Effect of the Combustor Wall on the Aerothermal Field of a Nozzle Guide Vane

2018 ◽  
Vol 140 (5) ◽  
Author(s):  
Ioanna Aslanidou ◽  
Budimir Rosic
Keyword(s):  
Gas Turbines ◽  
High Speed ◽  
Guide Vane ◽  
Leading Edge ◽  
Flow Structures ◽  
Experimental Facility ◽  
Aerodynamic Loss ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Processing Techniques

In gas turbines with can combustors, the trailing edge (TE) of the combustor transition duct wall is found upstream of every second vane. This paper presents an experimental and numerical investigation of the effect of the combustor wall TE on the aerothermal performance of the nozzle guide vane. In the measurements carried out in a high-speed experimental facility, the wake of this wall is shown to increase the aerodynamic loss of the vane. On the other hand, the wall alters secondary flow structures and has a protective effect on the heat transfer in the leading edge-endwall junction, a critical region for component life. The different clocking positions of the vane relative to the combustor wall are tested experimentally and are shown to alter the aerothermal field. The experimental methods and processing techniques adopted in this work are used to highlight the differences between the different cases studied.

Aerothermal Performance of Shielded Vane Design

2017 ◽  
Vol 139 (11) ◽  
Author(s):  
Ioanna Aslanidou ◽  
Budimir Rosic
Keyword(s):  
Heat Transfer ◽  
Total Pressure ◽  
High Speed ◽  
Guide Vane ◽  
Leading Edge ◽  
Experimental Facility ◽  
Numerical Studies ◽  
Nozzle Guide Vane ◽  
Inlet Conditions ◽  
Nozzle Guide

This paper presents an experimental investigation of the concept of using the combustor transition duct wall to shield the nozzle guide vane leading edge. The new vane is tested in a high-speed experimental facility, demonstrating the improved aerodynamic and thermal performance of the shielded vane. The new design is shown to have a lower average total pressure loss than the original vane, and the heat transfer on the vane surface is overall reduced. The peak heat transfer on the vane leading edge–endwall junction is moved further upstream, to a region that can be effectively cooled as shown in previously published numerical studies. Experimental results under engine-representative inlet conditions showed that the better performance of the shielded vane is maintained under a variety of inlet conditions.

A New Experimental Facility to Investigate Combustor-Turbine Interactions in Gas Turbines With Multiple Can Combustors

2014 ◽  
Author(s):  
S. Luque ◽  
V. Kanjirakkad ◽  
I. Aslanidou ◽  
R. Lubbock ◽  
B. Rosic ◽  
...  
Keyword(s):  
Gas Turbines ◽  
High Speed ◽  
Cooling System ◽  
Guide Vane ◽  
Experimental Facility ◽  
Blow Down ◽  
Guide Vanes ◽  
Wide Range ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

This paper describes a new modular experimental facility that was purpose-built to investigate flow interactions between the combustor and first stage nozzle guide vanes of heavy duty power generation gas turbines with multiple can combustors. The first stage turbine nozzle guide vane is subjected to the highest thermal loads of all turbine components and therefore consumes a proportionally large amount of cooling air that contributes detrimentally to the stage and cycle efficiency. It has become necessary to devise novel cooling concepts that can substantially reduce the coolant air requirement but still allow the turbine to maintain its aerothermal performance. The present work aims to aid this objective by the design and commissioning of a high-speed linear cascade which consists of two can combustor transition ducts and four first stage nozzle guide vanes. This is a modular non-reactive air test platform with engine realistic geometries (gas path and near gas path), cooling system, and boundary conditions (inlet swirl, turbulence level and boundary layer). The paper presents the various design aspects of the high pressure blow down type facility, and the initial results from a wide range of aerodynamic and heat transfer measurements under highly engine realistic conditions.

