A Turbulent Boundary Layer Flow Over an Open Shallow Cavity

2016 ◽  
Author(s):  
Khaled J. Hammad
Keyword(s):  
Boundary Layer ◽  
Large Scale ◽  
Vector Fields ◽  
Turbulence Models ◽  
Leading Edge ◽  
Field Measurements ◽  
Recirculation Region ◽  
Mean Flow ◽  
Trailing Edge ◽  
Cavity Depth

Particle Image Velocimetry (PIV) was used to study the flow structure and turbulence, upstream, over, and downstream a shallow open cavity. Three sets of PIV measurements, corresponding to a turbulent incoming boundary layer and a cavity length-to-depth ratio of four, are reported. The cavity depth based Reynolds numbers were 21,000; 42,000; and 54,000. The selected flow configuration and well characterized inflow conditions allow for straightforward assessment of turbulence models and numerical schemes. All mean flow field measurements display a large flow recirculation region, spanning most of the cavity and a smaller, counter-rotating, secondary vortex, immediately downstream of the cavity leading edge. The Galilean decomposed instantaneous velocity vector fields, clearly demonstrate two distinct modes of interaction between the free shear and the cavity trailing edge. The first corresponds to a cascade of vortical structures emanating from the tip of the leading edge of the cavity that grow in size as they travel downstream and directly interact with the trailing edge, i.e., impinging vortices. The second represents vortices that travel above the trailing edge of the cavity, i.e., non-impinging vortices. In the case of impinging vortices, a strong, large scale region of recirculation forms inside the cavity and carries the flow disturbances, arising from the impingement of vortices on the trailing edge of the cavity, upstream in a manner that interacts with and influences the flow as it separates from the cavity leading edge.

Characterization of Vortex Dynamics in the Near Wake of an Oscillating Flexible Foil

2016 ◽  
Vol 138 (10) ◽  
Author(s):  
Firas F. Siala ◽  
Alexander D. Totpal ◽  
James A. Liburdy
Keyword(s):  
Large Scale ◽  
Flow Characteristics ◽  
Leading Edge ◽  
Wake Flow ◽  
Small Scale ◽  
Mean Flow ◽  
Near Wake ◽  
Trailing Edge ◽  
Swirling Strength ◽  
The Mean

An experimental study was conducted to explore the effect of surface flexibility at the leading and trailing edges on the near-wake flow dynamics of a sinusoidal heaving foil. Midspan particle image velocimetry (PIV) measurements were taken in a closed-loop wind tunnel at a Reynolds number of 25,000 and at a range of reduced frequencies (k = fc/U) from 0.09 to 0.20. Time-resolved and phase-locked measurements are used to describe the mean flow characteristics and phase-averaged vortex structures and their evolution. Large-eddy scale (LES) decomposition and swirling strength analysis are used to quantify the vortical structures. The results demonstrate that trailing edge flexibility has minimal influence on the mean flow characteristics. The mean velocity deficit for the flexible trailing edge and rigid foils remains constant for all reduced frequencies tested. However, the trailing edge flexibility increases the swirling strength of the small-scale structures, resulting in enhanced cross-stream dispersion. Flexibility at the leading edge is shown to generate a large-scale leading edge vortex (LEV) for k ≥ 0.18. This results in a reduction in the swirling strength due to vortex interactions when compared to the flexible trailing edge and rigid foils. Furthermore, it is shown that the large-scale LEV is responsible for extracting a significant portion of energy from the mean flow, reducing the mean flow momentum in the wake. The kinetic energy loss in the wake is shown to scale with the energy content of the LEV.

Characterization of Vortex Dynamics in the Near Wake of an Oscillating Flexible Foil

2016 ◽  
Author(s):  
Firas F. Siala ◽  
Alexander D. Totpal ◽  
James A. Liburdy
Keyword(s):  
Large Scale ◽  
Flow Characteristics ◽  
Leading Edge ◽  
Mean Flow ◽  
Near Wake ◽  
Trailing Edge ◽  
Mean Velocity ◽  
Leading Edge Vortex ◽  
Edge Vortex ◽  
The Mean

