Experimental Study of the Flow Field Behind an Annular Turbine Nozzle Guide Vane With and Without Downstream Rotor

1984 ◽  
Author(s):  
E. Boletis ◽  
C. H. Sieverding
Keyword(s):  
Flow Field ◽  
Guide Vane ◽  
Three Dimensional ◽  
Rotor Blades ◽  
Pressure Probe ◽  
Flow Conditions ◽  
Nozzle Guide Vane ◽  
Contour Plots ◽  
Real Flow ◽  
Nozzle Guide

Measurements of the three dimensional flow field in annular turbine nozzle guide vanes present an important step in the simulation of the real flow conditions in turbomachinery bladings. This paper seeks to determine whether the installation of a rotor closely behind a high hub-to-tip ratio cascade (DH/DT=0.8) is indispensable for establishing correct flow conditions at the cascade exit or whether the use of an axial diffuser of a certain length is sufficient. Also, an attempt is made to separate the possible effects of the rotor blades from that of the rotating rotor disc. The tests are carried out on a low speed, low aspect ratio, high turning nozzle guide vane. The flow is explored by means of a double head four-hole pressure probe and the results are presented in the form of contour plots and spanwise pitch-averaged distributions of losses, flow angles and static pressure.

scholarly journals Effects of Tip Endwall Contouring on the Three Dimensional Flow Field in an Annular Turbine Nozzle Guide Vane: Part 2 — Numerical Investigation

1985 ◽  
Author(s):  
Tony Arts
Keyword(s):  
Flow Field ◽  
Numerical Investigation ◽  
Guide Vane ◽  
Three Dimensional ◽  
Three Dimensional Flow ◽  
Dimensional Flow ◽  
Nozzle Guide Vane ◽  
Contour Plots ◽  
Endwall Contouring ◽  
Nozzle Guide

This paper describes the numerical investigation of the three dimensional flow through a low speed, low aspect ratio, high turning annular turbine nozzle guide vane with meridional tip endwall contouring. This rotational flow field has been simulated using a finite volume discretization and a time marching technique to solve the three dimensional, time dependent Euler equations expressed in a cylindrical coordinates system. The results are presented under the form of contour plots, spanwise pitch-averaged distributions and blade static pressure distributions. Detailed comparisons with the measurements described in part I of the paper are also provided.

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.

The Extension of a Solution-Adaptive Three-Dimensional Navier–Stokes Solver Toward Geometries of Arbitrary Complexity

1993 ◽  
Vol 115 (2) ◽  
pp. 283-295 ◽  
Author(s):  
W. N. Dawes
Keyword(s):  
Guide Vane ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Centrifugal Impeller ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Wide Range ◽  
Nozzle Guide Vane ◽  
Recent Developments ◽  
Nozzle Guide

This paper describes recent developments to a three-dimensional, unstructured mesh, solution-adaptive Navier–Stokes solver. By adopting a simple, pragmatic but systematic approach to mesh generation, the range of simulations that can be attempted is extended toward arbitrary geometries. The combined benefits of the approach result in a powerful analytical ability. Solutions for a wide range of flows are presented, including a transonic compressor rotor, a centrifugal impeller, a steam turbine nozzle guide vane with casing extraction belt, the internal coolant passage of a radial inflow turbine, and a turbine disk cavity flow.

Flow Measurements in a Nozzle Guide Vane Passage With a Low Aspect Ratio and Endwall Contouring

2000 ◽  
Author(s):  
Steven W. Burd ◽  
Terrence W. Simon
Keyword(s):  
Aspect Ratio ◽  
Flow Field ◽  
Guide Vane ◽  
Flow Measurements ◽  
Vast Number ◽  
Aspect Ratios ◽  
Nozzle Guide Vane ◽  
Endwall Contouring ◽  
Nozzle Guide

The vast number of turbine cascade studies in the literature has been performed in straight-endwall, high-aspect-ratio, linear cascades. As a result, there has been little appreciation for the role of, and added complexity imposed by, reduced aspect ratios. There also has been little documentation of endwall profiling at these reduced spans. To examine the role of these factors on cascade hydrodynamics, a large-scale nozzle guide vane simulator was constructed at the Heat Transfer Laboratory of the University of Minnesota. This cascade is comprised of three airfoils between one contoured and one flat endwall. The geometries of the airfoils and endwalls, as well as the experimental conditions in the simulator, are representative of those in commercial operation. Measurements with hot-wire anemometry were taken to characterize the flow approaching the cascade. These measurements show that the flow field in this cascade is highly elliptic and influenced by pressure gradients that are established within the cascade. Exit flow field measurements with triple-sensor anemometry and pressure measurements within the cascade indicate that the acceleration imposed by endwall contouring and airfoil turning is able to suppress the size and strength of key secondary flow features. In addition, the flow field near the contoured endwall differs significantly from that adjacent to the straight endwall.

