Leading Edge Separation Vortex Research Articles

Separation control on NACA23012 using synthetic jet

A numerical study of separation control has been made to investigate aerodynamic characteristics of NACA23012 airfoil with synthetic jets. Computed results demonstrated that stall characteristics and control surface performance could be substantially improved by resizing separation vortices. The maximum lift was obtained when the separation point coincides with the synthetic jet location and the non-dimensional frequency is about 1. In addition, separation control effect was proportional to the peak velocity of the synthetic jet. It was observed that the actual flow control mechanism and flow structure is fundamentally different depending on the range of synthetic jet frequency. For low frequency range, small vortices due to synthetic jet penetrated to the large leading edge separation vortex, and as a result, the size of the leading edge vortex was remarkably reduced. For high frequency range, however, small vortex did not grow up enough to penetrate into the leading edge separation vortex. Instead, synthetic jet firmly attached the local flow and influenced the circulation of the virtual airfoil shape which is the combined shape of the main airfoil with the separation vortex. As a way to reduce the jet peak velocity, performance of a multi-array synthetic jet was investigated. Moreover, a high frequency multi-location synthetic jet was exploited to efficiently eliminate the unstable flow structure which was observed in low frequency range. Finally, by changing the phase angle in multi-location synthetic jets, highly controlled flow characteristics could be obtained with multi-array/multi-location synthetic jets. This shows efficiency of the current approach in separation control using synthetic jet.

Visualization by PIV of dynamic stall on a vertical axis wind turbine

The aerodynamic behavior of a vertical axis wind turbine (VAWT) is analyzed by means of 2D particle image velocimetry (PIV), focusing on the development of dynamic stall at different tip speed ratios. The VAWT has an unsteady aerodynamic behavior due to the variation with the azimuth angle θ of the blade’s sections’ angle of attack, perceived velocity and Reynolds number. The phenomenon of dynamic stall is then an inherent effect of the operation of a VAWT at low tip speed ratios, impacting both loads and power. The present work is driven by the need to understand this phenomenon, by visualizing and quantifying it, and to create a database for model validation. The experimental method uses PIV to visualize the development of the flow over the suction side of the airfoil for two different reference Reynolds numbers and three tip speed ratios in the operational regime of a small urban wind turbine. The field-of-view of the experiment covers the entire rotation of the blade and almost the entire rotor area. The analysis describes the evolution of the flow around the airfoil and in the rotor area, with special focus on the leading edge separation vortex and trailing edge shed vorticity development. The method also allows the quantification of the flow, both the velocity field and the vorticity/circulation (only the results of the vorticity/circulation distribution are presented), in terms of the phase locked average and the random component.

外翼前縁後退角変化がSST形態の低速ロール特性に及ぼす影響について

Wind tunnel tests were conducted to investigate static roll characteristics of three cranked arrow wings that have different outboard wing sweepback angles to investigate the effect of inboard and outboard vortices upon the wing roll characteristics. Results indicated that breakdown position of leading-edge separation vortex formed on the inboard wing moves upstream as the outboard wing sweepback angle is decreased. A large scale rolling moment hysteresis was observed for the wings with smaller outboard wing sweepback angle, whereas only a small scale hysteresis was observed for higher outboard wing sweepback angle. It was shown that these differences are caused by the different behaviors of inboard vortex breakdown.

Separation control mechanism of airfoil using synthetic jet

Numerical simulation of separation control using a synthetic jet was performed on NACA23012 airfoil. The computed results showed that stall characteristics and control surface performance could be improved substantially by resizing the separation vortices. It was observed that actual flow control mechanism was fundamentally different depending on the range of synthetic jet frequency. For low frequency range, small vortices due to synthetic jet penetrated to the large leading edge separation vortex flow, and as a result, the size of the leading edge separation vortex remarkably decreased. For high frequency range, however, the small vortex did not grow enough to penetrate into the large separation vortex, but the synthetic jet changed airfoil circulation directly. The synthetic jet conditions for effective lift increase are as follows: the non-dimensional frequency of the synthetic jet is 1; the location of the synthetic jet slot is the same as the separation point; and the jet velocity is large enough to perturb the separated flow. By exploiting these conditions, it was observed that the combination of the synthetic jet with a simple high lift device could be as good as a conventional fowler flap system.

