scholarly journals Comparison of Visualized Turbine Endwall Secondary Flows and Measured Heat Transfer Patterns

1984 ◽  
Vol 106 (1) ◽  
pp. 168-172 ◽  
Author(s):  
R. E. Gaugler ◽  
L. M. Russell
Keyword(s):  
Heat Transfer ◽  
Large Scale ◽  
Horseshoe Vortex ◽  
Secondary Flows ◽  
Turbine Vane ◽  
A New Technique ◽  
Exit Mach Number ◽  
Visualization Techniques ◽  
Neutrally Buoyant ◽  
First Time

Various flow visualization techniques were used to define the secondary flows near the endwall in a large scale turbine vane cascade. The cascade was scaled up from one used to generate endwall heat transfer data under a joint NASA-USAF contract. A comparison of the visualized flow patterns and the measured Stanton number distributions was made for cases where the inlet Reynolds number and exit Mach number were matched. Flows were visualized by using neutrally buoyant helium-filled soap bubbles, by using smoke from oil soaked cigars, and by a new technique using permanent marker pen ink dots and synthetic wintergreen oil. For the first time, details of the horseshoe vortex and secondary flows can be directly compared with heat transfer distributions. Near the cascade entrance, there is an obvious correlation between the two sets of data, but well into the passage the effect of secondary flow is not as obvious.

An Experimental Study Heat Transfer in a Large-Scale Turbine Rotor Passage

1994 ◽  
Vol 116 (1) ◽  
pp. 1-13 ◽  
Author(s):  
M. F. Blair
Keyword(s):  
Heat Transfer ◽  
Experimental Study ◽  
Large Scale ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Turbine Rotor ◽  
Secondary Flows ◽  
Vortex System ◽  
Tip Leakage ◽  
High Heat Transfer

An experimental study of the heat transfer distribution in a turbine rotor passage was conducted in a large-scale, ambient temperature, rotating turbine model. Heat transfer was measured for both the full-span suction and pressure surfaces of the airfoil and for the hub endwall surface. The objective of this program was to document the effects of flow three dimensionality on the heat transfer in a rotating blade row (versus a stationary cascade). Of particular interest were the effects of the hub and tip secondary flows, tip leakage, and the leading-edge horseshoe vortex system. The effect of surface roughness on the passage heat transfer was also investigated. Midspan results are compared with both smooth-wall and rough-wall finite-difference two-dimensional heat transfer predictions. Contour maps of Stanton number for both the rotor airfoil and endwall surfaces revealed numerous regions of high heat transfer produced by the three-dimensional flows within the rotor passage. Of particular importance are regions of local enhancement (as much as 100 percent over midspan values) produced on the airfoil suction surface by the secondary flows and tip-leakage vortices and on the hub endwall by the leading edge horseshoe vortex system.

An Experimental and Numerical Investigation of the Effect of Combustor-Nozzle Platform Misalignment on Endwall Heat Transfer at Transonic High Turbulence Conditions

2016 ◽  
Author(s):  
A. Arisi ◽  
D. Mayo ◽  
Z. Li ◽  
W. F. Ng ◽  
H. K. Moon ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Numerical Study ◽  
Horseshoe Vortex ◽  
Aerodynamic Characteristics ◽  
Secondary Flows ◽  
Navier Stokes ◽  
Vortex System ◽  
Ansys Fluent ◽  
Heat Transfer Augmentation ◽  
Exit Mach Number

