Direct Simulations of Wake-Perturbed Separated Boundary Layers

2011 ◽  
Author(s):  
Ayse G. Gungor ◽  
Mark P. Simens ◽  
Javier Jime´nez
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Separation Bubble ◽  
Turbine Blades ◽  
Plate Boundary ◽  
Maximum Velocity ◽  
Adverse Pressure Gradient ◽  
Suction Side ◽  
Single Experiment ◽  
The Stability

A wake-perturbed flat plate boundary layer with a stream-wise pressure distribution similar to those encountered on the suction side of typical low-pressure turbine blades is investigated by direct numerical simulation. The laminar boundary layer separates due to a strong adverse pressure gradient induced by suction along the upper simulation boundary, transitions and reattaches while still subject to the adverse pressure gradient. Various simulations are performed with different wake passing frequencies, corresponding to the Strouhal number 0.0043 < fθb/ΔU < 0.0496 and wake profiles. The wake profile is changed by varying its maximum velocity defect and its symmetry. Results indicate that the separation and reattachment points, as well as the subsequent boundary layer development, are mainly affected by the frequency, but that the wake shape and intensity have little effect. Moreover, the effect of the different frequencies can be predicted from a single experiment in which the separation bubble is allowed to reform after having been reduced by wake perturbations. The stability characteristics of the mean flows resulting from the forcing at different frequencies are evaluated in terms of local linear stability analysis based on the Orr-Sommerfeld equation.

Direct Numerical Simulations of Wake-Perturbed Separated Boundary Layers

2012 ◽  
Vol 134 (6) ◽  
Author(s):  
Ayse G. Gungor ◽  
Mark P. Simens ◽  
Javier Jiménez
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Separation Bubble ◽  
Turbine Blades ◽  
Plate Boundary ◽  
Maximum Velocity ◽  
Adverse Pressure Gradient ◽  
Single Experiment ◽  
Regeneration Time ◽  
Wake Passing

A wake-perturbed flat plate boundary layer with a streamwise pressure distribution similar to those encountered on the suction side of typical low-pressure turbine blades is investigated by direct numerical simulation. The laminar boundary layer separates due to a strong adverse pressure gradient induced by suction along the upper simulation boundary, transitions, and reattaches while still subject to the adverse pressure gradient. Various simulations are performed with different wake passing frequencies, corresponding to the Strouhal number 0.0043< fθb/ΔU <0.0496 and wake profiles. The wake profile is changed by varying its maximum velocity defect and its symmetry. Results indicate that the separation and reattachment points, as well as the subsequent boundary layer development, are mainly affected by the frequency, but that the wake shape and intensity have little effect, and that the forcing is effective as long as the wake-passing period is shorter than the bubble-regeneration time. Moreover, the effect of the different frequencies can be predicted from a single experiment in which the separation bubble is allowed to reform after having been reduced by wake perturbations. The stability characteristics of the mean flows resulting from the forcing at different frequencies are evaluated in terms of local linear stability analysis based on the Orr-Sommerfeld equation.

Turbine Cascade Flow Control Using a “Wake Filling” Pulsed Plasma Actuator

2005 ◽  
Author(s):  
H. Perez-Blanco ◽  
Robert Van Dyken ◽  
Aaron Byerley ◽  
Tom McLaughlin
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Laminar Boundary Layer ◽  
Separation Bubble ◽  
Adverse Pressure Gradient ◽  
Suction Side ◽  
Plasma Actuator ◽  
Pulsed Plasma ◽  
Power Cost ◽  
Hot Film

