An experimental study of static and dynamic vortex breakdown on slender delta wing planforms

Injection of jets from nozzles embedded in a delta wing were used to add momentum into the core of the leading edge vortices and thus delay their breakdown. The tests were performed on a delta wing with a 60° sweep angle at a Reynolds number of 260,000 and angles of attack varying from 12 to 20°. At a typical angle of attack of 15°, the vortex breakdown location was moved from its natural location at 55% of the wing chord to 90% near the trailing edge, using a blowing coefficient of Cu = 0.007. Though optimized for a = 15°, the jet blowing was effective in the entire range of angles of attack studied. Flow visualizations, which were performed by illuminating smoke with a laser sheet, showed the effect of the jet blowing on the vortex core and were used to determine the vortex trajectory and breakdown location. The mean axial velocity and turbulence intensity were mapped across the vortex at several axial locations using Laser Doppler Velocimetry. Vortex breakdown resulted in strong deceleration of the mean velocity in the core region, the appearance of negative flow and increased turbulence intensity. Fluctuations in the vortex breakdown location at a Strouhal number of 0.13 were recorded. The control jet injection accelerated the core velocity and stabilized the vortex as evidenced by a reduction in the turbulence intensity and a substantial shift of the vortex breakdown location towards the trailing edge. Nomenclature

An experimental study of the breakdown of leading-edge vortices on diamond, cropped, delta, and double delta wings during dynamic pitching

: The goal of this research is the control of leading-edge vortex breakdown location utilizing along-the-core blowing near the apex on the leeward surface of sharp-edged, slender, delta wings at high angles of attack. In the S2Ch subsonic wind tunnel at ONERA Chalais Meudon, two delta wing models with 70-deg sweep angles and root chords of 950 mm have been configured to collect qualitative and quantitative surface and flowfield data. First, an examination of the streamwise, time-dependent oscillation of the leading-edge vortex breakdown locations without active flow control is presented. These results further the understanding of the vortex breakdown phenomena and provide a more precise basis for evaluating the effectiveness of various flow control methods. Second, open-loop blowing along one of the vortex cores on the leeward surface of the delta wing demonstrates the ability to displace downstream the controlled, time-averaged, vortex breakdown location by 20% of the root chord.

The design, construction, manufacturing and geometric properties of a generic 65° Delta Wing for the International Vortex Flow Experiment 2 are described. Briefly, the measurement equipment implemented in the model and the external sensors are specified. A wind tunnel investigation has been performed and the results for two different leading edge geometries (sharp and rounded) at three angles of attack are discussed in detail. Further, the results are compared with the ones obtained by NASA and also with computational results. Partly developed (α = 13°) and fully developed (α = 18°) leading edge vortices as well as vortex breakdown (α = 23°) are considered.

The goal of this research is the control of leadingedge vortex breakdown location over sharp-edged, slender, delta wings at high angles of attack. A delta wing model with a 70” sweep angle and a root chord of 950 mm was used during these tests. Experiments utilizing open-loop blowing near the wing apex and along the vortex cores tangential to the leeward surface of the delta wing were accomplished in two subsonic wind tunnels at Onera. Asymmetric flow control along one vortex core demonstrates the ability to displace the controlled vortex breakdown location downstream. Asymmetric blowing, however, has no significant influence on the uncontrolled vortex breakdown location. Symmetric flow control along both vortex cores displaces both vortex breakdown locations downstream. The effectiveness of the flow control improves as the blowing mass flow rate is increased.

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 © 20 * Graduate Research Assistant, Student Member AIAA. † Associate Professor, Associate Fellow AIAA. Copyright ©2004 by the authors. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission.

