Aiming at the proximity operation of spacecraft, a collision-avoidance control algorithm based on improved Artificial Potential Fields (APF) method is proposed. According to the real-time state between the spacecraft and the target and obstacles, the APF method is used to calculate the real-time acceleration of the spacecraft, and the trajectory of the spacecraft is planned. In order to improve the applicability of the artificial potential function method, three improvement measures are proposed. First, in order to improve the accuracy of collision warning and reduce additional maneuvers, the collision probability combined with the relative distance, instead of only the relative distance, is used to evaluate the collision. Second, in order to increase the docking safety and slow down the approaching relative velocity, the safety boundary and control margin of relative velocity are used to calculate the target repulsion force. Third, in view of the fact that most spacecraft cannot provide any continuously varying thrust, two practical thrust forms, including the thrust with upper limit and constant variation rate and the bang-bang thrust, are used to substitute for the continuously variable thrust form. Numerical simulations are executed to validate the proposed method. The effects of the major mission parameters, such as the collision warning method, the target repulsion acceleration and acceleration forms, are successfully revealed by the comparison between different examples. The results show that the proposed method can improve the safety and efficiency of the spacecraft proximity operations, and has simple structure and strong real-time performance.
There are still lots of joint clearances that cannot be eliminated for large-scale flexible spacecraft in post-lock phase. Joint clearance directly affects the attitude maneuver of the flexible spacecraft as well as the pointing accuracy and stability of the payload, which has a great influence on the dynamic characteristics of the spacecraft. Aiming at this issue, a dynamic modelling and control method for the rigid-flexible coupling spacecraft with joint clearance is proposed in this paper. The accurate dynamic model of the joint with clearance is established firstly, thus the dynamic model of flexible structure with joint clearances is built. Then the discrete rigid-flexible coupling nonlinear dynamic model of the spacecraft with clearances is obtained by Hamilton principle and modal discrete method. The Newmark algorithm is used to solve the nonlinear equation. Based on macro fiber composite (MFC) actuator, the rigid-flexible-electrical coupling dynamic equation of the spacecraft is obtained and the control law is designed by the optimum control. The influences of joint parameters, moment of inertia of central rigid body, clearance size and clearance number on the dynamic characteristics of the spacecraft are analyzed. The effects of joint clearance on the attitude maneuver and structural vibration of the spacecraft are emphatically studied. Finally, the active control is applied to the spacecraft using MFC actuator. The results reveal that the joint parameters and moment of inertia of the central rigid body affect the natural frequency of the spacecraft. With the increase of the size of joint clearance and number of clearances, the overall stiffness of the spacecraft decreases gradually, while the attitude angle and vibration displacement response of the spacecraft increase. Through the active control based on MFC, the cooperative control of the attitude maneuver and structural vibration of the spacecraft with clearance can be realized, and the effects of clearance on the dynamic characteristics of the spacecraft can be alleviated.
During the deformation process, the dynamic modeling of the folding-wing aircraft presents the characteristics of multi-rigid、multi-degree of freedom and strong nonlinearity. At the same time, parameters such as aerodynamics/torque, pressure center, centroid and moment of inertia will also change greatly, which will seriously affect Flight stability. Therefore, this paper will mainly study the multi-rigid dynamics modeling and deformation stability control of the folding wing aircraft. The multi-rigid dynamic model of the folding wing aircraft is established based on the Kane method with the additional force and moment expressions. The functional relationship between the aerodynamic parameters and the folding angle is fitted through aerodynamic calculations. and the longitudinal dynamic characteristics of the aircraft at different folding angular speeds are analyzed. It is shown that the speed, height and pitch angle of the folding wing aircraft will change during the deformation process by analyzing the longitudinal dynamic characteristics, and the aircraft cannot maintain stable flight. A stability control method is proposed for the deformation process of the folding-wing aircraft based on the active disturbance rejection control theory. The nonlinear terms, coupling terms and parameter time-varying terms are regarded as the total internal and external disturbances in the longitudinal nonlinear dynamic model of the folding-wing aircraft, using the extended state observer to estimate and compensate the total disturbance in real time. The PD controller is proposed for compensated systems to realize decoupling control of speed channel and height channel. The stability of the system is proved by Lyapunov stability theory, and mathematical simulation is used to verified the stability of the folding wing aircraft. The simulation results show that the stability controller based on the active disturbance rejection control theory can solve the problems of strong nonlinearity and time-varying parameters caused by aircraft deformation, and ensure the high-precision and stable control of the aircraft.
Due to the high aero to inertia ratio and the presence of strong aerodynamic forces, the low Earth orbit nanosatellites are not very appropriate to depend on a set of momentum wheels for attitude controlling. A method of utilizing aerodynamic disturbance torque as control input based on mass moment technology is innovatively proposed for the Nano-satellite in the low Earth orbit to solve the problem of the external aerodynamic force. The exclusive use of moving mass actuator would lead to an underactuated as the aerodynamic torque was perpendicular to the relative flow vector. To achieve full three-axis stabilization, a three-axis magnetorquer is used to complement the moving mass system to generate a torque along the orbital velocity. The whole dynamic equations are derived, which describes the system with two actuators, the movable mass and the magnetorquer, actuating simultaneously. According to the influence of disturbance items, the equations are simplified. Considering the uncertainty of the aerodynamic forces, the error of system parameters, and unknown environmental disturbance, a sliding mode control scheme based on disturbance observer is designed for ideal control input. An optimal torque allocation strategy is designed in order to generate the torque determined by the aforementioned nonlinear control law by moving the masses and commanding the magnetotorquer, and therefore combining the subspace of two actuators. Finally, a semi-physical simulation platform was built for two actuators and the results indicate that, additional inertia torque, related to the mass acceleration, is the main disturbance torque during the attitude maneuver and can be significantly reduced by optimal torque decomposition strategy. Meanwhile, during the attitude maintenance, the disturbance observer can effectively observe the system disturbances and improve the attitude control accuracy. The error of attitude angle is less than $\pm $0.1$^\circ$. The results verify the feasibility of the use of the moving mass actuator to actively control the aerodynamic torque.