3 0 AIAA-91-0252 Effects of a Strake on the Flow Past a Wing-Body Junction W.J. Devenport, R.L. Simpson, M.B. Dewitz, and N.K. Agarwal Virginia Polytechnic lnstitute and State University, Blacksburg, VA

29th Aerospace Sciences Meeting , January 7-1 0, 1991 /Reno, Nevada

For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics 370 L'Enfant Promenade, S.W. Washington, D.C. 20024

EFFECTS OF A STRAKE ON THE FLOW PAST A WING-BODY JUNCTION

William J. ~even~ort*, Roger L. Simpsonff, Michael B. ~ewitz+ and Naval K. ~garwal*, Virqinia Polytechnic Institute and State University, 0

Abstract Pressure and velocity measurements and oil-

flow visualizations are presented to demonstrate the effects of a strake on the flow of a turbulent boundary layer past an idealized wing-body junction. The strake, a large fairing in the corner between the wing nose and body surface, modifies the flow in a way that is desirable in many applications. With the wing at zero angle of attack it eliminates leading-edge separation and thus the formation of a horseshoe vortex around the wing nose. At non-zero angle of attack it is less effective, leading-edge separation occuring on the .pressure side of the strake. However, the strake still eliminates much of the unsteadiness and flow non-uniformity associated with the horseshoe vortex.

Figure 1. The baseline wing.

Introduction This work forms the second part of an

experimental study of passive devices for altering or eliminating the horseshoe vortex formed in the flow past a wing-body junction.

The wing-body junction in question is shown in figure 1. It consists of an unswept cylindrical wing mounted on a flat wall on which an otherwise two-dimensional equilibrium turbulent boundary layer is growing. This type of junction creates a horseshoe vortex contained within a region of separation adjacent to the wing (figure 2). Upstream of the wing nose the vortex is formed by the separation and roll up of the boundary layer in the adverse pressure gradient imposed by the wing. Unsteadiness in the roll-up process generates large- scale low-frequency motions in the vortex'. These motions, which may be observed through the bimodal (double-peaked) velocity histograms they produce, are associated with locally high turbulence stresses and intense skin-friction and surface-pressure f lu~tuations'~~~~. The development of the vortex

pept. of Aerospace and Ocean Engineering ** Assistant Professor, Senior Member AIAA Jack E Cowling Professor, Associate Fellow AIAA +~raduate Assistant *~esearch Associate, Senior Member AIAA

Copyright 0 by W J Devenport, R L Simpson, M B Dewitz and N K Agarwal. Published by the American Institute of Astronautics and Aeronautics, Inc. with permission.

- - 24061, U.S.A.

downstream of the nose region is controlled by pressure gradients, turbulence diffusion and its entrainment of vorticity from the surrounding boundary layer. Much of this vorticity is oriented in the streamwise direction (because of the skewing of the boundary layer by cross-stream pressure gradients on either side of the wing) and thus tends to strengthen the vortex. The favourable pressure gradient experienced by the vortex between the nose and maximum thickness of the wing also tends to sustain its mean rotation and structure. The bimodal unsteadiness, however, decreases in intensity here4. Between the maximum thickness and the trailing edge, adverse pressure gradients, acting in concert with turbulence diffusion, dissipate the bimodal unsteadiness altogether. They also cause a rapid growth of the vortex and a slowing of its mean rotation rate. Although weakened, the legs of the horseshoe vortex persist far downstream of the trailing edge contributing substantially to the unsteadiness and non-uniformity of the wake. Along its entire length the mean motion of the vortex tends to bring high momentum fluid from the edge of the boundary layer into close contact with the wall. This increases the surface shear stresses adding to the drag and, when thermal effects are important, increasing heat transfer to the wall.

The motivation behind the search for a device that alters or eliminates the horseshoe vortex, is that several of its characteristics - the bimodal unsteadiness, the extra drag and heat transfer it produces, and the non-uniformity and unsteadiness it brings to the wake - are undesirable in many applications. In the first part of this study, described by Devenport et al.?, the effects of placing a circular fillet of constant radius around the base of the wing were studied. Unfortunately, this device did not modify the flow in a desirable way; it did not prevent the formation of the horseshoe vortex around the wing nose, it did not eliminate the bimodal unsteadiness, and it appeared to increase the size and strength of the vortex legs and their unsteadiness in the wake. The failure of the fillet to prevent separation upstream of the wing leading edge was identified as the principal cause of its w o r performance.

' Horseshoe v o r t e x

Figure 2. Flow around a cylindrical wing mounted normal to a wall.

