Document
AIAA-2002-3273
Virtual Shaping of a Two-dimensional NACA 0015 Airfoil
Using Synthetic Jet Actuator
Fang-Jenq Chen* and George B. Beeler* NASA Langley Research Center Hampton, VA 23681-0001 ABSTRACT ACI change of lift coefficient due to actuation c airfoil chord The Aircraft Morphing Program at NASA Langley envisions an aircraft without conventional control Eel, effective voltage surfaces. Instead of moving control surfaces, the Ein amplitude of sinusoidal input voltage vehicle control systems may be implemented with a f frequency combination of propulsive forces, micro surface h jet exit slot width effectors, and fluidic devices dynamically operated by I4y effective current an intelligent flight control system to provide aircraft P active power, Eq. (1) maneuverability over each mission segment. As a part Q reactive power, Eq. (2) of this program, a two-dimensional NACA 0015 airfoil Re_ Reynolds number based on chord length model was designed to test mild maneuvering U mean streamwise velocity capability of synthetic jets in a subsonic wind tunnel. u broadband rms streamwise fluctuating velocity The objective of the experiments is to assess the v phased-averaged jet velocity x streamwise or chordwise coordinate applicability of using unsteady suction and blowing to alter the aerodynamic shape of an airfoil with a purpose z spanwise coordinate to enhance lift and/or to reduce drag. Synthetic jet 0 phase angle of voltage with respect to current actuation at different chordwise locations, different forcing frequencies and amplitudes, under different Subscripts freestream velocities are investigated. The effect of max maximum value during actuator blowing cycle virtual shape change is indicated by a localized increase mean mean value of surface pressure in the neighborhood of synthetic jet rain minimum value during actuator suction cycle actuation. That causes a negative lift to the airfoil with rms root-mean-square value an upper surface actuation. When actuation is applied freestream near the airfoil leading edge, it appears that the stagnation line is shifted inducing an effect similar to that caused by a small angle of attack to produce an 1. INTRODUCTION overall lift change.
NOMENCLATURE New technologies and advanced materials, both nearly in hand and in early development, offer the potential to create revolutionary advances in aerospace Cp pressure coefficient vehicles with significantly greater performance and Cjp pressure drag coefficient maneuvering compared to conventional approaches.
C_,_ upper surface lift coefficient Conventional airfoils of aerospace vehicles have C_ jet momentum coefficient 2(h/c)*(v_/U_) 2 been designed for a single flight condition and then *Senior Aerospace Engineer.
modified to cover multiple flight conditions. This is Copyright © 2002 by the American Institute of Aeronautics and done through the use of control surfaces, such as Astronautics, Inc. No Copyright is asserted in the United States under ailerons and flaps, spoilers, and variable wing sweep.
Title 17, U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for Variable wing sweep affects changes in the local flow Governmental purposes. All other rights are reserved by the copyright field by altering the flow velocity normal to the leading owner.
American Institute of Aeronautics and Astronautics edge of theairfoil.Thecontrol surfaces andspoilers The work described inthepresent paper focuses on affect changes intheflowfieldbydirectly varying the theaerodynamic virtualshaping of anairfoilatzero camber oncertain regions ofthe airfoil, thereby causing angleof attack, usingonlysynthetic jets without changes in thebaseline aerodynamic characteristics of additional fixtures. This represents anexamination of
the entire airfoil. thefluidic modification ofaero-surfaces thatdoes not
The Aircraft Morphing Program atNASA Langley necessarily relyoncoupling toflowinstability such as hasbeeninitiatedwith the goalto develop next that inherently inseparated flows. The effects ofvirtual generation aerospace vehicles by mimicking Nature. shaping areinvestigated withsynthetic jet actuation at Effortsinclude providing muscle-like actuators to differentchordwise locations, differentforcing change aerodynamic form ondemand, nervous-system- frequencies and amplitudes, under different freestream likesensing, self-healing materials, and adaptive fault- velocities. Characteristics of electric power tolerant controls toenhance flightsafety. One approach consumption of the synthetic jet actuator arealso
envisions future vehicles without conventional control
reported inthis paper.
