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Showing papers on "Airfoil published in 2009"


Journal ArticleDOI
TL;DR: In this paper, the aerodynamic behavior of a vertical axis wind turbine is analyzed by means of 2D particle image velocimetry (PIV), focusing on the development of dynamic stall at different tip speed ratios.
Abstract: 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.

285 citations


Journal ArticleDOI
TL;DR: In this paper, the origins of transonic aerofoil buffet are linked to a global instability, which leads to shock oscillations and dramatic lift fluctuations, and the stability boundary, as a function of the Mach number and angle of attack, consists of an upper and a lower branch.
Abstract: Buffeting flow on transonic aerofoils serves as a model problem for the more complex three-dimensional flows responsible for aeroplane buffet. The origins of transonic aerofoil buffet are linked to a global instability, which leads to shock oscillations and dramatic lift fluctuations. The problem is analysed using the Reynolds-averaged Navier–Stokes equations, which for the foreseeable future are a necessary approximation to cover the high Reynolds numbers at which transonic buffet occurs. These equations have been shown to reproduce the key physics of transonic aerofoil flows. Results from global-stability analysis are shown to be in good agreement with experiments and numerical simulations. The stability boundary, as a function of the Mach number and angle of attack, consists of an upper and a lower branch – the lower branch shows features consistent with a supercritical bifurcation. The unstable modes provide insight into the basic character of buffeting flow at near-critical conditions and are consistent with fully nonlinear simulations. The results provide further evidence linking the transonic buffet onset to a global instability.

237 citations


Journal ArticleDOI
TL;DR: In this article, the simulation of the unsteady separated flows encountered by a plunging airfoil under low-Reynolds-number conditions (Rec 6 ◊ 10 4 ).
Abstract: This investigation addresses the simulation of the unsteady separated flows encountered by a plunging airfoil under low-Reynolds-number conditions (Rec 6 ◊ 10 4 ). The flow fields are computed employing a previously developed and extensively validated high-fidelity implicit large-eddy simulation (ILES) approach. In order to permit comparison with available experimental measurements, calculations are performed first for an SD7003 airfoil section at an angle of attack o = 4 plunging with reduced frequency k = 3.93 and nondimensional amplitude ho = 0.05. Under these conditions, it is demonstrated that for Rec = 10 4 , transitional effects are not significant and that the dynamic-stall vortices remain fairly coherent as they propagate along the airfoil. For Rec = 4 ◊ 10 4 , the dynamic-stall vortex system is laminar at is inception, however shortly afterwards, it experiences an abrupt breakdown associated with the onset of spanwise instability effects. A detailed description of this transition process near the leading edge is provided. The computed phased-averaged structures for both values of Reynolds number are found to be in good agreement with the experimental data. As a second example, the suppression of static stall at high angle of attack ( o = 14 ) is investigated using high-frequency small-amplitude vibrations (k = 10,ho = 0.005). At Rec = 6 ◊ 10 4 , separation is completely eliminated in a time-averaged sense, and the mean drag is reduced by approximately 40%. The instantaneous flow is characterized by the periodic generation of dynamic-stall vortices near the leading edge and by their subsequent transition as they convect close to the airfoil. For Rec = 10 4 , significant reduction of the timeaveraged separation region is still possible with transitional effects present in the aft-portion of the airfoil. For larger forcing amplitude (ho = 0.04,Rec = 10 4 ), a very intriguing regime emerges. The dynamic stall vortex moves around and in front of the leading edge and experiences a dramatic breakdown as it impinges against the airfoil. As a result, the phased-averaged flow displays no coherent vortices propagating along the airfoil upper surface. This new flow structure is also characterized in the mean by the existence of a strong jet in the near wake which manifests in a high value of net thrust. The present study demonstrates the importance of transitional effects for low-Reynolds-number maneuvering airfoils, as well as the suitability of the ILES approch for exploring such flow regime.

