Experimental and Computational Comparisons of Fan-Shaped Film-Cooling on a Turbine Vane Surface

2005 ◽  
Author(s):  
W. Colban ◽  
K. A. Thole ◽  
M. Haendler
Keyword(s):  
Turbulence Model ◽  
Film Cooling ◽  
Turbulence Models ◽  
Complex Flow ◽  
Ir Camera ◽  
Turbine Engine ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Film Cooling Holes

The flow exiting the combustor in a gas turbine engine is considerably hotter than the melting temperature of the turbine section components, of which the turbine nozzle guide vanes see the hottest gas temperatures. One method used to cool the vanes is to use rows of film-cooling holes to inject bleed air that is lower in temperature through an array of discrete holes onto the vane surface. The purpose of this study was to evaluate the row-by-row interaction of fan-shaped holes as compared to the performance of a single row of fan-shaped holes in the same locations. This study presents adiabatic film-cooling effectiveness measurements from a scaled-up, two-passage vane cascade. High resolution film-cooling measurements were made with an infrared (IR) camera at a number of engine representative flow conditions. Computational fluid dynamics (CFD) predictions were also made to evaluate the performance of some of the current turbulence models in predicting a complex flow such as turbine film-cooling. The RNG k-ε turbulence model gave a closer prediction of the overall level of film-effectiveness, while the v2-f turbulence model gave a more accurate representation of the flow physics seen in the experiments.

Experimental and Computational Comparisons of Fan-Shaped Film Cooling on a Turbine Vane Surface

2006 ◽  
Vol 129 (1) ◽  
pp. 23-31 ◽  
Author(s):  
W. Colban ◽  
K. A. Thole ◽  
M. Haendler
Keyword(s):  
Turbulence Model ◽  
Film Cooling ◽  
Turbulence Models ◽  
Infrared Camera ◽  
Complex Flow ◽  
Turbine Engine ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Film Cooling Holes

The flow exiting the combustor in a gas turbine engine is considerably hotter than the melting temperature of the turbine section components, of which the turbine nozzle guide vanes see the hottest gas temperatures. One method used to cool the vanes is to use rows of film-cooling holes to inject bleed air that is lower in temperature through an array of discrete holes onto the vane surface. The purpose of this study was to evaluate the row-by-row interaction of fan-shaped holes as compared to the performance of a single row of fan-shaped holes in the same locations. This study presents adiabatic film-cooling effectiveness measurements from a scaled-up, two-passage vane cascade. High-resolution film-cooling measurements were made with an infrared camera at a number of engine representative flow conditions. Computational fluid dynamics predictions were also made to evaluate the performance of some of the current turbulence models in predicting a complex flow such as turbine film-cooling. The renormalization group (RNG) k‐ε turbulence model gave a closer prediction of the overall level of film effectiveness, while the v2‐f turbulence model gave a more accurate representation of the flow physics seen in the experiments.

Interactions Between the Combustor Swirl and the High Pressure Stator of a Turbine

2012 ◽  
Author(s):  
Lucas Giller ◽  
Heinz-Peter Schiffer
Keyword(s):  
High Pressure ◽  
Flow Field ◽  
Film Cooling ◽  
Field Measurements ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Film Cooling Holes ◽  
The Impact

The interaction between the strongly swirling combustor outflow and the high pressure turbine nozzle guide vanes were investigated at the cascade test rig at Technische Universität Darmstadt. The test section of the rig consists of six swirl generators and five cascade vanes. The three middle vanes are equipped with film cooling holes at the leading edges. The swirler nozzles are aligned with the center of the cascade passages. The operating settings are defined by the swirl number, the distance between the swirler nozzles and the vanes, the blowing ratio and the radial angle of the film cooling holes. Flow field measurements using PIV downstream of the swirlers and five hole probe measurements at the inlet and outlet plane of the cascade were accomplished. Measurements using the ammonia diazo technique to determine the adiabatic film cooling effectiveness on the surface of the center cascade vane were also carried out. It is shown that a swirling inflow leads to a strong alteration of the flow field and the losses in the passages in comparison to an axial inflow. Furthermore, the impact of the swirl on the formation of the cooling film and it’s adiabatic film cooling effectiveness is presented.

