![]() ![]() Supercritical wings add a graceful appearance to the modified NASA F-8 test aircraft. The different separation structure would result in different pressure coefficient distribution and boundary layer thickness.An airfoil considered unconventional when tested in the early 1970s by NASA at the Dryden Flight Research Center is now universally recognized by the aviation industry as a wing design that increases flying efficiency and helps lower fuel costs. It is shown that with the attack angle increases, the separation bubble occurred on the upper surface first, then the trailing-edge separation occurred, the trailing-edge would separate totally at last. The computation attack angles of CH airfoil varied from 0oto 4o, Mach numbers varied from 0.74 to 0.82 while Reynolds numbers varied from 3×10 6 to 50×10 6 per airfoil chord. The two-dimensional Navier-Stokes equations were solved with structure grids by utilizing the S-A turbulence model. The transonic flows over a typical supercritical airfoil CH were numerically investigated in this paper, in order to analyses different shock-induced separation structure. But at transonic speeds, the complicated shock-induced separation on the upper surface of supercritical airfoil will change the aerodynamic characteristics. The shock wave location and intensity were affected by the three factors, and the boundary layer thickness was mainly affected by Reynolds number, while the separation structure was mainly determined by the attack angle and Mach number.Ībstract: The supercritical airfoil has been widely applied to large airplanes for sake of high aerodynamic efficiency. It was shown that the shock-induced separation was affected by attack angles, Mach numbers and Reynolds numbers, but the influence tendency and areas were quite different. The computation attack angles of CH airfoil varied from 0oto 7o, Reynolds numbers varied from 5×10 6 to 50×10 6 per airfoil chord while Mach number varied from 0.74 to 0.82. The Navier-Stokes equations were solved, in order to investigate influence of different attack angles, Mach numbers and Reynolds numbers. In this paper, the influencing factors of shock-induced separation for supercritical airfoil CH was analyzed at transonic speeds. The problem of the shock-induced separation was not solved completely for the complicated phenomena and flow mechanism. 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.Ībstract: The shock-induced separation easily occurred on the upper surface of supercritical airfoil at transonic speeds, which would change the aerodynamic characteristics. It is obvious that shock location moves afterward and shock intensity strengthens as Reynolds number increasing. ![]() 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 3×10 6 to 50×10 6 per airfoil chord. 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 shock characteristics such as location and intensity are sensitive to Reynolds number. But at transonic speeds, the shock wave on upper surface of supercritical airfoil may induce boundary layer separation, which would change the aerodynamic characteristics. Abstract: The supercritical airfoil has been widely applied to large airplanes for sake of high aerodynamic efficiency. ![]()
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