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.

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 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.

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.

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.

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

2020 ◽  
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 MFR (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.

Control of Secondary Flows in a Turbine Nozzle Guide Vane by Endwall Contouring

2000 ◽  
Author(s):  
T. I-P. Shih ◽  
Y.-L. Lin ◽  
T. W. Simon
Keyword(s):  
Film Cooling ◽  
Computational Study ◽  
Guide Vane ◽  
Three Dimensional ◽  
Secondary Flows ◽  
Navier Stokes ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide Vane ◽  
A Cell ◽  
Nozzle Guide

Computations were performed to study the three-dimensional flow and temperature distribution in a nozzle guide vane that has one flat and one contoured endwall with and without film cooling injected from two slots, one on each endwall located just upstream of the airfoil. For the contoured endwall, two locations of the same contouring were investigated, one with all contouring upstream of the airfoil and another with the contouring starting upstream of the airfoil and continuing through the airfoil passage. Results obtained show that when the contouring is all upstream of the airfoil, secondary flows on both the flat and the contoured endwalls are similar in magnitude. When the contouring starts upstream of the airfoil and continues through the airfoil passage, secondary flows on the contoured endwall are markedly weaker than those on the flat endwall. With weaker secondary flows on the contoured endwall, film-cooling effectiveness there is greatly improved. This computational study is based on the ensemble-averaged conservation equations of mass, momentum (compressible Navier-Stokes), and energy. Effects of turbulence were modeled by the low Reynolds number shear-stress transport k-ω model. Solutions were generated by a cell-centered, finite-volume method that uses third-order accurate flux-difference splitting of Roe with limiters and multigrid acceleration of a diagonalized ADI scheme with local time stepping on patched/embedded structured grids.

Film Cooling Experiments With Flow Introduced Upstream of a First Stage Nozzle Guide Vane Through Slots of Various Geometries

2002 ◽  
Author(s):  
Rohit A. Oke ◽  
Terrence W. Simon
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Suction Side ◽  
Main Flow ◽  
Secondary Flows ◽  
Base Line ◽  
Cooling Flow ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

This paper describes the advantages of introducing film cooling flow through the endwall upstream of the first stage nozzle guide vane. To perform these studies, a linear cascade is built. It consists of three vanes and two endwalls that form two passages. One endwall is flat and the other is contoured from upstream of the leading edge, continuing through the passage. The approach flow is of high turbulence and large length scale, representative of the engine combustor exit flow. Film cooling flow is introduced through two successive rows of slots, a single row of slots and slots that have particular area distributions in the pitchwise direction. Measurements are taken by heating the film cooling flow slightly above the main flow temperature and recording temperature distributions in the film cooling flow-main flow mixing zone at various axial planes. The single and double slot injection cases represent base-line injection geometries. They show that at lower ratios of coolant to main flow momentum fluxes, film cooling flow migrates toward the suction side due to secondary flow. At higher ratios, the pressure side endwall region is cooled more effectively. Observations are drawn by comparing the baseline injection cases with cases of different geometries for which slots are blocked partially to re-distribute mass and momentum injection rates of the emerging flow. The downstream evolutions of temperature contours are discussed. The idea is to utilize secondary flows to control pitchwise coolant distributions.

Numerical Analysis on Effects of the Contoured Endwall on the Three Dimensional Flow Characteristics in an Annular Turbine Nozzle Vane

2007 ◽  
Author(s):  
Wu Sang Lee ◽  
Jin Taek Chung ◽  
Dae Hyun Kim ◽  
Seung Joo Choe
Keyword(s):  
Guide Vane ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Cross Flow ◽  
Leading Edge ◽  
Secondary Flows ◽  
Three Dimensional Flow ◽  
Dimensional Flow ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

The three-dimensional flow in a turbine nozzle guide vane passage causes large secondary loss through the passage and increased heat transfer on the blade surface. In order to reduce or control these secondary flows, a linear turbine with contoured endwall configurations was used and changes in the three-dimensional flow field were analyzed and discussed. Contoured endwalls are installed at a location downstream of the saddle point near the leading edge of the pressure side blade and several positions along the centerline of the passage at constant distance. The objective of this study is to document the development of the three-dimensional flow in a turbine nozzle guide vane cascade with modified endwall. In addition, it proposes and appropriates endwall contouring which shows best overall loss reduction performance among the simulated contoured endwall. The results of this study show that the development of passage vortex and cross flow in the cascade composed of one flat and one contoured endwalls are affected by the acceleration which occurs in contoured endwall side. The overall loss is reduced near the flat endwall rather than contoured endwall, the best performance was shown for the case of 10–15% contoured for span-wise, 40–70% length of chord from trailing edge.

