基于相对轨道要素的航天器相对运动碰撞分析

采用基于相对轨道要素的方法,建立了计算碰撺概率的数学模型.在二体圆轨道条件下,推导验证了与基于Clohessy-wiltshire(C-W)方程的碰撞模型的等价关系,并解决了C-W方程存在的若干问题.在航天器近距离相对运动构型设计中提出了基于相对轨道要素的碰撞分析方法,突出了相对轨道要素在构型设计以及碰撞分析中的作用....     

基于中继测量的环月探测器定轨能力分析

嫦娥四号月球探测拟首次实现月球背面的软着陆,测控与数传依赖地月L2平动点的中继卫星,并有望获取四程测量与星间测量数据。对基于中继测量的环月探测器测定轨能力进行了仿真分析,结果表明,中继卫星可较好地实现环月探测器连续跟踪;在定轨能力方面,中继卫星自身轨道精度是制约环月探测器定轨精度的重要因素,当跟踪弧段达到5h以上时,定轨精度趋于稳定,但轨道精度较中继卫星的轨道精度相差1个量级;对于星间链路测量,除中继卫星自身的轨道精度外,星钟的稳定性是制约定轨精度的另一个重要因素,如果辅助以每天1h的地基跟踪亦可实现优于百m的定轨精度。     

Analytical long-term evolution and perturbation compensation models for BeiDou MEO satellites

The current paper establishes the analytical models of the long-term evolution and perturbation compensation strategy for Medium Earth Orbits(MEO)shallow-resonant navigation constellation,with application to the Chinese Bei Dou Navigation Satellite System(BDS).The long-term perturbation model for the relative motion is developed based on the Hamiltonian model,and the long-term evolution law is analyzed.The relationship between the control boundary of the constellation and the offset of the orbital elements is analyzed,and a general analytical method for calculating the offset of the orbit elements is proposed.The analytical model is further improved when the luni-solar perturbations are included.The long-term evolutions of the BDS MEO constellation within 10 years are illustrated,and the effectiveness of the proposed analytical perturbation compensation calculation approach is compared with the traditional numerical results.We found the fundamental reason for the nonlinear variations of the relative longitude of ascending node and the mean argument of latitude is the long-periodic variations of the orbital inclination due to the luni-solar perturbations.The proposed analytical approach can avoid the numerical iterations,and reveal the essential relationship between the orbital element offsets and the secular drifts of the constellation configuration.Moreover,there is no need for maintaining the BDS MEO constellation within 10 years while using the perturbation compensation method.     

Real-Time Shipboard Orbit Determination Using Kalman Filtering Techniques

The real-time tracking and orbit determination program used on board the NASA tracking ship, the USNS Vanguard, is described in this paper. The computer program uses a variety of filtering algorithms, including an extended Kalman filter, to derive real-time orbit determinations (position-velocity state vectors) from shipboard tracking and navigation data. Results from Appolo missions are given to show that orbital parameters can be estimated quickly and accurately using these methods.     

低轨卫星对高轨卫星仅测角初轨计算方法

针对我国地面测站对高轨卫星监视能力缺乏的问题,提出一种低轨卫星对高轨卫星仅测角初轨计算方法。该算法引入天基跟踪坐标系,消除测距信息影响,建立仅测角观测方程;引入法向运动,增加摄动因素影响,建立扩展拉普拉斯动力学模型;推导分析观测模型和动力学模型的关系方程,将初轨计算问题转换为非线性方程求解问题,利用高斯全主元消去法完成方程求解。通过实战和仿真测角数据对方法进行检验,结果表明,该方法能利用仅测角数据对非合作目标进行初轨确定,精度在公里量级,可为我国地基监视系统提供补充参考。     

Optical Calibration of Geostationary Satellite Tracking Systems

A precise calibration method for range and angle observation has been developed for eliminating the systematic error of tracking systems, thus improving the accuracy of orbit determination for geostationary satellites. The principle of calibration is based on an orbit determination employing a point of optical angle observation in addition to radio tracking observation, in which we estimate observation bias parameters simultaneously with orbital elements, including the effects of geodetic mismodelings. As shown by an actual calibration experiment in our ground station, orbit determinations is sufficiently accurate that the error of predicting satellite range falls within a few meters at four days after the day of orbit determination.     

