Abstract: Boundary layer transition plays a significant role in the prediction of aerodynamic and aerothermodynamics characteristics of hypersonic vehicle. The commonly used correlation based on γ-Re_θ transition model is implemented into a Reynolds-Averaged Navier-Stokes solver in order to evaluate its capability in hypersonic flows. The mean flow and turbulent flow equations are solved simultaneously based on the LU-SGS method. Three compressibility correction methods are adopted to simulate the hypersonic flows around a flat plate, a double wedge and a cone body. Results of Stanton number, heat flux on walls, contours of turbulence kinetic energy and intermittency factor are presented. The results of γ-Re_θ transition model show better flow essential than that of full laminar or full turbulent results. The transition onset positions obtained by different compressibility methods under the same free stream turbulence intensity varied significantly. Compressibility correction based on local Mach number delays the transition onset greatly, while the correction of source term of turbulent kinetic energy equation improves the accuracy of heat flux on turbulent section of the wall. Rational selection of compressibility correction method should rely on the type of flow. Elaborate and reliable compressibility correction method still need to be further investigated.