The Effects of Combustor Cooling Features on Nozzle Guide Vane Film Cooling Experiments

2018 ◽  
Vol 141 (1) ◽  
Author(s):  
Nicholas E. Holgate ◽  
Peter T. Ireland ◽  
Eduardo Romero
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Stagnation Region ◽  
Airfoil Surface ◽  
Cooling Performance ◽  
Nozzle Guide Vane ◽  
Effectiveness Measurement ◽  
Extraneous Radiation ◽  
Nozzle Guide ◽  
Endwall Film Cooling

Abstract Recent advances in experimental methods have allowed researchers to study nozzle guide vane (NGV) film cooling in the presence of combustor dilution ports and endwall films. The dilution injection creates nonuniformities in temperature, velocity, and turbulence, and an understanding of the vane film cooling performance is complicated by competing influences. In this study, dilution port temperature profiles have been measured in the absence of vane film cooling and compared to film effectiveness measurements in the presence of both films and dilution, illustrating the effects of the dilution port turbulence on film cooling performance. It is found that dilution port injection can create significant effectiveness benefits at the difficult-to-cool vane stagnation region due to the more turbulent hot mainstream enhancing the mixing of film coolant jets that have left the airfoil surface. Also explored are the implications of endwall film cooling for infrared (IR) vane surface temperature measurements. The reduced endwall temperatures reduce the thermal emissions from this surface, so reducing the amount of extraneous radiation reflected from the vane surface where measurements are being made. The results of a detailed calibration show that the maximum local film effectiveness measurement error could be up to 0.05 if this effect were to go unaccounted for.

scholarly journals The Effects of Combustor Cooling Features on Nozzle Guide Vane Film Cooling Experiments

2018 ◽  
Author(s):  
Nicholas E. Holgate ◽  
Peter T. Ireland ◽  
Eduardo Romero
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Stagnation Region ◽  
Airfoil Surface ◽  
Cooling Performance ◽  
Nozzle Guide Vane ◽  
Effectiveness Measurement ◽  
Extraneous Radiation ◽  
Nozzle Guide ◽  
Endwall Film Cooling

Recent advances in experimental methods have allowed researchers to study nozzle guide vane film cooling in the presence of combustor dilution ports and endwall films. The dilution injection creates nonuniformities in temperature, velocity, and turbulence, and an understanding of the vane film cooling performance is complicated by competing influences. In this study, dilution port temperature profiles have been measured in the absence of vane film cooling and compared to film effectiveness measurements in the presence of both films and dilution, illustrating the effects of the dilution port turbulence on film cooling performance. It is found that dilution port injection can create significant effectiveness benefits at the difficult-to-cool vane stagnation region, due to the more turbulent hot mainstream enhancing the mixing of film coolant jets that have left the airfoil surface. Also explored are the implications of endwall film cooling for infrared vane surface temperature measurements. The reduced endwall temperatures reduce the thermal emissions from this surface, so reducing the amount of extraneous radiation reflected from the vane surface where measurements are being made. The results of a detailed calibration show that the maximum local film effectiveness measurement error could be up to 0.05 if this effect were to go unaccounted for.

Aero-Thermal Aspects of Film Cooled Nozzle Guide Vane Endwalls: Part 2 — Thermal Measurements

2020 ◽  
Author(s):  
Mahmood H. Alqefl ◽  
Kedar P. Nawathe ◽  
Pingting Chen ◽  
Rui Zhu ◽  
Yong W. Kim ◽  
...  
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Three Dimensional ◽  
Secondary Flows ◽  
Complex Flow ◽  
Cooling Flows ◽  
Thermal Measurements ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Endwall Film Cooling

Abstract Flow over gas turbine endwalls is complex and highly three-dimensional. As boundaries for modern engine designs are pushed, this already-complex flow is affected by aggressive application of film cooling flows that actively interact. This two-part study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first stage nozzle guide vane passage. The approach flow conditions represent flow exiting a low-NOx combustor. The test section includes geometric and cooling details of a combustor-turbine interface in addition to endwall film cooling flows injected upstream of the passage. The first part of this study describes in detail, the passage aerodynamics as affected by injection of cooling flows. It reveals a system of secondary flows, including the newly-discovered Impingement Vortex, which redefines our understanding of the aerodynamics of flow in a modern, film-cooled, first-stage vane row. The second part investigates, through thermal measurements, the distribution, mixing and disruption of cooling flows over the endwall. Measurements are made with and without active endwall film cooling. Descriptions are made through adiabatic surface effectiveness measurements and correlations with in-passage velocity (presented in part one) and thermal fields. Results show that the newly-discovered impingement vortex has a positive effect on coolant distribution through passage vortex suppression and by carrying the coolant to hard-to-cool regions in the passage, including the pressure surface near the endwall.

