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.     

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.     

Optimization of transfer trajectories to the Apophis asteroid for spacecraft with high and low thrust

The problem of optimization of a spacecraft transfer to the Apophis asteroid is investigated. The scheme of transfer under analysis includes a geocentric stage of boosting the spacecraft with high thrust, a heliocentric stage of control by a low thrust engine, and a stage of deceleration with injection to an orbit of the asteroid’s satellite. In doing this, the problem of optimal control is solved for cases of ideal and piecewise-constant low thrust, and the optimal magnitude and direction of spacecraft’s hyperbolic velocity “at infinity” during departure from the Earth are determined. The spacecraft trajectories are found based on a specially developed comprehensive method of optimization. This method combines the method of dynamic programming at the first stage of analysis and the Pontryagin maximum principle at the concluding stage, together with the parameter continuation method. The estimates are obtained for the spacecraft’s final mass and for the payload mass that can be delivered to the asteroid using the Soyuz-Fregat carrier launcher.     

Chaotic oscillations of spacecraft with an elastic radially oriented tether

A mechanical system consisting of a spacecraft with weightless elastic tether and load is considered in the paper. The spacecraft motion under the action of tensile force of a radially oriented tether is investigated. It is shown that elastic tether oscillations can result in appearance of chaotic modes of spacecraft motion. By means of the Melnikov method a condition is obtained, allowing one to determine the measure of damping sufficient for prevention of these chaotic modes. The influence of system’s mass-geometric and elastic characteristics on the form of phase portrait and on the value of periodic disturbance, caused by oscillations of the elastic vertical tether, is studied.     

About one problem of trajectory optimization

The optimization problem for trajectories of spacecraft flight from the Earth to an asteroid is considered in this paper. The flight is realized in the central Newtonian gravitational field of the Sun with a possibility of gravitational maneuvers near planets. Perturbation maneuvers are taken into account using the method of point area of action with a limitation on the flyby altitude. The spacecraft is controlled by changing the value and direction of the engine thrust. The problem is solved taking into account constraints on the launch time, flight duration, and minimum distance to the Sun.     

Spacecraft Injection into a Heliocentric Terrestrial Orbit

We consider the problem of injection of a spacecraft into the heliocentric Earth's orbit ahead and/or behind the Earth by 60° and 120° in heliographic longitude. The range of solar and astrophysical problems for which these orbits are necessary is reviewed. The variants of injection into heliocentric orbits work from a low around-Earth orbit with one turn-on of the engine in this orbit and one turn-on at the end of the injection trajectory. In this case, it turns out to be more profitable to put spacecraft into orbit for three or even four revolutions of the Earth about the Sun. The velocities necessary for the start from a low around-Earth orbit, the velocities at the final point of injection, and the fuel mass (relative to the spacecraft mass) necessary for injection are estimated. The problems for which injection to similar orbits is executed, using the low-thrust engine and with a combined regime of injection, are also considered.     

Analytical solutions for the relative motion of spacecraft subject to Lorentz-force perturbations

A spacecraft capable of producing higher-than-natural electrostatic charges may achieve propellantless orbital maneuvering via the Lorentz-force interaction with a planetary magnetic field. Development of maneuver strategies for these propellantless vehicles is complicated by the fact that the perturbative Lorentz force acts along only a single line of action at any instant. Relative-motion dynamical models are developed that lead to approximate analytical solutions for the motion of charged spacecraft subject to the Lorentz force. These solutions indicate that the principal effects of the Lorentz force on a spacecraft in a circular orbit are to change the intrack position and to change the orbit plane. A rendezvous example is presented in which a spacecraft with a specific charge of ?3.81 $\times$ 10^?4 C/kg reaches a target vehicle initially 10 km away (on the same equatorial low-Earth orbit) in 1 day. Fly-around maneuvers may be achieved in low-Earth orbit with specific charges on the order of 0.001 C/kg.     

