Showing papers on "Airfoil published in 2008"

Parametric Study of an Oscillating Airfoil in a Power-Extraction Regime

Abstract: A wing that is heaving and pitching simultaneously may extract energy from an oncoming flow, thus acting as a turbine. The theoretical performance of such a concept is investigated here through unsteady two-dimensional laminar-flow simulations using the commercial finite volume computational fluid dynamics code FLUENT. Computations are performed in the heaving reference frame of the airfoil, thus leaving only the pitching motion of the airfoil to be dealt with through a rigid-body mesh rotation and a circular nonconformal sliding interface. Unsteady aerodynamics basics of the oscillating airfoil are first exposed, with a description of the operating regimes. Effects of unsteadiness are stressed and the inadequacy of a quasi-steady approach to take them into account is exposed. We present a mapping of power-extraction efficiency for a single oscillating airfoil in the frequency and pitching-amplitude domain: 0 55deg in which efficiencies are higher than 20%. Results from a parametric study are then provided and discussed. It is found that motion-related parameters such as heaving amplitude and frequency have the strongest effects on airfoil performances, whereas geometry and viscous parameters turn out to play a secondary role.

Direct numerical simulations of forced and unforced separation bubbles on an airfoil at incidence

Abstract: Direct numerical simulations (DNS) of laminar separation bubbles on a NACA-0012 airfoil at Re-c = 5 x 10(4) and incidence 5 degrees are presented. Initially volume forcing is introduced in order to promote transition to turbulence. After obtaining sufficient data from this forced case, the explicitly added disturbances are removed and the simulation run further. With no forcing the turbulence is observed to self-sustain, with increased turbulence intensity in the reattachment region. A comparison of the forced and unforced cases shows that the forcing improves the aerodynamic performance whilst requiring little energy input. Classical linear stability analysis is performed upon the time-averaged flow field; however no absolute instability is observed that could explain the presence of self-sustaining turbulence. Finally, a series of simplified DNS are presented that illustrate a three-dimensional absolute instability of the two-dimensional vortex shedding that occurs naturally. Three-dimensional perturbations are amplified in the braid region of developing vortices, and subsequently convected upstream by local regions of reverse flow, within which the upstream velocity magnitude greatly exceeds that of the time-average. The perturbations are convected into the braid region of the next developing vortex, where they are amplified further, hence the cycle repeats with increasing amplitude. The fact that this transition process is independent of upstream disturbances has implications for modelling separation bubbles.

Arrangement for controlling fluid jets injected into a fluid stream

Abstract: In an air mixing arrangement wherein a primary fluid is introduced through an opening in a wall to be mixed with a secondary fluid flowing along the wall surface, the opening is airfoil shaped with its leading edge being orientated at an attack angle with respect to the secondary fluid flow stream so as to thereby enhance the penetration and dispersion of the primary fluid stream into the secondary fluid stream. The airfoil shaped opening is selectively positioned such that its angle of attack provides the desired lift to optimize the mixing of the two streams for the particular application. In one embodiment, a collar is provided around the opening to prevent the secondary fluid from contacting the surface of the wall during certain conditions of operation. Multiple openings maybe used such as the combination of a larger airfoil shaped opening with a smaller airfoil shaped opened positioned downstream thereof, or a round shaped opening placed upstream of an airfoil shaped opening. Pairs of openings and associated collars maybe placed in symmetric relationship so as to promote mixing in particular applications, and nozzles maybe placed on the inner side of wall to enhance the flow characteristics of the primary fluid.

Concepts for morphing airfoil sections using bi-stable laminated composite structures

Abstract: The present paper investigates the potential of using bi-stable laminated composite structures for morphing an airfoil section. The objective of the paper is to identify geometries and lay-ups of candidate configurations that offer multiple stable shapes for the airfoil section. Carbon-fiber laminated composites with non-symmetric laminate configurations are used for morphing the airfoil section. Thermal curing is used to induce residual stresses into the structure in order to achieve bi-stability. Three concepts that focus on morphing a flap-like structure and the camber and chord of an airfoil section are proposed. Several geometries and laminate configurations are investigated using finite element nonlinear static analysis. The magnitude of loads required to actuate the airfoil section between the stable shapes is evaluated. The impact of manufacturability on producing viable morphing mechanisms within the airfoil section is also discussed.

