Ismael Heron

Gurney flap experiments on airfoils, wings, and reflection plane model

The effect of Gurney e aps on two-dimensional airfoils, three-dimensional wings, and a ree ection plane model were investigated. There have been a number of studies on Gurney e aps in recent years, but these studies have been limited to two-dimensional airfoil sections. A comprehensive investigation on the effect of Gurney e aps for a wide range of cone gurations and test conditions was conducted at Wichita State University. A symmetric NACA 0011 and a cambered GA (W)-2 airfoil were used during the single-element airfoil part of this investigation. The GA (W)-2 airfoil was also used during the two-element airfoil study with a 25% chord slotted e ap dee ected at 10, 20, and 30 deg. Straight and tapered ree ection plane wings with natural laminar e ow (NLF) airfoil sections were tested for the three-dimensional wing part of this investigation. A fuselage and engine were attached to the tapered NLF wing for the ree ection plane model investigation. In all cases the Gurney e ap improved the maximum lift coefe cient compared to the baseline clean cone guration. However, there was a drag penalty associated with this lift increase.

Vortex Burst Behavior Under Dynamic Freestream

A series of experiments on a 70-degree delta wing was conducted at Wichita State University to study the effect on the vortex burst position when simultaneous dynamic pitch and unsteady freestream velocity were used. The aim was to better understand the relationship between the freestream velocity and the time constants involved in the movement of the vortex burst point. Experiments indicated that a change in the freestream velocity resulted in a momentary pause in the forward progression of the vortex burst of the leading-edge vortex.

Delta Wing Vortex-Burst Behavior Under a Dynamic Freestream

A series of experiments was performed at Wichita State Universitys water tunnel on a 70-degree-sweep delta wing using a towing mount. A video camera captured dye-flow visualization images of the vortex burst that were subsequently analyzed using a computer-assisted image analysis software. The aim was to better understand the relationship between the freestream velocity and the time constants involved in the movement of the vortex-burst point. Experiments indicated that a change in the freestream velocity changed the forward progression of the vortex burst. Under pitch-up conditions, deceleration resulted in a momentary retardation in the forward progression ofthe burst, whereas acceleration resulted in a faster progression toward the apex.

Delta Wing Vortex Burst Behavior Under Dynamic Freestream, Part 1 - Fast Pitch-up During Deceleration

A series of experiments was performed at the Wichita State University’s Water Tunnel on a 12-inch root chord, 70-degree sweep delta wing, using a self-constructed towing mount. A video camera captured dye flow visualization images of the vortex burst which subsequently were analyzed using a computer-assisted image analysis software. The experiments were the main thrust of an investigation into the behavior of the leading edge vortex burst under a series of variable freestream velocity and fast pitch-up conditions. Results showed that there was a marked reduction in the forward propagation rate of the vortex burst when the delta wing was decelerated while being pitched up.

Flow Visualization Study on the Effect of a Gurney Flap in a Low Reynolds Number Compressor Cascade

The effect of a Gurney flap in a compressor cascade model at low Reynolds number was investigated using tuft flow visualization in a water table facility. Although small in scale, water tables have the advantage of low cost and the ease with which test conditions can be varied. In this experiment, tuft flow visualization was used to determine the outgoing flow angle for a NACA 65-(12)10 compressor cascade model with a solidity of 1.5 at a blade chord Reynolds number of 16,000. The baseline (no flap) results were found to be in good agreement compared to results in the literature for tests conducted at Reynolds number in the 250,000 + range. A second set of measurements were then taken for a Gurney flap with a height of 2% of the chord length attached to the trailing edge of the cascade blades. The results suggest that the Gurney flap energizes the flow and delays the stall at large incoming flow angles. Nomenclature c = chord length Re C = Reynolds number based on chord length, Uc/ν U = freestream velocity y = offset distance in the stagger direction βin = incoming flow angle, between the in-flow direction and a line perpendicular to the stagger line βout = outgoing flow angle, between the out-flow direction and a line perpendicular to the stagger line λ = stagger angle, between the chord line and a line perpendicular to the stagger line ν = kinematic viscosity σ = solidity of cascade, c/y

Impingement of a von Karman Vortex Street on a Delta Wing

Introduction H IGH-PERFORMANCE military aircraft oftentimes employ delta-shaped wings. It is well documented that delta wings at a fixed angle of attack generate lift by separating a shear layer of fluid (air or water) at the leading edge, and this shear layer forms two strong counter-rotating vortices on either side of the wing.1−5 These leading-edge (LE) vortices are critical to the generation of lift, as they produce a large suction peak on the surface. Under certain conditions, the LE vortices are prone to undergo a change in their coherent structure. The vortex expands around the core, slows down axially, and forms either a bubble or a spiral, with the spiral form being more predominant at Reynolds numbers of interest to aircraft designers. This change, called vortex burst or breakdown, is dependent on the aspect ratio of the wing, angle of attack, pressure gradient, yaw angle, and swirl angle of the vortex, among others.6−8 Delta wings have evolved over the years and are now used primarily in the form of leading-edge extensions on many fighter aircraft. As these aircraft become more and more maneuverable, the understanding of the physics of time-dependent unsteady flows is becoming more important. The vortex burst has such a detrimental effect on the lift generation of delta wings that it is important to correctly model the flow, predict it with accuracy, and control the movement of the burst location. If accurate computational models of these maneuvers are to be developed, it is necessary to understand the mechanisms involved in vortex bursting and mixing on the lee side of the delta wing. This is true of so-called “hyperagile” maneuvers (like the Cobra maneuver), where research has been intense.9−13 The behavior of the burst location is sensitive to yaw. Aircraft performing the Cobra maneuver (or other high-angle-of-attack maneuvers) can be unstable in yaw.14−17 Sometimes, the wake of an aircraft’s forebody can have vortices at very high angles of attack.18 In other cases, close proximity of the delta wing to the wake of another aircraft would subject the delta to an external vortical flow. This inspired the authors to consider the effect of an impingement from another set of vortices upon the delta-wing’s leading-edge vortices. Because a von Karman vortex street has a well-known behavior,

