大椭圆轨道气动变轨与轨道部署

提出了一种服务飞行器的大椭圆轨道气动变轨方案.根据速度轨道最优参数确定气动变轨动力学方程及参数,计算了变轨第一、二阶段的能力.研究了轨道部署方案,讨论了大椭圆轨道近地点幅角与轨道倾角,以及待机轨道覆盖性能.研究表明:3个空间飞行器就可100%覆盖低轨道主要目标群.     

空间交会对接寻的段三次变轨策略分析

根据第一次变轨采用地面或自主引导,以及后两次变轨策略,提出了采用Hohmann变轨与CW双脉冲最优变轨、CW双脉冲变轨 中途一次修正,以及CW三脉冲最优变轨的空间交会对接寻的段三次变轨方案.仿真分析了方案中Hohmann变轨椭圆转移轨道远地点高度、CW双脉冲转移时间、总转移时间或第二次变轨时间与总特征速度的关系,并给出寻的段变轨策略建议.     

一类椭圆轨道交会中的动力学方程研究

研究了一类追踪器和目标器轨道半长轴相差不大、轨道面外的距离相差不大的小偏心率椭圆交会的动力学问题.首先选择合适的圆轨道上的点建立参考系,推导出针对圆轨道参考系的无量纲化线性定常方程,并获得相应的相对状态;接着讨论该方程在小偏心椭圆轨道两冲量交会中的应用;最后进行数值仿真,验证动力学方程和制导策略,并与CW方程及制导策略的相关仿真进行比较.仿真结果表明本文给出的动力学方程的精度优于CW方程,能有效解决这类椭圆交会问题.     

共面变相位星座多目标交汇序列优化方法

针对空间多目标交汇时空约束繁杂、轨道能量影响变量多的特点,从空间几何解析法入手,提出基于穿越点的共面变相位交汇方法,在Lambert椭圆轨道机动的基础上,给出了机动装置共面变轨相位快速确定方法,总结了变轨速度需求以及运载段分离条件已知情况下相位角和转移时间的包络区,并基于覆盖效能对目标发射序列进行了优化,仿真验证了相关结论,得到的结果符合预期。     

GEO卫星电推进与化学推进组合变轨方案研究

针对GEO卫星采用电推进和化学推进系统进行变轨的问题,在给定轨道控制模型的基础上,给出实现轨道半长轴、偏心率和倾角3个参数单独调整和联合控制的控制律。结合假设的运载火箭发射能力约束,给出航天器不同入轨高度对应的初始质量。以此为约束,给出化学推进和电推进不同组合的6种GTO向GEO转移方案,并对比分析各方案完成变轨所需的时间、推进剂、速度增量以及完成变轨后的质量剩余情况,研究方法可为电推进系统在航天领域的应用提供参考。     

基于双曲线B平面参数的快速微分修正月地返回轨道精确设计

月地返回轨道设计是探月三期月球采样返回任务中的重要内容之一,其约束条件较地月转移轨道复杂.此外,微分修正算法对于初值有很强的敏感性,且不易搜索得到初值.本文提出选取月心段出口点的双曲线B平面参数作为第一次迭代的目标值,选取地心段约束值作为第二次迭代的目标值,可有效的减少迭代次数和迭代时间,完成搜索初值过程.针对直接返回型轨 道和间接返回型轨道的设计问题,使用基于双曲线B平面参数的快速微分修正月地返回轨道精确设计方法,满足了对应的约束条件,易于求取变轨点的位置矢量和速度矢量,得到标称返回轨道.最后针对2种返回轨道类型的算例说明该方法有效.     

探测器月地转移入射变轨策略优化设计

在探测器月球采样返回任务中,从环绕月球的轨道返回地球的转移入射考虑采用一次变轨或两次变轨的方案。其中,月地转移入射一次变轨方案要求采用大推力变轨发动机,难以实现高精度轨道控制。为此,文章提出将月地转移入射一次变轨机动改为分两次执行,两次间隔大约1天的优化策略。首先讨论了月地转移入射一次变轨的基本设计方法,然后给出了月地转移入射两次变轨方案的优化设计算法流程,并采用数值算例进行了说明。此外,还对月地转移入射三脉冲变轨方案和低能月地转移轨道展开了讨论,并在多方案比较分析的基础上,推荐了月球采样返回任务采用的月地转移入射变轨策略。     

非开普勒轨道下卫星交会变轨决策研究

针对卫星在非开普勒轨道下的运动特征,提出了研究卫星交会变轨决策的方法。首先,在Hill坐标下,采用比例导引方法求解卫星进行变轨机动时所需要的速度增量。然后,再通过Lagrange插值算法解出卫星在任意特征点上变轨机动所需要的速度增量,根据数字仿真得出大量的仿真数据并画出满足变轨决策条件的决策曲面。最后,对决策曲面的涵义做出分析,说明了该方法对于研究非开普勒轨道下卫星变轨决策问题具有一定的意义。     

