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本文用模拟计算的方法,对人造卫星照相和双频观测资料的定轨精度进行了初步分析。给出了在测站坐标误差100米,测时误差0.002秒,位置误差2角秒,测速误差15厘米/秒的条件下,轨道根数中误差的最佳理论值。同时对测站坐标误差和测站分布对定轨精度的影响,也进行了初步的探讨。 相似文献
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越来越多的LEO卫星装载了高精度的星载GPS接收机,星载GPS定轨已成为LEO卫星精密定轨的重要手段之一。星载GPS精密定轨精度依赖于GPS星历及钟差精度,采用CODE(Center for Orbit Determination in Europe)官方网站提供的GPS精密星历及钟差数据,基于瑞士伯尼尔大学开发的Bernese 5.0软件,采用非差减缩动力学定轨方法,解算了60天的CHAMP卫星和SAC-C卫星轨道,并将所得轨道与JPL和GFZ事后科学轨道比较,得出的轨道位置三维精度优于20 cm量级,速度三维精度约为0.20 mm/s。 相似文献
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SGP4/SDP4模型精度分析 总被引：2，自引：0，他引：2
本文基于最新发布的SGP4/SDP4(Simplified General Perturbation Version 4/Simplified Deep-space Perturbation Version 4)模型设计了一套定轨方案,从空间目标库中挑选出不同类型和轨道参数的1120个目标进行计算,定量给出了SGP4/SDP4模型处理不同类型空间目标的定轨预报精度.结果表明:近地目标定轨精度为百米量级;半同步和同步轨道定轨精度平均为0.7和1.9km.椭圆轨道目标的定轨精度与偏心率有关,除少数e>0.8的椭圆轨道目标,绝大多数椭圆轨道目标定轨误差均小于10km.用SGP4/SDP4模型对近地目标预报3天,半同步轨道预报30天,同步轨道预报15天,椭圆轨道预报1天,预报误差一般不超过40km. 相似文献
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根据我国广域差分GPS系统对GPS卫星定轨的要求,通过对国内GPS跟踪网实测数据的处理,分析和讨论了区域网定轨的数据处理方法和可能达到的定轨精度。为了提高所定轨道的稳定性和先进实时预报的精度,通过对所定轨道的误差分析提出在轨道沿迹方向引入经验加速度计算方案。计算结果表明,采用此方法后GPS卫星的定轨精度有了显提高。既论证了利用我国区域GPS网和广播星历进行独立定轨的可行性,也阐述了提高轨道预报精度的方法。 相似文献
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《中国科学院上海天文台年刊》2014,(0)
空间目标包括在轨卫星、空间碎片等,对其测定轨是空间攻防和空间利用的重要前提。由于地面测站资源有限,单站测量是目前对空间目标尤其是空间碎片测定轨较常用的方式。卫星激光测距(satellite laser ranging,SLR)技术测量精度很高,可达米级(非合作目标),甚至厘米级(合作目标),但不能单独用于单站短弧定轨;电荷耦合器件(charge coupled device,CCD)天文定位技术可观测距离较远的目标,但测量精度为角秒级,换算至空间距离不如SLR技术高。两者的联合为空间目标的高精度定位和跟踪提供了可能,并成为未来空间目标地基测量的发展方向。作为空间碎片单站监测的前期工作,对合作目标的单站定轨精度进行了评估。处理了1500 km高AJISAI低轨卫星的实测数据,分析了单站CCD测角和激光测距数据对低轨空间目标的联合定轨能力,并充分考虑两类不同类型观测数据的精度,数据综合时对其进行合理加权。利用全球激光站资料进行精密定轨,并以此作为参考轨道,采用上海佘山站AJISAI卫星2010年、2011年4天6圈的实测激光测距数据,以及CCD测角数据,开展了单站单圈和单站多圈定轨和预报试验。试验结果表明,测距数据的加入对定轨精度和24小时预报精度的改善非常明显,可提高至少一个数量级;单站单圈联合定轨和24小时预报的精度分别为20 m以内及数百米,单站多圈联合定轨和24小时预报的精度分别在米级及数十米。期望实验结果为中国未来的空间碎片望远镜建设提供参考。 相似文献
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采用重置参数的轨道改进算法 总被引：1，自引：1，他引：0
《天文研究与技术》2010,(4)
当使用精度差的初始根数作定轨计算时,被估值的模型参数会吸收初值中所含误差而偏离其合理数值(如CD约为2.2),使定轨计算过程的RMS已不再变化,但轨道收敛到与实际状态有偏离的轨道上。文中给出的算例采用重置被歪曲的估值模型参数方法,首先以TLE根数为初值用精密定轨程序解条件方程,然后以第一轮迭代计算结果作为初始根数并重置模型参数,再进行第二轮迭代计算,使定轨计算结果收敛到正确轨道上,文中还使用另一颗激光卫星的双行根数作初值验证了该方法的有效性。较好地解决了因初值不准所引起的定轨计算不收敛,或收敛到与实际状态有偏离的轨道上的问题。最终得出的RMS达到厘米级精度。文中图示了两次定轨计算的RMS变化曲线图、残差分布图,迭代过程的资料采用率及定轨计算结果。 相似文献
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《天文学报》2015,(5)
地球引力场模型是人造卫星轨道计算中最重要的动力学模型之一.近年来国际上空间重力卫星计划取得了极大成功,相继推出了一系列新的引力场模型.从近地卫星轨道计算的角度检验了2种传统引力场模型(JGM3,EGM96)和4种新引力场模型(EIGENCHAMP05S,GGM03S,GOCE02S,EGM2008)的精度,利用4颗近地卫星的激光测距资料进行精密定轨预报,统计比较了不同模型的定轨残差和预报误差.结果表明:(1)4种新引力场模型精度基本在同一水平,对于近地卫星定轨精度普遍优于9 cm,最高达到5cm,相对于JGM3和EGM96模型有明显改善;(2)以JGM3模型为基准,EGM96模型的精度有所提高,2000年以后的4种新模型的精度则普遍提高了12%～47%(定轨)和63%(预报).70阶之前定轨精度随着模型阶次增大而提高,70阶以后定轨精度基本保持稳定,这表明对于近地卫星轨道计算而言,70阶的引力场已经能够满足厘米级的精度需求. 相似文献
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利用VLBI数据确定"探测一号"卫星的轨道 总被引：5，自引：0，他引：5
双星计划的“探测一号”是中国首颗真正严格意义上的科学实验卫星,其运行轨道为中国迄今所发射的卫星中距地球最远,远地点地心距达7．8万公里．采用射电天文的VLBI技术可以对“探测一号”以及更远的深空目标,如探月飞行器实现跟踪．为了验证VLBI技术在我国探月计划中的作用,上海天文台组织了国内目前仅有的上海、乌鲁木齐和昆明3个台站对“探测一号”进行试跟踪,利用对“探测一号”约两天的VLBI观测数据,确定“探测一号”卫星的轨道,对VLBI的定轨能力做初步的探讨．按照测控部门提供的初轨 (其精度仅保证跟踪)推算的轨道与VLBI时延的拟合误差平均约2 km,时延率的拟合误差平均约15 cm／s．而利用VLBI数据定轨后的拟合程度相对于初轨有了很大的改善,结果表明,单独利用VLBI时延定轨,时延的拟合精度约5．5 m,作为外部检核的VLBI时延率的拟合精度在2 cm／s左右．单独利用VLBI时延率定轨,时延率的拟合精度约为1．3 cm／s,作为外部检核的VLBI时延的拟合精度约为29 m．而若将时延和时延率数据联合定轨,采用其内符精度加权,VLBI时延和时延率的残差分别为5．5 m和 2 cm／s．为了合理地评估VLBI定轨的真实精度,利用模拟数据进行误差协方差分析,结果表明VLBI定轨精度受动力学模型误差的影响较大,由于"探测一号”卫星的动力学模型难以精确确定,所以利用两天弧段的VLBI数据确定“探测一号”卫星轨道的位置误差为km量级,而速度误差可达cm／s量级．模拟计算还表明, VLBI和USB数据联合定轨可以大大提高定轨精度． 相似文献
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SONG Ye-zhi HUANG Yong HU Xiao-gong LI Pei-jia CAO Jian-feng 《Chinese Astronomy and Astrophysics》2014
