One optimization problem for trajectories of spacecraft rendezvous mission to a group of asteroids

The optimization problem is considered for the trajectory of a spacecraft mission to a group of asteroids. The ratio of the final spacecraft mass to the flight time is maximized. The spacecraft is controlled by changing the value and direction of the jet engine thrust (small thrust). The motion of the Earth, asteroids, and the spacecraft proceeds in the central Newtonian gravitational field of the Sun. The Earth and asteroids are considered as point objects moving in preset elliptical orbits. The spacecraft departure from the Earth is considered in the context of the method of a point-like sphere of action, and the excess of hyperbolic velocity is limited. It is required sequentially to have a rendezvous with asteroids from four various groups, one from each group; it is necessary to be on the first three asteroids for no less than 90 days. The trajectory is finished by arrival at the last asteroid. Constraints on the time of departure from the Earth, flight duration, and final mass are taken into account in this problem.     

Application of a magnetohydrodynamic element in the control loop of a rotating spacecraft with a spike antenna

The problem of stabilization of a rotating spacecraft with a flexible spike antenna located along the axis of spacecraft rotation is considered. A magnetohydrodynamic element is used as a final-control element in the control loop of spacecraft attitude, and the solar direction sensor serves as a measuring device. At the first stage of investigation, the problem of stability is considered for stationary and non-stationary modes of rotation of the spacecraft with a flexible antenna and with a cavity partially filled with a low-viscosity liquid.     

论航天型号项目的协作及项目群的管理

文章主要针对我国航天型号项目管理中存在的问题进行分析和研究。例如在大型环境试验和总装过程中,单一项目的管理已经日臻完善;但面对更加繁重的任务和巨大的发展机遇,单一项目的管理所暴露出的问题也越发明显。文章拟用项目群的管理思路来完善和丰富项目管理体系,从而促进环境试验和总装的项目管理水平。     

The Use of Quaternions in the Optimal Control Problems of Motion of the Center of Mass of a Spacecraft in a Newtonian Gravitational Field: II

The problem of a rendezvous in the central Newtonian gravitational field is considered for a controlled spacecraft and an uncontrollable spacecraft moving along an elliptic Keplerian orbit. For solving the problem, two variants of the equations of motion for the spacecraft center of mass are used, written in rotating coordinate systems and using quaternion variables to describe the orientations of these coordinate systems. In the first variant of the equations of motion a quaternion variable characterizes the orientation of an instantaneous orbit of the spacecraft and the spacecraft location in the orbit, while in the second variant it characterizes the orientation of the plane of the spacecraft instantaneous orbit and the location of a generalized pericenter in the orbit. The quaternion variable used in the second variant of the equations of motion is a quaternion osculating element of the spacecraft orbit. The problem of a rendezvous of two spacecraft is formulated as a problem of optimal control by the motion of the center of mass of a controlled spacecraft with a movable right end of the trajectory, and it is solved on the basis of Pontryagin's maximum principle.     

Optimal Control of a Spatial Turn of a Spacecraft with Minimal Loading of Construction

The problem of optimal control of a three-dimensional turn of a spacecraft is considered and solved. The turn is performed from an initial angular position into the required final angular position in a specified time and with a minimum value of the functional that represents the degree of loading of the construction. An analytical solution to the formulated problem is presented. It is demonstrated that the optimal (in this sense) control of the spacecraft reorientation can be determined in the class of a regular precession executed by the spacecraft. The instant when braking begins is determined based on the principles of terminal control using the actual kinematical parameters of the spacecraft motion, which substantially increases the accuracy of transferring the spacecraft to a specified position. Data of mathematical modeling are also presented that confirm the efficiency of the described method of controlling the spacecraft's three-dimensional turn.     

