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

2002 ◽  
Vol 124 (2) ◽  
pp. 167-175 ◽  
Author(s):  
G. A. Zess ◽  
K. A. Thole
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Gas Turbine ◽  
Total Pressure ◽  
Large Scale ◽  
Guide Vane ◽  
Leading Edge ◽  
Computational Design ◽  
Horseshoe Vortex ◽  
Laser Doppler Velocimeter

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, flowfield measurements were performed in a large-scale, linear, vane cascade. The flowfield 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 flow-field results showed that the turbulent kinetic energy levels were significantly reduced in the endwall region because of the absence of the unsteady horseshoe vortex.

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.

Improvement of Film Cooling Effectiveness in the Gas Turbine Endwall by Use of Optimization Framework

2019 ◽  
Author(s):  
Daisuke Hata ◽  
Kazuto Kakio ◽  
Yutaka Kawata ◽  
Masahiro Miyabe
Keyword(s):  
Pressure Gradient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Combined Cycle ◽  
Inlet Temperature ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness

Abstract Recently, the number of gas turbine combined cycle plants is rapidly increasing in substitution of nuclear power plants. The turbine inlet temperature (TIT) is constantly being increased in order to achieve higher effectiveness. Therefore, the improvement of the cooling technology for high temperature gas turbine blades is one of the most important issue to be solved. In a gas turbine, the main flow impinging at the leading edge of the turbine blade generates a so called horseshoe vortex by the interaction of its boundary layer and generated pressure gradient at the leading edge. The pressure surface leg of this horseshoe vortex crosses the passage and reaches the blade suction surface, driven by the pressure gradient existing between two consecutive blades. In addition, this pressure gradient generates a cross-flow along the endwall. This all results into a very complex flow field in proximity of the endwall. For this reason, burnouts tend to occur at a specific position in the vicinity of the leading edge. In this research, a methodology to cool the endwall of the turbine blade by means of film cooling jets from the blade surface and the endwall is proposed. The cooling performance is investigated using the transient thermography method. CFD analysis is also conducted to investigate the phenomena occurring at the endwall and calculate the film cooling effectiveness.

Research on Heat-Transfer and Cooling Performance of Gas Turbine Endwall

2017 ◽  
Author(s):  
Kazuto Kakio ◽  
Y. Kawata
Keyword(s):  
Heat Transfer ◽  
Pressure Gradient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Power Plants ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Combined Cycle ◽  
Inlet Temperature ◽  
Cooling Performance

Recently, the number of gas turbine combined cycle plants is rapidly increasing in substitution of nuclear power plants. The turbine inlet temperature (TIT) is being constantly increased in order to achieve higher efficiency. Therefore, the improvement of the cooling technology for high temperature gas turbine blades is one of the most important issue to be solved. In a gas turbine, the main flow impinging at the leading edge of the turbine blade generates a so called horseshoe vortex by the interaction of its boundary layer and generated pressure gradient at the leading edge. The pressure surface leg of this horseshoe vortex crosses the passage and reaches the blade suction surface, driven by the pressure gradient existing between two consecutive blades. In addition, this pressure gradient generates a crossflow along the endwall. This all results into a very complex flow field in proximity of the endwall. For this reason, burnouts tend to occur at a specific position in the vicinity of the leading edge. In this research, a methodology to cool the endwall of the turbine blade by means of film cooling jets from the blade surface is proposed. The cooling performance and heat transfer coefficient distribution is investigated using the transient thermography method. CFD analysis is also conducted to know the phenomena occurring at the end wall and calculate the heat transfer distribution.

