Analytical Estimation and Experimental Validation of Acceleration at Spacecraft Solar Array Latch-up Considering Differential Latching

Solar array deployment on-board a spacecraft is a mission critical activity. The torque margin over friction torque in the deployment hinge mechanism is an important parameter that ensures positive deployment. Higher torque margin results in higher latch-up moment that affects the design of panel substrates and hinges. Latch-up moment estimated earlier assumes simultaneous latching of hinges. However, in reality, hinges latch at different instants of time due to the presence of closed control loops which has been confirmed by on-orbit observations. The latch-up sequence of the panels influences the distribution of latch-up moments induced at the hinges. In this paper, a mathematical model for a solar array with a yoke two panel configuration is developed using ADAMS software which considers differential latching of hinges. The mathematical model includes the influence of panel flexibility, close control loops, harness, snubbers, ejectors and air drag. The acceleration and moment at latch-up are estimated as a function of time. Deployment time and acceleration at latch-up dictate the magnitude of latch-up moment at hinges and have been compared with experimental value obtained from accelerometer mounted on the outermost panel, during ground deployment test. The novelty in the present work is the formulation of a test validated methodology to analyse the deployment dynamics of a multi-panel solar array considering differential latching which has the potential to simulate the latching of solar array more accurately compared to previously adopted methods. It will be useful in carrying out deployment dynamics of futuristic solar arrays with larger number of panels.