Effect of Support Stem on a Dynamically Pitching Delta Wing Roy Y. Myose*
Department of Aerospace Engineering, Wichita State University, Wichita, KS 67260-0044
and
Ismael Heron†
Bombardier Learjet, Wichita, KS 67209
A series of experiments was conducted at Wichita State University to study the effect of support stem on the vortex burst position of a 70-degree sweepback delta wing. It is well known that a fighter aircraft’s performance at high angles of attack is greatly influenced by the development of leading edge vortices on a delta-shaped wing. The present investigation was motivated by a desire to understand how the design of specific model support structures can affect the delta wing vortex burst behavior. Results indicate that there is a slight influence on the burst location even if the support stem is located aft of the wing on the pressure side.
Introduction Recent interest in highly maneuverable military aircraft, capable of operating over a large range of angles of attack, has refocused attention on delta-shaped wings. One of the hallmark features of delta-shaped wings and strakes is the presence of a pair of vortices called leading-edge vortices. At non-zero angles of attack, there is a pressure difference between the upper suction surface and the lower pressure surface which causes a flow around the leading-edges. The flow detaches along the leading-edge into a shear layer that curls up into a spiral. The center of the spiral is tight enough that it forms, in essence, a pair of strong counter-rotating vortices. These leading-edge vortices induce velocities on the flow field and additional suction over the delta wing which can account for up to 30% of the total lift at moderate angles of attack.1 For example, a 70-degree swept delta wing continues to increase its lift until about 40-degrees angle of attack.2 In comparison, symmetric two dimensional airfoils typically stall out at around 10- to 15-degrees angle of attack. Unfortunately, there are limits to the benefits produced by the leading-edge vortices. As the angle of attack is increased, there is a sudden breakdown in vortex structure followed by degeneration into a non-coherent turbulent-like flow. This phenomena, also known as vortex bursting, can be defined as a sudden expansion in radial size and an abrupt decrease in the axial velocity of the vortex.3 Once this occurs, lift is no longer enhanced aft of the burst point. Thus, the development and subsequent breakdown of leading-edge vortices is crucial to the performance of delta wing aircraft. If the delta wing is pitched to a given angle of attack and then maintained at that angle until the transient flow features die down, it is said to be tested under "static" conditions. As this process is repeated at increasing values of α, the vortex burst will be located at "fixed" positions that are closer to the apex of the delta wing.
* Professor, Associate Fellow AIAA. † Senior Engineer Flight Sciences, Member AIAA.
9th AIAA Aviation Technology, Integration, and Operations Conference (ATIO) <br>and <br>Air21 - 23 September 2009, Hilton Head, South Carolina
In comparison, the dynamic case is the situation where the delta wing is continuously pitched, never allowing the flow features to become steady. When the delta wing is dynamically pitched upwards at a given rate, the location of the vortex burst is farther towards the trailing edge compared to the same angle of attack under static conditions.4-7 This produces a phase lag in the burst location, allowing transient values of lift to exceed those obtained during static testing. Similarly, pitching down the delta wing results in a vortex burst location forward of the static case. This produces a phase lead and a reduction in lift, compared to similar static angles of attack. Such phase lag or lead in the dynamic case means that a time delay or hysteresis effect occurs, where there is a difference in the measured lift coefficient value if the angle of attack is increasing or decreasing.8 The magnitude of the phase lag or lead increases as the pitch rate increases.7 The faster the pitch up rate, the higher the angle of attack before the vortex burst appears over the surface of the delta wing. Modern combat aircraft use either slender