Unsteady Transition Phenomena at a Compressor Blade Leading Edge

2008 ◽  
Vol 130 (2) ◽  
Author(s):  
Alan D. Henderson ◽  
Gregory J. Walker ◽  
Jeremy D. Hughes
Keyword(s):  
Boundary Layer ◽  
Numerical Solutions ◽  
Three Dimensional ◽  
Leading Edge ◽  
Inviscid Flow ◽  
Surface Boundary ◽  
Bypass Transition ◽  
Suction Surface ◽  
Film Sensor ◽  
Rotor Wake

Wake-induced laminar-turbulent transition is studied at the leading edge of a C4-section compressor stator blade in a 1.5-stage axial compressor. Surface hot-film sensor observations are interpreted with the aid of numerical solutions from UNSFLO, a quasi-three-dimensional viscous-inviscid flow solver. The passage of a rotor wake, with its associated negative jet, over the stator leading edge is observed to have a destabilizing effect on the suction surface boundary layer. This leads to transition closer to the stator leading edge than would have occurred under steady flow conditions. The strength of this phenomenon is influenced by the rotor-stator axial gap and the variability of individual rotor wake disturbances. A variety of transition phenomena is observed near the leading edge in the wake path. Wave packets characteristic of Tollmien-Schlichting waves are observed to amplify and break down into turbulent spots. Disturbances characteristic of the streaky structures occurring in bypass transition are also seen. Examination of suction surface disturbance and wake-induced transitional strip trajectories points to the leading edge as the principal receptivity site for suction surface transition phenomena at design loading conditions. This contrasts markedly with the pressure surface behavior, where transition at design conditions occurs remotely from leading-edge flow perturbations associated with wake chopping. Here, the local receptivity of the boundary layer to the wake passing disturbance and turbulent wake fluid discharging onto the blade surface may be of greater importance.

Unsteady Transition Phenomena at a Compressor Blade Leading Edge

2006 ◽  
Author(s):  
Alan D. Henderson ◽  
Gregory J. Walker ◽  
Jeremy D. Hughes
Keyword(s):  
Boundary Layer ◽  
Numerical Solutions ◽  
Three Dimensional ◽  
Leading Edge ◽  
Inviscid Flow ◽  
Surface Boundary ◽  
Bypass Transition ◽  
Suction Surface ◽  
Film Sensor ◽  
Rotor Wake

Wake-induced laminar-turbulent transition is studied at the leading edge of a C4-section compressor stator blade in a 1.5-stage axial compressor. Surface hot-film sensor observations are interpreted with the aid of numerical solutions from UNSFLO, a quasi three-dimensional viscous-inviscid flow solver. The passage of a rotor wake, with its associated negative jet, over the stator leading edge is observed to have a destabilizing effect on the suction surface boundary layer. This leads to transition closer to the stator leading edge than would have occurred under steady flow conditions. The strength of this phenomenon is influenced by the rotor-stator axial gap and the variability of individual rotor wake disturbances. A variety of transition phenomena are observed near the leading edge in the wake path. Wave packets characteristic of Tollmien–Schlichting waves are observed to amplify and break down into turbulent spots. Disturbances characteristic of the streaky structures occurring in bypass transition are also seen. Examination of suction surface disturbance and wake-induced transitional strip trajectories points to the leading edge as the principal receptivity site for suction surface transition phenomena at design loading conditions. This contrasts markedly with the pressure surface behavior, where transition at design conditions occurs remote from leading edge flow perturbations associated with wake chopping. Here the local receptivity of the boundary layer to the wake passing disturbance and turbulent wake fluid discharging onto the blade surface may be of greater importance.

