Airfoil Angle Research Articles

Active angular control of a sectioned airfoil using shape memory alloys and fuzzy controller

The present work illustrates an example of shape memory alloys and fuzzy controllers applied to active angular control of a sectioned airfoil. The main objective of the proposed control system is to modify the camber line of an airfoil based on the relative rotation between the different sections of the profile, given a reference angle. The changes of the sectioned airfoil angle result from the contraction/expansion of the shape memory alloy caused by heating of the wire with an electric current that changes its temperature by the Joule effect. Considering the presence of plant’s nonlinear effects, especially in the mathematical model of the alloy, this work proposes the application of a control system based on fuzzy logic theory. Through numerical and experimental tests, the performance of the fuzzy controller is compared with an on–off controller applied in an airfoil.

A study of long separation bubble on thick airfoils and its consequent effects

A parametric study has been performed to analyse the flow around the thick-symmetric NACA 0021 airfoil in order to better understand the characteristics and effects of long separation bubbles (LoSBs) that exist on such airfoils at low Reynolds numbers and turbulence intensities. In the article, the prediction capabilities of two recently-developed transition models, the correlation-based γ–Reθ model and the laminar-kinetic-energy-based κ–κL–ω model are assessed. Two-dimensional steady-state simulations indicated that the κ–κL–ω model predicted the separation and reattachment process accurately when compared with published experimental work. The model was then used to study the attributes and the effects of LoSBs as a function of the angle of attack, freestream turbulence intensity and Reynolds number. It was observed that LoSBs considerably degrade the aerodynamic performance of airfoils and lead to abrupt stall behaviour. It is, furthermore, illustrated that the presence of the LoSB leads to an induced camber effect on the airfoil that increases as the airfoil angle of attack increases due to the upstream migration of the bubble. An increase in the Reynolds number or turbulence levels leads to a reduction in the bubble extent, considerably improving the airfoil performance and leading to a progressive trailing-edge stall.

Turbulent intensity and Reynolds number effects on an airfoil at low Reynolds numbers

This work investigates the aerodynamics of a NACA 0012 airfoil at the chord-based Reynolds numbers (Rec) from 5.3 × 103 to 2.0 × 104. The lift and drag coefficients, CL and CD, of the airfoil, along with the flow structure, were measured as the turbulent intensity Tu of oncoming flow varies from 0.6% to 6.0%. The analysis of the present data and those in the literature unveils a total of eight distinct flow structures around the suction side of the airfoil. Four Rec regimes, i.e., the ultra-low (<1.0 × 104), low (1.0 × 104–3.0 × 105), moderate (3.0 × 105–5.0 × 106), and high Rec (>5.0 × 106), are proposed based on their characteristics of the CL-Rec relationship and the flow structure. It has been observed that Tu has a more pronounced effect at lower Rec than at higher Rec on the shear layer separation, reattachment, transition, and formation of the separation bubble. As a result, CL, CD, CL/CD and their dependence on the airfoil angle of attack all vary with Tu. So does the critical Reynolds number Rec,cr that divides the ultra-low and low Rec regimes. It is further noted that the effect of increasing Tu bears similarity in many aspects to that of increasing Rec, albeit with differences. The concept of the effective Reynolds number Rec,eff advocated for the moderate and high Rec regimes is re-evaluated for the low and ultra-low Rec regimes. The Rec,eff treats the non-zero Tu effect as an addition of Rec and is determined based on the presently defined Rec,cr. It has been found that all the maximum lift data from both present measurements and previous reports collapse into a single curve in the low and ultra-low Rec regimes if scaled with Rec,eff.

An implicit and adaptive nonlinear frequency domain approach for periodic viscous flows

An implicit nonlinear Lower–Upper symmetric Gauss–Seidel (LU-SGS) solver has been extended to the adaptive Nonlinear Frequency Domain method (adaptive NLFD) for periodic viscous flows. The discretized equations are linearized in both spatial and temporal directions, yielding an innovative segregate approach, where the effects of the neighboring cells are transferred to the right-hand-side and are updated iteratively. This property of the solver is aligned with the adaptive NLFD concept, in which different cells have different number of modes; hence, should be treated individually. The segregate analysis of the modal equations prevents assembling and inversion of a large left-hand-side matrix, when high number of modes are involved. This is an important characteristic for a selected flow solver of the adaptive NLFD method, where a high modal content may be required in highly unsteady parts of the flow field. The implicit nonlinear LU-SGS solver has demonstrated to be both robust and computationally efficient as the number of modes is increased. The developed solver is thoroughly validated for the laminar vortex shedding behind a stationary cylinder, high angle of attack NACA0012 airfoil, and a plunging NACA0012 airfoil. An order of magnitude improvement in the computational time is observed through the developed implicit approach over the classical modified 5-stage Runge–Kutta method.

