Investigation Of Aerothermal Environment Of Hypersonic Compression Corners

Shock wave/boundary layer interaction (SBLI) and shock/shock interaction areclassic phenomena in hypersonic flight vehicles that have compression corners in shapedesign. Interactions may lead to extraordinarily high heat transfer rate which can be tensof times of situations without interactions. This comes to be a severe problemconfronted by thermal protection design. As far as fligt vehicle design of aeronauticsand aerospace, it is thus of great significance to study the flow mechanism ofcompression corner, especially the aerothermal characteristics, and summarizecorresponding laws, and also provide usable heat transfer correlations for the designstage.Experimental and numerical methods are used to investigate the flow field of a12degree two-dimensional compression corner model. There are9kinds of free-streamconditions with nominal Mach number ranging from6to12and unit Reynolds numberranging from1.1×106/m to4.4×107/m.The influence of free-stream unit Reynolds number on aerothermal environmentare studied. Heat transfer rate is measured by thin film gauge, and whether the boundarylayer is laminar, transitional or turbulent is also obtained by the measurement. The400mm-diameter schileren system is used to get the schileren graph which contains thestructure of shock wave and boundary layer, and the separation region is captured.Additionally, pressure sensors with high natural frequency are utilized to measure theaverage and fluctuating pressure on compression surface. Results indicate that Reynoldsnumber have apparent effect on peak heating rate and transition process. HigherReynolds number leads to higher nondimensional peak heating rate, but imposes weakeffect on the location of peak.The experimental results also show that two local peak heating rate occurs whentransitional process initiates just downstream of the reattachment region. The relativelyupstream local peak heating is generated by the reattachment process, while thedownstream local peak heaing results in boudary layer transition. This is quite differentfrom that of situations of fully laminar intractions or turbulent interactiosn with singlepeak heating. The heat transfer results obtained under such experimental status iscomplementary for the “database” of the aerothermal environment of hypersoniccompression corner.By using laminar Navier-Stokes equations and Shear Stress Transport model fornumerical simulation, heat transfer, pressure and skin friction distributions are obtained.Basically, heat transfer results show certain discrepancy with experimental results and laminar simulation achieves the best agreement with experiment relatively.Finally, on the base of experimental results and former researchers’ achievement,the two-dimensional compression corner flow is classfied into seven different typesaccording to which the boundary layer status and whether flow separation occurs. Andfour types are included for investigation of experiment. Assisted by parameters of theinteraction region obtained by experimental and numerical tools, the correlation law forpredicting peak heating rate presented by Simeonides is assessed. Results show that thecorrelation law plays well under overall flow conditions selected for experimentspresented in this paper.