Impingement of a von Kàrmàn Vortex Street on a Dynamically Pitching Delta Wing

A series of experiments on a 70-degree delta wing was conducted at Wichita State University to study the effect on the vortex burst position of the leading-edge vortices when subjected to a von Karman wake from a cylinder while, simultaneously, the wing is dynamically pitched. The aim is to better understand how the impingement of the von Karman Vortex Street affects the vortex bursting location. Results indicate that there is a temporal correlation between the passage of the von Karman vortex and the leading edge vortex burst location, indicating modulation of the burst location. Nomenclature

On the Impingement of von Karman Vortex Street on a Delta Wing

It is well known that aircraft undergoing high angle of attack excursions (e.g., the Cobra maneuver) may experience asymmetric forebody vortex shedding under certain conditions. The detachment of these vortices from the forebody is similar to that observed from von Karman vortex shedding from cylinders. A series of experiments, where a von Karman vortex street wake was made to impinge upon a 70-degree delta wing, was conducted at Wichita State University. The aim is to better understand the mixing mechanism that occurs at these high angle of attack flight regimes in a less Reynolds number sensitive environment. A von Karman wake having a frequency similar to the forebody shedding process is used. As the von Karman vortex filaments are entrained by the shear layer, they appear to wrap themselves around the core. There is a temporal correlation to the burst location of the delta wing’s leading-edge vortices, indicating modulation of the burst location. Additionally, the core becomes distorted. It has been found that the vortex burst location is moved forward towards the apex when subjected to the von Karman wake. Nomenclature c Wing root chord d Cylinder diameter FOV Field of View LE Leading Edge Rec Chord Reynolds number, U∞c/ν ReD Diameter Reynolds number, U∞d/ν s Coordinate along root chord U∞ Freestream Velocity x Coordinate along horizontal direction α Angle of attack (deg) Introduction Delta wings have evolved over the years and are now used primarily in the form of leading edge extensions on many fighter aircraft. As these aircraft become more and more maneuverable, the understanding of the physics of time-dependent unsteady flows is becoming more important. In particular, if accurate computational models of these maneuvers are to be developed, it is necessary to understand the mechanisms involved in features such as vortex bursting and mixing on the lee side of the delta wing. This is true of maneuvers such as the Cobra and other “hyper-agile” maneuvers. To understand the situation, it is important to briefly discuss some of the experimental findings in the areas of vortex bursting behavior, the Cobra maneuver physics, and forebody shedding. On the Impingement of a von Karman Vortex Street on a Delta Wing Ismael Heron and Roy Y. Myose Department of Aerospace Engineering Wichita State University, Wichita KS 67260-0044 It is well documented that delta wings at a fixed angle of attack generate lift by separating a shear layer of air (or fluid, such as water) at the leading edge, and this shear layer forms two strong counter-rotating vortices on either side of the wing. These leadingedge (LE) vortices undergo small fluctuations in space, but remain relatively fixed over the suction side of the delta wing, and are critical to the generation of lift, as they produce a large suction peak on the surface. Two much smaller vortices, the secondary vortices, are also formed, as seen in Figure 1. In other words, LE vortices are the result of a balance between vorticity being generated at the leading edge, and the ability of the flow field to convect said vorticity along the vortex core. The LE vortices are not stable, and at some point their coherent structure will undergo a dramatic change, expanding around the core, slowing down axially, and either forming a bubble or a spiral, with the spiral form being more predominant at Reynolds numbers of interest to delta wing designers. This change, called vortex burst or breakdown, is dependent on the aspect ratio of the wing, angle of attack, pressure gradients, yaw angle, and swirl angle of the vortex, among others. The exact reason for this bursting is not known, but research has focused on two general areas. (a) The flows upstream and downstream of the vortex burst are two separate and very different flows, and the vortex burst is a necessary feature, similar to a hydraulic jump. (b) The core of the LE vortex serves as a mechanical waveguide for longitudinal waves; these waves either coalesce, or they become critical, thereby triggering the burst. Regardless of the burst triggering mechanism, the effect of the vortex burst is to reduce the lift generated by the delta wing. If the delta wing is pitched to a given angle of attack (α) and then maintained at that angle until the transient flow features die down, it is said to be tested under “static” conditions. As this 22nd Applied Aerodynamics Conference and Exhibit 16 19 August 2004, Providence, Rhode Island AIAA 2004-4731 Copyright

Impingement of a von Kàrmàn Vortex Street on a Dynamically Pitching Delta Wing, Part 1 - Large Cylinder Results

A series of experiments was conducted at Wichita State University to study the effect on the vortex burst position when a von Karman vortex street from a small cylinder was impinged upon a 70-degree sweep-back delta wing. Phase comparisons between different experimental runs were accomplished, to a limited degree, by controlling the circular cylinder’s vortex shedding. A small fin was attached on the cylinder’s downstream side, and the cylinder-fin arrangement was rotated at a frequency equal to the cylinder’s natural shedding frequency. The start of the delta wing’s pitch-up was then synchronized with the fin’s rotational position. Dye flow visualization showed that the vortex burst position appeared to jump forward towards the apex and then moved gradually back toward the trailing edge in sync with the passage of the von Karman vortices. The range of forward to rear-most variation in the burst position was about 10 to 30% of chord at an angle of attack of 35 degrees, and diminished to about 5 to 10% of chord at higher angles of attack.

Effect of Support Stem on a Dynamically Pitching Delta Wing

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.