带电粒子云的演变过程

根据带电粒子云从破碎点开始向空间扩散过程中粒子云密度和形状的变化规律,以几何形状和起主要作用的因素为特征,定义了球形、椭球形、绳形、螺旋线形、全方位弥漫直至球壳形6个演变阶段.论述了在各个阶段的主要特征和对演变过程起主要作用的因素.分析了在各个阶段电磁场对带电粒子的摄动影响,比较了带电粒子云与不带电粒子云在演化过程上的差异.在球形阶段起主要作用的是分离速度,带电碎片之间的排斥力加快了碎片扩散的速度.从椭球形阶段到球壳形阶段,带电粒子和不带电粒子的演化规律基本一致.带电粒子的演化过程加快或减慢取决于粒子带正电或带负电.将电场摄动力等效于改变地球引力的大小,罗列了阶段转换标志点时刻的计算公式.利用计算机仿真的方法,给出了各个阶段不带电碎片云和带电碎片云分布示意图,验证了演变过程阶段划分和电磁场摄动分析的合理性.      

考虑J_2摄动作用的异面椭圆轨道双冲量交会轨道修正方法

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.     

MDO approach for early design of aerobraking orbital transfer vehicles

This paper presents a new multidisciplinary design optimization (MDO) methodology for preliminary design of an aeroassisted orbital transfer vehicle (AOTV) performing a two-way transfer between a low-Earth “parking” orbit and a high-energy orbit. This work has been performed in the frame of Onera's CENTOR [N. Bérend, C. Jolly, F. Jouhaud, D. Lazaro, Y. Mauriot, C. Monjaret, J.M. Moschetta, M. Parlier, J.L. Pastre, Y. Servouze, J.L. Vérant, Project CENTOR: Preparing the design of future orbital transfer vehicles; IAC-07-D.2.3.07, in: 58th International Astronautical Congress, 24–28/09/2007, Hyderabad, India] project whose objective is to prepare tools and methodology for studying and designing future space transportation systems for new kinds of missions such as on-orbit servicing (OOS), payload ferrying, or in-situ observation of space-debris. Using simplified models and an appropriate low-dimension formulation for the optimization problem the method makes possible to obtain rapidly and easily a global view of the trade-off between the payload mass and the total mass. It also makes possible to discuss the feasibility of the vehicle with regard to different multidisciplinary constraints and technology hypotheses for the heat shield. This approach is illustrated by eight different AOTV design studies, considering two different missions (LEO–MEO and LEO–GEO), two different propulsion technologies (LOX-LH2 and LOX-CH4) and two different thermal protection system (TPS) characteristics. In each case, we discuss the feasibility and characteristics of the lightest vehicle carrying a prescribed 100 kg payload, and, conversely, a heavy vehicle with a prescribed 18 ton total mass, carrying the heaviest possible payload.     

光电干扰箔片云团运动规律的仿真研究

鉴于自燃箔片云团对红外成像导引头光电干扰有效的试验结果,针对气动阻力大的箔片云团,采取常用的气动阻力公式f=-kv2来描述运动规律,引入匀降速度u,分别推导出水平方向和垂直方向上升段和下降段的速度、位移的计算公式.结合图象分析对云团的运动规律进行仿真研究,还讨论了采用不同气动阻力公式对拟合系数和拟合相似度的影响.该方法可推广应用于大气层内亚声速质点运动轨迹的解析方法.     

