Determining Total Pressure Fields From Velocimetry Measurements in a Transonic Turbine Flowfield

2021 ◽  
Author(s):  
Alexander Rusted ◽  
Stephen Lynch
Keyword(s):  
Film Cooling ◽  
Compressible Flows ◽  
Guide Vane ◽  
Momentum Equation ◽  
Inlet Guide Vane ◽  
Pressure Fields ◽  
Image Velocimetry ◽  
Turbine Vane ◽  
Total Temperature ◽  
Transonic Turbine

Abstract This work describes a method for calculating pressure fields from temperature and velocity data in non-adiabatic compressible flows, such as the flow around a cooled turbine vane. Prior studies have demonstrated the ability to use particle image velocimetry methods to estimate the pressure gradient in the momentum equation, which is subsequently integrated to produce pressure fields. Due to changes in total temperature for non-adiabatic compressible flows, pressure fields cannot be computed from velocity measurements alone and temperature data must also be provided. In this work, a benchmarked steady 3D RANS simulation is used to generate velocity, temperature, and pressure fields in the transonic flow around a high-pressure turbine inlet guide vane. A procedure for solving the momentum equation and integrating for pressure is developed for non-adiabatic flows. Error is assessed by comparing calculated pressure to CFD predicted pressure, and the effects of PIV spatial resolution and measurement error are considered. The accuracy of the method on non-adiabatic flows is assessed using a vane with extensive film cooling. A clear benefit of incorporating temperature measurements in the pressure determination method is demonstrated, offering opportunities for deeper understanding of aerodynamic losses and entropy generation in cooled turbine flowfields.

scholarly journals Study of Mean- and Turbulent-Velocity Fields in a Large-Scale Turbine-Vane Passage

1980 ◽  
Vol 102 (1) ◽  
pp. 88-95 ◽  
Author(s):  
D. A. Bailey
Keyword(s):  
Boundary Layer ◽  
Laser Doppler Velocimetry ◽  
Large Scale ◽  
Guide Vane ◽  
Velocity Fields ◽  
Inlet Guide Vane ◽  
Test Conditions ◽  
Passage Vortex ◽  
Turbine Vane ◽  
The Mean

Laser-Doppler velocimetry was used to investigate the secondary flow in the endwall region of a large-scale turbine inlet-guide-vane passage. The mean and turbulent velocities were measured for three different test conditions. The different test conditions consisted of variations in the blade aspect ratio and the inlet boundary-layer thickness or all three cases, a passage vortex was identified and its development documented. The turbulent stresses within the vortex were found to be quite low in comparison with the turbulence in the inlet boundary layer.

Rotor Boundary Layer Response to an Impinging Wake

Volume 1 ◽  
2004 ◽  
Author(s):  
Francesco Soranna ◽  
Yi-Chih Chow ◽  
Oguz Uzol ◽  
Joseph Katz
Keyword(s):  
Boundary Layer ◽  
Guide Vane ◽  
Suction Side ◽  
Velocity Profiles ◽  
Inlet Guide Vane ◽  
Pressure Gradients ◽  
Reynolds Stresses ◽  
Mean Flow ◽  
Image Velocimetry ◽  
Entire Flow

This paper presents results of an experimental investigation on the response of a rotor boundary layer to an impinging Inlet Guide Vane (IGV) wake. High resolution two-dimensional Particle Image Velocimetry (PIV) measurements are conducted in a refractive index matched facility that provides an unobstructed view of the entire flow field. Data obtained at four different rotor phases, as the wake is chopped and passes by the rotor blade, allows us to examine the response of the rotor boundary layer to the mean flow and turbulence associated with the impinging wake. We focus on the suction side boundary layer in regions with adverse pressure gradients, from mid chord to the trailing edge. The phase-averaged velocity profiles are used for calculating the momentum and displacement thicknesses of the boundary layer, and for estimating the pressure gradients along the wall. Distributions of Reynolds stresses are also provided. The phase-averaged velocity profiles in the rotor boundary layer vary significantly with phase. During wake impingement the boundary layer becomes significantly thinner and more stable compared to other phases at the same location. Analysis of the possible causes for this trend suggests that the dominant contributors are unsteady, phase-dependent variation in pressure gradients along the wall.

