共查询到20条相似文献,搜索用时 187 毫秒
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针对人工太阳同步轨道的设计方法进行研究,通过施加法向连续推力调整升交点赤经(RAAN)变化率。首次推导了升交点赤经在变方向推力作用下的周期摄动平均值的精确计算公式,解决了已有近似方法对相关轨道参数的取值范围存在限制的问题,并给出了对应的轨道倾角周期摄动平均值计算公式。在分析J2项摄动对升交点赤经影响的基础上,给出了所需的法向连续推力幅值和一个轨道周期内对应的速度增量的计算方法。通过数值仿真,校验了计算公式的正确性,分析了实现人工太阳同步轨道的连续法向推力对轨道倾角的影响,给出了连续推力幅值随轨道参数的变化规律,并且提出了未来工程任务的应用建议。 相似文献
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针对现有的基于不变流形地月转移轨道设计方法存在的转移时间长、需要额外速度增量的缺点,本文利用平面圆型限制性三体问题下大幅值Lyapunov轨道的稳定流形,提出了一种地月低能转移轨道设计方法。首先计算与给定近月轨道相切的大幅值Lyapunov轨道作为参考轨道;然后根据小偏差值对稳定流形的影响确定Lyapunov轨道初始点的取值范围;再通过最近点截面确定与给定近地轨道相切的Lyapunov轨道的稳定流形;最后根据稳定流形的切点调整近月轨道半径,通过一条稳定流形直接连接近地轨道与近月轨道来实现地月转移。仿真结果表明,该方法仅需要两次速度增量,在能耗较低的前提下大大减少了探测器渐近Lyapunov轨道时的飞行时间,为地月低能转移轨道的设计提供了一种新思路。 相似文献
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首先给出三角平动点附近的高阶解析解,并计算了三种特殊的运动类型。以日–地+月系三角平动点附近无长周期运动分量的拟周期轨道作为目标轨道,探讨轨道保持问题。针对三角平动点任务的轨道保持问题,我们研究了两种轨道保持策略,分别为多点打靶轨道保持与重构目标轨道的策略。计算中,将轨道控制问题转化为非线性规划问题,并以优化方法求解。仿真表明优化方法在轨道保持问题求解方面非常有效。 相似文献
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针对大规模星座轨道预报存在卫星数量多、摄动方程强非线性等难题,提出一种基于平均速率矩阵的多星同步快速轨道预报算法。该算法首先基于哈密顿力学理论建立了二阶带谐项摄动下的轨道动力学模型,其次,利用无奇异轨道根数提出了一种近圆无奇异解析轨道预报模型,并基于Fourier-Bessel级数理论消除真近点角使模型只含有平近点角,简化预报计算过程,在此基础上,基于矩阵理论构造了多星轨道同步预报算法,实现多星轨道同步快速预报。以“星链”卫星星座为例进行仿真,结果表明:提出的方法能够将计算速度提高一个数量级,7天的轨道预报误差小于2.7 km。 相似文献
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面向航天器编队飞行的需求,对椭圆参考轨道航天器非线性周期相对运动条件进行研究,提出了确定椭圆参考轨道编队航天器非线性周期性相对运动条件的新方法。首先,考虑非线性、椭圆轨道等因素,通过哈密尔顿-雅可比(HJ)方程和正则摄动理论,推导了在任意非线性摄动下相对运动的模型和获得不需消耗任何燃料的周期性相对运动轨道的条件;然后,采用时域配点法,结合改进的列文伯格-马夸尔特(LM)法对周期性相对运动的初值进行求解;最后,设计数值仿真算例,利用上述条件,得到不消耗任何燃料的周期性绕飞轨道,由此验证了本文所提模型和方法的正确性。 相似文献
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Methods are proposed for constructing the orbits of spacecraft remaining for long periods of time in the vicinity of the L 2 libration point in the Sun-Earth system (so-called halo orbits), and the trajectories of uncontrolled flights from low near-Earth orbits to halo orbits. Halo orbits and flight trajectories are constructed in two stages: A suitable solution to a circular restricted three-body problem is first constructed and then transformed into the solution for a restricted four-body problem in view of the real motions of the Sun, Earth, and Moon. For a halo orbit, its prototype in the first stage is a combination of a periodic Lyapunov solution in the vicinity of the L 2 point and lying in the plane of large-body motion, with the solution for the linear second-order system describing small deviations of the spacecraft from this plane along the periodic solution. The desired orbit is found as the solution to the three-body problem best approximating the prototype in the mean square. The constructed orbit serves as a similar prototype in the second stage. In both stages, the approximating solution is constructed by continuation along a parameter that is the length of the approximation interval. Flight trajectories are constructed in a similar manner. The prototype orbit in the first stage is a combination of a solution lying in the plane of large-body motion and a solution for a linear second-order system describing small deviations of the spacecraft from this plane. The planar solution begins near the Earth and over time tends toward the Lyapunov solution existing in the vicinity of the L 2 point. The initial conditions of both prototypes and the approximating solutions correspond to the spacecraft’s departure from a low near-Earth orbit at a given distance, perigee, and inclination. 相似文献
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Low-thrust transfers between preset orbits are considered in the presence of perturbations of different origin. A simple method of finding the transfer trajectory is suggested, based on linearization of motion near reference orbits. The required accuracy of calculations is achieved by way of increasing the number of reference orbits. The method can also be used in the case of a large number of revolutions around the attracting center: no averaging of motion is required in this case. The suggested method is applicable as well, when the final orbit is specified partially and when there are constraints on the thrust direction. The optimal solution to the linearized problem is not optimal for the original problem; closeness of solutions to these two problems is estimated using a numerical example. Capabilities of the method are also illustrated by examples. 相似文献
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V. G. Petukhov 《Cosmic Research》2012,50(3):249-261
