When attempting dynamic stall control, compressibility must be seriously considered because typical full-scale Mach numbers on a rotorcraft retreating blade in ...
AIAA 2001-0734 th 39 Aerospace Sciences Meeting & Exhibit 12-15 January, 2001, Reno, NV
Enhanced Performance of Airfoils at Moderate Mach Numbers Using Zero-Mass Flux Pulsed Blowing Michael Hites, and Hassan Nagib, Illinois Institute of Technology, Chicago, IL Tomer Bachar and Israel Wygnanski* Tel Aviv University, Israel
Abstract Oscillatory wall-jets were introduced through spanwise slots along a flapped NACA 0015 airfoil to establish lift augmentation and drag reduction by the unsteady forcing of the separated flow. Pressure coefficient distributions, lift coefficients, and wake velocity profiles, to determine the drag coefficient, were measured over the test-section speed range of 25m/s < U∞ < 140m/s in the NDF. The present results demonstrated for the first time (Hites; 1997) the effectiveness of the oscillatory blowing technique as a separation control scheme at moderate Mach numbers, which exhibit compressibility effects. It is encouraging that lift-enhancement was observed over the entire range 0.1 < M < 0.4, even with the small amount of unsteady blowing applied in these experiments. As a result of the pulsed blowing, the lift coefficient increased by as much as 80%. Maximum pressure coefficients of nearly -5.0 for M = 0.4 experiments indicated the flow was supercritical near the leading edge of the airfoil, whereas it was not before the application of oscillatory blowing. The improvement in lift coefficient was found to be sensitive to the forcing frequency, even at the higher Mach numbers. Measurements at low angles of attack with a 20° flap showed that low amplitude pulsed blowing (0.02%) from the flap provided a 27% increasing in lift while steady blowing from the flap was detrimental to lift even at blowing coefficients as high as 3.5%. Oscillatory blowing with coefficients between 0.01% and 0.02%, based on RMS velocity, was shown to yield substantially better performance than steady blowing with Cµ in the range 0.5% to 3.5%. In is estimated that steady blowing of at least 10% would be required to reach the same levels of lift coefficient seen with the oscillatory blowing.
In general, oscillatory blowing has proven to be a robust means of separation control, consistently showing improvement in lift/drag ratio, while hysteresis does not appear to diminish the effectiveness of the technique. Again it appears very promising based on the present wind tunnel tests that the pulsed-blowing active flow technique should be effective for at least some airfoils oscillating in pitch and operating within the compressible regime.
Introduction and Background In modern design of military aircraft, separation control is vital to improving the flight characteristics of airfoils whether the application is highly maneuverable fighters, stealth bombers, or micro air vehicles. When air separates from a wing in flight, the result is loss of lift and increase in drag that threatens the stability of the aircraft and the safety of the pilot. Separation is typically avoided by geometric changes and by flying the aircraft within the flight envelope. Alternatives to fixed-geometry aircraft are flexible structures, where the wings and fuselage are constructed from materials that can be deformed to adapt to changing flow conditions, and flow control may be achieved through solid or fluidic actuators. In either case, the flow field is changed globally by locally acting on the flow. This may be achieved by arrays of actuators that energize the boundary layer with steady or pulsed momentum through the wing surface. These types of controlled actuation change the velocity, pressure, and vorticity field around the wing to achieve the desired objectives.
