Numerical Predictions of Turbine Cascade Secondary Flows and Heat Transfer With Inflow Turbulence

2020 ◽  
Author(s):  
Yousef Kanani ◽  
Sumanta Acharya ◽  
Forrest Ames
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Leading Edge ◽  
Suction Side ◽  
Lateral Spreading ◽  
Secondary Flows ◽  
Pressure Side ◽  
Initial Flow ◽  
Numerical Predictions ◽  
Inflow Turbulence

Abstract Turbine passage secondary flows are studied for a large rounded leading edge airfoil geometry considered in the experimental investigation of Varty et al. (J. Turbomach. 140(2):021010) using high resolution Large Eddy Simulation (LES). The complex nature of secondary flow formation and evolution are affected by the approach boundary layer characteristics, components of pressure gradients tangent and normal to the passage flow, surface curvature, and inflow turbulence. This paper presents a detailed description of the secondary flows and heat transfer in a linear vane cascade at exit chord Reynolds number of 5 × 105 at low and high inflow turbulence. Initial flow turning at the leading edge of the inlet boundary layer leads to a pair of counter-rotating flow circulation in each half of the cross-plane that drive the evolution of the pressure-side and suction side of the near-wall vortices such as the horseshoe and leading edge corner vortex. The passage vortex for the current large leading-edge vane is formed by the amplification of the initially formed circulation closer to the pressure side (PPC) which strengthens and merges with other vortex systems while moving toward the suction side. The predicted suction surface heat transfer shows good agreement with the measurements and properly captures the augmented heat transfer due to the formation and lateral spreading of the secondary flows towards the vane midspan downstream of the vane passage. Effects of various components of the secondary flows on the endwall and vane heat transfer are discussed in detail.

Numerical Predictions of Turbine Cascade Secondary Flows and Heat Transfer With Inflow Turbulence

2021 ◽  
pp. 1-34
Author(s):  
Yousef Kanani ◽  
Sumanta Acharya ◽  
Forrest Ames
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Leading Edge ◽  
Suction Side ◽  
Lateral Spreading ◽  
Secondary Flows ◽  
Pressure Side ◽  
Initial Flow ◽  
Numerical Predictions ◽  
Inflow Turbulence

Abstract Turbine passage secondary flows are studied for a large rounded leading edge airfoil geometry considered in the experimental investigation of Varty et al. (J. Turbomach. 140(2):021010) using high resolution Large Eddy Simulation. The complex nature of secondary flow formation and evolution are affected by the approach boundary layer characteristics, components of pressure gradients tangent and normal to the passage flow, surface curvature, and inflow turbulence. This paper presents a detailed description of the secondary flows and heat transfer in a linear vane cascade at exit chord Reynolds number of 500,000 at low and high inflow turbulence. Initial flow turning at the leading edge of the inlet boundary layer leads to a pair of counter-rotating flow circulation in each half of the cross-plane that drive the evolution of the pressure-side and suction side of the near-wall vortices such as the horseshoe and leading edge corner vortex. The passage vortex for the current large leading-edge vane is formed by the amplification of the initially formed circulation closer to the pressure side which strengthens and merges with other vortex systems while moving toward the suction side. The predicted suction surface heat transfer shows good agreement with the measurements and properly captures the augmented heat transfer due to the formation and lateral spreading of the secondary flows towards the vane midspan downstream of the vane passage. Effects of various components of the secondary flows on the endwall and vane heat transfer are discussed in detail.

scholarly journals Effects of Tip Clearance Gap and Exit Mach Number on Turbine Blade Tip and Near-Tip Heat Transfer

2013 ◽  
Author(s):  
K. Anto ◽  
S. Xue ◽  
W. F. Ng ◽  
L. J. Zhang ◽  
H. K. Moon
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Turbine Blade ◽  
Leading Edge ◽  
Suction Side ◽  
Tip Clearance ◽  
Pressure Side ◽  
Engine Blade ◽  
Blade Tip ◽  
Exit Mach Number