Aerodynamic Measurements and Analysis in a First Stage Nozzle Guide Vane Passage With Combustor Liner Cooling, Slot Film Cooling and Endwall Contouring

2018 ◽  
Author(s):  
Mahmood H. Alqefl ◽  
Yong W. Kim ◽  
Hee-Koo Moon ◽  
Luzeng Zhang ◽  
Terrence W. Simon
Keyword(s):  
Secondary Flow ◽  
Flow Rate ◽  
Film Cooling ◽  
Guide Vane ◽  
Secondary Flows ◽  
Cooling Flows ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Slot Film Cooling ◽  
Combustor Liner

Endwalls impose a challenge to cool because of the complex system of secondary flows and separation lines disrupting surface film coolant coverage. The interaction of film cooling flows with secondary flow structures is coupled. The momentum exchange of the film coolant with the mainstream affect the formation the secondary flows, which in turn affect the coolant coverage. Therefore, to develop better endwall cooling schemes, a good understanding of passage aerodynamics as affected by interactions with coolant flows is required. This study presents experimental and computational results for cascade representing the first stage nozzle guide vane of a high-pressure gas turbine. The cascade is subsonic, linear, and stationary with an axisymmetrically-contoured endwall. Two cooling flows are simulated; upstream combustor liner coolant-in the form of an aero-thermal profile simulated in the approach flow and endwall slot film cooling, which is injected immediately upstream of the passage inlet. The experiment is run with engine representative combustor exit flow turbulence intensity and integral length scales, with high turbine passage exit Reynolds number of 1.61 × 106. Measurements are performed with various slot film cooling mass flow rate to mainstream flow rate ratios (MFR). Aerodynamic effects are documented with five-hole probe measurements at the exit plane. Varying the slot film cooling MFR results in minimal effects on total pressure loss for the range tested. Vorticity distributions show a very thin, yet intense, cross-pitch flow on the contoured endwall side. Coolant distribution fields that were previously presented for the same cascade are discussed in context of the aerodynamic measurements. A coolant vorticity parameter presenting the advective mixing of the coolant due to secondary flow vorticity is introduced. This parameter gives developers a new prospective on aerodynamic-thermal performance associated with cooled turbine endwall. The numerical study is conducted for the same test section geometry and is run under the same conditions. The applicability of using RANS turbulence closure models for simulating this type of flow is discussed. The effects of including the combustor coolant in the approach flow is also briefly discussed in context of the numerical results.

Influence of Pre-History and Leading Edge Contouring on Aero-Performance of a 3D Nozzle Guide Vane

2013 ◽  
Author(s):  
Ranjan Saha ◽  
Boris I. Mamaev ◽  
Jens Fridh ◽  
Björn Laumert ◽  
Torsten H. Fransson
Keyword(s):  
Guide Vane ◽  
Three Dimensional ◽  
Leading Edge ◽  
Stagnation Line ◽  
Wide Range ◽  
Nozzle Guide Vane ◽  
Turbine Vane ◽  
Loss Coefficients ◽  
Exit Mach Number ◽  
Nozzle Guide

Experiments are conducted to investigate the effect of the pre-history in the aerodynamic performance of a three-dimensional nozzle guide vane with a hub leading edge contouring. The performance is determined with two pneumatic probes (5 hole and 3 hole) concentrating mainly on the endwall. The investigated vane is a geometrically similar gas turbine vane for the first stage with a reference exit Mach number of 0.9. Results are compared for the baseline and filleted cases for a wide range of operating exit Mach numbers from 0.5 to 0.9. The presented data includes loading distributions, loss distributions, fields of exit flow angles, velocity vector and vorticity contour, as well as, mass-averaged loss coefficients. The results show an insignificant influence of the leading edge fillet on the performance of the vane. However, the pre-history (inlet condition) affects significantly in the secondary loss. Additionally, an oil visualization technique yields information about the streamlines on the solid vane surface which allows identifying the locations of secondary flow vortices, stagnation line and saddle point.

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