Effect of amplitude and mean angle-of-attack on the boundary layer of an oscillating aerofoil

2008 ◽  
Vol 112 (1138) ◽  
pp. 705-713 ◽  
Author(s):  
M. R. Soltani ◽  
A. Bakhshalipour
Keyword(s):  
Boundary Layer ◽  
Transition Point ◽  
Angle Of Attack ◽  
Wind Tunnel Test ◽  
Leading Edge ◽  
Stream Velocity ◽  
Velocity Amplitude ◽  
Surface Pressure Distribution ◽  
Reduced Frequency ◽  
Hot Film

Abstract Extensive experiments were conducted to study the effect of various parameters on the surface pressure distribution and transition point of an aerofoil section used in a wind turbine blade. In this paper details of the variation of transition point on the aforementioned aerofoil are presented. The aerofoil spanned the wind-tunnel test section and was oscillated sinusoidally in pitch about the quarter chord. The imposed variables of the experiments were free stream velocity, amplitude of motion, mean angle-of-attack, and oscillation frequency. The spatial-temporal progressions of the leading-edge transition point and the state of the unsteady boundary-layer were measured using eight closely-spaced, hot-film sensors (HFS). The measurements show that: (i) Reduced frequency has a pronounced effect on the variations of the transition point. (ii) There exists a hysteresis loop in the dynamic transition location and its shape varies with the reduced frequency and mean angle-of-attack.

Experimental and Computational Investigation of a Simplified Geometry of an Engine/Pylon/Wing Installation at High-Lift Flight Conditions

2013 ◽  
Author(s):  
Matthieu Lucas ◽  
Yannick Bury ◽  
Cyril Bonnaud ◽  
Laurent Joly
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Longitudinal Vortex ◽  
Vortical Flow ◽  
Stream Velocity ◽  
Piv Measurements ◽  
Comprehensive Knowledge ◽  
Oil Flow

This paper focuses on the numerical and experimental characterization of the vortex structures that develop along a simplified geometry of a wing equipped with pylon-mounted engine at low speed/high angle of attack flight conditions. In these conditions, the presence of the engine installation under the wing induces a complex and unsteady vortical flow field at the nacelle/pylon/wing junctions which interacts with the upper wing boundary layer and leads to a drop of aircraft performances. In order to gain insight into the physics driving this interaction, it is proposed to isolate its fundamental mechanisms by simplifying the problem. The parameters of interest that led to the simplification of the model are first described. As a first step into a more comprehensive knowledge of this complex physics, this study is initially conducted at a Reynolds number of 200000, based on the chord wing and on the free stream velocity. Two configurations of angle of attack and sideslip angles (α = 8°/β = 0° and α = 8°/β = 30°) have been investigated. This work relies on unsteady RANS computations, oil flow visualizations and 3C-PIV measurements. The vortex dynamics thus produced is described in terms of vortex core position, intensity, size and turbulent intensity thanks to a vortex tracking post-processing algorithm. In addition, the analysis of the velocity flow field obtained from the PIV measurements will highlight the influence of the longitudinal vortex issued from the pylon/wing junction on the separation process of the boundary layer near the upper wing leading-edge.

Experimental Investigations of Laminar Separation Bubble for a Flow Past an Airfoil

2010 ◽  
Author(s):  
Deepakkumar M. Sharma ◽  
Kamal Poddar
Keyword(s):  
Boundary Layer ◽  
Flow Separation ◽  
Separation Bubble ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Boundary Layer Separation ◽  
Experimental Investigations ◽  
Stream Velocity ◽  
Laminar Separation Bubble ◽  
Laminar Separation

Wind tunnel experiments were carried out on NACA 0015 airfoil model to investigate the formation of laminar separation bubble on the upper surface of the airfoil by varying angle of attack from −5° to 25° with respect to the free stream velocity at constant Reynolds number varying from 0.2E06 to 0.6E06. Pressure signals were acquired from the pressure ports selected at the mid-span of the airfoil model along the chord. Static stall characteristics were obtained from the surface pressure distribution. The flow separation was found to be a trailing edge turbulent boundary layer separation preceded with a laminar separation bubble. Flow Visualizations were done by using Surface Oil flow Technique for qualitative analysis of the transition zone formed due to the presence of laminar separation bubble As the angle of attack is increased the separation bubble moves towards the leading edge of the airfoil and finally gets shredded or burst at a particular angle of attack resulting in leading edge turbulent flow separation which induces the static stall condition. The flow separation process is critically analyzed and the existence of laminar separation bubble is visualized and quantified with the increase in angle of attack and Re. Effect of Re and angle of attack on the various boundary layer and Separation bubble parameters are obtained and analyzed.