surfaces that could lead tosignificant weight reductions
2. APPARATUS
or decrease theaircraft signature. Instead of moving
control surfaces, theaerodynamic control systems may
beimplemented withacombination ofthrust vectoring 2.1 Wind Tunnel Facility
propulsion, adaptive micro-machined surface effectors, The experiment was conducted in the 2 ft x 3 ft and distributed fluidic devices dynamically operated by Low-Speed Wind Tunnel located at the NASA Langley an intelligent controlsystem to provideaircraft Research Center. The tunnel is a closed-loop type with a 10:1 contraction ratio. The test section is 91.4 cm
maneuverability under different flightconditions. Asa
partof thisprogram, a two-dimensional NACA0015 wide by 61.0 cm high by 6.1 m long. The turbulence- airfoilmodel wasdesigned totestmildmaneuvering reduction devices upstream of the contraction consist of capability ofsynthetic jetsin the NASA Langley 2 ft x a honeycomb followed by four stainless-steel screens.
3 ft Tunnel. Theobjective of theexperiments is to A vane-axial fan powered by a 30 horsepower DC assess theapplicability ofusing unsteady suction and motor is used to drive the tunnel. Speeds of blowing ofsynthetic jetstoalter the aerodynamic shape approximately 45 m/sec are attainable in the test section ofanairfoil toenhance liftand/or toreduce drag. with measured turbulence intensities, u/U_, of Previous preliminary testswith a piston-type approximately 0.1% in the range of 0.1 <f< 400 Hz.
actuator indicate that theactuator has aneteffect onthe
The test-section ceiling and floor are adjustable to boundary-layer flowofproducing alocal netincrease in achieve a desired streamwise pressure gradient and accommodate various test conditions. Further details of
the displacement thickness of the flow. Hence an
effective shape change could approximate theeffect of the facility have been previously published)
the synthetic jetactuator. Most recent works 1-6 onairfoil
2.2 Airfoil Model
aerodynamic manipulation usingsynthetic jets has
The two-dimensional NACA 0015 airfoil model
focused onthecontrol of flowseparation atmoderate
and large angles ofattack. The experimental workdone has the dimensions of 91.4 cm span and 91.4 cm chord.
byChatlynne etal. 7showed that it ispossible tomodify The model was made with a 0.3 cm thick lay-up of theapparent aerodynamic shape of anairfoilatlow fiberglass skin supported by four internal chordwise angles of attackwhenthe baseline flow is fully ribs machined from 6061-T6 aluminum alloy. There are attached. However, this virtual aero-shaping was six chordwise locations for installation of a synthetic jet achieved by combining the activation of a high- actuator with a long exhaust slot centered spanwise on frequency synthetic jet actuator placed downstream the model. Figures 1 (a), (b) and (c) show a side view, a froma miniature surface-mounted passive obstruction cut-away view and a photograph of the airfoil model.
(a little fence). Thiscombination introduces a small The model was mounted to external mounts through the stationary re-circulating flowregion next tothesurface, tunnel sidewalls by a spanwise-extended steel tube, as whichdisplaces thelocalstreamlines sufficiently to shown in the center of Figure 1(c). The surface finish on the airfoil surface was 1 micron or better.
modify thelocalpressure distribution. Themodified
flowresults in asignificant reduction in pressure drag Three rows of streamwise pressure taps are located witha minimal loss of lift. Physically, thefunction of on the midspan, the right- and the left-quarter-spans on thesynthetic jet actuator inthisapplication is toforce both upper and lower surfaces. The pressure taps are theseparated flowdownstream of thelittle fence to staggered in each row to minimize streamwise
reattach tothe surface. interference. The nominal orifice diameter of the
pressure taps is 0.05 cm. Three 32-port 10-inch water- American Institute of Aeronautics and Astronautics column ESP modules were connected topressure taps slot center on jet exit surface, are shown in Figure 5. It togetthesurface static pressure readings. Ten dynamic presents two distinguished peaks at the frequencies pressure transducers, twoKulite XCS-062-5D and eight around 600 Hz and 1300 Hz. These two peak Endevco 8510B-l, were distributed along themidspan frequencies correspond to the natural (mechanical) withnineontheupper surface andoneonthelower frequency of the piezoelectric diaphragm and the surface. Locations ofthe pressure taps and thedynamic Helmholtz (acoustic) frequency of the actuator cavity, pressure transducers areshown in the sketches of respectively. More detailed discussions of these Figures 2(a) and (b).One rowofspanwise pressure taps resonant frequencies were given in a previous paper. 9 distributed near the trailing edge approximately at x/c The characteristics of electric power used to drive 0.95, also shown in Figure 2(a), was used to check the piezoelectric diaphragms of the actuator at the possible tunnel-sidewall contamination of the surface forcing frequencies of 600 and 1300 Hz are shown in pressure readings. Figures 6 and 7, respectively. The electric power level was adjusted by varying the amplitude of sinusoidal 2.3 Synthetic Jet Actuator voltage input through the multi-channel attenuation unit A long two-dimensional synthetic jet actuator, as to a high-voltage power amplifier with a nominal gain shown in Figure 3, was installed across the airfoil span of 100 before sending to the actuator. The input at one of six different chordwise locations for the test.