Flow visualization and aerodynamic load calculation of three types of clap-fling motions in a Weis-Fogh mechanism

A flow visualization and aerodynamic load calculation of three types of clap-fling motions in a Weis-Fogh mechanism are performed through experiment and computation. In the experiment, the flow development is visualized using a smoke-wire technique for two wing motion types: ‘fling followed by clap and pause’, and ‘clap followed by fling and pause’. For the computation, the two-dimensional Navier–Stokes equations are solved for the same type of wing motions of the experiment. A good agreement between the experimental and numerical result is observed regarding the main flow features, such as the sequential development of the two families of separation vortex pairs and their movement. For the ‘fling followed by clap and pause motion’, a strong separation vortex pair of counter-rotation develops in the opening between the wings in the fling phase and then moves out from the opening in the following clap phase. For the ‘clap followed by fling and pause motion’, the separation vortex pair develops in the outside space in the clap phase, and then moves into the opening in the following fling phase. Results show that the leading edge separation vortices are dominant features of the flow pattern and cause a large negative pressure distribution near the leading edge. In the ‘cyclic fling and clap motion’ case, the aerodynamic loads of the fling phase after the second cycle of the wing motion when the periodicity of the flow pattern and the aerodynamic loads are established, are much greater than those of the first cycle. However, the aerodynamic loads of the clap phase in the later cycles are not much different from those of the first cycle.

Rolling Moment Characteristics of Supersonic Transport Configurations at High Incidence Angles

Wind-tunnel tests were performed to investigate rolling moment characteristics of a cranked arrow wing supersonic transport configuration at high incidence angles. Force measurements and surface static pressure measurements were conducted when the wing model was rolled statically at a Reynolds number of 9.21 x 10 3 based on the mean aerodynamic chord. To understand the behavior of the leading-edge separation vortices formed both on the inboard and outboard of the cranked arrow wing, a stereoscopic particle image velocimetry survey was performed at 55 and 83% root chord locations. As the wing roll angle was increased, linear stable rolling moments were observed at low incidence angles, whereas abrupt changes from stable to unstable rolling moments were observed at incidence angle of 20 deg. Variations in suction peaks of pressure distributions at the rear part of the wing contribute to destabilization of the rolling moment component. These variations in suction peaks that induce the unstable rolling moment are related to the vortex breakdown characteristics. Chordwise breakdown locations of the inboard and outboard vortices are different between the windward and leeward wings. This difference causes the sudden change of rolling moments.

Effects of Leading Edge Separation Vortex of Flexible Structure Delta Wing on Its Aerodynamic Characteristics

The aerodynamic forces and moments of a flexible delta wing in pitching motion were experimentally studied in a low-speed wind tunnel. Three types of flexible delta wing were investigated, the flexible parts of which were 44, 70 and 99% of the delta wing. Aerodynamic characteristics were different among the three types of flexible and completely hard delta wing, and it was found that the winding-up of the leading edge of the delta wing is the key factor for determining the leading edge vortex on the upper side of the wing and the pressure distribution on the windward side. Lift, drag, and pitching moment formed a hysteresis loop with an angle of attack in pitching motion, particularly in a region with a large attack angle, accompanied by leading edge vortex breakdown. The flow visualization of leading edge vortices was also carried out to explain the dynamic characteristics of the delta wings.