Nonlinear angle control of a sectioned airfoil by using shape memory alloys

The present work illustrates an application of shape memory alloys and nonlinear controller applied to the active angular control of a sectioned airfoil. The main objective of the proposed control system is to modify the shape of the profile based on a reference angle. The change of the sectioned airfoil angle is resultant by the effect of shape memory of the alloy due to heating of the wire caused by an electric current that changes its temperature by Joule effect. Considering the presence of plant’s nonlinear effects, especially in the mathematical model of the alloy, this work proposes the application of an on-off control system.

STUDY ON AERODYNAMIC PERFORMANCE OF THE BIONIC AIRFOIL BASED ON THE SWALLOW'S WING

This paper demonstrates the design of the airfoil of small wind turbines, the bionic airfoil was inspired by the morphology of the swallow's extended wing. The wind tunnel tests on the bionic and standard airfoils NACA4412 were conducted, and the aerodynamic performances of the airfoils were numerically investigated. The results show that the bionic airfoil has better aerodynamic performance, the lift coefficient and lift-drag ratio are larger than those of the NACA4412; with the angle of attack increases, both the bionic and standard airfoils stall, but the stall characteristics of the bionic airfoil are better.

Effect of Plasma Actuator Placement on the Airfoil Efficiency at Poststall Angles of Attack

One of the new techniques in active flow control is the single-dielectric barrier discharge plasma actuator, which postpones the separation with external momentum injection to the boundary layer flow. In this paper, the effect of the presence of a plasma actuator on the passing flow is investigated both numerically and experimentally at a poststall angle of attack NLF0414 airfoil. These investigations are performed for a 45-cm cord NLF0414 airfoil with an incompressible 25-m/s velocity airflow, which provides a turbulent flow. Both numerical and experimental studies are done under the same conditions for two different cases: 1) no plasma actuator is located on the airfoil and 2) a plasma actuator is located on the top surface of the airfoil. Simulating the flow over the airfoil with the presence of a plasma actuator at 9.6 mm from the leading edge of the airfoil showed that the numerical distribution of the body force induced by the plasma actuator (computed using the model presented by Suzen <etal xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink"/> ) is in an appropriate correlation with the experimental results. It has been also shown that the presence of the plasma actuator on the airfoil could delay the separation and subsequently increase the airfoil's efficiency. Initially, the experimental and numerical results were compared and the written code, for simulation of the body force produced by the plasma actuator was verified. Then, the effect of the plasma actuator's location on the separation's delay at different angles of attack was numerically studied. In this case, the place of the plasma actuator on the airfoil was changed and its effect on postponement of the separation point was investigated. The results show that when the actuator is placed exactly at the leading edge of the airfoil, it has the greatest influence on postponing the separation. In addition, at an angle of attack of 18 <formula formulatype="inline" xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink"><tex Notation="TeX">$^{\circ}$</tex></formula> , it transfers the separation point from 16 to 90 mm. In addition, it results in an increase of <formula formulatype="inline" xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink"> <tex Notation="TeX">${\sim}{100\%}$</tex></formula> in the efficiency (the ratio of lift to drag coefficient) of the airfoil at this angle of attack.

Active Flow Control by Dielectric Barrier Discharge to Increase Stall Angle of a NACA0012 Airfoil

Active flow control has great importance in design of flying objects in general and control flaps in particular. Using dielectric barrier discharge as plasma actuators over control flaps can cause a change in position of separation point or can eliminate it. The present study aims to investigate experimentally the effects of the plasma actuator over a NACA0012 airfoil in high angles of attack up to 6° above the stall in order to displace the separation point. Therefore the effects of the plasma actuator with unsteady sine wave excitation and different duty cycles on airfoil pressure coefficient distribution and lift coefficient are studied. The experiments were performed at Rec = 3.7 × 105. The results indicate that lift coefficient increases with unsteady plasma actuation at post-stall angles of attack up to 6°. The improvement of the pressure coefficient distribution in upper side of airfoil for lower duty cycles of about 14 % was shown to be better than higher duty cycles of above 50 %.