Separation bubbles in high-camber blades under part-load conditions have been addressed via continuous and pulsed jets, and also via plasma actuators. Numerous passive techniques have been employed as well. In this type of blades, the laminar boundary layer cannot overcome the adverse pressure gradient arising along the suction side, resulting on a separation bubble. When separation is abated, a common explanation is that kinetic energy added to the laminar boundary layer speeds up its transition to turbulent. In the present study, a plasma actuator installed in the trailing edge (i.e. “wake filling configuration”) of a cascade blade is used to excite the flow in pulsed and continuous ways. The pulsed excitation can be directed to the frequencies of the large coherent structures (LCS) of the flow, as obtained via a hot-film anemometer, or to much higher frequencies present in the suction-side boundary layer, as given in the literature. It is found that pulsed frequencies much higher than that of LCS reduce losses and improve turning angles further than frequencies close to those of LCS. With the plasma actuator 50% on time, good loss abatement is obtained. Larger “on time” values yield improvements, but with decreasing returns. Continuous high-frequency activation results in the largest loss reduction, at increased power cost. The effectiveness of high frequencies may be due to separation abatement via boundary layer excitation into transition, or may simply be due to the creation of a favorable pressure gradient that averts separation as the actuator ejects fluid downstream. Both possibilities are discussed in light of the experimental evidence.

Transition of streamwise streaks in zero-pressure-gradient boundary layers

2002 ◽  
Vol 472 ◽  
pp. 229-261 ◽  
Author(s):  
LUCA BRANDT ◽  
DAN S. HENNINGSON
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
High Speed ◽  
Plate Boundary ◽  
Transition Process ◽  
Free Stream Turbulence ◽  
Tollmien Schlichting ◽  
Streamwise Streaks ◽  
The Stability ◽  
Optimal Disturbances

A transition scenario initiated by streamwise low- and high-speed streaks in a flat-plate boundary layer is studied. In many shear flows, the perturbations that show the highest potential for transient energy amplification consist of streamwise-aligned vortices. Due to the lift-up mechanism these optimal disturbances lead to elongated streamwise streaks downstream, with significant spanwise modulation. In a previous investigation (Andersson et al. 2001), the stability of these streaks in a zero-pressure-gradient boundary layer was studied by means of Floquet theory and numerical simulations. The sinuous instability mode was found to be the most dangerous disturbance. We present here the first simulation of the breakdown to turbulence originating from the sinuous instability of streamwise streaks. The main structures observed during the transition process consist of elongated quasi-streamwise vortices located on the flanks of the low-speed streak. Vortices of alternating sign are overlapping in the streamwise direction in a staggered pattern. The present scenario is compared with transition initiated by Tollmien–Schlichting waves and their secondary instability and by-pass transition initiated by a pair of oblique waves. The relevance of this scenario to transition induced by free-stream turbulence is also discussed.

scholarly journals Turbulent Spot Development Under a Moderate Adverse Pressure Gradient

1993 ◽  
Author(s):  
J. P. Gostelow ◽  
G. Hong ◽  
N. Melwani ◽  
G. J. Walker
Keyword(s):  
Boundary Layer ◽  
Wave Packet ◽  
Pressure Gradient ◽  
Plate Boundary ◽  
Adverse Pressure Gradient ◽  
Spreading Rate ◽  
Lateral Spreading ◽  
Data Sampling ◽  
Turbulent Spot ◽  
Included Angle

Triggered turbulent spots are under investigation in a wind tunnel. A turbulent spot was initiated in a flat plate boundary layer under a moderate adverse pressure gradient. The spot was traversed at four streamwise locations using conventional hot-wire anemometry techniques. A triggering jet provided a phase reference for data sampling. Phase-averaged velocity traces, boundary layer integral properties and contours of velocity perturbation and disturbance level are presented. The central region resembles a zero pressure gradient spot but much of the span is dominated by the different behavioral stages of a stongly-amplified wave packet. The celerities of the spot leading and trailing edges under an adverse pressure gradient are significantly lower than those associated with zero pressure gradient spots but the lateral spreading rate is much higher. This combination of turbulent spot and wave packet, which spreads at an included angle as high as 60° is quite different from the well-documented zero pressure gradient spot spreading at an included angle of about 20°. Improvements in transition region predictions are therefore dependent on further detailed measurements of spots and wave packets under adverse pressure gradients.

scholarly journals Numerical Analysis of High-Altitude Inlet Air on Boundary Layer Flow Loss in an Aero-Engine Compressor

Energies ◽  
2020 ◽  
Vol 13 (16) ◽  
pp. 4155
Author(s):  
Feng Wu ◽  
Limin Gao ◽  
Lu Yang ◽  
Aqiang Lin ◽  
Hai Zhang
Keyword(s):  
Boundary Layer ◽  
Numerical Analysis ◽  
Mach Number ◽  
Pressure Gradient ◽  
High Altitude ◽  
Similarity Criterion ◽  
Adverse Pressure Gradient ◽  
Suction Side ◽  
Entropy Increase ◽  
Wall Boundary

