Quantitative MRI Measurements of Hot Streak Development in a Turbine Vane Cascade

2015 ◽  
Author(s):  
Sayuri D. Yapa ◽  
Christopher J. Elkins ◽  
John K. Eaton
Keyword(s):  
Film Cooling ◽  
Thermal Stresses ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Turbine Rotor ◽  
Quantitative Mri ◽  
Inlet Temperature ◽  
Cooling Design ◽  
Pass Through ◽  
Hot Streak

Hot streaks from the combustor and cool streaks from nozzle vane film cooling impose strong inlet temperature variations on high pressure turbine blades, which can lead to local hot or cold spots, high thermal stresses, and fatigue failures. Furthermore, the complex three dimensional flows around the vane may act to concentrate cool or hot fluid exiting the vane row. In order to optimize the cooling design of the turbine blades, the designer must be able to predict the temperature distribution entering the turbine rotor. Therefore, it is important to understand and predict how combustor hot streaks are dispersed as they pass through the vane row. The goal of the present work is to provide detailed three dimensional velocity and temperature data for simulated combustor hot streaks developing through a film cooled vane cascade using the Magnetic Resonance Velocity/Concentration experimental technique. The measurements show that the hot streaks are thinned by acceleration through the vane cascade and diffused by turbulence. The turbulent diffusivity is suppressed by acceleration and leaves significant temperature nonuniformity in the vane wake.

scholarly journals Recent Development in Turbine Blade Film Cooling

2001 ◽  
Vol 7 (1) ◽  
pp. 21-40 ◽  
Author(s):  
Je-Chin Han ◽  
Srinath Ekkad
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbines ◽  
High Performance ◽  
Pressure Ratio ◽  
Turbine Blades ◽  
Protective Layer ◽  
Turbine Rotor ◽  
Industrial Applications ◽  
Inlet Temperature

Gas turbines are extensively used for aircraft propulsion, land-based power generation, and industrial applications. Thermal efficiency and power output of gas turbines increase with increasing turbine rotor inlet temperature (RIT). The current RIT level in advanced gas turbines is far above the .melting point of the blade material. Therefore, along with high temperature material development, a sophisticated cooling scheme must be developed for continuous safe operation of gas turbines with high performance. Gas turbine blades are cooled internally and externally. This paper focuses on external blade cooling or so-called film cooling. In film cooling, relatively cool air is injected from the inside of the blade to the outside surface which forms a protective layer between the blade surface and hot gas streams. Performance of film cooling primarily depends on the coolant to mainstream pressure ratio, temperature ratio, and film hole location and geometry under representative engine flow conditions. In the past number of years there has been considerable progress in turbine film cooling research and this paper is limited to review a few selected publications to reflect recent development in turbine blade film cooling.

Experimental-Based Redesigns for Trailing Edge Film Cooling of Gas Turbine Blades

2013 ◽  
Vol 135 (4) ◽  
Author(s):  
Michael Benson ◽  
Sayuri D. Yapa ◽  
Chris Elkins ◽  
John K. Eaton
Keyword(s):  
Magnetic Resonance ◽  
Film Cooling ◽  
Large Scale ◽  
Thermal Stresses ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Flow Structures ◽  
Mean Flow ◽  
Trailing Edge ◽  
Mean Velocity

Magnetic resonance imaging experiments have provided the three-dimensional mean concentration and three component mean velocity field for a typical trailing edge film-cooling cutback geometry built into a conventional uncambered airfoil. This geometry is typical of modern aircraft engines and includes three dimensional slot jets separated by tapered lands. Previous analysis of these data identified the critical mean flow structures that contribute to rapid mixing and low effectiveness in the fully turbulent flow. Three new trailing edge geometries were designed to modify the large scale mean flow structures responsible for surface effectiveness degradation. One modification called the Dolphin Nose attempted to weaken strong vortex flows by reducing three dimensionality near the slot breakout. This design changed the flow structure but resulted in minimal improvement in the surface effectiveness. Two other designs called the Shield and Rounded Shield changed the land planform and added an overhanging land edge while maintaining the same breakout surface. These designs substantially modified the vortex structure and improved the surface effectiveness by as much as 30%. Improvements included superior coolant uniformity on the breakout surface which reduces potential thermal stresses. The utilization of the time averaged data from combined magnetic resonance velocimetry (MRV) and concentration (MRC) experiments for designing improved trailing edge breakout film cooling is demonstrated.