中国航天用几种固体火箭发动机简介(六)——FG-47长二丙改进型火箭变轨发动机

一、概况 FG-47发动机是发射铱星用的长征二号丙改进型运载火箭的变轨动力装置,由其将铱星从椭圆停泊轨道转移到630公里的圆形运行轨道,以建立低轨道全球     

嫦娥五号探测器多圈调相地月转移应急轨道设计与分析

针对嫦娥五号任务上升段末期火箭二级发动机可能出现的提前关机故障造成入轨半长轴偏差较大和中途修正速度增量超限问题,提出了多圈调相地月转移轨道应急控制策略.首先,分析了不同入轨半长轴偏差、中途修正时刻与中途修正速度增量消耗之间的关系;其次,针对半长轴偏差较大问题,基于微分改正算法与B平面参数,设计了解析窗口搜索与多圈调相地...     

载人月球极地探测地月转移轨道设计

针对载人月球极地探测任务,采用一种自由返回轨道与三脉冲机动轨道相结合的地月转移轨道方案。关于自由返回轨道部分的设计,建立了基于近月点伪参数的两段拼接模型,采用一种考虑地球扁率修正的改进多圆锥截线法进行求解,仿真结果显示改进的多圆锥截线法具有更高的求解精度,可为精确设计提供更好的初值;关于三脉冲机动轨道部分的设计,基于混合轨道模型,采用特殊点变轨和Lambert算法相结合的方法进行计算,仿真结果显示该方法能够有效地降低速度增量的消耗。最后,通过大量的仿真计算,对轨道的速度增量特性进行了分析。研究结论可为未来载人月球极地探测地月转移轨道方案的设计提供重要参考。     

Nonlinear feedback control of low-thrust orbital transfer in a central gravitational field

This paper investigates the problem of continuous-thrust orbital transfer using orbital elements feedback from a nonlinear control standpoint, utilizing concepts of controllability, feedback stabilizability and their interaction. Gauss's variational equations (GVEs) are used to model the state-space dynamics of motion under a central gravitational field. First, the notion of accessibility is reviewed. It is then shown that the GVEs are globally accessible. Based on the accessibility result, a nonlinear feedback controller is derived which asymptotically steers a spacecraft form an initial elliptic orbit to any given elliptic orbit. The performance of the new controller is illustrated by simulating an orbital transfer between two geosynchronous Earth orbits. It is shown that the low-thrust controller requires less fuel than an impulsive maneuver for the same transfer time. Closed-form, analytic expressions for the new orbital transfer controller are given. Finally, it is proven, based on a topological nonlinear stabilizability test, that there does not exist a continuous closed-loop controller that can transfer a spacecraft onto a parabolic escape trajectory.     

The optimization of the orbital Hohmann transfer

There are four bi-impulsive distinct configurations for the generalized Hohmann orbit transfer. In this case the terminal orbits as well as the transfer orbit are elliptic and coplanar. The elements of the initial orbit a₁, e₁ and the semi-major axis a₂ of the terminal orbit are uniquely given quantities. For optimization procedure, minimization is relevant to the independent parameter e_T, the eccentricity of the transfer orbit. We are capable of the assignment of minimum rocket fuel expenditure by using ordinary calculus condition of minimization for |ΔV_A|+|ΔV_B|=S.We exposed in detail the multi-steps of the optimization procedure. We constructed the variation table of S(e_T) which proved that S(e_T) is a decreasing function of e_T in the admissible interval [e_Tmin,e_Tmax]. Our analysis leads to the fact that e₂=1 for e_T=e_Tmax, i.e. the final orbit is a parabolic trajectory.     

Venus transfer design by combining invariant manifolds and low-thrust arcs

The design of interplanetary trajectories based on patched circular restricted three body models is gradually becoming a valuable alternative to the classical patched conic approach. The main advantage offered by such a model is the possibility to exploit the manifold dynamics to move naturally far from or toward a body. Generally, propulsive maneuvers are required to match these structures. Low-thrust arcs offer the possibility to have a significant propellant mass reduction when moving from manifold to manifold. The aim of this paper is to present a methodology to design low-thrust trajectories between two planetary orbits connecting the manifolds of two circular three body systems. The approach is based on a grid search on the main parameters governing the solution to identify those trajectories moving within the manifold images on given Poincarè sections. The value of the Jacoby constant of the target libration point periodic orbit is chosen as stop condition for the thrusting phases. Ballistic arcs follow up to the proper Poincarè section intersection. A grid search for an Earth to Venus transfer is presented as test case.     