scholarly journals Leading Edge Blowing to Mimic and Enhance the Serration Effects for Aerofoil

2021 ◽  
Vol 11 (6) ◽  
pp. 2593
Author(s):  
Yasir Al-Okbi ◽  
Tze Pei Chong ◽  
Oksana Stalnov
Keyword(s):  
Boundary Layer ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Boundary Layer Separation ◽  
Shear Instability ◽  
Vortical Flow ◽  
High Angle ◽  
High Angle Of Attack ◽  
Layer Separation ◽  
Aerodynamic Performances

Leading edge serration is now a well-established and effective passive control device for the reduction of turbulence–leading edge interaction noise, and for the suppression of boundary layer separation at high angle of attack. It is envisaged that leading edge blowing could produce the same mechanisms as those produced by a serrated leading edge to enhance the aeroacoustics and aerodynamic performances of aerofoil. Aeroacoustically, injection of mass airflow from the leading edge (against the incoming turbulent flow) can be an effective mechanism to decrease the turbulence intensity, and/or alter the stagnation point. According to classical theory on the aerofoil leading edge noise, there is a potential for the leading edge blowing to reduce the level of turbulence–leading edge interaction noise radiation. Aerodynamically, after the mixing between the injected air and the incoming flow, a shear instability is likely to be triggered owing to the different flow directions. The resulting vortical flow will then propagate along the main flow direction across the aerofoil surface. These vortical flows generated indirectly owing to the leading edge blowing could also be effective to mitigate boundary layer separation at high angle of attack. The objectives of this paper are to validate these hypotheses, and combine the serration and blowing together on the leading edge to harvest further improvement on the aeroacoustics and aerodynamic performances. Results presented in this paper strongly indicate that leading edge blowing, which is an active flow control method, can indeed mimic and even enhance the bio-inspired leading edge serration effectively.

The Evolution of Streamwise Counter-Rotating Vortices in Flat Plate Boundary Layer With Leading Edge Pattern

2015 ◽  
Author(s):  
Seyed Mohammad Hasheminejad ◽  
Hatsari Mitsudharmadi ◽  
S. H. Winoto ◽  
Kim Boon Lua ◽  
Hong Tong Low
Keyword(s):  
Boundary Layer ◽  
Flat Plate ◽  
Plate Boundary ◽  
Leading Edge ◽  
Streamwise Vortex ◽  
Streamwise Velocity ◽  
Stream Velocity ◽  
Hot Wire ◽  
Layer Flow ◽  
Different Characteristics

The evolution of streamwise counter-rotating vortices induced by different leading edge patterns is investigated quantitatively using hot-wire anemometer. A notched and triangular leading edge with the same wavelength and amplitude were designed to induce streamwise vortices over a flat plate at Reynolds number (based on the wavelength of the leading edge patterns) of 3080 corresponding to free-stream velocity of 3 m/s. The streamwise velocity at different streamwise locations collected and analyzed using a single wire probe hot-wire anemometer showed reveal different characteristics of boundary layer flow due to the presence of these two leading edge patterns. The major difference is the appearance of an additional streamwise vortex between the troughs of the notched pattern. Such vortices increase the mixing effect in the boundary layer as well as the velocity profile.

Periodic Transition on an Axial Compressor Stator — Incidence and Clocking Effects: Part I — Experimental Data

1998 ◽  
Author(s):  
G. J. Walker ◽  
J. D. Hughes ◽  
W. J. Solomon
Keyword(s):  
Guide Vane ◽  
Axial Compressor ◽  
Separated Flow ◽  
Leading Edge ◽  
Inlet Guide Vane ◽  
Turbulent Spot ◽  
Pressure Surface ◽  
Surface Pressure Distribution ◽  
Surface Transition ◽  
Hot Film

Periodic wake-induced transition on the outlet stator of a 1.5 stage axial compressor is examined using hot-film arrays on both the suction and pressure surfaces. The time-mean surface pressure distribution is varied by changing the blade incidence, while the freestream disturbance field is altered by clocking of the stator relative to an inlet guide vane row. Ensemble average plots of turbulent intermittency and relaxation factor (extent of calmed flow following the passage of a turbulent spot) are presented. These show the strength of periodic wake-induced transition phenomena to be significantly influenced by both incidence and clocking effects. The nature and extent of transition by other modes (natural, bypass and separated flow transition) are altered accordingly. Leading edge and mid-chord separation bubbles are affected in a characteristically different manner by changing freestream periodicity. There are noticeable differences between suction and pressure surface transition behavior, particularly as regards the strength and extent of calming. In Part II of this paper, the transition onset observations from the compressor stator are used to evaluate the quasi-steady application of conventional transition correlations to predict unsteady transition onset on the blading of an embedded axial compressor stage.

