Combustor Turbine Interface Studies—Part 2: Flow and Thermal Field Measurements

2003 ◽  
Vol 125 (2) ◽  
pp. 203-209 ◽  
Author(s):  
W. F. Colban ◽  
A. T. Lethander ◽  
K. A. Thole ◽  
G. Zess
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Film Cooling ◽  
Large Scale ◽  
Field Measurements ◽  
Pressure Profile ◽  
Stagnation Line ◽  
Turbine Engine ◽  
Turbine Vane ◽  
Cooling Methods

Most turbine inlet flows resulting from the combustor exit are nonuniform in the near-platform region as a result of cooling methods used for the combustor liner. These cooling methods include injection through film-cooling holes and injection through a slot that connects the combustor and turbine. This paper presents thermal and flow field measurements in the turbine vane passage for a combustor exit flow representative of what occurs in a gas turbine engine. The experiments were performed in a large-scale wind tunnel facility that incorporates combustor and turbine vane models. The measured results for the thermal and flow fields indicate a secondary flow pattern in the vane passage that can be explained by the total pressure profile exiting the combustor. This secondary flow field is quite different than that presented for past studies with an approaching flat plate turbulent boundary layer along the upstream platform. A counter-rotating vortex that is positioned above the passage vortex was identified from the measurements. Highly turbulent and highly unsteady flow velocities occur at flow impingement locations along the stagnation line.

Combustor Turbine Interface Studies: Part 2 — Flow and Thermal Field Measurements

2002 ◽  
Author(s):  
W. F. Colban ◽  
A. T. Lethander ◽  
K. A. Thole ◽  
G. Zess
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Film Cooling ◽  
Large Scale ◽  
Field Measurements ◽  
Pressure Profile ◽  
Stagnation Line ◽  
Turbine Engine ◽  
Turbine Vane ◽  
Cooling Methods

Most turbine inlet flows resulting from the combustor exit are non-uniform in the near-platform region as a result of cooling methods used for the combustor liner. These cooling methods include injection through film-cooling holes and injection through a slot that connects the combustor and turbine. This paper presents thermal and flow field measurements in the turbine vane passage for a combustor exit flow representative of what occurs in a gas turbine engine. The experiments were performed in a large-scale wind tunnel facility that incorporates combustor and turbine vane models. The measured results for the thermal and flow fields indicate a secondary flow pattern in the vane passage that can be explained by the total pressure profile exiting the combustor. This secondary flow field is quite different than that presented for past studies with an approaching flat plate turbulent boundary layer along the upstream platform. A counter-rotating vortex that is positioned above the passage vortex was identifed from the measurements. Highly turbulent and highly unsteady flow velocities occur at flow impingment locations along the stagnation line.

Measurements of Hub Flow Interaction on Film Cooled Nozzle Guide Vane in Transonic Annular Cascade

2015 ◽  
Vol 137 (8) ◽  
Author(s):  
Lamyaa A. El-Gabry ◽  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten Fransson
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Field Measurements ◽  
Co2 Concentration ◽  
Horseshoe Vortex ◽  
Mainstream Flow ◽  
Nozzle Guide

An experimental study has been performed in a transonic annular sector cascade of nozzle guide vanes (NGVs) to investigate the aerodynamic performance and the interaction between hub film cooling and mainstream flow. The focus of the study is on the endwalls, specifically the interaction between the hub film cooling and the mainstream. Carbon dioxide (CO2) has been supplied to the coolant holes to serve as tracer gas. Measurements of CO2 concentration downstream of the vane trailing edge (TE) can be used to visualize the mixing of the coolant flow with the mainstream. Flow field measurements are performed in the downstream plane with a five-hole probe to characterize the aerodynamics in the vane. Results are presented for the fully cooled and partially cooled vane (only hub cooling) configurations. Data presented at the downstream plane include concentration contour, axial vorticity, velocity vectors, and yaw and pitch angles. From these investigations, secondary flow structures such as the horseshoe vortex, passage vortex, can be identified and show the cooling flow significantly impacts the secondary flow and downstream flow field. The results suggest that there is a region on the pressure side (PS) of the vane TE where the coolant concentrations are very low suggesting that the cooling air introduced at the platform upstream of the leading edge (LE) does not reach the PS endwall, potentially creating a local hotspot.

