系统参数对电动力绳系动力学的影响

将电动力绳系(EDT)的主星质量、子星质量、绳系质量以及绳系中的电流视为系统参数，研究这些参数对系统的摆动动力学和轨道动力学的影响。哑铃模型下的电动力绳系摆动动力学方程存在不稳定的周期解，通过Floquet理论来衡量周期解的不稳定程度，从而研究各系统参数对摆动动力学的影响。建立了用春分点轨道元素的形式描述的电动力绳系轨道动力学方程，并以降轨时间来衡量电动力绳系的降轨效率，从而研究系统参数对轨道动力学的影响。运用算例对周期解迁移矩阵的特征值、降轨时间随各系统参数的变化关系进行了仿真，分别得出了各系统参数对系统摆动动力学和轨道动力学的影响。综合本文的仿真结果，并考虑实际发射及空间运行中的其它因素，对电动力绳系的设计和降轨策略提出了建议。     

火箭末级离轨系统中电动力绳系释放条件分析与控制设计

绳系释放过程是电动力绳系火箭末级离轨技术的关键,开展了绳系释放条件分析及控制方案设计。首先,基于"珠式"思想,建立离轨系统绳系释放过程的刚-柔耦合动力学模型。随后,基于数值仿真开展绳系释放条件分析。仿真结果分析发现,释放时火箭末级初始姿态对绳系释放稳定性影响较大,仅靠绳系拉力控制,系统稳定裕度较小,很难保证绳系释放稳定性。据此,进一步提出了一种绳系稳定释放控制方案,并通过数值仿真验证了控制方案的有效性。     

Experiments and numerical simulations of an electrodynamic tether deployment from a spool-type reel using thrusters

The amount of space debris is ever increasing, and pollution of the space environment has become a serious problem that can no longer be ignored. Consequently, the active removal of large space debris from crowded economically useful orbits should begin as soon as possible. The Japan Aerospace Exploration Agency has been investigating an active debris removal system that employs highly efficient electrodynamic tether (EDT) technology for orbital transfer. This study investigates the tether deployment from a spool-type reel using thrusters by means of numerical simulations of an EDT system. The thrusters are used in order to ensure the deployment of a tether with the length of several kilometers. In the simulations using a multiple mass tether model, the key parameters are estimated from various on-ground experiments. By means of the numerical simulations, the dynamics of tether deployment is studied and requirements of thruster needed for the deployment, such as the thrust forces and the periods of thruster activation, are clarified.     

Qualification and in-flight demonstration of a European tether deployment system on YES2

This paper highlights the design, qualification and mission performance of the tether deployer system on the second Young Engineers’ Satellite (YES2), that featured a tethered momentum transfer. The deployer is designed with a broad range of near-term tether applications in mind. The system contains the tether, including features to enhance safety and wound up in controlled manner onto a spool core, optical deployment sensors, a “barberpole” friction brake controlled by a stepper motor and a triple tether cutter system. To initiate the deployment a spring-based ejection system was developed, and to apply accurate momentum transfer a timer and release system is present on the subsatellite side. A small, 6 kg re-entry capsule was developed as subsatellite. On September 25th, 2007, YES2 deployed a 32 km tether in orbit and gathered a wealth of data. Confidence is gained from the mission results for use of the deployer in future missions.     

大气阻力摄动下的绳系辅助离轨系统的相对姿态跟踪控制研究

以绳系辅助离轨系统为背景,考虑到大气阻力摄动和子星--返回舱的非质点因素,返回舱与绳系辅助离轨系统的相对姿态将产生大幅周期性变化,这将影响或破坏系统的稳定性;同时三自由度下绳系系统的展开控制和绳系系统横向振动抑制控制也对返回舱姿态的稳定性和精度提出了更高的要求.基于此,本文建立了大气阻力摄动下的绳系系统的展开动力学和返回舱姿态动力学模型;并在此基础上,设计了绳系辅助离轨系统的相对姿态跟踪控制策略.通过数学仿真来验证大气阻力摄动下该姿态跟踪控制算法的有效性,结果表明,该控制律能够有效控制绳系辅助离轨系统的相对姿态,满足展开控制的需要.     