To prevent separation a device that better relieves the adverse pressure gradients experienced by the boundary layer here is needed. One possibility is a large fairing in the corner between the wing nose and wall upstream (figure 3), which we shall refer to as a strake. Some experimental work on such devices has been done by Sung et and Kubendran et a1.6. Kubendran et al. 's flow visualizations, performed with a laminar approach boundary layer, suggest that a suitably designed strake can eliminate leading-edge separation. Sung et al.'s velocity measurements suggest that it can also attenuate the strength of the vortex legs, decreasing the non-uniformity of the wake.

The purpose of this paper is to present measurements to clearly demonstrate the effects of adding a strake to an idealized wing-body junction with a turbulent approach boundary layer. The bimodal unsteadiness, the time-average structure of the flow around the wing and the non-uniformity and unsteadiness of the wake are examined. The effects of angle of attack and approach boundary layer thickness are also investigated.

Figure 3. The wing with strake. (a) Perspective view, (b) Side view of strake.

ADParatus and Measurement Techniques Only brief descriptions are given here, for

more details see Devenport et a1.2 or ~ewitz'.

Wins models The two wing models are shown in figures 1 and

3. Both were constructed from aluminum using a numerically-controlled milling machine. The baseline winq (fiqure 1) is cylindrical. Its cross - . - section consists of a 3:2 elliptical nose (major axis aligned with the chord) and a NACA 0020 tail. It has a maximum thickness (T) of 71.7mm, a chord of 305mm and a span of 229mm. The second wing (figure 3) is identical except for addition of the strake. The strake is shaped to provide a smooth transition from the vertical surface of the wing nose to the horizontal wall upstream. The radius of curvature of its surface is largest on the chord line where the strake extends 156mm upstream of the wing leading edge. The exact shape of the strake was specified numerically by the David Taylor Reeearch

Center, Bethesda, MD. For measurements at zero angle of attack a

boundary layer transition trip, a 6.4-mm wide strip of 120-grade sandpaper, was attached to both sides of the baseline wing, 28.2mm downstream of the leading edge. Surface oil-flow visualizations showed the trip to have no significant effect on the flow outside the wing boundary layer. No trips were used with the strake.

Test conditions, Wind tunnels The two wing models were tested under a

variety of conditions. A small boundary-layer tunnel was used to examine the effects of the strake with the wing at zero angle of attack with two different approach boundary layers, to be referred to as the "thick" and "thin" boundary layers. The Virginia Tech Stability Wind Tunnel (a large general purpose facility) was used for studies with the wing at 0°, 6' and 12' angle of attack. These studies were carried out with the "thin" approach boundary layer which was reproduced in this tunnel using a large flat plate. For measurements in this tunnel a cap was attached to the otherwise blunt top of the wing to avoid the complex flow that would otherwise be produced. The shape of the cap is that of a half body of revolution generated by rotating the cross section of the wing about its chord line.

In both tunnels blockage effects were minimal, free stream turbulence levels were low (less than 0.2%) and the approach boundary layers were closely two dimensional. The thick and thin boundary layers were produced with nominal approach free stream velocities of 27 and 32m/s respectively. Boundary- layer thicknesses, measured 2.15 wing thicknesses (T) upstream of the baseline wing at zero angle of attack, were 0.50T and 0.26T. These correspond to momentum-thickness Reynolds numbers of 6600 and 4500 respectively.

Measurements techniques Oil-flow visualizations and mean surface

pressure measurements were made on the wing models and surrounding walls. Visualizations were performed using a mixture of kerosine, titanium dioxide a d oleic acid using the procedure described k by Maltby . They were preserved using the method of Sutton9, described by Devenport and t imp son". Mean surface pressures were measured through an array of static pressure taps using a scanivalve system and two Setra Model 230 transducers interfaced to an IBM AT computer.

Hot-wire velocity measurements were made in the wakes using two rakes of 16 hot-wire sensors custom built by Dantec. These were operated using Miller1'-type constant-temperature anemometer bridges interfaced to the IBM AT. The rakes were calibrated in the wind tunnels against a pitot-static probe. To compensate for temperature variations in the stability tunnel (about 2' to 3% per hour) the rakes were calibrated frequently and corrections, determined using Bearman's'' method, were applied.

Surface pressure fluctuations were measured at the same streamwise location as the velocity measurements using Knowles Electronics BT1753 and Sennheiser MK 0 microphones. The method of Agarwal 13 and Simpson , employing two closely spaced microphones, was used to eliminate contributions to the microphone signals from vibration and acoustic noise.