Effects of Orientation and Position of the Combustor-Turbine Interface on the Cooling of a Vane Endwall

2011 ◽  
Author(s):  
A. A. Thrift ◽  
K. A. Thole ◽  
S. Hada
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Vortex Formation ◽  
Horseshoe Vortex ◽  
Time Resolved ◽  
Image Velocimetry ◽  
Thermal Measurements ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Stagnation Plane

First stage, nozzle guide vanes and accompanying endwalls are extensively cooled by the use of film cooling through discrete holes and leakage flow from the combustor-turbine interface gap. While there are cooling benefits from the interface gap, it is generally not considered as part of the cooling scheme. This paper reports on the effects of the position and orientation of a two-dimensional slot on the cooling performance of a nozzle guide vane endwall. In addition to surface thermal measurements, time-resolved, digital particle image velocimetry (TRDPIV) measurements were performed at the vane stagnation plane. Two slot orientations, 90° and 45°, and three streamwise positions were studied. Effectiveness results indicate a significant increase in area averaged effectiveness for the 45° slot relative to the 90° orientation. Flowfield measurements show dramatic differences in the horseshoe vortex formation.

A Technological Effect Modeling on Complex Turbomachinery Applications With an Overset Grid Numerical Method

2013 ◽  
Author(s):  
Gilles Billonnet ◽  
Lionel Castillon ◽  
Jacques Riou ◽  
Gilles Leroy ◽  
André Paillassa
Keyword(s):  
Film Cooling ◽  
Cooling System ◽  
Guide Vane ◽  
Grid Method ◽  
Industrial Test ◽  
Secondary Flows ◽  
Inlet Guide Vane ◽  
Overset Grid ◽  
Overlapping Grids ◽  
Wall Boundary Conditions

The modeling of technological effects on complex turbomachinery flow is described in the paper. The Chimera method based on structured overlapping grids is applied using the ONERA solver elsA. The application of the method on two industrial test cases are presented. The first investigated application is an experimental configuration of a turbine vane with film-cooling. The film cooling system is made up of a very large number of holes. The Chimera method enables simulating the interaction between the cooling jets and the vane flow and improves heat flux prediction compared to simulations modeling cooling flows with wall boundary conditions. The second investigated application is a variable Inlet Guide Vane of an experimental compressor. The application includes the main flow vane, the pivot linking the hub wall with the IGV blade, and the built-in turntable within the shroud which ensures the blade fitting. The benefit of the overset grid method is highlighted by comparisons with computation results obtained on the smooth end-walls. For three different stagger angles (0°, 30° and 60°) the patterns of the secondary flows are presented as well as the comparisons of the calculated flow field with the available experimental data.

Turbine Vane Endwall Film Cooling With Slashface Leakage and Discrete Hole Configuration

2017 ◽  
Vol 139 (6) ◽  
Author(s):  
Nafiz H. K. Chowdhury ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han ◽  
Luzeng Zhang ◽  
Hee-Koo Moon
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Secondary Flows ◽  
Blow Down ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Free Stream Turbulence ◽  
Leakage Flows ◽  
Turbine Vane

Turbine vanes are typically assembled as a section containing single or double airfoil units in an annular pattern. First stage guide vane assembly results in two common mating interfaces: a gap between combustor and vane endwall and another resulted from the adjacent sections, called slashface. High pressure coolant could leak through these gaps to reduce the ingestion of hot gas and achieve certain cooling benefit. As vane endwall region flow field is already very complicated due to highly three-dimensional secondary flows, then a significant influence on endwall cooling can be expected due to the gap leakage flows. To determine the effect of leakage flows from those gaps, film cooling effectiveness distributions were measured using pressure sensitive paint (PSP) technique on the endwall of a scaled up, midrange industrial turbine vane geometry with the multiple rows of discrete film cooling (DFC) holes inside the passages. Experiments were performed in a blow-down wind tunnel cascade facility at the exit Mach number of 0.5 corresponding to Reynolds number of 3.8 × 105 based on inlet conditions and axial chord length. Passive turbulence grid was used to generate free-stream turbulence (FST) level about 19% with an integral length scale of 1.7 cm. Two parameters, coolant-to-mainstream mass flow ratio (MFR) and density ratio (DR), were studied. The results are presented as two-dimensional film cooling effectiveness distribution on the vane endwall surface with the corresponding spanwise averaged values along the axial direction.