* Also at University of Arizona, Tucson, AZ
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It has been observed that relatively large quantities of steady blowing (Cµ = 10%) near the point of separation can reattach the flow and increase lift, but the steady blowing may also cause a thickening of both the boundary layer and the wake behind the airfoil which leads to increased drag. In contrast to steady blowing, the oscillatory blowing takes advantage of inherent local instabilities in the near-wall shear layer that causes the selective amplification of the input oscillation frequency. These amplified disturbances convect downstream along the airfoil as coherent large structures that serve to mix the boundary layer flow and delay separation. The efficiency of the mixing provides substantial increases in lift, while concomitantly reducing drag even when one assumes that the entire momentum added recovered as thrust. The unsteady forcing can also be tailored to affect some global instabilities of the separated flow and lead to reduction in the size of the region of separation and improved performance of the wing. Pulsed blowing has proven to be a reliable technique for separation control. Recent work by Seifert et al (1993, 1996, 1998, 1999), Wygnanski (1997), and Hites et al (1997) has shown repeatedly that low amplitude, oscillatory blowing can delay separation and enhance lift over a wide range of Reynolds numbers including those corresponding to aircraft takeoff and landing. These experiments have demonstrated several consistent results, including: 1) the most effective location for unsteady forcing is near the point of separation, 2) the optimum reduced frequency for the oscillations is about f+ = fc/U ≈1, and 3) the amplitude of the oscillations required for effective separation control is about two orders of magnitude lower than that for steady blowing. When oscillatory separation control is applied to an airfoil that is not separated, the effects are not detrimental to lift. McManus and Magill (1997) performed experiments on a 3-D, finite lambda wing model and demonstrated the effectiveness of pulsed vortex jets up to M = 0.2. The improvement in the lift coefficient was less then that of Seifert et al.; however, the method and the amplitude of the blowing were different. In the McManus and Magill experiments, four round jets were used to pulsate the flow locally just downstream of the leading edge, whereas the work by Seifert used a linear slot over the entire span of the airfoil at essentially x/C=0. McManus introduced streamwise vorticity into the boundary layer while Seifert introduced an unsteady vorticity with a predominantly spanwise component. In a numerical investigation by Towne and Buter (1994), subharmonic pulsed blowing applied from the
leading edge in an M = 0.2 simulation. Their results showed that the oscillations were ineffective at controlling dynamic stall, whereas steady blowing was effective. This contradicts recent experimental findings with pulsed blowing that demonstrate more efficient prevention of separation than steady blowing in airfoils. Although the Mach number was similar to the experimental studies, the chord Reynolds number was about three orders of magnitude lower than the experimental work of Seifert. Also the dimensionless blowing frequency, f+, was about 10-3 in Towne and Buter’s study, several orders of magnitude below the optimal pulsing frequency of other researchers. A comprehensive computational model has yet to be developed to describe the effects of oscillatory blowing on stall, although Wu et al. (1997) have begun to propose schemes which address non-linear mode interaction, resonance, and vortex dynamics. Very recently, extensive numerical simulations by various groups have achieved more success in representing the various flow modules of such complex flow, e.g., see Hassan and Munts (2000). Recently, periodic excitation has been demonstrated as an effective, efficient and practical method for controlling incompressible dynamic stall (e.g. Greenblatt & Wygnanski, 1999). Based on these results it is obvious to expect that the technique may also be effective in improving the performance of a wide range of airfoils used in the rotorcraft industry such as those carefully documented by McAlister et al. (1982). Carr (1988) shows, however, that compressibility can have a profound effect on dynamic stall, even at relatively moderate Mach numbers, i.e. M = 0.3, when the flow can be supersonic in the leadingedge region. Although in Carr’s case the airfoil stall was dominated by leading edge separation, where compressibility effects may be exaggerated by a shock boundary layer interaction, Carr and Chandrasekhara (1996) projected that such effects would lead to the failure of flow control methodologies in all unsteady airfoil applications for Mach numbers equal to or larger than about 0.3. When attempting dynamic stall control, compressibility must be seriously considered because typical full-scale Mach numbers on a rotorcraft retreating blade in the vicinity of dynamic stall are in the range from 0.3 to 0.5. (Apparently, the effect of Reynolds number is less understood due to the difficulty of varying Reynolds number significantly without introducing compressibility effects.) The review by Carr and Chandrasekhara (1996) shows that no definitive conclusions can be drawn concerning the impact of the supersonic flow in the leading-edge region. One of the problems is that various investigators report different findings for