delta wings or highly swept leading-edge extensions (i.e., strakes) that harness vorticity. These aircraft take advantage of the phase lag hysteresis effect in an attempt to increase the performance envelope. In many cases, the increase in performance has led to aircraft with "hyper-agility," or the ability to maneuver at very fast rates. Take, for example, the case of the Su-27 aircraft undergoing a Cobra maneuver. In this case, the aircraft enters the pitch up phase at 190 knots indicated airspeed (kias). During a 2- to 5-second time frame, the aircraft reaches 90-degrees angle of attack or more while the airspeed drops substantially to about 70 kias. The aircraft subsequently points nose down in order to accelerate and exit at a much lower angle of attack.9,10 In order to properly simulate "hyper-agile" maneuvers such as the Cobra, researchers have been focusing lately on experiments and simulations that are closer to the Reynolds number spectrum of full-scale aircraft.11 However, this involves the use of large, expensive, and heavily mechanized models and mounts that are necessary to replicate the high-rate maneuvers in pitch (and/or yaw and roll).12 Even with this, it is seldom possible to replicate the exact full-scale Reynolds numbers. Another problem with this approach is the control of the freestream velocity fluctuation that naturally occurs during many of these "hyper-agile" maneuvers. Many water and wind tunnels are not capable of decelerating their flow velocity rapidly enough, especially those large enough to accommodate the mounts necessary to accomplish the full-scale Reynolds numbers and the motions being simulated. At the opposite edge of the Reynolds number spectrum, renewed interest in MAV’s and UAV’s have promoted research into time-dependent methods of achieving high wing-loadings, high maneuverability, and small physical size. With these issues in mind, a low-cost towable delta wing mount was developed for use in the Wichita State University water tunnel, in order to study the effect of simultaneous pitch and freestream velocity changes. The delta wing vortex burst position was measured as a function of the pitch rate and acceleration or deceleration of the freestream velocity. Past work on this topic has described the development of the towing system13, some preliminary work,14 fast pitch-up under deceleration,15 and deceleration and acceleration at moderate and fast pitch-up and pitch-down rates.16 In a separate set of studies, the impingement of von Kàrmàn vortex street on the delta wing leading edge vortex along with their subsequent interaction was investigated using the same towing mount system. Past work on this topic have included the impingement of von Kàrmàn vortex street with the delta wing at fixed angle of attack17,18, preliminary results for the pitch-up case19, the full results for the large cylinder shedding case20, and results for the small cylinder shedding case21. Prior to these studies, results using the "new" towing mount system were obtained for some baseline cases (static angle of attack and three different dynamic pitch rates at constant freestream speed without von Kàrmàn vortices). This towing mount system was quite different from the support system used in the past at the same facility for dynamic pitch testing with fixed constant freestream speed.5,7,22 This brought up the issue of how comparable are the results using these two different model support systems. Taylor et al.23 and Taylor and Gursul24 found that the use of a support rod on the delta wing’s suction side affected the position of the vortex burst in two ways: the vortex burst was moved farther towards the apex and there was a hysteresis effect to this position. Previous dynamic testing at Wichita State University’s water tunnel utilized a vertical turntable with a support stem on the delta wing’s pressure side. Thus, the objective of the present paper is to discuss the effect of a pressure side support stem on the vortex bursting position for static angle of attack and dynamic pitching. (It should be noted that the actual testing and analysis for this support stem work was conducted prior to the series of work on dynamic freestream13-16 and impingement of von Kàrmàn vortex street17-21.)