Measurements of Hub Flow Interaction on Film Cooled Nozzle Guide Vane in Transonic Annular Cascade

2012 ◽  
Author(s):  
Lamyaa A. El-Gabry ◽  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten Fransson
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Field Measurements ◽  
Co2 Concentration ◽  
Pressure Side ◽  
Trailing Edge ◽  
Nozzle Guide

An experimental study has been performed in a transonic annular sector cascade of nozzle guide vanes to investigate the aerodynamic performance and the interaction between hub film cooling and mainstream flow. The focus of the study is on the endwalls, specifically the interaction between the hub film cooling and the mainstream. Carbon dioxide (CO2) has been supplied to the coolant holes to serve as tracer gas. Measurements of CO2 concentration downstream of the vane trailing edge can be used to visualize the mixing of the coolant flow with the mainstream. Flow field measurements are performed in the downstream plane with a 5-hole probe to characterize the aerodynamics in the vane. Results are presented for the fully cooled and partially cooled vane (only hub cooling) configurations. Data presented at the downstream plane include concentration contour, axial vorticity, velocity vectors, and yaw and pitch angles. From these investigations, secondary flow structures such as the horseshoe vortex, passage vortex, can be identified and show the cooling flow significantly impacts the secondary flow and downstream flow field. The results suggest that there is a region on the pressure side of the vane trailing edge where the coolant concentrations are very low suggesting that the cooling air introduced at the platform upstream of the leading edge does not reach the pressure side endwall, potentially creating a local hotspot.

Turbulence and Heat Transfer Measurements in an Inclined Large Scale Film Cooling Array: Part I—Velocity and Turbulence Measurements

2011 ◽  
Author(s):  
Lamyaa A. El-Gabry ◽  
Douglas R. Thurman ◽  
Philip E. Poinsatte ◽  
James D. Heidmann
Keyword(s):  
Heat Transfer ◽  
Flow Field ◽  
Film Cooling ◽  
Large Scale ◽  
Field Measurements ◽  
Scale Model ◽  
Liquid Crystal Thermography ◽  
Detailed Surface ◽  
Cooling Hole ◽  
Hotwire Anemometry

A large-scale model of an inclined row of film cooling holes is used to obtain detailed surface and flow field measurements that will enable future computational fluid dynamics code development and validation. The model consists of three holes of 1.9-cm diameter that are spaced 3 hole diameters apart and inclined 30° from the surface. The length to diameter ratio of the coolant holes is about 18. Measurements include film effectiveness using IR thermography and near wall thermocouples, heat transfer using liquid crystal thermography, flow field temperatures using a thermocouple, and velocity and turbulence quantities using hotwire anemometry. Results are obtained for blowing ratios of up to 2 in order to capture severe conditions in which the jet is lifted. This first part of the two-part paper presents the detailed velocity component and turbulence stresses along the centerline of the film-cooling hole and at various streamwise locations.

Detailed Velocity and Turbulence Measurements in an Inclined Large-Scale Film Cooling Array

2013 ◽  
Vol 135 (6) ◽  
Author(s):  
Lamyaa A. El-Gabry ◽  
Douglas R. Thurman ◽  
Philip E. Poinsatte ◽  
James D. Heidmann
Keyword(s):  
Fluid Dynamics ◽  
Computational Fluid Dynamics ◽  
Flow Field ◽  
Film Cooling ◽  
Large Scale ◽  
Field Measurements ◽  
Scale Model ◽  
Data Sets ◽  
Turbulent Dissipation ◽  
Hotwire Anemometry

