Research On Attitude Determination And Control Algorithm Of Fast Maneuver Satellite

The attitude determination algorithm of large angle and fast maneuver satellites has become the hot spot in recent years. The omitted items in Taylor expansion are not infinitesimal items since the satellite dynamic model is strong nonlinear during the maneuver process, thus the accuracy of the filter process descends dramatically. In order to solve this problem, the work done in this thesis is:Introduce the concept of model error in order to describe the strong nonlinearity of the system model. The estimation equation of the error model is given based on the minimum model error principle and the system dynamic model is modified in order to improve the accuracy of the filter. The complementary filter is introduced to integrate the high accuracy but long cycle character of star tracker and the low accuracy but short cycle character of gyroscope. The complementary filter for satellite attitude determination is established to enhance the accuracy of the filter system. The robust filter is introduced since the linearized model differs relatively much to the real model. The system model error is treated as the uncertain items and the error covariance matrix is constrained in an acceptable field by matrix unequal transformation, thus the state estimation with high accuracy is obtained.Moreover, the work done in this thesis in order to achieve the goal of large angle of fast maneuver is listed below:The path planning algorithm is presented based on Gauss pseudo spectral method and the point collocation method. The time optimal maneuver path is given based on the assumption that the time optimal condition occurs when the output torque is saturated. The time optimal trajectory is also derived according to the Bang-Bang control law for single axis maneuver.The Finite Time control is introduced in order to design the attitude tracking control law to track the expected trajectory in finite time. The Integration sliding mode and terminator function are introduced and based on these two method the attitude tracking law is designed. Considering the satellite will not always be in the fast maneuver mode, and in order to bring in the output torque saturation, the finite output control law is designed in this chapter.