从近地轨道入轨的载人登月发射窗口分析与设计

为了提高安全性和增加运载能力,现代载人登月将采用人货分运的方式.针对从近地轨道入轨的问题,建立从近地轨道出发,以自由返回轨道为设计目标,满足轨道约束、光照约束和测控约束的发射窗口的计算模型.提出多约束条件下发射窗口选择算法.最后,根据自由返回轨道的特点给出发射窗口仿真算例.研究结果表明,对于给定着月点和近地轨道,载人登月的发射窗口并不是每个月都会出现一次,但在1年中,仍有多次机会可供选择.     

关于月球低轨卫星运动的两个问题

对月球低轨卫星的轨道寿命特征和冻结轨道晶状态作了详尽的理论分析,给出它们与轨道倾角之间的关系以及它们相互之间的某种联系,并考虑低轨卫星的主要摄动源,在完整力模型下作了相应的模拟计算,不仅证实了理论分析的正确性,而且为环月运行探测器的轨道设计提供了极有参考价值的数值结果．     

月球卫星轨道力学综述

月球探测器的运动通常可分为3个阶段，这3个阶段分别对应3种不同类型的轨道：近地停泊轨道、向月飞行的过渡轨道与环月飞行的月球卫星轨道。近地停泊轨道实为一种地球卫星轨道；过渡轨道则涉及不同的过渡方式(大推力或小推力等)；环月飞行的月球卫星轨道则与地球卫星轨道有很多不同之处，它决不是地球卫星轨道的简单克隆。针对这一点，全面阐述月球卫星的轨道力学问题，特别是环月飞行中的一些热点问题，如轨道摄动解的构造、近月点高度的下降及其涉及的卫星轨道寿命、各种特殊卫星(如太阳同步卫星和冻结轨道卫星等)的轨道特征、月球卫星定轨等。     

谈谈印度首颗探月卫星

一、印度月球初航1号如何奔月 2008年10月22日,印度用极轨卫星运载火箭－XL（PSLV—XL）发射了其首颗探月卫星——月球初航1号（Chandrayaan-1,又叫月船1号）。探测器首先进入了近地点255千米、远地点22860千米的大椭圆轨道。在绕地运行两个星期的时间里,探测器上的液体发动机适时点火工作5次,     

2011年世界空间探测与科学观测活动回顾

2011年,美国、中国和俄罗斯等国都开展了多项空间探测与科学观测活动,发射了月球探测器、木星探测器等,尤其是由于2011年有火星探测器发射窗口,所以进行了两次重要的火星探测器发射,但有喜有悲：携带了世界最先进火星车“好奇号”的美国“火星科学实验室”顺利升空,并奔向火星;而携带了中国首个火星探测器“萤火-1”的俄罗斯的“火卫一-土壤”探测器虽然发射成功,但由于探测器本身故障而无法进入地火转移轨道,导致该任务失败。     

从近地轨道入轨的载入登月发射窗口分析与设计

为了提高安全性和增加运载能力,现代载人登月将采用人货分运的方式。针对从近地轨道入轨的问题,建立从近地轨道出发,以自由返回轨道为设计目标,满足轨道约束、光照约束和测控约束的发射窗口的计算模型。提出多约束条件下发射窗口选择算法。最后,根据自由返回轨道的特点给出发射窗口仿真算例。研究结果表明,对于给定着月点和近地轨道,载人登月的发射窗口并不是每个月都会出现一次,但在1年中,仍有多次机会可供选择。     

月球探测中转移轨道误差分析和中途修正计算

针对返回型月球探测器,基于月球探测中转移轨道的动力学特征和误差传递矩阵的性质,分别对地月转移轨道和月地转移轨道的误差传递特点进行了研究,并根据误差传递矩阵给出估算第1次中途轨道修正速度增量的线性公式.通过具体算例,给出在实际力学模型下月球探测中转移轨道误差传递性质,讨论了目标点和目标轨道两种不同的轨道修正方法的特点和适用情形,并结合再入约束条件对月地转移轨道第2次中途修正进行了分析和计算.     

满足约束条件的月球卫星飞行轨道的初步设计

讨论满足约束条件的月球卫星飞行轨道的设计问题，将约束条件分类为只与太阳，月球，地球，飞行器和观测站之间的相对位置有关的运行学约束条件以及涉及到飞行器轨道运行的动力学约束条件，在考虑月球卫星轨道的受力情况后，给出一种准确快速地计算和设计满足约束条件的标准飞行轨道的方法，并应用于不同约束条件下月球卫星的轨道预设计，初步讨论了轨道设计的误差分析，轨道跟踪及实时精密定轨等正在进行的其它相关工作。     