172 citations


Journal ArticleDOI
TL;DR: In this article, an experimental study of unsteady aerodynamics of two-dimensional membrane airfoils at low Reynolds numbers was conducted, where the amplitude and mode of the vibrations of the membrane depend on the relative location and the magnitude of the unsteadiness of the separated shear layer.
Abstract: Membrane wings are used both in nature and small aircraft as lifting surfaces. Separated flows are common at low Reynolds numbers and are the main sources of unsteadiness. Yet, the unsteady aspects of the fluid-structure interactions of membrane airfoils are largely unknown. An experimental study of unsteady aerodynamics of two-dimensional membrane airfoils at low Reynolds numbers has been conducted. Measurements of membrane shape with a high-speed camera were complemented with the simultaneous measurements of unsteady velocity field with a high frame-rate particle image velocimetry system and flow visualization. Vibrations of the membrane and mode shapes were investigated as a function of angle of attack and free stream velocity. While the mean membrane shape is not very sensitive to angle of attack, the amplitude and mode of the vibrations of the membrane depend on the relative location and the magnitude of the unsteadiness of the separated shear layer. The results indicate strong coupling of unsteady flow with the membrane oscillations. There is evidence of coupling of membrane oscillations with the vortex shedding in the wake, in particular, for the post-stall incidences. Comparison of rigid (but cambered) and flexible membrane airfoils shows that the flexibility might delay the stall. Hence this is a potential passive flow control method using flexibility in nature and engineering applications.

159 citations


Journal ArticleDOI
TL;DR: The analytical model of the trailing-edge noise of an airfoil derived in the first part of this study is assessed by first comparing the predictions with alternative analytical and numerical computations found in the literature as mentioned in this paper.

140 citations


Journal ArticleDOI
TL;DR: In this paper, a second-order Navier-Stokes solver is employed coupled with a membrane structural model suitable for the highly nonlinear structural response of the membrane, which results in a delay in stall with enhanced lift for higher angles of attack.

138 citations


Journal ArticleDOI
TL;DR: In this paper, the NACA0012 airfoil geometry is extended in chord so that its trailing edge is sharp and a family of grid-convergence trends of two-dimensional Euler solutions are investigated.
Abstract: Grid-convergence trends of two-dimensional Euler solutions are investigated. The airfoil geometry under study is based on the NACA0012 equation. However, unlike the NACA0012 airfoil, which has a blunt base at the trailing edge, the study geometry is extended in chord so that its trailing edge is sharp. The flow solutions use extremely- high-quality grids that are developed with the aid of the Karman-Trefftz conformal transformation. The topology of each grid is that of a standard O-mesh. The grids naturally extend to a far-field boundary approximately 150 chord lengths away from the airfoil. Each quadrilateral cell of the resulting mesh has an aspect ratio of one. The intersecting lines of the grid are essentially orthogonal at each vertex within the mesh. A family of grids is recursively derived starting with the finest mesh. Here, each successively coarser grid in the sequence is constructed by eliminating every other node of the current grid, in both computational directions. In all, a total of eight grids comprise the family, with the coarsest-to-finest meshes having dimensions of 32 x 32-4096 x 4096 cells, respectively. Note that the finest grid in this family is composed of over 16 million cells, and is suitable for 13 levels of multigrid. The geometry and grids are all numerically defined such that they are exactly symmetrical about the horizontal axis to ensure that a nonlifting solution is possible at zero degrees angle-of-attack attitude. Characteristics of three well-known flow solvers (FLO82, OVERFLOW, and CFL3D) are studied using a matrix of four flow conditions: (subcritical and transonic) by (nonlifting and lifting). The matrix allows the ability to investigate grid-convergence trends of 1) drag with and without lifting effects, 2) drag with and without shocks, and 3) lift and moment at constant angles-of-attack. Results presented herein use 64-bit computations and are converged to machine-level-zero residuals. All three of the flow solvers have difficulty meeting this requirement on the finest meshes, especially at the transonic flow conditions. Some unexpected results are also discussed. Take for example the subcritical cases. FLO82 solutions do not reach asymptotic grid convergence of second-order accuracy until the mesh approaches one quarter of a million cells. OVERFLOW exhibits at best a first-order accuracy for a central-difference stencil. CFL3D shows second-order accuracy for drag, but only first-order trends for lift and pitching moment. For the transonic cases, the order of accuracy deteriorates for all of the methods. A comparison of the limiting values of the aerodynamic coefficients is provided. Drag for the subcritical cases nearly approach zero for all of the computational fluid dynamics methods reviewed. These and other results are discussed.