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.

scholarly journals Wake Region Measurements of a Highly Three-Dimensional Nozzle Guide Vane Tested at DRA Pyestock Using Particle Image Velocimetry

1994 ◽  
Author(s):  
M. Funes-Gallanzi ◽  
P. J. Bryanston-Cross ◽  
K. S. Chana
Keyword(s):  
Flow Field ◽  
High Speed ◽  
Guide Vane ◽  
Three Dimensional ◽  
Light Sheet ◽  
Test Facility ◽  
Velocity Mapping ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Annular Cascade

The quantitative whole field flow visualization technique of PIV has over the last few years been successfully demonstrated for transonic flow applications. A series of such measurements has been made at DRA Pyestock. Several of the development stages critical to a full engine application of the work have now been achieved using the Isentropic Light Piston Cascade (ILPC) test facility operating with high inlet turbulence levels: • A method of seeding the flow with 0.5μm diameter styrene particles has provided an even coverage of the flow field. • A method of projecting a 1 mm thick high power Nd/YAG laser light sheet within the turbine stator cascade. This has enabled a complete instantaneous intra-blade velocity mapping of the flow field to be visualized, by a specially developed diffraction-limited optics arrangement. • Software has been developed to automatically analyze the data. Due to the sparse nature of the data obtained, a spatial approach to the extraction of the velocity vector data was employed. • Finally, a comparison of the experimental results with those obtained from a three-dimensional viscous flow program of Dawes; using the Baldwin-Lomax model for eddy viscosity and assuming fully turbulent flow. The measurements provide an instantaneous quantitative whole field visualization of a high-speed unsteady region of flow in a highly three-dimensional nozzle guide vane; which has been successfully compared with a full viscous calculation. This work represents the first such measurements to be made in a full-size transonic annular cascade at engine representative conditions.

Heat Transfer Studies on the Interior Surfaces of Cooled Nozzle Guide Vane in a Linear Cascade

2015 ◽  
Author(s):  
Arun Kumar Pujari ◽  
Bhamidi Prasad ◽  
Nekkanti Sitaram
Keyword(s):  
Heat Transfer ◽  
Surface Temperature ◽  
Guide Vane ◽  
Internal Heat ◽  
Leading Edge ◽  
Internal Surface ◽  
Transfer Coefficients ◽  
Computational Data ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Experimental and computational heat transfer investigations are reported in the interior side of a nozzle guide vane (NGV) subjected to combined impingement and film cooling. The domain of study is a two dimensional five-vane cascade having four passages. Each vane has a chord length of 228 mm and the pitch distance between the vanes is 200 mm. The vane internal surface is cooled by dry air supplied through the two impingement inserts: the front and the aft. The mass flow through the impingement chamber is varied, for a fixed spacing (H) to jet diameter (d) ratio of 1.2. The surface temperature distributions, at certain locations of the vane interior, are measured by pasting strips of liquid crystal sheets. The vane interior surface temperature distribution is also obtained by computations carried out by using Shear stress transport (SST) k-ω turbulence model in the ANSY FLUENT-14 flow solver. The computational data are in good agreement with the measured values of temperature. The internal heat transfer coefficients are thence determined along the leading edge and the mid span region from the computational data.

Effects of an Axisymmetric Contoured Endwall on a Nozzle Guide Vane: Convective Heat Transfer Measurements

2010 ◽  
Author(s):  
A. A. Thrift ◽  
K. A. Thole ◽  
S. Hada
Keyword(s):  
Heat Transfer ◽  
Guide Vane ◽  
Leading Edge ◽  
Critical Factor ◽  
Suction Side ◽  
Leakage Flow ◽  
Cross Sectional ◽  
Nozzle Guide Vane ◽  
Linear Slope ◽  
Nozzle Guide

Heat transfer is a critical factor in the durability of gas turbine components, particularly in the first vane. An axisymmetric contour is sometimes used to contract the cross sectional area from the combustor to the first stage vane in the turbine. Such contouring can lead to significant changes in the endwall flows thereby altering the heat transfer. This paper investigates the effect of axisymmetric contouring on endwall heat transfer of a nozzle guide vane. Heat transfer measurements are performed on the endwalls of a planar and contoured passage whereby one endwall is modified with a linear slope in the case of the contoured passage. Included in this study is upstream leakage flow issuing from a slot normal to the inlet axis. Each of the endwalls within the contoured passage presents a unique flowfield. For the contoured passage, the flat endwall is subject to an increased acceleration through the area contraction while the contoured endwall includes both increased acceleration and a turning of streamlines due to the slope. Results indicate heat transfer is reduced on both endwalls of the contoured passage relative to the planar passage. In the case of all endwalls, increasing leakage mass flow rate leads to an increase in heat transfer near the suction side of the vane leading edge.