voltage, E_n, is the peak-to-peak amplitude of the These six actuator locations are at x/c -0.015, 0.1, sinusoidal voltage generated by a universal function 0.15, 0.3, 0.5, and 0.75 (the negative value was on the generator. Figures 6(a) and 7(a) show variations of the bottom side of the airfoil). One of the test purposes was effective values (in terms of standard deviations) of to determine the optimal placement of the actuator on current, Iejj_ voltage, Ee_ and the phase angle, 0, of the airfoil for virtual shape change. The actuator voltage with respect to current, with the input voltage, consists of a continuous cavity enclosed by 14 pairs of E_n. Generally the effective current and voltage increase piezoelectric diaphragms. Each pair of piezoelectric with the input voltage but the phase angle decreases diaphragms was operated with a 180 ° phase differential slightly from +90 ° . It indicates that the piezoelectric at the same sinusoidal voltage and frequency. With diaphragm deviates from a perfect capacitance device actuation, a synthetic jet issued from a two-dimensional as the input voltage increases. Figures 6(b) and 7(b) slot that is 76.2 cm long by 0.5 mm wide on the top show relations between the jet energy, in terms of variance of jet velocity, (v,_s)2, the active (P) and the plate of the device. Six top plates with different curvatures were made to fair the actuator to different reactive (Q) electric powers with the input voltage, E_n.
chordwise locations on the airfoil model. A multi- Note that the reactive power is plotted as Q/4 in the channel attenuation unit was used to obtain the best figures in order to compile data in the same scale. The uniformity of the synthetic jet across the airfoil span by active power, P, and the reactive power, Q, are defined as 10 adjusting the amplitude of power input to each pair of piezoelectric diaphragms.
P= EeHIeHcosO (1) 3. CHARACTERISTICS OF Q = EeHI_HsinO (2) ACTUATOR The performance of the piezoelectric diaphragm (hence, The strength of synthetic jets generally is not the actuator) can be enhanced by an offset of the input voltage that is shown as a step jump for the input uniform along the 76.2 cm spanwise slot due to irregularity in the piezoelectric diaphragms. After voltage equal to and greater than 2.0 volts in the figures. The maximum performance condition of the careful tuning with the attenuation unit, the best uniformity of the synthetic jet at the maximum piezoelectric diaphragm and the actuator was achieved performance condition (defined later) was obtained and at the input voltage of 2.3 volts with an offset of 0.6 shown in Figure 4. It shows slight variations of jet volts. Data at the maximum performance condition are plotted in Figures 8(a) and 8(b) against the forcing velocities along the slot centerline on the jet exit surface in the quiescent environment. The jet is slightly frequency. The active power is closely correlated with the jet output as seen in the (b) plot of Figures 6, 7 and stronger in the position at top of the center of each piezoelectric diaphragm pair than that at top of the 8. The reactive power is due to the capacitive nature of the actuator. The magnitude of the reactive power is interval between two diaphragm pairs.