This study focuses on local heat transfer characteristics on the tip and near-tip regions of a turbine blade with a flat tip, tested under transonic conditions in a stationary, 2-D linear cascade with high freestream turbulence. The experiments were conducted at the Virginia Tech transonic blow-down wind tunnel facility. The effects of tip clearance and exit Mach number on heat transfer distribution were investigated on the tip surface using a transient infrared thermography technique. In addition, thin film gages were used to study similar effects in heat transfer on the near-tip regions at 94% height based on engine blade span of the pressure and suction sides. Surface oil flow visualizations on the blade tip region were carried-out to shed some light on the leakage flow structure. Experiments were performed at three exit Mach numbers of 0.7, 0.85, and 1.05 for two different tip clearances of 0.9% and 1.8% based on turbine blade span. The exit Mach numbers tested correspond to exit Reynolds numbers of 7.6 × 105, 9.0 × 105, and 1.1 × 106 based on blade true chord. The tests were performed with a high freestream turbulence intensity of 12% at the cascade inlet. Results at 0.85 exit Mach showed that an increase in the tip gap clearance from 0.9% to 1.8% translates into a 3% increase in the average heat transfer coefficients on the blade tip surface. At 0.9% tip clearance, an increase in exit Mach number from 0.85 to 1.05 led to a 39% increase in average heat transfer on the tip. High heat transfer was observed on the blade tip surface near the leading edge, and an increase in the tip clearance gap and exit Mach number augmented this near-leading edge tip heat transfer. At 94% of engine blade height on the suction side near the tip, a peak in heat transfer was observed in all test cases at s/C = 0.66, due to the onset of a downstream leakage vortex, originating from the pressure side. An increase in both the tip gap and exit Mach number resulted in an increase, followed by a decrease in the near-tip suction side heat transfer. On the near-tip pressure side, a slight increase in heat transfer was observed with increased tip gap and exit Mach number. In general, the suction side heat transfer is greater than the pressure side heat transfer, as a result of the suction side leakage vortices.

Platform Film Cooling Investigation on an HP Nozzle Vane Cascade With Discrete Shaped Holes and Slot Film Cooling

2020 ◽  
Author(s):  
G. Barigozzi ◽  
A. Perdichizzi ◽  
L. Abba ◽  
L. Pestelli
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Heat Transfer Performance ◽  
Leading Edge ◽  
Suction Side ◽  
Secondary Flows ◽  
Intensity Level ◽  
Transfer Performance ◽  
Film Cooling Effectiveness ◽  
Exit Mach Number

Abstract The present paper reports on an experimental investigation on the aerodynamic and heat transfer performance of different platform cooling schemes: two based on cylindrical and shaped holes and one featuring a slot located upstream of the leading edge plane simulating the combustor to stator interface gap. Tests were run on a 6-vane cascade operated at an isentropic cascade exit Mach number of 0.4 and a significant inlet turbulence intensity level of about 9%. The cooling schemes were first tested to quantify their impact on secondary flows and related losses for variable injection conditions. Heat transfer performance was then assessed through adiabatic film cooling effectiveness and heat transfer coefficient measurements. The Net Heat Flux Reduction parameter was then computed to critically assess the cooling schemes. When compared with the cylindrical hole scheme, shaped holes outperform for all tested injection rates, while the slot alone is able to thermally protect only the front of the passage. Discrete holes are required to cool the platform region along the whole pressure side and the suction side leading edge region.

Large Eddy Simulation of the Laminar Heat Transfer Augmentation on the Pressure Side of a Turbine Vane Under Freestream Turbulence

2019 ◽  
Vol 141 (4) ◽  
Author(s):  
Yousef Kanani ◽  
Sumanta Acharya ◽  
Forrest Ames
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Leading Edge ◽  
Transition To Turbulence ◽  
Experimental Measurements ◽  
Pressure Side ◽  
Freestream Turbulence ◽  
Heat Transfer Augmentation ◽  
Eddy Simulation ◽  
Large Eddy

