Active vibration control of flexible spacecraft using modified input shaping and adaptive positive position feedback

由于柔性航天器的柔性动力学与姿态/轨道动力学相互作用，使得高精度姿态/轨道控制变得更加困难，因此主动振动控制是现代柔性航天器的关键问题。针对这一问题，提出了一种将修正输入整形技术与自适应正位置反馈相结合的柔性航天器振动控制策略。改进的输入整形将集成到航天器作动器、动量轮和喷气推进器中，以抑制几种低阶振动模式，而自适应正位置反馈将应用于压电智能作动器，以控制主要的高阶振动模式。研究工作将进行如下。首先，研究了修改输入整形的基本原理，并开发了一种自适应算法，用于自动调整正位置反馈滤波器的参数。其次，基于柔性多体动力学，建立了柔性航天器的详细三轴动力学模型。研究执行机构的配置问题，确定最优配置。第三，开发了将修正输入整形与自适应正位置反馈相结合的混合振动控制器。还将开发最小化这两个振动控制器相互作用的方法。通过稳定性理论证明了混合振动控制器的稳定性。此外，还将考虑不同的航天器在轨运行模式。然后利用MATLAB/Simulink进行数值仿真，验证理论分析的正确性。最后，通过硬件实验验证所提出的振动控制策略的有效性。利用该策略，可以快速地将航天器的柔性振动抑制到允许的水平，从而实现高精度的姿态指向控制。

Active vibration control is a crucial issue for modern flexible spacecraft because the flexible dynamics may interact with the attitude/orbit dynamics, which makes high-precision attitude/orbit control more difficult. To address this issue, a strategy that combines modified input shaping technique and adaptive positive position feedback is proposed to control vibration of flexible spacecraft. Modified input shaping will be integrated with spacecraft actuators, momentum wheels and jet thrusters, to suppress several low-order vibration modes, while adaptive positive position feedback will be applied to piezoelectric smart actuators to control the dominant high-order vibration modes. The research work will be conducted as follows. Firstly, the basic principles of modified input shaping will be investigated and an adaptive algorithm will be developed for automatic tuning of parameters of the positive position feedback filters. Secondly, a detailed 3-axis dynamic model for flexible spacecraft will be developed based on flexible multi-body dynamics. The actuator configuration problem will be studied and the optimal configuration will be determined. Thirdly, the hybrid vibration controller that combines modified input shaping and adaptive positive position feedback will be developed. Method will also be developed to minimize the interaction of these two vibration controllers. The stability of the hybrid vibration controller will be proved through stability theory. Different spacecraft on-orbit operation modes will be considered as well. Then, numerical simulation will be conducted using MATLAB/Simulink to validate the theoretical analysis. Finally, hardware experiments will be conducted to verify the effectiveness of the proposed vibration control strategy. By using the proposed strategy, it is expected that the flexible vibration of spacecraft can be suppressed to an allowable level quickly so that high-precision attitude/pointing control can be achieved.