A large-scale model of an inclined row of film cooling holes is used to obtain detailed surface and flow field measurements that will enable future computational fluid dynamics code development and validation. The model consists of three holes of 1.9-cm diameter that are spaced three hole diameters apart and inclined 30 deg from the surface. The length to diameter ratio of the coolant holes is about 18. Measurements include film effectiveness using IR thermography and near wall thermocouples, heat transfer using liquid crystal thermography, flow field temperatures using a thermocouple, and velocity and turbulence quantities using hotwire anemometry. Results are obtained for blowing ratios of up to 2 in order to capture severe conditions in which the jet is lifted. For purposes of comparison with prior art, measurements of the velocity and turbulence field along the jet centerline are made and compare favorably with two data sets in the open literature thereby verifying the test apparatus and methodology are able to replicate existing data sets. In addition, a computational fluid dynamics model using a two-equation turbulence model is developed, and the results for velocity, turbulent kinetic energy and turbulent dissipation rate are compared with experimentally derived quantities.

Experimental Test Rig for Combustor-Turbine Interaction Research and Test Results Analysis

2015 ◽  
Author(s):  
Hong Yin ◽  
Shi Liu ◽  
Yongxin Feng ◽  
Mingfei Li ◽  
Jing Ren ◽  
...  
Keyword(s):  
Flow Field ◽  
Film Cooling ◽  
Cooling System ◽  
Leading Edge ◽  
Inlet Section ◽  
Test Results ◽  
Test Rig ◽  
Stagnation Line ◽  
Turbine Vane ◽  
Interaction Research

Lean premixed combustion technology has been adopted widely for heavy duty gas turbine application. At the combustor inlet section, the basic burner arrangement is multiple-swirl configuration. The multiple-swirl structure creates complicated swirling flow field downstream, which is characterized as non-uniform flow-field and has impacts on the turbine vane. The issue of combustor-turbine interaction effect has become quite prominent. This paper introduces a new test rig for the combustor-turbine interaction research, which is designed to investigate the influence of multiple-swirl on the turbine vane system. The test rig consists of a combustor simulator and a first stage turbine vane with cooling system. Measurement techniques including the Pressure Sensitive Paint and five-hole probe are applied. Preliminary test results show that the multiple-swirl combustor flow field has significant impact on the vane cooling system due to the residual swirl intensity at the combustor outlet. The stagnation line at the vane leading edge is obviously altered compared to uniform inflow. Film cooling effectiveness distribution has distinct characteristics under different conditions. The leading edge is most significantly influenced, while the pressure side film cooling system is affected slightly. Under certain condition, the suction side film cooling is influenced locally.

Secondary Flow Measurements in a Turbine Passage With Endwall Flow Modification

2000 ◽  
Vol 122 (4) ◽  
pp. 651-658 ◽  
Author(s):  
Nicole V. Aunapu ◽  
Ralph J. Volino ◽  
Karen A. Flack ◽  
Ryan M. Stoddard
Keyword(s):  
Secondary Flow ◽  
Film Cooling ◽  
Large Scale ◽  
Turbine Blades ◽  
Leading Edge ◽  
Suction Side ◽  
Flow Measurements ◽  
Turbine Engine ◽  
Flow Modification ◽  
Passage Vortex

A flow modification technique is introduced in an attempt to allow increased turbine inlet temperatures. A large-scale two half-blade cascade simulator is used to model the secondary flow between two adjacent turbine blades. Various flow visualization techniques and measurements are used to verify that the test section replicates the flow of an actual turbine engine. Two techniques are employed to modify the endwall secondary flow, specifically the path of the passage vortex. Six endwall jets are installed at a location downstream of the saddle point near the leading edge of the pressure side blade. These wall jets are found to be ineffective in diverting the path of the passage vortex. The second technique utilizes a row of 12 endwall jets whose positions along the centerline of the passage are based on results from an optimized boundary layer fence. The row of jets successfully diverts the path of the passage vortex and decreases its effect on the suction side blade. This can be expected to increase the effectiveness of film cooling in that area. The row of jets increases the aerodynamic losses in the passage, however. Secondary flow measurements are presented showing the development of the endwall flow, both with and without modification. [S0889-504X(00)01004-7]

Flow Field Measurements on a Large Scale Turbine Cascade With Leading Edge Film Cooling by Two Rows of Holes

