一种高超声速二元混压式进气道的研究

 针对飞行马赫数为6.00的二元进气道模型开展了高焓脉冲风洞试验研究，分析了进气道在不设置反压和设置反压两种情况下的激波结构、内通道皮托压分布及隔离段出口的性能，并结合数值仿真分析了通道内的流场特性。研究结果表明:在无反压情况下，进气道内通道激波反射明显，靠近下壁面的皮托压值均低于其他测点，在隔离段出口截面，靠近侧壁皮托压有所降低；在一定反压条件下，结尾激波系上传至隔离段内，结尾激波位置不对称；堵塞度为62%的反压条件下，结尾激波系位于喉道位置，隔离段出口截面下半部分已经是亚声速流动；在来流马赫数Ma=6.07，迎角α=4.5°无反压情况下，隔离段出口总压恢复系数为0.477，平均马赫数为2.72，增压比为44，流量系数为0.81,表明进气道性能良好。

Investigation of a Two-dimensional Mixed-compression Hypersonic Inlet

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.