It is known that the dynamical orbit determination is the most common way to get the precise orbits of spacecraft. However, it is hard to build up the precise dynamical model of spacecraft sometimes. In order to solve this problem, the technique of the orbit determination with the B-spline approximation method based on the theory of function approximation is presented in this article. In order to verify the effectiveness of this method, simulative orbit determinations in the cases of LEO (Low Earth Orbit), MEO (Medium Earth Orbit), and HEO (Highly Eccentric Orbit) satellites are performed, and it is shown that this method has a reliable accuracy and stable solution. The approach can be performed in both the conventional celestial coordinate system and the conventional terrestrial coordinate system. The spacecraft's position and velocity can be calculated directly with the B-spline approximation method, it needs not to integrate the dynamical equations, nor to calculate the state transfer matrix, thus the burden of calculations in the orbit determination is reduced substantially relative to the dynamical orbit determination method. The technique not only has a certain theoretical significance, but also can serve as a conventional algorithm in the spacecraft orbit determination. 相似文献
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讨论满足约束条件的月球卫星飞行轨道的设计问题，将约束条件分类为只与太阳，月球，地球，飞行器和观测站之间的相对位置有关的运行学约束条件以及涉及到飞行器轨道运行的动力学约束条件，在考虑月球卫星轨道的受力情况后，给出一种准确快速地计算和设计满足约束条件的标准飞行轨道的方法，并应用于不同约束条件下月球卫星的轨道预设计，初步讨论了轨道设计的误差分析，轨道跟踪及实时精密定轨等正在进行的其它相关工作。 相似文献
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Bruno Bertotti 《Celestial Mechanics and Dynamical Astronomy》2001,80(1):21-38
We show that, when a natural satellite like Titan is invisible (e.g., due to an opaque atmosphere) its planetary orbit and its mass can be determined by tracking a spacecraft in close flybys. This is an important problem in the Cassini mission to the Saturnian system, which will be greatly improved by a good astrometric model for all its main components; in particular, an accuracy of a few hundred meters for the orbit of Titan is necessary to allow a measurement of its moment of inertia. The orbit of the spacecraft is the union of elliptical arcs, joined by short hyperbolic transitions: a problem of singular perturbation theory, whose solution leads to a matching condition between the inner hyperbolic orbit and the elliptical orbital elements. Since the inner elements are given in terms of the relative position and velocity of the spacecraft, accurate Doppler measurements in both regions can provide a satisfactory determination of Titan's position and velocity, hence of its Keplerian elements. The errors in this determination are discussed on the basis of the expected Allan deviation of the Doppler method; it is found that the driving errors are those in the elliptical arcs; the fractional errors in Titan's orbital elements are expected to be 10–7. It is also possible to measure the mass of the satellite; however, when the eccentricity e of the flybys is large, the mass and a scaling transformation are highly correlated and the fractional error in the mass is expected to be e times worse. 相似文献
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Orbit Determination Using Satellite—to—Satellite Tracking Data 总被引：3，自引：0，他引：3
Ying－ChunLiu LinLiu 《中国天文和天体物理学报》2001,1(3):281-286
1 INTRODUCTIONThe tracking arc-length should be increased in order to approve the accuracy in orbitdetermination of LEO (Low Earth Orbit) satellites. The local ground-based tracking networkdoes not provide sufficient orbit coverage for the user satellites. The most promising methodis to use high orbiting satellites, such as GPS and TDRS, as trackers to observe the usersatellites. For examPle, tWO geosynclironous satellites could cover more than 85% of the orbitof any given user sate… 相似文献