Effects of Blowing Ratio on Multiple Shallow Angle Film Cooling Holes

2012 ◽  
Vol 225 ◽  
pp. 49-54 ◽  
Author(s):  
Kamil Abdullah ◽  
Ken Ichi Funazaki
Keyword(s):  
Temperature Field ◽  
Surface Temperature ◽  
Film Cooling ◽  
Cooling Effectiveness ◽  
Ir Camera ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Hole Configuration ◽  
The Common ◽  
Film Cooling Holes

This paper presents the investigation on the effects of the blowing ratio of multiple shallow angle film cooling holes. Multiple film cooling holes having a shallow hole angle (θ = 20°), arranged to perform in-line hole configuration has been considered in the present study. The investigation focuses on the effects of high blowing ratio of the film cooling effectiveness which have been carried out at ReD= 3100 and BR = 2.0, 3.0 and 4.0. The experiments make use of the IR camera in capturing the surface temperature to determine the film cooling effectiveness. The contours of the film cooling effectiveness distribution together with plots on laterally average film cooling effectiveness along the x/D are presented. The discussions have been made with a support of the temperature field captured at x/D = 3, 13, 23, and 33. The results clearly show the benefit of the employment of shallow hole angle (θ = 20°) at high blowing ratio which is much more superior in comparison to the common hole configuration (θ = 35°).

Experimental Characterization of Film-Cooling Effectiveness Near Combustor Dilution Holes

2005 ◽  
Author(s):  
J. J. Scrittore ◽  
K. A. Thole ◽  
S. W. Burd
Keyword(s):  
Film Cooling ◽  
High Momentum ◽  
Complex Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Flow ◽  
Thermal Fields ◽  
Flow Phenomena ◽  
Film Cooling Holes ◽  
Low Speed Wind Tunnel

Cooling combustor chambers for gas turbine engines is challenging, given the complex flow and thermal fields inherent to these modules. This complexity, in part, arises from the interaction of high-momentum dilution jets required to mix the fuel with film cooling jets that are intended to cool the combustor walls. This paper discusses the experimental results from a combustor simulator tested in a low-speed wind tunnel that includes both the dilution jets and film-cooling jets. The specific purpose of this study is to evaluate the influence that the dilution jets has on the film-cooling effectiveness. Infrared thermography was used to measure surface temperatures along a low thermal conductivity plate to quantify the adiabatic effectiveness from an array of film cooling holes with the presence of dilution holes. To further understand the flow phenomena, thermocouple probes and laser Doppler velocimetry were used to measure the thermal and flow fields, respectively. Parametric experiments indicate that the film cooling flow is disrupted along the combustor walls in the vicinity of the high-momentum dilution jets. In fact, a significant penalty in cooling effectiveness of the combustor is observed with increased dilution jet penetration.

Film Cooling Effectiveness in a Gas Turbine Engine: A Review

2014 ◽  
Vol 71 (2) ◽  
Author(s):  
Ehsan Kianpour ◽  
Nor Azwadi Che Sidik ◽  
Iman Golshokouh
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Future Research ◽  
Inlet Temperature ◽  
Turbine Engine ◽  
Film Cooling Effectiveness ◽  
Engine Efficiency ◽  
External Cooling ◽  
Critical Components ◽  
Film Cooling Holes

This study was carried out to extend database knowledge about the function of film cooling holes at the end of combustor and the inlet of turbine. Using the well-known Brayton cycle, raising the turbine inlet temperature is the key to obtain higher engine efficiency in gas turbine engines. However, high temperature of the combustor exit flow causes non-uniformities. These non-uniformities lead to the reduction of expected life of critical components. Therefore, an appropriate cooling technique should be designed to protect these parts. Film cooling is one of the most effective external cooling methods. Various film cooling techniques presented in the literature have been investigated. Moreover, challenges and future directions of film cooling techniques have been reviewed and presented in this paper. The aim of this review is to summarize recent development in research on film cooling techniques and attempt to identify some challenging issues that need to be solved for future research.

