Showing papers on "Airfoil published in 2010"

Measurement of the noise generation at the trailing edge of porous airfoils

Abstract: Owls are commonly known for their quiet flight, enabled by three adaptions of their wings and plumage: leading edge serrations, trailing edge fringes and a soft and elastic downy upper surface of the feathers. In order to gain a better understanding of the aeroacoustic effects of the third property that is equivalent to an increased permeability of the plumage to air, an experimental survey on a set of airfoils made of different porous materials was carried out. Several airfoils with the same shape and size but made of different porous materials characterized by their flow resistivities and one non-porous reference airfoil were subject to the flow in an aeroacoustic open jet wind tunnel. The flow speed has been varied between approximately 25 and 50 m/s. The geometric angle of attack ranged from −16° to 20° in 4°-steps. The results of the aeroacoustic measurements, made with a 56-microphone array positioned out of flow, and of the measurements of lift and drag are given and discussed.

Structural and Aerodynamic Models in Nonlinear Flight Dynamics of Very Flexible Aircraft

Abstract: An evaluation of computational models is carried out for flight dynamics simulations on low-speed aircraft with very-flexible high-aspect ratio wings. Structural dynamic models include displacement-based, strain-based, and intrinsic (first-order) geometrically-nonlinear composite beams, while thin-strip and vortex lattice methods are considered for the unsteady aerodynamics. It is first shown that all different beam finite element models (previously derived in the literature from different assumptions) can be consistently obtained from a single set of equations. This approach has been used to expand existing strain-based models to include shear effects. Comparisons are made in terms of numerical efficiency and simplicity of integration in flexible aircraft flight dynamics studies. On the structural modeling, it was found that intrinsic solutions can be several times faster than conventional ones for aircraft-type geometries. For the aerodynamic modeling, thin-strip models based on indicial airfoil response are found to perform well in situations dominated by small amplitude dynamics around large quasi-static wing deflections, while large-amplitude wing dynamics require three-dimensional descriptions (e.g. vortex lattice).

Stability and receptivity characteristics of a laminar separation bubble on an aerofoil

Abstract: Stability characteristics of aerofoil flows are investigated by linear stability analysis of time-averaged velocity profiles and by direct numerical simulations with time-dependent forcing terms. First the wake behind an aerofoil is investigated, illustrating the feasibility of detecting absolute instability using these methods. The time-averaged flow around an NACA-0012 aerofoil at incidence is then investigated in terms of its response to very low-amplitude hydrodynamic and acoustic perturbations. Flow fields obtained from both two- and three-dimensional simulations are investigated, for which the aerofoil flow exhibits a laminar separation bubble. Convective stability characteristics are documented, and the separation bubble is found to exhibit no absolute instability in the classical sense; i.e. no growing disturbances with zero group velocity are observed. The flow is however found to be globally unstable via an acoustic-feedback loop involving the aerofoil trailing edge as a source of acoustic excitation and the aerofoil leading-edge region as a site of receptivity. Evidence suggests that the feedback loop may play an important role in frequency selection of the vortex shedding that occurs in two dimensions. Further simulations are presented to investigate the receptivity process by which acoustic waves generate hydrodynamic instabilities within the aerofoil boundary layer. The dependency of the receptivity process to both frequency and source location is quantified. It is found that the amplitude of trailing-edge noise in the fully developed simulation is sufficient to promote transition via leading-edge receptivity.

Reduced-Order Nonlinear Unsteady Aerodynamic Modeling Using a Surrogate-Based Recurrence Framework

Abstract: A reduced-order nonlinear unsteady aerodynamic modeling approach suitable for analyzing pitching/plunging airfoils subject to fixed or time-varying freestream Mach numbers is described. The reduced-order model uses kriging surrogates to account for flow nonlinearities and recurrence solutions to account for time-history effects associated with unsteadiness. The resulting surrogate-based recurrence framework generates time-domain predictionsofunsteadylift,moment,anddragthataccuratelyapproximate computational fluiddynamicssolutions, but at a fraction of the computational cost. Results corresponding to transonic conditions demonstrate that the surrogate-based recurrence framework can mimic computational fluid dynamics predictions of unsteady aerodynamic responses when flow nonlinearities are present. For an unsteady aerodynamic modeling problem considered in this study, an accurate reduced-order model was generated by the surrogate-based recurrence framework approach with significantly fewer computational fluid dynamics evaluations compared to results reported in the literature for a similar problem in which a proper-orthogonal-decomposition-based approach was applied. Furthermore, the results show that the surrogate-based approach can accurately model time-varying freestream Mach number effects and is therefore applicable to rotary-wing applications in addition to fixed-wing applications.