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The paper analyzes the possibility for countering ballistic perturbations of the interplanetary transfer trajectory of the spacecraft with electric propulsion (EP) associated with the temporary impossibility of the normal use of the EP in phases of the heliocentric transfer. The main result of the present study is the method for the determination of a new nominal trajectory, at any point of which the allowed duration of the emergency shutdown of electric propulsion is large enough. The numerical analysis is given for one of the possible scenarios of spacecraft injection into the operational heliocentric orbit for solar research. 相似文献
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A. I. Nazarenko 《Solar System Research》2013,47(7):564-568
The article covers the outcome of the method for optimal filtering of measurements developed by the author and aimed at the determination of the time and place of reentry of the Phobos-Grunt spacecraft. So-called two-line elements (TLE) of the orbit of the American Space Surveillance System are used as measurements. 相似文献
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M. Vetrisano W. Van der Weg M. Vasile 《Celestial Mechanics and Dynamical Astronomy》2012,114(1-2):25-53
This paper presents a navigation strategy to fly to the Moon along a Weak Stability Boundary transfer trajectory. A particular strategy is devised to ensure capture into an uncontrolled relatively stable orbit at the Moon. Both uncertainty in the orbit determination process and in the control of the thrust vector are included in the navigation analysis. The orbit determination process is based on the definition of an optimal filtering technique that is able to meet accuracy requirements at an acceptable computational cost. Three sequential filtering techniques are analysed: an extended Kalman filter, an unscented Kalman filter and a Kalman filter based on high order expansions. The analysis shows that only the unscented Kalman filter meets the accuracy requirements at an acceptable computational cost. This paper demonstrates lunar weak capture for all trajectories within a capture corridor defined by all the trajectories in the neighbourhood of the nominal one, in state space. A minimum Δv strategy is presented to extend the lifetime of the spacecraft around the Moon. The orbit determination and navigation strategies are applied to the case of the European Student Moon Orbiter. 相似文献
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Trent Jansen‐Sturgeon Eleanor K. Sansom Philip A. Bland 《Meteoritics & planetary science》2019,54(9):2149-2162
Fireball networks establish the trajectories of meteoritic material passing through Earth's atmosphere, from which they can derive pre‐entry orbits. Triangulated atmospheric trajectory data require different orbit determination methods to those applied to observational data beyond the Earth's sphere of influence, such as telescopic observations of asteroids. Currently, the vast majority of fireball networks determine and publish orbital data using an analytical approach, with little flexibility to include orbital perturbations. Here, we present a novel numerical technique for determining meteoroid orbits from fireball network data and compare it to previously established methods. The re‐entry of the Hayabusa spacecraft, with its known pre‐Earth orbit, provides a unique opportunity to perform this comparison as it was observed by fireball network cameras. As initial sightings of the Hayabusa spacecraft and capsule were made at different altitudes, we are able to quantify the atmosphere's influence on the determined pre‐Earth orbit. Considering these trajectories independently, we found the orbits determined by the novel numerical approach to align closer to JAXA's telemetry in both cases. Using simulations, we determine the atmospheric perturbation to become significant at ~90 km—higher than the first observations of typical meteorite dropping events. Using further simulations, we find the most substantial differences between techniques to occur at both low entry velocities and Moon passing trajectories. These regions of comparative divergence demonstrate the need for perturbation inclusion within the chosen orbit determination algorithm. 相似文献