EXPERIMENTAL STUDY OF THE ENDWALL HEAT TRANSFER OF A TRANSONIC NOZZLE GUIDE VANE WITH UPSTREAM JET PURGE COOLING PART 1 - EFFECT OF DENSITY RATIO

2021 ◽  
pp. 1-36
Author(s):  
Shuo Mao ◽  
Ridge A. Sibold ◽  
Wing Ng ◽  
Zhigang LI ◽  
Bo Bai ◽  
...  
Keyword(s):  
Large Scale ◽  
Guide Vane ◽  
Flow Direction ◽  
Density Ratio ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Vital Role ◽  
Blowing Ratio ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Abstract Nozzle guide vane platforms often employ complex cooling schemes to mitigate the ever-increasing thermal loads on endwall. This study analyzes, experimentally and numerically, and describes the effect of coolant to mainstream blowing ratio, momentum ratio and density ratio for a typical axisymmetric converging nozzle guide vane platform with an upstream doublet staggered, steep-injection, cylindrical hole purge cooling scheme. Nominal flow conditions were engine-representative and as follows: Maexit = 0.85, Reexit,Cax = 1.5×106 and an inlet large-scale freestream turbulence intensity of 16%. Two blowing ratios were investigated, each corresponding to the design condition and its upper extrema at M = 2.5 and 3.5, respectively. For each blowing ratio, the coolant to mainstream density ratio was varied between DR=1.2, representing typical experimental neglect of coolant density, and DR=1.95, representative of typical engine conditions. The results show that with a fixed coolant-to-mainstream blowing ratio, the density ratio plays a vital role in the coolant-mainstream mixing and the interaction between coolant and horseshoe vortex near the vane leading edge. A higher density ratio leads to a better coolant coverage immediately downstream of the cooling holes but exposes the in-passage endwall near the pressure side. It also causes the in-passage coolant coverage to decay at a higher rate in the flow direction. From the results gathered, both density ratio and blowing ratio should be considered for accurate testing, analysis, and prediction of purge jet cooling scheme performance.

Combustor Turbine Interface Studies: Part 1 — Endwall Effectiveness Measurements

2002 ◽  
Author(s):  
W. F. Colban ◽  
K. A. Thole ◽  
G. Zess
Keyword(s):  
Film Cooling ◽  
Total Pressure ◽  
Large Scale ◽  
Guide Vane ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Profile ◽  
Secondary Flows ◽  
Adiabatic Wall ◽  
Combustor Liner

Improved durability of gas turbine engines is an objective for both military and commercial aeroengines as well as for power generation engines. One region susceptible to degradation in an engine is the junction between the combustor and first vane given that the main gas path temperatures at this location are the highest. The platform at this junction is quite complex in that secondary flow effects, such as the leading edge vortex, are dominant. Past computational studies have shown that the total pressure profile exiting the combustor dictates the development of the secondary flows that are formed. This study examines the effect of varying the combustor liner film-cooling and junction slot flows on the adiabatic wall temperatures measured on the platform of the first vane. The experiments were performed using large-scale models of a combustor and nozzle guide vane in a wind tunnel facility. The results show that varying the coolant injection from the upstream combustor liner leads to differing total pressure profiles entering the turbine vane passage. Endwall adiabatic effectiveness measurements indicate that the coolant does not exit the upstream combustor slot uniformly but instead accumulates along the suction side of the vane and endwall. Increasing the liner cooling continued to reduce endwall temperatures, which was not found to be true with increasing the film-cooling from the liner.

scholarly journals Streakline Flow Visualization Study of a Horseshoe Vortex in a Large-Scale, Two-Dimensional Turbine Stator Cascade

1980 ◽  
Author(s):  
R. E. Gaugler ◽  
L. M. Russell
Keyword(s):  
Boundary Layer ◽  
Large Scale ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Main Stream ◽  
Two Dimensional ◽  
Vortex System ◽  
Turbine Stator ◽  
Neutrally Buoyant ◽  
Limiting Streamlines

Neutrally buoyant helium-filled bubbles were observed as they followed the streamlines in a horseshoe vortex system around the vane leading edge in a large-scale, two-dimensional, turbine stator cascade. Inlet Reynolds number, based on true chord, ranged between 1.0 × 10 ⅝ to 3.0 × 10 ⅝. Bubbles were introduced into the endwall boundary layer through a slot upstream of the vane leading edge. The paths of the bubbles were recorded photographically as streaklines on 16-mm movie film. Individual frames from the film have been selected, and overlayed to show the details of the horseshoe vortex around the leading edge, the transport of the vortex across the passage near the leading edge is clearly seen when compared to the streaks formed by bubbles carried in the main stream. Limiting streamlines on the endwall surface were traced by the flow of oil drops.

