Using the Subsurface Soil Sounding Radar for Investigating the Structure and Total Electron Content of the Martian Ionosphere

The possibilities of using the Martian soil subsurface sounding radar for investigating the structure of the plasma shell surrounding the planet have been considered. Based on the numerical modeling results and actual soil sounding data, it has been shown that the soil sounding mode of the radio-locating MARSIS radar can be used to assess the structure of the Martian ionosphere. As the emitted signals pass to the planet’s surface, it is possible to use the reflected signals to estimate the total electron content of the Martian ionosphere along the flight track of the spacecraft.     

返回舱弹道重建与黑障区弹道再现研究

对于飞船返回舱，黑障区弹道再现是再人飞行试验气动分析工作的重要环节。利用舱上测量数据和有限的雷达测量数据重构飞行弹道，是黑障区弹道再现的有效方法。本文建立了返回舱弹道重建的数学模型，包括运动学模型、观测模型、测量误差模型，从而将返回舱弹道重建问题转化为一个非线性动态系统的参数辨识问题。给出了基于极大似然判据和Newton-Raphson迭代的弹道重建算法。对某飞船返回舱的飞行试验数据进行了计算和分析，结果证实了弹道重建数学模型的正确性和算法的可行性。通过弹道重建，不仅再现了黑障区的弹道，而且提供了可靠的、完整的弹道数据。     

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.     

Radiative Electrification of Spacecraft Construction Elements: Physical Modeling of Charge Accumulation and Neutralization

The basic principles of physical modeling of spacecraft radiative electrification in the ionosphere and the magnetosphere are formulated. It is shown that recorded cyclograms and current pulses of discharges provide complete information on the electrophysical properties of materials. The efficiency of plasma neutralization of electro-radiative effects depends on the density of low temperature plasma and the structure of the irradiated surface.     

The Use of a Solar Electrojet Propulsion System for Jupiter Satellite Injection

The problem of optimization of the interplanetary trajectory of flight for a multistage spacecraft with high- and low-thrust engines into the Jupiter satellite orbit is considered. Low-thrust engines (stationary plasma engines) are used on a heliocentric flight segment. Their operation is maintained with electric power supply from solar batteries. The principal feasibility of the realization of such a project is shown, and estimations of the mass of a spacecraft placed into Jupiter's satellite orbit are presented.     

Guidance of a spacecraft with low lift-to-drag ratio to a landing point

The problem of terminal control over a deorbiting spacecraft at the stage of its flight after leaving plasma (altitude of ∼40 km) is considered, the aim being to guide it to a preset landing point. The algorithm is based on a modification of the well-known method of proportional navigation, when a fixed point is the target. It is suggested to use satellite navigation systems (of the GLONASS or GPS types) and/or radio beacons, which should allow one to determine the spacecraft trajectory parameters with high precision. Single-channel control is performed by changing the roll angle according to current parameters of the trajectory, which ensures adaptability of the method. Examples of three-dimensional trajectories of flight are presented for a manned spacecraft with low lift-to-drag ratio (∼0.5), currently under design in Russia. The results of statistical modeling taking into account initial deviations of the trajectory parameters and wind disturbances are presented. A method of statistical choice of a reference trajectory for the guidance stage is suggested. A theoretical possibility of using the algorithm of spacecraft guidance (in case of in-light accident with a carrier launcher) to preset regions in the vicinity of launching route is demonstrated. A qualitative analysis of proportional navigation with a fixed target is presented.     

基于气动辅助变轨的变气动外形飞行器概念研究--最优控制律问题

从一个天地往返飞行器的上升轨道和再入返回轨道的优化，以及适用不同飞行任务的变轨要求的气动外形问题，提出一项基于气动力辅助变轨的变气动外形飞行器的新概念研究。对于一个固定气动外形飞行器要同时满足上升轨道有效载荷最大和再入轨道热流峰值、过载峰值及机动性能约束下的成本最低往往是困难的。若同时满足不同飞行任务：飞往太空站的运输任务，空间拦截和交会机动巡航任务及星际探测任务，则更为困难，实际上是不可能的。文章研究基于气动力辅助变轨，在热流约束下，气动外形参数变化对最优控制的影响。其结论为：热流约束下的最优控制解，包括考虑推力协同变轨，除了在非约束弧的滚转角不直接受气动外形影响外，其余的控制律，升力系数和滚转角都是气动外形参数和攻角的函数。因而变气动外形可作为一项新技术，即通过气动外形参数变化和相应的变轨策略而获得性能和成本都最佳的用途很广的一种新型飞行器。     