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The scientific objectives of a geodetic experiment based on a network of landers, such as NEIGE (NEtlander Ionosphere and Geodesy Experiment) are to improve the current knowledge of Mars' interior and atmosphere dynamics. Such a network science experiment allows monitoring the motions of the Martian rotation axis with a precision of a few centimeters (or milli-arc-seconds (mas)) over annual and sub-annual periods. Thereto, besides radio tracking of a Mars orbiter from the Earth, radio Doppler shifts between this orbiter and several landers at the planet's surface will be performed. From the analysis of these radio Doppler data, it is possible to reconstruct the orbiter motion and Mars' orientation in space. The errors on the orbit determination (position and velocity of the orbiter) have an impact on the geodetic parameters determination from the Doppler shifts and must be removed from the signal in order to achieve a high enough accuracy. In this paper, we perform numerical simulations of the two Doppler signals involved in such an experiment to estimate the impact of the spacecraft angular momentum desaturations on the determination of Mars' orientation variations. The attitude control of the orbiter needs such desaturation maneuvers regularly repeated. They produce velocity variations that must be taken into account when determining the orbit. For our simulations, we use a priori models of the Martian rotation and introduce the spacecraft velocity variations induced by each desaturation event. By a least-squares adjustment of the simulated Doppler signals, we then estimate the orbiter velocity variations and the parameters of the Mars' rotation model. We show that these velocity variations are ill resolved when the spacecraft is not tracked, therefore requiring a near-continuous tracking from the Earth to accurately determine the orbit. In such conditions we show that only 15- of lander-orbiter tracking per week allows recovering Mars' orientation parameters with a precision of a few mas over a period of 1 Martian year. 相似文献
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G. H. Born E. J. Christensen A. J. Ferrari J. F. Jordan S. J. Reinbold 《Celestial Mechanics and Dynamical Astronomy》1974,9(3):395-414
This paper presents a comprehensive analysis of the Mars orbital phase of the Mariner 9 trajectory as determined from Earth based radio data. Both the method and accuracy of the orbit determination process are reviewed. Analysis is presented to show the effects of Mars gravity model and node in the plane of the sky errors on the accuracy of orbit determination. In addition the long term evolution of the orbit from insertion through the first 500 revolutions is presented, and decomposed into effects from the Mars garvity field,n-body perturbations, and solar radiation pressure. Since the orbit period is nearly commensurable with the Mars rotational period, the orbit experiences significant resonance perturbations. The primary perturbation is in-track with a maximum amplitude of 1000 km and a wavelength of 39 spacecraft revolutions.This paper was presented at the AIAA/AAS Astrodynamics Conference, Palo Alto, California, September 11 and 12, 1972. At this time Mariner 9 operations were still underway. The operational life of Mariner 9 ended October 27, 1972, when the supply of nitrogen gas, used for attitude stabilization, was depleted. This paper represents one phase of research carried out at the Jet Propulsion Laboratory, California Institute of Technology, Pasadena, California, under NASA Contract No. NAS 7-100. 相似文献