CFD‐based optimization of aerofoils using radial basis functions for domain element parameterization and mesh deformation

Abstract: A novel domain element shape parameterization method is presented for computational fluid dynamics-based shape optimization. The method is to achieve two aims: (1) provide a generic 'wrap-around' optimization tool that is independent of both flow solver and grid generation package and (2) provide a method that allows high-fidelity aerodynamic optimization of two- and three-dimensional bodies with a low number of design variables. The parameterization technique uses radial basis functions to transfer domain element movements into deformations of the design surface and corresponding aerodynamic mesh, thus allowing total independence from the grid generation package (structured or unstructured). Independence from the flow solver (either inviscid, viscous, aeroelastic) is achieved by obtaining sensitivity information for an advanced gradient-based optimizer (feasible sequential quadratic programming) by finite-differences. Results are presented for two-dimensional aerofoil inverse design and drag optimization problems. Inverse design results demonstrate that a large proportion of the design space is feasible with a relatively low number of design variables using the domain element parameterization. Heavily constrained (in lift, volume, and moment) two-dimensional aerofoil drag optimization has shown that significant improvements over existing designs can be achieved using this method, through the use of various objective functions.

Two-Dimensional Low-Reynolds Number Wind Tunnel Results for Airfoil NACA 0018:

Abstract: The two-dimensional characteristics of airfoil NACA 0018 have been measured for Reynolds numbers between 0.15x106 and 1.0x106 to establish the lift, drag and moment curves that serve as input to performance calculations of vertical axis wind turbines. At the lower surface laminar separation occurs at low to medium angles of attack, which is of significant influence on the characteristics and the radiated noise. For the situation with a lower surface laminar separation bubble, span wise wake rake traverse measurements showed an irregular three-dimensional pattern. Noise reduction could be achieved with zigzag tape at the 70% to 80% lower surface chord station. Significant post-stall hysteresis loops occurred showing a high loss in lift.

Scaling Effects of an Aerodynamic Plasma Actuator

Abstract: We present experimental results to yield insight into the scalability and control effectiveness of single-dielectricbarrier-discharge plasma actuators for leading-edge separation control on airfoils. The parameters investigated are chord Reynolds number, Mach number, leading-edge radius, actuator amplitude, and unsteady frequency. This includes chord Reynolds numbers up to 1:0 � 106 and a maximum freestream speed of 60 m=s corresponding to a Mach number of 0.176. The main objective of this work is to examine the voltage requirements for the plasma actuators to reattach the flow at the leading edge of airfoils at poststall angles of attack for a range of flow parameters in order to establish scaling between laboratory and full-flight conditions. For the full range of conditions, an optimum unsteady actuator frequency f is found to minimize the actuator voltage needed to reattach the flow, such that F� � fLsep=U1 � 1. At the optimum frequencies, the minimum voltage required to reattach the flow is weakly dependent on chord Reynolds number and strongly dependent on the poststall angle of attack and leading-edge radius. The results indicate that the voltage required to reattach the flow scales as the square of the leading-edge radius.

A computational study of the aerodynamic performance of a dragonfly wing section in gliding flight.

Abstract: A comprehensive computational fluid-dynamics-based study of a pleated wing section based on the wing of Aeshna cyanea has been performed at ultra-low Reynolds numbers corresponding to the gliding flight of these dragonflies. In addition to the pleated wing, simulations have also been carried out for its smoothed counterpart (called the 'profiled' airfoil) and a flat plate in order to better understand the aerodynamic performance of the pleated wing. The simulations employ a sharp interface Cartesian-grid-based immersed boundary method, and a detailed critical assessment of the computed results was performed giving a high measure of confidence in the fidelity of the current simulations. The simulations demonstrate that the pleated airfoil produces comparable and at times higher lift than the profiled airfoil, with a drag comparable to that of its profiled counterpart. The higher lift and moderate drag associated with the pleated airfoil lead to an aerodynamic performance that is at least equivalent to and sometimes better than the profiled airfoil. The primary cause for the reduction in the overall drag of the pleated airfoil is the negative shear drag produced by the recirculation zones which form within the pleats. The current numerical simulations therefore clearly demonstrate that the pleated wing is an ingenious design of nature, which at times surpasses the aerodynamic performance of a more conventional smooth airfoil as well as that of a flat plate. For this reason, the pleated airfoil is an excellent candidate for a fixed wing micro-aerial vehicle design.

