American Institute of Aeronautics and Astronautics

1

Effects of an Aerodynamic Plasma Actuator on an HSNLF Airfoil

Martiqua L. Post1, Stephen L. Greenwade2, and Matthew H. Yan3 U.S. Air Force Academy, USAF Academy, CO 80840

Thomas C. Corke4 University of Notre Dame, Notre Dame, IN 46556

and

Mehul P. Patel5 Orbital Research Inc., Cleveland, OH 44103

Previous studies have shown that single-dielectric barrier discharge plasma actuators are effective in manipulating the flow around an airfoil. A virtual-wing concept has led to an interest in eliminating movable conventional control surfaces and high-lift devices. In this study the effectiveness of a plasma actuator is tested on a high-speed, natural laminar flow, HSNLF(1)-0213, airfoil. The 10 kVp-p actuator is designed to simulate an aileron-up or trailing-edge flap upward deflection at M=0.1 (Re=292K) and M=0.2 (Re=584K). The tests are performed at various angles of attack from α= -2° to 16°. Results at M=0.2 indicate a 2% increase in CL and up to an 8% increase in CD. The drag polar shows a shift up and to the right, which is supported by a slight decrease in the wake velocities. Findings reveal that inducing an upstream body force (against the freestream) with a plasma actuator does not effectively cause separation, but rather energizes the flow similar to the effects of inducing downstream flow. Hence, the effects of a plasma actuator when used as a trailing-edge flap are independent of body force orientation.

Nomenclature α = angle of attack αstall = stall angle of attack b = airfoil span c = chord length CD = coefficient of drag CL = coefficient of lift CP = coefficient of pressure H = test section height L = lower surface of the airfoil U = upper surface of the airfoil x = x distance along chord y = wake port location normalized by wake rake length

1 Assistant Professor, Department of Aeronautics, and AIAA Senior Member. 2 Cadet First Class, Department of Aeronautics, and AIAA Student Member. 3 Cadet First Class, Department of Aeronautics, and AIAA Student Member. 4 Clark Chair Professor, Aerospace and Mechanical Engineering Department, and AIAA Associate Fellow. 5 Director, Aerodynamics Group, and AIAA Senior Member.

45th AIAA Aerospace Sciences Meeting and Exhibit8 - 11 January 2007, Reno, Nevada

AIAA 2007-638

This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

American Institute of Aeronautics and Astronautics

2

I. Introduction variety of control and high-lift devices exist to manipulate longitudinal and lateral-directional motion on modern aircraft. The three primary control surfaces (aileron, rudder, and elevator) are typically a portion of a

wing which deflects in response to pilot actuation. These devices operate on the principle that a control surface deflection effectively alters the aerodynamic characteristics of a wing, such as lift and flow separation, by physically controlling the airflow. Though some high lift devices utilize ‘blowing’ techniques to control airflow and have been found to be effective, most stability and control systems rely on mechanically operated control surfaces and high-lift devices.

Several disadvantages afflict typical mechanical control devices. For military applications hinged control surfaces and high-lift devices are a major source of unwanted vibration, noise, and radar reflection.1 These control systems can also contribute significantly to the cost and weight of an aircraft. Present cost estimates reveal that replacing leading and trailing-edge control surfaces on a $30M aircraft would result in savings of nearly $450K.1 The ideal replacement for movable control surfaces is derived from a ‘virtual wing concept’.1 This ideal wing would consist of a generic shaped airfoil without camber and completely tiled with plasma actuators. Solely using plasma actuators’ response to govern the wing’s aerodynamic properties, the effects of devices such as ailerons, flaps, and cambered airfoils could be replicated and varied according to a flight profile.

Plasma research has been conducted since the early 50’s.2 With the advent of plasma research it was discovered that high voltage through two electrodes separated by an insulator would produce an electric field. Originally plasma was researched as a corollary to magnetic field research.2 As knowledge and understanding of the process grew, plasma’s effect on the surrounding fluid became more apparent. The electric field induced by the electrodes ionizes the air and consequently produces a body force on the air.