Secondary instability and subcritical transition of the leading-edge boundary layer

2016 ◽  
Vol 792 ◽  
pp. 682-711 ◽  
Author(s):  
Michael O. John ◽  
Dominik Obrist ◽  
Leonhard Kleiser
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
High Speed ◽  
Three Dimensional ◽  
Leading Edge ◽  
Reynolds Numbers ◽  
Secondary Instability ◽  
Bypass Transition ◽  
Wall Suction ◽  
Transition Mechanism

The leading-edge boundary layer (LEBL) in the front part of swept airplane wings is prone to three-dimensional subcritical instability, which may lead to bypass transition. The resulting increase of airplane drag and fuel consumption implies a negative environmental impact. In the present paper, we present a temporal biglobal secondary stability analysis (SSA) and direct numerical simulations (DNS) of this flow to investigate a subcritical transition mechanism. The LEBL is modelled by the swept Hiemenz boundary layer (SHBL), with and without wall suction. We introduce a pair of steady, counter-rotating, streamwise vortices next to the attachment line as a generic primary disturbance. This generates a high-speed streak, which evolves slowly in the streamwise direction. The SSA predicts that this flow is unstable to secondary, time-dependent perturbations. We report the upper branch of the secondary neutral curve and describe numerous eigenmodes located inside the shear layers surrounding the primary high-speed streak and the vortices. We find secondary flow instability at Reynolds numbers as low as$Re\approx 175$, i.e. far below the linear critical Reynolds number$Re_{crit}\approx 583$of the SHBL. This secondary modal instability is confirmed by our three-dimensional DNS. Furthermore, these simulations show that the modes may grow until nonlinear processes lead to breakdown to turbulent flow for Reynolds numbers above$Re_{tr}\approx 250$. The three-dimensional mode shapes, growth rates, and the frequency dependence of the secondary eigenmodes found by SSA and the DNS results are in close agreement with each other. The transition Reynolds number$Re_{tr}\approx 250$at zero suction and its increase with wall suction closely coincide with experimental and numerical results from the literature. We conclude that the secondary instability and the transition scenario presented in this paper may serve as a possible explanation for the well-known subcritical transition observed in the leading-edge boundary layer.

Measurements and Prediction of the Effects of Surface Roughness on Profile Losses and Deviation in a Turbine Cascade

1998 ◽  
Vol 120 (1) ◽  
pp. 20-27 ◽  
Author(s):  
R. J. Kind ◽  
P. J. Serjak ◽  
M. W. P. Abbott
Keyword(s):  
Surface Roughness ◽  
Three Dimensional ◽  
Leading Edge ◽  
Computational Method ◽  
Roughness Height ◽  
Surface Boundary ◽  
Turbine Cascade ◽  
Suction Surface ◽  
Roughness Effects ◽  
Profile Losses

Measurements of pressure distributions, profile losses, and flow deviation were carried out on a planar turbine cascade in incompressible flow to assess the effects of partial roughness coverage of the blade surfaces. Spanwise-oriented bands of roughness were placed at various locations on the suction and pressure surfaces of the blades. Roughness height, spacing between roughness elements, and band width were varied. A computational method based on the inviscid/viscous interaction approach was also developed; its predictions were in good agreement with the experimental results. This indicates that good predictions can be expected for a variety of cascade and roughness configurations from any two-dimensional analysis that couples an inviscid method with a suitable rough surface boundary-layer analysis. The work also suggests that incorporation of the rough wall skin-friction law into a three-dimensional Navier–Stokes code would enable good predictions of roughness effects in three-dimensional situations. Roughness was found to have little effect on static pressure distribution around the blades and on deviation angle, provided that it does not precipitate substantial flow separation. Roughness on the suction surface can cause large increases in profile losses; roughness height and location of the leading edge of the roughness band are particularly important. Loss increments due to pressure-surface roughness are much smaller than those due to similar roughness on the suction surface.

scholarly journals Direct numerical simulations of transition in a compressor cascade: the influence of free-stream turbulence