Effects of Leading Edge Separation Vortex of Flexible Structure Delta Wing on Its Aerodynamic Characteristics(Wing and Airfoil)

Effects of Leading Edge Separation Vortex of Flexible Structure Delta Wing on Its Aerodynamic Characteristics(Wing and Airfoil)

Stereoscopic PIV measurements of leading edge separation vortices on a cranked arrow wing

Three-component velocity measurement results via stereoscopic PIV for leeward flow of a cranked arrow wing with a centre body are presented in order to understand leeward flow phenomena, emphasizing angle-of-attack effects and Reynolds number effects. The tests are conducted in a low-speed wind tunnel with the Reynolds number range from 0.15 × 106 to 1.5 × 106 at angles of attack of 8° to 20°. Based on instantaneous results, statistical properties such as turbulence kinetic energy are derived. The test results indicate that with increase in angle of attack, interaction between two leading edge separation vortices and a vortex on the body becomes strong, merging into a single vortex. In results on angle of attack effects, vorticity magnitude and kinetic energy have a positive correlation in the cores of the separation vortices. Results of Reynolds number effects show that with increasing Reynolds number, the secondary vortex structure becomes small with the two primary separation vortices moving outward, and the core diameter and vorticity of the vortices become larger and smaller, respectively, due to boundary layer transition on the inboard wing. In general, these Reynolds number effects are similar to the effects of decrease in angle of attack.

Numerical Analysis of Aerodynamic Control of Delta Wing by Microflap

Computational fluid dynamics (CFD) simulation is performed to clarify the mechanism of rolling moment generation due to a delta wing with a microflap. As a result, it was found that a leading-edge separation vortex is displaced by changing the location of the microflap at the leading edge. The positive rolling moment was produced for installation of the flap in the lower part of the leading edge, whereas the negative rolling moment was produced in the upper part. These characteristics are similar to experimental data. Detailed observation of simulated results in turbulent case shows that the leading-edge separation vortex slightly moves toward the outboard direction in the former case, whereas it does not move toward inboard, but becomes diffusive, in the latter case

High-Lift Devices for a Delta Wing Installed Around a Trailing-Edge

Three devices to improve the aerodynamic characteristics of a delta wing at low speeds and high angles of attack are computationally examined. These are two types of blowing at the trailing edge and a slotted flap. All three are commonly employed on rectangle wings. They all show favorable effects, namely, chordwise blowing increases the lift considerably. The effectiveness in strengthening the leading-edge vortex is higher for chordwise blowing than lateral blowing. It is observed that, although the jet slots are installed at a trailing edge, the blowing strengthens the leading-edge separation vortex over the delta wing. This motivated an investigation of the use of a slotted flap. The flap also increases the lift of the wing, as well as the total lift due to both the wing and the flap, but the effect on the leading-edge vortices is smaller than both types of blowing

Frequency Characteristics of Fluctuating Pressure on Rotor Blade in a Propeller Fan.

A wavelet transform is introduced to analyze frequency characteristics of the fluctuating pressure on rotor blade in a propeller fan. The fluctuating pressure on the rotor blade is obtained by using the results of a large eddy simulation. The frequencies having high spectral peaks of the fluctuating pressure are determined by taking the time average of the local absolute modulus of the wavelet. The dominant frequency of the real-time pressure selected at the high pressure fluctuation region corresponds well to that of the fluctuating rotor torque and the experimental result of fan noise. It is mainly generated due to the unsteady behavior of the vortical flow, such as the tip vortex and the leading edge separation vortex, in the propeller fan. A frequency in the separation bubble region on the suction surface is higher than that of the dominant frequency caused by the vortical flow.

スピンするデルタ翼の空力特性について

The experiments of a delta wing’s spin phenomena were conducted at the low speed wind tunnel of Nagoya University with its exit test section inclined vertically. The spin rate of the delta wing in the free rotation mode was measured as well as surface pressure distributions on the spinning delta wing. At low angles of attack the spin rate increases with sideslip angle, whereas near the stall angle, it decreases. At low angles of attack, when the wing has a sideslip angle, the upper surface pressures of the windward wing-half become lower than those of the leeward wing-half, while near the stall angle the latter become lower than the former. These asymmetric pressure distributions cause the spinning motion of the delta wing. It is evident from these pressure distributions and flow visualizations that a leading edge separation vortex makes an important role in spin phenomena of the delta wing. It is also found that at the equilibrium spin rate, the spanwise distribution of the pressure difference between upper and lower surfaces, ΔCp , shows a nearly symmetric pattern regarding the wing centerline.

