Midspan Flow-Field Measurements for Two Transonic Linear Turbine Cascades at Off-Design Conditions

2001 ◽  
Author(s):  
D. B. M. Jouini ◽  
S. A. Sjolander ◽  
S. H. Moustapha
Keyword(s):  
Flow Field ◽  
Turbine Blade ◽  
Density Ratio ◽  
Leading Edge ◽  
Field Measurements ◽  
Reynolds Numbers ◽  
Experimental Results ◽  
Tabular Form ◽  
Data Set ◽  
Turbine Cascades

The paper presents detailed mid-span experimental results from two transonic linear turbine cascades. The blades for the two cascades were designed for the same service and differ mainly in their leading-edge geometries. One of the goals of the study was investigate the influence of the leading-edge metal angle on the sensitivity of the blade to positive off-design incidence. Measurements were made for incidence values of −10.0°, 0.0°, +4.5°, +10.0°, and +14.5° relative to design incidence. The exit Mach numbers varied roughly from 0.5 to 1.2 and the Reynolds numbers from about 4×105 to 106. The measurements include the midspan losses, blade loadings and base pressures. In addition, the axial-velocity-density ratio (AVDR) was extracted for each operating point The AVDR was found to vary from about 0.98 at −10.0° of incidence to about 1.27 at +14.5°. Thus, the data set also provides some evidence of the influence AVDR on axial turbine blade performance. Detailed experimental results for turbine blade performance at off-design incidence are very scarce in the open literature, particularly for transonic conditions. Among other things, the present results are intended to expand the database available in the open literature. To this end, the key aerodynamic results are presented in tabular form, along with the detailed geometry of the cascades. The results could be used in the development of new or improved correlations for use in the early stages of design. They could also be used to evaluate the ability of current CFD codes to capture reliably the variation in losses and other aerodynamic quantities with variations in blade incidence.

Midspan Flow-Field Measurements for Two Transonic Linear Turbine Cascades at Off-Design Conditions

2002 ◽  
Vol 124 (2) ◽  
pp. 176-186 ◽  
Author(s):  
D. B. M. Jouini ◽  
S. A. Sjolander ◽  
S. H. Moustapha
Keyword(s):  
Flow Field ◽  
Turbine Blade ◽  
Density Ratio ◽  
Leading Edge ◽  
Field Measurements ◽  
Reynolds Numbers ◽  
Experimental Results ◽  
Tabular Form ◽  
Data Set ◽  
Turbine Cascades

The paper presents detailed midspan experimental results from two transonic linear turbine cascades. The blades for the two cascades were designed for the same service and differ mainly in their leading-edge geometries. One of the goals of the study was to investigate the influence of the leading-edge metal angle on the sensitivity of the blade to positive off-design incidence. Measurements were made for incidence values of −10.0, 0.0, +4.5, +10.0, and +14.5 deg relative to design incidence. The exit Mach numbers varied roughly from 0.5 to 1.2 and the Reynolds numbers from about 4×105 to 106. The measurements include the midspan losses, blade loadings and base pressures. In addition, the axial-velocity-density ratio (AVDR) was extracted for each operating point. The AVDR was found to vary from about 0.98 at −10.0 deg of incidence to about 1.27 at +14.5 deg. Thus, the data set also provides some evidence of the influence AVDR on axial turbine blade performance. Detailed experimental results for turbine blade performance at off-design incidence are very scarce in the open literature, particularly for transonic conditions. Among other things, the present results are intended to expand the database available in the open literature. To this end, the key aerodynamic results are presented in tabular form, along with the detailed geometry of the cascades. The results could be used in the development of new or improved correlations for use in the early stages of design. They could also be used to evaluate the ability of current CFD codes to capture reliably the variation in losses and other aerodynamic quantities with variations in blade incidence.

Influence of Turbine Blade Leading Edge Profile on Film Cooling With Shaped Holes

2017 ◽  
Author(s):  
Mingjie Zhang ◽  
Nian Wang ◽  
Andrew F. Chen ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Edge Profile ◽  
Elliptical Cylinders ◽  
Blade Leading Edge