138 citations


Journal ArticleDOI
TL;DR: The actuator surface technique is applied to compute the flow past a two-bladed vertical axis wind turbine equipped with NACA 0012 airfoils and comparisons with experimental data show an encouraging performance of the method.
Abstract: This paper presents a new numerical technique for simulating two-dimensional wind turbine flow. The method, denoted as the 2D actuator surface technique, consists of a two-dimensional Navier-Stokes solver in which the pressure distribution is represented by body forces that are distributed along the chord of the airfoils. The distribution of body force is determined from a set of predefined functions that depend on angle of attack and airfoil shape. The predefined functions are curve fitted using pressure distributions obtained either from viscous-inviscid interactive codes or from full Navier-Stokes simulations. The actuator surface technique is evaluated by computing the two-dimensional flow past a NACA 0015 airfoil at a Reynolds number of 10 6 and an angle of attack of 10 deg and by comparing the computed streamlines with the results from a traditional Reynolds-averaged Navier-Stokes computation. In the last part, the actuator surface technique is applied to compute the flow past a two-bladed vertical axis wind turbine equipped with NACA 0012 airfoils. Comparisons with experimental data show an encouraging performance of the method.

137 citations


Journal ArticleDOI
TL;DR: A detailed analysis of the wing kinematics and wing deformations of desert locusts flying tethered in a wind tunnel implies tuning of the structural, morphological and kinematic parameters of the hindwing for efficient aerodynamic force production.
Abstract: Here, we present a detailed analysis of the wing kinematics and wing deformations of desert locusts (Schistocerca gregaria, Forskal) flying tethered in a wind tunnel. We filmed them using four high-speed digital video cameras, and used photogrammetry to reconstruct the motion of more than 100 identified points. Whereas the hindwing motions were highly stereotyped, the forewing motions showed considerable variation, consistent with a role in flight control. Both wings were positively cambered on the downstroke. The hindwing was cambered through an 'umbrella effect' whereby the trailing edge tension compressed the radial veins during the downstroke. Hindwing camber was reversed on the upstroke as the wing fan corrugated, reducing the projected area by 30 per cent, and releasing the tension in the trailing edge. Both the wings were strongly twisted from the root to the tip. The linear decrease in incidence along the hindwing on the downstroke precisely counteracts the linear increase in the angle of attack that would otherwise occur in root flapping for an untwisted wing. The consequent near-constant angle of attack is reminiscent of the optimum for a propeller of constant aerofoil section, wherein a linear twist distribution allows each section to operate at the unique angle of attack maximizing the lift to drag ratio. This implies tuning of the structural, morphological and kinematic parameters of the hindwing for efficient aerodynamic force production.

130 citations


Journal ArticleDOI
TL;DR: In this article, a dielectric barrier discharge (DBD) is mounted at the leading edge of a NACA 0015 airfoil model and the effects of steady and unsteady actuations on the lift and drag coefficients are investigated by time-averaged force measurements.

129 citations


Journal ArticleDOI
TL;DR: In this article, a set of experiments conducted on a NACA0012 airfoil undergoing stall flutter oscillations in a low-speed wind tunnel is presented, with the objective of characterizing the local bifurcation behavior of the system.
Abstract: Stall flutter is a nonlinear aeroelastic phenomenon that can affect several types of aeroelastic systems such as helicopter rotor blades, wind turbine blades, and highly flexible wings. Although the related aerodynamic phenomenon of dynamic stall has been the subject of many experimental studies, stall flutter itself has rarely been investigated. This paper presents a set of experiments conducted on a NACA0012 airfoil undergoing stall flutter oscillations in a low-speed wind tunnel. The aeroelastic responses are analyzed with the objective of characterizing the local bifurcation behavior of the system. It is shown that symmetric stall flutter oscillations are encountered as a result of a subcritical Hopf bifurcation, followed by a fold bifurcation. The cause of these bifurcations is the occurrence of dynamic stall, which allows the transfer of energy from the freestream to the wing. A second bifurcation occurs at the system's static divergence airspeed. As a consequence, the wing starts to undergo asymmetric stall flutter bifurcations at only positive (or only negative) pitch angles. The dynamic stall mechanism itself does not change but the flow only separates on one side of the wing.

