Dynamics and control of spacecraft solar array deployment considering physical contacts between locking mechanisms

Abstract

The numerical simulation of the deployment and locking process of flexible appendages is an important topic in the overall design of spacecraft. In the existing research, the connecting mechanism between the solar panels is usually regarded as an ideal revolute constraint, and the locking process is simulated by applying a virtual lock torque related to the deployment angle. However, this method cannot obtain the physical contact force history, and because the physical meaning of the stiffness term in the lock torque function is not clear, it is difficult to accurately illustrate the frequency characteristics of the flexural vibration of the panels in the post-lock phase. In fact, the locking process is a frequent contact-impact process between the lock pin and groove. In this study, a typical deployable spacecraft with torsion-spring-driven panels were investigated numerically. The locking mechanism is modeled as a physical entity, and the continuous contact force model between the lock pin and groove is established. The contact stiffness of cylindrical line-contact is given by preloading calculation using the finite element method. Based on the multibody dynamics, rigid body model and flexible body model of the solar panels are established respectively. Through the comparison between the numerical simulation results of the two models, it is found that the rigid body model will produce excessive contact force and false high-frequency vibration. Furthermore, the attitude PD controller of spacecraft is designed, and the difference of system dynamic response in the deployment process under the main-body free state and the main-body controlled state is studied. The vibration frequency under the main-body controlled state is close to the first-order natural frequency of the structural finite element analysis, which verifies the rationality of the calculation results of the model.