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experiments at nominally the same conditions on identical airfoils. Nevertheless, two main conclusions are apparent: Firstly, there is a limit to the magnitude of the suction peak that can be reached at high incidence in compressible flow and this limit is due to the local supersonic velocity on the surface of the airfoil. Secondly, compressibility effects can completely change the flow behavior compared to that observed at low Mach numbers (M < 0.3). The first attempt made to carefully and independently examine the conclusion by Carr and Chandrasekhara (1996), regarding the detrimental effects of compressibility on control of separation, and to examine the effectiveness of pulsed jets in the enhancement of airfoil performance at Mach numbers with compressibility effects, is contained in the thesis of Hites (1997). These experiments demonstrated effective control up to Mach numbers above 0.4. More recently, Seifert & Pack (1999) investigated the effect of periodic excitation on a static NACA 0015 airfoil under the conditions 0.28 ≤ Ma ≤ 0.55 . For all cases, excitation had a beneficial effect on lift and drag, with the most significant effects in the post-stall regime. In fact, Seifert & Pack showed that upper surface suction that is attenuated as a result of an increase in Mach number from 0.28 to 0.4 could be restored by excitation at low perturbation amplitudes ( Cµ < 0.1% ). On the other hand, Greenblatt et al. (1999) have shown that the net result of excitation in incompressible flows is essentially insensitive to whether an airfoil is stationary or oscillating in pitch. This is primarily due to the large disparity between the time-scales characterizing the airfoil pitch oscillations and those characterizing the excitation-generated large coherent structures (LCSs). It can therefore be speculated that periodic excitation is also effective in controlling dynamic stall under compressible conditions, even at low perturbation amplitudes.
Objectives This work extends the results reported by Hites (1997) and focuses on demonstrating the effectiveness of pulsed blowing at higher Reynolds and Mach numbers. The NACA 0015 model used here in experiments with M > 0.4, is the same model used by Siefert et al. (1993) in pulsed blowing experiments up to M = 0.15. It is well known that significant changes in the characteristics of stall occur due to compressibility for M > 0.3. For lower, incompressible Mach numbers, the stall behavior is initiated by trailing-edge stall where the stalled region migrates upstream until the
separated region covers nearly the entire chord. In contrast, for higher compressible Mach numbers (i.e., M > 0.3), leading-edge separation may be a precursor to complete stall. The leading-edge stall occurs abruptly as the airfoil is pitched beyond a critical angle of attack, while the former, trailing edge stall, evolves gradually as the angle of attack is increased. This work was aimed at assessing the effectiveness of the oscillatory-blowing flow control technique under both of these stall conditions.
Experimental Apparatus A NACA 0015 airfoil with a 24” span was mounted vertically in the NDF test section as shown in Figures 1 and 2. The airfoil could be pitched through ±30° and the flap could be adjusted from ±30° in 10° increments. The airfoil assembly is shown in Figures 3 and 4 along with the locations of the spanwise oscillatory blowing at 0%, 10%, and 75% chord. Details of the airfoil support design are available in Peterson (1995). The airfoil was tested in the National Diagnostic Facility at IIT, shortly after its completion. More details about the facility are available in Hites’ thesis (1997) as well as in Nagib et al. (1994) and Nagib and Hites (1994) The NACA 0015 airfoil had a single row of 36 static pressure taps around the perimeter of the airfoil at the center of the span. Each of the static pressure taps was connected with 0.063" vinyl tubing to a 48 port J9 Scanivalve. A Pitot-static probe located in the free stream provided the reference pressure for each pressure tap, and each pressure tap was sampled sequentially using a Validyne DP-103 pressure transducer calibrated to 1990 Pa/V. Oscillatory blowing was supplied via a compressed air source with a stagnation pressure of approximately 10 atmospheres. The amplitude of the of the blowing was modulated by a regulator, and the frequency of the blowing was controlled by computer with a variable frequency drive and 3 HP motor which drove a variable speed "daisy-wheel" valve. This valve was designed and supplied by Tel Aviv University and consisted of two stationary disks with a rotating disk between them. The mean and oscillatory components of the slot blowing were measured by a hot-wire at the exit of the leading edge slot and calibrated to the static pressure of the compressed air supply. Figure 4 shows the location and direction of the blowing relative to the airfoil cross-section. To obtain the total drag, the wake profile was measured by a fixed Pitot rake with one hundred 0.040” diameter (OD) tubes located at x/c = 2 behind the airfoil trailing edge. The pressure at each port was measured by a single Setra 205 pressure
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transducer calibrated to 3700 Pa/V with the help of the Scanivalve.
