Thomas Corke

Separation Control on HIgh Angle of Attack Airfoil Using Plasma Actuators

This work involves the documentation and control of leading-edge flow separation that occurs over an airfoil at high angles of attack, well above stall. A generic airfoil shape (NACA 663-018) was used because of its documented leading-edge stall characteristics. It was instrumented for surface-pressure measurements that were used to calculate lift coefficients. Mean-velocity profiles downstream of the airfoil were used to determine the drag coefficient. In addition to these, smoke streakline flow visualization was used to document the state of flow separation. The airfoil was operated over a range of freestream speeds from 10 to 30 m/s, giving chord Reynolds numbers from 77 × 10 3 to 333 × 10 3 .T wo types of plasma actuator designs were investigated. The first produced a spanwise array of streamwise vortices. The second produced a two-dimensional jet in the flow direction along the surface of the airfoil. The plasma actuators were found to lead to reattachment for angles of attack that were 8 deg past the stall angle (the highest investigated). This was accompanied by a full pressure recovery and up to a 400% increase in the lift-to-drag ratio.

Mechanisms and Responses of a Dielectric Barrier Plasma Actuator: Geometric Effects

The single dielectric barrier discharge plasma, a plasma sustainable at atmospheric pressure, has shown considerable promise as a flow control device operating at modest (tens of watts) power levels. Measurements are presented of the development of the plasma during the course of the discharge cycle, and the relevance of these measurements to the modeling of the actuators electrical properties is discussed. Experimental evidence is presented strongly pointing to the electric field enhancement near the leading edge of the actuator as a dominant factor determining the effectiveness of momentum coupling into the surrounding air

Mechanisms and Responses of a Single Dielectric Barrier Plasma Actuator: Plasma Morphology

We present simultaneous optical, electrical, and thrust measurements of an aerodynamic plasma actuator. These measurements indicate that the plasma actuator is a form of the dielectric barrier discharge, whose behavior is governed primarily by the buildup of charge on the dielectric-encapsulated electrode. Our measurements reveal the temporal and macroscale spatial structure of the plasma. Correlating the morphology of the plasma and the electrical characteristics of the discharge to the actuator performance as measured by the thrust produced indicates a direct coupling between the interelectrode electric field (strongly modified by the presence of the plasma) and the charges in the plasma. Our measurements discount bulk heating or asymmetries in the structure of the discharge as mechanisms for the production of bulk motion of the surrounding neutral air, although such asymmetries clearly exist and impact the effectiveness of the actuator.

Optimization of Dielectric Barrier Discharge Plasma Actuators for Active Aerodynamic Flow Control

This paper presents the results of a parametric experimental investigation aimed at optimizing the body force produced by single dielectric barrier discharge plasma actuators used for aerodynamic flow control. A primary goal of the study is the improvement of actuator authority for flow control applications at higher Reynolds number than previously possible. The study examines the effects of dielectric material and thickness, applied voltage amplitude and frequency, voltage waveform, exposed electrode geometry, covered electrode width, and multiple actuator arrays. The metric used to evaluate the performance of the actuator in each case is the measured actuator-induced thrust which is proportional to the total body force. It is demonstrated that actuators constructed with thick dielectric material of low dielectric constant produce a body force that is an order of magnitude larger than that obtained by the Kapton-based actuators used in many previous plasma flow control studies. These actuators allow operation at much higher applied voltages without the formation of discrete streamers which lead to body force saturation.

PLASMA FLAPS AND SLATS: AN APPLICATION OF WEAKLY-IONIZED PLASMA ACTUATORS

The experimental validation of an application of weakly-ionized plasma actuators for improved aerodynamic performance of multi-element wings and wings with movable control surfaces is presented. Two spanwise arrays of plasma actuators, configured to produce a wall-jet effect, were applied on the suction surface of a two-dimensional NACA 0015 wing model, one at the leading edge and the other near the trailing edge to mimic the effects of a wing leading-edge slat and a trailing-edge flap, respectively. Flow control tests were conducted at chord Reynolds numbers, corrected for blockage, of 0.217 x 10 6 and 0.307 x 10 6 in a low-speed wind tunnel at the University of Notre Dame. The leading-edge-separation control resulted in an increase in both the maximum lift coefficient and the stall angle of attack and a lift-to-drag improvement of as much as 340%. An optimum frequency was found to exist for unsteady excitation of the leading-edge separation. Under this condition, the power to the actuator was estimated to be only 2 W. The trailing-edge actuator was found to produce the same effect as a plain trailing-edge flap. This included a uniform shift at all angles of attack of the lift coefficient and a shift toward higher lift coefficients of the drag bucket. In addition, there was a slight decrease in the minimum drag coefficient. The obvious advantages of this approach are its simplicity, as there are no moving parts, and its lack of hinge gaps, which add drag. An example of their use as ailerons for roll control produces a comparable roll moment coefficient to a sample general aviation aircraft.