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Experimental Method An existing water tunnel facility was used to create a towing facility where dynamic pitching and dynamic freesteam could be obtained. The upper illustration of Figure 1 shows a schematic diagram of the water tunnel located in the Wichita State University’s National Institute for Aviation Research.25 This facility is a closed-loop tunnel containing 3500 gallons of water, and consists of a 2- by 3- by 6-feet test section. The facility has excellent optical access providing two side views, a bottom view, and an end view. The water tunnel is capable of producing flow velocities up to 1 ft/s using an impeller pump driven by a 5-hp variable-speed motor. For the purposes of the present investigation, however, the tunnel was used as a simple water tank with the pump turned off. The photograph in the lower half of Figure 1 shows the towing system with an inverted delta wing mounted on a carriage (which is not visible). The carriage itself rides on top of a track which is built into an aluminum frame. A lightweight "false ceiling" is built into the frame to suppress bow waves which may be created by towing the model support strut if there were a free surface. Towing speeds of 0.4 ft/s is obtained by pulling the carriage with a spring-tensioned nylon wire and a DC motor. The delta wing model is mounted upside down, such that "pitch-up" involves rotating the delta wing apex down towards the floor of the water tunnel. This arrangement reduces the likelihood of carriage derailment since the delta wing provides additional down force for the carriage wheels. The towing speed is kept constant at 0.4 ft/s + 5% (see Figure 2) in the present investigation so that the Reynolds number based on the delta wing’s 12-inch root chord length corresponds to Rec=33,000. Visible in the photograph of Figure 1 is the tandem strut mounting mechanism for the delta wing. The larger diameter strut is attached to a hinge on the delta wing’s pressure side. The smaller diameter strut located on the downstream side is attached to a cam mechanism on the carriage which is driven by a small DC motor. This allows the delta wing to be pivoted about the half-chord location at non-dimensional pitch rates up to κ=0.2. The angle of attack range which can be obtained using the towing mount system is 15°<α<55°. Additional details about the towing system are presented in Reference 13. The delta wing model (70-degree sweepback, 12-inch root chord, 1/8-inch thick, sharp leading edge beveled at a 30 degree angle) is the same model as used by Myose et al.7,22 For vortex burst location measurements, dye flow visualization was used. Dye ports (1/32 inch diameter stainless steel tubes) and delivery lines were symmetrically located on both the port and starboard sides of the delta wing’s pressure side. However, only dye from the side closest to the camera was used to visualize the vortex behavior, the assumption being that it is representative of the vortex not being observed. A schematic diagram of the experimental set-up is illustrated in Figure 3. The flow visualization images were video-taped and information such as the (towing carriage) freestream velocity U∞, delta wing angle of attack α, and time code were also recorded. The flow visualization images and relevant information were subsequently analyzed using a computer-assisted image analysis software tool.13
According to Ericsson26, sizeable errors in the observed vortex burst locations can occur for a wingspan to test section width (b/w) greater than 0.7. It was assumed that delta wings whose spans are smaller than this criterion exhibits a minimal error. In the present investigation, the wingspan to test section width ratio is 0.364. Blockage can be a particularly important source of interference, especially at high angles of attack. Blockage has the effect of accelerating the flow around the model as the cross sectional area is reduced. The ratio of the wing’s projected frontal area at its highest angle of attack to the tunnel cross-sectional area, ADelta/ATunnel, is an estimator of the severity of the blockage errors. An area ratio ADelta/ATunnel < 7% is considered acceptable according to the guidelines published in Reference 27. In the present investigation, the maximum cross-sectional area ratio (for a greater than expected maximum angle of attack of 60°) was 4%. There are a few limitations with this experimental set-up that need to be acknowledged: 1- The location of the vortex burst is accomplished by identifying where the core flares out (bubble burst), or the location of the first sharp kink (spiral burst). As such, the identified location may or may not coincide with the actual core stagnation point. The assumption that the two are close, if not coincident, has been done in the past. 2- Some of the features observed cannot be explored further using this system, thus a certain amount of interpretation must be exercised. Nevertheless, the uncertainty should be within ± 0.05c. Previous dynamic testing at Wichita State University’s water tunnel utilized a vertical turntable to pitch the model, which was supported by a stem from the trailing edge on the pressure side.5,7,22 The stem was then
4
connected by a cross-brace to the turntable (Figure 4, left-hand illustration). To simulate this stem and cross-brass, the delta wing was fitted with a wooden stem and cross-brace, 1/2 inch in diameter, resembling the mount system used in previous dynamic testing at this water tunnel. The simulated stem was only connected to the delta wing’s trailing edge. A slight clearance was left between the tunnel walls and the end of the brace. The test matrix for this investigation is presented in Table 1. Tests were performed under static (fixed angle of attack) and dynamic conditions of κ=0.1, 0.15, and 0.2 pitch rates. For each (static or dynamic) pitch condition, tests were performed with no stem and the stem located closest to the camera (on the port side where the leading edge vortex is visualized with dye) as well as with the stem located farthest from the camera (on the starboard side). Three experimental runs for each test condition were performed and the results presented below are the average of the three runs.
