带进气道的隔离段流场实验研究与数值模拟

Experimental and numerical investigation of isolator combined with hypersonic inlet

关键词:  非均匀流  隔离器  冲压喷气发动机  超音速燃烧  内流空气动力学

Key words:  Non uniform flow  Isolator  Ramjet engine  Supersonic combustion  Internal aerodynamics

摘要:

通过风洞实验与数值模拟研究了超燃冲压发动机带进气道的隔离段流动。结合RNGk ε模型隐式求解了三维N S方程,并将计算结果与实验结果进行了比较。研究发现,RNGk ε模型能较好地模拟出超声速流动的激波与分离现象;进气道喉道流动的非均匀性使隔离段内激波/附面层干扰流场与均匀来流条件下的流场有显著差别。风洞实验表明,同一个模型,风洞马赫数为3 85试验的隔离段入口压力畸变大于马赫5 3;但前者隔离段出口截面压力分布比后者更均匀;隔离段入口畸变度大,隔离段实际能达到的压升就低。研究表明隔离段内的总压损失在整个进气道 隔离段组合体总压损失中占了相当大的比重。

Abstract:

The experimental and numerical investigation of flow in isolator combined with hypersonic inlet was made. The isolator was designed by using Waltrup’s empirical theory. Results from a numerical investigation of a shock wave/turbulent boundary-layer interaction in isolator are compared to wall static pressure obtained from the experiment. The turbulence models are the two-equation RNG k-ε model, which is able to capture the major features of the flow with shock/boundary-layer interaction and separation. It is shown that the flow in the isolator with a non-uniform incoming airflow at the entrance is distinctly different from that with an uniform incoming airflow. The Mach numbers of wind tunnel were 5.3 and 3.85 respectively. For the same model, inlet distortion of isolator was more serious and outlet pressure contour of isolator was more uniform in the condition of lower Mach number. The isolator maximum back pressure capability is decreased with increasing inlet distortion. The total pressure loss in isolator is a quite large part of total pressure loss in inlet-isolator.