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1.
XuHai Yang ZhiGang Li ChuGang Feng Ji Guo HuLi Shi GuoXiang Ai FengLei Wu RongChuan Qiao 《中国科学G辑(英文版)》2009,52(3):333-338
A geostationary (GEO) satellite may serve as a navigation satellite, but there is a problem that maneuvers frequently occur
and the forces are difficult to model. Based on the technique of determining satellite orbits by transfer, a predicted orbit
with high accuracy may be achieved by the method of statistical orbit determination in case of no maneuver force. The predicted
orbit will soon be invalid after the maneuver starts, and it takes a long time to get a valid orbit after the maneuver ends.
In order to improve ephemeris usability, the method of rapid orbit forecasting after maneuvers is studied. First, GEO satellite
movement is analyzed in case of maneuvers based on the observation from the orbit measurement system by transfer. Then when
a GEO satellite is in the free status just after maneuvers, the short arc observation is used to forecast the orbit. It is
assumed that the common system bias and biases of each station are constant, which can be obtained from orbit determination
with long arc observations. In this way, only 6 orbit elements would be solved by the method of statistical orbit determination,
and the ephemeris with high accuracy may be soon obtained. Actual orbit forecasting with short arc observation for SINOSAT-1
satellite shows that, with the tracking network available, the precision of the predicted orbit (RMS of O-C) can reach about
5 m with 15 min arc observation, and about 3 m with 30 min arc observation.
Supported by the National High Technology Research and Development Program of China (Grant No. 2006AA12Z322), the National
Basic Research Program of China (Grant No. 2007CB815503), and the West Light Program of Chinese Academy of Sciences (Grant
No. 2007LH01) 相似文献
2.
Jingrui Zhang 《Acta Mechanica Sinica》2010,26(1):141-146
Studied in this paper are the attitude control law design and the output torque estimation problem of micro control moment gyros (MCMGs) for the agile satellites executing rapid attitude maneuver mission. An algorithm is proposed for estimating the output torques and the gimbal angular rates of MCMGs, which can help engineers to choose reasonable size for actuators so that the cost of satellite can be decreased. According to some special maneuver missions, a numerical example of attitude control system for a small satellite with MCMGs in pyramid configuration is studied, and the simulation results validate the proposed estimation algorithm. 相似文献
3.
有效估计阵形是提高机动条件下拖线阵声呐探测性能的关键,流体力学类阵形估计方法稳定性和可靠性欠佳,导致其难以应用于工程实际,该文针对此问题提出一种基于拖线阵运动特性的阵形估计方法。利用稳态振荡响应公式计算拖船回转机动时拖线阵稳态阵形特性,将转向机动过程中阵上各点运动状态划分为若干阶段,进而依据偏微分方程特征线理论计算各阶段的分界时刻,探究阵上相邻两点的沿阵切线方向差变化规律,最后通过计算当前阵上各点的沿阵方向实现阵形估计。计算机仿真和海上实验数据验证表明算法可行且有效,与传统的流体力学类阵形估计方法相比具有更高的稳定性和更好的工程应用前景。 相似文献
4.
Wheel-off-loadings and orbital maneuvers of the GEO satellite result in additional accelerations to the satellite itself.
Complex and difficult to model, these time varying accelerations are an important error source of precise orbit determination
(POD). In most POD practices, only non-maneuver orbital arcs are treated. However, for some applications such as satellite
navigation RDSS services, uninterrupted orbital ephemeris is demanded, requiring the development of POD strategies to be processed
both during and after an orbital maneuver. We in this paper study the POD for a maneuvered GEO satellite, using high precision
and high sampling rate ranging data obtained with Chinese Area Positioning System (CAPS). The strategy of long arc POD including
maneuver arcs is studied by using telemetry data to model the maneuver thrust process. Combining the thrust and other orbital
perturbations, a long arc of 6 days’ CAPS ranging data is analyzed. If the telemetry data are not available or contain significant
errors, attempts are made to estimate thrusting parameters using CAPS ranging data in the POD as an alternative to properly
account for the maneuver. Two strategies achieve reasonably good data fitting level in the tested arc with the maximal position
difference being about 20 m.
Supported by the National Natural Science Foundation of China (Grant No. 10703011) and the Science & Technology Commission
of Shanghai Municipality of China (Grant No. 06DZ22101) 相似文献
5.
XNAV/UVNAV/SINS组合导航在航天器轨道机动中的应用 总被引:1,自引:0,他引:1
针对X射线脉冲星导航在航天器轨道机动过程中精度不高甚至发散的问题,提出一种将X射线脉冲星导航结合惯性导航和紫外敏感器的组合导航方法。以航天器在惯性系中的位置、速度、姿态四元数和惯性导航设备误差作为系统状态变量,用X射线探测器测量X射线脉冲到达时间,用紫外敏感器测量中心天体质心相对于航天器的方向矢量和距离以及航天器在惯性系中的姿态四元数,用扩展卡尔曼滤波器估计组合导航系统状态。仿真结果验证了该组合导航方法的可行性,能够解决轨道机动中X射线脉冲星单独导航的误差过大(位置误差达107m)问题,且该组合导航具有较高的导航精度,在轨道机动前、机动中和机动后导航位置误差均在100 m以内。 相似文献
6.