Effects of Endwall Film Coolant Flowrate on Secondary Flows and Coolant Mixing in a First Stage Nozzle Guide Vane

2021 ◽  
Vol 143 (3) ◽  
Author(s):  
Mahmood H. Alqefl ◽  
Kedar P. Nawathe ◽  
Pingting Chen ◽  
Rui Zhu ◽  
Yong W. Kim ◽  
...  
Keyword(s):  
Test Section ◽  
Film Cooling ◽  
Gas Turbines ◽  
Thermal Loading ◽  
Guide Vane ◽  
Secondary Flows ◽  
Cooling Flows ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Endwall Film Cooling

Abstract Modern gas turbines are subjected to very high thermal loading. This leads to a need for aggressive cooling to protect components from damage. Endwalls are particularly challenging to cool due to a complex system of secondary flows near them that wash and disrupt the protective coolant films. This highly three-dimensional flow not only affects but is also affected by the momentum of film cooling flows, whether injected just upstream of the passage to intentionally cool the endwall or as combustor cooling flows injected further upstream in the engine. This complex interaction between the different cooling flows and passage aerodynamics has been recently studied in a first stage nozzle guide vane. The present paper presents a detailed study on the sensitivity of aero-thermal interactions to endwall film cooling mass flow to mainstream flow ratio. The test section represents a first stage nozzle guide vane with a contoured endwall and endwall film cooling injected just upstream of it. The test section also includes an engine-representative combustor–turbine interface geometry with combustor cooling flows injected at a constant rate. The approach flow conditions represent flow exiting a low-NOx combustor. Adiabatic surface thermal measurements and in-passage velocity and thermal field measurements are presented and discussed. The results show the dynamics of passage vortex suppression and the increase of impingement vortex strength as MFR changes. The effects of these changes of secondary flows on coolant distribution are presented.

Comparison of Film Cooling Performance for Different Purge Slot Configurations in a Cylindrical and State-of-the-Art Nozzle Guide Vane

2021 ◽  
Author(s):  
Christian Landfester ◽  
Gunther Müller ◽  
Robert Krewinkel ◽  
Clemens Domnick ◽  
Martin Böhle
Keyword(s):  
Secondary Flow ◽  
Film Cooling ◽  
State Of The Art ◽  
Guide Vane ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Abstract This comparative study is concerned with the advances in nozzle guide vane (NGV) design developments and their influence on the film cooling performance by injecting coolant through the purge slot. An experimental study compares the film cooling effectiveness as well as the aerodynamic effects for different purge slot configurations on both a flat and an axisymmetrically contoured endwall of a NGV. While the flat endwall cascade was equipped with four cylindrical vanes, the contoured endwall cascade consisted of four modern NGVs which represent state-of-the-art high-pressure turbine design standards. Geometric variations, e.g. the purge slot width and injection angle, as well as different blowing ratios (BR) at an engine-like density ratio (DR = 1.6) were realized to investigate the real-life effect of thermal expansion, design modifications and the interaction between secondary flow and coolant. The mainstream flow parameters were set to meet real engine conditions with regard to Reynolds and Mach numbers. The Pressure Sensitive Paint (PSP) technique was used to determine the adiabatic film cooling effectiveness. Five-hole probe measurements (DR = 1.0) were performed to measure the flow field with its characteristic vortex structures as well as the loss distribution in the vane wake region. For a more profound insight into the origin and development of the secondary flows, oil dye visualizations were carried out on both endwalls. The measurement results will be discussed based on a side-by-side comparison of the distribution of film cooling effectiveness on the endwall, its area-averaged values as well as the two-dimensional distribution of total pressure losses and the secondary flow field. The results of this study show that the advances in NGV design development have had a significantly positive influence on the distribution of the coolant. This has to be attributed to lesser disturbance of the coolant propagation by secondary flow for the optimized NGV design, since the design features are intended to suppress the formation of secondary flow. In contrast to the results of the cylindrical profile, sufficient cooling can be already provided with a perpendicular injection in the case of the modern NGV. It is therefore advisable to take these effects into account when designing the film cooling system of a modern high-pressure turbine.