General solution for the optimal trajectory of an AFE-type spacecraft

The aeroassisted flight experiment (AFE) refers to an experimental spacecraft to be launched and then recovered by the Space Shuttle. It simulates a transfer from a geosynchronous Earth orbit (GEO) to a low Earth orbit (LEO). In this paper, with reference to an AFE-type spacecraft, an actual GEO-to-LEO transfer is considered under the following assumptions: the GEO and LEO orbital planes are identical; both the initial and final orbits are circular; the initial phase angle is given, while the final phase angle is free. The aeroassisted orbital transfer trajectory involves three branches: a preatmospheric branch, GEO-to-entry; an atmospheric branch, entry-to-exit; a post-atmospheric branch, exit-to-LEO. The optimal trajectory is determined by minimizing the total characteristic velocity. The optimization is performed with respect to the velocity impulses at GEO, LEO, and the time history of the angle of bank during the atmospheric pass. It is assumed that the entry path inclination is free and that the angle of attack is constant, = 17.0 deg. The sequential gradient-restoration algorithm is used to compute the optimal trajectory and it is shown that the best atmospheric pass is to be performed with constant angle of bank. The resulting optimal trajectory constitutes an ideal nominal trajectory for the generation of guidance trajectories for two reasons: the fact that the low value of the characteristic velocity is accompanied by relatively low values of the peak heating rate and the peak dynamic pressure; and the simplicity of the control distribution, requiring constant angle of bank.     

Optimization of Flights between the Orbits of Artificial Satellites of Two Planets Using a Combination of High and Low Thrusts

The problem of optimization of the trajectory of an interplanetary flight of a multistaged spacecraft using jet engines with high and low thrust is considered. The issues concerning the problem of choosing the main design parameters of a multistage spacecraft are touched upon. A mathematical model of start-to-finish optimization of all segments of the interplanetary flight trajectory is proposed. Using this model the specific features of flights to the orbits of satellites of Jupiter and Mars are studied.     

Possibilities of Combining High- and Low-Thrust Engines in Flights to Mars

The use of combinations of chemical and electric jet engines in the spacecraft designs results in a multistage vehicle configuration and in related problems of the optimum distribution of masses between the stages, the problems of flight trajectory optimization, and the problems of choosing the design parameters of a spacecraft. The appropriate issues are considered using flights to Mars as an example. The conditions for the optimum matching of high and low thrust trajectory segments are presented. The model of the simultaneous optimization of the geocentric and heliocentric legs of the trajectory is proposed. One- and two-orbit optimum trajectories of flight are investigated and analyzed.     

Frequency fluctuations of coherent signals from spacecraft observed during dual-frequency radio sounding of the circumsolar plasma in 2004–2008

We have performed spectral processing of the data of experiments on radio sounding of circumsolar plasma by coherent S- and X-band signals from the spacecraft Ulysses, Mars Express, Rosetta, and Venus Express carried out from 1991 to 2009. The experiments were realized in the mode of coherent response, when a signal stabilized by the hydrogen standard is transmitted from the ground station to a spacecraft, received by the onboard systems, and retransmitted to the Earth with conserved coherence. Thus, the signal sounding the coronal plasma passes twice through the medium: on the propagation path ground station — spacecraft and on the same path in the opposite direction. The spectra of frequency fluctuations in both the bands are obtained and, using them, the radial dependences of fluctuation intensities are found, which can be approximated by a power law. It is shown that the ratio of intensities of frequency fluctuations in the S- and X-bands is comparable with the theoretical value and characterizes the degree of correlation of irregularities of the electron density along the propagation path ground station — spacecraft and back. Analysis of the correlation of frequency fluctuations on the two paths allows one to get a lower estimate of the outer scale of the circumsolar plasma turbulence. For heliocentric distances R = 10 solar radii (R _S ) the outer scale is larger than 0.25R _S .     