利用轨道瞬时根数预报的离轨制动控制方法

针对再入飞行器离轨制动问题，在考虑地球引力J2项摄动及有限推力影响下，设计了一种航天器自主离轨制动控制算法。该算法根据再入点状态约束，确定了离轨过渡轨道的平均轨道根数及其与离轨待命轨道平均轨道根数的关系，从而得到制动参数初值。通过在线数值递推轨迹，实时预报再入点瞬时轨道根数并计算再入点航迹倾角，当预报的航迹倾角满足约束条件时结束制动，并根据再入点纬度幅角误差修正制动起始点，从而修正制动参数。制动过程中，在考虑了J2项摄动影响下实时预报再入点瞬时轨道根数，依据实际任务需求确定关机时机。最后通过考虑初始状态误差、质量误差、推力误差以及姿态误差情况下的蒙特卡洛打靶仿真，分析了不同关机策略的落点散布特性，检验了该算法的自主决策和高精度再入点控制能力。     

Tethers and debris mitigation

In recent years, the use of tethers has been proposed for reduction of space debris either through momentum transfer or use of electrodynamic effects. Tethers have been shown to at least theoretically allow for quick, elegant and cost-effective deorbit of defunct satellites or spent stages. On the other hand, the large risk that tethers themselves may pose to other satellites in orbit has been recognized as well. The large collision area of tethers, combined with operational hazards and meteoroid risk may result in a large orbital exposure. For example, in 1997, the ESA/Dutch 35-km tether deployment of YES from TEAMSAT was inhibited after an analysis of the collision risk for the case the tether operation would fail. The question rises how these two points of view compare to eachother. This paper intends to highlight a representative selection of the proposed tether applications while taking into account the added risks caused by the tethers themselves.Typical applications from recent literature will be briefly described, such as an Ariane 502 spent stage re-entry from GTO and the concept of deboost of defunct satellites by interaction of a conductive tether with the Earth magnetic field.Mass savings of the tethered sytems versus conventional equivalents will be evaluated.Based on a crude risk analysis, involving elements such as mission complexity, dynamic stability, meteoroid risk and orbital life time, a general outline of limiting factors can be given for the various applications. Special attention is reserved for implementation of mechanisms that help reduce this tether risk, such as the DUtether (Tether Degradable by Ultraviolet), utilization of airdrag and solar pressure, the effect of residual current in bare tethers, tether retrieval etc.It is proposed how a net tether-induced mitigation can be compared to that of conventional alternatives, i.e. deboost by rocket engine or a completely passive approach.This comparison is put in the perspective of an ever-increasing occupation of the space environment.It is concluded that tethers can in fact help mitigate the debris risk and that for each application a useful niche can be defined. It is argued that eliminating pollution directly after use of the precious resource of space is not only good custom, but also an important way to make the risk of debris controllable and independent of future trends. Although tethers may have large exposure in terms of area-time product, they deliver a quick cleaning service that may be appreciated by the future users of space.     

Sounding rocket experiment of bare electrodynamic tether system

An overview of a sounding rocket, S-520-25th, project on space tether technology experiment is presented. The project is prepared by an international research group consisting of Japanese, European, American, and Australian researchers. The sounding rocket will be assembled by the ISAS/JAXA and will be launched in the summer of 2009. The sounding rocket mission includes two engineering experiments and two scientific experiments. These experiments consist of the deployment of bare electrodynamic tape tether in space, a quick ignition test of hollow cathode system in space, the demonstration of bare electrodynamic tether system in space, and the test of the OML (orbital-motion-limit) current collection theory.     

Magnetobraking for Mars return vehicles

It is proposed that magnetobraking may be used to dissipate hyperbolic excess velocity from a spacecraft returning from Mars to Earth orbit. In magnetobraking, an electrodynamic tether is deployed from the spacecraft. The Earth's magnetic field produces a force on electrical current in the tether, which can be used to either brake or accelerate the spacecraft without expenditure of reaction mass. The peak acceleration on the Mars return is 0.007 m/s², and the amount of braking possible is dependent on the density and current-carrying capacity of the tether, but is independent of length. Since energy is produced as the spacecraft velocity decreases, no on-board power source is required. As the spacecraft approaches the Earth, the magnetic field increases and the power produced by the tether increases, reaching a maximum of about 800 W per kg of spacecraft mass at closest approach.     

Optimal control of a spinning double-pyramid Earth-pointing tethered formation

The dynamics and control of a tethered satellite formation for Earth-pointing observation missions is considered. For most practical applications in Earth orbit, a tether formation must be spinning in order to maintain tension in the tethers. It is possible to obtain periodic spinning solutions for a triangular formation whose initial conditions are close to the orbit normal. However, these solutions contain significant deviations of the satellites on a sphere relative to the desired Earth-pointing configuration. To maintain a plane of satellites spinning normal to the orbit plane, it is necessary to utilize “anchors”. Such a configuration resembles a double-pyramid. In this paper, control of a double-pyramid tethered formation is studied. The equations of motion are derived in a floating orbital coordinate system for the general case of an elliptic reference orbit. The motion of the satellites is derived assuming inelastic tethers that can vary in length in a controlled manner. Cartesian coordinates in a rotating reference frame attached to the desired spin frame provide a simple means of expressing the equations of motion, together with a set of constraint equations for the tether tensions. Periodic optimal control theory is applied to the system to determine sets of controlled periodic trajectories by varying the lengths of all interconnecting tethers (nine in total), as well as retrieval and simple reconfiguration trajectories. A modal analysis of the system is also performed using a lumped mass representation of the tethers.     