An experimental investigation of hingeless flow control by vortex breakdown manipulation over a delta wing is described. Employed as a method of longitudinal control for slender-wing planforms, the performance of this active flow control technique was evaluated as a possible replacement to conventional moving or hinged control surfaces. Testing was conducted with a 60-deg delta-wing model incorporating pneumatic vortex control (PVC) actuators. These actuators delivered high-pressure jets of varying parameters on the upper surface of the model. The jets were used to alter the flowfield over the wing caused by leading-edge vortices. The natural location of vortex breakdown was shifted by three sets of PVC actuators strategically positioned on the model. This effect was intended to control the chordwise lift distribution over the wing by inducing both premature and delayed vortex breakdown, resulting in nose-up and nose-down pitch responses, respectively. A comparative analysis of this hingeless flow control concept was conducted in Texas A&M University's 3 × 4 ft low-speed wind tunnel and 2 x 3 ft water tunnel

A delta wing configuration with a sweep angle of 65 ° has been investigated numerically by detached eddy simulation and compared to experimental results. In the simulations different refinement settings of the grid as well a s time different time step sizes have been applied in order to reproduce the formation of the primary vortex correctly. For the comparison experimental data is available from steady and unsteady surface pressure measurements, hot wire anemometry and several other measurement techniques. I. Introduction avourable configuration for supersonic aircraft com monly are delta wings. There has been intense work on investigation of the complex flow field around delt a wings during the past decades. Two major projects concerning these flows were the international vorte x flow experiments (VFE) one and two. Both featured experimental as well as computational investigation s of delta wing wind tunnel models providing a larg e database of reference material for validation purposes. In VFE- 2 a delta wing with a sweep angle of 65° was invest igated featuring a sharp and several rounded leading edges 1 , of which only the sharp leading edge has been use d in the present study. In the simulations of the round lead ing edges the challenge is reproducing the correct separation behaviour behind the leading edge in order to achie ve the correct vortical flow field further downstre am. At higher angles of attack another phenomenon becomes present, the two primary vortices break down and thus the character of the flow field downstream of the b reakdown location changes drastically. The present work is focusing stronger on this vortex breakdown than on the flow separation. Thus it is limited to the shar p leading edge, which dictates the position for the separation, sin ce the flow cannot stay attached after passing the edge. Though its position thus is fixed the actual process of separa tion still needs to be handled carefully in the sim ulation, since it prescribes the formation of the vortices building t he fundament for the vortex breakdown. Common eddy viscosity RANS models without extensions for vortical flow usually are not able to reprodu ce these flows with vortex breakdown. Thus for the pre sent study an approach of detached eddy simulation (DES) has been selected. The underlying turbulence model is t he SST model 2 and the grid shield model is standard DES 3 . Several numerical meshes have been used to investigate the solution sensitivity to different refinemen t regions. Also the influence of time step size has been addressed. The simulations are compared with wind tunnel measurement results obtained in the low speed wind tunne l at the Technical University Munich.

Snapshot proper orthogonal decomposition is used to characterize the unsteady e ow past a delta wing, focusing on the structure of the leading-edge vortices. The decomposition is applied to two-dimensional velocity e elds obtained via a laser-scanning version of high-image-density particle image velocimetry to extract the most coherent structures in the e ow. Data are analyzed for cases involving both a stationary wing and a wing undergoing harmonic oscillations in roll. In each case, the velocity e elds are reconstructed as a truncated series expansion in terms of the computed eigenfunctions from which the corresponding contours of constant vorticity and sectional streamline patterns are calculated. Comparison with the original data demonstrates that the analysis provides an accurate representation of the velocity e elds while eliminating extraneous small-scale features. Inclusion of as few astwo eigenfunctionsin thereconstruction series reproduces the largest-scalefeaturesoftheleading-edge vortices, whereas the inclusion of half of the total number of eigenfunctions produces a reconstructed e eld that captures the majority of the e ow features. By appropriately combining the spatial and temporal components of the proper orthogonal decomposition analysis, one obtains the dynamical structures that evolve in both space and time, such as the e uctuations in the location of vortex breakdown. Nomenclature ak = modal amplitudes for velocity b = span C = correlation matrix co = centerline chord dI = interrogation window size E = mean total e uctuation energy M = total number of snapshots N = number of modes retained in Eq. (4) T = period of oscillation t = time U1 = freestream velocity u = x component of velocity V = velocity vector v = y component of velocity xvb = location of vortex breakdown, measured along wing centerline from apex x; y = coordinates ®k = kth eigenvector of the correlation matrix 1t = time interval between snapshots · = reduced frequency, .ob/=2U1T ¸k = kth eigenvalue of the correlation matrix 8 = roll angle Ak = kth empirical velocity eigenfunction N = time-averaged quantities 0 = e uctuation quantities