Results Figure 5 shows the vortex burst (VB) location results using the present towing mount system compared to results from an experimental set-up using a vertical turntable from Myose et al.7 It should be noted again that in the towing mount system, the delta wing is towed in a quiescent water tunnel and pitched up from the end of a vertical stem. In comparison, the vertical turntable set-up involves a stationary delta wing with water flowing through the tunnel and the delta wing pitched up using a horizontal stem attached to a vertically oriented turntable. Consequently, the two set-ups (illustrated in Figure 4) are quite different. Nevertheless, the VB locations are reasonably similar. At first glance, the dynamic case appears identical while the static results exhibit some differences. The curve’s general slope with the towing mount (solid line with open square symbols) appears initially shallower than the results of Myose et al7 (dashed line with grey-filled square symbols). As shown in Figure 6, however, comparison with other results28-30 reported for 70-degree sweepback delta wings show that there is wide scatter in the vortex burst locations, and the results obtained here are within the scatter that ranges about 20% of chord. Figures 7 through 10 show the results of vortex burst locations under static, κ=0.1, 0.15, and 0.2 dynamic pitch-up conditions, respectively. Each figure shows the results for the no stem condition (solid line with open symbols), stem located away from the vortex and burst being visualized (grey-filled symbols without any lines), and stem located on the side of the vortex and burst being visualized (dark-filled symbols without any lines). Also shown (in dashed line with grey-filled symbols) are the previously reported results from Myose et al.7 It should be noted that the angle of attack range with the towing mount does not reach as far up as Myose et al.’s due to physical constraints in the towing mount. Myose et al.’s results (i.e., with the stem located away from the vortex and burst being visualized) should be most comparable to the "far stem" results of the present investigation. The results show that there are some differences in the vortex burst locations, but they are well within the range of scatter found in investigations from various studies28-30 discussed earlier. Figures 11 through 14 present the same results as those presented earlier (in Figures 7 through 10), but with the range of observed vortex burst locations (from the three different runs) indicated for all three stem configurations in the bar chart. Removing the stem from the delta wing’s rear produced a fairly clear aftward shift in the vortex burst location. In general, the vortex burst is located slightly closer to the apex with the stem compared to the case without the stem. Furthermore, orienting the cross-brace on the same side as the vortex being visualized moved the burst even closer to the apex as shown in Figure 15. It should be noted that these trends are just at the outer edge of the measurement error band. Thus, these results should be viewed with some caution.
Summary The effect of support stem on the vortex burst location of a 70-degree sweepback delta wing was investigated. It is well known that a fighter aircraft’s performance at high angles of attack is greatly influenced by the development and subsequent bursting of the leading-edge vortices on a delta-shaped wing. The present investigation was motivated by a desire to understand how the design of specific model support structures can affect the delta wing vortex burst behavior. Dye flow visualization measurements of the vortex burst locations for the delta wing with and without a false support stem were made in the Wichita State University water tunnel. Results indicate that there is a slight influence on the burst location even if the support stem is located aft of the wing on the pressure side. Locating the cross-brace on the same side as the visualized vortex moved the burst slightly forward toward the apex compared to the no stem case.
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References
1. Polhamus, E. "A Concept of the Vortex Lift of Sharp-Edge Delta Wings Based on a Leading-Edge Suction Analogy," NASA TN D-3767, Dec. 1966.
2. Earnshaw, P.B., and Lawford, J.A., "Low-Speed Wind Tunnel Experiments on a Series of Sharp-Edged Delta Wings," Aeronautical Research Council, R M No. 3424, 1964.