绳系卫星系统中的周期运动

在椭圆轨道上飞行的绳系卫星系统的重要特征是系绳的方向角具有周期性运动，称为驻环。不稳定的驻环导致全系统的不稳定性。本文给出系统运动的数学模型、距离速率控制方法、求解驻环的数值迭代算法、驻环的稳定性和吸引域以及控制参数平面上驻环存在的区界和计算机仿真结果     

Benefits and risks of using electrodynamic tethers to de-orbit spacecraft

By using electrodynamic drag to greatly increase the orbital decay rate, an electrodynamic space tether can remove spent or dysfunctional spacecraft from low Earth orbit (LEO) rapidly and safely. Moreover, the low mass requirements of such tether devices make them highly advantageous compared to conventional rocket-based de-orbit systems. However, a tether system is much more vulnerable to space debris impacts than a typical spacecraft and its design must be proved to be safe up to a certain confidence level before being adopted for potential applications. To assess space debris related concerns, in March 2001 a new task (Action Item 19.1) on the “Potential Benefits and Risks of Using Electrodynamic Tethers for End-of-life De-orbit of LEO Spacecraft” was defined by the Inter-Agency Space Debris Coordination Committee (IADC). Two tests were proposed to compute the fatal impact rate of meteoroids and orbital debris on space tethers in circular orbits, at different altitudes and inclinations, as a function of the tether diameter to assess the survival probability of an electrodynamic tether system during typical de-orbiting missions. IADC members from three agencies, the Italian Space Agency (ASI), the Japan Aerospace Exploration Agency (JAXA) and the US National Aeronautics and Space Administration (NASA), participated in the study and different computational approaches were specifically developed within the framework of the IADC task. This paper summarizes the content of the IADC AI 19.1 Final Report. In particular, it introduces the potential benefits and risks of using tethers in space, it describes the assumptions made in the study plan, it compares and discusses the results obtained by ASI, JAXA and NASA for the two tests proposed. Some general conclusions and recommendations are finally extrapolated from this massive and intensive piece of research.     

TETHERED SATELLITE SYSTEM ANALYSIS (1) — TWO-DIMENSIONAL CASE AND REGULAR DYNAMICS

The study on tethered satellite system (TSS) in two-dimensional in-planar motion is restricted in that the tether is assumed to be massless. The equations of motion are given in a spherical coordinate system to describe the magnitude (tether length) and direction angle of the position vector between the satellites. A length rate control algorithm is adopted, and the controlled motion of the directional angle by the algorithm will have a stable equilibrium state. The equilibrium state is a fixed point if the orbit of the base-satellite is circular, and a limit cycle if the orbit is elliptic. The value and stability of the equilibrium state are determined by the parameters of the control algorithm, and the bifurcation analysis is also given. Two typical TSS missions have been simulated.     

Orbital and relative motion of a tethered satellite system

A simplified model for the orbital and relative motion of a tethered satellite system is presented. The tether acts as a light elastic string with small structural damping but without bending stiffness. Its mass is taken into account in the calculation of the total kinetic and potential energies of the tethered system. This formulation allows the inclusion of the complete gravity gradient influence on the dynamics of the system. The resulting three-dimensional motion is separated in the centre of mass orbital motion on the one hand and the relative motion of the end-bodies on the other. No restrictions on length of the tether or on mass ratio of the end-masses are imposed. It is found that the frequencies and amplitudes of the longitudinal tether oscillations are realistic as long as the tether mass is less than that of the subsatellite.     

