Experimental Investigation of Active Flow Control for Changing Stall Angle of a NACA0012 Airfoil, Using Plasma-Actuator

Active flow control has great importance in design of flying objects in general and control flaps in particular. Using dielectric barrier discharge as plasma actuators over control flaps can cause a change in position of separation point or can eliminate it. This new flow control device produces a body force that causes generating a micro wall jet. The present study aims to investigate experimentally the effects of the plasma actuator over a NACA0012 airfoil in high angle of attacks (up to 6 degrees above the stall) in order to displace the separation point. Therefore, the effects of the plasma actuator with sine wave unsteady excitation and different duty cycles on airfoil pressure coefficient distribution and lift coefficient were studied. The experiments were performed at  = 3.7 × 10. The plasma generation was performed, using a high voltage AC power supply at voltages up to 14. The results indicate that lift coefficient increases with unsteady plasma actuation at post stall angles of attack up to 6 degree. The improvement of the pressure coefficient distribution in upper side of the airfoil for lower duty cycles about 14% is better than higher duty cycles above 50%.