Trajectory optimization of multiple asteroids exploration with asteroid 2010TK7 as main target

In this study, the Earth’s Trojan asteroid 2010 TK₇ is selected as the rendezvous target. The multiple flyby sequence of asteroid exploration was proposed by optimizing the probe’s orbit. Impulsive maneuvers and low-thrust propulsion were used respectively to design the trajectories of the multiple asteroids exploration mission. Under impulsive maneuvers, gravity assist technique was adopted to reduce fuel consumption. First a reference orbit with only 2010 TK₇ as the rendezvous target was designed. Then five asteroids near the reference orbit were selected as candidates. Finally, we obtained a multiple asteroids exploration sequence of three asteroids based on gravity assist technique and genetic algorithm, and an additional velocity impulse of 0.4?km/s was required. In the subsequent section, a sixth-degree inverse polynomial shape-based method is applied to the low-thrust trajectory design of 2010 TK₇, and the exploration sequence under the action of low-thrust propulsion was provided.     

Accurate approximation of in-ecliptic trajectories for E-sail with constant pitch angle

Propellantless continuous-thrust propulsion systems, such as electric solar wind sails, may be successfully used for new space missions, especially those requiring high-energy orbit transfers. When the mass-to-thrust ratio is sufficiently large, the spacecraft trajectory is characterized by long flight times with a number of revolutions around the Sun. The corresponding mission analysis, especially when addressed within an optimal context, requires a significant amount of simulation effort. Analytical trajectories are therefore useful aids in a preliminary phase of mission design, even though exact solution are very difficult to obtain. The aim of this paper is to present an accurate, analytical, approximation of the spacecraft trajectory generated by an electric solar wind sail with a constant pitch angle, using the latest mathematical model of the thrust vector. Assuming a heliocentric circular parking orbit and a two-dimensional scenario, the simulation results show that the proposed equations are able to accurately describe the actual spacecraft trajectory for a long time interval when the propulsive acceleration magnitude is sufficiently small.     

Proposal for a multiple-asteroid-flyby mission with sample return

Asteroid exploration provides a new approach to study the formation of the solar system and the planetary evolution. Choosing a suitable target and designing of feasible profile for asteroid mission are challenging due to constraints such as scientific value and technical feasibility. This paper investigates a feasible mission scenario among the potential candidates of multiple flybys and sample return missions. First, a group of potential candidates are selected by considering the physical properties and accessibility of asteroids, for the sample return missions. Second, the feasible mission scenarios for multiple flybys and sample return missions to various spectral-type asteroids are investigated. We present the optimized design of preliminary interplanetary transfer trajectory for two kinds of missions. One is the single sample return mission to asteroids with various spectral types. The other is the multiple flybys and sample return mission to several asteroids. In order to find the optimal profiles, the planetary swing-by technique and Differential Evolution algorithm are used.     

Integrated design of solar thermal propulsion system for microsatellite with liquid ammonia as propellant

This paper presents a design of solar thermal propulsion (STP) system for microsatellite with liquid ammonia as propellant. The system was equipped with two concentrators, which were respectively placed in the tank and thrust chamber for propellant supply and heating. A platelet heat exchanger was adopted to heat the propellant in the chamber, and the fluid–solid coupling effect between the wall and the gas was considered. Meanwhile, the effects of satellite mass, initial orbit, nozzle size and target temperature on the performance of STP system were analyzed. The results show that for microsatellites with a total mass of 100 kg, the STP system can fully heat the propellant to more than 2050 K, generate an intermittent thrust of about 26 N, and enable the satellite to obtain a velocity increment of more than 1470 m/s within 19 days, consuming only 42 kg of propellant, which can directly meet the transfer mission from the geostationary transfer orbit (GTO) to the geostationary orbit (GEO). The maximum velocity increment could reach more than 1950 m/s when the propellant was completely consumed; Changing the mass and initial orbit of the satellite will not affect the thrust and specific impulse. Satellites with smaller mass will spend less time and propellant during orbit transfer. The lower is the perigee height of the initial orbit, the greater is the propellant consumption, while the shorter is the time of orbit transfer; The reduction of nozzle throat size and target temperature will lead to the increase of specific impulse and the decrease of orbital transfer time, but the reduction of thrust.     