High-lift airfoil trailing edge separation control using a single dielectric barrier discharge plasma actuator

Abstract: Control of flow separation from the deflected flap of a high-lift airfoil up to Reynolds numbers of 240,000 (15 m/s) is explored using a single dielectric barrier discharge (DBD) plasma actuator near the flap shoulder. Results show that the plasma discharge can increase or reduce the size of the time-averaged separated region over the flap depending on the frequency of actuation. High-frequency actuation, referred to here as quasi-steady forcing, slightly delays separation while lengthening and flattening the separated region without drastically increasing the measured lift. The actuator is found to be most effective for increasing lift when operated in an unsteady fashion at the natural oscillation frequency of the trailing edge flow field. Results indicate that the primary control mechanism in this configuration is an enhancement of the natural vortex shedding that promotes further momentum transfer between the freestream and separated region. Based on these results, different modulation waveforms for creating unsteady DBD plasma-induced flows are investigated in an effort to improve control authority. Subsequent measurements show that modulation using duty cycles of 50–70% generates stronger velocity perturbations than sinusoidal modulation in quiescent conditions at the expense of an increased power requirement. Investigation of these modulation waveforms for trailing edge separation control similarly shows that additional increases in lift can be obtained. The dependence of these results on the actuator carrier and modulation frequencies is discussed in detail.

Novel, Bidirectional, Variable-Camber Airfoil via Macro-Fiber Composite Actuators

Abstract: This study aims to enable solid-state aerodynamic force generation in high-dynamic-pressure airflow. A novel, high-load-output, bidirectional variable-camber airfoil employing a type of piezoceramic composite actuator known as a Macro-Fiber Composite is presented. The novel airfoil employs two active surfaces and a single four-bar (box) mechanism as the internal structure. The unique choice of boundary conditions allows variable and smooth deformation in both directions from a flat camber line. The paper focuses on actuation modeling and response characterization under aerodynamic loads. A parametric study of aerodynamic response is employed to optimize the kinematic parameters of the airfoil. The concept is fabricated by implementing eight Macro-Fiber Composite 8557-P1-type actuators in a bimorph configuration to construct the active surfaces. The box mechanism generates deflection and camber change as predicted. Wind-tunnel experiments are conducted on a 12.6% maximum thickness, 127 mm chord airfoil. Aerodynamic and structural performance results are presented for a flow rate of 15 m /s and a Reynolds number of 127,000. Nonlinear effects due to aerodynamic and piezoceramic hysteresis are identified and discussed. A lift coefficient change of 1.54 is observed purely due to voltage actuation. Results are compared with conventional, zero-camber NACA and other airfoils. A 72% increase in the lift-curve slope is achieved when compared with a NACA 0009 airfoil.

The ultra-low Reynolds number airfoil wake

Abstract: Lift force and the near wake of an NACA 0012 airfoil were measured over the angle (α) of attack of 0°–90° and the chord Reynolds number (Re c ), 5.3 × 103–5.1 × 104, with a view to understand thoroughly the near wake of the airfoil at low- to ultra-low Re c . While the lift force is measured using a load cell, the detailed flow structure is captured using laser-Doppler anemometry, particle image velocimetry, and laser-induced fluorescence flow visualization. It has been found that the stall of an airfoil, characterized by a drop in the lift force, occurs at Re c ≥ 1.05 × 104 but is absent at Re c = 5.3 × 103. The observation is connected to the presence of the separation bubble at high Re c but absence of the bubble at ultra-low Re c , as evidenced in our wake measurements. The near-wake characteristics are examined and discussed in detail, including the vortex formation length, wake width, spanwise vorticity, wake bubble size, wavelength of K–H vortices, Strouhal numbers, and their dependence on α and Re c .

