scholarly journals A Time Efficient Adaptive Gridding Approach and Improved Calibrations in Five-Hole Probe Measurements

2015 ◽  
Vol 2015 ◽  
pp. 1-14 ◽  
Author(s):  
Jason Town ◽  
Cengiz Camci
Keyword(s):  
Spatial Resolution ◽  
Large Scale ◽  
Guide Vane ◽  
Axial Flow ◽  
Grid Method ◽  
Flow Fields ◽  
Adaptive Grid ◽  
Current Approach ◽  
Secondary Flows ◽  
Complex Flow

Five-Hole Probes (FHP), being a dependable and accurate aerodynamic tool, are an excellent choice for measuring three-dimensional flow fields in turbomachinery. To improve spatial resolution, a subminiature FHP with a diameter of 1.68 mm is employed. High length to diameter ratio of the tubing and manual pitch and yaw calibration cause increased uncertainty. A new FHP calibrator is designed and built to reduce the uncertainty by precise, computer controlled movements and reduced calibration time. The calibrated FHP is then placed downstream of the nozzle guide vane (NGV) assembly of a low-speed, large-scale, axial flow turbine. The cold flow HP turbine stage contains 29 vanes and 36 blades. A fast and computer controllable traversing system is implemented using an adaptive grid method for the refinement of measurements in regions such as vane wake, secondary flows, and boundary layers. The current approach increases the possible number of measurement points in a two-hour period by 160%. Flow structures behind the NGV measurement plane are identified with high spatial resolution and reduced uncertainty. The automated pitch and yaw calibration and the adaptive grid approach introduced in this study are shown to be a highly effective way of measuring complex flow fields in the research turbine.

Uncertainty Propagation Analyses of Lean Burn Combustor Exit Conditions for a Robust Nozzle Cooling Design

2020 ◽  
Vol 142 (5) ◽  
Author(s):  
Marius Schneider ◽  
Heinz-Peter Schiffer ◽  
Knut Lehmann
Keyword(s):  
Guide Vane ◽  
Uncertainty Propagation ◽  
Flow Fields ◽  
Secondary Flows ◽  
Complex Flow ◽  
Lean Burn ◽  
Numerical Predictions ◽  
Cooling Design ◽  
Parameterized Model ◽  
Nozzle Guide

Abstract Knowing the flow conditions at the combustor turbine interface is a key asset for an efficient cooling design of high-pressure turbines. However, measurements and numerical predictions of combustor exit conditions are challenging due to the extreme temperatures and complex flow patterns in modern combustors. Even the time-averaged flow fields at the combustor exit which are commonly used as inlet condition for simulations of the turbine are therefore subject to uncertainty. The goal of this paper is to illustrate how aleatory uncertainties in the magnitude and position of residual swirl and hot spots at the combustor exit affect uncertainties in the prediction of cooling and heat load of the first nozzle guide vane. Also, it is identified which of these uncertain parameters have the greatest impact. An iso-thermal test rig and an engine realistic setup with lean burn inflow conditions are investigated. The analysis combines a parameterized model for combustor exit flow fields with uncertainty quantification methods. It is shown that the clocking position of turbine inlet swirl has a large effect on the formation of secondary flows on the vane surface and thus affects the uncertainty of thermal predictions on the hub and vanes.

An Experimental Investigation of the Flow Through an Axial-Flow Pump

1995 ◽  
Vol 117 (3) ◽  
pp. 485-490 ◽  
Author(s):  
W. C. Zierke ◽  
W. A. Straka ◽  
P. D. Taylor
Keyword(s):  
Large Scale ◽  
Three Dimensional ◽  
Axial Flow ◽  
Fast Response ◽  
Complex Flow ◽  
Penn State ◽  
Response Pressure ◽  
Axial Flow Pump ◽  
Scale Experiment ◽  
High Reynolds