The problem of local optimization of interplanetary low-thrust trajectories is considered with the use of the maximum principle
and continuation numerical methods. Two types of problems are analyzed: problems with limited power and problems with limited
thrust. The latter problem is generalized by introducing the dependence of thrust and specific impulse on available electric
power. In order to reduce the problem of optimal control to a boundary value problem, the Pontryagin maximum principle is
used, and then, using the continuation method, this boundary value problem is reduced to the Cauchy problem. Variants of the
continuation method for optimizing low-thrust trajectories are presented in the paper, including a new method of continuation
for the limited thrust problem, which does not require any choice of the initial approximation for boundary values of conjugate
variables. 相似文献
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V. K. Saulskiy 《Cosmic Research》2016,54(4):313-324
Single satellites and multisatellite constellations for the periodic coverage of the Earth are considered. The main feature is the use of several cameras with different swath widths. A vector method is proposed which makes it possible to find orbits minimizing the periodicities of coverage of a given area of Earth uniformly for all swaths. Their number is not limited, but the relative dimensions should satisfy the Fibonacci series or some new numerical sequences. The results apply to constellations of any number of satellites. Formulas were derived for calculating their structure, i.e., relative position in the constellation. Examples of orbits and the structure of constellations for the Earth’s multiswath coverage are presented. 相似文献
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This paper explores the possibility of developing a new attitude control method for satellites in elliptic, 24-hour orbits, in order to compensate for the effect of longitudinal periodic drift relative to the ground station. A simple solar attitude control technique has been proposed for achieving the fixed apparent satellite orientation with respect to the ground segment of the space mission. The proposed control approach appears to be quite attractive for various satellite applications as it can substantially overcome the problems created by the continual periodic angular drift as well as undesirable pitching excitation in the elliptic orbits. Generalizing the analytically developed open-loop control policy results in a significant improvement of the controller performance. 相似文献
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B. B. Kreisman 《Cosmic Research》2011,49(4):325-334
In the context of the restricted circular three-body problem a method for constructing families of periodic orbits is described.
Each orbit contains a segment of transfer from artificial satellite orbit of a smaller body to an orbit around L
1 or L
2 points of the Sun-Earth and Earth-Moon systems, a segment of multiple flyby of this libration point, and a segment of return
to the artificial satellite orbit. Dependences of velocities at the pericenter on the pericenter radius are given. 相似文献
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The application of dynamical systems techniques to mission design has demonstrated that employing invariant manifolds and resonant flybys enables previously unknown trajectory options and potentially reduces the ΔV requirements. In this investigation, planar and three-dimensional resonant orbits are analyzed and cataloged in the Earth–Moon system and the associated invariant manifold structures are computed and visualized with the aid of higher-dimensional Poincaré maps. The relationship between the manifold trajectories associated with multiple resonant orbits is explored through the maps with the objective of constructing resonant transfer arcs. As a result, planar and three-dimensional homoclinic- and heteroclinic-type trajectories between unstable periodic resonant orbits are identified in the Earth–Moon system. To further illustrate the applicability of 2D and 3D resonant orbits in preliminary trajectory design, planar transfers to the vicinity of L5 and an out-of-plane transfer to a 3D periodic orbit, one that tours the entire Earth–Moon system, are constructed. The design process exploits the invariant manifolds associated with orbits in resonance with the Moon as transfer mechanisms. 相似文献