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地球磁场对带电赤道卫星的轨道半长轴和偏心率的摄动影响 总被引:3,自引:0,他引:3
本文利用摄动理论研究了带电的赤道卫星在地球磁场中做椭圆轨道运动时地磁摄动力对轨道半长径和偏心率的摄动影响。研究结果表明:地磁场对带电赤道卫星的轨道半长径没有摄动影响,既无周期摄动,也无长期摄动,但对轨道偏心率有摄动影响,且只有周期性摄动,而无长期摄动。当卫星自身带电量较大时,这种摄动影响必须予以考虑 相似文献
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为了分析大气阻力摄动对卫星编队队形的影响,利用相对轨道根数法推导了包含大气阻力摄动的卫星编队相对运动的状态转移方程。仿真结果表明当几何形状和质量不同的两颗卫星在低轨道做编队飞行时,大气阻力摄动对编队队形的影响很大而不能忽略;当主卫星的半长轴相等时,主卫星轨道的偏心率越大编队飞行受大气阻力摄动的影响也越大;大气阻力摄动主要影响编队飞行迹向相对距离。 相似文献
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运用解析和数值计算两种方法,分析了各种摄动源对地球静止轨道(GEO)卫星弃置轨道近地点高度变化的影响,得到了近地点高度的变化规律。利用二阶日月引力摄动可造成GEO卫星弃置轨道近地点升高的特点,提出新的卫星离轨策略,在不满足机构间空间碎片协调委员会(IADC)"空间碎片减缓指南"中,GEO卫星弃置轨道偏心率小于0.003要求的情况下,还可以保证卫星不再进入GEO保护区域。 相似文献
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在近圆轨道编队飞行的假设条件下 ,根据动力学关系推导出了环绕卫星相对参考卫星的运动学简化模型 ,并以此简化模型为基础 ,研究了半长轴入轨偏差δa ,轨道倾角偏差δi ,轨道偏心率偏差δe在地球扁率摄动条件下对编队飞行星座相对构型稳定性影响 ,深入分析了各种因素的规律和特点 ,并以此对编队飞行星座轨道设计提出建议。 相似文献
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椭圆轨道编队飞行的典型模态与构型保持控制方法 总被引:6,自引:2,他引:6
基于T H方程推导了椭圆参考轨道编队飞行的周期性条件,讨论了椭圆参考轨道卫星编队飞行的典型模态。该模态是圆形参考轨道空间圆形模态的推广。论述了二阶带谐项(J2项)摄动对编队飞行构型保持的影响,基于相平面法提出了一种编队飞行构型保持控制方法。该控制方法不是消极的抵消干扰的影响,而是积极的利用干扰的作用达到节约燃料并精确保持构型的目的。仿真表明,采用该控制方法可对椭圆参考轨道卫星编队构型进行有效的保持。 相似文献
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用完全不同的方法推导出一种两颗卫星相对运动的表达式.首先,它描述的是一颗真实的卫星相对一颗在理想轨道上运动的虚拟卫星的运动,而这正是研究编队飞行所必须的;其次,这个表达式将真实卫星相对虚拟卫星的三维位置表示成虚拟卫星的Brouwer平根数及两卫星轨道根数之差的函数.据此便可以充分应用已有的关于轨道摄动的研究成果.作为一个基本模型,首先讨论了一颗真实卫星相对虚拟卫星形成椭圆型地面轨迹的飞行方式.基于轨道长期摄动的已有结果可以很容易地揭示这种飞行方式的长期变化,从而可以毫无困难地制定出轨道调整的策略.最后还研究了三颗卫星以同一条椭圆轨迹飞行的轨道设计和控制问题,该椭圆以虚拟卫星为中心. 相似文献
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两圆轨道之间的双共切转移轨道是其近地点和远地点分别在这两个圆轨道上的椭圆轨道。本文用两次冲量法给出了沿双共切椭圆轨道实现从一圆轨道向另一圆轨道转移的最优方案,并考虑到地球扁率造成的轨道摄动。文中的所谓圆轨道指的是变轨时刻的密切轨道为圆形的轨道,是对近圆轨道的近似替找。 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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E.A. Roth 《Acta Astronautica》1985,12(2):71-80
In this paper the stroboscopic method is applied to the equinoctial elements which avoid the singularities of circular and/or equatorial orbits. The Lagrange equations for the variation of parameters are formulated using respectively one of the three longitudes as fast angular variable. It is shown how the first-order theory of the stroboscopic method can be developed. The perturbation by the gravity potential of the central body and the third-body perturbation are considered in detail. The paper concludes with a few analytical results. 相似文献
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V. I. Prokhorenko 《Cosmic Research》2002,40(1):48-54
The geometric analysis applied to the solutions obtained by M.L. Lidov [1] of the satellite version of a restricted circular double-averaged three-body problem allowed us fresh view at the comparative analysis of evolution periods for various elliptical orbits. Based on the similarity theory, the parameters of similitude are invoked for orbits and perturbations. An expression for the period of evolution of three orbit elements through the appropriate similitude parameters is derived. On this basis, a regularity is formulated, the consequences of which are analyzed. 相似文献