Approximate Modeling of Unsteady Aerodynamics for Hypersonic Aeroelasticity

Abstract: DOI: 10.2514/1.C000190 Various approximations to unsteady aerodynamics are examined for the aeroelastic analysis of a thin doublewedge airfoil in hypersonic flow. Flutter boundaries are obtained using classical hypersonic unsteady aerodynamic theories: piston theory, Van Dyke’s second-order theory, Newtonian impact theory, and unsteady shock-expansion theory. The theories are evaluated by comparing the flutter boundaries with those predicted using computational fluid dynamics solutions to the unsteady Navier–Stokes equations. Inaddition, several alternative approaches to the classical approximations are also evaluated: two different viscous approximations based on effective shapes and combined approximate computational approaches that use steady-state computational-fluid-dynamics-based surrogatemodelsinconjunction withpistontheory.Theresultsindicatethat,with theexceptionof first-order piston theory and Newtonian impact theory, the approximate theories yield predictions between 3 and 17% of normalized root-mean-square error and between 7 and 40% of normalized maximum error of the unsteady Navier–Stokes predictions. Furthermore, the demonstrated accuracy of the combined steady-state computational fluid dynamics and piston theory approaches suggest that important nonlinearities in hypersonic flow are primarily due to steadystate effects. This implies that steady-state flow analysis may be an alternative to time-accurate Navier–Stokes solutions for capturing complex flow effects.

High-Fidelity Simulations of Moving and Flexible Airfoils at Low Reynolds Numbers (Postprint)

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 time-averaged 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(sub c) = 10(exp 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 motion-induced high angle of attack results in unsteady separation in the leading edge and in the formation of dynamic-stall-like vortices which convect downstream close to the airfoil.

Deformable trailing edge flaps for modern megawatt wind turbine controllers using strain gauge sensors

Abstract: The present work contains a deformable trailing edge flap controller integrated in a numerically simulated modern, variablespeed, pitch-regulated megawatt (MW)-size wind turbine. The aeroservoelastic multi-body code HAWC2 acts as a component in the control loop design. At the core of the proposed controller, all unsteady loads are divided by frequency content. Blade pitching and generator moment react to low-frequency excitations, whereas flaps deal with high-frequency excitations. The present work should be regarded as an investigation into the fatigue load reduction potential when applying trailing edge flaps on a wind turbine blade rather than a conclusive control design with traditional issues like stability and robustness fully investigated. Recent works have shown that the fatigue load reduction by use of trailing edge flaps may be greater than for traditional pitch control methods. By enabling the trailing edge to move independently and quickly along the spanwise position of the blade, local small flutuations in the aerodynamic forces can be alleviated by deformation of the airfoil flap. Strain gauges are used as input for the flap controller, and the effect of placing strain gauges at various radial positions on the blade is investigated. An optimization routine minimizes blade root fatigue loads. Calculations are based on the 5 MW reference wind turbine part of the UpWind project primarily with a mean turbulent wind speed close to rated power. A fatigue load reduction of 25% in the blade root moment was obtained for a continuous 6.3 m long flap. Copyright © 2009 John Wiley & Sons, Ltd.

Characteristics of Pitching and Plunging Airfoils Under Dynamic-Stall Conditions

Abstract: An experimental investigation into the dynamic-stall process of a pitching and plunging airfoil at low Reynolds numbers has been carried out using direct force measurements and smoke visualization in an Eiffel-type wind tunnel. The strong influence of reduced frequency (k = �fc/U∞) on the vortical wake of both pure-plunging and pure-pitching airfoils is revealed. Here a transition from a bluff-body to a mushroom-type wake has been observed at approximately k = 0.2. Some associated lift and moment hysteresis curves for combined pitching and plunging kinematics are then presented with an accompanying discussion on the nature of the dynamic-stall process. For these complex kinematics it is observed that both lift and moment phase lags grow with reduced frequency from k = 0.05 to k = 0.1. Despite substantial lift augmentation in the light- and deep-stall regimes, strong pitching-down moments are not avoided.

Broadband Noise Reduction With Trailing Edge Brushes

Abstract: Airfoil broadband trailing edge noise is reduced by modification of the trailing edge geometry. A brush made of a single row of flexible polypropylene fibers is integrated in the trailing edge of a cambered airfoil. Far field acoustic measurements show a noise reduction potential reaching 3 dB on a wide frequency range. Due to high curvature of the incident flow, a secondary acoustic source partly masks the trailing edge noise reduction. Hot wire correlation measurements in the very near wake of the airfoil show that longitudinal as well as transversal length scales are affected by the brush. Span wise coherence length of boundary layer eddies falls off by 25 % in the presence of a brush in the adequate frequency range, possibly explaining a 1.3 dB contribution to the noise reduction mechanism. Boundary layer turbulence exhibits a preferred coherence length l y v on a wide frequency range. l y v /d ≈ 2, is considered a proper brush design law, d being the diameter of the brush.