Typical variations of jet velocities at the maximum always greater than the active power in this device. It is very interesting to see that the performance peaks, performance condition with respect to the forcing frequency, in the range of 200 Hz to 1700 Hz, at the approximately at 600 and 1300 Hz, happen when the American Institute of Aeronautics and Astronautics phase angle between voltage andcurrent is small and cm wide tape with randomly distributed #35 Grit was the reactive power is low,asshown inFigures 8(a) and used to trip the boundary layer at x/c 0.12 on the 8(b). The active power isveryindicative ofthephase airfoil upper surface. Figure 12 shows that the matching of thedrivingfrequency to theresonance boundary-layer trip moved the transition boundary mode ofthesynthetic jet asseen bythepeaks in active (laminar data on left-hand side) to a lower freestream power andjet output in Figure 8(b).Thediffering velocity but no discernible difference on airfoil lift character of thepeak active power attheHelmholtz change due to synthetic jet actuation for data between resonance versus themechanical resonance maybe trip (turbulent data) and no-trip (laminar data) transition indicative ofthemore complex non-linear coupling in boundaries except normal data scattering.
the system. The data hints atthepotential ofmonitoring Typical variations of airfoil lift change with forcing theactive power asapotential tuning tooltomaintain frequency, when the actuator was driven at the the excitation near resonance ofthe actuator.
maximum performance condition (i.e., input 2.3 volts and offset 0.6 volts), are shown in Figure 13. The 4. VIRTUAL SHAPING USING change of lift coefficient, AC), is presented as a SYNTHETIC JETS percentage of the airfoil upper surface lift, Cl, u_, that has a nominal value of 0.23. Two negative lift peaks in The effect of virtual shape change on the airfoil Figure 13 correspond to the jet energy peaks as shown model was detected by comparing the measurements of in Figure 8(b). Variations of airfoil lift change, airfoil surface pressures with the synthetic jet actuation AC)/CI, u_, with jet momentum coefficient, C_, are on and oft'. Typical variations of the surface pressure presented in Figures 14(a) to 14(d) for the actuator located at x/c 0.1, 0.15, 0.3 and 0.5 on the airfoil coefficient, C_, with actuation on and off on the upper and the lower surfaces are shown in Figures 9(a) and upper surface. Data are presented for the actuator 9(b), respectively. Customarily the @ scale is inverted operated at peak forcing frequencies of 600 Hz and in the plot. It clearly indicates a localized increase of 1300 Hz with varying input voltages and freestream the surface pressure in the neighborhood of synthetic jet velocities. Generally the lift change increases with C_ and actuation chordwise coordinate after x/c 0.15.
actuation. That causes a negative lift on the airfoil with the actuator located on the upper surface. This negative However, the lift change decreases for actuation at x/c lift change is consistent with numerical simulations of 0.75 as shown in Figure 15. Numerical computations Hassan. 11 Note that, at zero angle of attack, there is zero verified that the laminar boundary layer separation lift for the baseline (without actuation) NACA 0015 occurred at x/c 0.53. It indicates that virtual shaping airfoil due to the geometric symmetry. The computed by synthetic jet actuation is much less effective in separated flow. In the range of data tested, the @ from the Reynolds Averaged Navier-Stokes solver CFL3D 12 for the baseline airfoil (also included in the maximum lift change was -0.015 that resulted in a -6% figures) show a good match with the measured no- change in ACI/CI,_ with actuation at x/c 0.5, as control data. There is no discernible spanwise variation shown in Figure 14(d). An attempt to correlate the data shown in Figures 14(a) to 14(d) is presented in Figure of Cp by comparing the three rows of streamwise 16 with a fitted line of Y -17")( °25 where X and Y pressure readings along the midspan, the right- and the left-quarter-spans on both upper and lower surfaces, represent the lumped variables for the x and y denoted as UpperC, UpperR, UpperL, LowerC, coordinates, respectively. The correlation result implies that virtual shaping by synthetic jet actuation is LowerR, and LowerL in the figures. Typical C_ distributions from spanwise pressure taps near the inversely proportional to the Reynolds number and the trailing edge with the synthetic jet actuation on/oft, local logarithmic pressure gradient but increases under the same flow conditions as Figure 9, are (negative lift) with the jet momentum coefficient.
presented in Figure 10. It indicates no discernible When synthetic jet actuation was near the leading edge (x/c -0.015), it appeared that the stagnation line sidewall contamination on the @ data.