针对加速度计、磁强计易受机动加速度和周围环境影响的问题,提出了一种单基线GPS/MIMU组合姿态估计方法。该方法利用单基线GPS天线代替磁强计提供航向信息,并利用GPS的速度信息对加速度计输出的机动加速度分量进行补偿,构成单基线GPS/MIMU组合姿态确定单元。同时考虑实时性要求,在SRUKF的基础上,引入加性非扩展形式和超球体单形采样,提出了简化超球体平方根UKF算法,通过减少状态维数和采样点的数量,降低算法计算量。建立加性噪声下的单基线GPS/MIMU姿态模型,并采用简化超球体平方根UKF算法进行姿态估计。实验结果表明,单基线GPS的速度信息可以有效提高加速度计对水平倾角的确定精度,姿态估计算法收敛后的最大误差小于0.8°,估计精度与UKF相当,且执行时间仅是UKF的42%。 相似文献
7.
针对BTT导弹非线性动力学模型中的参数不确定性及系统存在未知外部干扰的问题,提出了一种基于扩张状态观测器的滑模控制设计方法;将系统参数不确定性和未知外界干扰作为系统的复合干扰,采用扩张状态观测器估计系统的复合干扰,并实时补偿;利用Lyapunov方法证明系统的稳定性;该控制器算法简单、计算量小,更易于实际工程应用;仿真结果表明,设计的控制器对导弹控制系统受到的复合干扰具有较强的鲁棒性和稳定性,满足导弹姿态快速机动和高稳定度的控制要求,性能指标明显优于滑模控制器。 相似文献
8.
The control of flight forces and moments by flapping wings of a model bumblebee is studied using the method of computational fluid dynamics.Hovering flight is taken as the reference flight:Wing kinematic parameters are varied with respect to their values at hovering flight.Moments about(and forces along)x,y,z axes that pass the center of mass are computed.Changing stroke amplitude(or wingbeat frequency)mainly produces a vertical force.Changing mean stroke angle mainly produces a pitch moment.Changing wing angle of attack,when down-and upstrokes have equal change,mainly produces a vertical force,while when down-and upstrokes have opposite changes,mainly produces a horizontal force and a pitch moment.Changing wing rotation timing,when dorsal and ventral rotations have the same timing,mainly produces a vertical force,while when dorsal and ventral rotations have opposite timings,mainly produces a pitch moment and a horizontal force.Changing rotation duration has very small effect on forces and moments.Anti-symmetrically changing stroke amplitude(or wingbeat frequency)of the contralateral wings mainly produces a roll moment.Anti-symmetrically changing angles of attack of the contralateral wings,when down-and upstrokes have equal change,mainly produces a roll moment,while when down-and upstrokes have opposite changes,mainly produces a yaw moment.Anti-symmetrically changing wing rotation timing of the contralateral wings,when dorsal and ventral rotations have the same timing,mainly produces a roll moment and a side force,while when dorsal and ventral rotations have opposite timings,mainly produces a yaw moment.Vertical force and moments about the three axes can be separately controlled by separate kinematic variables.A very fast rotation can be achieved with moderate changes in wing kinematics. 相似文献
9.
Yue Bao-Zeng 《Nonlinear dynamics》2008,51(1-2):317-327
In this paper, the chaotic dynamics in an attitude transition maneuver of a rigid body with a completely liquid-filled cavity
in going from minor axis to major axis spin under the influence of viscous damping and a small flexible appendage constrained
to undergo only torsional vibration is investigated. The focus in this paper is on the way in which the dynamics of the liquid
and flexible appendage vibration are coupled. The equations of motion are derived and then transformed into a form suitable
for the application of Melnikov's method. Melnikov's integral is used to predict the transversal intersections of the stable
and unstable manifolds for the perturbed system. An analytical criterion for chaotic motion is derived in terms of the system
parameters. This criterion is evaluated for its significance to the design of spacecraft. The dependence of the onset of chaos
on quantities such as body shape and magnitude of damping values, fuel fraction and frequency of flexible appendage vibration
are investigated. 相似文献
10.
Walter Hohmann was a civil engineer who studied orbital maneuvers in his spare time. In 1925, he published an important book (Ref. 1) containing his main result, namely, that the most economical transfer from a circular orbit to another circular orbit is achieved via an elliptical trajectory bitangent to the terminal orbits. With the advent of the space program some three decades later, the Hohmann transfer maneuver became the most fundamental maneuver in space.In this work, we present a complete study of the Hohmann transfer maneuver. After revisiting its known properties, we present a number of supplementary properties which are essential to the qualitative understanding of the maneuver. Also, we present a simple analytical proof of the optimality of the Hohmann transfer and complement it with a numerical study via the sequential gradient-restoration algorithm. Finally, as an application, we present a numerical study of the transfer of a spacecraft from the Earth orbit around the Sun to another planetary orbit around the Sun for both the case of an ascending transfer (orbits of Mars, Jupiter, Saturn, Uranus, Neptune, and Pluto) and the case of a descending transfer (orbits of Mercury and Venus). 相似文献