A detailed experimental and numerical study has been conducted to investigate the endwall heat transfer characteristics on a nozzle platform that has been misaligned with the combustor exit, resulting in a backward facing step at the nozzle inlet. The study was carried out under transonic engine representative conditions with an exit Mach number of 0.85 (Reexit = 1.5 × 106), and an inlet turbulence intensity of 16%. A transient infrared thermography technique coupled with endwall static pressure ports, were used to map the endwall surface heat transfer and aerodynamic characteristics respectively. A numerical study was also conducted by solving the steady state Reynolds Averaged Navier Stokes (RANS) equations using the commercial CFD solver ANSYS Fluent v.15. The numerical results were then validated by comparing to experiment data and good agreement was observed. The results reveal that the classical endwall secondary flows (endwall crossflows, horseshoe and passage vortices) are weakened and a unique auxiliary vortex system develops within the passage and interacts with the weakened horseshoe vortex. It is observed that heat transfer in the first half of the passage endwall is heavily influenced by this auxiliary vortex system. Heat transfer augmentation of between 15% and 40% was also observed throughout the NGV endwall. Furthermore, the auxiliary vortex system results in a delayed cross-passage migration of the horseshoe vortex which consequently results in large lateral gradient in heat transfer downstream of the throat.

Platform Film Cooling Investigation on an HP Nozzle Vane Cascade With Discrete Shaped Holes and Slot Film Cooling

2020 ◽  
Author(s):  
G. Barigozzi ◽  
A. Perdichizzi ◽  
L. Abba ◽  
L. Pestelli
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Heat Transfer Performance ◽  
Leading Edge ◽  
Suction Side ◽  
Secondary Flows ◽  
Intensity Level ◽  
Transfer Performance ◽  
Film Cooling Effectiveness ◽  
Exit Mach Number

Abstract The present paper reports on an experimental investigation on the aerodynamic and heat transfer performance of different platform cooling schemes: two based on cylindrical and shaped holes and one featuring a slot located upstream of the leading edge plane simulating the combustor to stator interface gap. Tests were run on a 6-vane cascade operated at an isentropic cascade exit Mach number of 0.4 and a significant inlet turbulence intensity level of about 9%. The cooling schemes were first tested to quantify their impact on secondary flows and related losses for variable injection conditions. Heat transfer performance was then assessed through adiabatic film cooling effectiveness and heat transfer coefficient measurements. The Net Heat Flux Reduction parameter was then computed to critically assess the cooling schemes. When compared with the cylindrical hole scheme, shaped holes outperform for all tested injection rates, while the slot alone is able to thermally protect only the front of the passage. Discrete holes are required to cool the platform region along the whole pressure side and the suction side leading edge region.

A Cold Heat Transfer Tunnel for Gas Turbine Research on an Annular Cascade

1993 ◽  
Author(s):  
R. F. Martinez-Botas ◽  
A. J. Main ◽  
G. D. Lock ◽  
T. V. Jones
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Region Of Interest ◽  
Secondary Flows ◽  
Test Duration ◽  
Pressure Distributions ◽  
Exit Mach Number ◽  
Nozzle Guide ◽  
Turbine Design ◽  
Annular Cascade

The Oxford University Blowdown Tunnel has been substantially modified to test a large annular cascade of high pressure nozzle guide vanes (mean blade diameter of 1.11 m and axial chord of 0.0673 m). The new transonic facility has been constructed to obtain complete contours of heat transfer coefficient for both the end walls and blade surfaces using the transient liquid crystal technique, to measure pressure distributions and losses, and to study fundamental aspects of boundary layers and secondary flows. The facility allows an independent variation of Reynolds and Mach numbers, providing aerodynamic and heat transfer measurements in the region of interest for gas turbine design. The mass flow rate through the cascade at NGV design conditions (exit Mach number 0.96 and Reynolds number 2.0 × 106) is 38 kg/s and the pressure-regulated test duration exceeds 7 seconds.

