[american institute of aeronautics and astronautics 44th aiaa aerospace sciences meeting and exhibit...

American Institute of Aeronautics and Astronautics

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Delayed Feedback Control for Flow around Wing using Plasma Synthetic Jet Actuator

Kakuji OGAWARA*, Yamaguchi University, Dept of. Mechanical Engineering, 2-16-1 Tokiwadai, Ube, Yamaguchi, 755-8611 Japan

Naoki NAKATANI†, Japan Defense Agency, Central Contract Office, 5-1 Honmuracho, Ichigaya, Shinjuku,Tokyo, 162-8801 Japan

Takuya NAKA‡, Mie ISOBE§, and Takehiro HIGUCHI¶ Yamaguchi University, Dept. of Mechanical Engineering, 2-16-1 Tokiwadai, Ube, Yamaguchi, 755-8611 Japan

Flow on an airfoil was investigated with plasma synthetic jet actuator. Plasma synthetic jet actuator is a flow control device which is composed of electrodes with A.C. signal, using electrohydrodynamic effect and induces a flow around the electrodes. From the characteristics of electrical device, plasma synthetic jet actuator has some important advantages; for example, miniaturization, maintenance free, and easy to control. From these characteristics, it is effective to apply flow control around an airfoil with plasma synthetic jet actuator. In this study, the behavior of the plasma synthetic jet actuator around an airfoil is observed by experiment. A NACA0012 airfoil with plasma synthetic jet actuator is set in the wind tunnel. The wake velocity behind air foil is measured with and without the actuator. Experimental result with the Delayed Feedback Control for the flow control system is shown in this paper. The results show good improvement for the plasma synthetic jet actuator to reduce drag in the high angle of attacks.

Nomenclature Fb = body force per volume

0ε = permittivity of the free space E = electric field strength u = measured velocity U ∞ = main stream flow velocity Rec = Reynolds number K = weight τ = delayed time fa = basic frequency fc = control frequency

* Professor, Yamaguchi University, Dept. of Mechanical Engineering, 2-16-1 Tokiwadai, Ube, Yamaguchi, 755-8611 Japan, Member. † Officer, Japan Defense Agency Central Contract Office, 5-1 Honmuracho, Ichigaya, Shinjuku, Tokyo, 162-8801, Japan, Non-member. ‡, Student, Yamaguchi University, Dept. of Mechanical Engineering, 2-16-1 Tokiwadai, Ube, Yamaguchi, 755-8611 Japan, Student Member. §, Student, Yamaguchi University, Dept. of Mechanical Engineering, 2-16-1 Tokiwadai, Ube, Yamaguchi, 755-8611 Japan, Member. ¶, Research Associate, Yamaguchi University, Dept. of Mechanical Engineering, 2-16-1 Tokiwadai, Ube, Yamaguchi, 755-8611 Japan, Member.

44th AIAA Aerospace Sciences Meeting and Exhibit9 - 12 January 2006, Reno, Nevada

AIAA 2006-1406

Copyright © 2006 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.


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I. Introduction leading edge flap and a trailing edge slat is used for usual planes to delay the separation and to gain high lift

force to avoid stall. Besides, many flow control methods were proposed. Some of the control methods have ability to prevent separation in cases of high angle of attack over 20 1.

The actuator that has zero mass flux by mutual synthetic inlet and outlet are called synthetic jet actuator (SJA). SJA is effective device for separation control 2,3. In particular, McCormick et al. has shown a good result with changing the direction of jet4 Gillarranz et al. has shown a success by smaller and more efficient SJA, and gives a possibility of lift control without aileron using hinges5.

Recent years, plasma actuator is studied as the induced jet device for the application of the flow control similar to the SJA. This plasma actuator is the induced jet device that uses plasma generating with A.C. glow discharge. Sherman et al. has shown the drag reduction on a flat plate with plasma actuator6. Furthermore, Corke et al. has presented the suitable electrodes shape of plasma actuator for flow control on airfoil, and tested this actuator for separation control with NACA663-018, NACA0009, and NACA0015 airfoil7-9.

