快速衛星姿態旋轉之順滑模態控制

Abstract

本論文之重點在於針對實際的衛星姿態控制導航系統，設計出實用的快速衛星姿態旋轉控制器。論文包含了衛星動態方程式的推導，及按照衛星指向角度大小不同的情況來建立控制法則。小角度誤差控制器是採用最佳時間需求下之比例微分控制法則，其控制導航為中華衛星二號系統所採用。大角度控制則以順滑模態控制的設計方法，來解決欲追蹤具有最大轉矩限制之衛星最短路徑旋轉的目標軌跡。衛星以最大轉矩循最短路徑旋轉的好處是提供了最小時間的旋轉。順滑模態控制具有穩健控制的特性，可消除系統不確定性因素的影響，並且考慮了實務工程中最大轉矩的限制。接著提出以平滑模型順滑模態控制來改善順滑模態控制不理想的暫態響應。兩種已提出之大角度控制方法：最短路徑之四元素法迴授控制和近似最小時間旋轉之迴授控制的控制性能，被用來證明本論文所提出的快速衛星姿態旋轉之平滑模型順滑模態控制的優越性。

The thesis described the study of the practical spacecraft attitude guidance maneuvers to design useful rapid spacecraft attitude controller. We will derive the dynamic equation of the satellite and build its control law according to the different direction angles in this thesis. A PD control law will be used as a small-angle error controller which use optimized maneuvers. The guidance be used on ROCSAT-2. A sliding-mode control (SMC) design technique is proposed for tracking the desired trajectories that are based on spacecraft eigenaxis rotation with maximum torques. The major advantage of eigenaxis rotation with maximum torques is that it provides the minimum-time maneuver. Meanwhile, the SMC controller provides the robust control for removing the effects of model uncertainty. Practical control constraint on wheel torque is also taken into account in the determination of the maximum torques. To improve the undesirable transient response before reaching the sliding manifold, the smooth model sliding-mode control is developed. Two previous developed control schemes, quaternion feedback regulator for spacecraft eigenaxis rotation and near-minimum-time eigenaxis rotation maneuvers with rate-tracking feedback, were adapted to show the superiority of the proposed design method.