载人登月轨道研究综述

21世纪初叶,美国星座计划、欧洲曙光计划以及中国嫦娥工程计划的相继提出,彻底拉开了人类重返外太空的序幕,载人登月也被赋予了全新的涵义和使命.探测内容的多样化,轨道设计的复杂精细化,宇航员安全保障的突出化,进一步加大了载人登月工程任务设计的复杂度与难度.文中分别从载人登月轨道整体、地月转移轨道、空间交会对接以及全月面到达这4个方向介绍载人登月当前研究进展与现状,并列举了各研究领域的相关热点问题.基于对上述研究现状的分析,尝试展望了该研究领域未来的发展趋势.     

双月旁转向轨道的修正方法研究

双月旁转向轨道是深空探测中有应用潜力的一类非线性轨道. 在CR3BP模型下得到了双月旁转向轨道, 通过分析轨道的误差传播特性, 制定了双月旁转向轨道的修正原则, 提出了``关键节点'加``显式制导'的修正思路; 给出了考虑太阳引力作用的R4BP模型, 针对初始误差、导航误差、修正执行误差和太阳引力摄动偏差进行了轨道修正仿真, 得到了基于Monte-Carlo方法的轨道修正统计结果. 在此基础上, 采用了``初值速度补偿'与``显式制导'相结合的思路修正初始误差、导航误差、修正执行误差和太阳引力作用, 仿真表明: 这一方法能够较大程度地降低轨道修正所需的冲量. 结果和结论能够为双月旁转向轨道的工程应用提供参考.      

基于抱爪式对接机构捕获缓冲系统动力学仿真研究

现有在轨服务的对接机构由于其尺寸大、结构复杂、对接目标单一等局限性因素,无法很好地为后续我国探月工程任务提供有力支撑,且受限于运载能力,对接机构的轻量化也是必不可少的一项环节.为研究可服务于未来月球空间站以及载人登月等高轨道任务的对接机构,设计了一种新型抱爪式对接机构,其采用异体同构周边式构型,可以实现主/被动飞行器之间的互换.利用 V 型槽与爪钩等结构部件实现飞行器对接过程中的捕获以 及能量消耗功能,从而实现两飞行器之间的稳固联接.该对接机构具备尺寸小、重量轻、结构简单、功能易实现等优势. 对其捕获缓冲系统进行了动力学分析,计算了缓冲元器件的参数对其捕获性能的影响,在 ADAMS 完成了数字虚拟样机的建立,结合实际两种典型的对接初始条件工况进行了仿真研究.研究结果表明,两种工况下的对接过程能量消耗满足设计要求,能够以较小的 V 型槽的碰撞力完成捕获,结果证明了捕获缓冲系统的可行性以及该构型对接机构具备较好实现任务的能力.      

DYNAMIC SIMULATION ANALYSIS OF CAPTURE AND BUFFER SYSTEM BASED ON CLAW-TYPE DOCKING MECHANISM

现有在轨服务的对接机构由于其尺寸大、结构复杂、对接目标单一等局限性因素,无法很好地为后续我国探月工程任务提供有力支撑,且受限于运载能力,对接机构的轻量化也是必不可少的一项环节.为研究可服务于未来月球空间站以及载人登月等高轨道任务的对接机构,设计了一种新型抱爪式对接机构,其采用异体同构周边式构型,可以实现主/被动飞行器之间的互换.利用 V 型槽与爪钩等结构部件实现飞行器对接过程中的捕获以 及能量消耗功能,从而实现两飞行器之间的稳固联接.该对接机构具备尺寸小、重量轻、结构简单、功能易实现等优势. 对其捕获缓冲系统进行了动力学分析,计算了缓冲元器件的参数对其捕获性能的影响,在 ADAMS 完成了数字虚拟样机的建立,结合实际两种典型的对接初始条件工况进行了仿真研究.研究结果表明,两种工况下的对接过程能量消耗满足设计要求,能够以较小的 V 型槽的碰撞力完成捕获,结果证明了捕获缓冲系统的可行性以及该构型对接机构具备较好实现任务的能力.     