2010 ◽  
Vol 665 ◽  
pp. 57-98 ◽  
Author(s):  
TAMER A. ZAKI ◽  
JAN G. WISSINK ◽  
WOLFGANG RODI ◽  
PAUL A. DURBIN
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Free Stream ◽  
Transition To Turbulence ◽  
Surface Boundary ◽  
Bypass Transition ◽  
Suction Surface ◽  
Surface Boundary Layer ◽  
Laminar Separation ◽  
Free Stream Turbulence

The flow through a compressor passage without and with incoming free-stream grid turbulence is simulated. At moderate Reynolds number, laminar-to-turbulence transition can take place on both sides of the aerofoil, but proceeds in distinctly different manners. The direct numerical simulations (DNS) of this flow reveal the mechanics of breakdown to turbulence on both surfaces of the blade. The pressure surface boundary layer undergoes laminar separation in the absence of free-stream disturbances. When exposed to free-stream forcing, the boundary layer remains attached due to transition to turbulence upstream of the laminar separation point. Three types of breakdowns are observed; they combine characteristics of natural and bypass transition. In particular, instability waves, which trace back to discrete modes of the base flow, can be observed, but their development is not independent of the Klebanoff distortions that are caused by free-stream turbulent forcing. At a higher turbulence intensity, the transition mechanism shifts to a purely bypass scenario. Unlike the pressure side, the suction surface boundary layer separates independent of the free-stream condition, be it laminar or a moderate free-stream turbulence of intensityTu~ 3%. Upstream of the separation, the amplification of the Klebanoff distortions is suppressed in the favourable pressure gradient (FPG) region. This suppression is in agreement with simulations of constant pressure gradient boundary layers. FPG is normally stabilizing with respect to bypass transition to turbulence, but is, thereby, unfavourable with respect to separation. Downstream of the FPG section, a strong adverse pressure gradient (APG) on the suction surface of the blade causes the laminar boundary layer to separate. The separation surface is modulated in the instantaneous fields of the Klebanoff distortion inside the shear layer, which consists of forward and backward jet-like perturbations. Separation is followed by breakdown to turbulence and reattachment. As the free-stream turbulence intensity is increased,Tu~ 6.5%, transitional turbulent patches are initiated, and interact with the downstream separated flow, causing local attachment. The calming effect, or delayed re-establishment of the boundary layer separation, is observed in the wake of the turbulent events.

Control of separations in a highly loaded diffusion cascade by tailored boundary layer suction

2013 ◽  
Vol 228 (8) ◽  
pp. 1363-1374 ◽  
Author(s):  
Zhiyuan Cao ◽  
Bo Liu ◽  
Ting Zhang
Keyword(s):  
Boundary Layer ◽  
Static Pressure ◽  
Three Dimensional ◽  
Flow Fields ◽  
Separated Flows ◽  
Surface Boundary ◽  
Suction Surface ◽  
Surface Boundary Layer ◽  
Boundary Layer Suction ◽  
Highly Loaded

In order to explore the control mechanism of boundary layer suction on the separated flows of highly loaded diffusion cascades, a linear compressor cascade, which has separated flows on the whole span and three-dimensional separations over the suction surface/endwall corner, was investigated by tailored boundary layer suction. Three suction surface-slotted schemes and two combined suction surface/endwall-slotted schemes were designed. The original cascade and the cascade with part blade span suction were experimentally investigated on a high-subsonic cascade wind tunnel. In addition, numerical simulation was employed to study the flow fields of different suction schemes in detail. The results shows that while tailored boundary layer suction at part blade span can effectively remove the separations at the suction span, the flow fields of other spans deteriorated. The reasons are the ‘C’ shape or reverse ‘C’ shape spanwise distribution of static pressure after part blade span boundary layer suction. Suction surface boundary layer suction over the whole span can obviously eliminate the separation at the suction surface. However, because of the endwall boundary layer, suction surface boundary layer suction cannot effectively remove the corner three-dimensional separation. The separation over the whole span and the three-dimensional separation at the corner are completely eliminated by combined suction surface/endwall boundary layer suction. After combined boundary layer suction, the static pressure distribution over the blade span just like the shape of ‘C’ is good for the transport of the low-energy fluid near the endwall to the midspan.