An experimental study was conducted to explore the effect of surface flexibility at the leading and trailing edges on the near-wake flow dynamics of a sinusoidal heaving foil. Mid-span particle image velocimetry measurements were taken in a closed loop wind tunnel at a Reynolds number of 25,000 and at a range of reduced frequencies (k = fc/U) from 0.09–0.20. Time resolved and phase locked measurements were used to describe the mean flow characteristics and phase averaged vortex structures and their evolution throughout the oscillation cycle. Large eddy scale decomposition and swirl strength analysis were used to quantify the effect of flexibility on the vortical structures. The results demonstrate that flexibility at the trailing edge has a minimal influence on the mean flow characteristics when compared to the purely rigid foil. The mean velocity deficit for the flexible trailing edge and rigid foils is shown to remain constant for all reduced frequencies tested. However, the trailing edge flexibility increases the swirl strength of the small scale structures, which results in enhanced cross stream dispersion of the mean velocity profile. Flexibility at the leading edge is shown to generate a large scale leading edge vortex for k ≥ 0.18. This results in a reduction in the swirl strength due to the complex vortex interactions when compared to the flexible trailing edge and rigid foils. Furthermore, it is shown that the large scale leading edge vortex is responsible for extracting a significant portion of the energy from the mean flow, resulting in a substantial reduction of mean flow momentum in the wake. The kinetic energy loss in the wake is shown to scale well with the energy content of the leading edge vortex.

Characteristics of a separated flow past a semicircular leading-edge airfoil model under different imposed pressure gradient

2021 ◽  
pp. 095441002110212
Author(s):  
K Anand ◽  
KT Ganesh
Keyword(s):  
Boundary Layer ◽  
Particle Image Velocimetry ◽  
Pressure Gradient ◽  
Particle Image ◽  
Separation Bubble ◽  
Separated Flow ◽  
Leading Edge ◽  
Field Measurements ◽  
Bench Mark ◽  
Image Velocimetry

The effect of pressure gradient on a separated boundary layer past the leading edge of an airfoil model is studied experimentally using electronically scanned pressure (ESP) and particle image velocimetry (PIV) for a Reynolds number ( Re) of 25,000, based on leading-edge diameter ( D). The features of the boundary layer in the region of separation and its development past the reattachment location are examined for three cases of β (−30°, 0°, and +30°). The bubble parameters such as the onset of separation and transition and the reattachment location are identified from the averaged data obtained from pressure and velocity measurements. Surface pressure measurements obtained from ESP show a surge in wall static pressure for β = −30° (flap deflected up), while it goes down for β = +30° (flap deflected down) compared to the fundamental case, β = 0°. Particle image velocimetry results show that the roll up of the shear layer past the onset of separation is early for β = +30°, owing to higher amplification of background disturbances compared to β = 0° and −30°. Downstream to transition location, the instantaneous field measurements reveal a stretched, disoriented, and at instances bigger vortices for β = +30°, whereas a regular, periodically shed vortices, keeping their identity past the reattachment location, is observed for β = 0° and −30°. Above all, this study presents a new insight on the features of a separation bubble receiving a disturbance from the downstream end of the model, and these results may serve as a bench mark for future studies over an airfoil under similar environment.

Experimental and Theoretical Investigation of the Supersonic Boundary Layer Stability on the Swept Wing

2008 ◽  
Vol 3 (3) ◽  
pp. 34-38
Author(s):  
Sergey A. Gaponov ◽  
Yuri G. Yermolaev ◽  
Aleksandr D. Kosinov ◽  
Nikolay V. Semionov ◽  
Boris V. Smorodsky
Keyword(s):  
Boundary Layer ◽  
Three Dimensional ◽  
Leading Edge ◽  
Supersonic Boundary Layer ◽  
Mean Flow ◽  
Swept Wing ◽  
Wing Model ◽  
Dimensional Boundary ◽  
The Stability ◽  
Good Agreement

Theoretical and an experimental research results of the disturbances development in a swept wing boundary layer are presented at Mach number М = 2. In experiments development of natural and small amplitude controllable disturbances downstream was studied. Experiments were carried out on a swept wing model with a lenticular profile at a zero attack angle. The swept angle of a leading edge was 40°. Wave parameters of moving disturbances were determined. In frames of the linear theory and an approach of the local self-similar mean flow the stability of a compressible three-dimensional boundary layer is studied. Good agreement of the theory with experimental results for transversal scales of unstable vertices of the secondary flow was obtained. However the calculated amplification rates differ from measured values considerably. This disagreement is explained by the nonlinear processes observed in experiment

The structure of a highly decelerated axisymmetric turbulent boundary layer

2021 ◽  
Vol 929 ◽  
Author(s):  
N. Agastya Balantrapu ◽  
Christopher Hickling ◽  
W. Nathan Alexander ◽  
William Devenport
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Shear Layer ◽  
Layer Thickness ◽  
Boundary Layer Thickness ◽  
Large Scale ◽  
Convection Velocity ◽  
Mean Flow ◽  
Mean Velocity ◽  
The Mean