The Use of Quaternions in the Optimal Control Problems of Motion of the Center of Mass of a Spacecraft in a Newtonian Gravitational Field: III

The results of numerical solution of the problem of a rendezvous in the central Newtonian gravitational field of a controlled spacecraft with an uncontrollable spacecraft moving along an elliptic Keplerian orbit are presented. Two variants of the equations of motion for the spacecraft center of mass are used, written in rotating coordinate systems and using quaternion variables to describe the orientations of these coordinate systems. The problem of a rendezvous of two spacecraft is formulated [1, 2] as a problem of optimal control by the motion of the center of mass of a controlled spacecraft with a movable right end of the trajectory, and it is solved on the basis of Pontryagin's maximum principle. The paper is a continuation of papers [1, 2], where the problem of a rendezvous of two spacecraft has been considered theoretically using the two above variants of the equations of motion for the center of mass of the controlled spacecraft.     

The Use of Quaternions in the Optimal Control Problems of Motion of the Center of Mass of a Spacecraft in a Newtonian Gravitational Field: I

The problem of optimal control is considered for the motion of the center of mass of a spacecraft in a central Newtonian gravitational field. For solving the problem, two variants of the equations of motion for the spacecraft center of mass are used, written in rotating coordinate systems. Both the variants have a quaternion variable among the phase variables. In the first variant this variable characterizes the orientation of an instantaneous orbit of the spacecraft and (simultaneously) the spacecraft location in this orbit, while in the second variant only the instantaneous orbit orientation is specified by it. The suggested equations are convenient in the respect that they allow the general three-dimensional problem of optimal control by the motion of the spacecraft center of mass to be considered as a composition of two interrelated problems. In the first variant these problems are (1) the problem of control of the shape and size of the spacecraft orbit and (2) the problem of control of the orientation of a spacecraft orbit and the spacecraft location in this orbit. The second variant treats (1) the problem of control of the shape and size of the spacecraft orbit and the orbit location of the spacecraft and (2) the problem of control of the orientation of the spacecraft orbit. The use of quaternion variables makes this consideration most efficient. The problem of optimal control is solved on the basis of the maximum principle. Several first integrals of the systems of equations of the boundary value problems of the maximum principle are found. Transformations are suggested that reduce the dimensions of the systems of differential equations of boundary value problems (without complicating them). Geometrical interpretations are given to the transformations and first integrals. The relation of the vectorial first integral of one of the derived systems of equations (which is an analog of the well-known vectorial first integral of the studied problem of optimal control) with the found quaternion first integral is considered. In this paper, which is the first part of the work, we consider the models of motion of the spacecraft center of mass that employ quaternion variables. The problem of optimal control by the motion of the spacecraft center of mass is investigated on the basis of the first variant of equations of motion. An example of a numerical solution of the problem is given.     

Estimation of the Probability of Capture into a Resonance Mode of Motion for a Spacecraft during Its Descent in the Atmosphere

The estimation of the probability of capture into a resonance mode of motion is considered for a spacecraft with a small asymmetry during its entry into the atmosphere. It is assumed that the initial conditions of spacecraft motion are distributed uniformly in some sufficiently small domain. The problem is solved for the equations of spacecraft motion linear with respect to the angle of attack. An analytical estimate of the probability of the spacecraft capture into the resonance corresponding to an ascending branch of the velocity head is obtained. The emphasis in the analysis of the estimate is made on the effect of the spacecraft asymmetry type on the probability of capture. A comparison of the estimate with the results of numerical computation is carried out. A model problem concerning the construction of the domain of the spacecraft center of mass locations, most dangerous from the point of view of the realization of the stable resonant modes of motion, is solved.     

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...     

On stability of a rotating spacecraft partially filled with liquid: The case of a single cavity

The problem of stability of a rotating spacecraft with a cavity partially filled with liquid to a small depth is considered with regard to the distinction in angular velocities of spacecraft and liquid rotation and their variability (the modes of the spacecraft’s stationary rotation, spin-up, and rotation deceleration). The regions of stability (in space of the characteristic parameters of an object) are found, and mathematical simulating of the disturbed motion is carried out.     

Optimization Problem in Control over a Programmed Turn of a Spacecraft

The problem of spacecraft reorientation from its initial angular position into a desired final position within a given time interval with a minimum value of the angular moment is considered and solved analytically in this work. It is shown that the control over the spacecraft reorientation, optimal in this sense, might be defined in the class of a regular precession performed by the spacecraft. The moment of the start of deceleration is determined from the principles of the terminal control by using real kinematic parameters of apparatus motion, which increases significantly the accuracy of reorientation. The results of mathematical modeling are presented, showing a high efficiency of the proposed way of reorientation.     