Presumably virtual shaping is by nature an inviscid was shifted inducing an effect similar to that caused by local phenomena rather than a convective one such as a small positive angle of attack to produce an overall separation control. The status of the boundary layer, positive lift increase. Figures 17(a) and 17(b) show either laminar or turbulent, is assumed to be a minor typical @ distributions on the airfoil upper and lower effect on virtual shaping. Power spectra of pressure surfaces, respectively. Variations of lift increase with fluctuations obtained by dynamic pressure transducers, respect to the jet energy and the forcing frequency is such as the one shown in Figure 11, were used to verify compiled in Figure 18 with the actuator driven at the the boundary layer conditions. To investigate the effect maximum performance condition. The increase of lift of boundary-layer transition on virtual shaping, a 1.27 does not closely follow the peak of jet energy because American Institute of Aeronautics and Astronautics ACKNOWLEDGEMENT
the increase oflift issaturated at C_ _ 0.003, when U_
10 m/sec, as indicated in Figure 19. Variations of lift increase with jet momentum coefficient under different The authors would like to acknowledge the freestream velocities are shown in Figures 20(a) to assistance of Dr. Ponnampalam Balakumar of the 20(d) for data at forcing frequencies of 600 Hz and NASA Langley Research Center to provide numerical 1300 Hz only. Lift increase and jet momentum computational results used in this paper.
coefficient are correlated as a function of Reynolds REFERENCES number and presented in Figure 21 with two fitted lines, Y 230 + 72"1og Xand Y 260 + 72*logX, for forcing frequencies of 600 Hz and 1300 Hz, respectively. 1. Amitay, M., Smith, B. L., and Glezer, A., Where X and Y represent the lumped variables for the x "Aerodynamic Flow Control Using Synthetic Jet and y coordinates. Technology", AIAA Paper 98-0208, 1998.
The effect of virtual shaping on the airfoil pressure 2. Smith, D. R., Amitay, M., Kibens, V., Parekh, D., and Glezer, A., "Modification of Lifting Body drag coefficient, Co, is never greater than 0.002 under any circumstance of this investigation. Most of the time Aerodynamics Using Synthetic Jet Actuators", it is negligible therefore not included for discussion. AIAA Paper 98-0209, 1998.
3. Seifert, A., Eliahu, S., Greenblatt, D., and 5. SUMMARY Wygnanski, I., "Use of Piezoelectric Actuators for Airfoil Separation Control", AIAA Journal, Vol.
Aerodynamic virtual shaping of a two-dimensional 36, No. 8, pp. 1535-1537, 1998.
NACA 0015 airfoil were investigated in the NASA 4. Crook, A., Sadri, A. M., and Wood, N. J., "The Langley 2 ft x 3 ft Tunnel, using a synthetic jet actuator Development and Implementation of Synthetic Jets at different chordwise locations, different forcing for the Control of Separated Flow", AIAA Paper frequencies and amplitudes, under different freestream 99-3176, 1999.
velocities. Important characteristics of electric power 5. Amitay, M., Smith, D. R., Kibens, V., Parekh, D., consumption of the synthetic jet actuator were also and Glezer, A., "Aerodynamic Flow Control over reported. Results are summarized as follows: an Unconventional Airfoil Using Synthetic Jet 1. Synthetic jet energy is closely correlated to the Actuators", AIAA Journal, Vol. 39, No. 3, pp. 361- 370, 2001.
active power consumed by the actuator. The data hints at the potential of monitoring the active 6. Honohan, A. M., Amitay, M., and Glezer, A., power as a potential tuning tool to maintain the "Aerodynamic Control Using Synthetic Jets", excitation near resonance of the actuator.
AIAA Paper 2000-2401, 2000.
2. Performance of the synthetic jet actuator increases 7. Chatlynne, E., Rumigny, N., Amitay, M., and when the magnitude of reactive power and the Glezer, A., "Virtual Aero-Shaping of a Clark-Y phase lag of voltage to current are decreasing. Airfoil Using Synthetic Jet Actuators", AIAA 3. Synthetic jet actuation on the airfoil upper surface Paper 2000-0732, 2000.
causes a localized increase of surface pressure that 8. King, R. A., "Receptivity and Growth of Two- and Three-Dimensional Disturbances in a Blasius results in a negative lift to the airfoil. In the range of data tested, the maximum lift change was -0.015 Boundary Layer", Ph.D. thesis, Massachusetts Institute of technology, Cambridge, MA, 2000.
that resulted in a -6% change in AC l/Cl, u_ with actuation at 50% of chord.