地月低能转移的发生条件及轨迹构造

应用平动点理论研究了地月低能转移的发生条件和轨迹构造. 由于空间双圆模型存在周期性扰动,传统的平动点概念不再适用;根据Hill瞬时边界的连接情况定义了等效平动点LL_{1}和LL_{2}. 通过构造合适的Poincaré截面以获得所有可能的月球捕获轨道的近月距和偏心率分布,从而获得了完全不同于Hill和圆型限制性三体问题的月面捕获能量. 在空间双圆模型下,平动点和Halo轨道不变流形的渐近结构遭到破坏:到达或离开平动点的时间,由无穷转变为有限值;运动方向由双向变为单向. 经由LL_{1}点穿越获得了最小能量的低能转移,借助LL_{1}-Halo轨道穿越得到了(M,N)-圈穿越轨道,经由LL_{2}点穿越获得了最小能量的弱稳定边界转移,借助LL_{2}-Halo轨道穿越得到了弱稳定边界逃逸和捕获窗口. 最后,以地月转移和大幅值逆行轨道的切入为例,给出低能转移的小推力、脉冲和弱稳定边界等转移的实现方式.      

地球-火星引力辅助下的地-火无动力循环轨道

针对火星探测任务,提出一种地-火无动力循环轨道设计方法。该方法有望实现火星探测的自由返回,并能够多次无动力地火往返。本文首先求解地火之间转移的Lambert问题,考虑地球和火星引力辅助效应,并分别给出以地球和火星为中心的半圈转移轨道。其次,以该轨道作为辅助轨道,设计搜索能够实现无动力循环的发射窗口和转移轨道。最后,本文给出效率较高的几种无动力循环轨道,希望对未来火星探测的任务设计有一定指导意义。     

THE EVOLUTION CHARACTERIZATION OF LIBRATIN POINT ORBITS AND APPLICATION RESEARCH

三体问题中, 轨道的受力和运动规律非常复杂. 对于特定的任务, 如何选择轨道的初始解是一大难题.针对平面三体问题, 利 用近拱点庞加莱映射, 对平动点顺行轨道和逆行轨道的长期和短期演化性质进行分析.根据轨道的初始状态将其分为逃逸轨道和捕获轨道.对于逃逸轨道, 给出了同宿轨道和异宿轨道的设计方法, 并利用两级微分修正法消除了拼接点处的位置不连续问题.对于捕获轨道, 得到了几类典型的周期和准周期轨道.对逆行轨道的演化性质进行分析时发现, 逆行轨道通常为准周期轨道, 比顺行轨道更加稳定.利用近拱点庞加莱映射可以快速确定不同类型轨道对应的初始状态, 为特定任务需求下的轨道设计提供了一种快速而有效的选择方案.     

基于虚拟中心引力场方法的航天器转移轨道设计

空间机动技术是实现空间操作任务的基础,具有重要的研究价值. 研究了连续推力作用下航天器转移轨道设计问题,提出了一种基于虚拟中心引力场的轨道设计方法. 该方法有两大特点:(1) 能够将机动轨道设计问题转化为虚拟中心引力场参数的优化问题,简化了设计过程;(2) 对轨道形状或推力方向、大小不做任何假定,能够应用于一般情况下的机动轨道设计. 将该方法应用于航天器二维和三维的转移轨道设计,并和形状方法进行了对比分析. 仿真结果分析表明,采用该方法简化了轨道设计过程,为航天器快速轨道设计提供了新思路.      

气动力辅助变轨的数值模拟

采用了航天器在行星上层大气中进行高超声速飞行时的轨道动力学方程,针对航天器从地球静止轨道转移到一个共面圆形低地轨道的变轨过程,进行了气动力辅助变轨过程的模拟.在变轨过程中,航天器从地球静止轨道开始,经过8次大气路径,耗时43.7小时,到达圆形低地轨道,与霍曼转移进行对比,其所消耗的推进剂质量仅为霍曼转移的41%.研究结果表明:气动力辅助变轨技术能够在降低推进剂消耗的情况下实现航天器的轨道转移.     

基于偏转距离近似模型的动能撞击小行星防御任务脉冲轨道优化研究

小行星撞击对地球上的生命存在重大潜在威胁,动能撞击是目前最易实现且成熟度最高的防御方案.动能撞击任务的一种轨道优化指标为最大化偏转距离(即小行星被偏转前后近地距的改变量),若用数值积分的方法精确计算偏转距离, 会导致优化效率较低.在动能撞击任务的设计初期, 可以对动力学模型及偏转距离的计算方法进行简化,以提升优化效率. 本文首先将高精度模型简化为二体模型,分析了两种经典偏转距离解析模型的适用条件,同时提出一种基于近地点时刻预估的偏转距离近似模型; 考虑运载约束,将化学推进变轨简化为脉冲推力变轨,建立了直接转移(两脉冲及三脉冲)和行星借力飞行转移(单次及两次借力)的动能撞击轨道优化模型,利用遗传算法求解了优化问题. 以偏转小行星Apophis为例, 相比于解析模型,验证了本文提出的近似模型可以同时提升最优性、降低求解复杂性. 优化结果表明,三脉冲直接转移方案与两脉冲直接转移方案的最优偏转效果基本一致,借力飞行转移方案相比于直接转移方案对偏转距离的提升效果并不明显.在动能撞击任务的前期设计中, 可以基于二体模型进行防御效果的快速评估,虽然对计算偏转距离存在一定误差, 但对防御窗口的优化结果影响不大. 进一步,数值求解偏转距离时, 可通过引入主要引力摄动项(金星、地球、木星)修正二体模型,使其与高精度模型之间的求解误差在1%以下.      