A numerical analysis is performed to explore the high altitude and high Mach flight on the effect of wall boundary layer loss in the compressor. The accuracy for solution results by the application of the similarity criterion and parameter definition of the air inlet is compared with the existing experimental test result. The results indicate that the radial adverse pressure gradient in the rotor domain gradually increases along the span direction and decreases as flight Mach number increases; meanwhile, the circumferential adverse pressure gradient on the pressure side of the rotor blade is correspondingly larger and less than that on the suction side. In particular, the entropy increase along the streamwise shows a decreasing trend and an increasing trend inside the hub and shroud wall boundary layers, respectively. At 2.1 Ma, the entropy increase in the rotor domains enhances by 24.36–27.80% inside the shroud boundary layer, relative to the hub boundary layer; however, it decreases by 0.97–8.54% in the stator domain. With the increase in flight Mach number from 2.1 to 3.4, the average entropy increase reductions in the rotor domain decrease by 18.99–24.97% within the hub boundary layer and 5.71–8.1% within the shroud boundary layer. In the stator domain, it drops by 18.45–9.03% inside the hub boundary layer and 6.88–8.67% inside the shroud boundary layer. It was therefore found that, as Mach number increases from 2.1 to 3.4, the entropy increase reduction is larger inside the hub boundary layer than inside the shroud boundary layer.

scholarly journals Modeling and Prediction of the Laminar Leading-Edge Separation and Transition in a Blade-Cascade Flow

1996 ◽  
Author(s):  
Dimitri P. Tselepidakis ◽  
Sung-Eun Kim
Keyword(s):  
Boundary Layer ◽  
Separation Bubble ◽  
Plate Boundary ◽  
Leading Edge ◽  
Suction Side ◽  
Navier Stokes ◽  
Compressor Cascade ◽  
Real Flow ◽  
Side Flow ◽  
Edge Separation

This paper presents the computation of the flow around a controlled diffusion compressor cascade. Features associated with by-pass transition close to the leading edge — including laminar leading-edge separation — contribute significantly to the evolution of the boundary layer on the blade surface. Previous studies have demonstrated that conventional k-ε models, based on linear or non-linear Boussinesq stress-strain relations, are able to capture by-pass transition in simple shear, but are unable to resolve transitional features in complex strain, like the leading-edge separation bubble, which is of considerable influence to the suction-side flow at high inlet angle. Here, the k-ω turbulence model has been implemented in a nonstaggered, finite-volume based segregated Reynolds-Averaged Navier-Stokes solver. We demonstrate that this model, if properly sensitized to the generation of turbulence by irrotational strains, is capable of capturing the laminar leading-edge separation bubble. The real flow around the leading edge is laminar and the transition is only provoked on the reattachment region. Additional investigation of transition in a flat-plate boundary layer development has also produced reasonably promising results.

Flow Separation Control in an Annular to Conical Diffuser Using Two-Dimensional and Three-Dimensional Wall Steps

2013 ◽  
Vol 135 (4) ◽  
Author(s):  
Kin Pong Lo ◽  
Christopher J. Elkins ◽  
John K. Eaton
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Flow Control ◽  
Separation Bubble ◽  
Adverse Pressure Gradient ◽  
Boundary Layer Separation ◽  
Core Flow ◽  
Layer Separation ◽  
Wall Boundary Layer ◽  
Wall Boundary

Conical diffusers are often installed downstream of a turbomachine with a central hub. Previous studies showed that nonstreamlined hubs had extended separated wakes that reduced the adverse pressure gradient in the diffuser. Active flow control techniques can rapidly close the central separation bubble, but this restores the adverse pressure gradient, which can cause the outer wall boundary layer to separate. The present study focuses on the use of a step-wall diffuser to stabilize the wall boundary layer separation in the presence of core flow control. Three-component mean velocity data for a set of conical diffusers were acquired using magnetic resonance velocimetry. The results showed the step-wall diffuser stabilized the wall boundary layer separation by fixing its location. An axisymmetric step separation bubble was formed. A step with a periodically varying height reduced the reattachment length of the step separation and allowed the diffuser to be shortened. The step-wall diffuser was found to be robust in a range of core flow velocity profiles. The minimum distance between the core flow control mechanism and the step-wall diffuser as well as the minimum length of the step were determined.

