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.     

Analysis of the dynamics of the deployed aerodynamic space tether system

An analysis of the motion of a deployed space system that consists of two end bodies connected by a tether has been considered. One of the bodies has a relatively large ballistic coefficient that ensures aerodynamic braking or the stabilization of the motion of the entire system in relatively low near-Earth orbits. The deployment of this system mainly occurs due to the action of aerodynamic forces. Several ways of deploying the system have been analyzed, including (1) the uncontrolled release of the tether with hardly any braking; (2) deployment with constant braking force; (3) the dynamic control law without feedback, when the resistance force varies according to a set program; (4) a kinematic control law with feedback when programs are set for varying the velocity and length of the tether release. To analyze the dynamics of the system, a mathematical model of motion has been constructed in which the motion of the end bodies relative to their centers of mass is taken into account.     

Motion of a descent capsule relative to its center of mass when deploying the orbital tether system

The spatial motion relative to the center of mass is considered for a capsule on an elastic tether, when it is unrolled from a spacecraft by a special program. The spacecraft is in a circular orbit and oriented relative to the local vertical, which is guaranteed by operation of its own stabilization system. Angular motion of the capsule relative to the tether direction is studied, and the main factors influencing the stability of this motion are analyzed. An approximate quasi-linear mathematical model of the capsule attitude motion is obtained, which allows one to estimate the influence of major disturbances of its motion. The results of numerical simulations are presented for characteristic cases of the capsule motion.     

绳系空间太阳能电站动力学响应分析

利用绝对节点坐标方法研究绳系空间太阳能电站在轨飞行的太阳能电池板动力响应。通过勒让德变换引入广义动量,在约束哈密尔顿体系下建立轨道、姿态和弹性振动耦合的动力学方程。基于祖冲之类方法的思想,结合辛龙格-库塔方法对微分-代数方程进行数值求解。数值算例说明本文建模方法和数值算法都是有效的,能很好地保持系统约束和能量。最后分析了绳长、平台系统的质量、轨道高度对于梁中点挠度和轴向平均应变的影响。     

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.     

绳系卫星系统的运动与控制

绳系卫星系统的运动与控制分析,是绳系卫星应用的基础。本文的数学模型考虑了作用在子星、主星及系绳上的空气阻力以及在伸展与收回阶段主星与系绳之间的质量传递,给出了相应的具体的计算式。应用该数学模型,本文对现有的几种主要控制法,进行了模拟计算,作出评审意见。     

电动力绳系研究进展

介绍了电动力绳系的主要构造及工作原理,回顾了国内外关于电动力绳系的研究历程,并总结了相关在轨试验进展情况。从绳系模型、轨道动力学以及绳系展开动力学与控制等几个方面对电动力绳系进行了分析,最后对电动力绳系技术未来的发展和应用做了展望。     

微重力环境中重力梯度加速度引起的流体晃动运动

本文研究了关于旋转轴在贮箱的非对称轴上且远离贮箱的几何中心情况下，流体在微重力环境中由重力梯度加速度诱发的晃动特性，建立了问题的数学模型并对模型进行了数值模拟。以高级Ｘ射线天文物理实验卫星（简称ＡＸＡＦ─Ｓ）作为研究对象，获得了由旋转运动引起的重力梯度加速度的数学表达式。晃动问题的数值计算以与卫星固连的非惯性坐标系为基础，目的是寻求一种较为易处理且适合于流体力学方程的边界和初始条件。通过数值计算获得了流体作用于卫星贮箱上的力和力矩。     

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.     

Geostationary station keeping control of a space elevator during initial cable deployment

This paper presents a basic concept to derive an orbital control strategy to achieve the full deployment and the geostationary station keeping of a space elevator during its initial cable deployment. The space elevator model is composed of a main spacecraft, a sinker mass and a massive cable connecting them. The cable elasticity, flexibility and taper of the cross-sectional area are omitted for simplification. A reference trajectory is designed so that the space elevator and its center of mass ascend vertically along the geostationary position with keeping the geostationary orbital rate. From the reference trajectory analyses, an orbital control that leads the space elevator orbit to the reference one is derived. However it is found that the reference trajectory is unstable throughout the deployment and a linear feedback control is introduced for stabilization. It is also clarified that the libration destabilizes the orbital control because the orbital acceleration caused by the libration always acts in the opposite direction to the orbital control. Therefore, a libration control is also introduced to stabilize the coupled orbital and librational motions. Numerical simulation result clearly shows that these controls facilitate the full deployment and the geostationary station keeping of the space elevator within the feasible thrust force and amount of propellant.     

