Abstract:

To investigate the working capability of liquid rocket engines in the high temperature environment on the lunar surface， the high-altitude hot fire simulation test of a 120 N bipropellant rocket engine using hypergolic propellants is conducted. The thermal control assembly and the proportion integration derivative control device are used to control the heating and holding temperature of the propellant supply lines and solenoid valve housing. The effects of the holding temperature at normal atmospheric temperature and 80–135 ℃ on the steady-state and pulse working performance of the engine ignition are investigated. The feasibility of discharging the vaporized propellant through the long-pulse ignition procedure after holding is verified. The test results show that the engine can operate in both the steady state and the pulsed mode under all high temperature conditions， with comparable steady-state working performance after stabilization. The start-up response time based on thrust measurement is significantly longer than that under normal atmospheric temperature conditions， and returns to normal level at shutdown. Affected by the combined effect of the propellant density and the nitrogen tetroxide gas-liquid flow state， the first pulse thrust impulse at 30 ms pulse width relative to the normal atmospheric temperature condition is 80%， 66%， 31%， 16% and 17% at holding temperatures of 80 ℃， 90 ℃， 105 ℃， 120 ℃ and 135 ℃， respectively. When the pulse width is smaller or larger than the start-up response time， the deviation value of the thrust impulse compared with that at normal atmospheric temperature gradually decreases with increasing pulse width. The 7 sets of 128 ms pulse width can basically drain the vaporized propellant in the propellant supply lines， and the engine pulse working performance returns to the normal level after the exhaust procedure.

Key words: liquid rocket engine, high temperature environment, thermal control assembly, propellant vaporization, working performance, pulse impulse value