关键词: 航空航天推进系统, 吸气式高超声速飞行器, 二元进气道, 风洞试验, 数值模拟

Abstract: The performance of a fixedgeometry twodimensional mixedcompression hypersonic inlet is investigated at Mach number 600 in the highenthalpy impulse wind tunnel. The investigation focuses on the structure of the shocks, the distribution of Pitotpressure ratio at the duct and the performance of the inlet at the exit of the isolator. Characteristics of the internal flow are analyzed in combination with numerical simulation. Results indicate: (1) With no pressure imposed by the mass flow meter, there is a distinct reflected shock structure in the isolator, and the Pitotpressure of the probe at the exit section of the isolator near the ramp side is lower than that of all other probes. Meanshile, the Pitotpressure of the rake near the sidewall is lower than others. (2) The terminal shock is propagated in the isolator when the back pressure ratio increases to a certain value. The terminal shock is asymmetric in the isolator. When the throttling ratio increases to 62%, the terminal shock is located at the section of the throat and the Pitotpressure profiles at the exit of the isolator show that the flow speed for at least a half region in the isolator height is already subsonic. (3) At Ma=6.07，α=4.50° and with no back pressure, the total pressure recovery coefficient and the mass flow ratio are 0.477 and 0.81 respectively at the exit of the isolator, while the average Mach number is 2.72 and the pressure ratio between the flow at the exit of the isolator and free stream is 44.

Key words: aerospace propulsion system, airbreathing hypersonic vehicle, twodimensional inlet, wind tunnel test, numerical simulation