Research On Multi-body Dynamics And Vibration Control Of Spacecraft Solar Array System

With the development of astronautical science and technology,modern spacecraft often deploy large-scale,complex and lightweight appendages for new requirements in functionality for a a variety of space missions.Whether the solar arrays can be successfully deployed and locked after entering orbit,and quickly reach a stable and normal working state directly determines the success or failure of the space mission.The deployment and locking process of large-scale flexible solar arrays present the typical coupling characteristics of large rotation motions and large elastic deformations.Moreover,the composite laminated solar panel further induces complex nonlinear dynamic behaviors.Therefore,the deployment of the solar array is a highly dynamic movement that involving strong coupling characteristics and nonlinearity.In addition,the spacecraft will periodically enter and exit the shadow and sunshine area during orbiting the Earth.The thermally vibration of solar array will severely affect the performance of spacecraft.This thesis establishes the rigid-flexible coupling dynamic model of spacecraft solar array multibody system,and studies the deployment dynamics,the thermally induced vibration and the vibration suppression of solar panels.The main contents and conclusions are listed as follows.The rigid-flexible coupling dynamical model of the spacecraft system with solar array is established based on the Natural Coordinate Formulation and the Absolute Nodal Coordinate Formulation.The deployment of the solar array is numerically simulated based on Hilber-Hughes-Taylor-I3 method and dynamic responses of spacecraft system are analyzed.The effectiveness of the rigid-flexible coupling model is verified by comparing with the co-simulation results from ADAMS and ABAQUS software.Solar panels are deployed synchronously under the combined action of the torques generated by the driving and closed cable loop mechanisms.The deployment of the solar array causes the positon deviations in the direction of the sailboard deployment and the vertical direction of the deployment plane,and the attitude disturbances in the direction of the revolute joint’s rotation axis on spacecraft main-body.After the solar panels are locked,the locking mechanisms generate strong impact moments which cause strong structural vibrations on the flexible solar panels.The flexibility of solar panels significantly affects the vibration responses of solar panels.A three-dimensional thermal-structural element is proposed based on the Absolute Nodal Coordinate Formulation,and the rigid-flexible-thermal coupling dynamic model of spacecraft solar array system is established considering space thermal environment.The results show that the thermally induced vibrations of the solar panels caused by dramatic temperature changes are relatively stable when spacecraft system traveled through the umbra and sunlight regions.These are sudden changes of vibration amplitude when spacecraft system traveled through the penumbra region,which corresponds to the trend of solar radiation heat flux received by the solar panels.The solar array experiences sudden space heating when the spacecraft travels through the penumbra region.The proposed model can well-reveal the coupling effect of the rigid-body motion,elastic deformation,and temperature distribution for the solar array deployed on the attitude-maneuvering spacecraft.The coupling effect presents the change of the space radiation flux causes the thermos-elastic deformation of the solar panel,which leads to the attitude disturbance of spacecraft while the space radiation flux absorbed by the solar array depends on the elastic deformation of the solar panel and the attitude of spacecraft.In addition,the orbital height of spacecraft is an important factor for the temperature distribution of the solar panels.The heat fluxes of the Earth-reflected radiation and Earth-emitted radiation have a non-ignorable effect on the temperature change in the solar array deployed on the low-earth-orbital spacecraft.The piezoelectric actuator-sensor coupling element is formulated based on the Absolute Nodal Coordinate Formulation and the effectiveness of the element is verified by comparing with the analytical solution.The intelligent vibration control scheme using piezoelectric actuators and sensors for a deployable solar array is designed.The results show that the electric potential value of the sensors present a consistent trend with the element strain,which indicates that the electric potential of the sensor induce the strain state of the element,and then reflects the structural vibration of the solar panel.After the solar array is deployed and locked,the electric potential value of the sensors near spacecraft main-body fluctuates greatly which indicates that the solar panels produce strong structural vibration near the spacecraft main-body.By comparing the feedback signals of the piezoelectric sensors of the solar panel before and after control,it is found that the piezoelectric stress generated by the actuator can offset the interference excitation force and bending moment acting on the solar panel,thus effectively suppressing the vibration of flexible solar panels.The positon deviations and attitude disturbances of the spacecraft main-body are analyzed during the deployment of the solar array based on the rigid-flexible coupling dynamical model of spacecraft system.On this basis,a fuzzy PD control scheme is proposed for eliminating the positon deviations and attitude disturbances of spacecraft main-body during the deployment of the solar array.The results show that the proposed fuzzy PD controller adjusts the gain coefficient online according to the dynamic responses of spacecraft main-body,which effectively improves the control performance of the classical PD controller.