Dynamic Modeling And Attitude Control For Flexible Spacecraft

With the development of aerospace techniques, more and more large scale spacecrafts adopt flexible structure, e.g. large scale flexible antenna, solar array and large space etc. Rigid body motion of flexible spacecraft in orbit will excite elastic deformation of flexible body, and these two kinds of motion are coupled and affected by each other which results in the significant complexity of rigid-flexible coupling. Hence, research on dynamic model of flexible spacecraft and designing its attitude control system has important significance. Supported by CASC project"Complex framework satellite intelligent adaptive inverse control technique research"and NSFC project"Adaptive inverse control based on passivity approach and its application in a class of complex system", we take the flexible spacecraft with cantilever beam where the beam appendage is connected with body through flexible joints as the research object in this thesis, and do a deep research on modeling of flexible spacecraft, attitude control and suppressing of appendage vibration in term of theory. It mainly consists of the following aspects:1. Dynamics model of flexible spacecraft with cantilever beam is derived using Lagrange Dynamics Equation after analyzing rigid model and model of spacecraft with clamped support cantilever beam appendage. Model analysis method is adopted and vibration equation of beam appendage is derived. Coupling dynamics equation of flexible spacecraft is normalized to adapt the analysis and design of attitude control system. Furthermore, models of flexible spacecraft with multiple cantilever beam appendage and multiple cantilever beam appendage with joints are derived. Finally, characteristic of spacecraft actuator is analyzed simply, and parameter of PWPF is designed.2. Attitude control methods of flexible spacecraft include output feedback, state feedback and decentralized control, in which the output feedback has the advantage of simple structure. The input shaping is introduced to suppress low frequency vibration of spacecraft based on the idea of traditional output feedback, and variable structure output feedback controller is designed using tracking control model. It was compared with the traditional PD control strategy. The simulation results indicated the presented method had the obvious advantage in suppress the low frequency vibration exited by maneuver of spacecraft body. At the same time, similarities and differences of attitude control requirement between this kind of spacecraft and generals is discussed based on the simulation results.3. State observer is designed to utilize less sensors on beam estimating vibration modes of appendage. By this way, the requirement of state feedback is approved. On the basis of this, optimal guaranteed cost feedback control law of uncertain parameter system is analyzed and designed when control is constrained. This control law has good robustness via solving LMI and extreme value optimization problem.4. Piezoelectric elements are used to suppress vibration which makes system model be a multi input system. At first, optimal guaranteed cost feedback control law of control constrained multi input system is studied, which is a centralized control method. Local state observer and decentralized feedback control law are designed based on the design idea of decentralized control for local state feedback. The simulation results indicate that both of these two methods can approve the attitude control precise requirement and suppressed the vibration of flexible appendage efficiently.