Analysis of Hot Streak Effects on Turbine Rotor Heat Load

1997 ◽  
Vol 119 (3) ◽  
pp. 544-553 ◽  
Author(s):  
T. Shang ◽  
A. H. Epstein
Keyword(s):  
Heat Load ◽  
Three Dimensional ◽  
Axial Flow ◽  
Turbine Rotor ◽  
Inlet Temperature ◽  
Pressure Surface ◽  
Physical Mechanisms ◽  
Design Variables ◽  
Hot Streak ◽  
Turbine Design

The influence of inlet hot streak temperature distortion on turbine blade heat load was explored on a transonic axial flow turbine stage test article using a three-dimensional, multiblade row unsteady Euler code. The turbine geometry was the same as that used for a recently reported testing of hot streak influence. Emphasis was placed on elucidating the physical mechanisms by which hot streaks affect turbine durability. It was found that temperature distortion significantly increases both blade surface heat load nonuniformity and total blade heat load by as much as 10–30 percent (mainly on the pressure surface), and that the severity of this influence is a strong function of turbine geometry and flow conditions. Three physical mechanisms were identified that drive the heat load nonuniformity: buoyancy, wake convection (the Kerrebrock–Mikolajczak effect), and Rotor–Stator interactions. The latter can generate significant nonuniformity of the time-averaged relative frame rotor inlet temperature distribution. Dependence of these effects on turbine design variables was investigated to shed light on the design space, which minimizes the adverse effects of hot streaks.

Novel Turbine Rotor Shroud Film-Cooling Design and Validation: Part 2

2009 ◽  
Author(s):  
K. S. Chana ◽  
B. Haller
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Cooling System ◽  
Three Dimensional ◽  
Poor Performance ◽  
Turbine Rotor ◽  
Test Facility ◽  
Test Results ◽  
Film Cooling Effectiveness ◽  
Cooling Design

This paper is part two of a two part paper which considers a shroud film-cooling system design. The design was carried out using test results from a previous two-dimensional (2D) design and optimisation using three-dimensional (3D) CFD. The first cooling design was carried out using a streamline boundary layer approach and tested in the QinetiQ turbine test facility (TTF). The test results showed the design did not function as well as had been predicted and gave a poor performance in terms of film cooling effectiveness. Lessons learnt from the 2D design as well as understanding gained from heat transfer and pressure data taken on the rotor casing led to the formulation of a completely new design philosophy. Accepting, cooling films would not survive rotor passing and therefore concentrating on localised cooling as well as the re-establishment of cooling films between rotor passings. The design concept was validated/optimised with the aid of 3D CFD. Heat transfer instrumentation was implemented in a cooling insert fitted over the test rotor to evaluate the performance of the design. Tests carried out with and without cooling showed an improvement in cooling performance, leading to a 40% reduction in heat transfer rate to the rotor casing across the rotor overtip region. A significant improvement was achieved with the new design over the original with reductions in casing heat transfer rates of up to 44%, with a design coolant mass flow of 1.85% of core flow. Heat transfer data were successfully processed to Nusselt number, allowing the results to be translated to a gas turbine engine design.