对接动力学仿真试验台再现频率指标验证

为验证对接动力学仿真试验台再现对接动力学过程中频率0~6 Hz相对运动的能力,用弹性棒方案再现频率0~6 Hz的碰撞过程,由建模与设计计算选定弹性棒的悬臂长度、刚度和航天器质量等设计参数,仿真分析了碰撞过程的接触碰撞力、变形量与速度,以及再现频率等的影响。结果表明:设计参数合理,精度考核方案可行。     

Design feasibility via ascent optimality for next-generation spacecraft

This paper deals with the optimization of the ascent trajectories for single-stage-sub-orbit (SSSO), single-stage-to-orbit (SSTO), and two-stage-to-orbit (TSTO) rocket-powered spacecraft. The maximum payload weight problem is studied for different values of the engine specific impulse and spacecraft structural factor.The main conclusions are that: feasibility of SSSO spacecraft is guaranteed for all the parameter combinations considered; feasibility of SSTO spacecraft depends strongly on the parameter combination chosen; not only feasibility of TSTO spacecraft is guaranteed for all the parameter combinations considered, but the TSTO payload is several times the SSTO payload.Improvements in engine specific impulse and spacecraft structural factor are desirable and crucial for SSTO feasibility; indeed, aerodynamic improvements do not yield significant improvements in payload.For SSSO, SSTO, and TSTO spacecraft, simple engineering approximations are developed connecting the maximum payload weight to the engine specific impulse and spacecraft structural factor. With reference to the specific impulse/structural factor domain, these engineering approximations lead to the construction of zero-payload lines separating the feasibility region (positive payload) from the unfeasibility region (negative payload).     

Optical effects of the operation of the onboard engine of the <Emphasis Type="Italic">Progress M-17M</Emphasis> spacecraft at thermospheric heights

This paper presents the results of optical observations in the active space experiment “Radar-Progress” on April 17, 2013, after switching on the approach-correction engine of the Progress M-17M cargo spacecraft at thermospheric heights (412 km), are presented in this paper. During engine operation, a region of enhanced emission intensity has been recorded. It was presumably related to the scatter of twilight solar emission at the engine exhausts in the cargo spacecraft orbit and, probably to the occurrence of an additional emission in the atomic oxygen line [OI] 630 nm. The maximum observed dimensions of the emission region were ~350 and ~250 km along the orbit and across it, respectively. The velocity of the expansion of the emission region at the first moments after the initiation of engine operation was ~7 and ~3.5 km/s along the orbit and across it, respectively. The maximum intensity of the disturbed region is estimated to be a value equivalent to ~40–60 R within the spectral band of 2 nm. No optical manifestation, which would exceed the natural variations in brightness of the night airglow and which would be related to possible large-scale modification of the ionosphere, was detected in the natural emission lines [O] 557.7 and 630.0 nm in a zone remote from the place of injection of engine exhausts.     

超低轨航天器气动力分析与减阻设计

轨道降低,航天器受到的气动力增大,气动力对航天器影响显著。考虑自由分子流态 下的超低轨航天器,利用分割法把简单外形的航天器分割为几部分,分别计算各部分的气动 力,然后相加获得总的气动力效果;通过对平面的气动力进行计算分析,提出了超低 轨航天器的减阻设计方法;结果表明：当轨道高度降低到250 km左右时,航天器受到的气动 阻力比500 km高出约2个数量级;一般情况下,超低轨航天器应采用细长体构型,减小迎风 面积;侧面积引起的航天器阻力已经不可忽略,应采用侧面光滑技术,减少侧面阻力;当超 低轨航天器长细比超过一定限度后,随着长细比增大,大气阻力升高。
     

航天器在轨释放安全性分析与仿真研究

针对航天器舱内在轨释放任务,文章建立了大型航天器与伴随卫星的动力学模型和运动学模型,并分析了影响释放安全性的主要因素;从保证航天器在轨安全的角度,得到描述释放安全性的最小相对距离计算方法,并利用ADAMS软件对航天器在轨释放过程进行了仿真研究。     

