Small spacecraft formation using potential functions

A group of small spacecraft able to change its orbital formation through using the potential function is discussed. Spacecraft shapes, sizes, and maneuvering capabilities in general are not identical. All objects are assumed to maneuver under discrete thruster effects. A hyperbolic form of attractive potential function is then used to reduce the control intervention by using the natural orbital motion for approaching goal configuration. A superquadric repulsive potential with 3D rigid object representation is then used to have more accurate mutual sensing between objects. As spacecraft start away from their goals, the original parabolic attractive potential becomes inefficient as the continuous control force increases with distance linearly. The hyperbolic attractive potential offers good representation of the control force independent of the distance to goal, ensuring global stability as well.     

The Investigation of Possible Approaches of Cataloged Space Objects to Manned Spacecraft

Initial orbital parameter errors are used to examine the miss distance between a spacecraft and an ensemble of tracked objects by a Monte Carlo-type analysis. The radial separation between orbits is evaluated and a keep-out zone is determined, which reduces the risk of collision to an acceptable level.An operational prediction methodology is suggested based on a catalog database, which identifies potentially hazardous approaches and computes the probability of collision for selected spacecraft. An example for the Mir Space Station is presented, which estimates the collision probability and the cross-sectional flux of cataloged objects for the time frame of interest. The results appear to be in good agreement with those of other space debris models.     

A Proposal for Constructing a Group of Space Objects for Solving Transport and Energy Problems

We propose to develop a scheme for a group of space objects which includes a set of orbital tethers and exchange masses. These objects are injected into circumterrestrial orbits. A variety of problems can be solved with the help of this space system, namely, transport problems and problems of converting the electric power generated onboard spacecraft into mechanical energy of motion of the space objects. In the future, natural celestial bodies (in particular, the Moon) can be considered as elements of the system. This opens up the possibility of using the energy of motion of the Moon both for solving transport problems and for generating electric power onboard spacecraft.     

6-DOF robust adaptive terminal sliding mode control for spacecraft formation flying

This paper addresses the tracking control problem of the leader–follower spacecraft formation, by which we mean that the relative motion between the leader and the follower is required to track a desired time-varying trajectory given in advance. Using dual number, the six-degree-of-freedom motion of the follower spacecraft relative to the leader spacecraft is modeled, where the coupling effect between the translational motion and the rotational one is accounted. A robust adaptive terminal sliding mode control law, including the adaptive algorithms, is proposed to ensure the finite time convergence of the relative motion tracking errors despite the presence of model uncertainties and external disturbances, based on which a modified controller is furthermore developed to solve the dual-equilibrium problem caused by dual quaternion representation. In addition, to alleviate the chattering, hyperbolic tangent function is adopted to substitute for the sign function. And by theoretical analysis, it is proved that the tracking error in such case will converge to a neighborhood of the origin in finite time. Finally, numerical simulations are performed to demonstrate the validity of the proposed approaches.     

Space escalator,semi-perpetual motion in space

A novel orbital transfer system using the mutual action-reaction of two spacecraft as the main propulsive force is proposed. A set of space platforms which are deployed on circular orbits of various altitude and have mobility to transfer to other orbits by themselves can work on a very little propellant.     

环月飞行器返回轨道设计与优化

针对传统月地返回轨道设计方法在从初始轨道直接转移的局限性，提出一种基于Lambert问题的月地转移轨道设计方法。该方法重新选取了轨道约束参数，并在笛卡尔坐标系下建立了返回轨道动力学模型，并基于该模型全面分析了返回轨道关于约束参数的变化特性。此外，针对带有约束条件的返回轨道优化设计问题，提出了快速收敛的函数构造法，并给出了有效的目标函数形式。以某一在轨环月飞行器的轨道参数作为初始轨道进行返回轨道计算，结果验证了所提出返回轨道设计方法以及函数构造法的有效性。     

基于智能Agent的航天器MEMS 自主热控系统研究

以航天器MEMS热控系统为对象,将系统层次的Agent智能体决策体系与热控系统自主控制任务相结合,形成一种无需外界过多干预的自主决策控制机制,即系统自身能依据所辨识出的外部轨道热环境及内部热负荷变化进行控制变量自主调节,达到优化协调多个控制变量且能自适应调节控制器参数的新型智能热控系统。     

Lorentz-force-perturbed orbits with application to J2-invariant formation

The Lorentz force acting on an electrostatically charged spacecraft in the Earth's magnetic field provides a new propellantless means for controlling a spacecraft's orbit. Assuming that the Lorentz force is much smaller than the gravitational force, the perturbation of a charged spacecraft's orbit by the Lorentz force in the Earth's magnetic field, which is simplified as a titled rotating dipole, is studied in this article. Our research starts with the derivation of the equations of motion in geocentric equatorial inertial Cartesian coordinates using Lagrange mechanics, and then derives the Gauss variational equations involving Lorentz-force perturbation using a set of nodal inertial coordinates as an intermediate step. Subsequently, the approximate averaged changes in classical orbital elements, including single-orbit-averaged and one-day-averaged changes, are obtained by employing orbital averaging. We have found that the approximate analytic one-day-averaged changes in semi-major axis, eccentricity, and inclination are nearly zero, and those in the other three angular orbital elements are affected by J₂ and Lorentz-force perturbations. This characteristic is applied to model bounded relative orbital motion in the presence of the Lorentz force, which is termed Lorentz-augmented J₂-invariant formation. The necessary condition for J₂-invariant formation is derived when the chief spacecraft's reference orbit is either circular or elliptical. It is shown that J₂-invariant formation is easier to implement if the deputy spacecraft is capable of establishing electric charge. All conclusions drawn from the approximate analytic solutions are verified by numerical simulation.     