Comparison of pneumatic jets and tabs for Active Aerodynamic Load Control

ABSTRACTA fast, efficient way to control loads on utility scale wind turbines is important for the growth of the wind industry. Microtabs and microjets are two Active Aerodynamic Load Control devices, which address this need. Both act perpendicular to the surface of the airfoil, and these actively controlled devices are used to mitigate changes in aerodynamic loading experienced by wind turbine rotors due to wind gusts, wind shear, or other atmospheric phenomena. This work explores the aerodynamic effects of microjets and then compares them to those of microtabs. Flow around an airfoil with an activated microjet at the trailing edge has been simulated using the Reynolds‐averaged Navier–Stokes solver OVERFLOW‐2. Using a Chimera overset grid topology, a microjet has been placed near the trailing edge of the lower surface of a NACA 0012 airfoil. For a jet velocity about half of the freestream velocity, the microjet can change the lift up to ΔCL = 0.2, but the amount of change varies with the momentum coefficient of the jet. The change in lift is not symmetric for positive and negative angles of attack due to changes in the boundary layer thickness with angle of attack. Increasing the Reynolds number reduces the effectiveness of the microjet only slightly. The effects of jet velocity, jet activation time, and airfoil angle of attack on airfoil lift, drag, and pitching moment are compared with previous work, which illustrates the deployment of a microtab at the 95% chord location of a NACA 0012 airfoil. This study shows that microjets and microtabs have very similar responses in lift and pitching moment, but the drag for the microjet is noticeably lower. Copyright © 2013 John Wiley &amp; Sons, Ltd.

Variation of Aerodynamic Load Engineering Analysis during Wind Turbine Run

In order to improve design reliability of wind turbine, it is needed that calculating method of aerodynamic load during wind turbine run. In paper from the angle of the project, NACA special airfoil of wind turbine is analyzed. Combined with thin-theory, airfoil angle of attack variation is deduced, meanwhile wind turbine actual force is calculated in each blade location point when blade of wind turbine is running based on wind shear theory and tower shadow effect. According to actual condition calculation method is engineering amplified, aerodynamic load calculation method of wind turbine blade is obtained. By this method aerodynamic load which is calculated match with experiment result, it fits better.

1225 EFFECT OF TURBULENCE INTENSITY ON THE LOW REYNOLDS NUMBER AIRFOILWAKE

The work investigates the effect of turbulent intensity (Tu) on the force and wake of a NACA0012 airfoil at chord Reynolds number Rec = 5.3X103 and 2X104. Lift and drag coefficients (CL and CD) on and flow fields around the airfoil were measured with Tu varied from 0.6% to 6.0%. Four Rec regimes are proposed based on the characteristics of the maximum lift coefficient (CL,max), i.e., ultra-low (Rec < 104), low (104 〜 3X105), moderate (3X105 〜 5X106) and high (> 5X106) Rec. It is noted that at Rec = 5.3X103 (ultra-low Rec regime) the stall is absent for Tu = 0.6% but occurs for Tu = 2.6% and 6.0%. In the low Rec regime, the Tu influence wanes and the critical Rec, above which stall occurs, decreases with increasing Tu. When the airfoil angle (α) of attack varies from 0° to 30°, at low Tu (0.6%) the shear layers over the airfoil are mostly laminar in the ultra-low Rec regime but involve transition in the low Rec regime. The effect of increasing Tu on flow bears similarity to that of increasing Rec, albeit with a difference. The flow separation shifts upstream with increasing Rec but downstream with increasing Tu.