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.

Research of Aerodynamic Performance of HP-Turbine with Coolant Injections for Variable Cycle Engine

2011 ◽  
Vol 110-116 ◽  
pp. 1047-1053
Author(s):  
Zhi Gang Liu ◽  
Xiang Jun Fang ◽  
Si Yong Liu ◽  
Ping Wang ◽  
Zhao Yin
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Supersonic Nozzle ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Thermodynamic Cycle ◽  
Aerodynamic Performance ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Transonic Rotor

A highly loaded high-pressure turbine with a supersonic nozzle guide vane and a transonic rotor for a Variable Cycle Engine (VCE) has been investigated. Film cooling strategies were designed for the whole stage, during which the positions, injection orientations and arrangements of cooling holes were confirmed. Three-dimensional steady numerical simulations have been performed in the two operation modes of low and high bypass ratio with different thermodynamic cycle parameters according to the VCE and the coolant injections have been simulated by means of additional source term method. The influences of coolant injections in the fully cooled turbine stage on aerodynamic performance and flow characteristics have been analyzed. The results indicate that, the supersonic nozzle guide vane, over-expansion degree of main flows, fluctuations of static pressure and intensity of corner vortex are lessened or alleviated. In the transonic rotor, expansion and doing work capabilities in the mixed fluid are strengthened. Proper coolants injections are beneficial to the flow characteristics in the blade passage.

The Influence of the Boundary Layer State and Reynolds Number on Film Cooling and Heat Transfer on a Cooled Nozzle Guide Vane

2000 ◽  
Author(s):  
Hans Reiss ◽  
Albin Bölcs
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Film Cooling ◽  
Guide Vane ◽  
Suction Side ◽  
Flow Conditions ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Film cooling and heat transfer measurements were carried out on a cooled nozzle guide vane in a linear cascade, using a transient liquid crystal technique. Three flow conditions were realized: the nominal operating condition of the vane with an exit Reynolds number of 1.47e6, as well as two lower flow conditions: Re2L = 1.0e6 and 7.5e5. The vane model was equipped with a single row of inclined round film cooling holes with compound angle orientation on the suction side. Blowing ratios ranging form 0.3 to 1.5 were covered, all using foreign gas injection (CO2) yielding an engine-representative density ratio of 1.6. Two distinct states of the incoming boundary layer onto the injection station were compared, an undisturbed laminar boundary layer as it forms naturally on the suction side, and a fully turbulent boundary layer which was triggered with a trip wire upstream of injection. The aerodynamic flow field is characterized in terms of profile Mach number distribution, and the associated heat transfer coefficients around the uncooled airfoil are presented. Both detailed and spanwise averaged results of film cooling effectiveness and heat transfer coefficients are shown on the suction side, which indicate considerable influence of the state of the incoming boundary layer on the performance of a film cooling row. The influence of the mainstream flow condition on the film cooling behavior at constant blowing ratio is discussed for three chosen injection regimes.

Toward Improved Throughflow Capability: The Use of Three-Dimensional Viscous Flow Solvers in a Multistage Environment

1992 ◽  
Vol 114 (1) ◽  
pp. 8-17 ◽  
Author(s):  
W. N. Dawes
Keyword(s):  
Experimental Evidence ◽  
Guide Vane ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Flow Properties ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Blade Row ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

A methodology is presented for simulating turbomachinery blade rows in a multistage environment by deploying a standard three-dimensional Navier–Stokes solver simultaneously on a number of blade rows. The principal assumptions are that the flow is steady relative to each blade row individually and that the rows can communicate via inter-row mixing planes. These mixing planes introduce circumferential averaging of flow properties but preserve quite general radial variations. Additionally, each blade can be simulated in three-dimensional or axisymmetrically (in the spirit of throughflow analysis) and a series of axisymmetric rows can be considered together with one three-dimensional row to provide, cheaply, a machine environment for that row. Two applications are presented: a transonic compressor rotor and a steam turbine nozzle guide vane simulated both isolated and as part of a stage. In both cases the behavior of the blade considered in isolation was different to when considered as part of a stage and in both cases was in much closer agreement with the experimental evidence.

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.

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