1997 ◽  
Author(s):  
Sabine Ardey ◽  
Leonhard Fottner
Keyword(s):  
Flow Field ◽  
Film Cooling ◽  
Gas Turbines ◽  
Large Scale ◽  
Vortex Pair ◽  
Leading Edge ◽  
Field Measurements ◽  
Suction Side ◽  
Pressure Side ◽  
Turbine Cascade

To increase the understanding of the aerodynamic processes dominating the flow field of turbine bladings with leading edge film cooling, isothermal investigations were carried out on a large scale high pressure turbine cascade. Near the stagnation point the blades are equipped with one row of film cooling holes on the suction side and one on the pressure side. Blowing ratio, turbulence intensity, Mach number, and Reynolds number are set to values typically found in modern gas turbines. Experimental data of the cascade flow were obtained by pneumatic probes and static pressure tappings. The flow field was visualized by Schlieren and oil flow techniques. For detailed investigations near the blowing holes the Laser Transit Velocimetry and the three dimensional Hot Wire Anemometry were used. The flow field measurements in the near hole region of the suction side show the typical kidney shaped vortex pair. A local suction peak on the pressure side causes a large recirculation area behind the holes on the pressure side and induces separation bubbles in between the pressure side holes. This leads to the generation of two pairs of vortices: The kidney-vortex is located on top of a second vortex pair and a trough flow that fills up the deficit of the recirculation. Thus the film cooling air is detached from the pressure side surface. In addition to the mean flow vectors Reynolds stress components are a good means to judge the propagation of the jet. In spite of the complex flow pattern occurring on each single jet, the surveyed loss-increase due to the leading edge blowing can be predicted by the mixing layer model.

Flow Field Computations of Combustor-Turbine Interactions Relevant to a Gas Turbine Engine

2004 ◽  
Vol 126 (1) ◽  
pp. 122-129 ◽  
Author(s):  
Sarah Stitzel ◽  
Karen A. Thole
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Gas Turbine ◽  
High Performance ◽  
Computational Study ◽  
Pressure Profile ◽  
Turbine Engines ◽  
Turbine Engine ◽  
Stator Vane ◽  
Secondary Flow Field

The current demands for high-performance gas turbine engines can be reached by raising combustion temperatures to increase power output. High combustion temperatures create a harsh environment that leads to the consideration of the durability of the combustor and turbine sections. This paper presents a computational study of a flow field that is representative of what occurs in a combustor and how that flow field convects through the first downstream stator vane. The results of this study indicate that the development of the secondary flow field in the turbine is highly dependent on the incoming total pressure profile. The endwall heat transfer is also found to depend strongly on the secondary flow field.

Secondary Flow Measurements in a Turbine Passage With Endwall Flow Modification

2000 ◽  
Author(s):  
Nicole V. Aunapu ◽  
Ralph J. Volino ◽  
Karen A. Flack ◽  
Ryan M. Stoddard
Keyword(s):  
Secondary Flow ◽  
Film Cooling ◽  
Large Scale ◽  
Turbine Blades ◽  
Leading Edge ◽  
Suction Side ◽  
Flow Measurements ◽  
Turbine Engine ◽  
Flow Modification ◽  
Passage Vortex

A flow modification technique is introduced in an attempt to allow increased turbine inlet temperatures. A large-scale two half-blade cascade simulator is used to model the secondary flow between two adjacent turbine blades. Various flow visualization techniques and measurements are used to verify that the test section replicates the flow of an actual turbine engine. Two techniques are employed to modify the endwall secondary flow, specifically the path of the passage vortex. Six endwall jets are installed at a location downstream of the saddle point near the leading edge of the pressure side blade. These wall jets are found to be ineffective in diverting the path of the passage vortex. The second technique utilizes a row of 12 endwall jets whose positions along the centerline of the passage are based on results from an optimized boundary layer fence. The row of jets successfully diverts the path of the passage vortex and decreases its effect on the suction side blade. This can be expected to increase the effectiveness of film cooling in that area. The row of jets increases the aerodynamic losses in the passage, however. Secondary flow measurements are presented showing the development of the endwall flow, both with and without modification.

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