月球探测器过渡轨道的短弧定轨方法

论述的短弧定轨,是指在无先验信息情况下又避开多变元迭代的初轨计算方法,它需要相应的动力学问题有一能反映短弧内达到一定精度的近似分析解．探测器进入月球引力作用范围后接近月球时可以处理成相对月球的受摄二体问题,而在地球附近,则可处理成相对地球的受摄二体问题,但在整个过渡段的力模型只能处理成一个受摄的限制性三体问题．而限制性三体问题无分析解,即使在月球引力作用范围外,对于大推力脉冲式的过渡方式,相对地球的变化椭圆轨道的偏心率很大(超过Laplace极限),在考虑月球引力摄动时亦无法构造摄动分析解．就此问题,考虑在地球非球形引力(只包含J2项)和月球引力共同作用下,构造了探测器飞抵月球过渡轨道段的时间幂级数解,在此基础上给出一种受摄二体问题意义下的初轨计算方法,经数值验证,定轨方法有效,可供地面测控系统参考．     

2009年科学卫星和空间探测器大盘点

2009年,美国继续领跑,在科学卫星和空间探测器领域都有突出的表现,尤其是成功进行了又一次哈勃望远镜修复、“开普勒”、“月球勘测轨道器”和“月球坑观测与感知卫星”等重要航天发射,但也经历了一次发射惨败。     

Luni-solar perturbations of an Earth satellite

Luni-solar perturbations of the orbit of an artificial Earth satellite are given by modifying the analytical theory of an artificial lunar satellite derived by the author in recent papers. Expressions for the first-order changes, both secular and periodic, in the elements of the geocentric Keplerian orbit of the earth satellite are given, the moon's geocentric orbit, including solar perturbations in it, being found by using Brown's lunar theory.The effects of Sun and Moon on the satellite orbit are described to a high order of accuracy so that the theory may be used for distant earth satellites.     

Simulation of precise orbit determination of lunar orbiters

Based on the ongoing Chinese lunar exploration mission, i.e. the “Chang'e 1” project, precise orbit determination of lunar orbiters is analyzed for the actual geographical distribution and observational accuracy of the Chinese united S-band (USB) observation and control network as well as the very long baseline interferometry (VLBI) tracking network. The observed data are first simulated, then solutions are found after including the effects of various error sources and finally compared. We use the space data analysis software package, GEODYN, developed at Goddard Space Flight Center, NASA, USA. The primary error source of the flight orbiting the moon is the lunar gravity field. Therefore, the (formal) error of JGL165P1, i.e. the model of the lunar gravity field with the highest accuracy at present, is first discussed. After simulating the data of ranging and velocity measurement as well as the VLBI data of the time delay and time delay rate, precise orbit determination is carried out when the error of the lunar gravity field is added in. When the orbit is determined, the method of reduced dynamics is adopted with the selection of appropriate empirical acceleration parameters to absorb the effect of errors in the lunar gravity field on the orbit determination. The results show that for lunar missions like the “Chang'e 1” project, that do not take the lunar gravity field as their main scientific objective, the method of reduced dynamics is a simple and effective means of improving the accuracy of the orbit determination of the lunar orbiters.     

Error Analysis and Trajectory Correction Maneuvers of Lunar Transfer Orbit

For a returnable lunar probe, this paper studies the characteristics of both the Earth-Moon transfer orbit and the return orbit. On the basis of the error propagation matrix, the linear equation to estimate the ?rst midcourse trajectory correction maneuver (TCM) is ?gured out. Numerical simulations are performed, and the features of error propagation in lunar transfer orbit are given. The advantages, disadvantages, and applications of two TCM strategies are discussed, and the computation of the second TCM of the return orbit is also simulated under the conditions at the reentry time.     

Effects of the physical libration of the moon on lunar orbiters

Lunar physical libration, which is true oscillation of lunar equator in the space, alters the lunar gravitational field in the space coordinate system and affects the orbiting motion of lunar orbiters (hereafter called as lunar satellites) correspondingly. The effect is very similar to that of the precession and nutation on the earth satellites, and a similar treatment can be used. The variations in the gravitational force and in the orbit perturbation solution are clearly given in this paper together with numerical illustrations.     

The Effect of GPS Ephemeris on the Accuracy of Precise Orbit Determination of LEO Satellite-borne GPS

The satellite-borne GPS receivers dedicated to precise orbit determination are now being carried by more and more low earth orbit (LEO) satellites and the satellite-borne GPS has become one of the main means for the precise orbit determination of low earth orbit satellites. The accuracy of satellite-borne GPS precise orbit determination depends on the accuracies of the GPS ephemeris and the clock error. Based on the orbit determination function of SHORDEIII zero-difference dynamics and using the observational data obtained by the GRACE satellites for the week from 2005 August 1 to 7 as an example, three versions of GPS ephemerides (igs, igr and igu) are used to carry out orbit determination under the same conditions and to estimate the effect of the GPS ephemeris accuracy on the accuracy of orbit determination of low earth orbit satellites. Our calculated results show that the two ephemerides, igs and igr, are equivalent to each other in orbit determination accuracy (about 9.5 cm), while igu is slightly less accurate, at about 10.5 cm. The effect produced by the data of the high frequency GPS satellite clock error on the accuracy of orbit determination is 1–6 cm.     