A single dielectric barrier discharge (SDBD) plasma actuator is a construction of electrodes separated by a dielectric isolative barrier. High AC voltage (10-20 kVp-p) is supplied to the electrodes at low current (~mA). In the region of highest potential, plasma forms on the surface of the dielectric surface as a result of ionized air between the electrodes. The electron discharge and buildup on the dielectric in the presence of an electric field gradient creates a body force in the direction away from the exposed electrode as shown in Figure 1. The buildup of charge on the dielectric and its low conductivity allow for a self-limiting discharge, and are the primary reasons the plasma does not arc. Most plasma actuators in current research use an asymmetric electrode configuration such that the induced flow is in the direction of the freestream.

Figure 1. Schematic of an Aerodynamic Plasma Actuator.

The application of weakly-ionized plasma actuators to high-lift devices such as flaps and slats has been found to

be effective, specifically with increases in maxLC and stallα as well as a 340% increase in lift-to-drag ratio.1 The application of plasma to wing sections without flow separation, investigated by Corke, Jumper, Post, Orlov, and McLaughlin, have shown that plasma actuators are effective in decreasing the zero lift angle of attack by approximately one half degree while maintaining the same lift curve slope.4 Furthermore, these plasma applications have shown plasma to effectively delay separation when the actuator is applied to the leading edge of an airfoil at high angles of attack normally associated with completely separated flow. 4 Previous flow control research found plasma actuators to be effective at delaying flow separation on a NACA 0015 airfoil up to angles of attack of 22°, where the airfoil would normally be stalled.6

Recent research3 indicates that several efforts to optimize the SDBD plasma actuator include various dielectric and electrode geometries. The extent of the plasma is found to be directly related to the input voltage. Therefore, for a given voltage input, a maximum effective electrode width exists.5 Decreasing the thickness of the exposed electrode is found to effectively increase the momentum transfer to the air. 5 It is also found that a thicker insulated electrode produced greater body forces at higher voltages.3

A

American Institute of Aeronautics and Astronautics

3

The present research intends to produce similar but with an oppositely oriented body force with respect to the freestream. When simulating a trailing-edge up aileron deflection or flap up configuration, the desired flow control includes a decrease in lift, an increase in drag, and induced flow separation at the plasma actuator location. The arrangement of electrodes on the plasma actuator is with the induced body force aligned against the freestream flow.

II. Set-up and Procedure A. Equipment and Data Collection The facility used for this investigation is the United States Air Force Academy’s (USAFA) Subsonic Wind Tunnel as shown in Figure 2. The wind tunnel is a closed-loop subsonic wind tunnel with a 3 ft by 3 ft test section, capable of Mach numbers in excess of M = 0.5. The testing is performed at an altitude of approximately 7,000 ft.

Figure 2. Schematic of the U.S. Air Force Academy Subsonic Wind Tunnel. A high-speed natural laminar flow airfoil, HSNLF(1)- 0213, was selected as the model for this investigation.

The HSNLF airfoils are characterized by their laminar flow and favorable pressure gradients over most of the chord. This type of airfoil allows for a relatively large separation ramp ideal for manipulating separation control at the trailing-edge of the airfoil. The same designation airfoil was used in the plasma actuator flow control work of Corke, Mertz, and Patel.7 The airfoil used in this investigation has a chord of 6 inches and a span of 18 inches. The model is constructed out of Sika block material using a numerical-controlled milling machine.

The airfoil is instrumented with 16 static pressure ports on both suction and pressure surfaces of the airfoil, shown in Figure 3. The chordwise port locations and relative spacing of the pressure ports is described in Table 1. Oval endplates with major axis diameters of 6 in. and minor axis diameters of 4 in., constructed using a Plexiglas polymer, minimize 3D wingtip effects.

Figure 3. HSNLF(1)-0213 Airfoil in the wind tunnel (photo taken from downstream).

