基于时间最短的SGKW共面打击轨道优化设计

天基对地打击动能武器(SGKW)用于从太空对地面高价值战略目标进行快速、准确的打击。针对最短打击时间要求,研究了SGKW共面打击轨道的优化设计方法。首先建立了SGKW的无量纲化平面运动模型,然后利用庞特里亚金极大值原理将时间最短共面打击轨道的最优控制问题转化为两点边值问题。由于约束条件中存在优化参数,一种基于"遗传算法 序列二次规划"的组合优化算法被用于求解未知参数。仿真结果验证了上述方法的有效性。     

SGKW轨道在线生成技术研究

天基对地打击动能武器(SGKW)用于从太空对地面高价值战略目标进行快速、准确的打击.针对作战实时性要求,探讨了SGKW的轨道在线生成技术.首先,利用庞特里亚金极大值原理将时间最短打击轨道的最优控制问题转化为两点边值问题.由于约束条件中存在优化参数,一种基于"遗传算法+序列二次规划"的组合优化算法被用于求解伴随变量初值和再入点参数.为了提高轨道生成的速度,在大量离线优化数据的基础上建立了BP神经网络预测模型,其预测值通过序列二次规划算法稍加修正,即可满足相应任务的落点精度要求.仿真结果验证了上述方案的有效性.     

伪谱法在SGKW轨道快速优化中的应用

天基对地打击动能武器(SGKW)用于从太空对地面高价值战略目标进行快速、准确的打击.针对作战实时性要求,探讨了基于伪谱法的SGKW轨道快速优化技术,该方法的实质是将最优控制问题转化为非线性规划问题.为了提高优化计算的快速性,提出了轨道分段生成的策略,即首先根据参考配点获得满足落点要求的再入点参数,再将求得的再入点参数作为终端约束并运用序列二次规划算法对过渡段进行优化.仿真结果验证了上述方案的有效性.     

一种预测校正制导中攻角剖面的在线规划方法

提出一种预测校正制导中基于可达性量化评估的攻角剖面在线规划方法。首先采用最大纵程、最大横程、最小纵程3个关键参数构建可达区模型,并通过大量的离线弹道计算构建上述3个关键参数与升力修正因子、阻力修正因子和常值攻角的三维插值数据库。提出一种可达性量化因子用于反应期望终端位置的可达性。在飞行中对升力修正因子和阻力修正因子进行在线辨识,在攻角剖面的在线规划时以可达性量化因子最小为优化目标。仿真表明采用该方法可以获得较高的可达性判断正确率,攻角剖面的在线规划可以提高预测校正制导的成功率。本方法对于无动力升力式飞行器的制导策略在线决策和优化具有参考和应用价值。     

天基对地打击武器飞行过程动力学分析与建模仿真

天基对地打击武器用于从太空对地面战略目标和高价值目标进行快速、准确的打击。本文运用空间飞行器动力学,在特定条件下,对天基对地打击武器的作战飞行过程分大气层外和大气层内2个阶段进行动力学分析,并利用建模仿真的方法得出在不同再入角和不同运行轨道高度的情况下飞行过程的相关数据,验证了天基对地打击武器作战迅速、打击时间短的特性,表明其在未来潜在的军事应用价值。     

再入飞行器可达区域近似算法及地面覆盖研究

在无旋圆地球假设下,利用飞行器最大纵程、最小纵程和最大横程三个典型性能指标,建立了飞行器纵程和横程之间近似椭圆分布的解析关系式,通过该关系式能够快速得到纵程和横程对应的边界曲线,并根据球面三角形将纵程和横程的对应关系转化为经纬度关系,从而得到飞行器地球表面可达区域的边界线.通过与Legendre伪谱法计算所得最优解的比较发现,在不同经纬度,以不同航向角再入后,最优化方法计算得到的边界点与解析方法计算的边界曲线分布基本一致,并仿真分析了不同弧段再入后飞行器地面可达区域的变化特点,针对顺行轨道和逆行轨道完成了再入飞行器地面覆盖范围的计算.该解析方法通过3个典型指标就能够快速计算飞行器再入可达区域,有助于飞行器着陆场快速选择和初期轨道快速设计.     