Numerical study of boundary layer transition using intermittency model

2019 ◽  
Vol 91 (8) ◽  
pp. 1156-1168 ◽  
Author(s):  
Massoud Tatar ◽  
Mojtaba Tahani ◽  
Mehran Masdari
Keyword(s):  
Boundary Layer ◽  
Shear Stress ◽  
Transport Equation ◽  
Transition Point ◽  
Static Pressure ◽  
Boundary Layer Transition ◽  
Layer Transition ◽  
Content Type ◽  
Reduced Frequency ◽  
Pressure Suction

Purpose In this paper, the applicability of shear stress transport k-ω model along with the intermittency concept has been investigated over pitching airfoils to capture the laminar separation bubble (LSB) position and the boundary layer transition movement. The effect of reduced frequency of oscillations on boundary layer response is also examined. Design/methodology/approach A two-dimensional computational fluid dynamic code was developed to compute the effects of unsteadiness on LSB formation, transition point movement, pressure distribution and lift force over an oscillating airfoil using transport equation of intermittency accompanied by the k-ω model. Findings The results indicate that increasing the angle of attack over the stationary airfoil causes the LSB size to shorten, leading to a rise in wall shear stress and pressure suction peak. In unsteady cases, both three- and four-equation models are capable of capturing the experimentally measured transition point well. The transition is delayed for an unsteady boundary layer in comparison with that for a static airfoil at the same angle of attack. Increasing the unsteadiness of flow, i.e. reduced frequency, moves the transition point toward the trailing edge of the airfoil. This increment also results in lower static pressure suction peak and hence lower lift produced by the airfoil. It was also found that the fully turbulent k-ω shear–stress transport (SST) model cannot capture the so-called figure-of-eight region in lift coefficient and the employment of intermittency transport equation is essential. Practical implications Boundary layer transition and unsteady flow characteristics owing to airfoil motion are both important for many engineering applications including micro air vehicles as well as helicopter blade, wind turbine and aircraft maneuvers. In this paper, the accuracy of transition modeling based on intermittency transport concept and the response of boundary layer to unsteadiness are investigated. Originality/value As a conclusion, the contribution of this paper is to assess the ability of intermittency transport models to predict LSB and transition point movements, static pressure distribution and aerodynamic lift variations and boundary layer flow pattern over dynamic pitching airfoils with regard to oscillation frequency effects for engineering problems.

scholarly journals Experimental and Computational Analysis of Model–Support Interference in Low-Speed Wind-Tunnel Testing of Fuselage-Boundary-Layer Ingestion

2019 ◽  
Vol 304 ◽  
pp. 02020
Author(s):  
Biagio Della Corte ◽  
André A.V. Perpignan ◽  
Martijn van Sluis ◽  
Arvind Gangoli Rao
Keyword(s):  
Boundary Layer ◽  
Wind Tunnel ◽  
Wind Tunnel Test ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Junction Flow ◽  
Tunnel Model ◽  
Aerodynamic Interaction ◽  
Support Interference ◽  
Model Support

Junction flow caused by the aerodynamic interaction between a wind-tunnel model and the support structure can largely influence the flowfield and hence the experimental results. This paper discusses a combined numerical and experimental study which was carried out to mitigate the model–support interference in a wind-tunnel test setup for the study of fuselage boundary-layer ingestion. The setup featured an axisymmetric fuselage mounted through a support beam, covered by a wing-shaped fairing. The junction flow appearing at the fuselage–fairing connection produced undesired flow distortions at the fuselage aft section, due to the formation of an horseshoe vortex structure at the fairing leading edge. Numerical and experimental analysis were performed with the aim of reducing the distortion intensity by improving the fairing design. Results show that modifying the leading-edge shape of the fairing effectively decreased the flowfield distortions. Moreover, the addition of a dummy fairing diametrically opposed to the first one was found to be beneficial due to the enhancement of the configuration symmetry.