Journal ArticleDOI
TL;DR: In this paper, the formation process of leading-edge vortices has been investigated experimentally using Particle Image Velocimetry and various airfoil kinematics have been tested, including asymmetric and peak-shifted plunging motions, and evaluated for Re = 30,000 and a reduced frequency range of 0.2 ≤ k ≤ 0.33.
Abstract: The formation process of leading-edge vortices has been investigated experimentally using Particle Image Velocimetry. Various airfoil kinematics have been tested, including asymmetric and peak-shifted plunging motions, and are evaluated for Re = 30,000 and a reduced frequency range of 0.2 ≤ k ≤ 0.33. By measuring the growth in the leading-edge vortex during the dynamic-stall process, the vortex pinch-off process is examined based on the concept of an optimal vortex formation time. The various kinematics are then evaluated with respect to their associated vortex strength, timing and convection into the wake.

Journal ArticleDOI
TL;DR: A computational framework for simulating structural models of varied fidelity and a Navier-Stokes solver, aimed at simulating flapping and flexible wings, and implications of fluid density on aerodynamic loading are explored.
Abstract: Because of their small size and flight regime, coupling of aerodynamics, structural dynamics, and flight dynamics are critical for micro aerial vehicles This paper presents a computational framework for simulating structural models of varied fidelity and a Navier-Stokes solver, aimed at simulating flapping and flexible wings The structural model uses either 1) the in-house developed UM/NLABS, which decomposes the equations of 3-D elasticity into cross-sectional and spanwise analyses for slender wings, or 2) MSCMarc, which is a commercial finite-element solver capable of modeling geometrically nonlinear structures of arbitrary geometry The flow solver employs a well-tested pressure-based algorithm implemented in STREAM A NACA0012 cross-sectional rectangular wing of aspect ratio 3, chord Reynolds number of 3 x 10 4 , and reduced frequency varying from 04 to 182, with prescribed pure plunge motion is investigated Both rigid and flexible wing results are presented, and good agreement between experiment and computation are shown regarding tip displacement and thrust coefficient Issues related to coupling strategies, fluid physics associated with rigid and flexible wings, and implications of fluid density on aerodynamic loading are also explored in this paper

Journal ArticleDOI
TL;DR: In this article, a V-22 wing/nacelle combination with discrete jets pointing in the direction of streaming and sweeping side to side along the span was used to delay flow separation.
Abstract: Experiments aimed at delaying flow separation through discrete jets pointing in the direction of streaming and sweeping side to side along the span were conducted on a V-22 airfoil with and without deflected trailing-edge flaps. The results indicated substantial drag reduction and lift increase at moderately low inputs of mass and momentum. Additional experiments were carried out on a semispan V-22 wing/nacelle combination, and they too provided an increase in lift-to-drag ratio L/D of approximately 60% (although active flow control was applied to the wing only). The effectiveness of the sweeping jets on reducing the download force acting on a V-22 full-span powered model in hover was also examined. A 29% reduction in download was realized using the embedded sweeping jets, corresponding approximately to a 2000 1b increase in hover lift.

Journal ArticleDOI
TL;DR: In this paper, the correlation-based transition model has been applied to flow over a flat plate, flow over the S809 and the NACA63-415 airfoils, and finally to the NREL Phase VI wind turbine rotor for the zero yaw upwind cases from NREL/NASA Ames wind tunnel test.
Abstract: When predicting the flow over airfoils and rotors, the laminar-turbulent transition process can be important for the aerodynamic performance. Today, the most widespread approach is to use fully turbulent computations, where the transitional process is ignored and the entire boundary layer on the wings or airfoils is handled by the turbulence model. The correlation based transition model has lately shown promising results, and the present paper describes the effort of deriving the two non-public empirical correlations of the model to make the model complete. To verify the model it is applied to flow over a flat plate, flow over the S809 and the NACA63-415 airfoils, flow over a prolate spheroid at zero and thirty degrees angle of attack, and finally to the NREL Phase VI wind turbine rotor for the zero yaw upwind cases from the NREL/NASA Ames wind tunnel test. Copyright © 2009 John Wiley & Sons, Ltd.

Journal ArticleDOI
TL;DR: In this article, the aerodynamic properties of two-dimensional membrane airfoils were experimentally investigated in a wind tunnel and the effects of the membrane pre-strain and excess length on the unsteady aspects of the fluid-structure interaction were studied.