Time-optimal trajectories to circumsolar space using solar electric propulsion

The aim of this paper is to explore the capabilities of a solar electric propelled spacecraft on a mission towards circumsolar space. Using an indirect approach, the paper investigates minimum time of transfer (direct) trajectories from an initial heliocentric parking orbit to a desired final heliocentric target orbit, with a low perihelion radius and a high orbital inclination. The simulation results are then collected into graphs and tables for a trade-off analysis of the main mission parameters. Finally, a comparison of the performance between a solar electric and a (photonic) solar sail based spacecraft is discussed.     

End-to-end trajectory design for a solar-sail-only pole-sitter at Venus,Earth, and Mars

The concept of a pole-sitter has been under investigation for many years, showing the capability of a low-thrust propulsion system to maintain a spacecraft at a static position along a planet’s polar axis. From such a position, the spacecraft has a view of the planet’s polar regions equivalent to that of the low- and mid-latitudes from geostationary orbit. Previous work has hinted at the existence of pole-sitters that would only require a solar sail to provide the necessary propulsive thrust if a slight deviation from a position exactly along the polar axis is allowed, without compromising on the continuous view of the planet’s polar region (a so-called quasi-pole-sitter). This paper conducts a further in-depth analysis of these high-potential solar-sail-only quasi-pole-sitters and presents a full end-to-end trajectory design: from launch and transfer to orbit design and orbit control. The results are the next steppingstone towards strengthening the feasibility and utility of these orbits for continuous planetary polar observation.     

全电推进卫星轨道优化的推力同伦解法

全电推进卫星的入轨过程是一个典型的多圈小推力轨道优化问题,由于其推力器加速度小,变轨圈数多,造成其最优理论解的求解较困难。为解决该问题,利用最优控制理论建立了全电推进卫星变轨优化的间接法模型,将变轨优化问题转化为协态变量初值猜测的两点边值问题。从大推力问题开始,通过遗传算法获得大范围猜测值并结合系列二次规划方法获得大推力的精确解。采用推力同伦思想,使用逐渐缩小推力的方式完成小推力问题的求解。仿真算例表明,采用推力同伦的方法,通过数十次的推力缩减即可有效解决多达上百圈变轨的静止轨道全电推进卫星入轨优化问题。     

太阳帆推进任务的快速仿真方法

研究太阳帆的力学特性和轨道控制设计方法,导出太阳帆的无奇点控制律.提出通过STK中MATLAB语言编写的嵌入式脚本（Plug in Script）来将由控制律得到的光压力加速度矢量,添加到STK轨道计算力学模型中,从而进行轨道控制的方法.仿真结果表明,对于常规方法难以进行仿真分析的航天器动力学模型（如太阳帆）,所提出的方法能快速灵活地支持其相应的任务,并增强任务场景的可视化,从而实现利用STK丰富的功能特性进行复杂航天任务的设计、分析和验证.     

Contingency target assessment,trajectory design,and analysis for NASA’s NEA scout solar sail mission

The presented study examines contingency target selection and trajectory design for NASA’s Near-Earth Asteroid Scout mission under the assumption of a missed lunar gravity assist. Two previously considered asteroids are selected as potential targets for the given scenario based on favorable orbital characteristics for launch dates ranging from June 27, 2020 through July 26, 2020. Initially, a simplified circular restricted 3-body problem + ideal solar sail model is utilized to survey trajectory options for a month-long launch window. Selected solutions from this data set are then converged in an N-body ephemeris + non-ideal sail model. Results suggest that NEA Scout can still perform asteroid rendezvous mission under the missed lunar gravity assist scenario with new targets, 2019 GF1, 2018 PK21, and 2007 UN12, based on the target launch dates. Further target assessment is carried out for 165 days beyond the current June 27, 2020 launch date.     