9. Chen, F.-J., Yao, C., Beeler, G. B., Bryant, R. G., 4. Effect of virtual shaping is drastically decreased and Fox, R. L., "Development of Synthetic Jet Actuators for Active Flow Control at NASA when synthetic jet actuation is applied under separated flow. Langley", AIAA Paper 2000-2405, 2000.
5. The stagnation line is shifted by synthetic jet 10. Fitzgerald, E., Higginbotham, D. E., and Grabel, actuation near the airfoil leading edge, inducing an A., "Basic Electrical Engineering", Fifth Edition, effect similar to that caused by a small angle of McGraw-Hill Book Company, 1981.
attack to produce an overall lift change. 11. Hassan, A. A., "Numerical Simulations and Synthetic jets represent a breakthrough in actuator Potential Applications of Zero-Mass Jets for technology, but further development is required to Enhanced Rotorcraft Aerodynamic performance", substantially increase jet momentum output before AIAA Paper 98-0211, 1998.
applying them to real flight vehicles. 12. Rumsey, C., Biedron, R., and Thomas, J., "CFL3D: Its History and Some Recent Applications", NASA TM-112861, May 1997.
American Institute of Aeronautics and Astronautics (a)Side view ofairfoil model.
(b)Cut-away viewofairfoil model. (c)Photograph ofairfoil model onbench top.
Figure 1.NACA 0015 two-dimensional airfoil model forvirtual shaping test.
50 , , , i , , , i , , , i , , , i , , , 50 40 4O 30 3O UpperR o o o o o o S 20 2O LowerR 10 10 LowerC UppelC oooo o o o o o o o o 01 0 _oooo_O _o_ _o_o _o o -10 -10 UpperL -20 _bo o o o o o o -20 -30 -30 • Kulite • Endevco o ESP -40 Airfoil -40 , I , , I , , , I , , , I , -50 -50 20 40 60 80 100 0 20 40 60 80 100 x, cm x, cm (a) Airfoil upper surface. (b) Airfoil lower surface.
Figure 2. Distribution of ESP pressure ports and dynamic pressure transducers on the airfoil model surface.
Figure 3. Photograph of the two-dimensional synthetic jet actuator on bench top.
American Institute of Aeronautics and Astronautics 50 i i i i i i i i i i i i i i , i , i , i , i , i , i , i , i , i , i , i , i , i , 0 v_ear,
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Figure 6. Characteristics of electric power consumption and jet energy output at forcing frequencyf 600 Hz.
0.10 100 .... i .... i .... i .... i ....
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Figure 7. Characteristics of electric power consumption and jet energy output at forcing frequencyf 1300 Hz.
American Institute of Aeronautics and Astronautics 1.0 0.10 .... 80 700 ' ' ' Pl_ase_ngle' ' ' 'I'_' ' + (v_ 2 0.09 600 ...Fq-.. acdve P O. 9 • reactive Q/4 0.08 0.8 ,_ 0.07 0.7 _._ 400 50 g 0.06 0.6
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0.05 40 g "_ 0.5 _" 200 0.04 0.4 0.03 , , , , I , , , , I , , , , I , , , , 0.3 0.02 2O 0 500 1000 1500 2000 0 500 1000 1500 2000 Forcing Frequency f, Hz Forcing Frequency f, Hz (a) Electric current, voltage and phase angle. (b) Jet energy, active and reactive electric powers.
Figure 8. Variations of electric power consumption and jet energy output with synthetic jet forcing frequency.
-0.8 -0.6 -0.4 -0.4 -0.2 -0.2 0.0 0.0 0.2 @ 0.2 -- CFL3D 0.4_ 0.41 --CFL3D UppelC, 011 _LowelC, On ---V--- UpperC, Off ""_'"LowelC, Off 0.6 0.6 • UpperR, 011 • LowerR, On • UpperR, Off • LowerR, Off [X UpperL, Oll o.81 o.8_ LowerL, On /1 UpperL, Off Lowe_,Off 1.0 1.0 , , , I , , , I , , , I , , , I , , , ii,lll,lll,lll,lll, 0.0 0.2 0.4 0.6 0.8 1.0 0.0 0.2 0.4 0.6 0.8 1.0 x/c (a) @ on airfoil upper surface. (b) 6_ on airfoil lower surface.