Results The amplitude of steady and/or pulsed blowing is characterized by the momentum coefficient h u Cµ = 2 c U ∞
2
where h is the height of the blowing slot, c is the chord of the control surface, and u is either the mean (U) or rms (u’) velocity depending on whether steady or pulsed blowing is used. Pulsed blowing of amplitude = 0.01% to 0.02% (±0.005%) was used for separation control Mach numbers ranging from 0.07 to 0.41. Steady blowing of up to Cµ = 3.5% was used for comparison to the oscillatory blowing. The pressure distributions of Figure 5 show the typical effects of oscillatory blowing on a stalled airfoil for the low Mach number of 0.18. In Figure 6 similar results are presented for the same airfoil at M= 0.32 with pulsed blowing form the leading edge slot. The airfoil was initially stalled in the unforced case, and two scenarios for the application of the pulsed blowing were used and are compared to the baseline case in the figure to examine any possible hysteresis effects. The application of oscillatory blowing from the leading edge results in the re-establishment of the high suction peak, good pressure recovery along the airfoil, and a trailing edge pressure that more closely matches the Kutta condition. The outcome of these changes in pressure distributions is that the lift coefficient is increased by about 80% over the stalled airfoil lift coefficient. Figure 6 also shows that the improvement in lift is independent of the manner in which the forcing is applied. The square symbols (α sweep) were obtained when the forcing initiated at α = 0° then slowly pitching the airfoil to α = 14.5°, which was stalled in absence of forcing. The diamond symbols are the Cp values when the forcing is turned on instantaneously while the airfoil is stalled at α = 14.5°. The pressure distribution is identical for both of the methods, indicating that hysteresis is not a major factor in the effectiveness of separation control by this technique. At the highest Mach number tested, M = 0.41, the pressure coefficients and the lift coefficient were measured for 0° flap angle with and without f+ = 2.8 (840 Hz) introduced at leading edge slot. Figure 7 shows re-attachment of the flow and an increase in lift
of 50% over the baseline-stalled airfoil. The maximum pressure coefficient, -Cpmax, of nearly 5.0 while the flow control is applied indicates that the flow is supercritical near the leading edge, whereas it was not before the application of oscillatory blowing. This is a clear indication of the effect of the control method on improving the flow attachment near the leading edge, and hence the improved lift values. Similar results are presented in Figure 8 demonstrating the effectiveness of the technique under the compressible flow conditions not only with no flap deflection, but also under the flap setting of 10 degrees. The lift capabilities of the airfoil are summarized for two Mach numbers in Figure 9, where it is clear that similar effects of the unsteady blowing are achieved both at low and moderate Mach numbers. Nearly identical results were obtained at 0.41 Mach number and are given by Hites (1997). The effects of the dimensionless forcing frequency, f+, were examined by measuring the change in lift coefficient, ∆Cl/Cl, as a function of Mach number and the corresponding f+. Figure 10 indicates the sensitivity to forcing frequency over 0.1 < M < 0.42 with constant amplitude forcing at a single frequency of 840 Hz. A nearly constant lift increment of about ∆Cl/C = 0.85 was observed for the range 2.5 < f+ < 4. Outside of this range, the effect of the pulsed blowing was diminished. Another experiment compared the effect of oscillatory blowing under similar stall conditions over the range 0.17 < M < 0.41 where the initial “depth into stall” was kept constant at each Mach number before the application of oscillatory blowing; i.e., α − α stall = constant. In Figure 11, the