Passive shock wave/boundary layer control for a complete NACA 0012 airfoil, with a porous surface was investigated in the center of the RPI 3 in. x 15.4 in. Transonic Wind Tunnel at various angles of attack. A porous surface with a cavity underneath it was positioned on an area of the airfoil where the shock wave occurs. Schlieren photographs for solid and porous surfaces were taken and the photographs indicated that the normal shock was changed to a lambda shock wave over the porous surface. The static pressure distributions over the airfoil and the wake impact pressure data were used to calculate the flow Mach number over the airfoil and the profile drag of the airfoil, respectively. Over the porous surface the Mach number decreases from a supersonic flow to nearly sonic velocity ahead of the terminating normal shock wave, which minimizes the boundary layer separation; in addition, the lambda shock wave also minimizes the entropy losses through the shock wave. Both of these effects tend to decrease the transonic drag. At a freestream Mach number of 0.82 and angles of attack of 2 deg and 4 deg, introduction of a porous surface resulted in drag reductions of approximately 19 % and 25 %, respectively.
Passive drag reduction on a complete NACA 0012 airfoil at transonic Mach numbers
Passive Verringerung des Luftwiderstands eines kompletten NACA 12-Tragfluegels bei Mach-Zahlen im Ueberschallbereich
AIAA-Papers ; Jul ; 1-10
1987
10 Seiten, 10 Bilder, 13 Quellen
Aufsatz (Konferenz)
Englisch
Viscous Drag Reduction Using Riblets on NACA 0012 Airfoil to Moderate Incidence
Online Contents | 1996
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