高超声速风洞马赫数4.5喷管气动设计和数值验证

Aerodynamic design and numerical validation of Mach 4.5 nozzle in hypersonic wind tunnel

常规跨超声速风洞进行马赫数4.5试验时,常常伴有空气液化现象,造成试验数据可信度低,在高超声速风洞研制马赫数4.5喷管,具有对气流加热的能力,可以提供更加准确的试验数据。目前国内0.5 m量级高超声速风洞还不具备马赫数4.5的试验能力。通过无黏流计算方法计算轴对称喷管型面,并采用Sivells-Payne方法进行附面层修正,然后进行数值验证,证明了计算出的型面满足国军标对马赫数的设计要求,可以投入加工生产。

The Mach 4.5 test in a conventional tran-supersonic wind tunnel is often accompanied by the phenomenon of air liquefaction,resulting in low reliability of the test data.The Mach 4.5 nozzle developed in a hypersonic wind tunnel has the ability to heat the airflow,which can provide more accurate test data.At present,the test capability of Mach 4.5 is not available in China for the 0.5-meter hypersonic wind tunnel.The axisymmetric nozzle profile was calculated by the inviscid flow calculation method,and the boundary layer was modified by the Sivells-Payne method.Then,numerical simulation was carried out.It was proved that the calculated profile met the GJB design requirements of Mach number and can be put into production.