Experiments were performed over a body of revolution at a length-based Reynolds number of 1.9 million. While the lateral curvature parameters are moderate ( $\delta /r_s < 2, r_s^+>500$ , where $\delta$ is the boundary layer thickness and r s is the radius of curvature), the pressure gradient is increasingly adverse ( $\beta _{C} \in [5 \text {--} 18]$ where $\beta_{C}$ is Clauser’s pressure gradient parameter), representative of vehicle-relevant conditions. The mean flow in the outer regions of this fully attached boundary layer displays some properties of a free-shear layer, with the mean-velocity and turbulence intensity profiles attaining self-similarity with the ‘embedded shear layer’ scaling (Schatzman & Thomas, J. Fluid Mech., vol. 815, 2017, pp. 592–642). Spectral analysis of the streamwise turbulence revealed that, as the mean flow decelerates, the large-scale motions energize across the boundary layer, growing proportionally with the boundary layer thickness. When scaled with the shear layer parameters, the distribution of the energy in the low-frequency region is approximately self-similar, emphasizing the role of the embedded shear layer in the large-scale motions. The correlation structure of the boundary layer is discussed at length to supply information towards the development of turbulence and aeroacoustic models. One major finding is that the estimation of integral turbulence length scales from single-point measurements, via Taylor's hypothesis, requires significant corrections to the convection velocity in the inner 50 % of the boundary layer. The apparent convection velocity (estimated from the ratio of integral length scale to the time scale), is approximately 40 % greater than the local mean velocity, suggesting the turbulence is convected much faster than previously thought. Closer to the wall even higher corrections are required.

Velocity and Pressure Measurements Along a Row of Confined Cylinders

2006 ◽  
Author(s):  
Barton L. Smith ◽  
Jack J. Stepan ◽  
Donald M. McEligot
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Symmetry Plane ◽  
Field Measurements ◽  
Reynolds Numbers ◽  
Boundary Layer Transition ◽  
Recirculation Region ◽  
Pressure Measurements ◽  
Layer Transition ◽  
Flow Experiments

The results of flow experiments performed in a cylinder array designed to mimic a VHTR Nuclear Plant lower plenum design are presented. Pressure drop and velocity field measurements were made. Based on these measurements, five regimes of behavior are identified that are found to depend on Reynolds number. It is found that the recirculation region behind the cylinders is shorter than that of half cylinders placed on the wall representing the symmetry plane. Unlike a single cylinder, the separation point is found to always be on the rear of the cylinders, even at very low Reynolds number. Boundary layer transition is found to occur at much lower Reynolds numbers than previously reported.

Measurements of Secondary Flows Downstream of a Turbine Cascade at Off-Design Incidence

2004 ◽  
Author(s):  
M. W. Benner ◽  
S. A. Sjolander ◽  
S. H. Moustapha
Keyword(s):  
Large Scale ◽  
Prediction Method ◽  
Leading Edge ◽  
Field Measurements ◽  
Inviscid Flow ◽  
Secondary Flows ◽  
Subsequent Paper ◽  
Empirical Prediction ◽  
Pressure Distributions ◽  
Passage Vortex

This paper presents experimental results of the secondary flows from two large-scale, low-speed, linear turbine cascades for which the incidence was varied. The aerofoils for the two cascades were designed for the same inlet and outlet conditions and differed mainly in their leading-edge geometries. Detailed flow field measurements were made upstream and downstream of the cascades and static pressure distributions were measured on the blade surfaces for three different values of incidence: 0, +10 and +20 degrees. The results from this experiment indicate that the strength of the passage vortex does not continue to increase with incidence, as would be expected from inviscid flow theory. The streamwise acceleration within the aerofoil passage seems to play an important role in influencing the strength of the vortex. The most recent off-design secondary loss correlation (Moustapha et al. [1]) includes leading-edge diameter as an influential correlating parameter. The correlation predicts that the secondary losses for the aerofoil with the larger leading-edge diameter are lower at off-design incidence; however, the opposite is observed experimentally. The loss results at high positive incidence have also high-lighted some serious shortcomings with the conventional method of loss decomposition. An empirical prediction method for secondary losses has been developed and will be presented in a subsequent paper.