力限技术在航天器振动试验中的应用

为了在正弦振动试验中模拟真实的飞行环境,防止因“过试验”而导致航天器结构发生不必要的破坏,需要对试验条件进行下凹控制。在以往的加速度下凹控制方法的基础上,引入力限控制方法可以提高航天器主频处下凹控制的精度和有效性。文章分析了传统加速度下凹控制方法的局限性,并以某结构星力限控制试验为基础,阐述了结构星力限试验条件的制定方法,介绍了力限双控试验平台设计,并分析了试验结果。经分析表明,力限控制具有较高的控制精度,在加速度控制的基础上引入力限控制的“双控”试验方法,能够有效解决航天器振动试验中的“过试验”和“欠试验”问题。     

Trajectories for missions of low-thrust spacecraft aimed at delivery of soil samples from main belt asteroids and Phobos

Trajectories of spacecraft with electro-jet low-thrust engines are studied for missions planning to deliver samples of matter from small bodies of the Solar System: asteroids Vesta and Fortuna, and Martian moon Phobos. Flight trajectories are analyzed for the mission to Phobos, the limits of optimization of payload spacecraft mass delivered to it are determined, and an estimate is given to losses in the payload mass when a low-thrust engine with constant outflow velocity is used. The model of an engine with ideally regulated low thrust is demonstrated to be convenient for calculations and analysis of flight trajectories of a low-thrust spacecraft.     

振动试验中力限控制技术

文章阐述了在地面模拟航天器经历的力学环境时,应充分合理地考核航天器的结构；回顾了国内外解决振动试验中过试验问题的方法；全面介绍了振动试验中的力限控制技术。该技术是利用加速度控制的同时,通过限制振动试验中受试产品与工装连接面作用力的大小,来避免振动试验中的过试验问题。     

Prediction of Absorbed Doses for Spacecraft Irradiated by Particle Fluxes of Solar Cosmic Rays

A new method of evaluating the absorbed dose onboard spacecraft under the action of particle fluxes of solar cosmic rays (SCR) is described. The method is based on the concepts and formulas of a probabilistic model of SCR proton fluxes beyond the Earth's magnetosphere (GOST R 25645.165-2001) and takes into account variations of particle fluxes at orbits of spacecraft and behind their protective shields.     

空间热物理技术发展的探讨

文章以航天器和火箭发动机中的工程热物理问题为背景,探讨了航天器热控技术的现状和发展趋势,介绍了火箭发动机热设计、推进剂燃烧等问题的发展方向.     

Analysis of evolution of multi-orbit perturbed trajectories of a spacecraft in a central field of attraction

Properties of differential equations of multi-orbit trajectory motion of a spacecraft are investigated analytically. The spacecraft moves under the action of small perturbations (in particular, low thrust) in the plane of a central Newtonian field of attraction. The conditions are specified for existence of a partial singular aperiodic solution, in the neighborhood of which the behavior of osculating elements changes sharply. In this case, phase variables (the angular position of the pericenter and the true anomaly) are found to undergo the sharpest changes. The exact superposition of solutions is suggested for the equations of motion transformed to the form of a quasi-linear, weakly non-stationary system: a partial singular aperiodic solution and fast solutions oscillating around it. Asymptotic representations are obtained for both components of the superposition. They are fairly exact in the region of smallness of perturbing terms at a long variation of the argument.     

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.     

The Impact of Variable Accelerations on Crystal Growth onboard Spacecraft by the Floating Zone Method

The numerical investigation of the impact of time-dependent accelerations (vibrations) on the flow and heat and mass transfer in the melt is carried out for the case of modeling the crystal growth by the floating zone method under conditions of microgravity that exist onboard spacecraft. The effects of the Archimedean buoyancy force and of vibrations of the free surface of fluid are considered separately. When solving the problem of the effect of the vibrations of the free surface of fluid, the previously obtained data were used. It is shown that vibrations of the free surface have a much stronger effect on the processes under consideration than the buoyancy. Some problems that are related to the newly discovered effects are discussed. The use of vibroprotected systems and a rotating magnetic field can help solve these problems. We plan to continue our investigations in future spacecraft experiments, in particular, at the International Space Station, which is under construction at the moment.