This paper presents the turbine blade leading edge model film cooling effectiveness with shaped holes, using the pressure sensitive paint (PSP) mass transfer analogy method. The effects of leading edge profile, coolant to mainstream density ratio and blowing ratio are studied. Computational simulations are performed using the realizable k-ε turbulence model. Effectiveness obtained by CFD simulations are compared with experiments. Three leading edge profiles, including one semi-cylinder and two semi-elliptical cylinders with an after body, are investigated. The ratios of major to minor axis of two semi-elliptical cylinders are 1.5 and 2.0, respectively. The leading edge has three rows of shaped holes. For the semi-cylinder model, shaped holes are located at 0 degrees (stagnation line) and ± 30 degrees. Row spacing between cooling holes and the distance between impingement plate and stagnation line are the same for three leading edge models. The coolant to mainstream density ratio varies from 1.0 to 1.5 and 2.0, and the blowing ratio varies from 0.5 to 1.0 and 1.5. Mainstream Reynolds number is about 100,900 based on the diameter of the leading edge cylinder, and the mainstream turbulence intensity is about 7%. The results provide an understanding of the effects of leading edge profile and on turbine blade leading edge region film cooling with shaped-hole designs.

Experimental Study of the Effects of an Oscillating Approach Flow on Overall Cooling Performance of a Simulated Turbine Blade Leading Edge

2009 ◽  
Author(s):  
Ross Johnson ◽  
Jonathan Maikell ◽  
David Bogard ◽  
Justin Piggush ◽  
Atul Kohli ◽  
...  
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Stagnation Line ◽  
Cooling Performance ◽  
Blowing Ratio ◽  
Oscillation Frequencies ◽  
Overall Effectiveness ◽  
Blade Leading Edge

When a turbine blade passes through wakes from upstream vanes it is subjected to an oscillation of the direction of the approach flow resulting in the oscillation of the position of the stagnation line on the leading edge of the blade. In this study an experimental facility was developed that induced a similar oscillation of the stagnation line position on a simulated turbine blade leading edge. The overall effectiveness was evaluated at various blowing ratios and stagnation line oscillation frequencies. The location of the stagnation line on the leading edge was oscillated to simulate a change in angle of attack between α = ± 5° at a range of frequencies from 2 to 20 Hz. These frequencies were chosen based on matching a range of Strouhal numbers typically seen in an engine due to oscillations caused by passing wakes. The blowing ratio was varied between M = 1, M = 2, and M = 3. These experiments were carried out at a density ratio of DR = 1.5 and mainstream turbulence levels of Tu ≈ 6%. The leading edge model was made of high conductivity epoxy in order to match the Biot number of an actual engine airfoil. Results of these tests showed that the film cooling performance with an oscillating stagnation line was degraded by as much as 25% compared to the performance of a steady flow with the stagnation line aligned with the row of holes at the leading edge.

Computational Design and Experimental Evaluation of Using a Leading Edge Fillet on a Gas Turbine Vane

2001 ◽  
Author(s):  
G. A. Zess ◽  
K. A. Thole
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Gradient ◽  
Gas Turbine ◽  
Total Pressure ◽  
Guide Vane ◽  
Leading Edge ◽  
Field Measurements ◽  
Computational Design ◽  
Horseshoe Vortex

With the desire for increased power output for a gas turbine engine comes the continual push to achieve higher turbine inlet temperatures. Higher temperatures result in large thermal and mechanical stresses particularly along the nozzle guide vane. One critical region along a vane is the leading edge-endwall juncture. Based on the assumption that the approaching flow to this juncture is similar to a two-dimensional boundary layer, previous studies have shown that a horseshoe vortex forms. This vortex forms because of a radial total pressure gradient from the approaching boundary layer. This paper documents the computational design and experimental validation of a fillet placed at the leading edge-endwall juncture of a guide vane to eliminate the horseshoe vortex. The fillet design effectively accelerated the incoming boundary layer thereby mitigating the effect of the total pressure gradient. To verify the CFD studies used to design the leading edge fillet, flow field measurements were performed in a large-scale, linear, vane cascade. The flow field measurements were performed with a laser Doppler velocimeter in four planes orientated orthogonal to the vane. Good agreement between the CFD predictions and the experimental measurements verified the effectiveness of the leading edge fillet at eliminating the horseshoe vortex. The flowfield results showed that the turbulent kinetic energy levels were significantly reduced in the endwall region because of the absence of the unsteady horseshoe vortex.