3. Erickson, G.E., "Water-Tunnel Studies of Leading-Edge Vortices," AIAA Journal of Aircraft, Vol. 19, No. 6, 1982, pp. 442-448.
4. LeMay, W.P., Batill, S.M., and Nelson, R.C., "Vortex Dynamics on a Pitching Delta Wing," AIAA Journal of Aircraft, Vol. 27, No. 2, 1990, pp. 131-138.
5. Miller, L.S., and Gile, B.E., "Effects of Blowing on Delta Wing Vortices During Dynamic Pitching," AIAA Journal of Aircraft, Vol. 30, No. 3, 1993, pp. 334-339.
6. Rediniotis, O.K., Klute, S.M., Hoang, N.T., and Telionis, D.P., "Dynamic Pitch-Up of a Delta Wing," AIAA Journal, Vol. 32, No. 4, 1994, pp. 716-725.
7. Myose, R.Y., Hayashibara, S., Yeong, P.C., and Miller, L.S., "Effect of Canards on Delta Wing Vortex Breakdown during Dynamic Pitching," AIAA Journal of Aircraft, Vol. 32, No. 2, 1997, pp. 168-173.
8. Al-Garni, A.Z., Ahmed, S.A., Sahin, A.Z., and Al-Garni, A.M., "An Experimental Study of a 65-Degree Delta Wing with Different Pitching Rates," Canadian Aeronautics and Space Journal, Vol. 47, No. 2, 2001, pp. 85-93.
9. Skow, A.M., "An Analysis of the Su-27 Flight Demonstration at the 1989 Paris Airshow," SAE Paper 90-1001, 1990.
10. Kolano, E., "Flying the Flanker," Flight Journal, Vol. 4, No. 4, Aug. 1999. 11. Ericsson, L.E., "Effect of Fuselage Geometry on Delta-Wing Vortex Breakdown," AIAA Journal of
Aircraft, Vol. 35, No. 6, Nov. 1998. 12. Schaeffler, N.W., and Telionis, D.P., "Apex Flap Control of Dynamic Delta Wing Maneuvers," AIAA
Paper 96-0662, 1996. 13. Heron, I., and Myose, R.Y., "Development of a Dynamic Pitch and Unsteady Freestream Delta Wing
Mount For A Water Tunnel," AIAA Paper 2003-3527, June 2003. 14. Heron, I., and Myose, R.Y., "Vortex Burst Behavior Under Dynamic Freestream," AIAA Paper 2005-
0063, January 2005. 15. Heron, I., and Myose, R.Y., "Delta Wing Vortex Burst Behavior Under Dynamic Freestream, Part 1 - Fast
Pitch-up During Deceleration," AIAA paper 2007-6725, August 2007. 16. Myose, R.Y., and Heron, I., "Delta Wing Vortex Burst Behavior Under Dynamic Freestream, Part 2 -
Deceleration and Acceleration at Moderate and Fast Rates," AIAA paper 2007-6726, August 2007. 17. Heron, I., and Myose, R.Y., "Impingement of a von Kàrmàn Vortex Street on a Delta Wing," AIAA paper
2004-4731, August 2004. 18. Heron, I. and Myose, R.Y., "Impingement of a von Kàrmàn Vortex Street on a Delta Wing," AIAA
Journal of Aircraft, Vol. 42, No. 4, 2005, pp. 1084-1087. 19. Heron, I., and Myose, R.Y., "Impingement of a von Kàrmàn Vortex Street on a Dynamically Pitching
Delta Wing," AIAA paper 2006-0057, January 2006. 20. Heron, I., and Myose, R.Y., "Impingement of a von Kàrmàn Vortex Street on a Dynamically Pitching
Delta Wing, Part 1 - Large Cylinder Results," AIAA paper 2009-0095, January 2009. 21. Myose, R.Y., and Heron, I., "Impingement of a von Kàrmàn Vortex Street on a Dynamically Pitching
Delta Wing, Part 2 – Small Cylinder Results," AIAA paper 2009-0096, January 2009. 22. Myose, R.Y., Lee, B.K., Hayashibara, S., and Miller, L.S., "Diamond, Cropped, Delta, and Double Delta
Wing Vortex Breakdown During Dynamic Pitching," AIAA Journal, Vol. 35, No. 3, 1997, pp. 567-569. 23. Taylor, G., Gursul, I., and Greenwell, D., "Investigationn of Support Interference in High-Angle-of-
24. Taylor, G., and Gursul, I., "Support Interference for a Maneuvering Delta Wing," AIAA Journal of Aircraft, Vol. 42, No. 6, November-December 2005.