Numerical study of flows past airfoils with wavy surfaces

The paper presents three-dimensional numerical studies on the aerodynamic characteristics of two different modified NACA0012 airfoils with different wavy surfaces using the large eddy simulations. Two types of wavy airfoils are investigated with wavy airfoil-A having a sinusoidal waviness on upper and lower surfaces with a constant chord length, while wavy airfoil-B having sinusoidal variation in both of the leading and trailing edges as well as on the upper and lower surfaces along the spanwise direction. The force characteristics and the flow structures are captured and compared with a corresponding standard NACA0012 airfoil with a Reynolds number of Re=1.6×105. The flow structures and surface pressure distributions on wavy airfoils were found to be significantly different from those on a conventional NACA0012 airfoil. For angles of attack less than the baseline stall angle of a NACA0012 airfoil, a slight decrease of lift coefficient was observed for both types of wavy airfoils, while the lift coefficient for the wavy airfoil-B increases up to 20% greater than that of a NACA0012 airfoil when the angle of attack is larger than the baseline stall angle of 13°. The flow over the leading edge of wavy airfoil-B remained attached at post stall angles of attack. In general, the wavy airfoil-A just exhibits a suppression the airfoil’s fluctuation force, while the wavy airfoil-B demonstrates an advantageous aerodynamic effect on the control of loss of lift in the post stall regime of a conventional NACA0012 airfoil.

116 4枚羽根垂直軸型風力発電用風車の翼形状の最適化(G01-8流体抵抗とその応用,G01 熱流体部門)

The present study investigate the effect of flap angle of airfoil and wing attack angle to obtain the amount of power generation. We also examine the minimum starting wind velocity of the windmill. Flap wing generate electricity efficiently by using the high wing attack angle. And using high flap angle make it possible to obtain minimum starting velocity. To obtain high output power, it is better to keep flap angle of 0° and using high wing attack angle we have obtain maximum output power with wing attack angle of a =29.9° 〜-37.1° .

Stall Onset on Airfoils at Moderately High Reynolds Number Flows

The inception of leading-edge stall on two-dimensional smooth thin airfoils at moderately high Reynolds number flows [in the range O(104) to O(106)] is investigated by an asymptotic approach and numerical simulations. The asymptotic theory is based on the work of Rusak (1994) and demonstrates that a subsonic flow about a thin airfoil can be described in terms of an outer region, around most of the airfoil chord, and an inner region, around the nose, that asymptotically match each other. The flow in the outer region is dominated by the classical thin airfoil theory. Scaled (magnified) coordinates and a modified (smaller) Reynolds number are used to correctly account for the nonlinear behavior and extreme velocity changes in the inner region, where both the near stagnation and high suction areas occur. It results in a model (simplified) problem of a uniform flow past a semi-infinite parabola with a far-field circulation governed by a parameter à that is related to the airfoil’s angle of attack, nose radius of curvature, and camber and to the flow Mach number. The model parabola problem consists of a compressible and viscous flow described by the steady Navier-Stokes equations. This problem is solved numerically for various values of à using a Reynolds-averaged Navier-Stokes flow solver, and utilizing the Spalart-Allmaras viscous turbulent model to account for near-wall turbulence. The value Ãs where a large separation zone first appears in the nose flow concurrent with a sudden increase in the minimum pressure coefficient is determined. The change of Ãs with the modified Reynolds number is determined. These values indicate the stall onset on the airfoil at various flow conditions. The predictions according to this approach show good agreement with results from both numerical simulations and available experimental data of the stall of thin airfoils. This simplified approach provides a criterion to determine the stall angle of airfoils with a parabolic nose and the effect of airfoil’s thickness ratio, nose radius of curvature, camber and flaps, and flow compressibility on the onset of stall. This approach also presents an analysis method that can be used to predict the stall of airfoils with alternative nose geometry.

Numerical Investigation of Flow past a Wavy Airfoil

The effect of the wavy surface on the aerodynamic characteristics of an airfoil is studied using the large eddy simulations. A more gentle lift characteristic is obtained during stall. For angles of attack less than the baseline stall angle of a NACA0012 airfoil, a lift coefficient reduction was observed for the wavy airfoils, while the lift coefficient increases up to 23% greater than that of a NACA0012 airfoil when the angle of attack is larger than the baseline stall angle of the NACA0012 airfoil.

Aeroacoustic beamforming for airfoil trailing edge tonal noise sources.

Airfoils produce tonal and broadband noise at low to moderate Reynolds number flow conditions. The effect of variation in flow Reynolds number and airfoil angle of attack on the quantification and localization of the acoustic sources responsible for the observed tonal noise component is still uncertain. An investigation of the effect of Reynolds number, angle of attack, and acoustic frequency on the tonal noise sources will be presented. This study was performed using aeroacoustic beamforming in a small anechoic wind tunnel. Accurately locating the tonal noise source locations of an airfoil will improve knowledge of how airfoils produce tonal noise in this flow regime which will help the implementation of noise reduction techniques. In addition, some of the problems associated with beamforming low frequency noise sources and close proximity coherent tonal noise sources using conventional beamforming techniques and deconvolution algorithms will be addressed. [The authors would like to thank the School of Mechanical Engineering at the University of Adelaide for access and use of the facilities and equipment detailed in this presentation.]