Vane pressure side heat transfer is studied numerically using large eddy simulation (LES) on an aft-loaded vane with a large leading edge over a range of turbulence conditions. Numerical simulations are performed in a linear cascade at exit chord Reynolds number of Re = 5.1 × 105 at low (Tu ≈ 0.7%), moderate (Tu ≈ 7.9%), and high (Tu ≈ 12.4%) freestream turbulence with varying length scales as prescribed by the experimental measurements of Varty and Ames (2016, “Experimental Heat Transfer Distributions Over an Aft Loaded Vane With a Large Leading Edge at Very High Turbulence Levels,” ASME Paper No. IMECE2016-67029). Heat transfer predictions on the vane pressure side are in a very good agreement with the experimental measurements and the heat transfer augmentation due to the freestream turbulence is well captured. At Tu ≈ 12.4%, freestream turbulence enhances the Stanton number on the pressure surface without boundary layer transition to turbulence by a maximum of about 50% relative to the low freestream turbulence case. Higher freestream turbulence generates elongated structures and high-velocity streaks wrapped around the leading edge that contain significant energy. Amplification of the velocity streaks is observed further downstream with max rms of 0.3 near the trailing edge but no transition to turbulence or formation of turbulence spots is observed on the pressure side. The heat transfer augmentation at the higher freestream turbulence is primarily due to the initial amplification of the low-frequency velocity perturbations inside the boundary layer that persist along the entire chord of the airfoil. Stanton numbers appear to scale with the streamwise velocity fluctuations inside the boundary layer.

scholarly journals Scale-Resolving Simulations of Bypass Transition in a High-Pressure Turbine Cascade Using a Spectral Element Discontinuous Galerkin Method

2017 ◽  
Vol 140 (3) ◽  
Author(s):  
Anirban Garai ◽  
Laslo T. Diosady ◽  
Scott M. Murman ◽  
Nateri K. Madavan
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Guide Vane ◽  
Suction Side ◽  
Spectral Element ◽  
Bypass Transition ◽  
Heat Transfer Characteristics ◽  
High Pressure Turbine ◽  
Computational Capability ◽  
Inflow Turbulence

The application of a new computational capability for accurate and efficient high-fidelity scale-resolving simulations of turbomachinery is presented. The focus is on the prediction of heat transfer and boundary layer characteristics with comparisons to the experiments of Arts et al. (1990, “Aero–Thermal Investigation of a Highly Loaded Transonic Linear Turbine Guide Vane Cascade,” von Karman Institute for Fluid Dynamics, Rhode St. Genese, Belgium, Technical Note No. 174.) for an uncooled, transonic, linear high-pressure turbine (HPT) inlet guide vane cascade that includes the effects of elevated inflow turbulence. The computational capability is based on an entropy-stable, discontinuous Galerkin (DG) spectral element approach that extends to arbitrarily high orders of spatial and temporal accuracy. The suction side of the vane undergoes natural transition for the clean inflow case, while bypass transition mechanisms are observed in the presence of elevated inflow turbulence. The airfoil suction-side boundary layer turbulence characteristics during the transition process thus differ significantly between the two cases. Traditional simulations based on the Reynolds-averaged Navier–Stokes (RANS) fail to predict these transition characteristics. The heat transfer characteristics for the simulations with clean inflow agree well with the experimental data, while the heat transfer characteristics for the bypass transition cases agree well with the experiment when higher inflow turbulence levels are prescribed. The differences between the clean and inflow turbulence cases are also highlighted through a detailed examination of the characteristics of the transitional and turbulent flow fields.

Flowfield Measurements in the Endwall Region of a Stator Vane

1999 ◽  
Author(s):  
M. B. Kang ◽  
K. A. Thole
Keyword(s):  
Heat Transfer ◽  
Suction Side ◽  
Secondary Flows ◽  
Laser Doppler Velocimeter ◽  
Transfer Coefficients ◽  
Suction Surface ◽  
Numerical Predictions ◽  
High Heat Transfer ◽  
Passage Vortex ◽  
Stator Vane

A first stage stator vane experiences high heat transfer rates particularly near the end wall where strong secondary flows occur. In order to improve numerical predictions of the complex endwall flow at low speed conditions, benchmark quality experimental data are required. This study documents the flowfield in the endwall region of a stator vane that has been scaled up by a factor of nine while matching an engine exit Reynolds number of Reex = 1.2·106. Laser Doppler velocimeter (LDV) measurements of all three components of the mean and fluctuating velocities are presented for several flow planes normal to the turbine vane. Measurements indicate that downstream of the minimum static pressure location on the suction surface of the vane, an attenuated suction side leg of the horseshoe vortex still exists. At this location, the peak turbulent kinetic energy coincides with the center of the passage vortex location. These flowfield measurements were also related to previously reported convective heat transfer coefficients on the endwall showing that high Stanton numbers occur where the passage vortex brings mainstream fluid towards the vane surface.