Experimental Investigation of Jet Mixing Mechanism of Co-Flow Jet Airfoil

Abstract: The jet mixing of a co-flow jet (CFJ) airfoil is investigated to understand the mechanism of lift enhancement, drag reduction, and stall margin increase. Digital Particle Image Velocimetry, flow visualization and aerodynamic forces measurements are used to reveal the insight of the CFJ airfoil mixing process. At low AoA and low momentum coefficient, the mixing between the wall jet and mainflow is dominant with large structure coherent structures for the attached flows. When the momentum coefficient is increased, the large vortex structure disappears. At high AoA with flow separation, the CFJ creates a upstream flow strip between two counter rotating vertical shear layer, i.e., the outer shear layer and inner flow induced by CFJ. The UFS is characterized with large vortex free region. The co-flow wall jet is deflected normal to the airfoil surface characterized with a saddle point. With increased momentum coefficient of the CFJ, the saddle point moves downstream and eventually disappears when the flow is attached. Turbulence plays a key role in mixing the CFJ with mainflow to transport high kinetic energy from the jet to mainflow so that the mainflow can remain attached at high AoA to generate high lift. When the flow is separated, increased CFJ momentum coefficient also increases the turbulence intensity at jet injection mixing region.

High-lift airfoil separation with dielectric barrier discharge plasma actuation

Abstract: The efficacy of a single dielectric barrier discharge plasma actuator for controlling turbulent boundary-layer separation from the deflected flap of a high-lift airfoil is investigated between Reynolds numbers of 240,000 (15 m/s) and 750,000 (45 m/s). Momentum coefficients for the dielectric barrier discharge plasma actuator are approximately an order of magnitude lower than those usually employed for such studies, yet control authority is still realized through amplification of natural vortex shedding from the flap shoulder, which promotes momentum transfer between the freestream and separated region. This increases dynamic loading on the flap and further organizes turbulent fluctuations in the wake. The measured lift enhancement is primarily due to upstream effects from increased circulation around the entire model, rather than full reattachment to the deflected flap surface. Lift enhancement via instability amplification is found to be relatively insensitive to changes in angle of attack, provided that the separation location and underlying dynamics do not change. The modulation waveform used to excite low-frequency perturbations with a high-frequency plasma-carrier signal has a considerable effect on the actuator performance. Control authority decreases with increasing Reynolds number and flap deflection, highlighting the necessity for further improvement of plasma actuators for use in realistic takeoff and landing transport aircraft applications. These findings are compared to studies on a similar high-lift platform using piezoelectric-driven zero-net-mass flux actuation.

Parametric Study of Sweeping Jet Actuators for Separation Control

Abstract: In this paper, studies of separation control over a generic Multiple Flap Airfoil (MFA) using sweeping jet actuator arrays are presented. These jets, exiting from millimeter-scale nozzles, oscillate from side-to-side in a sweeping manner similar to windshield wipers and are henceforth referred to as sweeping jet actuators. Different flap sizes and flap deflections up to 45 o were investigated in the experiments. The MFA, with integrated rows of sweeping jet actuators at several chordwise locations on the flaps, enabled an extensive variation of geometrical and fluid dynamical parameters to study separation control. The effect of flap size, actuator location and actuation parameters on the lift and drag coefficients of the airfoil are discussed.

Design Optimization Utilizing Gradient/Hessian Enhanced Surrogate Model

Abstract: In this paper, gradient/Hessian-enhanced surrogate models have been developed based on Kriging approaches. The gradient/Hessian-enhanced Kriging methods have been developed based on direct and indirect formulations. The efficiencies of these methods are compared by analytical function fitting, aerodynamic data modeling and 2D airfoil drag minimization problems. For the aerodynamic problems, efficient CFD gradient/Hessian calculation methods are utilized that make use of adjoint and automatic differentiation techniques. The gradient/Hessian-enhanced surrogate models are shown to be useful in the development of efficient design optimization, aerodynamic database construction, and uncertainty analysis.