Journal ArticleDOI
TL;DR: In this paper, the effect of grid resolution and grid quality on aerodynamic drag prediction has been examined in detail and recommendations for improvements in mesh generation technology which have the potential to impact the state-of-the-art of aerodynamic prediction are given.
Abstract: The drag prediction workshop series (DPW), held over the last six years, and sponsored by the AIAA Applied Aerodynamics Committee, has been extremely useful in providing an assessment of the state-of-the-art in computationally based aerodynamic drag prediction. An emerging consensus from the three workshop series has been the identification of spatial discretization errors as a dominant error source in absolute as well as incremental drag prediction. This paper provides an overview of the collective experience from the worksho series regarding the effect of grid-related issues on overall drag prediction accuracy. Examples based on workshop results are used to illustrate the effect of grid resolution and grid quality on drag prediction, and grid convergence behavior is examined in detail. For fully attached flows, various accurate and successful workshop results are demonstrated, while anomalous behavior is identified for a number of cases involving substantial regions of separated flow. Based on collective workshop experiences, recommendations for improvements in mesh generation technology which have the potential to impact the state-of-the-art of aerodynamic drag prediction are given.

Journal ArticleDOI
TL;DR: In this article, a high-fidelity simulation technique was applied to the analysis of low-Reynolds-number transitional flows over moving and flexible canonical configurations motivated by small natural and man-made flyers.
Abstract: The present paper highlights results derived from the application of a high-fidelity simulation technique to the analysis of low-Reynolds-number transitional flows over moving and flexible canonical configurations motivated by small natural and man-made flyers. This effort addresses three separate fluid dynamic phenomena relevant to small fliers, including: laminar separation and transition over a stationary airfoil, transition effects on the dynamic stall vortex generated by a plunging airfoil, and the effect of flexibility on the flow structure above a membrane airfoil. The specific cases were also selected to permit comparison with available experimental measurements. First, the process of transition on a stationary SD7003 airfoil section over a range of Reynolds numbers and angles of attack is considered. Prior to stall, the flow exhibits a separated shear layer which rolls up into spanwise vortices. These vortices subsequently undergo spanwise instabilities, and ultimately breakdown into fine-scale turbulent structures as the boundary layer reattaches to the airfoil surface. In a timeaveraged sense, the flow displays a closed laminar separation bubble which moves upstream and contracts in size with increasing angle of attack for a fixed Reynolds number. For a fixed angle of attack, as the Reynolds number decreases, the laminar separation bubble grows in vertical extent producing a significant increase in drag. For the lowest Reynolds number considered \((Re_c = 10^4)\), transition does not occur over the airfoil at moderate angles of attack prior to stall. Next, the impact of a prescribed high-frequency small-amplitude plunging motion on the transitional flow over the SD7003 airfoil is investigated. The motioninduced high angle of attack results in unsteady separation in the leading edge and in the formation of dynamic-stalllike vortices which convect downstream close to the airfoil. At the lowest value of Reynolds number \((Re_c = 10^4)\), transition effects are observed to be minor and the dynamic stall vortex system remains fairly coherent. For \(Re_c = 4 \times 10^4\), the dynamic-stall vortex system is laminar at is inception, however shortly afterwards, it experiences an abrupt breakdown associated with the onset of spanwise instability effects. The computed phased-averaged structures for both values of Reynolds number are found to be in good agreement with the experimental data. Finally, the effect of structural compliance on the unsteady flow past a membrane airfoil is investigated. The membrane deformation results in mean camber and large fluctuations which improve aerodynamic performance. Larger values of lift and a delay in stall are achieved relative to a rigid airfoil configuration. For \(Re_c = 4.85 \times 10^4\), it is shown that correct prediction of the transitional process is critical to capturing the proper membrane structural response.

Journal ArticleDOI
TL;DR: The Brandenburg University of Technology at Cottbus is a newly commissioned research facility for the experimental study of sound generation from bodies immersed in a fluid flow as mentioned in this paper, where the design criteria for the open jet wind tunnel that provides a maximum wind speed of 72m/s at continuous operation and may be operated with nozzles of different dimension between 35cm diameter (circular nozzle) and 12 cm by 14.7 cm (rectangular nozzle).