baseline Cl was maintained constant over the full range of Mach numbers by varying the post-stall AOA for each Mach number until Cl was the same for each velocity. Then, the oscillatory blowing was applied at 270Hz or 840Hz at a constant momentum coefficient Cµ = (0.02; 0.02)%; i.e., consisting of equal steaduy and oscillating components. This resulted in producing a nearly constant increase in Cl for the 270Hz forcing and an f+ dependent increase in Cl for the 840Hz case, as was demonstrated in Figure 5. The peak in the 840 Hz blowing case does not appear in the 270Hz measurements because the Mach number corresponding to the frequency of maximum efficiency (f+= 2.9) is about 3 times lower than that of the 840Hz measurements, which is outside of the range of the plot in Figure 11. Experiments were also performed over the range 0.07 < M < 0.3 with the flap set to 20° with the airfoil at low angle of attack, which was below that of complete stall. In contrast to the previous
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measurements, the focus here was on the separation over the flap, not the leading edge. In addition to pressure and lift coefficients, the wake velocity profile and the total drag were measured using a Pitot tube rake located two chord lengths behind the airfoil. In these experiments the forcing was introduced from a slot near the leading edge of the deflected flap. Figures 12 and 13 compare the pressure distributions and the wake profiles for the baseline (unforced) case to cases with steady or pulsed blowing when δ = 20° and α = -2°. It was observed that pulsed blowing at = 0.01 produced an increase in lift of 27% over the baseline case and that steady blowing from the same location reduced lift by as much as 15%. Figure 12 shows that the pressure distribution over the entire airfoil, and not just over the flap, was enhanced due to the oscillatory blowing from the flap. In contrast, steady blowing attenuated the pressure distribution. Figure 13 shows the wake profiles that correspond to the pressure and lift coefficients of Figure 12 and demonstrate that drag is reduced due to the pulsed blowing and that steady blowing increased drag. Figures 14 and 15 show the lift and drag data for similar experiments at M = 0.2; however the magnitude of the steady blowing is seven times greater than that of the M = 0.07 experiments. The pressure distributions in Figure 14 show that the high levels of steady blowing (Cµ = 3.5%) are not detrimental to the lift as they were for Cµ = 0.5% for M = 0.07. Nonetheless, the steady blowing did not show any improvements to the pressure coefficients. Pulsed blowing at = 0.01% achieved the same effect at M = 0.2 as at 0.07, namely the pressure distribution was enhanced over the entire airfoil, and the lift coefficient increased due to the application of the oscillatory separation control. Finally, the pressure on the airfoil and wake velocity profile were measured for δ = 20° and α = -2° at M = 0.3 using steady blowing at both 0.5% and 3.5% and pulsed blowing at 0.02%. Figure 16 demonstrates that the improvement in the pressure and lift coefficients is independent of the Mach number range examined. The effects of the steady blowing are consistent with the lower Mach number cases, that is, it is detrimental to lift when the steady forcing is applied from the flap. The wakes are also similar for all Mach numbers, as shown in Figure 17. The velocity deficit is smallest for the experiments where oscillatory forcing is on. The lift/drag ratios for the M = 0.3 experiment are summarized in Table 1 and show that the application of oscillatory blowing at the flap doubles the lift/drag ratio over the basic airfoil. Table 1 clearly
demonstrates the effectiveness of pulsed blowing as a tool to increase lift and reduce drag, especially when compared to the relative ineffectiveness of steady blowing under similar conditions.