The form drag of two-dimensional bluff-plates immersed in turbulent boundary layers

1968 ◽  
Vol 31 (3) ◽  
pp. 547-582 ◽  
Author(s):  
M. C. Good ◽  
P. N. Joubert
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Separation Bubble ◽  
Adverse Pressure Gradient ◽  
Gradient Flows ◽  
Mean Flow ◽  
Plate Height ◽  
Form Drag ◽  
Law Of The Wall ◽  
History Of

Measurements of the distributions of pressure on a bluff flat plate (fence) have been correlated with the characteristics of the smooth-wall boundary layer in which it is immersed. For zero pressure-gradient flows, correlations are obtained for the variation of form drag with plate heighthwhich are analogous in form to the ‘law of the wall’ and the ‘velocity-defect law’ for the boundary-layer velocity profile. The data for adverse pressure-gradient flows is suggestive of a ‘law of the wake’ type correlation. Pressures on the upstream face of the bluff-plate are determined by a wall-similarity law, even forh/δ > 1, and are independent of the pressure-gradient history of the flow; the separation induced upstream is apparently of the Stratford-Townsend type. The effects of the history of the boundary layer are manifested only in the flow in the rear separation bubble, and then only forh/δ > ½. The base pressure is also sensitive to free-stream pressure gradients downstream of the bluff-plate. The relative extent of upstream influence of the bluff-plate on the boundary layer is found to increase rapidly ash/δ decreases. One set of measurements of the mean flow field is also presented.

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

2008 ◽  
Vol 130 (5) ◽  
Author(s):  
Hui Hu ◽  
Zifeng Yang
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Laminar Boundary Layer ◽  
Flow Separation ◽  
Separation Bubble ◽  
Adverse Pressure Gradient ◽  
Laminar Separation Bubble ◽  
Airfoil Surface ◽  
Laminar Separation ◽  
Turbulence Transition

An experimental study was conducted to characterize the transient behavior of laminar flow separation on a NASA low-speed GA (W)-1 airfoil at the chord Reynolds number of 70,000. In addition to measuring the surface pressure distribution around the airfoil, a high-resolution particle image velocimetry (PIV) system was used to make detailed flow field measurements to quantify the evolution of unsteady flow structures around the airfoil at various angles of attack (AOAs). The surface pressure and PIV measurements clearly revealed that the laminar boundary layer would separate from the airfoil surface, as the adverse pressure gradient over the airfoil upper surface became severe at AOA≥8.0deg. The separated laminar boundary layer was found to rapidly transit to turbulence by generating unsteady Kelvin–Helmholtz vortex structures. After turbulence transition, the separated boundary layer was found to reattach to the airfoil surface as a turbulent boundary layer when the adverse pressure gradient was adequate at AOA<12.0deg, resulting in the formation of a laminar separation bubble on the airfoil. The turbulence transition process of the separated laminar boundary layer was found to be accompanied by a significant increase of Reynolds stress in the flow field. The reattached turbulent boundary layer was much more energetic, thus more capable of advancing against an adverse pressure gradient without flow separation, compared to the laminar boundary layer upstream of the laminar separation bubble. The laminar separation bubble formed on the airfoil upper surface was found to move upstream, approaching the airfoil leading edge as the AOA increased. While the total length of the laminar separation bubble was found to be almost unchanged (∼20% of the airfoil chord length), the laminar portion of the separation bubble was found to be slightly stretched, and the turbulent portion became slightly shorter with the increasing AOA. After the formation of the separation bubble on the airfoil, the increase rate of the airfoil lift coefficient was found to considerably degrade, and the airfoil drag coefficient increased much faster with increasing AOA. The separation bubble was found to burst suddenly, causing airfoil stall, when the adverse pressure gradient became too significant at AOA>12.0deg.

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