Air drag bearing friction and harness bending torque that influence the simulation were included in the theoretical modelling of deployment, and a correlation was obtained with test performance. 2.2.2 Thermo Vacuum Endurance Test. Some of the performance degradations expected in the environment were increased co-efficient of friction and increased harness bending torque at low temperatures, increased friction torque in vacuum, thermal distortion, figure and other effects. A full-scale model of the mechanism with spacecraft simulated side wall was used as the test article and the predicted extreme in- orbit operating conditions were duplicated, as shown in Fig. 4. 2.2.3 Vibration Test. The initial vibration test of the solar array mechanism along with a fully fledged spacecraft revealed a problem of high amplification factor, causing cell damage and mechanical interference with the nearest interfaces. The clampdown latches were then modified with spring dampers and the set-up was vibrated again. A good improvement was indicated when the transmissibility got reduced by 50% which is illustrated in Fig. 5(A) and (5B) and Table I. 2.2.5 Spacecraft Level Tests. A zero-g fixture was developed with a vertical suspension type support for conducting a partial deployment test with the solar array mounted on to the spacecraft. With its own articulate mechanism the fixture enables observation of the release of latches and first motion of array. 2.2.6 Summary of Achieved Mechanism Characteristics. Deployment time = 4.5 Sec Latchup shock = 3-5 “G” Mass —Array substrate = 3.100 Kg each panel —Cell network = 2.400 Kg each panel —Deployment mechanism = 2.675 Kg each wing —Total (2 array wings) = 16.350 Kg
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