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A new set of relative orbit elements is strictly defined through spherical geometry. The exact transformation equations between the new relative orbit elements and classical-orbital elements are derived. A new relative motion model with no singularity problem is derived based on the relative orbit elements, which are suitable for both elliptical and circular reference orbits. The in-plane and out-of-plane relative motion can be completely decoupled based on the new model. The inverse transformation of state transfer matrix is obtained to analyze perturbation effects and control strategy. The geometric characteristics of relative motion can be easily described using the relative eccentricity/inclination vector method. The proposed method and conclusions are validated by simulation through some typical examples. This paper improves the basic theory of relative orbit elements and unifies the expressions of the elliptical and near-circular close relative motion. 相似文献
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重点研究了地球扁率J2摄动作用对固定时间双冲量异面椭圆轨道交会运动的影响以及相应的轨道修正方法.根据摄动理论,利用Eneke法计算了J2摄动作用引起的实际飞行轨迹与理论标称轨迹之间的位置偏差.在此基础上,提出了通过修正预定的交会位置来修正转移轨道偏差的控制方法.通过仿真计算,对修正前后的位置偏差进行了仿真分析,结果表明,文中提出的轨道修正方法能够有效地减小摄动作用引起的制导误差,且方法简洁、易于实现. 相似文献
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The present paper is concerned with the search for orbits that have potential to require low fuel consumption for station-keeping maneuvers for constellations of satellites. The method used to study this problem is based on the integral over the time of the undesired perturbing forces. This integral measures the change of velocity caused by the perturbation forces acting on the satellite, so mapping orbits that are less perturbed, which generates good candidates for orbits that requires low fuel consumption for station-keeping maneuvers. The integral over the time depends only on the orbit of the spacecraft and the dynamical system considered. The type of engine and the control technique applied to the spacecraft are not considered to search for those orbits. It can be a good strategy to be applied for a first mapping of orbits. For this search, it is analyzed the integral of orbits with different values of the Keplerian elements in order to find the best ones with respect to this criterion. The perturbations considered are the ones caused by the third body, which includes the Sun and the Moon, and the J2 term of the geopotential. The results presented here show numerical simulations to obtain the integral of those perturbing forces for different orbits. The GPS and the Molniya constellations are used as examples for those calculations. 相似文献
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The costate along a coast arc on an optimal space trajectory contains critically important information about the trajectory. For free-time fuel-optimal flight, the costate at the start of the coast determines completely the optimal length of the coast. Yet most closed-form solutions for costate under various coordinate systems available in the literature are only for two-dimensional flight. In this paper complete three-dimensional closed-form costate solutions in flight-path coordinate system are derived for all conic orbits. These results, as an example of their practical usefulness, enable the optimal duration of any non-circular Keplerian coast arc to be accurately determined from the appropriate root of a polynomial of 5th degree in true anomaly, and a 4th degree polynomial for circular orbits. The value of the development in the paper is demonstrated by solving two relatively difficult multi-finite-burn orbital transfer problems. 相似文献