Nature of Wakelike and Jetlike Axial Tip Vortex Flows

The mechanisms responsible for the wakelike and jetlike axial flows of a tip vortex were investigated experimentally at Re = 3.07 x 10 5 . Both square and round tips were tested. Along the wing tip, a pocket of a higher- than-freestream, or jetlike, fluid was continually present, regardless of the tip condition and airfoil angle of attack α. Its maximum velocity increased with α and always exhibited a peak value around the trailing edge. In the near field, this jetlike fluid pocket was entrained by the shear layers and the wing wake, and it resulted in a wakelike axial flow for smaller α. For higher α, the jetlike fluid pocket was, however, surrounded by the shear layers that, in turn, protected it from the destructive effects of the wing wake and remained jetlike. The switchover angle at which the axial flow switched from wakelike to jetlike appeared around 7 deg, regardless of the tip condition, which also coincided with the angle at which the lift-to-drag ratio was a maximum. Finally, the round-tip-produced vortex was more concentrated and had a higher peak tangential velocity and core circulation but a smaller core radius compared with those of the square tip.

Panel methods for airfoils in turbulent flow

This paper describes how panel methods can be used to calculate the unsteady loading and radiated noise from airfoils in incompressible turbulent flow, while completely accounting for the mean flow distortion of the turbulence in the vicinity of the blade. Formulations based on the velocity and on the stagnation enthalpy are discussed. In three-dimensional flows, care must be taken with the velocity-based formulation to avoid singular behavior associated with vortex stretching by the mean flow. The velocity-based method is implemented in two dimensions to illustrate application of these methods, and is validated against Amiet's theory. Calculations showing the effect of blade thickness and angle of attack on the unsteady loading spectra are given. It is concluded that airfoil angle of attack has only a small effect on the unsteady loading, but that blade thickness reduces the spectral levels at high frequencies.

Bubble Burst Control for Stall Suppression on a NACA 631-012 Airfoil

Based on the novel concepts of the burst control plate, the current article aims to demonstrate the effectiveness of the burst control plate attached at different locations (5%, 7.5% and 10% of the chord of the airfoil) on a NACA 631-012 airfoil section for laminar separation bubble burst delay, stall suppression, lift augmentation and drag reduction at a chord Reynolds number of 1.3x10 5 . The effectiveness of the burst control plate is governed by its height, width, geometry and the distance between the leading-edge of the airfoil and the trailing-edge of the plate. Flow patterns around the airfoil with and without the plate attachment were studied with an aid of smoke flow visualization. The results suggest that the stall angle of the original airfoil (without the burst control plate) occurs at 10 o and it is successfully postponed to approximately 13 o when the burst control plate is attached onto the airfoil. Lift and drag forces were measured for a range of angle-of-attack from 0 o to +20 o . The experimental results have shown that the lift generated on the NACA 631-012 airfoil with the plate attachment and above 9 o angle-of-attack is significantly higher than the original airfoil, while the drag is sufficiently reduced. The overall results demonstrated the application of the burst control plate on the NACA 631-012 airfoil can be effective means of bubble burst control and airfoil stall suppression in low speed flows.

Airfoil Stall Suppression by Use of a Bubble Burst Control Plate

To suppress the stall on an NACA 0012 airfoil, a thin plate (hereafter referred to as a burst control plate) was attached on the airfoil to delay the burst of laminar separation bubbles formed at the stall angle. The burst control plate is used to enhance the vortical structures in the separated shear layer. Wind-tunnel tests were conducted at a chord Reynolds number of 1.3 x 10 5 . Flow visualization tests and surface pressure measurements showed that the burst control plate placed on the airfoil suppresses the bubble burst at a wide range of angle of attack and therefore, both the stall angle and the maximum lift coefficient are increased. The particle image velocimetry measurements indicated that when the angle of attack of the airfoil with the burst control plate is set higher than the stall angle of the original airfoil, a similar flow structure of the original airfoil was observed.