Experimental Study of Plasma Flow Control on Highly Swept Delta Wing

Abstract

A ERODYNAMIC flow control based on plasma actuators is now in expansion, because plasma actuators are fully electronic with nomoving parts; they have an extremely fast response, very low mass, low input power, and the easy ability to simulate their effect in numerical flow solvers [1]. In particular, they areflexible, so that they can be formed to various shapes and located on the air vehicles with relative ease. There are no other known actuators that have such flexibility [2]. There are many different plasma actuators, including dielectric barrier discharge (DBD), direct current glow discharge, radio frequency glow discharge, and filamentary arc discharges. Suchomel et al. [3] provided an overview of different plasma generation technologies currently under investigation for aeronautical applications. The actuator used here is based on surface DBD. This new discharge was invented by Roth et al. and protected by a U.S. patent since 1995 [4]. The surface plasma has considerably influenced research on airflow control by plasmas, because the simplicity of its use allowed many researchers in aerodynamics to work on this subject (without necessarily being specialists in plasma generation [5]). Velocity measurements indicate that theprimary result of the averagedplasmainduced body forces is the formation of a wall jet that imparts momentum to the fluid [1]. The plasma actuators for aerodynamic flow control can be applied for different purposes. Examples include boundary layer control [6], lift augmentation and separation control for airfoils [7,8], and control of the dynamic stall vortex on oscillating airfoils [9]. More recently, the plasma actuator has been demonstrated on application in three-dimensional vortical flow controls on delta wings and unmanned aerial vehicle (UAVs) [2,10–12]. Patel et al. [10] used the DBD plasma actuators for hingeless flow control over the 1303 unmanned combat air vehicle wing. Control was implemented at the wing leading edge to provide longitudinal control without the hinged control surfaces. Force balance results showed considerable changes in the lift characteristics of thewing for the plasma-controlled cases, when comparedwith the baseline cases. Compared with the conventional traditional trailing-edge devices, the plasma actuators were demonstrated to have a significant improvement in the control authority in the 15 to 35 deg angle-ofattack range, thereby extending the operational flight envelope of the wing. In addition to the lift modification, Nelson et al. [11] studied the plasma actuator to provide roll control at high angles of attack on a scaled 1303 UAV configuration. It was found to have excellent roll control capability, which was very responsive. Greenblatt et al. [2] investigated the DBD plasma actuator active control of a leadingedge vortex on a semispan delta wing at typical micro aerial vehicle Reynolds numbers. The plasma actuator produced a plasma-induced jet inward from the leading edge. The maximum CL ( 36 deg) increased by 0.2 in the poststall region at the optimum reduced frequency F 1 1. Visbal and Gaitonde [12] deployed asymmetric DBD plasma actuators on the apex of a 75 deg swept delta wing to control the vortical flows by numerical simulation. The actuator was placed at x=c 0:08 and extended in a spanwise direction. The strength of the plasma actuator was chosen to be Dc 2400. The changes near the apex of the wing, induced by the plasma, resulted in significant downstream displacements of the vortex breakdown location. This effect could be potentially beneficial for roll control authority at high angles of attack. Following actuation, the unsteady shear layer evolved into a steady pattern, characterized by stationary helical subvortices. The origin of these steady substructures appeared to be linked to the increase in axial velocity within the secondary vortex, induced by the momentum injection of the actuator. In this technical note, the configurations of the plasma actuator on the delta wing, similar to that used in Visbal and Gaitonde’s simulation [12], are used. The plasma actuators are mounted in different chordwise locations on the leeward side of the delta wing. The effect of the location of the plasma actuator on the aerodynamic performance of the delta wing is studied by the balanced force measurement in the wind tunnel. At last, the smokewire visualization is used to present the flow structure variation induced by the plasma actuator.