摘要
轨道确定是航天器飞行和控制的基础。星载GPS技术因其能够提供相对经济、连续、准确的定位服务而备受关注,成为航天器轨道确定方面的研究热点。本文以星载GPS技术在航天器编队定轨方面的应用为研究对象,分别在星载GPS测量数据处理、基于星载GPS的航天器绝对定轨、基于星载GPS的航天器编队相对定轨等方面展开研究,具体工作如下:
通过统计分析星载GPS接收机关于导航星的观测情况,提出了采用降低观测截止高度角的方法增加参与定轨的测量数据量,以提高定轨连续性。针对降低观测高度角所引起的测量误差增大问题,分析了测量误差对周跳探测的影响。在此基础上对TurboEdit算法周跳探测部分提出改进,通过引入加权系数的方式,根据观测数据质量相应地调整周跳判断条件,从而实现减少周跳误判发生,提高周跳探测效率。并分别设计了两种构造加权系数的方法。
在载波相位整周模糊度LAMBDA解算算法的实数解估计环节,设计并引入加权矩阵。根据测量时刻的观测高度角对测量数据质量进行评价,并据此在估计过程中分配以不同的权重,达到降低相关性和减小估算误差的目的。
考虑到未来GPS系统将同时提供L1、L 2和L 5三频载波信号,针对目前三频逐级模糊度解算方法存在的缺陷,提出了一套检验、修正方法。该方法综合利用载波相位宽巷、窄巷技术对三频逐级模糊度解算结果进行检验,并根据检验结果,对解算过程进行反复修正,直至解算出正确解。经过分析和讨论,证明该方法是有效的。最后给出了关于三频载波相位整周模糊度解算、检验和修正的完整工作流程。
在基于星载GPS的航天器绝对定轨研究中,对运动学最小二乘法、动力学Kalman法、弱动力学最小二乘法和弱动力学Kalman法等四种定轨方法进行了分析比较。在弱动力学定轨中,设计了一种经验加速度相关参数的选取方法。通过对定轨过程中由未建模误差和观测量误差所引起的定轨偏差进行充分估算,根据估算结果选择恰当的经验加速度相关参数,从而最大程度抵消这两种误差对定轨结果的影响。并采用GRACE卫星编队实测数据对以上方法进行了仿真验证。
在航天器绝对定轨结果的基础上,设计了两种航天器编队相对定轨方法。在“基于C-W相对动力学方程的相对定轨方法”中,提出根据绝对定轨数据计算C-W方程摄动加速度,从而使相对动力学方程由非线性褪变成线性,简化了相对定轨过程的计算。在“基于绝对轨道差分的相对定轨方法”中,采用差分的方式根据参考航天器的绝对定轨结果和成员航天器的绝对轨道动力学方程,构造相对定轨的状态方程,从而避免了C-W方程的参与,为航天器编队相对定轨增加了一条途径。同时,为了提高估计精度并减少计算量,这两种定轨方法均采用球形单边UKF滤波器。
根据星载GPS测量的特点,为充分利用观测数据,提出在两种相对定轨方法中,均同时采用参考与成员航天器之间、成员与成员航天器之间的双差载波相位组合作为观测量。从而使参与相对定轨的观测数据量显著增加,对提高相对定轨估计精度具有积极意义。最后,引入联邦滤波器,对上述两种方法的相对定轨结果进行融合,实现最优的相对定轨,并进行仿真验证。
Spacecraft orbit determination is the key technology for flying and control. Due to the ability to provide relative economic, accurate, continuous position services, spaceborne GPS technology for spacecraft gradually becomes to be the research hotspot of orbit determination technology. In this paper, the spaceborne GPS technology applied in spacecraft formation orbit determination is choose as the research object. And the spaceborne GPS measurement data processing, the absolute orbit determination, the relative orbit determination of spacecraft formation based on spaceborne GPS are researched in this paper. The specific research work are as follows:
Through the statistical analysis of GPS navigation satellite observations situation, propose that in order to increase the amount of measured data, which will participated in orbit determination calculation, by reducing the observation elevation angle. And in the same time improve the continuity of low orbit spacecraft Orbit Determination. Subsequently, as the lower elevation angle observations will lead to increased measurement error, the impact of measurement error on cycle slip detection is analyzed. On this basis, the cycle slip detection part of TurboEdit algorithm is modified. In the process of cycle slip judgment process, accordance with the quality of observable data the weighting factor is introduced. And two kinds of mathematical models are proposed to construct the weighting factors. This relates the judgment of cycle slip detection with the carrier phase measurement error and observation elevation angle. Thereby reducing the risk of misjudgment of justice cycle slip.
In ambiguity solution process of spaceborne GPS carrier phase measurements, the LAMBDA algorithm is introduced in this paper. In the least squares estimation of LAMBDA method a weighting matrix is designed. The construction of weighting matrix is depened on the elevation angle of carrier phase measurement and the quality of measurement data. According the data quality to assign different weighting factor in ambiguity in estimation process, in order to reduce the relation between different ambiguity and reduce the estimation error.
Considering that GPS system will provide three kind of frequencies carrier signal in future, in this paper the shortcoming of CAR three frequencies ambiguity solution algorithm is analyzed. Then a verification method to verify the results that given by CAR is proposed. This verification method combines both wide-lane and narrow–lane carrier phase techniques to verify the CAR results. After the verification, the revise method for the CAR ambiguity algorithm is also given. Mathematical analysis and discussion show that the test method is feasible. Finally, the whole flow of three frequencies carrier phase integer ambiguity resolution, verification and revise process is given.
In the research on spacecraft absolute orbit determination based on spaceborne GPS, four kinds of spacecraft absolute orbit determination method, which are the kinematics least-square method, dynamics Kalman method, reduced dynamic least-square method and the reduced Kalman method, are analyzed and compared. In the reduced dynamics orbit determination method, a select method of relevant parameters for empirical acceleration is proposed. Through fully estimating the unmodeling error and measurement error in orbit determination process, the appropriate parameters of empirical acceleration are chosen. With the more accurate relevant parameters, the effect of unmodeling error and measurement error on absolute orbit determination is reduced. Then a simulation of the above method which is based on the GRACE formation measurement is given.
In the research on the spacecraft formation relative orbit determination, base on the absolute orbit determination results, two kinds of spacecraft formation relative orbit determination method are designed in this paper. In the“relative orbit determination method based on C-W dynamics equation”, proposed that comput the perturbation acceleration of C-W equation in advance according to the result of absolute orbit results. So that the relative dynamics equation is changed into a linear, which simplify the process of relative orbit determination calculations. In the“relative orbit determination method based on absolute orbit difference”, the relative orbit is composed by the difference of absolute spacecraft orbit results. Thus avoids the application of C-W equation and provides a another way for spacecraft formation relative orbit determination. In order to improve the accuracy of estimation, the two relative orbit determination methods both use the spherical UKF as the filter.
In order to take full advantage of double difference observation data in spaceborne GPS, in two kinds of relative orbit determination method proposed by this paper, choose both double difference carrier phase measurement combinations of the refer spacecraft to member spacecraft and the member spacecraft to member spacecraft as observation. So that, in relative orbit determination process, the quantity of observation data has been significantly increased, and this will improve the accuracy of the relative orbit determination.Finally, the federal filter technology is introduced to fuse the results of two relative orbit determination methods, in order to get the optimal relative orbit. Then the simulations on the above methods are given.
引文
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