Dielectric Barrier Discharge Flow Control at Very Low Flight Reynolds Numbers

Abstract: Experiments were performed on a flat-plate airfoil and an Eppler E338 airfoil at very low flight Reynolds numbers (3000 ≤ Re ≤ 50,000), in which dielectric barrier discharge plasma actuators were employed at the airfoil leading edges to effect flow control. The actuators were driven in a high-frequency (kilohertz) steady mode and a pulsed mode in which pulse frequency and duty cycle were varied in a systematic fashion. Optimum reduced frequencies for generating poststall lift were approximately between 0.4 and 1, and this was broadly consistent with zero-mass-flux slot-blowing data acquired at Reynolds numbers that were approximately 200 times higher. Nevertheless, profound differences in the response to reduced frequency and duty cycle were observed between the flat-plate and E338 airfoils. In general, actuation produced considerable performance improvements, including an increase in maximum lift coefficient of 0.4 to 0.8 and maintained elevated endurance at significantly higher lift coefficients. Actuation in the steady mode resulted in circulation control at Re = 3000. Pulsed actuation also exerted a significant effect on the wake at prestall angles of attack, in which control of the upper-surface flat-plate bubble shedding produced significant differences in wake spreading and vortex shedding. The flat plate was also tested in a semispan-wing configuration (AR = 6), and the effect of control was comparable with that observed on the airfoil.

Skin design studies for variable camber morphing airfoils

Abstract: This paper identifies the desirable attributes of a flexible skin of a morphing wing. The study is conducted using airfoil camber morphing as an example. The ideal flex-skin would be highly anisotropic, having a low in-plane axial stiffness but a high out-of-plane flexural stiffness. Reduced skin axial stiffness allows morphing at low actuation cost. However, for some substructure and actuation designs, a lower limit on the skin's in-plane axial stiffness may be required to prevent unacceptable global camber deformation under aerodynamic loads. High flexural stiffness prevents local deformation of skin sections between supports due to aerodynamic pressure loads, and avoids buckling of skin sections under compression as the airfoil cambers under actuation force. For the camber morphing application the strain levels in the flex-skin are not expected to exceed around 2%. If the axial stiffness of the flex-skin is reduced significantly, it may be necessary to consider aerodynamic stiffness (negligible vis-a-vis structural stiffness for classical airfoils) to accurately calculate deformation under loading. The approach followed in the study can be used to identify specifications for the skin and then reverse engineer and design highly anisotropic composite skins that meet the specifications.

A Modified Dynamic Stall Model for Low Mach Numbers

Abstract: The Leishman–Beddoes dynamic stall model is a popular model that has been widely applied in both helicopter and wind turbine aerodynamics. This model has been specially refined and tuned for helicopter applications, where the Mach number is usually above 0.3. However, experimental results and analyses at the University of Glasgow have suggested that the original Leishman–Beddoes model reconstructs the unsteady airloads at low Mach numbers less well than at higher Mach numbers. This is particularly so for stall onset and the return from the fully stalled state. In this paper, a modified dynamic stall model that adapts the Leishman–Beddoes dynamic stall model for lower Mach numbers is proposed. The main modifications include a new stall-onset indication, a new return modeling from stalled state, a revised chordwise force, and dynamic vortex modeling. The comparisons to the Glasgow University dynamic stall database showed that the modified model is capable of giving improved reconstructions of unsteady aerofoil data in low Mach numbers.

An Experimental Study of the Laminar Flow Separation on a Low-Reynolds-Number Airfoil

Abstract: An experimental study was conducted to characterize the transient behavior of laminar flow separation on a NASA low-speed GA (W)-1 airfoil at the chord Reynolds number of 70,000. In addition to measuring the surface pressure distribution around the airfoil, a high-resolution particle image velocimetry (PIV) system was used to make detailed flow field measurements to quantify the evolution of unsteady flow structures around the airfoil at various angles of attack (AOAs). The surface pressure and PIV measurements clearly revealed that the laminar boundary layer would separate from the airfoil surface, as the adverse pressure gradient over the airfoil upper surface became severe at AOA ≥8.0 deg. The separated laminar boundary layer was found to rapidly transit to turbulence by generating unsteady Kelvin-Helmholtz vortex structures. After turbulence transition, the separated boundary layer was found to reattach to the airfoil surface as a turbulent boundary layer when the adverse pressure gradient was adequate at AOA 12.0 deg.