On aerodynamic modelling of an insect–like flapping wing in hover for micro air vehicles

This theoretical paper discusses recent advances in the fluid dynamics of insect and micro air vehicle (MAV) flight and considers theoretical analyses necessary for their future development. The main purpose is to propose a new conceptual framework and, within this framework, two analytic approaches to aerodynamic modelling of an insect-like flapping wing in hover in the context of MAVs. The motion involved is periodic and is composed of two half-cycles (downstroke and upstroke) which, in hover, are mirror images of each other. The downstroke begins with the wing in the uppermost and rearmost position and then sweeps forward while pitching up and plunging down. At the end of the half-cycle, the wing flips, so that the leading edge points backwards and the wing's lower surface becomes its upper side. The upstroke then follows by mirroring the downstroke kinematics and executing them in the opposite direction. Phenomenologically, the interpretation of the flow dynamics involved, and adopted here, is based on recent experimental evidence obtained by biologists from insect flight and related mechanical models. It is assumed that the flow is incompressible, has low Reynolds number and is laminar, and that two factors dominate: (i) forces generated by the bound leading-edge vortex, which models flow separation; and (ii) forces due to the attached part of the flow generated by the periodic pitching, plunging and sweeping. The first of these resembles the analogous phenomenon observed on sharp-edged delta wings and is treated as such. The second contribution is similar to the unsteady aerodynamics of attached flow on helicopter rotor blades and is interpreted accordingly. Theoretically, the fluid dynamic description is based on: (i) the superposition of the unsteady contributions of wing pitching, plunging and sweeping; and (ii) adding corrections due to the bound leading-edge vortex and wake distortion. Viscosity is accounted for indirectly by imposing the Kutta condition on the trailing edge and including the influence of the vortical structure on the leading edge. Mathematically, two analytic approaches are proposed. The first derives all the quantities of interest from the notion of circulation and leads to tractable integral equations. This is an application of the von Kármán-Sears unsteady wing theory and its nonlinear extensions due to McCune and Tavares; the latter can account for the bound leading-edge vortex and wake distortion. The second approach uses the velocity potential as the central concept and leads to relatively simple ordinary differential equations. It is a combination of two techniques: (i) unsteady aerodynamic modelling of attached flow on helicopter rotor blades; and (ii) Polhamus's leading-edge suction analogy. The first of these involves both frequency-domain (Theodorsen style) and time-domain (indicial) methods, including the effects of wing sweeping and returning wake. The second is a nonlinear correction accounting for the bound leading-edge vortex. Connections of the proposed framework with control engineering and aeroelasticity are pointed out.

デルタ翼周りの流れとその可視化

Leading-edge separation vortex formed on a delta wing surface is the main cause of poor low-speed aerodynamic characteristics of the delta wing. Some results of flow visualizations around delta wings are summarized in this paper.

Analysis of Vortical Flow Field in a Propeller Fan by LDV Measurements and LES—Part I: Three-Dimensional Vortical Flow Structures

Three-dimensional structures of the vortical flow field in a propeller fan with a shroud covering only the rear region of its rotor tip have been investigated by experimental analysis using laser Doppler velocimetry (LDV) measurements and by numerical analysis using a large eddy simulation (LES) in Part I of the present study. The propeller fan has a very complicated vortical flow field near the rotor tip compared with axial fan and compressor rotors. It is found that three vortex structures are formed near the rotor tip: the tip vortex, the leading edge separation vortex, and the tip leakage vortex. The tip vortex is so strong that it dominates the flow field near the tip. Its formation starts from the blade tip suction side near the midchord. Even at the design condition the tip vortex convects nearly in the tangential direction, thus impinging on the pressure surface of the adjacent blade. The leading edge separation vortex develops close along the tip suction surface and disappears in the rear region of the rotor passage. The tip leakage vortex is so weak that it does not affect the flow field in the rotor.

