The Influence of Propellant Feed Temperature on Hydrazine Thruster

• Main Authors: Lin, Chih-Jung, 林志榮
• Other Authors: Tony Yuan
• Format: Others
• Language: zh-TW
• Online Access: http://ndltd.ncl.edu.tw/handle/91151464264126410994

碩士 === 國立成功大學 === 航空太空工程學系 === 85 === Hydrazine/Shell 405 thruster is a simple, steady, and repeatable system.It has been used on attitude control and orbital transfer systems forspacecraft. Because the starting characteristic of the thruster affectsthe effectiveness of thruster''s pulse operation, this thesis researchfocused on the change of the response time of the thruster(ignition delay) by variating the temperature of the feed propellant. The ideal and steady-state performance of the thruster based on the chamber pressure andtemperature measurements are also calculated. Hydrazine was heated by a thermostatically control system to keep thetemperature of the feed propellant at steps of 30, 40, 50, 60℃. Inreaction chamber, a peizo- electric pressure transducer was used to observethedynamic pressures at the starting period. The voltage .vs. time signalfrom the piezo-electric pressure transducer was adopted to define theignition delay and compared to the static chamber pressure traces. Inaddition, a series of experiments with different configurations of thecylindrical catalyst bed were performed to observe the changes of chamberpressure, temperature, and the extent of ammonia dissociation (x). Thesedata were used to estimate the thrust and specific impulse of the thruster.The results showed :(a) With a constant catalytic bed load, increasing the diameter of the bed shortened the ignition delay. The delay time of the 15mm bed diameter was 16% longer than that of the 20mm''s.(b) A higher propellant feed temperature showed a shorter ignition delay. The delay time decreased about 80% at 60℃than that at 30℃.(c) When the propellant temperature was increased from 30℃ to 60℃, the extent of ammonia dissociation (x) was increased about 3∼5%. The values of x in the present experimental configurations were ranged in 0.52∼0.58.(d) Increasing the propellant feed temperature decreased the chamber pressure and temperature slightly, as a result, the calculated thrust and specific impulse were also decreased. The calculated values decreases about 4% at 60℃than that at 30℃.