全太阳系深空探测轨迹全局优化(英文)

针对太阳系中全部的248997颗行星的探测问题,给出了一种关于探测飞行器的深空探测全局四维轨迹(t,x,y,z)优化方案,即飞行器从地球发射进入太阳系并采用小推力控制,优化方案的性能指标为飞行器与太阳系中全部行星中相遇和交会的星的数量最多并且燃料消耗最少。本方案给出了四维飞行轨迹进行全局优化的一套算法,该算法由搜索算法和四维轨迹优化算法组成。此搜索算法从太阳系的248997颗行星中寻找获得尽可能多的经过近地球3维走廊内的行星;而四维轨迹优化算法由改进的动态规划算法、基于最优控制理论的共轭梯度算法和静态参数优化算法组成,其中静态参数优化算法用于搜索最优发射时间窗口。基于该组合算法,通过长时间的大规模的飞行数字仿真,最终计算出探测器的四维最优飞行轨迹,在一年内路过了太阳系中全部行星中的12颗行星。     

遗传算法在载人飞船返回轨道一体化设计中的应用

载人飞船返回再入轨道设计的主要任务是确定出标准返回轨道和再入制导规律。遗传算法具有全局寻优的特点，适合于工程设计的要求。文章探讨采用遗传算法实现飞船返回轨道的一体化设计，并给出飞船返回轨道设计步骤。     

Optimization of passive orbit with the use of gravity maneuver

A high-precision method of calculating gravitational interactions is applied in order to determine optimal trajectories. A number of problems, necessary for determination of optimal parameters at a launch of a spacecraft and during its flyby near celestial bodies, are considered. The spacecraft trajectory was determined by numerical integration of the equations of passive motion of the spacecraft and of the equations of motion for planets, the Sun, and the Moon. The optimal trajectory of the spacecraft approaching the Sun is determined by fitting its initial conditions.     

Mechanisms of generation and dynamics of HF and LF emissions in the experiments with electron beams injected into the earth’s ionosphere

Mechanisms of generation and dynamics of HF and LF emissions related to an electron beam injection from a spacecraft into the ionosphere are studied. The choice of the source and region of emission is substantiated. The plasma current neutralizing the spacecraft charge and the region of spatial charge in the vicinity of the spacecraft are chosen. Mechanisms of the generation of HF and LF emissions are proposed and their characteristics are determined. For the considered ranges of emission, the explanation of the facts related to the difficulties in interpretation of the amplitude value observed experimentally and the behavior of its envelope are presented. It is shown that the results obtained in this work agree well with the data of the conducted experiments.     

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。     

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.     

基于雷达跟踪仿真的滑翔式再入弹道突防性能分析

从防御雷达对再入弹道跟踪效果的角度,对滑翔式再入飞行器的突防性能进行了初步分析。基于Unscented卡尔曼滤波和考虑气动模型的目标跟踪模型,加入雷达测量噪声,计算目标跟踪误差,比较弹道式与滑翔式再入弹道的跟踪效果。仿真结果表明：(1) 滑翔式再入弹道有横向机动且飞行高度较低,雷达的发现时间较晚,可观测时间比较短;(2) 由于雷达的跟踪动力学模型中很难对气动力准确建模,滑翔式弹道借助气动力产生机动,雷达对其跟踪的误差要大于弹道式再入弹道; (3) 在相近的雷达观测条件下,弹道机动性越强,雷达跟踪精度越差;(4) 尽量使防御方所假设的目标动力学模型失配是增加突防能力的有效措施。     

The Electric Potential of a Spacecraft Injecting an Electron Beam into the Earth's Ionosphere

The problem of the electric potential value for the spacecraft injecting an electron beam into the Earth's ionosphere is considered. The formulas defining the spacecraft's electric potential in the equilibrium state as a function of experimental parameters are derived. It is shown that the presented theoretical results are in a good agreement with both the results of numerical modeling and experimental data. A comparison of different models of spacecraft charge neutralization is performed. An explanation is given for the facts related to the difficulties of interpretation of the mechanisms of spacecraft charge neutralization within a period less than the time of ionization of neutral particles and plasma heating by the injected beam.     

Exploring the Kuiper Belt: An extended Pluto mission

A robotic flyby mission to the planet Pluto is being planned for launch early in the next decade. The spacecraft will continue on out of the solar system in an almost radial direction traveling at about four AU per year and begin transiting the Kuiper Belt shortly after Pluto encounter. Recent discoveries and observations of Kuiper Belt objects have generated increased interest in the characteristics of these bodies. This paper examines the opportunities and requirements for extending the Pluto mission to include the search for, and encounters with, objects in the Kuiper Disk at 40+ AU. The trajectory and ΔV requirements will be presented. An automated, on-board sky survey will be proposed to inventory the Kuiper objects in the vicinity of the flight path and to identify which objects are candidates for altering the trajectory for a close flyby. A possible Kuiper object encounter science scenario will be described.