Results of the parabolic flight tests of the rapunzel deployer

The tether assisted re-entry of small payloads is a highly interesting tool for space transportation especially for the return of small payloads from Space Station ISSA. The small tether mission Rapunzel was initiated in 1991 by the Institute of Astronautics, TU München and the Kayser-Threde Company, to design a low cost and feasible tether experiment for the verification of the tether assisted re-entry. Together with the Samara State Aerospace University, Russia, a mission concept on a Russian Resurs or Photon capsule was developed. Based on this mission a deployer has been designed, mainly based on technology of the textile industry, which insures high reliability at low cost. Recently a similar configuration is being discussed for the ESA-TSE mission.The main work during the recent time was the development and test of the breadboard model of the deployer system. After successfully completing initial ground tests with the deployer, further tests during the ESA Parabolic Flight campaign in November 1995 were conducted. After a short introduction of the overall mission scenario, the planned configuration in orbit, this paper will present the results of the microgravity test campaign onboard the KC-135 aircraft and compare them with the ground test. The deployer showed a good performance during all tests, including ejection of the end-mass, deployment, and braking. Problems that occurred during the tests will be discussed, and solutions for the detected flaws and the results of the redesign now in progress will be presented. These verifications have shown the feasibility of the concept and will lay the base for the planned development of the flight model of the deployer.     

空间碎片绳网捕获拖曳恒张力控制

针对空间绳网系统捕获空间碎片后，在轨道转移过程中的精确控制问题，提出一种使用常值拉力将空间碎片拖曳至坟墓轨道的方法。首先，采用牛顿欧拉法建立绳系组合体动力学模型；其次，通过李雅普诺夫方法证明仅使用恒张力即可实现拖曳过程的稳定控制；再次，提出采用常值拉力切换控制律抑制空间碎片的姿态章动，采用基于相平面控制原理的控制律抑制绳系组合体面内面外摆动，规避在轨道转移过程中系绳松弛造成缠绕、系绳张力过大造成断裂或两星接近发生碰撞等风险。最后，通过拖曳离轨全过程仿真分析，校验了所提出控制方法的有效性。     

辐射开环空间绳系机器人编队动力学及控制

提出了一种新型辐射开环空间绳系机器人编队系统,其在编队稳定性、任务灵活性及燃料消耗等方面具有明显的优势。针对辐射开环空间绳系机器人编队系统自旋运动过程中的构型误差控制问题,首先建立了编队系统的自旋动力学模型;然后分析了空间绳系机器人的绳系拉力和空间平台的自旋扭矩对编队系统自旋运动中出现的构型误差的控制能力;设计了一种依靠空间绳系机器人绳系拉力和空间平台自旋扭矩作为控制量,对构型误差进行控制的协调控制方法;最后通过数字仿真进行了校验和分析。仿真结果表明：设计的协调控制方法能够明显改善编队系统自旋运动中构型误差的控制效果。     

电动绳系在低极轨卫星中的应用研究

低极轨卫星具有轨道周期短、对地观测分辨率高等优点,但由于所在轨道大气阻力大,其使用寿命受到较大限制。文章提出采用水平结构电动绳系抵消低极轨卫星大气阻力的方法,通过系绳电流与地球磁场相互作用产生洛仑兹力进行推进,进而在无燃料消耗的情况下实现对低极轨卫星轨道高度的维持。初步分析了该方法在低极轨不同尺寸卫星中的应用潜力,计算了160 、400 和800 km 典型高度低极轨卫星所经历的地球磁场、电离层和高层大气环境相关参数变化,比较了不同条件下电动绳系推力与大气阻力大小随轨道位置的变化。分析结果表明,该方法适用于400 km 轨道高度以上大卫星;在满足一定系绳长度和轨道高度的条件下,电动绳系可以有效延长低极轨卫星的轨道寿命。     

电动系绳离轨装置的设计与分析

电动系绳离轨效果显著,无需消耗推进剂,是处理LEO(低地球轨道)废弃航天器的有效手段。文章介绍了利用电动系绳进行离轨的原理,提出了电动系绳离轨装置所需的硬件及其设计要求;给出了电动系绳离轨装置的控制策略和工作过程,并利用系绳运动方程对电动系绳离轨的效能进行了分析。结果显示,利用电动系绳可以大大减小离轨的面积时间积。     

美国小行星俘获任务及其启示

小行星俘获(ACR)任务是美国Keck空间研究中心发起的一项深空探测任务。该任务计划选定一颗近地小行星,通过口袋式抓捕系统对其实施抓捕,并于2025年左右将其带回近月空间。文章介绍了ACR任务的内容和系统设计,具体包括:航天器总体构型、抓捕分系统、探测识别分系统和控制与推进分系统;对小行星抓捕的目标探测与识别、旋转匹配、抓捕、消旋、轨道转移等核心操作。基于ACR任务,提出了空间目标俘获技术的需求与应用、抓捕航天器系统设计的启示;基于我国目前的技术研究情况,总结分析了发展空间目标俘获任务所需的关键技术,如大功率柔性太阳翼、长时间大范围轨道机动、目标探测与识别、快速机动、目标抓捕与消旋。