Film Cooling Jet Injection Effect in Heat Transfer Coefficient Augmentation for the Pressure Side Cooling of Turbine Vane

2014 ◽  
Author(s):  
Hossein Nadali Najafabadi ◽  
Matts Karlsson ◽  
Mats Kinell ◽  
Esa Utriainen
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Time Resolved ◽  
Turbine Vane

Improving film cooling performance of turbine vanes and blades is often achieved through application of multiple arrays of cooling holes on the suction side, the showerhead region and the pressure side. This study investigates the pressure side cooling under the influence of single and multiple rows of cooling in the presence of a showerhead from a heat transfer coefficient augmentation perspective. Experiments are conducted on a prototype turbine vane working at engine representative conditions. Transient IR thermography is used to measure time-resolved surface temperature and the semi-infinite method is utilized to calculate the heat transfer coefficient on a low conductive material. Investigations are performed for cylindrical and fan-shaped holes covering blowing ratio 0.6 and 1.8 at density ratio of about unity. The freestream turbulence is approximately 5% close to the leading edge. The resulting heat transfer coefficient enhancement, the ratio of HTC with to that without film cooling, from different case scenarios have been compared to showerhead cooling only. Findings of the study highlight the importance of showerhead cooling to be used with additional row of cooling on the pressure side in order to reduce heat transfer coefficient enhancement. In addition, it is shown that extra rows of cooling will not significantly influence heat transfer augmentation, regardless of the cooling hole shape.

Direct Numerical Simulations of Transitional Flow in Turbomachinery

2006 ◽  
Vol 128 (4) ◽  
pp. 668-678 ◽  
Author(s):  
J. G. Wissink ◽  
W. Rodi
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Free Stream ◽  
Separation Bubble ◽  
Suction Side ◽  
Direct Numerical Simulations ◽  
Pressure Side ◽  
Turbulent Fluctuations ◽  
First Case ◽  
Turbine Cascades

An overview is provided of various direct numerical simulations (DNS) of transitional flows in turbine-related geometries. Two flow cases are considered: the first case concerns separating flow over a flat plate and the second case flows in turbine cascades. In the first case, in which Re=60,000, either an oscillating oncoming flow (1) or a uniform flow with and without oncoming turbulent free-stream fluctuations (2) is prescribed at the inlet. In both subcases (1) and (2), separation is induced by a contoured upper wall. In (1), the separated boundary layer is found to roll up due to a Kelvin-Helmholtz (KH) instability. This rolled-up shear layer is subject to spanwise instability and disintegrates rapidly into turbulent fluctuations. In (2), a massive separation bubble is obtained in the simulation without oncoming free-stream fluctuations. A KH instability is eventually triggered by numerical round-off error and is followed again by a rapid transition. With oncoming turbulent fluctuations, this KH instability is triggered much earlier and transition is enhanced, which leads to a drastic reduction in size of the separation bubble. The second case, concerning flow in turbine cascades, includes (1) flow in the T106 turbine cascade with periodically oncoming wakes at Re=51,800 and (2) flow and heat transfer in a MTU cascade with oncoming wakes and background turbulence at Re=72,000. In the simulation of flow in the T106 cascade with oncoming wakes, the boundary layer along the downstream half of the suction side is found to separate intermittently and subsequently rolls up due to a KH instability leading to separation-induced transition. At times when the wakes impinge separation is suppressed. In the simulations of flow around a MTU turbine blade, evidence of by-pass transition in the suction-side boundary-layer flow is observed while the pressure-side boundary layer remains laminar in spite of significant fluctuations present. In agreement with the experiments, the impinging wakes cause the heat transfer coefficient to increase significantly in the transitional suction-side region close to the trailing edge and by about 30% on the pressure side. The large increase in heat transfer in the pre-transitional suction-side region observed in the experiments could not be reproduced. The discrepancy is explained by differences in spectral contents of the turbulence in the oncoming wakes.

