitude determination. The proposed method is characterized by taking the ground signal source as the attitude reference and acquiring attitude information from DOA estimation. Firstly, an attitude measurement equation with DOA estimation is derived in detail. Then, the error of the measurement equation is analyzed. Finally, an attitude determination algorithm is presented using a dynamic model, the attitude measurement equation, and measurement errors. A developing low Earth orbit （LEO） satellite which tests mobile communication technology with smart antennas can be stabilized in three axes by corporately using a magnetometer, reaction wheels, and three-axis magnetorquer rods. Based on the communication satellite, simulation results demonstrate the effectiveness of the method. The method could be a backup of attitude determination to prevent a system failure on the satellite. Its precision depends on the number of snapshots and the input signal-to-noise ratio （SNR） with DOA estimation.     

交会对接任务轨道控制规划设计与实施简

针对我国空间交会对接轨道控制规划技术,研究了轨道交会优化、应急轨道控制、安全轨道防护和发射窗口规划等一系列关键问题.设计了全寿命周期交会对接任务轨道控制规划方案,从目标飞行器发射到飞船返回,对轨道控制进行了全程协同、全局优化.设计了相位、高度、圆化度多目标融合控制算法;建立了规划变量对远距离导引终点六自由度的独立控制方程;设计了标称整体规划与动态逐级规划相结合的多模式规划策略;基于导引终点整体调整和局部调整的方式,实现了正常和应急条件下天地导引交接点的动态规划;提出了基于飞行控制过程建模的导引终点精度分析方法,确定了地面导引向自主导引切换的关键判据;建立了多约束交会对接发射窗口模型,构建了多任务多年度发射窗口集合.交会对接轨道控制规划技术成功应用于神舟八号、神舟九号和神舟十号交会对接任务.     

轨控标定方法研究及在交会对接中的应用简

在航天测控任务中,对轨控效果进行标定并合理利用可以实现更为精准的轨道控制目标.提出一种用虚拟等效推力来替代姿态偏差所带来的影响,利用控前控后精密轨道同时标定轨控执行过程轨道切向、径向、法向三方向速度增量的方法,介绍该方法在我国交会对接任务中的应用情况,及其标定结果对定轨精度的敏感程度,试图在理论上进一步说明利用精密轨道进行轨控标定的可行性.     

载人航天轨道大气密度模式修正研究简

热层大气密度模式的误差,是影响载人航天定轨精度的关键因素.分析载人航天工程所用Jacchia、MSISE、DTM三类大气密度模式的误差特点,通过比较精度和稳定性,基于现有空间天气参数,选取MSISE模式作为基础模式.研究利用星载加速度计数据反演载人航天轨道大气密度的方法,以验证我国载人航天轨道实测数据的精度;同时利用天宫一号以及神舟二号、三号、四号实测密度数据,以及相应的航天测控数据,分析模式误差与地方时、纬度和高度因素之间的关系,讨论建立合适的三维误差库来存储模式误差的方法,研究平均误差修正法和加权误差修正法,建立NRLMSISE-00的误差修正模式.修正结果应用于交会对接任务,与完全不修正时模式平均11.44％的误差相比较,两种修正方法的误差均明显减小,分别为5.41％和4.99％;其中平均误差修正法和加权误差修正法在未来1天、2天、3天的修正结果的误差分别是4.06％、3.73％,6.06％、5.78％,6.13％、5.72％,表明提前1天的修正效果最好;同时比较累积1-5天的误差库滑动也可以看出,误差库累积1天的效果相对较好;比较两种方法的预测效果显示,加权误差修正法优于平均误差修正法.研究表明基于三维误差库的模式修正方法显著提高了载人航天轨道大气密度预测精度,可为交会对接等载人航天任务提供技术支持.     

基于沿迹差修正部分轨道参数的两步法定轨

通过分析预报误差与轨道确定参数误差的关系,提出一种使用二次多项式对沿迹差随时间变化的函数进行拟合,并根据拟合结果修正半长轴、面质比参数误差的定轨方法.仿真计算表明：对于稀疏观测数据,该方法的处理结果优于常规的最小二乘轨道确定方法;对于高度在400 km以上的低轨目标,根据该方法得到的定轨结果,预报5d的位置误差小于22 km,与SGP4（Simplified General Perturbations Version 4,简化普适摄动4）/SDP4（Simplified Deep space Perturbations Version 4,简化深空摄动4）水平相当.该方法是一种适用于双屏电子篱笆稀疏观测数据的批量数据轨道确定方法.     