一种近似解析的再入滑翔飞行器可达域快速生成方法

针对再入滑翔飞行器可达域快速计算问题，研究了一种近似解析求解方法。首先，将可达域求解问题归结为两类纵、横程极值轨迹规划问题，利用极大值原理推导出满足最大纵程/横程要求时飞行攻角对应当前最大升阻比攻角的结论，通过仿真验证了该结论的正确性。其次，结合再入滑翔飞行器动力学特性，设计了不同极值轨迹的倾侧角变化规律，并通过梯度下降方法优化得到相关参数。最后，在不同飞行器初始状态条件下，通过数值仿真得到了所提方法的可达域。其与伪谱法优化结果基本一致，证明了所提方法的有效性。     

天基对地手打击动能武器再入解析预测制导技术

天基对地打击动能武器用于从太空对地面高价值战略目标进行快速、准确的打击.为了在各种干扰因素的作用下仍能保证足够的命中精度,动能弹头必须实施再入制导.针对实时性要求,探讨了一种解析预测制导方法.首先详细推导了零攻角再入弹道参数的三维解析解,在此基础上借鉴牛顿迭代法的思想设计了速度倾角与航向角的迭代修正算法,并最终将其用于制导指令的生成.仿真结果表明,解析预测制导方法能有效提高再入弹头的落点精度,且实时性强.此外,通过对不同制导参数下的制导性能进行分析还发现,制导步长取450～600米、精度参数取20～50米最为合适.     

升力式再入飞行器两种可达区域计算方法的探讨

对2种计算升力式再入飞行器可达区域的方法进行了分析和探讨。首先建立了飞行器三自由度运动方程模型和升力式飞行器基于准平衡假设下的简化模型;然后基于最优控制方法,考虑过程约束和终端约束推导了闭环倾侧角控制律,采用双参数搜索法求取不同纵程下的最大横程;最后,以一种升力式飞行器为例分别用优化方法和常值倾侧角法进行仿真分析,结果表明常值倾侧角法可以代替优化方法进行分析与设计。     

天基对地打击动能武器再入解析预测制导技术

天基对地打击动能武器用于从太空对地面高价值战略目标进行快速、准确的打击。为了在各种干扰因素的作用下仍能保证足够的命中精度,动能弹头必须实施再入制导。针对实时性要求,探讨了一种解析预测制导方法。首先详细推导了零攻角再入弹道参数的三维解析解,在此基础上借鉴牛顿迭代法的思想设计了速度倾角与航向角的迭代修正算法,并最终将其用于制导指令的生成。仿真结果表明,解析预测制导方法能有效提高再入弹头的落点精度,且实时性强。此外,通过对不同制导参数下的制导性能进行分析还发现,制导步长取450-600米、精度参数取20-50米最为合适。     

The Use of Quaternions in the Optimal Control Problems of Motion of the Center of Mass of a Spacecraft in a Newtonian Gravitational Field: I

The problem of optimal control is considered for the motion of the center of mass of a spacecraft in a central Newtonian gravitational field. For solving the problem, two variants of the equations of motion for the spacecraft center of mass are used, written in rotating coordinate systems. Both the variants have a quaternion variable among the phase variables. In the first variant this variable characterizes the orientation of an instantaneous orbit of the spacecraft and (simultaneously) the spacecraft location in this orbit, while in the second variant only the instantaneous orbit orientation is specified by it. The suggested equations are convenient in the respect that they allow the general three-dimensional problem of optimal control by the motion of the spacecraft center of mass to be considered as a composition of two interrelated problems. In the first variant these problems are (1) the problem of control of the shape and size of the spacecraft orbit and (2) the problem of control of the orientation of a spacecraft orbit and the spacecraft location in this orbit. The second variant treats (1) the problem of control of the shape and size of the spacecraft orbit and the orbit location of the spacecraft and (2) the problem of control of the orientation of the spacecraft orbit. The use of quaternion variables makes this consideration most efficient. The problem of optimal control is solved on the basis of the maximum principle. Several first integrals of the systems of equations of the boundary value problems of the maximum principle are found. Transformations are suggested that reduce the dimensions of the systems of differential equations of boundary value problems (without complicating them). Geometrical interpretations are given to the transformations and first integrals. The relation of the vectorial first integral of one of the derived systems of equations (which is an analog of the well-known vectorial first integral of the studied problem of optimal control) with the found quaternion first integral is considered. In this paper, which is the first part of the work, we consider the models of motion of the spacecraft center of mass that employ quaternion variables. The problem of optimal control by the motion of the spacecraft center of mass is investigated on the basis of the first variant of equations of motion. An example of a numerical solution of the problem is given.     

