Coupled Analysis And Cooperative Control For Spacecraft Attitude Motion And Solar Panel Vibration

Flexible spacecraft are usually fitted with solar panels to provide energy for the normal operation.They are typical rigid-flexible coupling system because of the coupling between the solar panel elastic vibration and the spacecraft attitude motion.To achieve stronger function and longer life of modern spacecraft,the solar panel size is growing,which leads to more conspicuous rigid-flexible coupling effect.At present,most of the researches on dynamics and control of flexible spacecraft rely on the modes of the individual solar panel(i.e.,assumed mode).This kind of mode cannot reflect the rigid-flexible coupling effect,which results that those studies have a certain shortage when used to study the spacecraft with large-span solar panels.Therefore,it is of great theoretical and engineering value to obtain the rigid-flexible coupling modes(RFCMs)of flexible spacecraft to conduct coupled analysis and cooperative control of spacecraft attitude motion and solar panel vibration.In this dissertation,the spacecraft with large-span solar panels is studied in detail by analytical and numerical approaches.Approaches are firstly proposed to obtain rigid-flexible coupling modes.Then,investigations are conducted relevant to the natural characteristics,dynamic modeling and dynamic characteristics,and cooperative control of attitude-structure motion.The main contents and achievements derived in this dissertation are listed as following.An analytical method is developed to directly obtain the RFCMs of flexible spacecraft.Describing the spacecraft attitude motion and solar panel vibration with a uniform set of generalized coordinates,the analytical expressions and frequency values of RFCMs can be accurately solved from the rigid-flexible coupling dynamic equations of the continuous system of flexible spacecraft by using the boundary conditions.The convergence study reveals that,as the degree of freedom of discrete models based on assumed modes or finite elements increases,the frequencies calculated from the discrete models converge to those from the proposed analytical method.Compared with the modeling approach with assumed modes,those RFCMs can be used to derive the low-dimensional but high-precise discrete dynamic model of the flexible spacecraft,and the low-dimensional controller designed by using this model can achieve the cooperative control-index for spacecraft attitude motion and solar panel vibration in a shorter time with less input energy.The Rayleigh-Ritz method widely used in free vibration analysis of elastic structures is extended to investigate the modal characteristics of the rigid-flexible coupling hub-beam/plate system by representing the spacecraft rigid-body motion with the product of rigid-body mode and generalized coordinate.The extended Rayleigh-Ritz method can compute the analytical expressions and frequencies of RFCMs with high accuracy,excellent convergence and high efficiency.Using this method,numerical simulations are conducted for a three-axis attitude stabilized spacecraft installed with a pair of solar panels.Based on the analyses of those simulation results,the characteristics of RFCMs are investigated comprehensively,and then a criterion is given to determine the rigid-flexible coupling property of each mode.Finally,the influences of the solar panel length and the moment of inertia of the rigid body on the system’s modes are studied.Modeling the solar panel composed of honeycomb sandwich panel by a three-layer laminated plate,i.e.,the honeycomb core and two facesheets,rather than a single-layer isotropic plate,a high-precise discrete dynamic model which is closer to the real engineering is established by using RFCMs of a flexible spacecraft.Subsequently,the effects of honeycomb panel parameters and spacecraft flexibility on the system’s natural characteristics are investigated.Applied the discrete dynamic model and combining the input shaping(IS)technique with the proportional-derivative(PD)controller,a real-time control scheme with simple algorithm is designed to achieve the cooperative control of attitude maneuver and panel vibration for spacecraft with different flex ibility and to inhibit the possible occurrence of unstable thermally induced vibration for flexible spacecraft subjected to solar radiation.For an attitude maneuvering flexible spacecraft under solar radiation,an explicit recursive scheme is derived with finite difference method to calculate the cross-sectional temperature distribution of solar panels composed of honeycomb panel,and the system’s rigid-flexible-thermal coupling dynamic model is established in terms of RFCMs by considering the thermal stress and strain.Then,this model is used to compute the spacecraft thermally induced responses by employing a coupled solving procedure.The numerical results reveal that: the solar panel of a maneuvering spacecraft is subjected to time-varying thermal loading,and the thermally induced responses consist of quasi-static displacement and superimposed vibration;if the final incident angle of heat flux(the sum of initial incident angle and the maneuver attitude angle)is large,the thermally induced responses of a spacecraft with small damping may be unstable and the thermal flutter may occur in this case;the thermally induced vibration decays with time and only the quasi-static deformation exists finally in the case of large damping.