American Institute of Aeronautics and Astronautics

4

Static pressures on the surface of the airfoil are acquired from pressure ports at the locations in Table 1. These

pressure ports are connected to a PSI 16TC transducer (1 psid) using Tygon tubes with an inner diameter of 0.04 in. A wake rake consisting of 32 pressure ports, spaced vertically apart at 0.125 in. between each port, is placed one chord length (6 in.) behind the airfoil. The wake rake pressure ports are connected to a PSI ESP-48 pressure transducer with a range of ±2.5 psid. Absolute and differential pressure measurements are taken from the wind tunnel system for each test. The entire pressure system is calibrated using a PSI System 8400.

Table 1. Locations of the airfoil’s surface mounted static pressure ports.

The basic configuration for a dielectric barrier plasma actuator consists of two asymmetrically placed

electrodes, constructed from 2.8-mil Saint-Gobain C665 copper tape, that are separated by a 5-mil Kapton/Polyimide tape with silicone adhesive dielectric. The exposed electrode (0.25 in. width) is arranged asymmetrically atop the embedded electrode (0.5 in. width) with a 0.0625 in. overlap and is placed further downstream then the embedded electrode as shown in Figure 1. This will allow for a body force opposite to the freestream.

Figure 4. Circuit diagram of the plasma actuator electronics.

A 5 kHz sine wave is supplied by an Agilent 33220A 20 MHz function generator. Its output is magnified by a Crown CE1000 1000 Watt amplifier and a CMI-5530-0212 step-up transformer. This provides the actuator with a 10 kVp-p sine function signal at a frequency of 5 kHz. A Tektronix P6015A high voltage probe and a Pearson Model 150 coil current monitor measure the output voltage and current respectively, while a Tektronix 2246 100 MHz oscilloscope allows observation of the outputs from the voltage probe and current monitor, as well as the original input signal. The plasma electrical equipment is connected using Silicones 3239 electrical wire rated up to 25 kV. The exposed electrode is grounded, while the embedded electrode subjected to the high voltage lead wire. This process allows for generation of the plasma on the surface of the airfoil over the embedded electrode.

Placement of the actuator is determined by examining the pressure distribution of the base cases (plasma off) at each Mach number. The pressure plots provides the location of the laminar separation bubble and/or the location of the inflection point of the pressure distribution. By placing the actuator in this vicinity, the actuator may be most efficient simulating the aileron-up deflection or flap up configuration desired. Taking this into consideration, the leading edge of the exposed electrode is placed at the x/c = 0.75 position for tests at M = 0.1 (Re=292K), and at the x/c = 0.70 chord position for M = 0.2 (Re=584K).

When constructing the actuator, special care was given to the application of the electrodes and the dielectric, to ensure high-quality plasma production and to prevent arcing. The airfoil was constructed with a small indented

Port # x/c Port # x/c 1 0.040 9 0.550 2 0.070 10 0.600 3 0.100 11 0.650 4 0.160 12 0.700 5 0.245 13 0.750 6 0.330 14 0.800 7 0.420 15 0.850 8 0.500 16 0.900

American Institute of Aeronautics and Astronautics

5

recess precisely the thickness of the dielectric in order to minimize flow obstruction and aerodynamic effects due to the presence of the actuator. Tests were conducted at M=0.1 and M=0.2 for number for a range of angles of attack varying from -2° to 16° in 2° increments, with a steady plasma actuator at 10 kV.

B. Theory To determine the effectiveness of the plasma actuator, the coefficients of lift and drag were calculated for each case. The coefficient of pressure, CP, was determined using the static pressure measurements from each of the static ports along the surface of the airfoil in conjunction with Equation 1,

2

21

∞∞

∞−=

U

ppC p

ρ

, (1)

where p is the local static pressure at a given port, p∞ is the freestream static pressure and 2

21

∞∞Uρ is the freestream

dynamic pressure. The coefficient of lift, CL, is calculated using Equation 2,

αα sincos anL CCC −= , (2)

where Cn is the normal component of lift, represented by Equation 3, and Ca is the axial component of lift, quantified using Equation 4

