The supercritical airfoil has been widely applied to large airplanes for sake of high aerodynamic efficiency. But at transonic speeds, the shock wave on upper surface of supercritical airfoil may induce boundary layer separation, which would change the aerodynamic characteristics. The shock characteristics such as location and intensity are sensitive to Reynolds number. In order to predict aerodynamic characteristics of supercritical airfoil exactly, the Reynolds number effects of shock wave must be investigated.The transonic flows over a typical supercritical airfoil CH were numerically simulated with two-dimensional Navier-Stokes equations, and the numerical method was validated with test results in ETW(European Transonic Windtunnel). The computation attack angles of CH airfoil varied from 0oto 8o, Mach numbers varied from 0.74 to 0.82 while Reynolds numbers varied from 3×106 to 50×106 per airfoil chord. It is obvious that shock location moves afterward and shock intensity strengthens as Reynolds number increasing. The similar curves of shock location and intensity is linear with logarithm of Reynolds number, so that the shock location and intensity at flight condition could be extrapolated from low Reynolds number.
Numerical Investigation on the Reynolds Number Effects of Supercritical Airfoil