Numerical Investigation of Three-Dimensional Separation in an Axial Flow Compressor: The Influence of Free-Stream Turbulence Intensity and Endwall Boundary Layer State

2016 ◽  
Author(s):  
Ashley D. Scillitoe ◽  
Paul G. Tucker ◽  
Paolo Adami
Keyword(s):  
Boundary Layer ◽  
Turbulence Intensity ◽  
Free Stream ◽  
Separation Bubble ◽  
Three Dimensional ◽  
Boundary Layer Transition ◽  
Surface Boundary ◽  
Suction Surface ◽  
Corner Separation ◽  
Free Stream Turbulence

Regions of three-dimensional separations are an inherent flow feature of the suction surface - endwall corner in axial compressors. These corner separations can cause a significant total pressure loss and reduce the compressor’s efficiency. This paper uses wall-resolved LES to investigate the loss sources in a corner separation, and examines the influence of the inflow turbulence on these sources. Different subgrid scale (SGS) models are tested and the choice of model is found to be important. The σ SGS model, which performed well, is then used to perform LES of a compressor endwall flow. The time-averaged data is in good agreement with measurements. The viscous and turbulent dissipation are used to highlight the sources of loss, with the latter being dominant. The key loss sources are seen to be the 2D laminar separation bubble and trailing edge wake, and the 3D flow region near the endwall. Increasing the free-stream turbulence intensity (FST) changes the suction surface boundary layer transition mode from separation induced to bypass. However, it doesn’t significantly alter the transition location and therefore the corner separation size. Additionally, the FST doesn’t noticeably interact with the corner separation itself, meaning that in this case the corner separation is relatively insensitive to the FST. The endwall boundary layer state is found to be significant. A laminar endwall boundary layer separates much earlier leading to a larger passage vortex. This significantly alters the endwall flow and loss. Hence, the need for accurate boundary measurements is clear.

Low Pressure Turbine Relaminarization Bubble Characterization using Massively-Parallel Large Eddy Simulations

2012 ◽  
Vol 134 (2) ◽  
Author(s):  
Shriram Jagannathan ◽  
Markus Schwänen ◽  
Andrew Duggleby
Keyword(s):  
Boundary Layer ◽  
Reverse Flow ◽  
Leading Edge ◽  
Suction Side ◽  
Low Pressure ◽  
Surface Boundary ◽  
Suction Surface ◽  
Hairpin Vortices ◽  
Low Pressure Turbine ◽  
Large Eddy

The separation and reattachment of suction surface boundary layer in a low pressure turbine is characterized using large-eddy simulation at Ress = 69000 based on inlet velocity and suction surface length. Favorable comparisons are drawn with experiments using a high pass filtered Smagorinsky model for sub-grid scales. The onset of time mean separation is at s/so = 0.61 and reattachment at s/so = 0.81, extending over 20% of the suction surface. The boundary layer is convectively unstable with a maximum reverse flow velocity of about 13% of freestream. The breakdown to turbulence occurs over a very short distance of suction surface and is followed by reattachment. Turbulence near the bubble is further characterized using anisotropy invariant mapping and time orthogonal decomposition diagnostics. Particularly the vortex shedding and shear layer flapping phenomena are addressed. On the suction side, dominant hairpin structures near the transitional and turbulent flow regime are observed. The hairpin vortices are carried by the freestream even downstream of the trailing edge of the blade with a possibility of reaching the next stage. Longitudinal streaks that evolve from the breakdown of hairpin vortices formed near the leading edge are observed on the pressure surface.