The authors have been studying the plasma actuator with electrode shape to induce 3-D jet for separation control. This actuator works like SJA and is named plasma synthetic jet actuator (PSJA). In experimental investigation of this PSJA, the NACA0012 airfoil had effect of drag reduction up to 29%10. In numerical study, the optimum position of electrode and the combination of PSJA was studied11-13.

In this paper, the wake velocity behind the NACA0012 airfoil is measured with and without PSJA proposed by Corke et al. The effect of the PSJA is observed, and closed active control using delayed feedback control is tested.

II. Plasma Synthetic Jet Actuator Schematic view of plasma synthetic jet actuator induced 3-D jet is shown in Fig. 1. This actuator is composed of

anode with nichrome which is plated cord in the shape of needle, and cathode with 0.05 [mm] thick aluminum tape in the shape of washer named Needle – Washer type ( N – W type ). Anode is set in the center of cathode making a 4 [mm] diameter hole. Schematic view of PSJA for 2-D jet proposed by Corke et al. is shown in Fig. 2. This actuator is composed of anode and cathode with 0.02 [mm] thick copper tape and polyimide film for insulator between anode and cathode. This is called Strip - Strip type ( S - S type ). The electrodes and insulator is arranged perpendicular to the flow direction, and anode is set in the upper side. In this paper, experimental results are demonstrated about S – S type. Fig.3 is schematic view of circuit to operate PSJA. This actuator is supplied by high voltage A.C. input of rectangle wave using a signal generator (MEGURO, MAS – 412), the input voltage is amplified by power amplifier (KENWOOD, COMPACT DISC STEREO SYSTEM DG99) and transformer (UNION, 1 : 75) is used to operate PSJA. In the experiment, the input voltage is 1300 [V] and the frequency is 500 [Hz].Mechanisms to induce flow including electric force is shown in equation (1),

A

Nichrom-plated cord Aluminum tape

1mm

4mm

50μm

Airfoil

Induced Flow

Fig.1 N – W type PSJA

Aluminum tapesKapton film(Insulator)

Induced Flow

Fig. 2 S – S type PSJA

Cupper Tapes

Top View

Fig.1 N-W type PSJA Fig.2 S-S type PSJA


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2

021 EFb ∇= ε (1)

where Fb is the body force per volume, 0ε is the permittivity of the free space, and E is the electric field strength. This enables N – W type and S – S type PSJA to produce the flow shown in Fig. 1 and Fig. 2.

III. Experiment Setup The NACA0012 airfoil, with PSJA is set in wind tunnel to observe behavior of the actuator in this experiment.

Schematic view of wind tunnel is shown in Fig. 4. The wind tunnel is subsonic wind tunnel with the size of 2400 × 1450 × 800 [mm] (total length × total height × total width). This wind tunnel consists of bell mouse, honeycomb, contraction, test section, diffuser, damping screen, axial blower, and control box. The air is drawn into the facility through a flow manipulation section consists of a bell mouse (W 750 × H 775 × L 80 [mm] ) and 85 [mm] thick honeycomb ( 3 / 8 inch cell ). The combination of the flow manipulation and contraction is designed to give u/ U ∞ within 1 %. The test section is a straight section with a W 150 [mm] × H 250 [mm] × L 500 [mm].