Investigation of Flow Pattern and Losses in Transonic Turbine Vane and Blade Cascades by Means of Laser Anemometer Measurements and Navier-Stokes Analysis

2003 ◽  
Author(s):  
Andrei V. Granovski ◽  
Alexei N. Kolesov
Keyword(s):  
Flow Structure ◽  
Navier Stokes ◽  
Mean Flow ◽  
Static Pressure Distribution ◽  
Pressure Fields ◽  
Profile Losses ◽  
Turbine Vane ◽  
Laser Anemometer ◽  
Transonic Turbine ◽  
Numerical Approaches

This paper presents results of a complex numerical and experimental investigation of the flow structure and losses in the Vane (Ma2is = 1.0, Re = 9.8*105) and Blade (Ma2is = 1.12, Re = 7.8*105) straight cascades on transonic modes. The measurements of turbulence pulsation and mean flow velocity upstream, within and downstream of the cascades were made by means of laser anemometer (LA). The static and total pressure fields were measured upstream and downstream of the cascades to determine profile losses. The static pressure distribution on the suction and pressure surfaces was measured as well. Comparisons were made with predictions using 2D Navier–Stokes analysis. The use of the both experimental and numerical approaches allowed to eliminate separated zones more accurately, to define how these zones affected on flow structure and losses.

Fluid Thermal Network Studies on Cooled Nozzle Guide Vane

2019 ◽  
Author(s):  
Pol Reddy Kukutla ◽  
B. V. S. S. S. Prasad
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Flow Distribution ◽  
One Dimensional ◽  
Dimensional Network ◽  
Commercial Code ◽  
Turbine Vane ◽  
Dimensional Simulation ◽  
Nozzle Guide ◽  
Network Studies

Abstract The aerothermal analysis is reported with the help of one-dimensional network modeling for the impingement cum film cooled gas turbine vane. The purpose of this one-dimensional simulation is to obtain the optimized film hole diameters of each row by analyzing the coolant flow distribution and overall effectiveness variations. FlownexR2017 commercial code is used to determine the detailed steady-state performance of the cooling vane. The results show that it is a useful simulation tool to obtain improved effectiveness of film cooling rows in a relatively short turn around time.

Upstream Film Cooling on the Contoured Endwall of a Transonic Turbine Vane in an Annular Cascade

2021 ◽  
Author(s):  
Daniel Salinas ◽  
Izhar Ullah ◽  
Lesley Wright ◽  
Je-Chin Han ◽  
John Mcclintic ◽  
...  
Keyword(s):  
Film Cooling ◽  
Turbine Vane ◽  
Transonic Turbine ◽  
Annular Cascade

Effects of Axial Row-Spacing for Double-Jet Film-Cooling on the Cooling Effectiveness

2011 ◽  
Author(s):  
Zhan Wang ◽  
Jian-Jun Liu ◽  
Bai-tao An ◽  
Chao Zhang
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Inlet Guide Vane ◽  
Row Spacing ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Cylindrical Holes ◽  
Compound Angle

The effects of axial row-spacing for double jet film-cooling (DJFC) with compound angle on the cooling characteristics under different blowing ratios were investigated numerically. First, the flow fields and cooling effectiveness of DJFC on flat plate with different axial row-spacing were calculated. Film-cooling with fan-shaped or cylindrical holes was also calculated for the comparison. The results indicate that a larger axial row-spacing is helpful to form the anti-kidney vortex and to improve the cooling effectiveness. The DJFC was then applied to the suction and pressure surface of a real turbine inlet guide vane. Comparisons of film-cooling effectiveness with the cylindrical and fan-shaped holes were also conducted. The results for the guide vane show that on the suction surface the DJFC with a larger axial row-spacing leads to better film coverage and better film-cooling effectiveness than the cylindrical or fan-shaped holes. On the pressure surface, however, the film-cooling with fan-shaped holes is superior to the others.

Shower Head and Trailing Edge Cooling Influence on Transonic Vane Aero Performance

2014 ◽  
Vol 136 (11) ◽  
Author(s):  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten H. Fransson ◽  
Boris I. Mamaev ◽  
Mats Annerfeldt ◽  
...  
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Trailing Edge ◽  
Aerodynamic Loss ◽  
Nozzle Guide Vane ◽  
Turbine Vane ◽  
Engine Component ◽  
Annular Sector ◽  
Exit Mach Number ◽  
Nozzle Guide

An experimental investigation on a cooled nozzle guide vane (NGV) has been conducted in an annular sector to quantify aerodynamic influences of shower head (SH) and trailing edge (TE) cooling. The investigated vane is a typical high pressure gas turbine vane, geometrically similar to a real engine component, operated at a reference exit Mach number of 0.89. The investigations have been performed for various coolant-to-mainstream mass–flux ratios. New loss equations are derived and implemented regarding coolant aerodynamic losses. Results lead to a conclusion that both TE cooling and SH film cooling increase the aerodynamic loss compared to an uncooled case. In addition, the TE cooling has higher aerodynamic loss compared to the SH cooling. Secondary losses decrease with inserting SH film cooling compared to the uncooled case. The TE cooling appears to have less impact on the secondary loss compared to the SH cooling. Area-averaged exit flow angles around midspan increase for the TE cooling.

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