Large flexible appendages of spacecraft are prone to prolonged vibration under external excitation, which degrades on-orbit attitude control precision and stability. Model reduction of the finite element model for flexible appendages is essential for efficient dynamic and control analysis of spacecraft with complex configurations. The interfacial dynamic stiffness matrix for flexible appendages, with statically determinate interfaces and weak damping, is investigated to establish a symmetric order-reduced model through theoretical derivation. A demonstration case study of a spring–mass system with multiple degrees of freedom is performed to illustrate the model reduction process and to verify the accuracy of the proposed method. A finite element model of a full-scale solar array is further presented to verify the practical feasibility of this model reduction approach in real engineering applications.
Wang et al. (Sun,) studied this question.