Conclusions and Recommendations The present results demonstrated for the first time the effectiveness of the oscillatory blowing technique as a separation control scheme at moderate Mach numbers, which exhibit compressibility effects. It is encouraging that lift-enhancement was observed over the entire range 0.1 < M < 0.4, even with the small amount of unsteady blowing applied in these experiments. A primary benefit of oscillatory blowing compared to steady blowing is that at least two orders of magnitude less momentum coefficient are required. Steady blowing has been used for separation control, but in general it is an expensive method because high velocity bleed is required for the technique to be beneficial. In our measurements, oscillatory blowing with coefficients between 0.01% and 0.02%, based on RMS velocity, was shown to yield substantially better performance than steady blowing with Cµ in the range 0.5% to 3.5%. In is estimated that steady blowing of at least 10% would be required to reach the same levels of lift coefficient seen with the oscillatory blowing. The observation that supercritical flow was achieved due to the application of the oscillatory blowing demonstrates that the technique can be extended successfully to higher Mach numbers. Similar measurements have been performed up to M = 0.15 by Seifert et al (1993, 1996), and up to M = 0.55 by Seifert and Pack (1999), which documented the benefits of pulsed blowing separation control. Hysteresis appears not to be an important factor in oscillatory forcing, and shows that the technique may be practical for maneuvering high-speed aircraft. It was demonstrated that the separation control was independent of the sequence of its application, that is, whether the airfoil was stalled or not before applying the forcing. This insensitivity to time of application is encouraging for use on real aircraft for both cruise and emergency conditions. It also strongly suggests that the technique is likely to be effective for rotorcraft applications with periodically oscillating airfoils. The recent work of Greenblatt et al. (1999) confirms this speculation. This leads us to recommend that this active flow control should be evaluated for oscillating airfoils under the Mach number range of interest to the rotorcraft applications between 0.3 and 0.6. Contrary to the earlier conclusions of Carr and Chandrasekhara
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(1996), it appears very promising based on the present wind tunnel tests that the pulsed-blowing active flow technique should be effective for at least some oscillating airfoils operating within the compressible regime.
Acknowledgments The authors greatly appreciate the support of the Air Force Office of Scientific Research (AFOSRF49620-96-1-0459) monitored by Drs. Jim McMichael and Mark Glauser. Thanks also are due to Dr. Lawrence Carr of NASA Ames for his assistance with preliminary high-speed measurements and tuft visualization of the airfoil.
References Carr, L. W., “Progress in the analysis and prediction of dynamic stall” AIAA Journal of Aircraft, Vol. 25, No. 1, 1988, pp. 6-17. Carr, L.W. and Chandrasekhara, M.S., "An Assessment of the Impact of Compressibility on Dynamic Stall," AIAA-95-0779, 33rd Aerospace Sciences Meeting, Reno, NV, 1995. Carr, L. W. and Chandrasekhara, M. S., “Compressibility effects on dynamic stall”, Progress in Aerospace Sciences, Vol. 32, pp. 523-573, 1996. Gault, D.E., "A Correlation of Low-Speed, Airfoil-Section Stalling Characteristics With Reynolds Number and Airfoil Geometry," NACA TN 3963, March 1957. Greenblatt, D. and Wygnanski, I. “Parameters affecting dynamic stall control by oscillatory excitation”, AIAA Paper 99-3121, 17th AIAA Applied Aerodynamics Conference, Norfolk, VA, 28 June – 1 July 1999. Greenblatt, D., Darabi, A., Nishri, B. and Wygnanski, I. “Some factors affecting stall control with particular emphasis on dynamic stall”, AIAA Paper 99-3504, 30th AIAA Fluid Dynamics Conference, Norfolk, VA, 28 June – 1 July, 1999. Hassan, A. and Munts E. “Transverse and Near Tangent Synthetic Jets for Aerodynamic Flow Control,” AIAA Paper 2000-4334, Applied Aerodynamic Conference, Denver, CO, 14 – 17 August 2000. Hites, M.H., Scaling of High-Reynolds-Number Turbulent Boundary Layers in the National Diagnostic Facility, Ph.D. Thesis, Illinois Institute of Technology, 1997.