This paper presents an experimental study on the effects of roll on the loading and rolling moment of a 60 deg swept delta wing. The parameters studied were the incidence angle (including those in which vortex breakdown was present), roll angle, and ground proximity. Results indicate that when vortex breakdown is not present, the effects of roll on the wing loading and rolling moment are in reasonable agreement with predictions. When vortex breakdown is present, the magnitude of the rolling moment coefficient is reduced considerably and its sign can even be altered. The changes in the rolling moment characteristics due to vortex breakdown can be attributed to the roll-induced spanwise displacement of the leading-edge vortices. Effects of ground proximity on roll characteristics were shown to be small.

Normal force and rolling moment acting on a delta wing in forced sinusoidal rolling oscillation are measured in low speed wind tunnels. Flow visualization revieled details of the separated vortices. With increasing the rolling frequency, the normal force and the rolling moment show, within a cycle of motion, large hysteresises which due to time lag in leading edge separated vortices to follow the moving wing surface. The amplitude of hysteresis increases with angle of attack, up to the angle at which the vortex breakdown reaches the trailing edge of the wing. Beyond this angle nonlinearities are introduced by the vortex breakdown. The effect of breakdown is studied using a mathematical model. The result shows that the time lag in vortex movement ends in positive energy supply from free streams and the one in asymmetric vortex breakdown suppresses this.

An experimental investigation into the physical phenomena associated with limit cycle wing rock on slender delta wings has been conducted. The model used was a slender flat plate delta wing with 80-deg leading edge sweep. The investigation concentrated on three main areas: motion characteristics obtained from time history plots, static and dynamic flow visualization of vortex position, and static and dynamic flow visualization of vortex breakdown. The flow visualization studies are correlated with model motion to determine the relationship between vortex position and vortex breakdown with the dynamic rolling moments. Dynamic roll moment coefficient curves reveal rate-dependent hysteresis, which drives the motion. Vortex position correlated with time and model motion show a time lag in the normal position of the upward moving wing vortex. This time lag may be the mechanism responsible for the hysteresis. Vortex breakdown is shown to have a damping effect on the motion.

The three-dimensional, Reynolds-averaged, Navier-Stokes (RANS) equations are used to numerically simulate vortical flow about a 65 degree sweep delta wing. Subsonic turbulent flow computations are presented for this delta wing at 30 degrees angle of attack and static roll angles up to 42 degrees. This work is part of an on going effort to validate the RANS approach for predicting high-incidence vortical flows, with the eventual application to wing rock. The flow is unsteady and includes spiral-type vortex breakdown. The breakdown positions, mean surface pressures, rolling moments, normal forces, and streamwise center-of-pressure locations compare reasonably well with experiment. In some cases, the primary vortex suction peaks are significantly underpredicted due to grid coarseness. Nevertheless, the computations are able to predict the same nonlinear variation of rolling moment with roll angle that appeared in the experiment. This nonlinearity includes regions of local static roll instability, which is attributed to vortex breakdown.

Numerical investigation of vortical flows over a close-coupled delta canard-wing configuration

The effect of stream-wise fences on the twin-tail buffet responses of generic fighter aircraft is presented. The generic fighter aircraft consists of a sharp-edged delta wing of aspect ratio one and swept-back, flexible, vertical twin tail of taper ratio of 0.23. Twin streamwise fences are located at the 30% chord-station of the delta wing. The fences are used to alter the vortical flow and vortex breakdown location above the delta wing in order to alleviate the twin-tail buffet. The effect of the geometrical shape of the fences and the performance of the fences over wide range of angles of attack are investigated. This complex multi-disciplinary problem is solved using the multi-disciplinary computing environment (MDICE). The results indicated that the fences are effective in reducing the aeroelastic loads and responses, especially at angles of attack less than 30°. The fences of taper ratio of 0.7 produced the most favorable results.