卫星定轨综述

对卫星定轨中的热点问题作一综述,内容包括短弧定轨和长弧精密定轨。对于短弧定轨(通常指初轨),不仅从原理上分析其关键之处,还介绍如何将适用于无摄运动的特殊定轨方法推广用于同时考虑各种摄动的受摄运动,并可同时适用于圆锥曲线的各种情况：变化的椭圆轨道、抛物线轨道和双曲线轨道,后者在深空探测器的运动过程中将会遇到。对于长弧精密定轨,阐述当今提高精密定轨精度的力模型补偿问题以及几何定轨的提法,天地基网联合定轨和星-星相对测量的自主定轨问题。     

影响环月飞行器定轨精度的误差源分析

以我国正在实施的探月计划“嫦娥1号”工程为背案,在现有测控网分布、观测弧段以及尽可能接近真实情况的误差源等前提下,利用仿真模拟的方法对影响环月飞行器定轨精度的误差源进行了初步探讨和分析。重点考察了月球重力场误差、观测量精度、初始时刻的先验轨道误差以及观测资料类型等对环月飞行器定轨精度的影响。     

基于地月三角平动点的卫星自主定轨

近年提出了利用地月系平动点建立深空导航星座的设想。在受太阳摄动的真实力模型下,地月系平动点是不稳定的,从而会导致导航星座必须通过控制才能定点在特定区域。针对此问题,引入一种特殊的平动点轨道,即动力学替代轨道。平动点轨道卫星星座可利用星间测距数据自主定轨,由于动力学替代轨道具有长期稳定性,整个自主定轨过程不需要来自地面的测控支持,且定轨精度可达到观测精度。研究结果表明,观测资料的长短、导航卫星垂直白道面的运动分量都将影响到导航星间的自主定轨精度。该研究成果可以应用在以后的地月系导航星座中。     

Dynamical behaviors and relative trajectories of the spacecraft formation flying

To describe the relative motion of spacecraft formation flying, this paper presents a method based on relative orbital elements, which is suitable to elliptical orbit with arbitrary eccentricity. The long time formation flying conditions are theoretically derived taken into account the relationship between relative motion and relative orbital elements. These conditions include that both the orbital periods of all participating spacecrafts should be the same and other relative orbital elements should be small enough. The expected relative distance of the spacecrafts would determine the magnitudes of such relative orbital elements. Theoretical analysis and numerical simulation results show that the spacecrafts with sufficient small relative orbital elements can keep long time formation flying without any active control when the orbital perturbations are not considered. The results also show that Hill's equation is only suitable for describing a short time formation flying.     

Spacecraft orbit determination using GPS navigation solutions

The orbit determination using the GPS navigation solutions for the KOMPSAT-1 spacecraft has been studied. The Cowell method of special perturbation theories was employed to develop a precision orbit propagation, and the perturbations due to geopotential, the gravity of the Sun and the Moon, solid Earth tides, ocean tides, the Earth's dynamic polar motion, solar radiation pressure, and atmospheric drag were modeled. Specifically, the satellite box-wing macro model was applied to minimize the drag errors at low altitude. The estimation scheme consisted of an extended Kalman filter and Bayesian least square method. To investigate the applicability of the method to the KOMPSAT-1 spacecraft, the orbit determination was accomplished using the GPS navigation solutions for the TOPEX/POSEIDON and TAOS satellites. The orbit determination results were compared with NASA POE generated by global laser tracking. The position and velocity accuracy was estimated about 16∼7 m and 0.0157∼0.0074 m·s⁻¹ RMS, respectively, for the two satellites in the presence of SA. These results verify that an orbit determination scheme using GPS navigation solutions can provide the static orbit information and reduce conspicuously the position and velocity errors of navigation solutions. It can be suggested that the sequential and batch orbit determination using the GPS navigation solutions be the most appropriate method in the KOMPSAT-1 type mission.     

基于加权最小二乘法的椭圆轨道编队飞行设计

提出了一种不依赖于相对运动描述的偏心率编队飞行设计方法。利用加权最小二乘法，根据事先所确定的理想相对运动轨迹，就能设计出与之最接近的相对运动轨道。这种方法适用于任意偏心率，并能很容易地加入各种摄动力的影响。仿真验证了其有效性。     

A geopause satellite system concept

The forthcoming 10 cm range tracking accuracy capability holds much promise in connection with a number of Earth and ocean dynamics investigations. These include a set of earthquake-related studies of fault motions and the Earth's tidal, polar and rotational motions, as well as studies of the gravity field and the sea surface topography which should furnish basic information about mass and heat flow in the oceans. The state of the orbit analysis art is presently at about the 10 m level, or about two orders of magnitude away from the 10 cm range accuracy capability expected in the next couple of years or so. The realization of a 10 cm orbit analysis capability awaits the solution of four kinds of problems, namely, those involving orbit determination and the lack of sufficient knowledge of tracking system biases, the gravity field, and tracking station locations. The Geopause satellite system concept offers promising approaches in connection with all of these areas. A typical Geopause satellite orbit has a 14 hour period, a mean height of about 4.6 Earth radii, and is nearly circular, polar, and normal to the ecliptic. At this height only a relatively few gravity terms have uncertainties corresponding to orbital perturbations above the decimeter level. The orbit s, in this sense, at the geopotential boundary, i.e., the geopause. The few remaining environmental quantities which may be significant can be determined by means of orbit analyses and accelerometers. The Geopause satellite system also provides the tracking geometery and coverage needed for determining the orbit, the tracking system biases and the station locations. Studies indicate that the Geopause satellite, tracked with a 2 cm ranging system from nine NASA affiliated sites, can yield decimeter station location accuracies. Five or more fundamental stations well distributed in longitude can view Geopause over the North Pole. This means not only that redundant data are available for determining tracking system biases, but also that both components of the polar motion can be observed frequently. When tracking Geopause, the NASA sites become a two-hemisphere configuration which is ideal for a number of Earth physics applications such as the observation of the polar motion with a time resolution of a fraction of a day. Geopause also provides the basic capability for satellite-to-satellite tracking of drag-free satellites for mapping the gravity field and altimeter satellites for surveying the sea surface topography. Geopause tracking a coplanar, drag-free satellite for two months to 0.03 mm per second accuracy can yield the geoid over the entire Earth to decimeter accuracy with 2.5° spatial resolution. Two Geopause satellites tracking a coplanar altimeter satellite can then yield ocean surface heights above the geoid with 7° spatial resolution every two weeks. These data will furnish basic boundary condition information about mass and heat flows in the oceans which are important in shaping weather and climate.     