Nonlinear feedback control of low-thrust orbital transfer in a central gravitational field

This paper investigates the problem of continuous-thrust orbital transfer using orbital elements feedback from a nonlinear control standpoint, utilizing concepts of controllability, feedback stabilizability and their interaction. Gauss's variational equations (GVEs) are used to model the state-space dynamics of motion under a central gravitational field. First, the notion of accessibility is reviewed. It is then shown that the GVEs are globally accessible. Based on the accessibility result, a nonlinear feedback controller is derived which asymptotically steers a spacecraft form an initial elliptic orbit to any given elliptic orbit. The performance of the new controller is illustrated by simulating an orbital transfer between two geosynchronous Earth orbits. It is shown that the low-thrust controller requires less fuel than an impulsive maneuver for the same transfer time. Closed-form, analytic expressions for the new orbital transfer controller are given. Finally, it is proven, based on a topological nonlinear stabilizability test, that there does not exist a continuous closed-loop controller that can transfer a spacecraft onto a parabolic escape trajectory.     

Use of Optical Characteristics to Identify Orbiting Material

Knowledge of the observable properties of orbital debris is necessary to validate debris models for both the low Earth orbit (LEO) and the geosynchronous Earth orbit (GEO). Current methods determine the size and mass of orbital debris based on knowledge or assumption of the material type of the piece. Improvement in the knowledge of material is the goal of the research described herein. The process of using spectral absorption features to determine the material type is explored. A review of the optical measurements of orbital debris as well as current research in the area is discussed. Reflectances of common spacecraft materials are compared. The need for, and advances made possible by obtaining real data are explored. The prospects of the venture are investigated.     

远程导引脉冲变轨方案的有限推力修正

对远程导引脉冲变轨方案的有限推力修正进行了研究。将多脉冲变轨方案设计结果转换为有限推力式并进行修正,以消除转换误差和摄动模型误差。分析了由转换算法和摄动模型产生的初值误差,比较了不同工况下的修正效果。算例表明:将初始多脉冲变轨方案转换为有限推力后进行修正,所得结果能实现真实飞行环境中追踪航天器对目标航天器的拦截或交会。     

航天器自主控制与智能信息处理技术

首先从系统的体系结构出发 ,将自主控制系统划分为分层递阶型、包容型、混合型三种基本形式和多智能代理结构的复杂形式 ,然后分别讨论对一般轨道飞行航天器、多个航天器所组成的空间网络、星球表面探测系统和空间机器人等几类不同对象进行自主控制的特点 ,最后论述自主控制对智能信息处理技术的基本需求并提出有关的关键技术     

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.     

在轨物体接近算法研究

根据NORAD公布的在轨物体(空间碎片或航天器) 的轨道数据,给出了计算在轨物体两两间发生接近事件的方法。该方法结合了计算在轨物体 接近的运动学筛选法和相对距离函数法,以提高计算效率。以该方法为基础,提出了特定航 天器与全体在轨物体间的接近事件算法。算法采用了改进后的几何筛选法。该方法并不求解 两轨道的最近点,而是求解包含最近点的时间窗口,从而解决了原方法漏报接近事件的问题 。该方法在理论上可以检测到所有的接近事件,数值仿真验证支持了该结论。在计算效率 上新方法相对于原算法也有约10倍的提高。     

主从式编队航天器连通性保持与碰撞规避

针对主从式航天器编队过程中存在的通信距离约束、航天器之间的碰撞以及空间干扰等问题，提出一种基于非线性干扰观测器和人工势函数的分布性协同控制方法。当初始通信网络连通时，通过在分布式协同控制器中引入吸引势函数，保证整个编队过程中通信网络始终是连通的。针对主航天器速度仅有部分从航天器直接可知的情况，为每一个从航天器设计分布式的速度观测器估计主航天器的速度，从而实现航天器之间的速度协同。此外，通过在控制器中引入非线性干扰观测器对外界干扰进行观测，显著增强了航天器编队的精度。仿真结果表明，本文提出的分布式协同控制方法不但能够实现对主航天器的速度跟踪以及航天器之间的队形保持，而且能够在编队过程中实现通信网络的连通性保持和航天器之间的碰撞规避。     

航天器机械臂柔性力控辅助装配方法

文章从装配的实际需求出发,提出一种航天器机械臂柔性力控辅助装配方法:通过在机械臂末端法兰与负载之间安装的六维力传感器感知作用力与力矩信息,人手直接作用于机械臂末端的负载,系统通过负载的重力补偿算法获得人手作用的力与力矩信息,而后以力/力矩信息作为输入来控制机械臂进行移动或转动,使负载柔性跟随人手运动。文章给出了装配方法的详细设计方案与相关算法,并进行了初步试验验证。试验表明,该装配方法可以有效提高在航天器狭小空间内进行大重量工件安装的效率,且安全可靠。     

考虑输入饱和的多航天器系统姿轨耦合分布式协同跟踪控制

基于一致性理论,研究了多航天器系统相对轨道及姿态耦合的分布式协同控制问题。在仅有部分跟随航天器可获取领航航天器信息的情形下,针对各跟随航天器存在未建模动态以及外部环境干扰等问题,利用双曲正切函数的性质,提出了考虑输入饱和的分布式自适应协同控制律。首先,对于领航航天器具有时变状态的情形,为每个跟随航天器设计了3个滑模估计器,对领航航天器的状态进行估计。其次,针对跟随航天器间相对速度和角速度难以测量的问题,设计仅需领航航天器状态的切比雪夫神经网络自适应更新律。最后,设计考虑输入饱和的分布式自适应协同控制律保证各跟随航天器跟踪动态领航航天器。仿真结果表明了该算法的有效性、可行性。     

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。     

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.