The airfoil used in this study is a NACA0012 made of acrylic fiber. This airfoil is shown in Fig. 5. The airfoil is symmetric airfoil with 60[mm] chord150 [mm] span, and 7.2 [mm] thick at 18 [mm] from leading edge. PSJA is

Amp. Input Signal

10Ω

PSJA1 :625

Fig.3 Schematic view of Circuit

① Bell Mouse ② Honeycomb ③ Contraction ④ Test Section ⑤ Diffuser ⑥ DampingScreen ⑦ Axial Blower ⑧ Airfoil ⑨ Control Box

① ② ③

④ ⑤ ② ⑥ ⑦⑧

⑨

Fig.4 Schematic view of Wind Tunnel

Fig.5 NACA0012 Airfoil & Actuator (unit : mm)

60

150

18 7.2 4.2 P S J A

Copper Tape

Polyimide Film

Hot Wire x

y

PSJA Flow

Fig.6 Wind Tunnel Test


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embedded on upper side of the airfoil. The airfoil is supported at the position of the maximum airfoil thickness and is possible to change angle of attack(A.O.A.). Schematic view of experimental setup in the wind tunnel test section is shown in Fig. 6. In this experiment, direction of span (z axis) and direction of support axis is the same. The wake velocity distribution behind the airfoil is measured with or without PSJA by changing A.O.A. from –15 to 15 around z axis. A I type hot wire is used to measure wake velocity at x / C ≅ 1.2, 2.3, 3.5. In measurement point, as interval of measurement in y axis∆ y [mm], ∆ y = 2 [mm] within y = ± 36 [mm], ∆ y = 5 [mm] is from ± 36 [mm] to ± 56 [mm]. The measurement time is 10 [s], and sampling rate is 10 [kHz] in each measurement point. Reynolds number Rec with representative velocity U ∞ = 5 [m/s], and representative length C = 60 [mm] is Rec ≅ 20000.

IV. Experimental Results The wake velocity distribution ( u / U ∞ ) and the root mean square ( rms ) distribution behind airfoil at A. O.

A. -15 , -10 , -8 , -5 , 0 , 5 , 8 , 10 and 15 are shown in Fig. 7. In the condition at A. O. A. 0 , 5 , 8 and 10 , changes in averaged velocity distribution and rms distribution was observed. This shows that PSJA added turbulence to a flow. This change remarkably appeared in upper side of airfoil, and the velocity loss and rms was reduced. The most notable effect appeared at A.O.A. 10 , where drag reduction was observed. On the other hand, the effect did not appear at A. O. A. 15 , -15 , -10 , -8 , -5 as shown in Fig. 7. In this experiment, the stall angle is around 7 , the effect didn’t appear in complete separating flow around airfoil at A. O. A. 10 . Here, the FFT of the frequency is analyzed for the raw velocity data of 10 where the effect was notably observed. The analysis is done at y = 6 [mm] where the remarkable effect appeared. The FFT results of 10 and 15 is compared. The result of x / C≅ 1.2, 2.3, and 3.5 was much the same, the result of x / C≅ 2.3 is shown in Fig. 8. In the case of 10 , FFT analysis without PSJA has various frequency in low frequency region. But, in the result with PSJA, the frequency ingredient in the low frequency did not exist. The results show that only the frequency of the ingredient to which some were restricted is contained. In the case of 15 FFT analysis show that the various frequency existed in the low frequency region. Especially at frequency around 60 [Hz] has a pointed peak. In the case of A.O.A.15 , this separation which flow separated from leading edge of NACA0012 airfoil is considered that a frequency of Karman vortex emitted from leading edge is around 60 [Hz]. The change of FFT analysis could not be observed with or without operating PSJA in case of 15 .

V. Delayed Feedback Control As next approach, the delayed feedback control (DFC)15 is introduced to see the flow condition in case of A.O.A

15 using the following equation.

( ) ( ){ } 5.0+−−= tutuKduty τ (2) Figure 9 shows the input signal for PSJA using the DFC. The input PWM signal with basic frequency of

fa=500[Hz] is given by using Digital Signal Processor ( DSP; dSPACE, DS1104 R&D Controller Board ). duty in Eq.(2) gives the on-off time ratio shown in Fig.9. The control frequency fc is given as 60[Hz] and 170[Hz]. From this input, it is possible to put the jet with the lower frequency of the basic frequency. u(t) is the velocity of the wake flow. K is the gain and the τ is the delay time. Using this K and τ , the control is designed heuristically.

