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We numerically examine the mechanisms that describe the shock-boundary layerinteractions in transonic flow past an oscillating wing section. At moderateand high angles of incidence but low amplitudes of oscillation, shock inducedflow separation or shock-stall is observed accompanied by shock reversal. Eventhough the power input to the airfoil by the viscous forces is three orders ofmagnitude lower than that due to the pressure forces on the airfoil, theboundary layer manipulates the shock location and shock motion andredistributes the power input to the airfoil by the pressure forces. The shockmotion is reversed relative to that in an inviscid flow as the boundary layercannot sustain an adverse pressure gradient posed by the shock, causing theshock to move upstream leading to an early separation. The shock motion shows aphase difference with reference to the airfoil motion and is a function of thefrequency of the oscillation. At low angles of incidence, and low amplitudes ofoscillation, the boundary layer changes the profile presented to the externalflow, leads to a slower expansion of the flow resulting in an early shock, anda diffused shock-foot caused by the boundary layer.

Viscosity, shock-induced separation, and shock reversal -- Oscillating wing section in transonic flow