Nonlinear H∞ based underactuated attitude control for small satellites with two reaction wheels

Underactuated attitude control is supposed to be used on spacecraft when failure happens with onboard actuators. One main problem with existing underactuated attitude control designs is their limited capabilities against disturbances. In order to solve this problem, an approach based on the theory of H∞$H\infty$ is proposed in this paper. Two propositions are derived from the H∞$H\infty$ theory to improve the robustness of one popular underactuated attitude control design, which was presented by Tsiotras et al. It is proved mathematically that the controller satisfying these two propositions respectively can stabilize the underactuated attitude system locally or globally. The numerical simulations show that the improved controllers based on the H∞$H\infty$ theory could provide higher pointing accuracy for small satellites against disturbances. This validates the effectiveness of the proposed H∞$H\infty$ based approach to improve existing underactuated attitude control designs.     

Effects of uncertainties and flexible dynamic contributions on the control of a spacecraft full-coupled model

One of the most important problems for performing a good design of the spacecraft attitude control law is connected to its robustness when some uncertainty parameters are present on the inertial and/or on the elastic characteristics of a satellite. These uncertainties are generally intrinsic on the modeling of complex structures and in the case of large flexible structures they can be also attributed to secondary effects associated to the elasticity. One of the most interesting issues in modeling large flexible space structures is associated to the evaluation of the inertia tensor which in general depends not only on the geometric ‘fixed’ characteristic of the satellite but also on its elastic displacements which of course in turn modify the ‘shape’ of the satellite. Usually these terms can be considered of a second order of magnitude if compared with the ones associated to the rigid part of a structure. However the increasing demand on the dimension of satellites due to the presence for instance of very large solar arrays (necessary to generate power) and/or large antennas has the necessity to investigate their effects on their global dynamic behavior in more details as a consequence. In the present paper a methodology based on classical Lagrangian approach coupled with a standard Finite Element tool has been used to derive the full dynamic equations of an orbiting flexible satellite under the actions of gravity, gravity gradient forces and attitude control. A particular attention has been paid to the study of the effects of flexibility on the inertial terms of the spacecraft which, as well known, influence its attitude dynamic behavior. Furthermore the effects of the attitude control authority and its robustness to the uncertainties on inertial and elastic parameters has been investigated and discussed.     

A hybrid attitude controller consisting of electromagnetic torque rods and an active fluid ring

In this paper, a novel hybrid actuation system for satellite attitude stabilization is proposed along with its feasibility analysis. The system considered consists of two magnetic torque rods and one fluid ring to produce the control torque required in the direction in which magnetic torque rods cannot produce torque. A mathematical model of the system dynamics is derived first. Then a controller is developed to stabilize the attitude angles of a satellite equipped with the abovementioned set of actuators. The effect of failure of the fluid ring or a magnetic torque rod is examined as well. It is noted that the case of failure of the magnetic torque rod whose torque is along the pitch axis is the most critical, since the coupling between the roll or yaw motion and the pitch motion is quite weak. The simulation results show that the control system proposed is quite fault tolerant.     

载人航天器放电调节器稳定性研究

电源系统的稳定性是保证载人航天器在轨飞行可靠安全的要素之一。文章研究了应用混合型功率调节技术的电源系统放电调节器的建模方法,获得了系统开环传递函数,进行了仿真验证,在此基础上分析了影响放电调节器稳定性的因素,表明系统中存在过大的延迟会使系统的稳定性大幅降低甚至造成不稳定,地面试验结果验证了此结论的正确性。     

四轴飞行器姿态监控系统设计

设计了一套四轴飞行器上位机监控系统。针对准确掌握四轴飞行器的实时姿态、位置等信息和对四轴飞行器进行精准控制两大核心问题,开发了由上位机实现四轴飞行器姿态数据采集、通信及控制等功能的监控系统。监控系统良好的功能设计和友好的人机界面使得用户能够准确掌握飞行器的飞行信息,从而能够更精准地控制四轴飞行器飞行姿态。实验验证,监控系统简单实用,性能稳定,精度高,使整个飞行系统抗干扰能力强,鲁棒性、稳定性好。     

星载SAR轨道及姿态误差模型分析

平台轨道及姿态误差模型是星载SAR数据模拟及相应的模拟系统的关键组成部分。简要分析了机载SAR和星载SAR的差异,重点对星载SAR的平台轨道及姿态模型进行了分析,最后对星载SAR回波信号仿真方法进行了分析。     

基于低成本陀螺和倾角仪的姿态估计

针对动中通测控系统中低成本陀螺和倾角仪的姿态估计和陀螺误差校正问题,提出一种利用UKF（Uncented Kalman Filter）滤波器对载体姿态角进行估计,然后利用互补滤波器对陀螺漂移进行估计的算法。该算法通过设计三维完全可观测UKF滤波方程和陀螺误差校正模型对姿态角和陀螺漂移分别进行估计,有效避免了利用卡尔曼滤波进行姿态估计和陀螺漂移误差估计时由于误差模型不准确而产生的发散问题,同时降低了滤波器维数。试验中姿态角估计误差在1°以内,x轴陀螺漂移估计误差为0.0148°/s,y轴陀螺漂移估计误差为0.0017°/s,试验结果表明该算法能有效提高姿态角估计和陀螺漂移估计的精度。     

挠性航天器的鲁棒多目标姿态控制器设计

航天器挠性附件的振动难以直接测量且对航天器的姿态控制精度产生不利影响。为解决这一问题,提出一种基于观测器的鲁棒多目标姿态控制系统设计方法,借助特征结构配置的参数化方法,给出了含有自由参向量的控制器以及观测器的完整参数化形式。通过对自由参向量的多目标优化选取,使控制系统具有良好的鲁棒性和干扰抑制能力,并避免了控制输出饱和现象的发生。对一个挠性航天器的设计实例表明了方法的有效性。
     

具侧向脉冲力制导炮弹的非线性稳定性分析

研究了制导炮弹受侧向脉冲推力作用后处于大攻角情况下的非线性运动稳定性问题.应用首次积分描述了弹丸的运动方程,得到了描述攻角余弦变化的多项式,通过对多项式的分析将弹丸角运动稳定性的问题转化为求解多项式的负根问题,最后运用霍尔维茨定理给出了弹丸非线性运动稳定的充分条件,为非线性运动条件下弹丸的飞行形态的预测、脉冲执行机构的设计等问题提供了一定的理论依据.      

计入畸变修正的旋翼尾迹前飞状态稳定性分析

现有的旋翼尾迹稳定性分析方法在涡线过近处会得出不现实的过大发散率结果,因而存在一定的局限性。为了消除这种前飞状态下尾迹高度畸变对其稳定性分析所造成的不合理干扰,提出了一个旋翼尾迹在正弦扰动下,结合涡线畸变修正的线性化稳定性分析方法。在该方法中,将尾迹涡线离散为一系列涡元,并对其端点施加沿尾迹涡线分布的正弦扰动;考虑了桨尖涡的自诱导、互诱导以及与桨叶的干扰作用,并创新地引入在伪涡核对稳定性分析中,因尾迹高度畸变产生的不合理干扰而进行的修正。以H-34型直升机旋翼模型为例,在前飞状态下,重点计算分析了尾迹向后飘移、涡核大小和前飞速度等对旋翼尾迹稳定性的影响。研究结果表明:引入伪涡核模型可以有效地修正在涡线过近处所产生的不现实稳定性分析结果;前飞速度、涡核大小等因素会在一定条件下有限地影响尾迹的不稳定性,但都不改变前飞状态旋翼尾迹的本质不稳定性。