The Effect of Step Misalignment on Purge Flow Cooling of Nozzle Guide Vane at Transonic Conditions

2020 ◽  
Vol 142 (10) ◽  
Author(s):  
Luke Luehr ◽  
Ridge Sibold ◽  
Shuo Mao ◽  
Wing F. Ng ◽  
Zhigang Li ◽  
...  
Keyword(s):  
Nusselt Number ◽  
Film Cooling ◽  
Guide Vane ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Purge Flow ◽  
Exit Mach Number ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

Abstract This study describes a detailed investigation on the effects that upstream step misalignment and upstream purge film cooling have on the endwall heat transfer for first stage nozzle guide vanes (NGVs) in a gas turbine at transonic conditions. Endwall Nusselt number and adiabatic film-cooling effectiveness distributions were experimentally measured and compared with flow visualization. Tests were conducted in a transonic linear cascade blowdown facility at an inlet freestream turbulence intensity of 16%, an exit Mach number of 0.85, and an exit Re = 1.5 × 106 based on axial chord. Varied upstream purge blowing ratios (BRs) and a no-blowing case were tested for three different upstream step geometries, the baseline (no misalignment), a span-wise upstream step of +4.86% span, and a step of −4.86% span. Experimentation shows that compared with no-blowing case, the addition of upstream purge film cooling increases the Nusselt number at injection upward of 50% but lowers it in the passage throat by approximately 20%. The backward facing step induces more turbulent mixing between the coolant and mainstream flows, thus reducing film effectiveness coverage and increasing Nusselt number by nearly 40% in the passage throat. In contrast, the presence of a forward step creates a more stable boundary layer for the coolant flow aiding to help keep the film attached to the endwall. Increasing the blowing ratio increases film-cooling effectiveness and endwall coverage up to a certain point, beyond which, the high momentum of the coolant results in poor cooling performance due to jet liftoff.

Effects of an Axisymmetric Contoured Endwall on a Nozzle Guide Vane: Adiabatic Effectiveness Measurements

2010 ◽  
Author(s):  
A. A. Thrift ◽  
K. A. Thole ◽  
S. Hada
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Cooling Effectiveness ◽  
Flow Rates ◽  
Guide Vanes ◽  
Film Cooling Effectiveness ◽  
Adiabatic Effectiveness ◽  
Improved Performance ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

Gas turbine designs seek improved performance by modifying the endwalls of nozzle guide vanes in the engine hot section. Within the nozzle guide vanes these modifications can be in the form of an axisymmetric contour as the area contracts from the combustor to the turbine. This paper investigates the effect of axisymmetric endwall contouring on the cooling performance of a film cooled endwall. Adiabatic effectiveness measurements were performed in a planar passage for comparison to a contoured passage whereby the exit Reynolds numbers was matched. For the contoured passage, measurements were performed on both the flat endwall and on the contoured endwall. Fully expanded film cooling holes were distributed on the endwall surface preceded by a two-dimensional slot normal to the inlet axis. Results indicated that the coolant coverage from the upstream leakage slot was spread over a larger area of the contoured endwall in comparison to the flat endwall of the planar passage. Film cooling effectiveness on the flat endwall of the contoured passage showed minimal differences relative to the planar passage results. The contracting endwall of the contoured passage, however, showed a significant reduction with average film cooling effectiveness levels approximately 40% lower than the planar passage at low film cooling flow rates. In the case of all endwalls, increasing leakage and film cooling mass flow rates led to an increase in cooling effectiveness and coolant coverage.

The Effect of Step Misalignment on Purge Flow Cooling of Nozzle Guide Vane at Transonic Conditions

2019 ◽  
Author(s):  
Luke Luehr ◽  
Ridge Sibold ◽  
Shuo Mao ◽  
Wing F. Ng ◽  
Zhigang Li ◽  
...  
Keyword(s):  
Nusselt Number ◽  
Film Cooling ◽  
Guide Vane ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Purge Flow ◽  
Exit Mach Number ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