Figure 9. Variations of Cp on airfoil surfaces with actuator located at x/c 0.5 on upper surface, operated at forcing frequencyf 600 Hz and freestream velocity U_ -_ 10 m/sec.
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Figure 11. Power spectra of pressure fluctuations on
Figure 10. Variations of spanwise 6_ at x/c 0.95
on airfoil upper surface.
airfoil upper surface for U_ -_10 m/sec.
American Institute of Aeronautics and Astronautics
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10 20 30 40 50 Forcing Frequency f, Hz U_, m/sec Figure 12. Effects ofbouaadary-layer transition Figure 13. Variations of airfoil lift change with on virtual shaping when the actuator forcing frequency when the actuator located at x/c 0.15 on upper surface. located at x/c 0.1 on upper surface.
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(a) Actuator at x/c 0.1. (b) Actuator at x/c 0.15.
1 ........ i ........ i ........ i ........ i ........ i ........ 1 ........ i ........ i ........ i ........ i ........ i ........
0 o %11o -1 • O _i o o <> -2 -z O f= 600 Hz, U_ ;, 10 m/sec o f= 600_,U_lOn_se¢ ,o o -3 • f= 1300Hz, U_ ;, 10m/sec f= 1300Hz, U_ _, 10 m/sec _2 <> f= 600Hz, U_;,20m/sec <> f = 600 Hz, U_ _, 20 m/sec -4 • f= 1300Hz, U_ ;, 20 m/sec -4 • f= 1300Hz, U_ _, 20 m/sec A f= 600Hz, U_;,30m/sec A f = 600 Hz, U_ _, 30 m/sec -5 -5 • f= 1300Hz, U_ ;, 30 m/sec • f= 1300Hz, U_ _, 30 m/sec [] f= 600 Hz, U_ _, 40 m/sec [] f = 600 Hz, U_ _, 40 m/sec -6 -6 • f= 1300Hz, U_ ;, 40 m/sec • f= 1300Hz, U_ _, 40 m/sec 8 _7 ' ....... t ........ t ........ I ........ I ........ I ..... _7 ........ t ........ t ........ t ........ i ........ i ......
'1
0 7 10 6 10 5 10 4 10 3 10 2 10 1 0 7 10 6 10 5 10 4 10 3 10 2 10 1 c. c.
(c) Actuator at x/c 0.3. (d) Actuator at x/c 0.5.
Figure 14. Variations of airfoil lift change with jet momentum coefficient when the actuator located at different chordwise locations on the airfoil upper surface.
American Institute of Aeronautics and Astronautics 1 ........ i ........ i ........ i ........ i ........ i ........
o -1 -1 -2 -2 0 O f= 600 Hz, U_ _, 10 m/sec _" -3 • f= 1300Hz, U_ _, 10 m/sec • f= 1300Hz @ x/c= 0.1 Q_ C' f= 600 Hz, U_ _, 20 m/sec O f= 600Hz@x/c=0.15 "_ -4 • f= 1300Hz, U_ _, 20 m/sec -4 • f= 1300Hz @ x/c= 0.15 ZX f= 600Hz@x/c=0.3 L_ ZX f= 600 Hz, U_ _, 30 m/sec -5 • f= 1300Hz @ x/c= 0.3 -5 • f= 1300Hz, U_ _, 30 m/sec _2 [] f= 600Hz@x/c=0.5 -3 Of= 600 Hz @ x/c= 0.1 [] f= 600 Hz, U_ _, 40 m/sec • f= 1300Hz @ x/c = 0.5 -6 -6 Y = -17 * X 025 • f= 1300Hz, U_ _, 40 m/sec _ , ....... i ........ i ........ i ........ i ........ i .....
0 10 6 10 5 10 4 10 3 10 2 10 1 -_0 11' "1';:10' '"1'; 9'' '"1'; 8'' '"1'; 7' '"i'; 6' '"i'; 5' '"i'; 4'' "10 3 c'_/ l o_(Re_/6ooooo) Figure 15. Variations of airfoil lift change with jet Figure 16. Correlation of airfoil lift change with jet momentum coefficient when the actuator momentum coefficient for data presented located at x/c 0.75 on upper surface. in Figures 14(a)_(d).