Optimum low thrust transfer to geosynchronous orbit

The low thrust transfer for geosynchronous mission has been studied by many investigators from the viewpoint of optimization in case of continuous thrust. This paper discusses the possibility of fuel saving to attain a geosynchronous orbit by introducing coast phases during each revolution. In advance of optimizing the whole transfer mission, optimization during a single revolution is treated, and it is shown that the entrance and the exit of optimal coasting arcs are expressed by a sixth order equation, which, in case of coplanar transfer, degenerates into a cubic equation, with respect to the cosine of true longitude. Then an optimum transfer to a geosynchronous orbit, including coast phases in each revolution, is simulated. Computational results for typical initial conditions are shown to be compared with those for all-propulsion cases.     

轨道机动过程中推力加速度的实时最小方差估计

飞行器轨道机动过程中，为跟踪、定位机动目标和干预机动控制过程，需要统计处理离散的雷达观测量实时估计推进发动机的推力，进而确定飞行器的瞬时轨道参数。本文所述算法是该工程问题的探讨和解决方案。文章建立了轨道机动过程中连续变质量运动模型和离散雷达量测模型，推进发动机的质量秒耗量作为表征推力加速度的一个近似常量，应用扩展卡尔曼滤波对离散的雷达测量数据进行顺序统计处理给出秒耗量的最小方差估计；文章详细地推导了线性化量模型的变分方程和观测矩阵；仿真结果表明该算法能快速、准确地估计推进发动机的质量秒耗量和向机动目标施加的实际推力。     

基于解析梯度的经典Lambert问题迭代求解方法

针对现有求解模型复杂、收敛速度慢等问题,在将经典Lambert转移问题转化为超越方程的基础上,提出一种基于解析梯度的Lambert问题迭代求解算法。选择转移轨道的真近点角为迭代变量,导出转移时间关于真近点角的解析梯度,构造一种基于解析梯度的牛顿迭代算法,降低了算法计算复杂度。理论分析表明该算法具有二阶以上的收敛速度。依据偏心率向量与转移轨道形状的关系,通过几何方法分析得到转移轨道在初始位置处的速度约束条件,推导转移轨道真近点角的最大值和最小值的解析表达式,并采用线性插值方法确定迭代初值,进一步提高了迭代算法的收敛速度。数学仿真结果表明在各种转移条件下算法均能快速收敛,采用所给出的初值选取方法初值确定精度高,进而能够加快收敛速度,而与较割线法相比较收敛速度快、计算量小,验证了所提出算法的有效性。     

Orbit determination and control for the European Student Moon Orbiter

This paper presents the preliminary navigation and orbit determination analyses for the European Student Moon Orbiter. The severe constraint on the total mission Δv and the all-day piggy-back launch requirement imposed by the limited available budget, led to the choice of using a low-energy transfer, more specifically a Weak Stability Boundary one, with a capture into an elliptic orbit around the Moon. A particular navigation strategy was devised to ensure capture and fulfil the requirement for the uncontrolled orbit stability at the Moon. This paper presents a simulation of the orbit determination process, based on an extended Kalman filter, and the navigation strategy applied to the baseline transfer of the 2011–2012 window. The navigation strategy optimally allocates multiple Trajectory Correction Manoeuvres to target a so-called capture corridor. The capture corridor is defined, at each point along the transfer, by back-propagating the set of perturbed states at the Moon that provides an acceptable lifetime of the lunar orbit.     

转移轨道航天器深层充放电效应仿真分析

为研究转移轨道卫星介质深层动态充电规律特征,针对卫星动态辐射环境的特点,基于FLUMIC思想,建立辐射带动态电子环境模式。针对动态辐射环境下星上介质深层充电的特征,使用辐射诱导电导率(RIC)模型和Geant4建立了适用于转移轨道卫星动态环境下的介质深层充电应用模型。首次对地球同步转移轨道(GTO)和嫦娥一号卫星调相轨道运行过程中介质深层充电情况进行了仿真分析。结果表明:转移轨道卫星在运行时会多次穿越辐射带区域,电子通量存在明显波动,这种波动性反映在材料的充电电位变化中。材料峰值充电电位分别为-2 846V和-4 110V,介质内部平衡电场均超过106 V/m,存在内放电风险,需要在工程设计中进行针对性防护。     

基于模型预测的轨道自主快速大机动控制

为改善太阳同步轨道(SSO)卫星轨道大机动经典控制方法的性能,提出了一种基于卫星动力学模型预测的轨道自主机动控制算法,用大推力轨控发动机开环控制加小推力器闭环修正进行控制。给出了制导策略和算法流程。理论分析和仿真结果表明,该控制算法适于卫星轨道自主快速机动控制,其轨控精度较高,收敛时间可满足快速性要求。