Mazanec, M.J., Design of Research Equipment for Use in the National Diagnostic Facility, M.S. Thesis, Illinois Institute of Technology, May 1996. McAlister, K. W., Pucci, S. L., McCroskey, W. J. and Carr, L. W. “An experimental study of dynamic stall on advanced airfoil sections. Volume 2. Pressure and force data” NASA TM 84245, 1982. McManus, K. and Magill, J., “Airfoil performance enhancement using pulsed jet separation control”, AIAA-1971, 1997. Nagib, H.M., and Hites, M.H., "Measurement of Disturbance Levels in the National Diagnostic Facility", AIAA 94-0770, 32nd Aerospace Sci. Mtg., Reno, NV, 1994. Nagib, H., Hites, M., Gravante, S., and Won, J., "Flow Quality Documentation of the National Diagnostic Facility," AIAA 94-2499, 18th AIAA Aerospace Ground Testing Conference, 1994. Peterson, B.A., Design of Experimental Equipment for Use in the National Diagnostic Facility, M.S. Thesis, Illinois Institute of Technology, December 1995. Seifert, A., Bachar, T., Koss, D., Shepshelovich, M. and Wygnanski, I., "Oscillatory Blowing: A Tool to Delay Boundary-Layer Separation," AIAA Journal, Vol. 31, No. 11, p. 2052, 1993. Seifert, A., Darabi, B. Nishri, B, and Wygnanski, I, “The effects of forced oscillations on the performance of airfoils”, AIAA 93-3264, 1993. Seifert, A. and Pack, L.G., “Oscillatory control of separations at high Reynolds numbers”, AIAA 980214, 1998. Seifert, A, Darabi, A., and Wygnanski, “Delay of airfoil stall by periodic excitation”, Journal of Aircraft, Vol. 33, No. 4, 1996. Seifert, A. and Pack, L. G., “Oscillatory excitation of unsteady compressible flows over airfoils at flight Reynolds numbers” AIAA Paper 99-0925, 37th Aerospace Sciences Meeting and Exhibit, Reno, NV, Jan. 11-14, 1999. Wygnanski, I. and Seifert, A., "The Control of Separation by Periodic Oscillations," AIAA-94-2608, 18th AIAA Aerospace Ground Testing Conference, Colorado Springs, CO, 1994. Wygnanski, I, Boundary Layer and Flow Control by Period Addition of Momentum (Invited)”, AIAA 97-2117, 1997. Wu, J-Z., Lu, X-Y., and Wu, J-M., “Post-stall lift enhancement on an airfoil by local unsteady control”, AIAA 97-2064.
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Figure 1. NDF Test-Section Arrangement for Testing Airfoils at Moderate Mach Numbers.
Figure 2. Schematic of Experimental Setup and Instrumentation.
Configuration
Cµ or (%)
Cl
Cd
L/D
Basic Steady blowing Pulsed blowing
0 3.5 0.02
0.52 0.46 0.66
0.062 0.051 0.038
8.4 9.0 17.4
Table 1: Comparison of lift/drag ratio for the basic airfoil, steady blowing from flap, and pulsed blowing from flap at δ = 20° and α = -2° and M = 0.3. 7 American Institute of Aeronautics and Astronautics
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Figure 5. Pressure Distribution on Airfoil With and Without Active Flow Control at M = 0.18. Figure 3. Photograph of Model in Test Section.
Figure 4. Schematic of Hardware to Support and Pitch the Airfoil Model and Sketch of Flow Control Scheme.
Figure 6. Pressure Distribution on Airfoil With and Without Active Flow Control at M = 0.32.
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Figure 9. Lift Performance of NACA Airfoil With and Without Active Flow Control at M = 0.18 and 0.31. Figure 7. Pressure Distribution on Airfoil With and Without Active Flow Control at M = 0.41.
Figure 10. Optimum Lift Enhancement as a Function of Mach Number With Varying Forcing Frequency.
Figure 8. Pressure Distribution on Airfoil With and Without Active Flow Control at M = 0.41 in Presence of Deflected Flap.
Figure 11. Lift Performance as a Function of Mach Number With Fixed Forcing Frequency.
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Figure 12. Pressure Distribution on Airfoil With and Without Active Flow Control at M = 0.07 In Presence 20-Degree Flap Deflection.
Figure 13. Wake Surveys for Airfoil With and Without Active Flow Control, at M = 0.07, In Presence 20-Degree Flap Deflection.
Figure 14. Pressure Distribution on Airfoil With and Without Active Flow Control at M = 0.2 In Presence 20-Degree Flap Deflection.
Figure 15. Wake Surveys for Airfoil With and Without Active Flow Control, at M = 0.2, In Presence 20-Degree Flap Deflection.
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Figure 16. Pressure Distribution on Airfoil With and Without Active Flow Control at M = 0.07 In Presence 20-Degree Flap Deflection.
Figure 17. Wake Surveys for Airfoil With and Without Active Flow Control, at M = 0.3, In Presence 20-Degree Flap Deflection.
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