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.

Aerodynamic Influence of Endwall Fences Located in the Vicinity of the Leading Edge-Endwall Junction of Nozzle Guide Vanes

2011 ◽  
Author(s):  
Zeki Ozgur Gokce ◽  
Cengiz Camci
Keyword(s):  
Total Pressure ◽  
Numerical Study ◽  
Guide Vane ◽  
Flow Characteristics ◽  
Cross Flow ◽  
Vertical Cylinder ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Horseshoe Vortices ◽  
Nozzle Guide

Secondary flow characteristics like horseshoe vortices and related total pressure losses decrease turbine efficiency. Computerized simulations of potentially favorable modifications in turbine systems could provide a fast, numerical and inexpensive method of evaluating their effects on flow properties: This paper consists of a comparative numerical study of the flow characteristics of a domain containing a vertical cylinder subjected to cross flow and upstream endwall modifications. Analyzing the flow around a turbine nozzle guide vane (NGV) could be simplified by modeling it as a vertical cylinder with a diameter proportional to the leading edge diameter of the blade, and adding upstream endwall fences of varying dimensions and alignments could attenuate the development of a horseshoe vortex. A commercial computational fluid dynamics (CFD) software package, Fluent, was used for the numerical analysis. To validate the modeling strategy, experimental data previously reported in the literature for conventional cylinders in cross flow were compared to the current predictions. A grid independence study was also performed. The lateral distance between the two legs of the horseshoe vortex downstream of the cylinder was decreased by 7% to 14%. All fence types effectively changed the location of the main horseshoe vortex roll-up. The height of the fence was more influential than the length of the fence in modifying flow characteristics. The existence of the fences slightly increased the mass-averaged total pressure loss far downstream of the cylinder; however, beneficial near-fence flow characteristics were observed in all cases. Also, it was noted that an endwall fence could possibly result in decreased interaction between the horseshoe vortices created by consecutive blades in a row of NGV blades, which would be expected to result in improved flow conditions within actual turbine passages.

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.

Combustor Turbine Interface Studies—Part 1: Endwall Effectiveness Measurements

2003 ◽  
Vol 125 (2) ◽  
pp. 193-202 ◽  
Author(s):  
W. F. Colban ◽  
K. A. Thole ◽  
G. Zess
Keyword(s):  
Film Cooling ◽  
Total Pressure ◽  
Large Scale ◽  
Guide Vane ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Profile ◽  
Secondary Flows ◽  
Adiabatic Wall ◽  
Combustor Liner

Improved durability of gas turbine engines is an objective for both military and commercial aeroengines as well as for power generation engines. One region susceptible to degradation in an engine is the junction between the combustor and first vane given that the main gas path temperatures at this location are the highest. The platform at this junction is quite complex in that secondary flow effects, such as the leading edge vortex, are dominant. Past computational studies have shown that the total pressure profile exiting the combustor dictates the development of the secondary flows that are formed. This study examines the effect of varying the combustor liner film-cooling and junction slot flows on the adiabatic wall temperatures measured on the platform of the first vane. The experiments were performed using large-scale models of a combustor and nozzle guide vane in a wind tunnel facility. The results show that varying the coolant injection from the upstream combustor liner leads to differing total pressure profiles entering the turbine vane passage. Endwall adiabatic effectiveness measurements indicate that the coolant does not exit the upstream combustor slot uniformly, but instead accumulates along the suction side of the vane and endwall. Increasing the liner cooling continued to reduce endwall temperatures, which was not found to be true with increasing the film-cooling from the liner.

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