The Use of Quaternions in the Optimal Control Problems of Motion of the Center of Mass of a Spacecraft in a Newtonian Gravitational Field: II

The problem of a rendezvous in the central Newtonian gravitational field is considered for a controlled spacecraft and an uncontrollable spacecraft moving along an elliptic Keplerian orbit. For solving the problem, two variants of the equations of motion for the spacecraft center of mass are used, written in rotating coordinate systems and using quaternion variables to describe the orientations of these coordinate systems. In the first variant of the equations of motion a quaternion variable characterizes the orientation of an instantaneous orbit of the spacecraft and the spacecraft location in the orbit, while in the second variant it characterizes the orientation of the plane of the spacecraft instantaneous orbit and the location of a generalized pericenter in the orbit. The quaternion variable used in the second variant of the equations of motion is a quaternion osculating element of the spacecraft orbit. The problem of a rendezvous of two spacecraft is formulated as a problem of optimal control by the motion of the center of mass of a controlled spacecraft with a movable right end of the trajectory, and it is solved on the basis of Pontryagin's maximum principle.     

On solving the problems of optimization of trajectories of many-revolution orbit transfers of spacecraft

The problem of optimal control over many-revolution spacecraft orbit transfers between circular coplanar orbits of satellites is considered. The spacecraft flight is controlled by a thrust vector of a jet engine with restricted thrust (JERT). The mass expenditure is minimized at a limited time of flight. The optimal control problem is solved based on the maximum principle. The boundary value problem of the maximum principle is solved numerically using the shooting method. A modified computation scheme of the shooting method is suggested (multi-point shooting), as well as a method (correlated with the scheme) of choosing the initial approximation with the use of a solution to the optimization problem in the impulse formulation. The scheme and method allow one to construct many-revolution spacecraft orbit transfers.     

轨道机动的时间-能量综合最优控制

围绕航天器快速精确轨道机动问题,探讨在持续小推力作用下,航天器轨道机动中时 间和能量综合最优控制的技术和方法。基于Pontryagin最小(大)值原理,针对目标轨道为平 面和空间椭圆的情况,推导了时间-能量综合最优控制的Hamilton正则方程组、终端条件 、横截条件和最优控制的表达式,应用数值方法求解正则微分方程组的两点边值问题,得到 了最优控制的数值解,包括最小时间、最小能量、最优轨道、最优控制时变曲线和最优反馈 控制曲线等,实现轨道机动最优控制的精确数值模拟。从数值结果的对比分析中得出了一些 有意义的结论,可供工程实际参考。
     

Solution of the Problem of Optimal Spacecraft Launching into Orbit Using Reactive Acceleration and Solar Sail in Kustaanheimo–Stiefel Variables

Cosmic Research - Using the Pontryagin maximum principle and the Kustaanheimo–Stiefel variables, the spatial problem of optimal launching into a given orbit of a spacecraft (SC) controlled by...     

多基站ISAR多运动目标成像与横向定标研究

在多基站ISAR多运动目标回波模型的基础上,采用时间-调频斜率分布估计各基站多个目标的信号参数,基于CLEAN算法实现了各基站多个目标的信号分离。利用多个基站获得的目标的信号参数和多个ISAR图像,估计目标的运动参数,实现多个目标的横向尺度标定,同时给出了多基站ISAR多目标成像和横向定标的约束条件。仿真实验验证了多基站ISAR多目标成像和横向定标算法。

     

The use of three-impulse transfer to insert the spacecraft into the high Moon Artificial Satellite orbits

The problem of the optimal spacecraft’s insertion from the Earth into the high circular polar Moon Artificial Satellite’s orbit (MAS) with a radius of 4000–8000 km has been investigated. A comparison of single- and three-impulse insertion schemes has been performed. The analysis was made taking into account the disturbances from the lunar gravity field harmonics and the gravity fields of the Earth and the Sun, as well as the engine’s limited thrust. It has been shown that the three-impulse transfer from the initial selenocentric hyperbola of the approach into the considered final high MAS orbit is noticeably better with respect to the final mass than the ordinary single-impulse deceleration. The control parameters that implement this maneuver and provide nearly the same energy expenses as in the Keplerian case have been presented. It was found that, in contrast to the Keplerian case, in the considered case of the real gravity field, there is the optimal maximum distance of the maneuver. Recently, the Moon exploration problem became actual again.