( )∫ ⎟⎠⎞

⎜⎝⎛−=

1

0 cxdCCC U

PLPn (3)

⎟⎠⎞

⎜⎝⎛

⎟⎟⎠

⎞⎜⎜⎝

⎛⎟⎠⎞

⎜⎝⎛

∂∂

−⎟⎠⎞

⎜⎝⎛

∂∂

= ∫ cxdL

xyCU

xyCC L

pupa

1

0

. (4)

As seen in Equations 3 and 4, the normal and axial components of lift are determined by integrating the

differential pressure existing between upper and lower surfaces of the airfoil. Equation 2 implies that the normal component has a contributes more to lift for the angles of attack investigated in this research, while the axial component ensures accurate results as angle of attack becomes large.

The coefficient of drag, CD, is calculated using the momentum deficit method. Local velocities in the wake of the airfoil were determined using the pressures acquired by the wake rake and isentropic relations. Equation 5,

( ) ( )( )∫ ⎟⎠⎞

⎜⎝⎛−= ∞∞

1

0 HydyUUyUD ρ , (5)

provides an expression for drag developed by Von Karman, where U(y) is the local velocity at each wake rake port, and H is the total test section height. An expression for the coefficient of drag per unit chord is then developed and is Equation 6,

( ) ( )( )∫ ⎟⎠⎞

⎜⎝⎛−= ∞

∞

1

022

HydyUUyU

cUCD , (6)

where c is the chord length of the airfoil in feet (for this investigation, 0.5 ft). Therefore Equation 7,

( ) ( )( )∫ ⎟⎠⎞

⎜⎝⎛−= ∞

∞

1

024

HydyUUyU

UCD , (7)

is the equation used to calculate coefficient of drag. The uncertainty analysis performed in this experiment consists of two primary considerations: the equipment manufacturer uncertainty and the random uncertainty of the pressure measurements. Total uncertainty in all the data is approximately 1.0%.

American Institute of Aeronautics and Astronautics

6

III. Results The pressure distributions for the baseline cases with the actuator off for the M=0.1 and M=0.2 cases are shown in Figures 6 and 7. These distributions depict the pressure over the suction surface of the airfoil at various angles of attack. In the M=0.1 case, the location of a separation bubble is visible at approximately x/c=0.75. In the M=0.2 case there is no separation bubble; but the inflection point of the pressure distribution occurs at x/c=0.70. These two characteristics in the pressure distributions determined the placement of the upstream orientated steady plasma actuator. The short pressure recovery region along the aft part of the airfoil and a large favorable pressure towards the leading edge indicates that natural laminar flow is being preserved.

M=0.1-2.5

-2

-1.5

-1

-0.5

0

0.50 0.2 0.4 0.6 0.8 1

x/c

CP

AoA=-2AoA=0AoA=2AoA=4AoA=6AoA=8AoA=10

M=0.2-3.5

-3

-2.5

-2

-1.5

-1

-0.5

00 0.2 0.4 0.6 0.8 1x/c

CP

AoA=-2AoA=0AoA=2AoA=4AoA=6AoA=8AoA=10

Figure 5. M=0.1 - CP with Actuator Off Figure 6. M=0.2 - CP with Actuator Off

By integrating the pressure distributions, like those in Figures 5 and 6, over the surface of the airfoil, the coefficient of lift was calculated. Figure 7 shows the coefficient of lift versus angle of attack for the M=0.1 case with the actuator on and off. A 2.2% increase in lift coefficient is observed at α = 8°. Figure 8 presents results of coefficient of lift versus angle of attack for the M=0.2 case with the actuator on and off. Here the effect of the plasma actuator is to uniformly increase the coefficient of lift for a given angle of attack. This is the effect of a traditional trailing-edge flap. In this investigation the body force generated by the actuator is against the freestream velocity. Since there is a slight increase in coefficient of lift, and not a decrease, the plasma actuator may be acting as a boundary layer trip forcing the flow to become turbulent. This would cause a delay in separation or a reattachment of the boundary layer aft of the actuator, both leading to an increase in lift.