Journal ArticleDOI
TL;DR: In this paper, a large-eddy simulation (LES) is used to predict wall-pressure fluctuations and noise of a low-speed flow past a thin cambered airfoil using an open-jet anechoic windtunnel at Ecole Centrale de Lyon.
Abstract: This paper discusses the prediction of wall-pressure fluctuations and noise of a low-speed flow past a thin cambered airfoil using large-eddy simulation (LES). The results are compared with experimental measurements made in an open-jet anechoic wind-tunnel at Ecole Centrale de Lyon. To account for the effect of the jet on airfoil loading, a Reynolds-averaged Navier-Stokes calculation is first conducted in the full wind-tunnel configuration, and the mean velocities from this calculation are used to define the boundary conditions for the LES in a smaller domain within the potential core of the jet. The LES flow field is characterized by an attached laminar boundary layer on the pressure side of the airfoil and a transitional and turbulent boundary layer on the suction side, in agreement with experimental observations. An analysis of the unsteady surface pressure field shows reasonable agreement with the experiment in terms of frequency spectra and spanwise coherence in the trailing-edge region. In the nose ...

Journal ArticleDOI
TL;DR: In this paper, the authors presented a theoretical model to predict the frequency of laminar boundary layer instability noise produced by an arbitrary aerofoil, which was validated against a number of well-known published experiments and also against the results of an experimental investigation.

Journal ArticleDOI
TL;DR: In this article, a numerical study of separation control has been made to investigate aerodynamic characteristics of NACA23012 airfoil with synthetic jets, and it was observed that the actual flow control mechanism and flow structure is fundamentally different depending on the range of synthetic jet frequency.

Journal ArticleDOI
TL;DR: In this article, a procedure to extract the aerodynamic loads and pressure distribution on an airfoil in the transonic flow regime from particle image velocimetry (PIV) measurements is presented.
Abstract: The present investigation assesses a procedure to extract the aerodynamic loads and pressure distribution on an airfoil in the transonic flow regime from particle image velocimetry (PIV) measurements. The wind tunnel model is a two-dimensional NACA-0012 airfoil, and the PIV velocity data are used to evaluate pressure fields, whereas lift and drag coefficients are inferred from the evaluation of momentum contour and wake integrals. The PIV-based results are compared to those derived from conventional loads determination procedures involving surface pressure transducers and a wake rake. The method applied in this investigation is an extension to the compressible flow regime of that considered by van Oudheusden et al (2006 Non-intrusive load characterization of an airfoil using PIV Exp. Fluids 40 988–92) at low speed conditions. The application of a high-speed imaging system allows the acquisition in relatively short time of a sufficient ensemble size to compute converged velocity statistics, further translated in turbulent fluctuations included in the pressure and loads calculation, notwithstanding their verified negligible influence in the computation. Measurements are performed at varying spatial resolution to optimize the loads determination in the wake region and around the airfoil, further allowing us to assess the influence of spatial resolution in the proposed procedure. Specific interest is given to the comparisons between the PIV-based method and the conventional procedures for determining the pressure coefficient on the surface, the drag and lift coefficients at different angles of attack. Results are presented for the experiments at a free-stream Mach number M = 0.6, with the angle of attack ranging from 0? to 8?.

Journal ArticleDOI
TL;DR: The present approach achieves sufficient accuracy at the immersed boundary and overcomes deficiencies in previous IB methods by introducing additional constraints - a compatibility for the interpolated velocity boundary condition related to mass conservation and the formal decoupling of the pressure on this surfaces.

Proceedings ArticleDOI
31 Dec 2009
TL;DR: The NACA 63 and 64 6-digit series of airfoils tested in the NACA LTPT in view to verify the RFOIL calculated airfoil characteristics for high Reynolds numbers was investigated in this article.
Abstract: This paper investigates the NACA 63 and 64 6-digit series of airfoils tested in the NACA LTPT in view to verify the RFOIL calculated airfoil characteristics for high Reynolds numbers. Some anomalies in the zero-lift angles of 15% and 18% thick airfoils from these series are identified, both in the airfoil clean case and in case of wrap-around roughness. It is found that RFOIL predicts the maximum lift coefficient at a Reynolds number of 3 million well, but consistently under predicts the Cl,max for Reynolds numbers of 6 and 9 million. It is, however, based on other comparisons at high Reynolds numbers unclear if this is due to an inability of the prediction code or to a deviation in the test results. The drag coefficient is under predicted with about 9% for a wide range of airfoils and Reynolds numbers. Due to wrap-around roughness the maximum lift coefficient decreases with 18% to 20%.