Influence of Turbine Blade Leading Edge Profile on Film Cooling With Shaped Holes

2018 ◽  
Vol 10 (5) ◽  
Author(s):  
Mingjie Zhang ◽  
Nian Wang ◽  
Andrew F. Chen ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Edge Profile ◽  
Elliptical Cylinders ◽  
Blade Leading Edge

This paper presents the turbine blade leading edge model film cooling effectiveness with shaped holes, using the pressure sensitive paint (PSP) mass transfer analogy method. The effects of leading edge profile, coolant to mainstream density ratio, and blowing ratio are studied. Computational simulations are performed using the realizable k–ɛ (RKE) turbulence model. Effectiveness obtained by computational fluid dynamics (CFD) simulations is compared with experiments. Three leading edge profiles, including one semicylinder and two semi-elliptical cylinders with an after body, are investigated. The ratios of major to minor axis of two semi-elliptical cylinders are 1.5 and 2.0, respectively. The leading edge has three rows of shaped holes. For the semicylinder model, shaped holes are located at 0 deg (stagnation line) and ±30 deg. Row spacing between cooling holes and the distance between impingement plate and stagnation line are the same for three leading edge models. The coolant to mainstream density ratio varies from 1.0 to 1.5 and 2.0, and the blowing ratio varies from 0.5 to 1.0 and 1.5. Mainstream Reynolds number is about 100,000 based on the diameter of the leading edge cylinder, and the mainstream turbulence intensity is about 7%. The results provide an understanding of the effects of leading edge profile on turbine blade leading edge region film cooling with shaped hole designs.

Enhancement of Film Cooling Performance at the Leading Edge of Turbine Blade

2006 ◽  
Author(s):  
K.-S. Kim ◽  
Youn J. Kim ◽  
S.-M. Kim
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Cylindrical Body ◽  
Spanwise Direction ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio

To enhance the film cooling performance in the vicinity of the turbine blade leading edge, the flow characteristics of the film-cooled turbine blade have been investigated using a cylindrical body model. The inclination of the cooling holes is along the radius of the cylindrical wall and 20 deg relative to the spanwise direction. Mainstream Reynolds number based on the cylinder diameter was 1.01×105 and 0.69×105, and the mainstream turbulence intensities were about 0.2% in both Reynolds numbers. CO2 was used as coolant to simulate the effect of density ratio of coolant-to-mainstream. Furthermore, the effect of coolant flow rates was studied for various blowing ratios of 0.4, 0.7, 1.1, and 1.4, respectively. In experiment, spatially-resolved temperature distributions along the cylindrical body surface were visualized using infrared thermography (IRT) in conjunction with thermocouples, digital image processing, and in situ calibration procedures. This comparison shows the results generated to be reasonable and physically meaningful. The film cooling effectiveness of current measurement (0.29 mm × 0.33 min per pixel) presents high spatial and temperature resolutions compared to other studies. Results show that the blowing ratio has a strong effect on film cooling effectiveness and the coolant trajectory is sensitive to the blowing ratio. The local spanwise-averaged effectiveness can be improved by locating the first-row holes near the second-row holes.

Measurements of Hub Flow Interaction on Film Cooled Nozzle Guide Vane in Transonic Annular Cascade

2015 ◽  
Vol 137 (8) ◽  
Author(s):  
Lamyaa A. El-Gabry ◽  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten Fransson
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Field Measurements ◽  
Co2 Concentration ◽  
Horseshoe Vortex ◽  
Mainstream Flow ◽  
Nozzle Guide

An experimental study has been performed in a transonic annular sector cascade of nozzle guide vanes (NGVs) to investigate the aerodynamic performance and the interaction between hub film cooling and mainstream flow. The focus of the study is on the endwalls, specifically the interaction between the hub film cooling and the mainstream. Carbon dioxide (CO2) has been supplied to the coolant holes to serve as tracer gas. Measurements of CO2 concentration downstream of the vane trailing edge (TE) can be used to visualize the mixing of the coolant flow with the mainstream. Flow field measurements are performed in the downstream plane with a five-hole probe to characterize the aerodynamics in the vane. Results are presented for the fully cooled and partially cooled vane (only hub cooling) configurations. Data presented at the downstream plane include concentration contour, axial vorticity, velocity vectors, and yaw and pitch angles. From these investigations, secondary flow structures such as the horseshoe vortex, passage vortex, can be identified and show the cooling flow significantly impacts the secondary flow and downstream flow field. The results suggest that there is a region on the pressure side (PS) of the vane TE where the coolant concentrations are very low suggesting that the cooling air introduced at the platform upstream of the leading edge (LE) does not reach the PS endwall, potentially creating a local hotspot.