-0.6 ............
-0.2 0.0 0.2
o
0.4 -- CFL3D --CFL3D UpperC, On --z2r-- LowerC, On ---_7--- LowerC, Off ---_--- UpperC, Off 0.6 • UpperR, On • LowerR, On • LowerR, Off • UpperR, Off 0.8 jl Ix UpperL, On IX LowerL, On Z] UpperL, Off Z] LowerL, Off 1.0_ , , , I , , , I , , , I , , , I , , , 0.0 0.2 0.4 0.6 0.8 1.0 0.4 0.6 0.8 x/c x/c (a) @ on airfoil upper surface. (b) @ on airfoil lower surface.
Figure 17. Variations of @ on airfoil surfaces with actuator located at x/c -0.015 on upper surface, operated at forcing frequencyf 600 Hz and freestream velocity U_ -_ 10 m/sec.
7OO 0.07 0.05 .... i .... i .... i .... i .... i .... i .... i ....
6OO 0.06
o o4iooze
5OO 0.05
Oo ° oo°o o °
0.03
°o oa8 o-
4OO 0.04 o L7 o eq _ 3OO 0.03 0.02
@
,b 2OO 0.02 o O.Ol O IO0 0.01 _,,,,i .... i .... i .... i .... i .... i .... i,,, 0 0.00 0 500 1000 1500 2000 0 0.001 0.002 0.003 0.004 0.005 0.006 0.007 0.008
Forcing Frequency f, Hz c3
Figure 18. Variations of jet energy and airfoil lift change Figure 19. Variations of airfoil lift change with jet momentum coefficient when the actuator with forcing frequency when the actuator located atx/c -0.015 and U_ -_ 10 m/sec. located at x/c -0.015 and U_ -_ 10 m/sec.
American Institute of Aeronautics and Astronautics 0.05 ........ , ........ , ........ , ........ , ........ 0.05 ........ , ........ , ........ , ........ , ........
0.04 O f= 1300Hz O f= 600Hz I
004i o f=13OO l
lof ooHz
0.03
003 S 'l
C _2
0.02
o 44F
d_
,Oo O.Ol 0 O.Ol 0 o -0.01 ........ ' ........ ' ........ ' ........ ' ......
- - 0 " 0 4 0 1 7 . . . . . . . . l . . . . . . . . l . . . . . . . . l . . . . . . . . l . . . . . . .
10 7 10 6 10 5 10 4 10 3 10 2 10 6 10 5 10 4 10 3 10 2 (a) U_ -_ 10 m/sec. (b) U_ -_ 20 m/sec.
0.05 ........ i ........ i ........ i ........ i ........ 0 , 0 5 ........ i ........ i ........ i ........ i . . . . . . . .
0.04 0.04
O f= I f= 1300Hz
O f= 600Hz I f= 1300Hz 0.03 0.03
_2 _2
0.02 0.02 0.01 0.01 O 0
ooo °
69oO
0 _o, -0.0_ ........ l ........ l ........ l ........ l ...... -0"0 "I 01 7 ........................ I ........ I .......
0 7 10 6 10 5 10 4 10 3 10 2 10 6 10 5 10 4 10 3 10 2 (c) U_ -_ 30 m/sec. (d) U_ -_ 40 m/sec.
Figure 20. Variations of airfoil lift change with jet momentum coefficient when the actuator located at x/c -0.015 on airfoil lower surface.
IO0 ,£ 0 f= 600Hz, U_ 10mJsec % • f= 1300Hz, U_ 10 mJsec O f= 600Hz, U_ 20m/sec _2 • f= 1300Hz, U_ 20 nl/sec 4O ZX f= 600Hz, U_ 30m/sec • f= 1300Hz, U_ 30 m/sec 2O [] f= 600Hz, U_ 40m/sec • f= 1300Hz, U_ 40 m/sec -- = 600Hz,Y=230+72*log f 1300Hz,Y=260+72*log_ ..... f= O .....
........ i ........ i , ,, -2_0 5 10 4 10 3 10 2 10 1 100 * r, 0 75 c'_ lo (ReE600000) Figure 21. Correlation of airfoil lift change with jet momentum coefficient when the actuator located at x/c -0.015 on airfoil lower surface.
American Institute of Aeronautics and Astronautics