Spacecraft dynamics with regard to elastic gravitational stabilizer deployment

The present paper deals with the study the dynamics of the spacecraft with gyro-gravitational system of stabilization. The deployment of the boom of the gravitational stabilizer commences after placing the spacecraft into the orbit and completion of the preliminary damping, when the gyroscopes are uncaged. Primarily the boom is the pre-stressed tape wound on the special drum. When the drum starts deploying the tape, it turns into the elastic cylindrical rod with the mass at its tip. The objective of the study is the creation of the generalized mathematical model and the conducting of the computer modelling of the spacecraft dynamics. The equations of motion are worked out with the use of the Lagrangian formalism. The numerical simulation of typical modes of system functioning is conducted. It is shown that the folding and the following deployment of the boom result in the turn of the spacecraft by 180° about the axis of the pitch. The results illustrate the behaviour of the main system variables.     

Space flight visual simulation.

In this paper, based on the scenes of stars seen by astronauts in their orbital flights, we have studied the mathematical model which must be constructed for CGI system to realize the space flight visual simulation. Considering such factors as the revolution and rotation of the Earth, exact date, time and site of orbital injection of the spacecraft, as well as its orbital flight and attitude motion, etc., we first defined all the instantaneous lines of sight and visual fields of astronauts in space. Then, through a series of coordinate transforms, the pictures of the scenes of stars changing with time-space were photographed one by one mathematically. In the procedure, we have designed a method of three-times "mathematical cutting." Finally, we obtained each instantaneous picture of the scenes of stars observed by astronauts through the window of the cockpit. Also, the dynamic conditions shaded by the Earth in the varying pictures of scenes of stars could be displayed.     

绳系卫星系统交会对接的应用与设计

文章介绍了绳系系统交会对接这项新技术在空间中的应用。主要包括:空间站利用系绳与航天器交会对接,实现为空间站提供各种供给;利用绳系系统与空间碎片对接,可回收或转移空间碎片,保护空间环境;利用一级或多级的绳系系统组成轨道转移系统,实现向地球同步轨道或火星轨道上转移和运送有效载荷。文章还介绍了绳系交会对接系统的设计,包括系统的一般控制方法和算法以及系统的结构设计。随着各项相关技术的发展,绳系卫星系统交会对接将发挥更大作用。     

可变构型复合柔性结构航天器动力学建模研究

针对中心刚体加复合柔性结构类航天器采用混合坐标法和子结构模态综合法，建立了可变构型复合柔性结构航天器低阶动力学模型。获得的柔性动力学方程及其各类耦合系数矩阵，适用于全星级可变构型系统和部件级复合柔性附件系统的控制系统设计与仿真，该模型具有阶数低和工程实用的特点。     

基于逼近理想解排序法的航天器初始轨道选优

针对航天器入轨段，如何从多组初始轨道中选择一组最优或最符合客观实际轨道根数问题，进行了深入地理论分析，提出了基于逼近理想解排序法的航天器初始轨道根数选优算法。描述了逼近理想解排序法的设计思想以及求解的六个基本步骤；针对航天器初始轨道的设计特点，巧妙地利用标称轨道作为理想解，偏差轨道作为负理想解，成功地应用逼近理想解排序法建立了初始轨道根数选优的决策数学模型，并以一颗典型的太阳同步轨道卫星为例，验证了其数学模型和选优算法的合理性和正确性。     

一种基于激光雷达的航天器隐藏点测量方法

为保证航天器各系统的仪器安装满足在轨姿态、轨道控制等要求,需对其安装位置及姿态进行检测,但一些隐藏的目标却无法直接测得。文章提出了一种基于激光雷达测量系统的隐藏点测量方法,针对激光雷达测量坐标系的建立以及应用高精度平面镜进行隐藏点测量进行了数学模型的分析,并对测量系统的不确定度进行了评估。结果表明,在保持测量环境稳定的情况下,应用激光雷达进行航天器隐藏点测量的标准不确定度优于0.1071 mm。     

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。     

辐射器展开角度对航天器热控能力影响的研究

为了在有限的结构尺寸下提高航天器热控系统的散热能力,提出与流体回路耦合的可展开式辐射器热控方案,建立可展开式辐射器空间散热模型,分析辐射器不同展开角度下系统的热控特性。结果表明,随着辐射器展开角度的变化,辐射器吸收的空间热流也随之发生变化,并最终决定热控回路的流量分配。在工程应用中,基于热控流体回路,通过调节可展开式辐射器的展开角度,可以有效提高航天器热控系统的能力范围。     

Resonance Effects when Moving a Small Spacecraft around the Center of Mass in a Deployable Tether System

Cosmic Research - The resonance motions of a small spacecraft relative to the center of mass when deploying a tether system are analyzed. The tether system is deployed from a base spacecraft moving...     

航天飞行器相对运动动力学及其控制方法

采用本地轨道从标系对两邻飞船间相对运动动力学展开研究，指出在相对运动中也存在平衡状态，用相平面方法分析了其稳定性，基于此分析，综合出相对运动控制方法，即距离速率控制方法，受控运动轨迹是一条稳定的稳态直线，进而建立了全方位距离速率控制方法。最后以系绳卫星系统和飞船安全为例完成了计算机模拟。