Time-resolved and volumetric PIV measurements of a transitional separation bubble on an SD7003 airfoil

Abstract: To comprehensively understand the effects of Kelvin–Helmholtz instabilities on a transitional separation bubble on the suction side of an airfoil regarding as to flapping of the bubble and its impact on the airfoil performance, the temporal and spatial structure of the vortices occurring at the downstream end of the separation bubble is investigated. Since the bubble variation leads to a change of the pressure distribution, the investigation of the instantaneous velocity field is essential to understand the details of the overall airfoil performance. This vortex formation in the reattachment region on the upper surface of an SD7003 airfoil is analyzed in detail at different angles of attack. At a Reynolds number Re c 4°. Due to transition processes, turbulent reattachment of the separated shear layer occurs enclosing a locally confined recirculation region. To identify the location of the separation bubble and to describe the dynamics of the reattachment, a time-resolved PIV measurement in a single light-sheet is performed. To elucidate the spatial structure of the flow patterns in the reattachment region in time and space, a stereo scanning PIV set-up is applied. The flow field is recorded in at least ten successive light-sheet planes with two high-speed cameras enclosing a viewing angle of 65° to detect all three velocity components within a light-sheet leading to a time-resolved volumetric measurement due to a high scanning speed. The measurements evidence the development of quasi-periodic vortex structures. The temporal dynamics of the vortex roll-up, initialized by the Kelvin–Helmholtz (KH) instability, is shown as well as the spatial development of the vortex roll-up process. Based on these measurements a model for the evolving vortex structure consisting of the formation of c-shape vortices and their transformation into screwdriver vortices is introduced.

Dynamics of Airfoil Separation Control using Zero-Net Mass-Flux Forcing

Abstract: Zero-net mass-flux jet based control of flow separation over a stalled airfoil is examined using numerical simulations. Two-dimensional simulations are carried out for a NACA 4418 airfoil at a chord Reynolds number of 40,000 and angle of attack of 18 deg. Results for the uncontrolled flow indicate the presence of three distinct natural time scales in the flow corresponding to the shear layer, separation bubble, and wake regions. The natural frequencies are used to select appropriate forcing frequencies, and it is found that forcing frequencies closer to the separation bubble frequency elicit the best response in terms of reduction of separation extent and an improvement in aerodynamic performance. In contrast, higher forcing frequencies closer to the natural shear layer frequency tend to enhance separation. The vortex dynamics and frequency response of flow are examined in detail to gain insight into mechanisms underlying the observed behavior.

Flexible-Membrane Airfoils at Low Reynolds Numbers

Abstract: An experimental study was conducted to assess the benefits of using flexible-membrane airfoils/wings at low Reynolds numbers for micro air vehicle applications compared with using a conventional rigid airfoil/wing. In addition to measuring aerodynamic forces acting on flexible-membrane airfoils/wings, a high-resolution particle image velocimetry system was used to conduct flowfield measurements to quantify the transient behavior of vortex and turbulent flow structures around the flexible-membrane airfoils/wings to elucidate the associated underlying fundamental physics. The aerodynamic force measurements revealed that flexible-membrane airfoils could provide better aerodynamic performance compared with their rigid counterpart at low Reynolds numbers. The flexibility (or rigidity) of the membrane skins of the airfoils was found to greatly affect their aerodynamic performance. Particle image velocimetry measurements elucidated that flexible-membrane airfoils could change their camber (i.e., crosssectional shape) automatically to adapt incoming flows to balance the pressure differences on the upper and lower surfaces of the airfoils, therefore suppressing flow separation on the airfoil upper surfaces. Meanwhile, deformation of the flexible-membrane skins was found to cause significant airfoil trailing-edge deflection (i.e., lift the airfoil trailing edge up from its original designed position), which resulted in a reduction of the effective angles of attack of the flexible-membrane airfoils, thereby delaying airfoil stall at high angles of attack. The nonuniform spanwise deformation of the flexible-membrane skins of the flexible-membrane airfoils was found to significantly affect the characteristics of vortex and turbulent flow structures around the flexible-membrane airfoils.