Prior research at the U.S. Air Force Academy by Grothe, Wood, Post, and Thomspon8 noticed similar results on the same HSNLF airfoil with a plasma actuator whose body force was aligned with the freestream direction. Corke, He, and Patel1 observed, with a NACA 0015 airfoil and a trailing-edge actuator at x/c=0.9, a comparable increase in the coefficient of lift both in magnitude and direction; however, in that study the body force generated by the plasma flap was aligned in the flow direction. This suggests that when the plasma actuator is used as trailing-edge flap, the effect of the plasma actuator may be independent of body force orientation.

Wake velocity profiles for the Mach 0.1and 0.2 cases are shown in Figures 9 and 10, respectively. For the M=0.1 case the decrease in the wake velocities at α = 10° suggests that the drag is being increased with the actuator on. Likewise, a similar effect is seen in the M=0.2 case at α = 10°. By integrating the wake, the coefficient of drag can be calculated using the wake deficit method.

American Institute of Aeronautics and Astronautics

7

M=0.1

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

-4 -2 0 2 4 6 8 10 12 14α (deg)

CL

Plasma OffPlasma On

M=0.2

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

-4 -2 0 2 4 6 8 10 12 14

α (deg)

CL

Plasma OffPlasma On

Figure 7. M=0.1 – CL with Plasma Off and On Figure 8. M=0.2 - CL with Plasma Off and On

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0.85

0.90 0.95 1.00 1.05

V/V∞

Plasma OffPlasma On

M = 0.1α = 10°

0.60

0.65

0.70

0.75

0.80

0.85

0.90

0.95

1.00

0.90 0.95 1.00 1.05

V/V∞

Plasma OffPlasma On

M = 0.2α = 8°

Figure 9. M=0.1 – Wake Velocity Profile Figure 10. M=0.2 - Wake Velocity Profile with Plasma Off and On with Plasma Off and On

American Institute of Aeronautics and Astronautics

8

Figures 11 and 12 show the drag polar for the M=0.1 and M=0.2 cases. With the actuator on, the drag polar shifts up and to the right for both Mach numbers. This indicates that both lift and drag are increased with the actuator on. The upward drag polar shift appears to be accentuated in the Mach 0.2 case.

The shifting of the drag polar to the right is consistent with traditional trailing-edge flaps; however, the upward shift is not. Therefore, the upstream oriented actuator does not produce the classic effect of a flap, due to the observed increase in drag in this experiment. It produces a shift in the drag polar that is equivalent to increasing the camber effect of a plane flap. A type of flap that produces more drag and a similar shift in the drag polar was the so-called “split flap.” This result produced by this orientation and operation of the actuator is most analogous to the split flap. Split flaps were used in World War II aircraft because of their simplicity; however, the increase in drag they generate makes them less popular today. Corke, Jumper, Post, Orlov, and McLaughlin2 found a similar increase in drag with a single actuator near the trailing-edge. When more actuators were added in series (up to four), the increase in drag was eliminated. A single actuator separated the flow, while multiple actuators prevented this while still providing lift control.

0.02

0.03

0.04

0.05

0.06

0.07

0.08

0.09

0.10

0.11

0.12

-0.2 0.0 0.2 0.4 0.6 0.8 1.0 1.2

CL

CD

Plasma OffPlasma OnPoly. (Plasma On)Poly. (Plasma Off)

M=0.1

0.020

0.025

0.030

0.035

0.040

0.045

0.050

0.055

0.060

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7

CL

CD

Plasma OffPlasma OnPoly. (Plasma On)Poly. (Plasma Off)

M=0.2

Figure 11. M=0.1 – Drag Polar Figure 12. M=0.2 – Drag Polar with Plasma Off and On with Plasma Off and On

The pressure distributions showed that the boundary layer was laminar as it approached the trailing-edge ramp.