On the effects of each term of the geopotential perturbation along the time I: Quasi-circular orbits

This paper provides a useful new method to determine minimum and maximum range of values for the degree and order of the geopotential coefficients required for simulations of orbits of satellites around the Earth. The method consists in a time integration of the perturbing acceleration coming from each harmonic of the geopotential during a time interval T. More precisely, this integral represents the total velocity contribution of a specific harmonic during the period T  . Therefore, for a pre-fixed minimum contribution, for instance 1×10^-8$1\times {10}^{-8}$ m/s during the period of time T, any harmonic whose contribution is below this value can, safely, be neglected. This fact includes some constraints in the degree and order of the terms which are present in the geopotential formula, saving computational efforts compared to the integration of the full model. The advantage of this method is the consideration of other perturbations in the dynamics (we consider the perturbations of the Sun, the Moon, and the direct solar radiation pressure with eclipses), since these forces affect the value of the perturbation of the geopotential, because these perturbations depend on the trajectory of the spacecraft, that is dependent on the dynamical model used. In this paper, we work with quasi-circular orbits and we present several simulations showing the bounds for the maximum degree and order (M) that should be used in the geopotential for different situations, e. g., for a satellite near 500 km of altitude (like the GRACE satellites at the beginning of their mission) we found 35?M?198$35?M?198$ for T=1$T=1$ day. We analyzed the individual contribution of the second order harmonic (_J2$J_2$) and we use its behavior as a parameter to determine the lower limit of the number of terms of the geopotential model. In order to test the accuracy of our truncated model, we calculate the mean squared error between this truncated model and the “full” model, using the CBERS (China-Brazil Earth Resources Satellite) satellite in this test.     

非线性相对运动方程的分析解

阐述追踪航天器相对目标航天器的运动方程的建立,以线性相对运动方程分析解为基础,推导出非线性相对运动方程分析解,这里目标航天器沿一般椭圆轨道（包括圆）运动,追踪航天顺受常值推力作用或在自由状态下运动。     

飞行器轨道交会的H_∞次优控制问题

研究了近圆轨道上两个飞行器轨道交会的控制问题。通过对最优控制理论设计最优交会轨线的研究 ,提出了利用 H∞控制理论设计实际控制器的方法 ,给出了仿真结果 ,验证了这种方法的正确性。说明了该方法具有鲁棒的控制效果以及次优的轨道特性。     

GNSS广域差分改正数估计方法研究

随着我国北斗三号基本系统的正式运行,基于地面监测站的广域差分增强系统成为进一步提升卫星导航定位精度的手段之一。在码噪声多径误差修正(CNMC)的基础上,使用等效钟差方法实现GNSS卫星轨道与钟差误差的解耦,并依据卫星轨道运动的动力学特性,引入希尔差分方程描述卫星轨道误差变化,实现对轨道误差的实时卡尔曼滤波估计。基于GPS实测数据,对改正前后的用户等效测距误差(UERE)和定位精度进行了对比研究。实验结果表明,采用该方法,UERE标准差由改正前的0.456 m减小至0.227 m,降幅达到50.22%;整体水平定位误差(95%置信区间)由0.981 m减小至0.782 m,垂直定位误差(95%置信区间)由1.991 m减小至1.131 m,分别提升了20.29%和43.19%,差分改正效果明显。     

一种顾及轨道误差的实时GPS钟差估计方法

为了提高实时卫星钟差估计的精度和稳定性,提出了一种顾及轨道误差的实时GPS钟差估计方法。基于超快速轨道产品,分析了轨道标准差与绝对轨道误差的相关性。通过线性插值获得实时轨道误差信息,以优化先验残差的随机模型。基于先验轨道标准差阈值,采用分段方式剔除实时轨道异常卫星对应的观测值。实验结果表明：轨道标准差和轨道误差的相关性高达0.82。与常用的高度角相关的随机模型相比,GPS卫星钟差估计精度最大提高了15.2%,平均提高了8.1%,钟差误差的时间序列更加平稳,所有GPS卫星的平均钟差STD均在0.15ns以内。因此,超快速轨道产品中提供的轨道标准差与绝对轨道误差表现出较强相关。采用顾及轨道误差的实时钟差估计方法可提高GPS卫星钟差估计精度,准确识别并剔除GPS实时轨道异常的卫星,从而提高GPS实时钟差估计的稳定性。     

地球同步卫星的发射轨道的选择

本文对地球同步卫星的发射轨道作了一些分析。讨论了有关发射点纬度、停泊轨道、非共面发射轨道对发射轨道选择的影响。