Deployment Dynamics for a Flexible Solar Array Composed of Composite-Laminated Plates

Modern spacecraft often deploy large-scale and lightweight solar arrays consisting of composite-laminated plates because of their high reliability, superior mechanical properties, and low manufacturing cost. This paper presents the deployment dynamics of a flexible solar array composed of composite-laminated plates undergoing large rotation and large deformation motions. The subject is a constrained rigid–flexible coupling spacecraft consisting of a rigid main body and a flexible solar array composed of composite-laminated plates. The rigid main body and the flexible solar array are described by the natural coordinate formulation and the absolute nodal coordinate formulation, respectively. A constitutive model of a laminated shell formed of fiber-reinforced composite material is established, and the corresponding generalized elastic force is derived. Thus, the spacecraft’s equations of motion are derived as a set of differential algebraic equations and a computational scheme based on the Hilber-Hughes-Taylor (HHT)-I3 method is illustrated. A series of numerical calculations and simulations are conducted to investigate the solar array deployment dynamics. The results show that the solar panel flexibility and fiber orientation of fiber-reinforced composite materials have obvious effects on spacecraft dynamic responses during solar array deployment.