深远空连续推进动力与施图林格解的解析

针对太阳系远深距离的探测将是人类下一阶段深空探测活动的主要目标。这一目标的实现依赖于探测器连续推进动力技术的突破。从描述连续常值推力下太空飞行的施图林格解出发,对其中反映的深远空飞行任务有效载荷比、任务时间、飞行距离等关键参数与发动机性能之间的关系进行了深入分析。给出了在特定任务时长、特定飞行距离要求下发动机比冲、功率需要满足的条件及其对有效载荷比、最终飞行速度等指标的影响。此外,基于二体轨道动力学对太阳系行星探测的大椭圆转移轨道和转移能量进行了推导,并对连续推力的太阳帆任务方案涉及的关键技术指标做了理论性的计算。这些结论是对深空探测连续推力方案基础理论的归纳,可以为我国未来开展深远空探测活动提供重要的启发和指导。     

利用频谱分析确定空间目标太阳翼定向模式

分析对比了空间目标与可控飞行器轨道确定的不同,论述了太阳翼定向模式的不同对轨道运动特性的影响。通过地面站的测距与测角数据,推算出测量数据的地心距频谱。假定一种太阳翼定向模式,利用测量数据进行轨道确定,利用定轨数据计算地心距频谱并与测量数据的频谱比对,由此可确定太阳翼定向模式。利用数值仿真对近地太阳同步轨道和地球静止轨道两种情况进行了验证,仿真结果表明该思路和方法是有效的。     

基于电推力器进行南北位置保持的一种地球同步轨道注入参数方法

摘要： 电推力器在静止轨道卫星上应用越来越广泛,特别是基于电推力器进行南北位置保持,可以有效节省推进剂.提出改进的GPS星历参数解析算法,在此基础上考虑包含电推力模型在内多摄动项模型进行地面精密轨道计算,采用微分修正法,提出一种地球同步轨道注入参数方法,该方法可应用于星上自主完成基于电推力器的南北位置保持.仿真算例表明使用该方法得到的轨道注入参数,卫星能够在保证姿态确定精度的同时,完成南北位置保持任务.     

基于金星共振借力的太阳抵近探测任务轨道设计

对于太阳抵近探测任务,从地球直接发射探测器至太阳附近需要消耗巨大能量,通过多次金星借力飞行,可有效降低地球发射能量C₃及中途变轨的燃料消耗.本文研究基于金星共振借力的太阳抵近探测任务轨道优化设计,建立了连续共振借力和混合共振借力的转移轨道优化设计模型,并针对2025—2028年的发射窗口开展太阳抵近探测任务轨道优化设计.仿真结果表明,相比连续共振借力,混合共振借力可以有效缩短太阳抵近探测任务的轨道转移时间,对于地球发射能量C₃和中途变轨燃料消耗的影响未见明显的规律性,能量降低与序列中的共振比相关.      

环月轨道交会的载人登月任务轨道与窗口规划

针对环月轨道(Low Lunar Orbit,LLO)共面交会支持的"人货分离"载人登月任务,提出了一种任务窗口与轨道一体化规划方法。分析了基于LLO共面交会的"人货分离"载人登月任务的基本流程和工程约束;针对任务各阶段窗口与轨道求解问题,提出了以动力下降时刻为迭代初值的窗口规划策略,并建立了高精度模型下的环月轨道、双二体模型下的人员和货物运输轨道规划模型。以载人月球探测中国科学家命名的环形山为假想背景,给出仿真实例,仿真结果验证了文章所提方法的正确性,为探月工程任务提供了一种有效的窗口与轨道设计工具。     