基于轨道方程线性化的中段飞行轨道修正方法

对于深空探测、大范围远程交会、抵近观察等中段飞行时间较长的航天任务,中段轨道修正是完成任务的重要保证.基于摄动制导的思路,通过对轨道方程进行线性化,研究了二体条件下的中段轨道修正方法.在近地航天任务中,各种摄动力特别是J2项对终端位置精度的影响不可忽略,因此研究了考虑J2摄动影响时线性化方程的补偿方法.最后通过数值仿真...     

绕飞慢自旋小天体的航天器运动分析

针对可以用三轴椭球体近似建模的小天体,给出了非球形引力势函数,建立了航天器绕飞小天体的轨道动力学方程。利用Jacobi积分常数绘制了探测器在小天体周围的零速度曲线,并分析了探测器的可能运动区域,给出了航天器不碰撞小天体的边界条件。针对绕飞慢自旋小天体的情况,基于平均轨道根数的近似解分析了小天体扁率和椭率的摄动影响,并给出了几条冻结轨道及其稳定条件。     

摄动对编队飞行星座相对构型的影响分析

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，分别研究了大气阻力摄动、日月引力摄动、太阳光压摄动和地球扁率摄动对编队飞行星座的构型影响，着重分析了地球扁率摄动周期项和长期项对构型的影响，并以此对编队轨道设计提出建议。     

直接解算轨道根数的轨道约束自校准定轨方法

应用轨道约束自校准技术能够有效地修正系统误差和提高定轨精度。本文为了适应各种航天试验任务要求,将轨道约束自校准技术推广应用到直接解算轨道根数的定轨方法中,并导出了解算开普勒根数的轨道约束自校准公式。     

Controlling Stationary Satellite Orbits Minimum Fuel Consumption Rate

The orbital angular velocity of a stationary satellite expresses the perturbation in the orbit of the satellite. The minimum variation in direction of this velocity corresponds to the minimum fuel consumption rate to maintain a stationary satellite within allocated bounds. The directional variation of the orbital angular velocity is minimized by maintaining the ascending node of the orbit near the direction of the vernal equinox. The direction of the ascending node to maintain the orbit with minimum fuel consumption rate is given over the 18.6 yr nodal period of the moon. Over that period the inclination variation of the orbit and the angular speed proportional to the necessary fuel consumption rate to maintain the orbit are also given. An example of fuel consumption is given with a comparison with fuel savings over the standard stationkeeping method. The method here is applicable to the geostationary communication satellite, UHF broadcasting satellite, solar power satellite, etc.     

三体系统中引力拖车的偏置非开普勒轨道研究

为了研究太阳-小行星-引力拖车三体系统中引力拖车的轨道运动问题,采用柱坐标系下的Hill方程描述了三体系统中引力拖车的运动情况,应用平均化方法消除周期项的影响,得到平均偏置非开普勒轨道的表达式,并研究了轨道稳定性与引力拖车最大有效拉力等问题.研究表明:三体系统中,在小行星飞行方向（或反方向）上存在偏置非开普勒轨道;与二...     

嫦娥二号再拓展试验测定轨精度研究

基于＂嫦娥二号＂卫星再拓展试验的设计轨道,研究各种摄动力对轨道确定精度的影响,得出的结论是：若要达到km量级的轨道确定精度,必须考虑除天王星和海王星之外所有大行星以及日月的质点引力。文章进一步利用数值分析法研究再拓展任务的轨道确定精度,分析结果表明：基于目前的测控条件,使用30 d以上的测轨弧段可以得到稳定可靠的轨道解,而短弧（小于20 d）稳定轨道的获取需要VLBI（甚长基线干涉）测轨数据支持;当＂嫦娥二号＂距离地球700万km时,测控精度可优于30 km;虽然每天测轨弧段的增加可以改善轨道精度,但是当增加到8 h以上时,定轨精度将不再有明显改善。