25. Johnson, B.L., "Facility Description of the Walter H. Beech Memorial 7 x 10 foot Low-Speed Wind Tunnel," AR93-1, National Institute for Aviation Research, Wichita State University, June 1993.
26. Ericsson. L.E., and Beyers, M.E., "Aspects of Ground Facility Interference on Leading-Edge Vortex Breakdown," AIAA Journal of Aircraft, Vol. 38, No. 2, March 2001, pp. 310-314.
27. Rae, W.H., and Pope, A., Low-Speed Wind Tunnel Testing, 2nd Ed., J. Wiley, 1984. 28. Wentz, W.H., and Kohlman, D. L., "Vortex Breakdown on Slender Sharp-Edged Wings," AIAA Journal
of Aircraft, Vol. 8, No. 3, Mar. 1971, pp. 156-161 29. Thompson, D.H. "A water tunnel study of vortex breakdown over wings with highly swept leading
edges," Australian Research Laboratories Note ARL/A 356, May 1975. 30. Kegelman, J.T., and Roos, F.W., "Effects of leading edge shape and vortex burst on the flowfield of a 70-
degree-sweep delta wing," AIAA paper 89-0086, January 1984.
7
Figure 1 - Schematic diagram of the Wichita State University 2- by 3-feet water tunnel (top)25 and photograph of test section showing the towing system with an inverted delta wing (bottom)13.
Carriage
Drive
Water Level
8
0123456
0 2 4 6 8 10 12
Elapsed time (sec)
Vel
ocity
(in/
s)
Figure 2 - Velocity Ramp for Carriage.13
Figure 3 - Overview of the Data Gathering Set-Up.17
Left FOV
Right FOV
Carriage Travel (inches)
0 31 6145
Center FOV
Carriage Travel Analyzed
Average
White Backboard
Delta Wing
Fluorescent Lamps
Tunnel Test Section
3 ft
Field of View
(FOV)
1½ ft
Camera
α and U∞ voltages
A/D converter
(VisualBASIC® Program)
Video Mixer
Time Code Generator
S-VHS VCR
9
Figure 4 - Dynamic mount used by Myose et al.7,22 (left) and current towing mount with false stem (right).
Table 1 - Test Matrix for Support Stem Effect Investigation
Type of Test Angle of Attack α (deg)
Non-Dimensional Pitch Rate κ
Stem Points towards No. Runs per Test Case
Static 25, 30, 35, 40, 45, 50 0 3 (runs per α) x 6 (α) = 18
Dynamic Pitch 15 to 55 0.1, 0.15, 0.2 None 3 (runs per κ) x 3 (κ) = 9
Static 25, 30, 35, 40, 45, 50 0 3 (runs per α) x 6 (α) = 18
Dynamic Pitch 15 to 55 0.1, 0.15, 0.2 Far Wall 3 (runs per κ) x 3 (κ) = 9
Static 25, 30, 35, 40, 45, 50 0 3 (runs per α) x 6 (α) = 18
Dynamic Pitch 15 to 55 0.1, 0.15, 0.2 Near Wall 3 (runs per κ) x 3 (κ) = 9
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