Three-Dimensional Flow Field in the Tip Region of a Compressor Rotor Passage—Part I: Mean Velocity Profiles and Annulus Wall Boundary Layer

1982 ◽  
Vol 104 (4) ◽  
pp. 760-771 ◽  
Author(s):  
B. Lakshminarayana ◽  
M. Pouagare ◽  
R. Davino
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Three Dimensional ◽  
Leading Edge ◽  
Leakage Flow ◽  
Suction Surface ◽  
Mean Velocity ◽  
Wall Boundary Layer ◽  
Compressor Rotor ◽  
Wall Boundary

The flow field in the annulus wall and tip region of a compressor rotor was measured using a triaxial, hot-wire probe rotating with the rotor. The flow was surveyed across the entire passage at five axial locations (leading edge, 1/4 chord, 1/2 chord, 3/4 chord, and trailing edge locations) and at six radial locations inside the passage. The data derived include all three components of mean velocity. Blade-to-blade variations of the velocity components, pitch and yaw angles, as well as the passage-averaged mean properties of the annulus wall boundary layer, are derived from this data. The measurements indicate that the leakage flow starts beyond a quarter-chord and tends to roll up farther away from the suction surface than that observed in cascades. Substantial velocity deficiencies and radial inward velocities are observed in this region. The annulus wall boundary layer is well behaved up to half a chord, beyond which interactions with the leakage flow produce complex profiles.

scholarly journals Experimental Investigation on Spatial Development of Streamwise Vortices in a Turbine Inlet Guide Vane Cascade

1985 ◽  
Author(s):  
Toyotaka Sonoda
Keyword(s):  
Boundary Layer ◽  
Guide Vane ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Spatial Development ◽  
Inlet Guide Vane ◽  
Surface Boundary ◽  
Streamwise Vortices ◽  
Suction Surface ◽  
Turbine Inlet

In order to obtain a better understanding of secondary flow in a turbine cascade, spatial development of a leading-edge horseshoe vortex has been investigated experimentally in a large-scale, low-speed, high-accelerated, plane turbine inlet guide vane cascade. Flow has been visualized by issuing kerosene vapor into the inlet boundary layer and the vane suction surface boundary layer, respectively. Based on many cross-sectional photographs normal to the flow and supplemental measurements of the wall static pressure on the vane and the endwall, the evolution of a leading-edge horseshoe vortex into streamwise vortices and the generation of a new type streamwise vortex pair on the suction surface near the endwall are discussed.

Large Eddy Simulation of Tandem Blade Stator Cascades

2017 ◽  
Author(s):  
Pratik Mitra ◽  
Jahnavi Kantharaju ◽  
Rohan Rayan ◽  
Joseph Mathew
Keyword(s):  
Boundary Layer ◽  
Leading Edge ◽  
Immersed Boundary ◽  
Surface Boundary ◽  
Suction Surface ◽  
Surface Boundary Layer ◽  
Substantial Impact ◽  
Eddy Simulation ◽  
Large Eddy ◽  
Rapid Transition

Large eddy simulations of tandem blade compressor cascades have been performed with an explicit filtering method. A low speed case was simulated using the public domain code Incompact3d which solves incompressible flow with an immersed boundary method for embedded solid bodies, obviating the effort expended on preparing good quality meshes around blading. The LES successfully captures transition on the front blade and yields a significantly different loading compared with RANS solutions obtained before. The less substantial impact on the rear blade is traced to rapid transition forced by the turbulent wake of the front blade. LES with a refined grid was found to shorten the transition width due to the crucial role of small scales during transition. A complementary study with an in-house compressible LES solver was conducted for a transonic tandem cascade at the inlet Mach number of 0.89. Flow expands around the leading edge of the front blade and is terminated by a shock which interacts with the suction surface boundary layer. The beneficial effect of tandem blading was found to be achieved by limiting this separation. The shock-induced separation also marks a rapid transition of the suction surface boundary layer that is readily captured in the LES, showing pre-transitional streaks, but could prove difficult even for current transition-sensitive RANS.

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