Analysis of Vortical Flow Field in a Propeller Fan by LDV Measurements and LES—Part II: Unsteady Nature of Vortical Flow Structures Due to Tip Vortex Breakdown

The unsteady nature of vortex structures has been investigated by a large eddy simulation (LES) in a propeller fan with a shroud covering only the rear region of its rotor tip. The simulation shows that the tip vortex plays a major role in the structure and unsteady behavior of the vortical flow in the propeller fan. The spiral-type breakdown of the tip vortex occurs near the midpitch, leading to significant changes in the nature of the tip vortex. The breakdown gives rise to large and cyclic movements of the tip vortex, so that the vortex impinges cyclically on the pressure surface of the adjacent blade. The movements of the tip vortex cause the leading edge separation vortex to oscillate in a cycle, but on a small scale. The movements of the vortex structures induce high-pressure fluctuations on the rotor blade and in the blade passage.

Flow field measurements of leading-edge separation vortex formed on a delta wing with vortex flaps

Wind tunnel tests are carried out using a 70 · delta wing model with leading-edge vortex flaps. The structure of the leading-edge separation vortex over the leading-edge vortex flap is measured by use of a 5 holes pitot probe, surface pressure measurement technique and oil flow visualization technique. Separation vortices formed on a plain delta wing, on a vortex flap and inboard the vortex flap hinge line are clearly visualized. Results indicate that the flow around the vortex flaps is classified into several different cross flow patterns. The streamwise flap deflection angle is defined to discuss the vortex flap performance. The optimum lift to drag ratio is attained when the amount of the wing angle of attack is not far different from that of the streamwise flap deflection angle, as long as the vortex flap is deflected modestly.

Unsteady flow structure and vorticity convection over the airfoil oscillating at high reduced frequency

Abstract Unsteady vortex structures and vorticity convection over the airfoil (NACA 0012), oscillating in the uniform inflow, are studied by flow visualization and velocity measurements. The airfoil, pivoting at one-third of the chord, oscillates periodically near the static stalling angle of attack (AOA) at high reduced-frequency. The phase-triggering and modified phase-averaged techniques are employed to reconstruct the pseudo instantaneous velocity field over the airfoil. During the down stroke cycle, the leading-edge separation vortex is growing and the vortex near the trailing edge begins to shed into the wake. During the upstroke cycle, the leading-edge separation vortex is matured and moves downstream, and the counter clockwise vortex is forming near the trailing edge. Convection speeds and wavelength of the unsteady vortex structure over the airfoil equal to that of the counter clockwise vortex shed into the wake. This kind of vortex structure is termed as “synchronized shedding” type. The wavelength of unsteady vortex structure over the airfoil is significantly different from that at low reduced-frequency. Consistent convection speeds of the leading-edge separation vortex are acquired from the spatial-temporal variations of local circulation and local surface vorticity generation, and equals that predicted from flow visualization. Spatial-temporal variations of the local surface vorticity generation clearly reveal the formation and passage of the leading-edge separation vortex only in the region where the flow does not separate completely from the surface. Significant amounts of the surface vorticity are generated within the leading-edge region of the airfoil during the upstroke cycle. Only negligible amount of surface vorticity is produced within the region of complete flow separation. During the down stroke cycle, the surface vorticity generation is mild along the airfoil surface, except the leading-edge region where a small scale leading-edge separation vortex is forming and growing.

Parallel Interaction of Incident Vortex Array with Oscillating Airfoil

Parallel interaction of an incident vortex array with a NACA 0012 airfoil in the form of sinusoidal oscillation was investigated experimentally, based on the concept of vorticity balance. It was found that the temporal ‐spatial variations of the local circulation can clearly reveal the convection of the highly distorted vorticity e eld over the airfoil. The phase-averaged near-surface e ow structure cone rmed that the temporal ‐spatial variations of the surface vorticity generation are closely related to the formation and subsequent motion of the leading-edge separation vortex along the airfoil. During the upstroke portion of the oscillation cycle, signie cant amounts of the surface vorticity were generated from the leading edge of the airfoil. The phase information of the local circulation relative to that of the surface vorticity generation clearly indicated that the leading-edge separation vortex was induced immediately after the passage of the clockwise-rotating incident vortex. When compared with the oscillating airfoil in uniform e ow, the impingement of the incident vortex array on the oscillating airfoil increased the mean overall circulation around the airfoil over the oscillation cycle.