Effects of Deposits on Film Cooling of a Vane Endwall Along the Pressure Side

2008 ◽  
Vol 130 (4) ◽  
Author(s):  
N. Sundaram ◽  
M. D. Barringer ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Secondary Flows ◽  
Pressure Side ◽  
Surface Deposition ◽  
Hot Gas ◽  
Turbine Vane ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes ◽  
Effect Of Surface

Film cooling is influenced by surface roughness and depositions that occur from contaminants present in the hot gas path, whether that film cooling occurs on the vane itself or on the endwalls associated with the vanes. Secondary flows in the endwall region also affect the film-cooling performance along the endwall. An experimental investigation was conducted to study the effect of surface deposition on film cooling along the pressure side of a first-stage turbine vane endwall. A large-scale wind tunnel with a turbine vane cascade was used to perform the experiments. The vane endwall was cooled by an array of film-cooling holes along the pressure side of the airfoil. Deposits having a semielliptical shape were placed along the pressure side to simulate individual row and multiple row depositions. Results indicated that the deposits lowered the average adiabatic effectiveness levels downstream of the film-cooling rows by deflecting the coolant jets toward the vane endwall junction on the pressure side. Results also indicated that there was a steady decrease in adiabatic effectiveness levels with a sequential increase in the number of rows with the deposits.

An Experimental Study of Heat Transfer and Film Cooling on Large-Scale Turbine Endwalls

1974 ◽  
Vol 96 (4) ◽  
pp. 524-529 ◽  
Author(s):  
M. F. Blair
Keyword(s):  
Heat Transfer ◽  
Thermal Conductivity ◽  
Film Cooling ◽  
Large Scale ◽  
Convective Heat Transfer Coefficient ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Test Models ◽  
Turbine Vane ◽  
Flow Vortex

Experiments were conducted to determine the film cooling effectiveness and convective heat transfer coefficient distributions on the endwall of a large-scale turbine vane passage. The vane test models employed simulated the passage geometry and upstream cooling slot geometry of a typical first stage turbine. The test models were constructed of low thermal conductivity foam and foil heaters. The tests were conducted at a typical engine Reynolds number but at lower than typical Mach numbers. The film cooling effectiveness distribution for the entire endwall and the heat transfer distribution for the downstream one-half of the endwall were characterized by large gapwise variations which were attributed to a secondary flow vortex.

scholarly journals Direct Numerical Simulations of a High-Pressure Turbine Vane

2016 ◽  
Vol 138 (7) ◽  
Author(s):  
Andrew P. S. Wheeler ◽  
Richard D. Sandberg ◽  
Neil D. Sandham ◽  
Richard Pichler ◽  
Vittorio Michelassi ◽  
...  
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Surface Flow ◽  
Surface Heat ◽  
Suction Surface ◽  
Data Set ◽  
Turbulent Dissipation ◽  
Surface Heat Transfer ◽  
Turbine Vane ◽  
Exit Mach Number

In this paper, we establish a benchmark data set of a generic high-pressure (HP) turbine vane generated by direct numerical simulation (DNS) to resolve fully the flow. The test conditions for this case are a Reynolds number of 0.57 × 106 and an exit Mach number of 0.9, which is representative of a modern transonic HP turbine vane. In this study, we first compare the simulation results with previously published experimental data. We then investigate how turbulence affects the surface flow physics and heat transfer. An analysis of the development of loss through the vane passage is also performed. The results indicate that freestream turbulence tends to induce streaks within the near-wall flow, which augment the surface heat transfer. Turbulent breakdown is observed over the late suction surface, and this occurs via the growth of two-dimensional Kelvin–Helmholtz spanwise roll-ups, which then develop into lambda vortices creating large local peaks in the surface heat transfer. Turbulent dissipation is found to significantly increase losses within the trailing-edge region of the vane.

An Experimental Study of Heat Transfer in a Large-Scale Turbine Rotor Passage

1992 ◽  
Author(s):  
Michael F. Blair
Keyword(s):  
Heat Transfer ◽  
Experimental Study ◽  
Large Scale ◽  
Leading Edge ◽  
High Heat ◽  
Turbine Rotor ◽  
Secondary Flows ◽  
Suction Surface ◽  
Tip Leakage ◽  
High Heat Transfer