The solution of heat transfer problems by the Wiener-Hopf technique. I. Leading edge of a hot film

1973 ◽  
Vol 333 (1594) ◽  
pp. 347-362 ◽  
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Asymptotic Expansion ◽  
Half Plane ◽  
Leading Edge ◽  
Infinite Plane ◽  
Longitudinal Diffusion ◽  
Boundary Layer Equations ◽  
Leading Term ◽  
Hot Film

An incompressible fluid of constant thermal diffusivity flows with velocity Sy in the x -direction over the infinite plane wall y = 0. The half-plane y = 0, x > 0 is maintained at a uniform temperature T 1 greater than the temperature T 0 of the oncoming fluid. The adiabatic boundary condition T y = 0 is imposed on the half-plane y = 0, x < 0. An exact solution for the dimensionless heat transfer from the heated half-plane x > 0, incorporating longitudinal diffusion, is obtained by the Wiener-Hopf technique, and is reduced to a single convergent real integral which is evaluated numerically. An asymptotic expansion is made in inverse powers of x , whose leading term is Lévêque’s (1928) boundary-layer solution. Subsequent terms in the expansion lead to a determination of the coefficients of the eigenfunctions of the boundary-layer equations which would remain arbitrary in a direct asymptotic expansion of the governing equation.

Influence of Boundary Layer Behavior on Aerodynamic Coefficients of a Swept-Back Wing

2006 ◽  
Vol 129 (6) ◽  
pp. 674-681 ◽  
Author(s):  
S. C. Yen ◽  
C. M. Hsu
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Bluff Body ◽  
Separation Bubble ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Cross Sectional ◽  
Wing Model ◽  
Oil Flow ◽  
Back Angle

The effects of the Reynolds number and angle of attack on the boundary layer and the aerodynamic performance of a finite swept-back wing are studied experimentally. The cross-sectional profile of the wing is NACA 0012 (aspect ratio=10), and the sweep-back angle is 15 deg. The Reynolds number is set in the range of 30,000–130,000. The boundary layer field is visualized with surface oil-flow techniques. Six characteristic flow regimes—laminar separation, separation bubble, leading-edge bubble, bubble burst, turbulent separation, and bluff-body wake—are categorized and studied by considering the Reynolds numbers and angles of attack. The characteristic behaviors of boundary layer significantly affect the lift, drag, and moment coefficients. The bubble length shortens significantly in the separation bubble and leading-edge bubble regimes as the angle of attack rises. The aerodynamic performances demonstrate that the swept-back wing model has no hysteresis.

Experimental investigation of the leading edge vortex formation on unsteady boundary layer

2017 ◽  
Vol 232 (18) ◽  
pp. 3263-3280
Author(s):  
Ali R Davari ◽  
Rezvan Abdollahi ◽  
Ehsaneddin Azimizadeh
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Leading Edge ◽  
Adverse Pressure Gradient ◽  
Unsteady Boundary Layer ◽  
Leading Edge Vortex ◽  
Reduced Frequency ◽  
Blade Section ◽  
Edge Vortex ◽  
The Impact

Extensive experimental studies have been performed to investigate the unsteady boundary layer behavior over a plunging wind turbine blade section. The studies have been undertaken at various combinations of reduced frequencies, Reynolds numbers, and locations. A boundary layer rake has been carefully manufactured and utilized for velocity measurements inside the unsteady boundary layer. The measurement has been conducted in pre-static stall conditions. The reduced frequency and free stream velocity have varied from 0.005 to 0.1, and 30 to 60 m/s, respectively. To cover all possible scenarios, the streamwise positions of measurements have been chosen to be in favorable (x/c = 0.37), almost zero (x/c = 0.47), and adverse pressure gradient (x/c = 0.57) regions, on the blade section. The velocity inside the boundary layer has shown high sensitivity to the reduced frequency in the different pressure gradient regions. In some definite test cases, velocity inside boundary layer has shown beating phenomena, which is the result of the periodical appearance of the leading edge vortex. The impact of the leading edge vortex on the velocity has been observed to be more evident, in some cases, in the form of signal beating. This signature has been more evident, as the rake entered the adverse pressure gradient region. In order to quantify this observed phenomenon, the time-dependent velocity data have been transformed into the frequency domain, utilizing the discrete Fourier transformation. Even though the leading edge vortex has been continuously developed on the profile, and then has shed toward the leading edge, during each cycle on a plunging profile, the dominant frequency throughout this process has been measured to be about 4 Hz for this blade section.

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