Patent
18 Mar 2009
TL;DR: In this paper, a turbine airfoil (12) is disclosed, which includes one of a turbine shroud (54), liner (52), vane or blade (50), including a film-cooling hole.
Abstract: A turbine airfoil (12) is disclosed. The airfoil (12) includes one of a turbine shroud (54), liner (52), vane or blade (50), including an airfoil sidewall (18) having a film-cooling hole that extends between an airfoil cooling circuit (24) and an airfoil surface (21). The airfoil (12) also includes an insert disposed in the film-cooling channel having a body (110). The body (110) has a proximal end (112) configured for disposition proximate the airfoil surface (21) and a distal end (114). The body (110) is also configured to define a passageway (116) that extends between the distal end (114) and proximal end (112) upon disposition in the film-cooling hole.

Journal ArticleDOI
TL;DR: In this article, the authors present the design and performance of an open jet, blow down wind tunnel that was newly commissioned in the anechoic chamber at the ISVR, University of Southampton, UK.

Journal ArticleDOI
TL;DR: In this paper, the authors developed an implicit solver for the harmonic balance equations, which is tested on two transonic test cases and evaluation is made against the unsteady simulation results.
Abstract: The computation of the aerodynamic forces arising from forced periodic motions is required for the generation of dynamic terms in models for flight simulation. The periodicity can be used to avoid using fully unsteady calculations by using the harmonic balance method. The current paper develops an implicit solver for the harmonic balance equations. The method is tested on two transonic test cases and evaluation is made against the unsteady simulation results. The first caseis for the pitching NACA 0012aerofoil. The second is for forced pitching of the F-5 wing with a wing tip launcher and missile. A reduction in computational time by one order of magnitude compared with the unsteady solver is obtained. Nomenclature A = matrix in frequency domain equation c = chord D = matrix in harmonic balance equation E = transformation matrix between frequency and time domains e = energy F, G, H = convective fluxes I = residual of semidiscrete system I = identity matrix k = reduced frequency nH = number of harmonics p = pressure R = residual vector T = period t = time u, v, w = Cartesian velocity components W = conserved variables � = angle of attack � t = pseudo time step

Journal ArticleDOI
TL;DR: In this article, a type of piezoceramic composite actuator known as Macro-Fiber Composite (MFC) is used for changing the camber of the wings.
Abstract: The purpose of the research presented here is to exploit actuation via smart materials to perform shape control of an aerofoil on a small aircraft and to determine the feasibility and advantages of smooth control surface deformations. A type of piezoceramic composite actuator known as Macro-Fiber Composite (MFC) is used for changing the camber of the wings. The MFC actuators were implemented on a 30° swept wing, 0·76m wingspan aircraft. The experimental vehicle was flown using two MFC patches in an elevator/aileron (elevon) configuration. Preliminary flight and wind-tunnel testing has demonstrated the stability and control of the concept. Flight tests were performed to quantify roll control using the MFC actuators. Lift and drag coefficients along with pitch and roll moment coefficients were measured in a low-speed, open-section wind tunnel. A vortex-lattice analysis complemented the database of aerodynamic derivatives used to analyse control response. The research, for the first time, successfully demonstrated that piezoceramic devices requiring high voltages can be effectively employed in small air vehicles without compromising the weight of the overall system.

Journal ArticleDOI
01 Mar 2009
TL;DR: In this paper, a multi-element aerofoil including NACA2415 and NACA22 leading-edge slat is experimentally and computationally investigated at a transitional Reynolds number of 2×105.
Abstract: In this study, a multi-element aerofoil including NACA2415 aerofoil with NACA22 leading-edge slat is experimentally and computationally investigated at a transitional Reynolds number of 2×105. In the experiment, the single-element aerofoil experiences a laminar separation bubble, and a maximum lift coefficient of 1.3 at a stall angle of attack of 12° is obtained. This flow has been numerically simulated by FLUENT, employing the recently developed, k—kL—ω and k—ω shear—stress transport (SST) transition models. Both transition models are shown to accurately predict the location of the experimentally determined separation bubble. Experimental measurements also illustrate that the leading-edge slat significantly delays the stall up to an angle of attack of 20°, with a maximum lift coefficient of 1.9. The fluid dynamics governing this improvement is the elimination of the separation bubble by the injection of high momentum fluid through the slat over the main aerofoil — an efficient means of flow contr...