Nozzle Passage Endwall Effectiveness Values With Various Combustor Coolant Flowrates—Part 2: Endwall and Vicinity Surface Effectiveness Measurements

2021 ◽  
Vol 143 (4) ◽  
Author(s):  
Kedar P. Nawathe ◽  
Rui Zhu ◽  
Enci Lin ◽  
Yong W. Kim ◽  
Terrence W. Simon
Keyword(s):  
Gas Turbines ◽  
Guide Vane ◽  
Suction Surface ◽  
Turbine Engine ◽  
Pressure Surface ◽  
Coolant Injection ◽  
Measurement Results ◽  
Nozzle Guide Vane ◽  
Effectiveness Measurement ◽  
Nozzle Guide

Abstract Effective coolant schemes are required for providing cooling to the first-stage stator vanes of gas turbines. To correctly predict coolant performance on the endwall and vane surfaces, these coolant schemes should also consider the effects of coolant streams introduced upstream in the combustor section of a gas turbine engine. This two-part paper presents measurements taken on a first-stage nozzle guide vane cascade that includes combustor coolant injection. The first part of this paper explains how coolant transport and coolant-mainstream interaction in the vane passage is affected by changing the combustor coolant and endwall film coolant flowrates. This paper explains how those flows affect the coolant effectiveness on the endwall and vane surfaces. Part one showed that a significant amount of coolant injected upstream of the endwall is present along the pressure surface of the vanes as well as over the endwall. Part two shows effectiveness measurement results taken in this study on the endwall and pressure and suction surfaces of the vanes. Sustained endwall coolant effectiveness is observed along the whole passage for all cases. It is uniform in the pitch-wise direction. Combustor coolant flow significantly affects cooling performance even near the trailing edge. The modified flowfield results in the pressure surface being cooled more effectively than the suction surface. While the effectiveness distribution on the pressure surface varies with combustor and film coolant flowrates, the distribution along the suction surface remains largely unchanged.

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.

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.

A new experimental approach for heat transfer coefficient and adiabatic wall temperature measurements on a Nozzle Guide Vane with inlet temperature distortions

2021 ◽  
pp. 1-33
Author(s):  
Tommaso Bacci ◽  
Alessio Picchi ◽  
Bruno Facchini ◽  
Simone Cubeda
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Wall Temperature ◽  
Gas Turbines ◽  
Experimental Approach ◽  
Guide Vane ◽  
Adiabatic Wall ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Abstract Modern gas turbines lean combustors are used to limit NOx pollutant emissions; on the other hand, their adoption presents other challenges, especially concerning the combustor-turbine interaction. Turbine inlet conditions are generally characterized by severe temperature distortions and swirl degree, which is responsible for very high turbulence intensities. Past studies have focused on the description of the effects of these phenomena on the behavior of the high pressure turbine. Nevertheless, very limited experimental results are available when it comes to evaluate the heat transfer coefficient (HTC) on the nozzle guide vane surface, since relevant temperature distortions present a severe challenge for the commonly adopted measurement techniques. The work presented in this paper was carried out on a non-reactive, annular, three-sector rig, made by a combustor simulator and a NGV cascade. It can reproduce a swirling flow, with temperature distortions at the combustor-turbine interface plane. This test apparatus was exploited to develop an experimental approach to retrieve heat transfer coefficient and adiabatic wall temperature distributions simultaneously, to overcome the known limitations imposed by temperature gradients on state-of-the-art methods for HTC calculation from transient tests. A non-cooled mockup of a NGV doublet, manufactured using low thermal diffusivity plastic material, was used for the tests, carried out using IR thermography with a transient approach. In the authors' knowledge, this presents the first experimental attempt of measuring a nozzle guide vane heat transfer coefficient in the presence of relevant temperature distortions and swirl.

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