Histograms of velocity fluctuations were measured upstream of the wing with strake using part of the two-component laser-Doper velocimeter (LDV) described by Simpson and Chew . The aim. here was to see if the strake eliminated the bimodal unsteadiness found around the nose of the unmodified wing.

Results and Discussion Results will be presented using the coordinate

system (X, Y, Z) and (X , Y, Z ) . The system (X, Y, Z), shown in figure 1, 2s fixe& in the wind tunnel, the spanwise and streamwise coordinates X and 2 being measured from a line coincident with the wing leading edge at zero angle of attack. The system (XI, Y, Z1) is fixed in the wing, X' and Z' being measured from the leading edge. In both systems Y is measured normal to the test wall. Distances will be

-- I

Leading-edge separation

- -Leading-edge separation

Figure 4. Surface oil-flow visualizations performed on the test wall with the thin approach boundary layer. (a) Baseline wing, 0' angle of attack, (b) Wing with strake, 0' angle of attack, (c) Wing with strake, 12' angle of attack.

normalized on the maximum thickness of the wing T or the wing chord C.

The measurements made are summarized in table 1. Because of the volume of results only a representative sample will be presented here. Uncertainties in measurements calculated using the method of Kline and ~cclintock'~, are listed in table 2.

Surface oil-flow visualizations Surface oil-flow visualizations performed with

the thin approach boundary layer on the test wall surrounding the wing models are presented in figure 4. With the baseline wing at zero angle of attack (figure 4(a)) the visualization clearly reveals the line of leading-edge separation associated with the junction vortex. This originates at a separation point in the plane of symmetry 0.45T upstream of the leading edge. Another prominent feature is a line of low wall shear stress formed between the wing

Boundary layer Thick Thin Thin Thin Reynolds nuber Re 6600 4500 4500 4500 Angle of attack ofj wing 0' 0' 6' 12'

Oil-flowvisualizations J J J J on test wall and wings

zero angle of attack. On the pressure side, however, a leading-edge separation line, originating in the corner between the strake and wall, is clearly visible. Measurements presented below suggest that this separation is associated with the roll up of a streamwise vortex.

Mean surface pressures on test wall and wings

Hot-wire surveys of J J J wake at X/C=3

LDV measurements near J J nose at Z/T=O

Microphone measurements J J J / in wake at X/C=3

Table 1. Summary of measurements made with the baseline wing and the wing with strake.

Quantity Uncertainty

Mean pressure coefficient3 f 0. 006Cp

Turbulence normal stress u2/u; f 10.0%

Y location of hot wire +0. 006T

Z location of hot wire f0.05T

RMS pressure coefficient C ' P 28.0%

Table 2. Uncertainty intervals calculated using the method of Kline and Mc~lintock'~.

nose and separation line. This feature, discussed in more detail by Devenport and ~im~son', merges with the separation line just downstream of the maximum- thickness location. Separation occurs again towards the trailing edge of the wing. Secondary separation lines emerge from the corner between the wing and wall at about the 88% chord location and then spread out, creating a fishtail-like pattern in the wake.

Flgure 4(b) shows tne corresponalng visualization performed with the strake. This picture, and visualizations performed on the strake surface, show that at zero angle of attack this device eliminates all evidence of leading-edge separation and of the formation of the horseshoe vortex around the wing nose. Many of the features present in the baseline visualization are therefore absent here. On feature that remains is the secondary separation towards the trailing edge and the fishtail pattern it produces. It is surprising that trailing-edge separation occurs almost as far downstream (X = 85%C) as in the baseline case. Given that the strake almost entirely eliminates the horseshoe vortex we would expect it also to greatly reduce the momentum transfer to the corner between the wing and wall, hastening separation here.

Oil-flow visualizations performed at zero angle of attack with the thick approach boundary layer are qualitatively and quantitatively almost indistinguishable from those shown in figures 4 (a) and (b). A weak dependence of the junction flow on the properties of the approach boundary layer is consistent with the observations of ~aker'~ on the flow past a circular cylinder mounted normal to a wall.

Apart from shifting the leading-edge separation point around towards the oncoming flow, placing the baseline wing at 6' and 12' angle of attack produced no qualitative changes in oil-flow pattern. In fact the oil-flow pattern retained a surprising degree of symmetry. Significant changes were observed, however, with the strake. Figure 4(c) shows a visualization performed around the wing with strake at 12' angle of attack, the visualization at 6' being qualitatively identical. On the suction side of the wing the visualization is much as at

Mean surface pressure measurements Contours of mean surface pressure coefficient

C constructed from measurements made on the test wsil surrounding the wing models with the thin approach boundary layer are presented in figure 5. C defined as (p - pref)/(p - p ) , where p, and p a!; the stagnation and &atirGf pressures of tKi approach free stream, has an uncertainty of t0.006.