Trailing-edge stall

1970 ◽  
Vol 42 (3) ◽  
pp. 561-584 ◽  
Author(s):  
S. N. Brown ◽  
K. Stewartson
Keyword(s):  
Boundary Layer ◽  
Complete Solution ◽  
Leading Edge ◽  
Asymptotic Solutions ◽  
Angle Of Incidence ◽  
Trailing Edge ◽  
Boundary Layer Equations ◽  
Order Of Magnitude ◽  
Stall Angle ◽  
Fundamental Equations

A study is made of the laminar flow in the neighbourhood of the trailing edge of an aerofoil at incidence. The aerofoil is replaced by a flat plate on the assumption that leading-edge stall has not taken place. It is shown that the critical order of magnitude of the angle of incidence α* for the occurrence of separation on one side of the plate is$\alpha^{*} = O(R^{\frac{1}{16}})$, whereRis a representative Reynolds number, for incompressible flow, and α* =O(R−¼) for supersonic flow. The structure of the flow is determined by the incompressible boundary-layer equations but with unconventional boundary conditions. The complete solution of these fundamental equations requires a numerical investigation of considerable complexity which has not been undertaken. The only solutions available are asymptotic solutions valid at distances from the trailing edge that are large in terms of the scaled variable of orderR−⅜, and a linearized solution for the boundary layer over the plate which gives the antisymmetric properties of the aerofoil at incidence. The value of α* for which separation occurs is the trailing-edge stall angle and an estimate is obtained from the asymptotic solutions. The linearized solution yields an estimate for the viscous correction to the circulation determined by the Kutta condition.

Computational Design and Experimental Evaluation of Using a Leading Edge Fillet on a Gas Turbine Vane

2001 ◽  
Author(s):  
G. A. Zess ◽  
K. A. Thole
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Gradient ◽  
Gas Turbine ◽  
Total Pressure ◽  
Guide Vane ◽  
Leading Edge ◽  
Field Measurements ◽  
Computational Design ◽  
Horseshoe Vortex

With the desire for increased power output for a gas turbine engine comes the continual push to achieve higher turbine inlet temperatures. Higher temperatures result in large thermal and mechanical stresses particularly along the nozzle guide vane. One critical region along a vane is the leading edge-endwall juncture. Based on the assumption that the approaching flow to this juncture is similar to a two-dimensional boundary layer, previous studies have shown that a horseshoe vortex forms. This vortex forms because of a radial total pressure gradient from the approaching boundary layer. This paper documents the computational design and experimental validation of a fillet placed at the leading edge-endwall juncture of a guide vane to eliminate the horseshoe vortex. The fillet design effectively accelerated the incoming boundary layer thereby mitigating the effect of the total pressure gradient. To verify the CFD studies used to design the leading edge fillet, flow field measurements were performed in a large-scale, linear, vane cascade. The flow field measurements were performed with a laser Doppler velocimeter in four planes orientated orthogonal to the vane. Good agreement between the CFD predictions and the experimental measurements verified the effectiveness of the leading edge fillet at eliminating the horseshoe vortex. The flowfield results showed that the turbulent kinetic energy levels were significantly reduced in the endwall region because of the absence of the unsteady horseshoe vortex.

scholarly journals Numerical Analysis of an Instrumented Turbine Blade Cascade

2018 ◽  
Author(s):  
Bryn N. Ubald ◽  
Jiahuan Cui ◽  
Rob Watson ◽  
Paul G. Tucker ◽  
Shahrokh Shahpar
Keyword(s):  
Numerical Analysis ◽  
Measurement Accuracy ◽  
Free Stream ◽  
Separation Bubble ◽  
Leading Edge ◽  
Pressure Probe ◽  
Mean Flow ◽  
Trailing Edge ◽  
Free Stream Turbulence ◽  
Jet Trajectory

The measurement accuracy of the temperature/pressure probe mounted at the leading edge of a turbine/compressor blade is crucial for estimating the fuel consumption of a turbo-fan engine. Apart from the measurement error itself, the probe also introduces extra losses. This again would compromise the measurement accuracy of the overall engine efficiency. This paper utilizes high-fidelity numerical analysis to understand the complex flow around the probe and quantify the loss sources due to the interaction between the blade and its instrumentation. With the inclusion of leading edge probes, three dimensional flow phenomena develop, with some flow features acting in a similar manner to a jet in cross flow. The separated flow formed at the leading edge of the probe blocks a large area of the probe bleed-hole, which is one of the reasons why the probe accuracy can be sensitive to Mach and Reynolds numbers. The addition of 4% free stream turbulence is shown to have a marginal impact on the jet trajectory originated from the probe bleedhole. However, a slight reduction is observed in the size of the separation bubble formed at the leading edge of the probe, preceding the two bleedhole exits. The free stream turbulence also has a significant impact on the size of the separation bubble near the trailing edge of the blade. With the addition of the free stream turbulence, the loss observed within the trailing edge wake is reduced. More than 50% of the losses at the cascade exit are generated by the leading edge probe. A breakdown of the dissipation terms from the mean flow kinetic energy equation demonstrates that the Reynolds stresses are the key terms in dissipating the counter rotating vortex pairs with the viscous stresses responsible for the boundary layer.

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