Measurements of Hub Flow Interaction on Film Cooled Nozzle Guide Vane in Transonic Annular Cascade

2012 ◽  
Author(s):  
Lamyaa A. El-Gabry ◽  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten Fransson
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Field Measurements ◽  
Co2 Concentration ◽  
Pressure Side ◽  
Trailing Edge ◽  
Nozzle Guide

An experimental study has been performed in a transonic annular sector cascade of nozzle guide vanes to investigate the aerodynamic performance and the interaction between hub film cooling and mainstream flow. The focus of the study is on the endwalls, specifically the interaction between the hub film cooling and the mainstream. Carbon dioxide (CO2) has been supplied to the coolant holes to serve as tracer gas. Measurements of CO2 concentration downstream of the vane trailing edge can be used to visualize the mixing of the coolant flow with the mainstream. Flow field measurements are performed in the downstream plane with a 5-hole probe to characterize the aerodynamics in the vane. Results are presented for the fully cooled and partially cooled vane (only hub cooling) configurations. Data presented at the downstream plane include concentration contour, axial vorticity, velocity vectors, and yaw and pitch angles. From these investigations, secondary flow structures such as the horseshoe vortex, passage vortex, can be identified and show the cooling flow significantly impacts the secondary flow and downstream flow field. The results suggest that there is a region on the pressure side of the vane trailing edge where the coolant concentrations are very low suggesting that the cooling air introduced at the platform upstream of the leading edge does not reach the pressure side endwall, potentially creating a local hotspot.

Influence of Scaling Parameters and Gas Properties on Overall Effectiveness on a Leading Edge Showerhead

2018 ◽  
Author(s):  
Connor J. Wiese ◽  
Carol E. Bryant ◽  
James L. Rutledge ◽  
Marc D. Polanka
Keyword(s):  
Mass Flux ◽  
Momentum Flux ◽  
Density Ratio ◽  
Leading Edge ◽  
Reynolds Numbers ◽  
Scale Model ◽  
Freestream Velocity ◽  
Scaling Parameters ◽  
Overall Effectiveness ◽  
Engine Condition

Testing new turbine cooling schemes at engine conditions becomes increasingly cost prohibitive as the desired gas-path temperatures increase. As a result, the turbine component is simulated in a laboratory with a large-scale model that is sized and constructed out of a selected material so that the Biot number is matched between the laboratory and engine conditions. Furthermore, the experimental temperatures are lower, so the surface temperature that the metal component would experience in the engine is scaled via the overall cooling effectiveness, ϕ. Properly measuring ϕ requires that the relevant flow physics must be matched, thus the relevant Reynolds numbers be matched-both those of the freestream and the coolant, as well as the other scaling parameters, such as the mass flux, momentum flux, and velocity ratios. However, if the coolant-to-freestream density ratio does not match that of the engine condition, the mass flux, momentum flux, coolant and freestream Reynolds numbers, and coolant-to-freestream velocity ratios cannot be matched simultaneously to the engine condition. Furthermore, the coolant thermal transfer properties are unaccounted for in these parameters, despite their large influence on the resultant overall effectiveness. While a good deal of research has focused on the effects of the coolant-to-freestream density ratio, this study specifically examines the influence of other thermodynamic properties, in particular the specific heat, which differ substantially between experimental and engine conditions. This study demonstrates the influence of various coolant properties on the overall effectiveness distribution on a leading edge by selectively matching M, I, and ACR with air, argon and carbon dioxide coolants.

The Effects of an Excited Impinging Jet on the Local Heat Transfer Coefficient of Aircraft Turbine Blades

1988 ◽  
Author(s):  
P. J. Disimile ◽  
D. M. Paule
Keyword(s):  
Heat Transfer ◽  
Flow Field ◽  
Turbine Blade ◽  
Turbine Blades ◽  
Leading Edge ◽  
Local Heat Transfer ◽  
Impinging Jet ◽  
Local Heat ◽  
Primary Objective ◽  
Inlet Temperature

The primary objective of this paper is to present the results of research into the effects of periodic excitation upon the local heat transfer characteristics of a turbine blade cooled by an impinging jet of air. A curved plate (used to simulate the inner leading edge of a turbine blade) was subjected to a two-dimensional jet flow field (Re = 10,000) with a superimposed periodic acoustic disturbance. When compared to the naturally disturbed flow, the excited flow field was found to reduce the local Nusselt number and cool the blade less efficiently (by as much as ten percent in the extreme cases). The results of the study appear to indicate that harmonic disturbances present a serious controlling factor in the quest for optimization of turbine blade cooling techniques. By isolating dominant frequencies in gas turbine engines and working to suppress them, the authors believe it possible to make significant contributions towards the desired increase in turbine inlet temperature.

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