Aerodynamic and Endwall Film-Cooling Investigations of a Gas Turbine Nozzle Guide Vane Applying Temperature-Sensitive Paint

2011 ◽  
Vol 133 (3) ◽  
Author(s):  
Martin Kunze ◽  
Konrad Vogeler ◽  
Glenn Brown ◽  
Chander Prakash ◽  
Kenneth Landis
Keyword(s):  
Heat Conduction ◽  
Gas Turbine ◽  
Film Cooling ◽  
Guide Vane ◽  
Temperature Sensitive ◽  
Test Plate ◽  
Nozzle Guide Vane ◽  
Temperature Sensitive Paint ◽  
Nozzle Guide ◽  
Endwall Film Cooling

Endwall film-cooling investigations are conducted with a single row of fan-shaped holes in a low-speed, six-bladed linear cascade. The incidence of the inlet flow was changed between −5 deg and 40 deg to achieve higher loading conditions, which results in an intensification of the secondary flow and enhanced interaction with the injected coolant. The investigated profile is based on a near-hub section of the nozzle guide vane of a highly loaded gas turbine. The aerodynamic performance was investigated using pneumatic probes. The film-cooling effectiveness distribution is determined using the temperature-sensitive paint technique. Carbon dioxide was used as coolant to provide elevated density ratios of about 1.4. Although low thermal conductivity material is used for the endwall test plate, the measured temperature fields show influences of 3D-heat conduction inside the test plate. To measure film effectiveness and the heat transfer separately, an adiabatic test surface is needed. Therefore, the effects of heat conduction are modeled using the finite-element-method. With the resulting convective heat flux pattern derived from the computations, the endwall film-cooling measurements are corrected. Furthermore, this approach is applied to evaluate the heat loss inside the holes and the film discharge temperature at the hole exit.

AERO-THERMAL ASPECTS OF FILM COOLED NOZZLE GUIDE VANE ENDWALLS - PART 2: THERMAL MEASUREMENTS

2021 ◽  
pp. 1-39
Author(s):  
Mahmood H. Alqefl ◽  
Kedar P. Nawathe ◽  
Pingting Chen ◽  
Rui Zhu ◽  
Yong Kim ◽  
...  
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Three Dimensional ◽  
Secondary Flows ◽  
Complex Flow ◽  
Cooling Flows ◽  
Thermal Measurements ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Endwall Film Cooling

Abstract Flow over gas turbine endwalls is complex and highly three-dimensional. As boundaries for modern engine designs are pushed, this already-complex flow is affected by aggressive application of film cooling flows that actively interact. This two-part study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first stage nozzle guide vane passage. The approach flow conditions represent flow exiting a low-NOx combustor. The test section includes geometric and cooling details of a combustor-turbine interface in addition to endwall film cooling flows injected upstream of the passage. The first part of this study describes in detail, the passage aerodynamics as affected by injection of cooling flows. It reveals a system of secondary flows, including the newly-discovered Impingement Vortex, which redefines our understanding of the aerodynamics of flow in a modern, film-cooled, first-stage vane row. The second part investigates, through thermal measurements, the distribution, mixing and disruption of cooling flows over the endwall. Measurements are made with and without active endwall film cooling. Descriptions are made through adiabatic surface effectiveness measurements and correlations with in-passage velocity (presented in part one) and thermal fields. Results show that the newly-discovered impingement vortex has a positive effect on coolant distribution through passage vortex suppression and by carrying the coolant to hard-to-cool regions in the passage, including the pressure surface near the endwall.