With the baseline wing at zero angle of attack (figure 5(a)) the pressure distribution shows many of the features one would expect from inviscid considerations; a strong adverse pressure gradient upstream of the wing nose (which causes leading-edge separation), a strong favourable pressure gradient between the nose and a pressure minimum near the maximum-thickness location, and a region of pressure recovery over the downstream half of the wing. Perhaps the most interesting feature is the shallow "valley" in pressure wrapped around the wing nose about 0.3T from its surface. This is most likely due to a lowering of the pressure in the vicinity of the vortex core. Placing the baseline wing at 6' or 12' angle of attack did not qualitatively alter the pressure distribution. It did, however, increase the magnitude of the pressure gradients on the suction side of the wing and reduce them pressure side, as would be expected.

Figure 5(b) shows the pressure distribution at zero angle of attack with the strake. (Note that figures 5(b) and (c) include measurements made on the strake surface.) The strake appears to have two important effects on the pressure distribution which help explain why it prevents leading-edge separation. First, it greatly reduces the magnitude of the adverse pressure gradient experienced by the test-wall boundary layer upstream of the nose. Steep adverse pressure gradients are only felt close to the wing leading edge where the strake surface is elevated above the boundary layer. Second, it substantially reduces the magnitude of the pressure minimum found on the test wall close to the maximum thickness of the wing. Spanwise pressure gradients around the wing, which contribute to the streamwise vorticity available to the horseshoe vortex by skewing the boundary-layer structure, are thus also lowered.

At 6' and 12' angle of attack (figure 5(c)) the strake still relieves some of these spanwise gradients. Upstream of the wing nose, however, it is less effective than at zero angle of attack. As the stagnation point on the wing nose moves around towards the oncoming flow, steep pressure gradients, adverse to the local flow direction, appear on the pressure side of the strake. These presumably are responsible for the one-sided separation observed here in the oil-flow visualizations. No other evidence of this separation is visible in the pressure distribution, suggesting that it does not greatly disturb the velocity field, as a strong streamwise vortex would. At angle of attack extremely steep, but locally favourable, pressure gradients also appear on the strake surface.

Velocitv measurements in the wake reaion Measurements of mea~treamwise velocity U and

turbulence normal stress u2 were made in a spanwise plane at X/C = 3 (X/T = 12.7) to give a cross- sectional view of the wake of the wing-body junction. Measurements were made at all conditions except with the wing at 12' angle of attack.

As t representative sample of the results, contours of u deduced from measurements made in the Stability Tunnel with the thin approach boundary lay- are presented in figures 6(a) through 6.(d). u2 is normalized on U , the maln stream veloclty at X/C = 3, and has an incertainty of *lo%. Note that with the wings at zero angle of attack measurements were made only on one side of the wake, the flow being assumed symmetrical. The complete cross-sections shown in figures 6(a) and 6(c) were thus obtained by reflecting these data about Z = 0.

Figure 5 . Contours of mean surface-pressure coefficient C on the test wall with the thin approach boundary layer. (a) Baseline wing, 0' angle of attack, (%) Wing with strake, 0' angle of attack, (c) Wing with strake, 12' angle of attack.

In all the figures the approximately horizontal and vertical contours of the test-wall boundary layer and wing wake are clearly visible. At O0 angle of attack (figures 6(a) and 6(c)) the wake has a height of about 3.6T and closes in a manner that suggests that the cap (used for the Stability- Tunnel measurements) was sucessful in preventing flow separation over the top of the wing. At 6' angle of attack (figures 6(b) and (d)) the wake has almost the same height but is dominated by the wing-tip vortex shed towards the suction side of the wing.

With the baseline wing (figures 6(a) and 6(c)) the velocity contours on either side of the wake show clear evidence of the horseshoe-vortex legs shed by the junction. The buckling of the contours in the vicinity of the vortex legs is indicative of

wall. Adding the strake at zero angle of attack

(figure 6(b)! greatly reduces, but does not eliminate entirely, these features. This suggests that although the strake prevents leading edge separation and thus the formation of the horseshoe vortex around the wing nose, vortex legs are formed downstream. However, in the time averaged views of figure 6 these legs are clearly much smaller than those produced with the baseline wing. Mean-velocity

Figure 6. Contours of u2/ue in a spanwise plane at x/C = 3 with the thin approach boundary layer. (a) Baseline wing, 0' angle of attack, (b) Wing with strake, O0 angle of attack, (c) Baseline wing, 6' angle of attack, (d) Wing with strake, 6O angle of attack. Arrows on Z axis mark locations of microphone measurements, see table 3.

contours, not presented here, also suggest that the vortex legs are closer together with the strake. This implies that they are weaker, since weaker legs would move apart less rapidly under the action of their images in the test wall.