地月系L1点的月球借力间接转移轨道设计

针对地月系L1点（LL1）的轨道转移问题,在平面圆型限制性三体问题模型下,提出了利用月球借力的间接转移设计方法.转移设计分为地月转移轨道段和月球至LL1流形段.首先,通过改变LL1点初始机动速度,逆向积分LL1点的拟流形,以寻找初始速度、月心会合坐标系下的轨道高度和相位角这三者之间的关系,确定月球—LL1流形段微分改正的初始条件.然后,借助地月系不同转移时间的霍曼转移轨道所对应的近月点高度和相位角的关系,获得使2段转移在近月点相拼接的地月转移轨道段.这种设计方法给出了一系列LL1点间接转移轨道,将此设计结果与其他文献中的转移设计方法进行比较,此间接转移轨道比低能量转移轨道节省时间,比直接转移轨道节约能量.     

基于嫦娥一号高能粒子数据的地球磁层屏蔽效应研究

月球绕地球运行轨道约有1/4位于地球磁层内,因此,地球磁层是否会为月球轨道附近高能粒子提供足够的磁场屏蔽对于探索月球活动具有重要影响.嫦娥一号是中国首颗绕月人造卫星,其绕月飞行的工作轨道距离月球表面200 km.通过对嫦娥一号高能粒子探测器(HPD)的探测数据进行分析,比较了当月球位于地球磁层内外6个不同能道(能量范围4～400 MeV)时质子通量的变化,发现当月球位于地球磁层内时,这些能道的质子通量并没有发生显著减少,结果表明地球磁层不能为月球轨道附近高能粒子提供显著的磁屏蔽.     

Pulsed laser interactions with space debris: Target shape effects

Among the approaches to the proposed mitigation and remediation of the space debris problem is the de-orbiting of objects in low Earth orbit through irradiation by ground-based high-intensity pulsed lasers. Laser ablation of a thin surface layer causes target recoil, resulting in the depletion of orbital angular momentum and accelerated atmospheric re-entry. However, both the magnitude and direction of the recoil are shape dependent, a feature of the laser-based remediation concept that has received little attention. Since the development of a predictive capability is desirable, we have investigated the dynamical response to ablation of objects comprising a variety of shapes. We derive and demonstrate a simple analytical technique for calculating the ablation-driven transfer of linear momentum, emphasizing cases for which the recoil is not exclusively parallel to the incident beam. For the purposes of comparison and contrast, we examine one case of momentum transfer in the low-intensity regime, where photon pressure is the dominant momentum transfer mechanism, showing that shape and orientation effects influence the target response in a similar, but not identical, manner. We address the related problem of target spin and, by way of a few simple examples, show how ablation can alter the spin state of a target, which often has a pronounced effect on the recoil dynamics.     

Integrated orbit and attitude hardware-in-the-loop simulations for autonomous satellite formation flying

Development and experiment of an integrated orbit and attitude hardware-in-the-loop (HIL) simulator for autonomous satellite formation flying are presented. The integrated simulator system consists of an orbit HIL simulator for orbit determination and control, and an attitude HIL simulator for attitude determination and control. The integrated simulator involves four processes (orbit determination, orbit control, attitude determination, and attitude control), which interact with each other in the same way as actual flight processes do. Orbit determination is conducted by a relative navigation algorithm using double-difference GPS measurements based on the extended Kalman filter (EKF). Orbit control is performed by a state-dependent Riccati equation (SDRE) technique that is utilized as a nonlinear controller for the formation control problem. Attitude is determined from an attitude heading reference system (AHRS) sensor, and a proportional-derivative (PD) feedback controller is used to control the attitude HIL simulator using three momentum wheel assemblies. Integrated orbit and attitude simulations are performed for a formation reconfiguration scenario. By performing the four processes adequately, the desired formation reconfiguration from a baseline of 500–1000 m was achieved with meter-level position error and millimeter-level relative position navigation. This HIL simulation demonstrates the performance of the integrated HIL simulator and the feasibility of the applied algorithms in a real-time environment. Furthermore, the integrated HIL simulator system developed in the current study can be used as a ground-based testing environment to reproduce possible actual satellite formation operations.