Direct Noise Computation of the Turbulent Flow Around a Zero-Incidence Airfoil

Abstract: A large eddy simulation of the flow around a NACA 0012 airfoil at zero incidence is performed at a chord-based Reynoldsnumber of500,000 anda Machnumberof 0.22.Theaim istoshow thathigh-order numericalschemes can successfully be used to perform direct acoustic computations of compressible transitional flow on curvilinear grids. AtaReynoldsnumberof500,000,theboundarylayersaroundtheairfoiltransition fromaninitially laminarstateto a turbulent state before reaching the trailing edge. Results obtained in the large eddy simulation show a well-placed transition zone and turbulence levels in the boundary layers that are in agreement with experimental data. Furthermore, the radiated acoustic field is determined directly by the large eddy simulation, without the use of an acoustic analogy. Third-octave acoustic spectra are compared favorably with experimental data.

Scanning PIV investigation of the laminar separation bubble on a SD7003 airfoil

Abstract: A laminar separation bubble occurs on the suction side of the SD7003 airfoil at an angle of attack α = 4–8° and a low Reynolds number less than 100,000, which brings about a significant adverse aerodynamic effect. The spatial and temporal structure of the laminar separation bubble was studied using the scanning PIV method at α = 4° and Re = 60,000 and 20,000. Of particular interest are the dynamic vortex behavior in transition process and the subsequent vortex evolution in the turbulent boundary layer. The flow was continuously sampled in a stack of parallel illuminated planes from two orthogonal views with a frequency of hundreds Hz, and PIV cross-correlation was performed to obtain the 2D velocity field in each plane. Results of both the single-sliced and the volumetric presentations of the laminar separation bubble reveal vortex shedding in transition near the reattachment region at Re = 60,000. In a relatively long distance vortices characterized by paired wall-normal vorticity packets retain their identities in the reattached turbulent boundary layer, though vortices interact through tearing, stretching and tilting. Compared with the restricted LSB at Re = 60,000, the flow at Re = 20,000 presents an earlier separation and a significantly increased reversed flow region followed by “huge” vortical structures.

ACOUSTICS2008/3539 LES prediction of wall-pressure fluctuations and noise of a low-speed airfoil

Abstract: The wall-pressure fluctuations and noise of a low-speed airfoil are computed 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 eect of the jet on airfoil loading, a RANS calculation is conducted in the full wind-tunnel configuration, which provides velocity boundary conditions for the LES in a smaller domain within the potential core of the jet. The flow field is characterized by an attached laminar boundary layer on the pressure side and a transitional and turbulent boundary layer on the suction side. The predicted unsteady surface pressure field shows reasonable agreement with the experimental data in terms of frequency spectra and coherence in the trailing-edge region. In the nose region, characterized by unsteady separation and transition to turbulence, the wall-pressure fluctuations are highly sensitive to small perturbations and dicult to predict or measure with certainty. The LES, in combination with the Ffowcs Williams and Hall solution to the Lighthill equation, also predicts well the radiated trailing-edge noise. A finite-chord correction is derived and applied to the noise prediction, which is shown to improve the overall agreement with the experimental sound spectra.

Discrete conservation principles in large-eddy simulation with application to separation control over an airfoil

Abstract: An unstructured-grid large-eddy simulation (LES) technique is used to investigate the turbulent flow separation over an airfoil with and without synthetic-jet control. Numerical accuracy and stability on arbitrary shaped mesh elements at high Reynolds numbers are achieved using a finite-volume discretization of the incompressible Navier–Stokes equations based on higher-order conservation principles—i.e., in addition to mass and momentum conservation, kinetic energy conservation in the inviscid limit is used to guide the selection of the discrete operators and solution algorithm. Two different stall configurations, which consist of flow over a NACA 0015 airfoil at 16.6° and 20° angles of attack, are simulated at Reynolds number of 896 000 based on the airfoil chord length and freestream velocity. In the case of 16.6° angle of attack where flow separates around a midchord location, LES results show excellent agreement with the experimental data for both uncontrolled and controlled cases. LES confirms the ex...