The high Reynolds number pump (HIREP) facility at ARL Penn State has been used to perform a low-speed, large-scale experiment of the incompressible flow of water through a two-blade-row turbomachine. The objectives of this experiment were to provide a database for comparison with three-dimensional, turbulent flow computations, to evaluate engineering models, and to improve our physical understanding of many of the phenomena involved in this complex flow field. This summary paper briefly describes the experimental facility, as well as the experimental techniques—such as flow visualization, static-pressure measurements, laser Doppler velocimetry, and both slow- and fast-response pressure probes. Then, proceeding from the inlet to the exit of the pump, the paper presents highlights of experimental measurements and data analysis, giving examples of measured physical phenomena such as endwall boundary layers, separation regions, wakes, and secondary vortical structures. In conclusion, this paper provides a synopsis of a well-controlled, larger scope experiment that should prove helpful to those who wish to use the database.

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.

Numerical Validations of Secondary Flows and Loss Development Downstream of a Highly Loaded Low Pressure Turbine Outlet Guide Vane Cascade

2007 ◽  
Author(s):  
Johan Hja¨rne ◽  
Valery Chernoray ◽  
Jonas Larsson ◽  
Lennart Lo¨fdahl
Keyword(s):  
Numerical Simulations ◽  
Secondary Flow ◽  
Incompressible Flows ◽  
Guide Vane ◽  
Low Pressure ◽  
Flow Fields ◽  
Secondary Flows ◽  
Test Facility ◽  
Low Pressure Turbine ◽  
Highly Loaded

In this paper 3D numerical simulations of turbulent incompressible flows are validated against experimental data from the linear low pressure turbine/outlet guide vane (LPT/OGV) cascade at Chalmers in Sweden. The validation focuses on the secondary flow-fields and loss developments downstream of a highly loaded OGV. The numerical simulations are performed for the same inlet conditions as in the test-facility with engine-like properties in terms of Reynolds number, boundary-layer thickness and inlet flow angles with the goal to validate how accurately and reliably the secondary flow fields and losses for both on- and off-design conditions can be predicted for OGV’s. Results from three different turbulence models as implemented in FLUENT, k-ε Realizable, kω-SST and the RSM are validated against detailed measurements. From these results it can be concluded that the RSM model predicts both the secondary flow field and the losses most accurately.

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.

Computational Validation of the Flow Through a Turbine Stage and the Effects of Rim Seal Cavity Leakage on Secondary Flows

2012 ◽  
Author(s):  
Özhan H. Turgut ◽  
Cengiz Camcı
Keyword(s):  
Validation Study ◽  
Guide Vane ◽  
Pressure Coefficient ◽  
Axial Flow ◽  
Turbine Rotor ◽  
Secondary Flows ◽  
Transitional Flow ◽  
Interface Plane ◽  
Exit Plane ◽  
Computational Validation

A computational validation study related to aerodynamic loss generation mechanisms is performed in an axial flow turbine. The 91.66 cm diameter axial flow turbine research facility has a stationary nozzle guide vane assembly and a 29 bladed HP turbine rotor. The NGV inlet and exit Reynolds numbers based on midspan axial chord are around 300000 and 900000, respectively. GRIDPRO is used as the structured grid generator. y+ values are kept below unity. The finite-volume flow solver ANSYS CFX with SST k–ω turbulence model together with the transitional flow model is employed. Experimental flow conditions are imposed at the boundaries. The computational predictions are compared to experimental data at NGV exit plane and rotor inlet plane. NGV exit plane measurements come from a previous experimental study with a five-hole probe and the data at rotor inlet plane is taken by the current authors using a Kiel probe with 3.175mm head diameter. The comparison of rotor-stator interface models shows that the stage model, which calculates the circumferentially averaged fluxes and uses as the boundary condition at the interface plane, agrees well with the experimental total pressure coefficient data at the NGV exit. The difference between the NGV only simulation and the rotor-stator simulation is emphasized. The effect of rim seal flow on the mainstream aerodynamics is investigated. This validation study shows that the effect of future geometrical modifications on the endwalls and the vane will be predicted reasonably accurately.