The controlled results are shown in Fig. 10 and 11. The right figure of Fig. 10 shows the wake velocity distribution and left figure is the root mean square distribution.in the figure shows the case with PSJA and the white circle shows the case without PSJA. The results show that the rms and wake velocity distribution did change when the PSJA is control with the DFC. This shows that the control worked to change the condition of the wake flow. Especially (c) and (d) shows that the velocity loss of the wake flow has decreased which shows that the drag has reduced. Fig.11 shows the FFT analysis of the frequency. The peaks around 60[Hz] in Fig.11 has decreased by the effect of PSJA whichdefines that the Karman vortex from the leading edge has decreased using new control. From these results, the PSJA is able to control the cases with high angle of attacks that were difficult without the DFC.


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VI. Conclusion Investigation results for flow around airfoil with plasma synthetic jet actuator are explained in this paper. The

actuator with electrode shape to induce 3-D jet was proposed for separation control, and the wake velocity behind the NACA0012 in which this PSJA was embedded was measured with or without operating PSJA proposed by Corke et al. In the condition at A. O. A. 0 , 5 , 8 and 10 , change of the averaged velocity distribution and rms distribution was observed with operating PSJA. This change remarkably appeared in upper side of airfoil, and the velocity loss and rms was reduced. And closed active control using Delayed Feedback Control was tested. This result of closed loop control system has shown improvement when it was compared to open loop control result. The Delayed Feedback Control has shown good improvement on condition of high angle of attack around 15 .

References 1Gad – el - Hak, “Separation Control: Review”, ASME J. of Fluid Engr., Vol. 113, pp 5 – 30, 1991 2Seifent A., et al., “Oscillatory Blowing: Atool to Delay Boundary-Layer Separation”, AIAA J., Vol. 31, No. 11, pp. 2052-

2060, 1993 3Smith. L. and Glezer. A, “The Formation and Evolution of Synthetic Jets“, Phys. Fluids, Vol. 10, N0. 9, pp. 2281 – 2297,

1998 4D. McCormick, et al., “Rotorcraft Retreating Blade Stall Control”, AIAA 2000 – 2475, Fluids 2000 Conference and Exhibit,

Denver Colorado, 19 – 20 June 2000 5Gilarranz, et al., “Compact High-Power Synthetic Jets for Flow Separation Control”,AIAA paper, 2001 – 0737, 2002 6Daniel M. Sherman, et al, “Electrohydrodynamic Flow Control with a Glow – Discharge Surface Plasma”, AIAA J., Vol. 38,

No. 7, July 2000 7Corke, T. C., Jumper, E., Post, M., Orlov, D. and McLaughlin, T. 2002. “Application of weakly – ionized plasmas as wing

flow – control devices”, AIAA paper, 2002 – 0350, Jan 2002 8Post, M. and T. Corke.., “Separation Control on high angle of attack airfoil using plasma actuator”, AIAA paper, 2003 –

1024, 2003 9Post, M., Corke, T. C., “Separation Control using plasma actuators stationary and oscillating airfoils“, AIAA paper, 2004 –

0841, 2002 10Kakuji Ogawara, Naoki Nakatani, Souichi Saeki, “Flow Control to the Surface on Airfoil with Plasma Synthetic Jet

Actuator”, JSME, 02 – 1494, Oct 2003 11Tadao Ueda, Kakuji Ogawara, “Numerical 3 – D Simulation for Control of Flow around an Airfoil Using Plasma Synthetic

Jet Actuator”, JSME, 03 – 0853, Aug 2004 12Tadao Ueda, Kakuji Ogawara, “Numerical Simulation of Flow Induced by Plasma Synthetic Jet Actuator”, JSME, 04 –

0297 13Tadao Ueda, Kakuji Ogawara, “Study on Optimum Location of Plasma Synthetic Jet Actuator Using Pseudo Viscosity