Reduction in Secondary Losses in Turbine Cascade Using Contoured Boundary Layer Fence

2014 ◽  
Author(s):  
Srikanth Deshpande ◽  
Marcus Thern ◽  
Magnus Genrup
Keyword(s):  
Boundary Layer ◽  
Kinetic Energy ◽  
Pressure Gradient ◽  
Total Pressure ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Turbine Cascade ◽  
Flow Passage ◽  
Horse Shoe Vortex

Present work deals with reducing secondary losses in turbine cascade by using boundary layer fences in two ways. Firstly, to reduce the strength of vortex which is incident at the leading edge of airfoil and hence reduce the strength of horse shoe vortex, and secondly, to reduce the pressure gradient between the pressure side and the suction side in the flow passage region between airfoils. In previous works, the boundary layer fence followed the profile of airfoil. In this publication, boundary layer fence does not follow the profile of airfoil i.e stagger and camber of boundary fence is different when compared to airfoil. A profiled boundary layer fence is proposed in the present work which reduces the incident vorticity and also reduces pressure gradient from pressure side to suction side. Such boundary layer fence was checked on T106 test cascade which is available as open literature. Numerical work is carried out using commercial software Ansys CFX. Viscous RANS simulations are carried out using k-ω SST turbulence model with yplus value around unity on all walls. Coefficient of secondary kinetic energy (CSKE) and Secondary Kinetic energy helicity (SKEH) are used as target functions. Total pressure loss is also monitored. All the three functions show a reduction in secondary loss. The strength of horse shoe vortex is reduced by the fence protruding in front of leading edge. The converging flow passage created by the fence near the pressure side of airfoil reduces the pressure gradient from pressure side to suction side. The total pressure loss was reduced by 1.5 % and CSKE was improved by 36 % when the boundary layer fence was adopted.

Formation and Transport of Secondary Flows Caused by Vortex-Blade Interaction in a High Pressure Turbine

2015 ◽  
Author(s):  
Zuo-Jun Wei ◽  
Wei-Yang Qiao ◽  
Ping-Ping Chen ◽  
Jian Liu
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Pressure Gradient ◽  
Suction Side ◽  
Secondary Flows ◽  
Pressure Side ◽  
High Pressure Turbine ◽  
Suction Surface ◽  
Blade Loading ◽  
Passage Vortex

As modern turbines are designed with low aspect ratio and high blade loading, secondary flow interactions become more important. In the present work, numerical simulation is performed in a two-stage high-pressure turbine with divergent meridional passage to investigate the transport and interaction of secondary vortex from the first stage rotor within the second stage’s stator. Scale-Adaptive Simulation model coupled with Shear Stress Transport model (SAS-SST turbulence model) is used to capture the flow structures caused by the interaction in the second stator. Coupled with the passage vortex of the first rotor, the shed vortex rotates opposite in the direction and has comparable strength. As both of these vortices convect downstream to the stator bladerow, each deforms into two legs on the pressure and suction sides in the passage. In the passage due to the cross pressure gradient by blade loading, all the low-momentum fluid contained in these vortices moves towards the suction side. Besides, with the existing static pressure gradient in radial direction and vortex dynamics, the suction-side leg and the pressure-side leg move in different radial directions. The suction side leg of incoming passage vortex moves towards the endwall along the suction surface and interacts with the developing passage vortex of the second stator. The incoming shed vortex moves towards the midspan and rolls up the boundary layer fluid from suction surface. Due to the interactions between the incoming shed vortices from the hub and casing and the boundary layer of second stator, two counter-rotating vortices are formed near the midspan. Additional high loss is found there at the outlet plane, which has a comparable magnitude to the endwall secondary loss. The pressure side leg of the incoming passage vortex remains in a certain span with that of the incoming shed vortex and is not engulfed by the developing passage vortex.

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