Aero-Thermal Aspects of Film Cooled Nozzle Guide Vane Endwalls: Part 1 — Aerodynamics

2020 ◽  
Author(s):  
Mahmood H. Alqefl ◽  
Kedar P. Nawathe ◽  
Pingting Chen ◽  
Rui Zhu ◽  
Yong W. Kim ◽  
...  
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Three Dimensional ◽  
Leading Edge ◽  
Secondary Flows ◽  
Protective Film ◽  
Cooling Flows ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Endwall Film Cooling

Abstract The first stage turbine of a modern gas turbine is subjected to high thermal loads which lead to a need for aggressive cooling schemes to protect its components from melting. Endwalls are particularly challenging to cool due to the complex system of secondary flows near them that wash the protective film coolants into the mainstream. This paper shows that without including combustor cooling, the complex secondary flow physics are not representative of modern engines. Aggressive injection of all cooling flows upstream of the passage is expected to interact and change passage aerodynamics and, subsequently, mixing and transport of coolants. This study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first stage nozzle guide vane passage. The test section involves an engine-representative combustor-turbine interface geometry, combustor coolant flow and endwall film cooling flow injected upstream of a linear cascade. The approach flow conditions represent flow exiting a cooled, low-NOx combustor. This first part of this two-part study aims to understand the complex aerodynamics near the endwall through detailed measurements of passage three-dimensional velocity fields with and without endwall film cooling. The aerodynamic measurements reveal a dominant vortex in the passage, named here as the Impingement Vortex, that opposes the passage vortex formed at the airfoil leading edge plane. This Impingement Vortex completely changes our description of flow over a modern film cooled endwall.

Aero-thermal Aspects of Film Cooled Nozzle Guide Vane Endwall - Part 1: Aerodynamics

2021 ◽  
pp. 1-54
Author(s):  
Mahmood H. Alqefl ◽  
Kedar P. Nawathe ◽  
Pingting Chen ◽  
Rui Zhu ◽  
Yong W. Kim ◽  
...  
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Three Dimensional ◽  
Leading Edge ◽  
Secondary Flows ◽  
Protective Film ◽  
Cooling Flows ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Endwall Film Cooling

Abstract The first stage turbine of a modern gas turbine is subjected to high thermal loads which lead to a need for aggressive cooling schemes to protect its components from melting. Endwalls are particularly challenging to cool due to the complex system of secondary flows near them that wash the protective film coolants into the mainstream. This paper shows that without including combustor cooling, the complex secondary flow physics are not representative of modern engines. Aggressive injection of all cooling flows upstream of the passage is expected to interact and change passage aerodynamics and, subsequently, mixing and transport of coolants. This study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first stage nozzle guide vane passage. The test section involves an engine-representative combustor-turbine interface geometry, combustor coolant flow and endwall film cooling flow injected upstream of a linear cascade. The approach flow conditions represent flow exiting a cooled, low-NOx combustor. This first part of this two-part study aims to understand the complex aerodynamics near the endwall through detailed measurements of passage three-dimensional velocity fields with and without endwall film cooling. The aerodynamic measurements reveal a dominant vortex in the passage, named here as the Impingement Vortex, that opposes the passage vortex formed at the airfoil leading edge plane. This Impingement Vortex completely changes our description of flow over a modern film cooled endwall.

Nozzle Guide Vane Film Cooling Effectiveness for Radial Showerheads With Restricted Cooling Hole Surface Angles

2017 ◽  
Author(s):  
Nicholas E. Holgate ◽  
Peter T. Ireland ◽  
Kevin P. Self
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Stagnation Region ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide Vane ◽  
Cooling Hole ◽  
Leading Edges ◽  
Nozzle Guide ◽  
Surface Angle

Adiabatic film cooling effectiveness measurements are made on nozzle guide vane leading edges in an engine-realistic flow environment. The tested leading edges feature radial showerheads with different spanwise distributions of hole surface angle. The showerheads blow towards the midspan, except for one model with showerhead holes orthogonal to the vane surface. The results show that low surface angle radial showerhead holes generate high effectiveness within their rows and further downstream, but neglect the stagnation region lying between the two most upstream cooling hole rows. This downstream effectiveness gain is due to both the continued surface attachment of this coolant as it progresses downstream, and its beneficial interactions with downstream cooling jets. Moderate radial showerhead surface angles cause moderate coolant jet penetration into the mainstream, which promotes near-surface mixing of the coolant with the mainstream, increasing stagnation region effectiveness. The mixing effect is enhanced by the intense turbulence generated by combustor dilution jets. High surface angles may cause the stagnation region coolant to penetrate too far for either of these gains to be realised. Considering also the presence of endwall film cooling, these effects, taken together, suggest the superiority of radial showerheads which blow towards the midspan, as against those which blow towards each endwall. Surface temperature data is acquired by a novel infrared thermography technique which permits measurement of both heat transfer coefficient and film effectiveness from a single heated test.

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