Analytical design methods for transfer trajectories between the Earth and the Lunar Orbital Station

Lunar Orbital Station (LOS) is proposed as support of manned lunar exploration missions. A fast-converging iteration method for determining the initial conditions of two-impulse transfer trajectories between the Earth and the LOS is proposed based on the patched conic approach. In the Earth phase, near Earth state is connected with the state at the lunar sphere of influence (LSOI) based on the relationship between the initial and terminal orbital state. Then, an analytical algorithm is proposed to find the state vector at LSOI, such to satisfy the LOS orbital constraint. An iterative process is finally adopted to generate favorable initial solutions that satisfy the constraint near the Earth and at the perilune. The algorithm convergence is investigated, and two types of transfer trajectories are found for both Earth-LOS and LOS-Earth transfer. Based on the algorithm, orbital transfer windows, velocity impulse and time of flight are analyzed in the typical years 2025 and 2034. At last, the initial solution is corrected with a high fidelity model based on the active-set method, which shows the precision of this algorithm. The novel procedure for the transfer trajectories design and the analytic result can be used as a basis for rapid mission evaluation and design for future manned lunar missions based on the LOS.     

Dynamical History of the Earth–Moon System

Differential equations describing the tidal evolution of the earth's rotation and of the lunar orbital motion are presented in a simple close form. The equations differ in form for orbits fixed to the terrestrial equator and for orbits with the nodes precessing along the ecliptic due to solar perturbations. Analytical considerations show that if the contemporary lunar orbit were equatorial the evolution would develop from an unstable geosynchronous orbit of the period about 4.42 h (in the past) to a stable geosynchronous orbit of the period about 44.8 days (in the future). It is also demonstrated that at the contemporary epoch the orbital plane of the fictitious equatorial moon would be unstable in the Liapunov's sense, being asymptotically stable at early stages of the evolution. Evolution of the currently near-ecliptical lunar orbit and of the terrestrial rotation is traced backward in time by numerical integration of the evolutional equations. It is confirmed that about 1.8 billion years ago a critical phase of the evolution took place when the equatorial inclination of the moon reached small values and the moon was in a near vicinity of the earth. Before the critical epoch t _cr two types of the evolution are possible, which at present cannot be unambiguously distinguished with the help of the purely dynamical considerations. In the scenario that seems to be the most realistic from the physical point of view, the evolution also has started from a geosynchronous equatorial lunar orbit of the period 4.19 h. At t < t _cr the lunar orbit has been fixed to the precessing terrestrial equator by strong perturbations from the earth's flattening and by tidal effects; at the critical epoch the solar perturbations begin to dominate and transfer the moon to its contemporary near-ecliptical orbit which evolves now to the stable geosynchronous state. Probably this scenario is in favour of the Darwin's hypothesis about originating the moon by its separation from the earth. Too much short time scale of the evolution in this model might be enlarged if the dissipative Q factor had somewhat larger values in the past than in the present epoch. Values of the length of day and the length of month, estimated from paleontological data, are confronted with the results of the developed model.     

借助光压将探测器推向月球

若采用圆型限制性三体问题模型，从近地停泊轨道上发射一个月球探测器，其最小初始速度必须使相应的Jacobi常数C小于某一临界值C2。但这仅仅是探测器可能飞向月球的必要条件，而且这样飞向月球耗时过长。若采用Hohmann转移轨道，则需要获得较大的变轨冲量，能量消耗较大。如果需要仔细探测地月空间环境，而又不必很快地飞往月球，那么采用较大的太阳帆板，并使其法向有一特殊指向，可借助太阳光压加速引导探测器在不长的时间内飞向月球。利用相应的分析和计算，证实上述考虑是有效的，而且若使太阳帆板截面积大到一定程度（如果技术上能实现），则无需任何动力，也可借助光压将探测器推向月球，就像一条太空帆船（简称太空帆）。     

Reconstructing the landing trajectory of the CE-3 lunar probe by using images from the landing camera

An accurate determination of the landing trajectory of Chang'e-3(CE-3)is significant for verifying orbital control strategy, optimizing orbital planning, accurately determining the landing site of CE-3 and analyzing the geological background of the landing site. Due to complexities involved in the landing process, there are some differences between the planned trajectory and the actual trajectory of CE-3. The landing camera on CE-3 recorded a sequence of the landing process with a frequency of 10 frames per second. These images recorded by the landing camera and high-resolution images of the lunar surface are utilized to calculate the position of the probe, so as to reconstruct its precise trajectory. This paper proposes using the method of trajectory reconstruction by Single Image Space Resection to make a detailed study of the hovering stage at a height of 100 m above the lunar surface. Analysis of the data shows that the closer CE-3 came to the lunar surface, the higher the spatial resolution of images that were acquired became, and the more accurately the horizontal and vertical position of CE-3 could be determined. The horizontal and vertical accuracies were7.09 m and 4.27 m respectively during the hovering stage at a height of 100.02 m. The reconstructed trajectory can reflect the change in CE-3's position during the powered descent process. A slight movement in CE-3 during the hovering stage is also clearly demonstrated. These results will provide a basis for analysis of orbit control strategy,and it will be conducive to adjustment and optimization of orbit control strategy in follow-up missions.