Separation control using plasma actuators : Dynamic stall vortex control on oscillating airfoil

A plasma actuator was used to control leading-edge flow separation and dynamic stall vortex on a periodically oscillated NACA 0015 airfoil. The effectiveness of the actuator was documented through phase-conditioned surface pressure measurements and smoke flow visualization records. The airfoil was driven in a periodic cycle corresponding to α = 15 deg+10deg sinwt. The results presented here are for a reduced frequency of k = ωc/2U ∞ = 0.08. Three cases of control with the plasma actuator were investigated: open-loop control with steady plasma actuation, open-loop control with unsteady plasma actuation, and closed-loop control with steady plasma actuation. For closed-loop control, the actuator was activated in selected portions of the oscillatory cycle based on angle-of-attack feedback. All of the cases investigated exhibited an increase in cycle-integrated lift with improvements in the lift-cycle hysteresis. In two cases, the pitch-moment stall angle was delayed and in one of these, the adverse negative moment peak was significantly reduced.

Plasma Actuators for Cylinder Flow Control and Noise Reduction

In this paper, the results of flow-control experiments using single dielectric barrier discharge plasma actuators to control flow separation and unsteady vortex shedding from a circular cylinder in crossflow are reported. This work is motivated by the need to reduce landing gear noise for commercial transport aircraft via an effective streamlining created by the actuators. The experiments are performed at Re D = 3.3 x 104. Using either steady or unsteady actuation, Karman shedding is totally eliminated, turbulence levels in the wake decrease significantly, and near-field sound pressure levels associated with shedding are reduced by 13.3 dB. In the case of unsteady plasma actuation, an actuation frequency of St D = 1 is found to be most effective. The unsteady actuation has the advantage that total suppression of shedding is achieved for a duty cycle of only 25%. However, because unsteady actuation is associated with an unsteady body force and produces a tone at the actuation frequency, steady actuation is more suitable for noise-control applications.

Scaling Effects of an Aerodynamic Plasma Actuator

We present experimental results to yield insight into the scalability and control effectiveness of single-dielectricbarrier-discharge plasma actuators for leading-edge separation control on airfoils. The parameters investigated are chord Reynolds number, Mach number, leading-edge radius, actuator amplitude, and unsteady frequency. This includes chord Reynolds numbers up to 1:0 � 106 and a maximum freestream speed of 60 m=s corresponding to a Mach number of 0.176. The main objective of this work is to examine the voltage requirements for the plasma actuators to reattach the flow at the leading edge of airfoils at poststall angles of attack for a range of flow parameters in order to establish scaling between laboratory and full-flight conditions. For the full range of conditions, an optimum unsteady actuator frequency f is found to minimize the actuator voltage needed to reattach the flow, such that F� � fLsep=U1 � 1. At the optimum frequencies, the minimum voltage required to reattach the flow is weakly dependent on chord Reynolds number and strongly dependent on the poststall angle of attack and leading-edge radius. The results indicate that the voltage required to reattach the flow scales as the square of the leading-edge radius.

Plasma actuators for hingeless aerodynamic control of an unmanned air vehicle

The use of dielectric barrier discharge plasma actuators for hingeless flow control over a 47-deg 1303 unmanned combat air vehicle wing is described. Control was implemented at the wing leading edge to provide longitudinal control without the use of hinged control surfaces. Wind-tunnel tests were conducted at a chord Reynolds number of 4.12 x 105 and angles of attack ranging from 15 to 35 deg to evaluate the performance of leading-edge plasma actuators for hingeless flow control. Operated in an unsteady mode, the actuators were used to alter the flowfield over the lee-side wing to modify the aerodynamic lift and drag forces on the vehicle. Multiple configurations of the plasma actuator were tested on the lee side and wind side of the wing leading edge to affect the wing aerodynamics. Data acquisition included force-balance measurements, laser fluorescence, and surface flow visualizations. Flow visualization tests mainly focused on understanding the vortex phenomena over the baseline uncontrolled wing to aid in identifying optimal locations for plasma actuators for effective flow manipulation. Force-balance results show considerable changes in the lift and drag characteristics of the wing for the plasma-controlled cases compared with the baseline cases. When compared with the conventional traditional trailing-edge devices, the plasma actuators demonstrate a significant improvement in the control authority in the 15- to 35-deg angle-of-attack range, thereby extending the operational flight envelope of the wing. The study demonstrates the technical feasibility of a plasma wing concept for hingeless flight control of air vehicles, in particular, vehicles with highly swept wings and at high angles of attack flight conditions in which conventional flaps and ailerons are ineffective.

Autonomous Sensing and Control of Wing Stall Using a Smart Plasma Slat

DOI: 10.2514/1.24057 The concept of a self-governing smart plasma slat for active sense and control of flow separation and incipient wing stall is presented. The smart plasma slat design involves the use of an aerodynamic plasma actuator on the leading edge of a two-dimensional NACA 0015 airfoil in a manner that mimics the effect of a movable leading-edge slat of a conventional high-lift system. The self-governing system uses a single high-bandwidth pressure sensor and a feedback controller to operate the actuator in an autonomous mode with a primary function to sense and control incipient flow separation at the wing leading edge and to delay incipient stall. Two feedback control techniques are investigated. Wind tunnel experiments demonstrate that the aerodynamic effects of a smart actuator are consistent with the previously tested open-loop actuator, in that stall hysteresis is eliminated, stall angle is delayed by 7 deg, and a significant improvement in the lift-to-drag ratio is achieved over a wide range of angles of attack. These feedback control approaches provide a means to further reduce power requirements for an unsteady plasma actuator for practical air vehicle applications. The smart plasma slat concept is well suited for the design of low-drag, quiet, highlift systems for fixed-wing aircraft and rotorcraft.