Design of low-energy transfer from lunar orbit to asteroid in the Sun-Earth-Moon system

Asteroid exploration trajectories which start from a lunar orbit are investigated in this work.It is assumed that the probe departs from lunar orbit and returns to the vicinity of Earth,then escapes from the Earth by performing a perigee maneuver.A low-energy transfer in Sun-EarthMoon system is adopted.First,the feasible region of lowenergy transfer from lunar orbit to perigee within 5 000 km height above the Earth surface in Sun-Earth-Moon system is calculated and analyzed.Three transfer types are found,i.e.,large maneuver and fast transfers,small maneuver and fast transfers,and disordered and slow transfers.Most of feasibility trajectories belong to the first two types.Then,the lowenergy trajectory leg from lunar orbit to perigee and a heliocentric trajectory leg from perigee to asteroid are patched by a perigee maneuver.The optimal full-transfer trajectory is obtained by exploiting the differential evolution algorithm.Finally,taking 4179 Toutatis asteroid as the target,some low-energy transfer trajectories are obtained and analyzed.     

First-round design of the flight scenario for Chang’e-2’s extended mission:take off from lunar orbit

Chang'e-2, Chinese second lunar probe, was inserted into a 100 km altitude low lunar orbit on October 9th, 2010, its purpose is to continuously photograph the lunar surface and possibly chosen landing sites for future lunar missions. The probe will still carry considerable amount of propellant after completing all prescribed tasks in about six months. After the successful launch of Chang'e-2, we began designing the probe's subsequent flight scenario, considering a total impulse of 1 100 m/s for takeoff from low lunar orbit and a maximum 3×10 6 km distance for Earth-probe telecommunication. Our first-round effort proposed a preliminary flight scenario that involves consecutive arrivals at the halo orbits around the Earth-Moon L1/L2 and Sun-Earth L1/L2 points, near-Earth asteroid flyby, Earth return, and lunar im- pact. The designed solution of Chang'e-2's subsequent flight scenario is a multi-segment flight trajectory that serves as a reference for making the final decision on Chang'e-2's extended mission, which is a flight to the Sun-Earth L2 point, and a possible scheme of lunar impact via Earth flyby after remaining at the Sun-Earth L2 point was also presented. The proposed flight trajectory, which possesses acceptable solution accuracy for mission analysis, is a novel design that effectively exploits the invariant manifolds in the circular restricted three-body problem and the patched-manifold-conic method.     

Lunar landing trajectory design based on invariant manifold

The low-energy lunar landing trajectory design using the invariant manifolds of restricted three-body problem is studied.Considering angle between the ecliptic plane and lunar orbit plate the four-body problem of sun-earth-moon-spacecraft is divided into two three-body problems,the sun-earth-spacecraft in the ecliptic plane and the earth- moon-spacecraft in the lunar orbit plane.Using the orbit maneuver at the place where the two planes and the invariant manifolds intersect,a general method to design low energy lunar landing trajectory is given.It is found that this method can save the energy about 20% compared to the traditional Hohmann transfer trajectory,The mechanism that the method can save energy is investigated in the point of view of energy and the expression of the amount of energy saved is given.In addition,some rules of selecting parameters with respect to orbit design are provided.The method of energy analysis in the paper can be extended to energy analysis in deep space orbit design.     

Systematic design and development of a flexible wheel for low mass lunar rover

Low mass compact rovers provide cost effective means to explore extra-terrestrial terrains. Use of flexible wheels in such applications where the wheel size is restricted, improves traction at reduced slip and sinkage. Design of a flexible wheel for a given mission is a challenging task requiring consideration of stiffness of rim and spokes, stress induced in the wheel, chassis movement during wheel rotation and the operating mode of the wheel. Also, accurate mathematical models are required to save design and development time and reduce the number of prototypes for selection. It is observed that most of the research papers deal with performance testing of flexible wheels and information on analytical formulation is scarce. Therefore, in the present work, a methodology has been formulated to systematically design a flexible wheel for a low mass lunar rover. The prototype performance is tested and compared with analytical estimates and reasons for difference are investigated. Paper contains details of design criteria, mathematical modelling, realisation of wheel prototype, test fixture and analysis test comparison. Authors believe that this work provides a useful aid to the designer to systematically design flexible wheels for low mass lunar rovers.     