With the upstream oriented steady actuator on, an increase in both lift and drag were observed in this investigation. Similarly, Huang, Corke, and Thomas9 found that for separation control on the Pak-B blades simulating turbine blades in the low-pressure turbine stage, that for a steady plasma actuator they obtained the same results in terms of the reattachment location if the actuator were oriented upstream or downstream. From this result, they concluded the steady actuator was working as a turbulence transition trip. However, the same was not the case if the actuator was operating at an unsteady frequency near the optimum. They also found that the unsteady forcing only worked when the actuator was oriented downstream.

On the basis of this investigation and other comparative experiments1,2,8,9 the conclusion is that a steady plasma actuator with the body force aligned with or against the freestream when operated as a trailing-edge flap acts as a boundary layer trip. However, an optimized, downstream oriented, unsteady plasma actuator acts upon flow instabilities as its flow control mechanism.

American Institute of Aeronautics and Astronautics

9

IV. Summary The experiment analyzed the effect of a plasma actuator in simulating the effect of an upward aileron deflection

or trailing-edge flap upward configuration, via inducing flow opposite to the freestream direction. For all test cases, the effect of the plasma actuator resembled more closely the effect of a flap. Lift was increased an average of 2% with a higher increase at angles of attack where separation would normally exist (between 6°and14°). Drag also increased between 3% and 8%, showing a clear upward shift in the drag polar with a slight shift to the right. These results are most analogous with a split flap.

The steady plasma actuator with the upstream oriented body force acted as a trip in the boundary layer, causing a delay in separation or possibility a turbulent reattachment of the flow. On the basis of this investigation and other work in the literature1,2,8,9, the steady plasma actuator in the trailing-edge flap configuration acts as a trip in the boundary layer and its effect is independent of the induced body force direction. This combination of a laminar flow airfoil and a plasma actuator is an example of how a wing profile can be selected to make the flow more receptive to the actuator to achieve a desired outcome, in this case lift control.

V. Acknowledgments The authors would like to thank (1) Mr. Ken Ostasiewski, USAFA Subsonic Wind Tunnel technician, for

providing assistance with the experimental setup and data acquisition, and (2) Mr. Jeff Falkenstein for constructing the airfoil model and wake rake used in the experiment.

References 1Corke, T.C., He, C., and Patel, M.P., “Plasma Flaps and Slats: An Application of Weakly Ionized Plasma

Actuators,” AIAA Paper 2004-2127, July 2004. 2Gross, E.P., “Plasma Oscillations in a Static Magnetic Field,” Physics Review, Vol. 82, Issue 2, 1951, pp. 232,

242. 3VanDyken, R.D., McLaughlin, T.E., and Enloe, C.L., “Parametric Investigations of a Single Dielectric Barrier

Plasma Actuator,” AIAA Paper 2004-0846, Jan. 2004. 4Corke, T.C., Jumper, E., Post, M.L., Orlov, D., and McLaughlin, T.E., “Application of weakly-ionized plasmas

as wing flow-control devices,” AIAA Paper 2002-0350, 2002. 5Enloe, C.L., McLaughlin, T.E., VanDyken, R.D., Kachner, K.D., Jumper, E.J., Corke, T.C., Post, M.L., and

Haddad, O., “Mechanisms and Responses of a Single Dielectric Barrier Plasma Actuator: Geometric Effects,” AIAA Journal, Vol. 42, No. 3, March 2004.

6Post, M.L., and Corke, T.C., “Separation Control Using Plasma Actuators: Stationary and Oscillating Airfoils,” AIAA Journal, p. 9432-9444, Jan. 2004.

7Corke, T., Mertz, B., and Patel, M., “Plasma Flow Control Optimized Airfoil,” AIAA Paper 2006-1208, Jan. 2006.

8Grothe, R., Wood, K., Post, M., and Thompson, B.., “Scaling Effects of Aerodynamic Plasma Actuators on a HSNLF Airfoil,” USAFA Internal Report, May 2004.

9Huang, J., Corke, T., and Thomas, F., “Unsteady plasma actuators for separation control on lower pressure turbine blades,” Submitted to AIAA Journal.