A universal on-orbit servicing system used in the geostationary orbit

The geostationary orbit (GEO), a unique satellite orbit of the human beings, is a very precious orbit resource. However, the continuous increasing of GEO debris makes the GEO orbit more and more crowded. Moreover, the failures of GEO spacecrafts will result in large economic cost and other bad impacts. In this paper, we proposed a space robotic servicing system, and developed key pose (position and orientation) measurement and control algorithm. Firstly, the necessity of orbit service in GEO was analyzed. Then, a servicing concept for GEO non-cooperative targets was presented and a universal space robotic servicing system was designed. The system has a 2-DOF docking mechanism, a 7-DOF redundant manipulator and a set of stereo vision, in addition to the traditional subsystems of a spacecraft. This system can serve most existing satellites in GEO, not requiring specially designed objects for grappling and measuring on the target. The servicing contents include: (a) visual inspecting; (b) target tracking, approaching and docking; (c) ORUs (Orbital Replacement Units) replacement; (d) Malfunctioned mechanism deploying; (e) satellites life extension by taking over its control, or re-orbiting the abandoned satellites. As an example, the servicing mission of a malfunctioned GEO satellite with three severe mechanical failures was designed and simulated. The results showed the validity and flexibility of the proposed system.     

中低轨红外预警星座基于自由段跟踪的单星覆盖性能

推导了中低轨天基红外星座中单颗卫星基于自由段跟踪时的空域覆盖体积与星下黑体积,研究了其随轨道高度变化的规律,分析了存在切地角影响时的体积变化,并进行了计算。     

Resonant quasi-periodic near-rectilinear Halo orbits in the Elliptic-Circular Earth-Moon-Sun Problem

Motivated by the near-future re-exploration of the cislunar space, this paper investigates dynamical substitutes of the Earth-Moon’s resonant Near-Rectilinear Halo Orbits (NRHOs) under the Elliptic-Circular Restricted Four-Body Problem formulation of the Earth-Moon-Sun system. This model considers that the Earth and Moon move in elliptical orbits about each other and that a third body, the Sun, moves in a circular orbit about the Earth-Moon barycenter. By making use of this higher-fidelity dynamical model, we are able to incorporate the Sun’s influence and the Moon’s eccentricity, two of the most significant perturbations of the cislunar environment. As a result of these perturbations, resonant periodic NRHOs of the Earth-Moon Circular Restricted Three-Body Problem (CR3BP) are hereby replaced by two-dimensional quasi-periodic tori that better represent the dynamical evolution of satellites near the vicinity of the Moon. We present the steps and algorithms needed to compute these dynamical structures in the Elliptic-Circular model and subsequently assess their utility for spacecraft missions. We focus on the planned orbit for the NASA-led Lunar Gateway mission, a 9:2 synodic resonant L₂ southern NRHO, as well as on the 4:1 synodic and 4:1 sidereal resonances, due to the proximity to the nominal orbit and their advantageous dynamical properties. We verify that the dynamical equivalents of these orbits preserve key dynamical attributes such as eclipse avoidance and near-linear stability. Furthermore, we find that the higher dimensionality of quasi-periodic solutions offers interesting alternatives to mission designers in terms of phasing maneuvers and low-altitude scientific observations.     

“用户故事”在姿轨控软件需求分析中的应用

需求分析是软件开发过程中的重要环节.在国内卫星姿轨控软件设计过程中,需求分析阶段描述和定义用户需求的工作多数仍采用传统方法,过于关注软件的设计过程,而忽略了软件需要实现的功能,常常引发需求分析结果与任务方期望不一致的情况,影响开发进度.针对姿轨控软件的特点,在软件需求分析工作中引入敏捷开发所采用的＂用户故事（＂User Story）方法,可以高效清晰地描述和定义用户所需要的软件功能,提高任务方在需求分析阶段的参与程度,显著提高需求分析的准确性.     

Assessing attitude error of FORMOSAT-3/COSMIC satellites and its impact on orbit determination

An attitude determination and control system (ADCS) is critical to satellite attitude maneuvers and to the coordinate transformation from the inertial frame to the spacecraft frame. This paper shows specific sensors in the ADCS of the satellite mission FORMOSAT-3/COSMIC (F3/C) and the impact of the ADCS quality on orbit accuracy. The selection of main POD antenna depends on the beta angles of the different F3/C satellites (for FM2 and FM4) during the inflight phase. In particular, under the eclipse, alternative attitude sensors are activated to replace the Sun sensors, and such a sensor change leads to anomalous GPS phase residuals and a degraded orbit accuracy. Since the nominal attitude serves as a reference for ADCS, the 3-dimensional attitude-induced errors in reduced dynamic orbits over selected days in 2010 show 9.35, 10.78, 4.97, 5.48, 7.18, and 6.89 cm for FM1–FM6. Besides, the 3-dimensional velocity errors induced by the attitude effect are 0.10, 0.10, 0.07, 0.08, 0.09, and 0.10 for FM1–FM6. We analyze the quality of the observed attitude transformation matrix of F3/C and its impact on kinematic orbit determination. With 249 days of GPS in 2008, the analysis leads to the following averaged 3-dimensional attitude-induced orbit errors: 2.72, 2.62, 2.37, 1.90, 1.70, and 1.99 cm for satellites FM1–FM6. Critical suggestions of geodetic payloads for the follow-on mission of F3/C are presented based on the current result.     

On the problem of optimal control of the thrust value of the electric propulsion rocket with solar energy source

Under consideration is the optimal control problem on a spacecraft motion in Newtonian central gravity field. With the use of the mathematical model of electrojet propulsion device (EPD) with solar energy source, proposed earlier in paper [1], the dependence of the EPD working substance choice on both the duration of the given dynamic maneuver and the propellant expenditures for its fulfillment is investigated. The efficiency evaluation is carrying out of optimal control of variable valued thrust as well as that for relay mode thrust and relay mode thrust with optimal fixed thrust value.     

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.