Spacecraft aerodynamics and trajectory simulation during aerobraking

This paper uses a direct simulation Monte Carlo （DSMC） approach to simulate rarefied aerodynamic characteristics during the aerobraking process of the NASA Mars Global Surveyor （MGS） spacecraft. The research focuses on the flowfield and aerodynamic characteristics distribution under various free stream densities. The vari- ation regularity of aerodynamic coefficients is analyzed. The paper also develops an aerodynamics-aeroheating-trajectory integrative simulation model to preliminarily calculate the aerobraking orbit transfer by combining the DSMC technique and the classical kinematics theory. The results show that the effect of the planetary atmospheric density, the spacecraft yaw, and the pitch attitudes on the spacecraft aerodynamics is significant. The numerical results are in good agreement with the existing results reported in the literature. The aerodynamics-aeroheating-trajectory integrative simulation model can simulate the orbit transfer in the complete aerobraking mission. The current results of the spacecraft trajectory show that the aerobraking maneuvers have good performance of attitude control.     

Design and characterization of GRC-1: A soil for lunar terramechanics testing in Earth-ambient conditions

Earth experiments must be carried out on terrain that deforms similarly to the lunar terrain to assess the tractive performances of lunar vehicles. Most notably, terrain compaction and shear response underneath the lunar vehicle wheels must represent that of the Moon. This paper discusses the development of a new lunar soil simulant, Glenn Research Center lunar soil simulant #1 (GRC-1), which meets this need. A semi-empirical design approach was followed in which the soil was created by mixing readily available manufactured sands to a particle size distribution similar to the coarse fraction of lunar soil. By varying terrain density, a broad range of in situ cone penetration measurements collected by the Apollo mission astronauts can be replicated. An extensive set of characterization data is provided in this article to facilitate the use of this material. For reference, the index and geotechnical properties of GRC-1 are compared to the lunar soil and existing lunar soil simulants.     

Pendulum motions of extended lunar space elevator

In the usual everyday life, it is well known that the inverted pendulum is unstable and is ready to fall to “all four sides,” to the left and to the right, forward and backward. The theoretical studies and the lunar experience of moon robots and astronauts also confirms this property. The question arises: Is this property preserved if the pendulum is “very, very long”? It turns out that the answer is negative; namely, if the pendulum length significantly exceeds the Moon radius, then the radial equilibria at which the pendulum is located along the straight line connecting the Earth and Moon centers are Lyapunov stable and the pendulum does not fall in any direction at all. Moreover, if the pendulum goes beyond the collinear libration points, then it can be extended and manufactured from cables. This property was noted by F. A. Tsander and underlies the so-called lunar space elevator (e.g., see [1]). In the plane of the Earth and Moon orbits, there are some other equilibria which turn out to be unstable. The question is, Are there equilibria at which the pendulum is located outside the orbital plane? In this paper, we show that the answer is positive, but such equilibria are unstable in the secular sense. We also study necessary conditions for the stability of lunar pendulum oscillations in the plane of the lunar orbit. It was numerically discovered that stable and unstable equilibria alternate depending on the oscillation amplitude and the angular velocity of rotation. The study of the lunar elevator dynamics originates in [2]. The concept of lunar elevator was developed in detail in [3, 4]. Several classes of equilibria with the finiteness of the Moon size taken into account were studied in [5]. The possibility of location of an orbital station fixed to the Moon surface by a pair of tethers was investigated in [6]. The problem of orientation of the terminal station of the lunar space elevator was studied in [7]. The influence of the tether length variations on the motion of the lunar tether system was considered in [8]. The alternation of stable and unstable flat oscillations is well known in the problem of satellite oscillations in a circular orbit [9, 10].     

Formation around planetary displaced orbit

The paper investigates the relative motion around the planetary displaced orbit. Several kinds of displaced orbits for geocentric and martian cases were discussed. First, the relative motion was linearized around the displaced orbits. Then, two semi-natural control laws were investigated for each kind of orbit and the stable regions were obtained for each case. One of the two control laws is the passive control law that is very attractive for engineering practice. However, the two control laws are not very suitable for the Martian mission. Another special semi-natural control law is designed based on the requirement of the Martian mission. The results show that large stable regions exist for the control law.