Model”, JSME, 04 – 0298 14Tomonori Nakano, Nobuyuki Fujisawa, “Simultaneous measurement of noise and velocity field under discrete tone noise

from a symmetrical aerofoil”, Journal of Visualization , Vol. 24, No. 1, (2004), 337 – 340, 15K. Pyragas, “Continuous control of chaos by self – controlling feedback”, Phys. Lett. A, Vol. 170, 421 – 428


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0.6 0.7 0.8 0.9 1.0

U/U∞ [-]

Y Axis [mm]

WithPSJA WithoutPSJA

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0.2 0.4 0.6 0.8 1.0

rms [m/s]

Y Axis [mm]


0.2 0.4 0.6 0.8 1.0 0.2 0.4 0.6 0.8 1.0X X

Fig. 7 (a) Averaged velocity and rms profile at -15

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rms [m/s]

Y Axis [mm]


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Fig. 7 (b) Averaged velocity and rms profile at -10

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rms [m/s]

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Fig. 7 (c) Averaged velocity and rms profile at -8


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Fig. 7 (f) Averaged velocity and rms profile at 3

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0.78 0.84 0.90 0.96 1.02

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0.200.250.300.350.400.45

r.m.s. [m/s]

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0.200.250.300.350.400.45 0.200.250.300.350.400.45

Fig. 7 (d) Averaged velocity and rms profile at -5

XX

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0.84 0.88 0.92 0.96 1.00

Y Axis [mm]


0.84 0.88 0.92 0.96 1.00

U/U∞[-]

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0.1 0.2 0.3 0.4 0.5 0.6 0.7

Y Axis [mm]


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r.m.s. [m/s]

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Fig. 7 (e) Averaged velocity and rms profile at 0


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Fig. 7 (g) Averaged velocity and rms profile at 8

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rms [m/s]

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Fig. 7 (h) Averaged velocity and rms profile at 10

-60

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0.6 0.7 0.8 0.9 1.0 1.1

U/U∞[-]

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rms [m/s]

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Fig. 7 (i) Averaged velocity and rms profile at 15


9

10 100 1000

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Power

Frequency [Hz]

WithPSJA

Fig. 8(b) FFT analysis at 15 ( y = 6 [mm] x / C ≅ 2.3 )

10 100 1000

1000

10000

100000

Power

Frequency [Hz]

WithoutPSJA

A(t)

fa = 500 [Hz]

Ta = 1 / fa

t

Tc = 1 / fc

duty = C / B

C

B

Fig. 9 Schematic view of Wave for Duty Modulating Input with Delayed Feedback Control

10 100 1000

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Power

Frequency [Hz]

WithoutPSJA

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WithPSJA

Fig.8 (a) FFT analysis at 10 ( y = 6 [mm] x /C ≅ 2.3 )


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Fig. 10 (a) Averaged velocity profile and rms profile with DFC (K = 0.3, τ = 0.008, fc = 170 [Hz])

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Fig. 10 (b) Averaged velocity profile and rms profile with DFC (K = 0.3, τ = 0.016, fc = 170 [Hz])

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Fig. 10 (c) Averaged velocity profile and rms profile with DFC (K = 0.3, τ = 0.004, fc = 60 [Hz])


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Fig. 10 (d) Averaged velocity profile and rms profile with DFC (K = 0.3, τ = 0.008, fc = 60 [Hz])

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Fig. 11 (a) FFT Result at y = 6 [mm] x / C ≅ 2.3(K = 0.3, τ = 0.008, fc = 170 [Hz])

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Fig. 11 (c) FFT Result at y = 6 [mm] x / C ≅ 2.3(K = 0.3, τ = 0.004, fc = 60 [Hz])

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Power

Frequency [Hz]

Fig. 11 (d) FFT Result at y = 6 [mm] x / C ≅ 2.3(K = 0.3, τ = 0.008, fc = 60 [Hz])

Fig. 11 (b) FFT Result at y = 6 [mm] x / C ≅ 2.3(K = 0.3, τ = 0.016, fc = 170 [Hz])

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