Abstract This study describes a detailed investigation on the effects that step misalignment and upstream purge film cooling have on the endwall heat transfer for 1st stage nozzle guide vanes in a land-based power generating gas turbine at transonic conditions. Endwall Nusselt number and adiabatic film cooling effectiveness distributions were experimentally measured and compared with flow visualization. Tests were conducted in a transonic linear cascade blowdown facility where data were gathered at an exit Mach number of 0.85 with a freestream turbulence intensity of 16% at an exit Re = 1.5 × 106 based on axial chord. Varied upstream purge blowing ratios and a no blowing case were tested for 3 different upstream step geometries, the baseline (no misalignment), a span-wise upstream step of +4.86% span, and a step of −4.86% span. Experimentation shows that the addition of upstream purge film cooling increases the Nusselt number at injection upwards of 50% but lowers it in the throat of the passage by approximately 20% compared to no blowing case. The addition of a backward facing step induces more turbulent mixing between the coolant and mainstream flows, thus reducing film effectiveness coverage and increasing Nusselt number by nearly 40% in the passage throat. In contrast, the presence of a forward step creates a more stable boundary layer for the coolant flow aiding to help keep the film attached to the endwall at higher blowing ratios. Increasing the blowing ratio increases film cooling effectiveness and endwall coverage up to a certain point, beyond which, the high momentum of the coolant results in poor cooling performance due to jet liftoff.

Effects of an Axisymmetric Contoured Endwall on a Nozzle Guide Vane: Adiabatic Effectiveness Measurements

2011 ◽  
Vol 133 (4) ◽  
Author(s):  
A. A. Thrift ◽  
K. A. Thole ◽  
S. Hada
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Cooling Effectiveness ◽  
Flow Rates ◽  
Guide Vanes ◽  
Film Cooling Effectiveness ◽  
Adiabatic Effectiveness ◽  
Improved Performance ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

Gas turbine designs seek improved performance by modifying the endwalls of nozzle guide vanes in the engine hot section. Within the nozzle guide vanes, these modifications can be in the form of an axisymmetric contour as the area contracts from the combustor to the turbine. This paper investigates the effect of axisymmetric endwall contouring on the cooling performance of a film cooled endwall. Adiabatic effectiveness measurements were performed in a planar passage for comparison to a contoured passage, whereby the exit Reynolds numbers were matched. For the contoured passage, measurements were performed both on the flat endwall and on the contoured endwall. Fully expanded film cooling holes were distributed on the endwall surface preceded by a two-dimensional slot normal to the inlet axis. Results indicated that the coolant coverage from the upstream leakage slot was spread over a larger area of the contoured endwall in comparison to the flat endwall of the planar passage. Film cooling effectiveness on the flat endwall of the contoured passage showed minimal differences relative to the planar passage results. The contracting endwall of the contoured passage, however, showed a significant reduction with average film cooling effectiveness levels approximately 40% lower than the planar passage at low film cooling flow rates. In the case of all endwalls, increasing leakage and film cooling mass flow rates led to an increase in cooling effectiveness and coolant coverage.

Numerical Studies on Combustor-Turbine Interaction at the Large Scale Turbine Rig (LSTR)

2017 ◽  
Author(s):  
Jonathan Hilgert ◽  
Martin Bruschewski ◽  
Holger Werschnik ◽  
Heinz-Peter Schiffer
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Guide Vane ◽  
Turbulence Models ◽  
Detailed Examination ◽  
Computational Domain ◽  
Complex Flow ◽  
Lean Burn ◽  
Swirl Generator ◽  
Nozzle Guide

In order to fully understand the physical behavior of lean burn combustors and its influence on high pressure turbine stages in modern jet engines, the use of Computational Fluid Dynamics (CFD) promises to be a valuable addition to experimental techniques. The numerical investigations of this paper are based on the Large Scale Turbine Rig (LSTR) at Technische Universität Darmstadt, Germany which has been set up to explore the aerothermal combustor turbine interaction. The underlying numerical grids of the simulations take account of the complex cooling design to the fullest extent, considering coolant cavities, cooling holes and vane trailing edge slots within the meshing process. In addition to the k-ω-SST turbulence model, Scale-Adaptive Simulation (SAS) is applied for a computational domain comprising swirl generator and nozzle guide vanes in order to overcome the shortcomings of eddy viscosity turbulence models with regard to streamline curvature. The numerical results are compared with Five Hole Probe measurements at different streamwise locations showing good agreement and allowing for a more detailed examination of the complex flow physics caused by the interaction of turbine flow with lean-burn combustion and advanced film-cooling concepts. Moreover, numerically predicted Nu-contours on the hub end wall of the nozzle guide vane are validated by means of Infrared Thermography measurements.

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