Combustor Turbine Interface Studies: Part 1 — Endwall Effectiveness Measurements

2002 ◽  
Author(s):  
W. F. Colban ◽  
K. A. Thole ◽  
G. Zess
Keyword(s):  
Film Cooling ◽  
Total Pressure ◽  
Large Scale ◽  
Guide Vane ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Profile ◽  
Secondary Flows ◽  
Adiabatic Wall ◽  
Combustor Liner

Improved durability of gas turbine engines is an objective for both military and commercial aeroengines as well as for power generation engines. One region susceptible to degradation in an engine is the junction between the combustor and first vane given that the main gas path temperatures at this location are the highest. The platform at this junction is quite complex in that secondary flow effects, such as the leading edge vortex, are dominant. Past computational studies have shown that the total pressure profile exiting the combustor dictates the development of the secondary flows that are formed. This study examines the effect of varying the combustor liner film-cooling and junction slot flows on the adiabatic wall temperatures measured on the platform of the first vane. The experiments were performed using large-scale models of a combustor and nozzle guide vane in a wind tunnel facility. The results show that varying the coolant injection from the upstream combustor liner leads to differing total pressure profiles entering the turbine vane passage. Endwall adiabatic effectiveness measurements indicate that the coolant does not exit the upstream combustor slot uniformly but instead accumulates along the suction side of the vane and endwall. Increasing the liner cooling continued to reduce endwall temperatures, which was not found to be true with increasing the film-cooling from the liner.

Improving turbine endwall cooling uniformity by controlling near-wall secondary flows

2016 ◽  
Vol 231 (14) ◽  
pp. 2689-2705 ◽  
Author(s):  
Mitra Thomas ◽  
Thomas Povey
Keyword(s):  
Mass Flux ◽  
Large Scale ◽  
Guide Vane ◽  
Pressure Ratio ◽  
Experimental Studies ◽  
Pressure Profile ◽  
Design Guidelines ◽  
Secondary Flows ◽  
The Impact ◽  
Endwall Cooling

In this paper, we propose a design philosophy for cooling high-pressure nozzle guide vane endwalls, which exploits the momentum of cooling jets to control vane secondary flows thereby improving endwall cooling uniformity. The impact of coolant-to-mainstream pressure ratio, hole inclination angle, hole diameter, vane potential field, and overall mass flux ratios are considered. Arguments are developed by examining detailed experimental studies conducted in a large-scale low-speed cascade tunnel with engine-realistic combustor geometry and turbulence profiles. Computational fluid dynamics predictions validated by the same are used to extend the parameter space. We show that the global flow field is highly sensitive to the inlet total pressure profile, which in turn can be modified by introducing relatively low mass flow rates of cooling gas at engine realistic coolant-to-mainstream pressure ratios and mass flux ratios. This interaction effect must be understood for successful design of optimised endwall cooling schemes, an effect which is not sufficiently emphasized in much of the literature on this topic. Design guidelines are given that we hope will be of use in industry.

scholarly journals Flow Measurement in a Multistage Large Scale Low Speed Axial Flow Research Compressor

1998 ◽  
Author(s):  
P. Boos ◽  
H. Möckel ◽  
J. M. Henne ◽  
R. Seimeler
Keyword(s):  
Large Scale ◽  
Measurement Data ◽  
Axial Flow ◽  
Field Measurements ◽  
Secondary Flows ◽  
Navier Stokes ◽  
Test Configuration ◽  
Low Speed ◽  
Design Concepts ◽  
University Of Technology

In this paper the newly built large scale low speed axial flow research compressor at Dresden University of Technology is presented. This compressor rig serves three main purposes. Firstly, it shall improve the understanding of compressor aerodynamics (especially secondary flows) by allowing detailed flow field measurements without heavily disturbing the flow. Secondly, it will be used to examine new design concepts. Thirdly, the detailed measurements in the absolute and relative system will be used for the calibration of existing CFD-codes. The design and the construction of the test rig which will allow an easy variation of the test configuration is described. A short view of the different data acquisition units for steady and unsteady measurement in the stationary and rotating system will be given. The blading of the compressor in the first series of test runs simulates a middle stage of a contemporary high-pressure compressor. Measurement data will be compared with results of 3D-Navier-Stokes calculations that were performed at MTU München.

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