基于分段常值的全电推进GEO卫星制导策略

电推进技术因其比冲高的技术特点在GEO轨道转移中应用可大大减少燃料质量,提高有效载荷质量比,延长任务寿命等。针对全电推进GEO卫星入轨的轨迹优化和制导问题,首先利用间接法获得小推力燃料最优GEO轨道转移的数值解,提出一种多项式曲线拟合最优轨迹的方法,多项式曲线形式简单,可作为参考轨道在星上存储和使用。在多项式参考轨道的基础上,建立了一种分段常值推力跟踪参考轨道的闭环制导策略,在常值推力条件下,轨道要素控制量与控制力有解析关系,简化了制导律设计;将多圈轨道转移问题分解为多个单圈轨道优化问题。结果显示,本文提出的分段常值跟踪制导策略跟踪精度高,和最优轨道相比多消耗7%的燃料。本制导策略控制结构简单,易于工程实施。     

基于三体系统平动点轨道的环日全景任务轨道设计

在空间开展太阳观测是研究太阳活动周、太阳爆发、极端天气等事件起源的重要手段。环日全景探测计划是为实现从黄道面360°全方位观察太阳行星际空间而提出的。本文针对环日全景探测计划,构建了基于三体系统平动点低能量轨道的环日全景轨道部署方法。该方法以日–地L1/L2点Halo轨道幅值及Halo轨道离轨点为变量,以转移轨道飞行时间、入轨机动大小为评价指标,基于三体系统不变流形构建环日全景的转移轨道,并开展轨道优化设计。采用等高线图对设计变量及任务成本进行全局分析。仿真计算表明,轨道部署无法同时满足飞行时间最短与入轨机动最小的要求。设计了轨道机动约束条件下的最优飞行时间解,并给出了基于长三甲运载火箭的一箭双星发射及入轨方案。      

载人火星探测飞行方案

对世界各国载人火星探测的研究情况进行了简要综述,研究了国内外有关载人火星探测飞行方案,提出了载人火星探测方案确定的原则和方案基本思想.给出了一种载人火星探测飞行方案的总体设计,包括飞行轨道方案和载人火星飞船方案等.尤其对轨道设计的重要的两个参数——速度增量和飞行时间进行了详细计算.最后给出了飞行轨道选择、火星飞船从地球到火星和从火星返回地球等的轨道方案和火星飞船各组成部分方案的详细设计结果.     

地月拉格朗日L2点中继星轨道分析与设计

地月L2点位于地月连线的延长线上,在地月L2点运行的卫星可以连续观测月球背面,解决月球背面与地球之间的通讯问题,在月球背面着陆探测任务中起着至关重要的作用。对从地球出发、利用月球引力辅助变轨、形成地月L2点的轨道进行了研究,分析了发射窗口、地月转移时间、近月点高度、近月点倾角、轨道振幅等多项因素对转移轨道和使命轨道特性的影响,寻求满足地月L2点中继任务需求的飞行轨道。通过分析研究,文章明确了转移和使命轨道的相关特性,可为中继星任务轨道的参数设计和优化提供有益参考。     

Solar sails for perturbation relief: Application to asteroids

A major cause of spacecraft orbital variation comes from natural perturbations, which, in close proximity of a body, are dominated by its non-spherical nature. For small bodies, such as asteroids, these effects can be considerable, given their uneven (and uncertain) mass distribution. Solar sail technology is proposed to reduce or eliminate the net secular effects of the irregular gravity field on the orbit. Initially, a sensitivity analysis will be carried out on the system which will show high sensitivity to changes in initial conditions. This presents a challenge for optimisation methods which require an initial guess of the solution. As such, the Genetic Algorithm (GA) is proposed as the preferred optimisation method as this requires no initial guess from the user. A multi-objective optimisation is performed which aims to achieve a periodic orbit whilst also minimising the effort required by the sail to do so. Given the system sensitivity, the control law for one orbit is not necessarily applicable for any subsequent orbit. Therefore, a new method of updating the control law for subsequent orbits is presented, based on linearisation and use of a Control Transition Matrix (CTM). The techniques will later find application in a multiple asteroid rendezvous mission with a solar sail as the primary propulsion system.