Features of rotational motion of a spacecraft descending in the Martian atmosphere

Angular motion at atmospheric entry is studied in the paper for a spacecraft with a bi-harmonic moment characteristic. Special attention is given to the case when the spacecraft possesses two stable balanced positions, and, hence, it can oscillate in dense atmospheric layers in the ranges of small or large angles of attack. The averaged equations of spacecraft motion are derived, which allow one to increase the speed of calculations by several orders of magnitude. A real example is presented, which concerns a spacecraft specially designed for descending in the Martian atmosphere.     

Some issues of time-optimal control over a spacecraft programmed turn

The problem of optimal turn of a spacecraft from an arbitrary initial position to a final specified angular position in a minimum time is considered and solved. A case is investigated, when the constraint on spacecraft’s angular momentum during the turn is essential. Based on the quaternion method a solution to the posed problem has been found, and an optimal control program is constructed taking the constraints on controlling moment into account. The optimal control is found in the class of regular motions. A condition (calculation expression) is presented for determining the moment to begin braking with the use of measurements of current motion parameters, which considerably improves the accuracy of putting the spacecraft into a preset position. For a dynamically symmetrical spacecraft the solution to the problem of optimal control by the spacecraft spatial turn is presented in analytical form (expressions in elementary functions). An example of mathematical modeling of the spacecraft motion dynamics under optimal control over reorientation is given.     

编队卫星姿态的自适应协同控制

研究了无向信息交互条件下的多卫星姿态同步和跟踪问题,分别在模型参数精确已知和存在模型参数不确定性的情形下设计分散式协同控制算法。当模型参数精确已知时,基于滑动模态变量设计的协同控制力矩,实现了编队卫星姿态的同步机动,即保证编队卫星姿态的同步,并同时跟踪期望的时变参考姿态轨迹。当在轨卫星存在模型参数不确定性和常值未知扰动力矩时,提出了一种自适应协同控制律,仍能实现多卫星姿态的同步机动。基于Lyapunov-Krasovskii的分析表明,当星间链路存在任意未知的定常通讯时延时,本文设计的控制器有效。数值仿真算例验证了本文提出的控制算法。     

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.     

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。     

载人航天器空间碎片防护与减缓设计

为降低空间碎片撞击所带来的威胁,文章针对某载人航天器进行了空间碎片撞击风险分析,并设计了空间碎片专用防护结构,制定了轨道规避策略;同时针对重点设备进行钝化设计,并采用寿命末期有控离轨制动方案。以上措施可有效实现空间碎片减缓。     

Experimental Flight of a Spacecraft with Solar Sail

We analyze the prospects of launching an experimental spacecraft with a solar sail. The problems hindering the implementation of this task are discussed. The ways to solve these problems are pointed out. The feasibility of organization of a system of controlling this spacecraft with a small consumption of propellant mass is briefly investigated.     

Influence of disturbances on the angular motion of a spacecraft in the powered section of its descent

The motion of a variable-mass spacecraft is considered in the powered section of a descending trajectory. Approximate analytical solutions are obtained for the angles of spatial orientation of the spacecraft, which allows one to analyze the nutation motion and to develop recommendations on the spacecraft’s mass configuration, providing the smallest possible deviations of the longitudinal axis and thrust vector from specified directions. The errors of stabilization of the spacecraft’s longitudinal axis are calculated by means of numerical integration of complete models and using the obtained analytical solutions, the results being in good agreement.     

航天器公用空间搭载试验平台设计

我国航天器搭载试验包括空间材料实验、航天医学实验和航天新技术验证试验,具有数据量大、需要实时传输、工作模式多样等特点。文章根据试验项目的需求,设计了应用于低轨道航天器的支持各类空间搭载试验项目的公用试验平台。地面测试验证和在轨运行情况表明,该平台在不影响飞行器平台安全的同时,满足了对试验项目的工作控制、遥测监视及试验数据传输管理的任务要求。