Porous airfoils: noise reduction and boundary layer effects

Abstract: permeable) materials. The objective of the research is the analysis of the turbulent boundary layer properties of porous airfoils and, subsequently, of the noise generated at the trailing edge. The inuence of the porous materials, characterized by their air ow resistivity, is discussed. The acoustic measurements were performed using a planar 56{channel microphone array and the boundary layer properties were measured using constant temperature anemometry. The recorded acoustic data underwent further processing by application of an advanced beamforming algorithm. A noticeable reduction of the emitted trailing edge noise was measured for the porous airfoils over a large range of frequencies. At high frequencies, some of the porous airfoils were found to generate more noise than the reference airfoil which might be due to the surface roughness noise contribution. It is found that the turbulent boundary layer thickness and the boundary layer displacement thickness of the airfoils increase with decreasing ow resistivities for both suction and pressure side. Both boundary layer thickness and displacement thickness of the non{porous airfoil are below those of the porous airfoils.

Real Time Morphing Wing Optimization Validation Using Wind-Tunnel Tests

Abstract: In this paper, wind-tunnel results of a real time optimization of a morphing wing in the wind tunnel for delaying the transition toward the trailing edge are presented. A morphing rectangular finite aspect ratio wing, having a wind tunnel experimental airfoil reference airfoil cross section, was considered, with its upper surface made of a flexible composite material and instrumented with Kulite pressure sensors and two smart memory alloys actuators. Several wind-tunnel test runs for various Mach numbers, angles of attack, and Reynolds numbers were performed in the 6' x 9' wind tunnel at the Institute for Aerospace Research at the National Research Council Canada. Unsteady pressure signals were recorded and used as feedback in real time control while the morphing wing was requested to reproduce various optimized airfoils by changing automatically the two actuators' strokes. This paper shows the optimization method implemented into the control software code that allows the morphing wing to adjust its shape to an optimum configuration under the wind-tunnel airflow conditions.

Data mining of Pareto-optimal transonic airfoil shapes using proper orthogonal decomposition

Abstract: A new approach to extract useful design information from Pareto-optimal solutions of optimization problems is proposed and applied to an aerodynamic transonic airfoil shape optimization. The proposed approach enables an analysis of line, face, or volume data of all Pareto-optimal solutions such as shape and flow field by decomposing the data into principal modes and corresponding base vectors using proper orthogonal decomposition (POD). Analysis of the shape and surface pressure data of the Pareto-optimal solutions of an aerodynamic transonic airfoil shape optimization problem showed that the optimized airfoils can be categorized into two families (low drag designs and high lift designs), where the lift is increased by changing the camber near the trailing edge among the low drag designs while the lift is increased by moving the lower surface upward among the high lift designs.

Modeling and testing of a morphing wing in open-loop architecture

Abstract: This paper presents the modeling and experimental testing of the aerodynamic performance of a morphing wing in open-loop architecture. We show the method used to acquire the pressure data from the external surface of the flexible wing skin, using incorporated Kulite pressure sensors and the instrumentation of the morphing controller. The acquired pressure data are analyzed through fast Fourier transforms to detect the magnitude of the noise in the surface airflow. Subsequently, the data are filtered by means of high-pass filters and processed by calculating the root mean square of the signal to obtain a plot diagram of the noise in the airflow. This signal processing is necessary to remove the inherent noise electronically induced from the Tollmien-Schlichting waves, which are responsible for triggering the transition from laminar to turbulent flow. The flexible skin is required to morph the shape of the airfoil through two actuation points to achieve an optimized airfoil shape based on the theoretical flow conditions similar to those tested in the wind tunnel. Two shape memory alloy actuators with a nonlinear behavior drive the displacement of the two control points of the flexible skin toward the optimized airfoil shape. Each of the shape memory actuators is activated by a power supply unit and controlled using the Simulink/MATLAB® software through a self-tuning fuzzy controller. The methodology and the results obtained during the wind-tunnel test proved that the concept and validity of the system in real time are discussed in this paper. Real-time acquisition and signal processing of pressure data are needed for further development of the closed-loop controller to obtain a fully automatic morphing wing system.

Aerodynamic Characterization of a NACA 0018 Airfoil at Low Reynolds Numbers

Abstract: 3to 200x10 3 and angles of attack from 0° to 18°. These data were used to characterize the separation bubble and determine lift coefficients. From these results, two distinct regions in the lift curves can be identified: a region of rapid and linear growth of the lift coefficients at low angles of attack and a region of more gradual and linear growth at higher pre-stall angles. Furthermore, the slope of the lift curve in each region is found to be linked to the rates of change in separation, transition, and reattachment locations with the angle of attack. These findings are substantiated by an analysis of the available experimental data for a NACA 0012 airfoil.