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Unsteady Stall Penetration Experiments at High Reynolds Number

机译：高雷诺数下的非定常失速侵彻实验

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An experiment was performed to examine the unsteady aerodynamics of stall penetration at constant pitch rate and high Reynolds number, in an attempt to more accurately model conditions during aircraft post-stall maneuvers and during helicopter high speed forward flight. The model spanned the 8 ft wind tunnel and consisted of a 17.3 in. chord wing with a Sikorsky SSC-AOQ airfoil section. Two forms of pitching motion were used: constant pitch rate ramps and sinusoidal oscillations. Ramp data were obtained for 36 test points at pitch rates between 0.001 and 0.020, Mach numbers between 0.2 and 0.4, and Reynolds numbers between 2 and 4 million. Sinusoidal data were obtained for an additional 9 conditions. The results demonstrate the influence of the leading edge stall vortex on the unsteady aerodynamic response during and after stall. The vortex-related unsteady increments to the lift, drag, and pitching moment increase with pitch rate; the maximum delta C sub L is 1.2 at A =0.02. Angular delays in stall events also increase with pitch rate. Vortex strength and propagation velocity were determined from pressures induced on the airfoil surface. The vortex is strengthened by increasing the pitch rate, and is weakened both by increasing the Mach number and by starting the motion close to the steady-state stall angle. Propogation velocity increases linearly with pitch rate.

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