At 6' angle of attack with the thin approach boundary layer (figure 6(d)) the strake again reduces the distortion of the velocity contours associated with the vortex legs. The reduction is, however, greater on the suction side (towards negative Z) than on the pressure side of the wake. The apparently stronger and larger pressure-side vortex leg presumably originates in the corner between the strake and wall where the oil-flow visualizations showed the formation of a region of separated flow.

Changing the thickness of the approach boundary layer did not qualitatively alter the structure of the wake and vortex legs or the effects of the strake upon them. Quantitatively changes in the apparent size of the vortex legs were small in comparison to the factor of two change in approach boundary-layer thickness. This insensitivity to the approach boundary layer is consistent with the surface oil-flow visualizations.

Microphone measurements in the wake reaion Measurements of surface pressure flu~tuati0nS

were made at X/C = 3 in an attempt to quantify the effect of the strake on the wake unsteadiness. Microphones were located under those regions where the velocity contours appeared most distorted by streamwise vorticity. Exact measurement positions are indicated by the arrows on figure 6. Results are summarized i~ -table 3 in terms of the coefficient C ' = Jp' '1 (.%p~,ef2), uncert3inty fa%.

Except 'in the thlck approach boundary layer (where C ' is almost the same with and without the strake) ?he strake appears to significantly reduce pressure fluctuation levels. Spectra show this reduction to be fairly uniform across the spectrum. At angle of attack the strake is less effective at reducing pressure fluctuations on the pressure side than on the suction side of the wake. This suggests that the streamwise vortex, apparently generated by separation on the pressure side of the wing with strake, is a significant source of unsteadiness in the wake.

Angle of attack/ Boundary Baseline Wing with locat ion layer wing strake

0' Thick 0.0063 0.0064 0' Thin 0.0095 0.0088 6'/suction side Thin 0.0108 0.0098 6'/pressure side Thin 0.0083 0.0078 12'/suction side Thin 0.0109 0.0092 12'/pressure side Thin 0.0113 0.0102

Table 3. Total wall pressure fluctuation levels at X/C = 3 expressed in terms of C '. See figure 6 for spanwise locations of measuremehs.

LDV measurements near the nose The LDV was used to search for double-peaked

histograms of velocity fluctuations in the vicinity of the strake at zero angle of attack. Histograms of this type are indicative of the large-scale unsteadiness present around the nose of the baseline wing. Not surprisingly (since this device eliminates leading-edge separation at zero angle of attack) none were found. It remains an open question as to whether the pressure-side separation produced by the wing with strake at 6' and 12' angle of attack is associated with unsteadiness of this type.

Conclusions The effects of a strake on the flow of a

turbulent boundary past an idealized wing body junction have been studied experimentally. The strake consisted of a large fairing in the corner between the wing nose and the body surface upstream. Measurements made with two approach boundary layers of different thickness and with the wing at o', 6' and 12' angle of attack suggest the following conclusions.

At zero angle of attack in both boundary layers the strake prevents leading edge separation and thus the formation of a horseshoe vortex around the wing nose. It also eliminates the bimodal unsteadiness that characterizes the flow in the nose region of the unmodified wing. The strake greatly reduces the non-uniformity of the wake downstream of the wing. Velocity contours measured here do show some evidence of horseshoe vortex legs presumably formed downstream of the nose region. However, these are very much weaker than with the unmodified wing.

At 6' and 12' angle of attack the strake also greatly improves the characteristics of the flow. However, at angle of attack separation occurs on the pressure side of the strake because of adverse pressure gradients here. Separation appears to cause the roll up of a streamline vortex. While this vortex is much weaker than either leg of the horseshoe vortex generated by the unmodified wing, it does lower somewhat the effectiveness of the strake in reducing the non-uniformity and unsteadiness of the wake.

Overall the strake appears to modify the flow past the wing body junction in a way that would be desirable in most applications. An obvious topic for future research is the structure of the one- sided leading-edge separation region produced with the strake at angle of attack. Whether or not this separation is associated with bimodal unsteadiness like that observed with the baseline wing remains an open question.

Acknowledaments This work was supported by ONR under contract

N00014-88-C-0291.

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