Visualization Techniques Research In Short Duration Tunnels For Shock Wave- Boundary Layer Interaction On 2-D Hypersonic Inlet Forebody

In the airbreathing hypersonic vehicles research, shock wave-boundary layer interaction (SWBLI) problems have been investigated for a long time but are far from solved nowadays. The shock wave structure, boundary layer separation and transition process on the external compression surfaces of the hypersonic inlet are influenced by SWBLI, and then the integration performances of inlet could be seriously affected. Researches on SWBLI are abundant, but few relating to hypersonic inlet external compression surfaces. The research on the effect of the multi-compression corners and the leading edge radius (nominated SR) are especially not systemic and comprehensive.The short duration facilities, which can simulate the total temperature and pressure at high Mach number, are important complementarities for the conventional wind tunnels. The combination of the two kinds of wind tunnels can provide larger range of Mach number, Re number and wall temperature ratio, therefor more aboundant data on SWBLI can be obtained with high Mach number test. In the SWBLI research, both shock waves and boundary layer behavior should be investigated, and measurement techniques are very important to the investigation. However, in hypersonic situation, the techiniques for getting the thickness and separation state of boundary layer are quite infrequent, as the temperature and speed of the flow are quite high, and the test time is very short on the short duration facilities.Two kinds of test techniques are developed on the short duration facilities for the study of SWBLI in hypersonic inlet. Firtly, hypersonic density boundary layer thickness measure technique by focusing schlieren system is developed in Shock Wave Tunnel. To do this work, the lamp-house system is improved, and the influence of the focusing schlieren system components on focus thickness is experimented. The relationship between density boundary layer thickness, velocity boundary layer thickness, and temperature boundary layer thickness is analysed, it is proved that the density grads boundary layer thickness is equal to density boundary layer thickness; the density boundary layer thickness is the same as the temperature boundary layer thickness, and velocity boundary layer thickness should be less than the density boundary layer thickness.and the method to get boundary layer thickness based on density from the focusing schlieren photos is put forword. Secondly, the oil flow visualization technique is developed in Pulse Combustion Wind Tunnel to show the boundary layer separation state. The difficulties for appling the thchnique in Pulse Combustion Wind Tunnel are analysed, based on the running characteristic of the tunnel. The forces acting on the oil film are analyzed according to the flow field condition, and the influence factors of the oil film velocity are deduced, then a guide line is provided for film type oil flow technique, through which the oil viscosity and film thickness can be selected. The oil film is provided by spray gun, and the thickness of the film is unity. The problem of high temperature test gas shine is discussed, and real time high speed vidicon system is designed. The results show that the film type oil flow patterns can be recorded by a high speed camera, with a LED lamp. The oil flow patterns are obtained in very short valid test time, displaying the separation state of compression ramps.The character of SWBLI on the multi-compression corners and the leading edge radius affection are investigated in Shock Wave Tunnel,under the conditions of Mach number 6 and Renault number 3.4×107/m The wall pressure distribution, heat flux distribution, the boundary layer thickness distribution and shock wave structure are obtained and the character of SWBLI is displayed correctly. The value of pressure increases gradually from the upstream of the corner, reaches to a pressure platform downstream the corner. With the increasing of the leading edge radius, the value of pressure platform decreases and the distance needed to arrive the platform increases, which means that the SWBLI region is extended. The heat flux value increases gradually from the upstream of the corner, reaches to a local maximum and then decreases in the downstream compression surface. With increasing of the leading edge radius, the heat flux of local maximum at the corner decreases obviously, and the peak heat flux position moves forward. The boundary layer thickness increases firstly to a local maximum at upstream of the corner, and then decreases across the interaction region, then it increases again in the next compression surface. The results show that the wall static pressure from CFD by three different viscous models are similar, all of which are coincident. with experiment results, but the wall heat flux and boundary layer thickness is different by different viscous models. With increasing of the leading edge radius, the first shock wave angle increases, while the second and third decreases, the shock wave root bends more seriously and the interaction region extends.