An experimental study of the heat transfer distribution in a turbine rotor passage was conducted in a large–scale, ambient temperature, rotating turbine model. Meat transfer was measured for both the full–span suction and pressure surfaces of the airfoil as well as for the hub endwall surface. The objective of this program was to document the effects of flow three–dimensionality on the heat transfer in a rotating blade row (vs. a stationary cascade). Of particular interest were the effects of the hub and tip secondary flows, tip leakage and the leading–edge horseshoe vortexsystem. The effect of surface roughness on the passage heat transfer was also investigated. Midspan results are compared with both smooth–wall and rough–wall finite–difference two dimensional heat transfer predictions. Contour maps of Stanton number for both the rotor airfoil and endwall surfaces revealed numerous regions of high heat transfer produced by the three dimensional flows within the rotor passage. Of particular importance are regions of local enhancement (as much as 100% over midspan values) produced on the airfoil suction surface by the secondary flows and tip–leakage vortices and on the hub endwall by the leading–edge horseshoe vortex system.

The Augmentation of Internal Blade Tip-Cap Cooling by Arrays of Shaped Pins

2007 ◽  
Author(s):  
Ronald S. Bunker
Keyword(s):  
Heat Transfer ◽  
Surface Area ◽  
Smooth Surface ◽  
Large Scale ◽  
Secondary Flows ◽  
Scale Model ◽  
Transfer Coefficients ◽  
Inlet Channel ◽  
Liquid Crystal Thermography ◽  
Aspect Ratios

The objective of the present study is to demonstrate a method to provide substantially increased convective heat flux on the internal cooled tip cap of a turbine blade. The new tip cap augmentation consists of several variations involving the fabrication or placement of arrays of discrete shaped pins on the internal tip cap surface. Due to the nature of flow in a 180-degree turn, the augmentation mechanism and geometry have been designed to accommodate a mixture of impingement-like flow, channel flow, and strong secondary flows. A large-scale model of a sharp 180-degree tip turn is used with the liquid crystal thermography method to obtain detailed heat transfer distributions over the internal tip cap surface. Inlet channel Reynolds numbers range from 200,000 to 450,000 in this study. The inlet and exit passages have aspect ratios of 2:1, while the tip turn divider-to-cap distance maintains nearly the same hydraulic diameter as the passages. Five tip cap surfaces were tested including a smooth surface, two different heights of aluminum pin arrays, one more closely spaced pin array, and one pin array made of insulating material. Effective heat transfer coefficients based on the original smooth surface area were increased by up to a factor of 2.5. Most of this increase is due to the added surface area of the pin array. However, factoring this surface area effect out shows that the heat transfer coefficient has also been increased by about 20 to 30%, primarily over the base region of the tip cap itself. This augmentation method resulted in negligible increase in tip turn pressure drop over that of a smooth surface.

scholarly journals Streakline Flow Visualization Study of a Horseshoe Vortex in a Large-Scale, Two-Dimensional Turbine Stator Cascade

1980 ◽  
Author(s):  
R. E. Gaugler ◽  
L. M. Russell
Keyword(s):  
Boundary Layer ◽  
Large Scale ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Main Stream ◽  
Two Dimensional ◽  
Vortex System ◽  
Turbine Stator ◽  
Neutrally Buoyant ◽  
Limiting Streamlines

Neutrally buoyant helium-filled bubbles were observed as they followed the streamlines in a horseshoe vortex system around the vane leading edge in a large-scale, two-dimensional, turbine stator cascade. Inlet Reynolds number, based on true chord, ranged between 1.0 × 10 ⅝ to 3.0 × 10 ⅝. Bubbles were introduced into the endwall boundary layer through a slot upstream of the vane leading edge. The paths of the bubbles were recorded photographically as streaklines on 16-mm movie film. Individual frames from the film have been selected, and overlayed to show the details of the horseshoe vortex around the leading edge, the transport of the vortex across the passage near the leading edge is clearly seen when compared to the streaks formed by bubbles carried in the main stream. Limiting streamlines on the endwall surface were traced by the flow of oil drops.

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