Experimental-Based Redesigns for Trailing Edge Film Cooling of Gas Turbine Blades

2012 ◽  
Author(s):  
Michael Benson ◽  
Sayuri Yapa ◽  
Chris Elkins ◽  
John K. Eaton
Keyword(s):  
Magnetic Resonance ◽  
Film Cooling ◽  
Large Scale ◽  
Thermal Stresses ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Flow Structures ◽  
Mean Flow ◽  
Trailing Edge ◽  
Mean Velocity

Magnetic resonance imaging experiments have provided the three-dimensional mean concentration and three component mean velocity field for a typical trailing edge film-cooling cutback geometry built into a conventional uncambered airfoil. This geometry is typical of modern aircraft engines and includes three dimensional slot jets separated by tapered lands. Previous analysis of these data identified the critical mean flow structures that contribute to rapid mixing and low effectiveness in the fully turbulent flow. Three new trailing edge geometries were designed to modify the large scale mean flow structures responsible for surface effectiveness degradation. One modification called the Dolphin Nose attempted to weaken strong vortex flows by reducing three dimensionality near the slot breakout. This design changed the flow structure but resulted in minimal improvement in the surface effectiveness. Two other designs called the Shield and Rounded Shield changed the land planform and added an overhanging land edge while maintaining the same breakout surface. These designs substantially modified the vortex structure and improved the surface effectiveness by as much as 30%. Improvements included superior coolant uniformity on the breakout surface which reduces potential thermal stresses. The utilization of the time averaged data from combined magnetic resonance velocimetry (MRV) and concentration (MRC) experiments for designing improved trailing edge breakout film cooling is demonstrated.

Analysis of Hot Streak Effects on Turbine Rotor Heat Load

1996 ◽  
Author(s):  
Tonghuo Shang ◽  
Alan H. Epstein
Keyword(s):  
Heat Load ◽  
Three Dimensional ◽  
Axial Flow ◽  
Turbine Rotor ◽  
Inlet Temperature ◽  
Pressure Surface ◽  
Physical Mechanisms ◽  
Design Variables ◽  
Hot Streak ◽  
Turbine Design

The influence of inlet hot streak temperature distortion on turbine blade heat load was explored on a transonic axial flow turbine stage test article using a three-dimensional, multi-blade row unsteady Euler code. The turbine geometry was the same as that used for a recently reported testing of hot streak influence. Emphasis was placed elucidating the physical mechanisms by which hot streaks affect turbine durability. It was found that temperature distortion significantly increases both blade surface heat load nonuniformity and total blade heat load by as much as 10–30% (mainly on the pressure surface), and that the severity of this influence is a strong function of turbine geometry and flow conditions. Three physical mechanisms were identified which drive the heat load nonuniformity — buoyancy, wake convection (the Kerrebrock-Mikolajczak effect), and rotor-stator interactions. The latter can generate significant nonuniformity of the time-averaged relative frame rotor inlet temperature distribution. Dependence of these effects on turbine design variables was investigated to shed light on the design space which minimizes the adverse effects of hot streaks.

scholarly journals Comparison of Two-Equation Turbulence Models for Prediction of Heat Transfer on Film-Cooled Turbine Blades

1997 ◽  
Author(s):  
Vijay K. Garg ◽  
Ali A. Ameri
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Turbulence Models ◽  
Computer Time ◽  
Suction Surface ◽  
The Heat Transfer Coefficient

A three-dimensional Navier-Stokes code has been used to compute the heat transfer coefficient on two film-cooled turbine blades, namely the VKI rotor with six rows of cooling holes including three rows on the shower head, and the C3X vane with nine rows of holes including five rows on the shower head. Predictions of heat transfer coefficient at the blade surface using three two-equation turbulence models, specifically, Coakley’s q-ω model, Chien’s k-ε model and Wilcox’s k-ω model with Menter’s modifications, have been compared with the experimental data of Camci and Arts (1990) for the VKI rotor, and of Hylton et al. (1988) for the C3X vane along with predictions using the Baldwin-Lomax (B-L) model taken from Garg and Gaugler (1995). It is found that for the cases considered here the two-equation models predict the blade heat transfer somewhat better than the B-L model except immediately downstream of the film-cooling holes on the suction surface of the VKI rotor, and over most of the suction surface of the C3X vane. However, all two-equation models require 40% more computer core than the B-L model for solution, and while the q-ω and k-ε models need 40% more computer time than the B-L model, the k-ω model requires at least 65% more time due to slower rate of convergence. It is found that the heat transfer coefficient exhibits a strong spanwise as well as streamwise variation for both blades and all turbulence models.

Numerical and Experimental Investigation for Internal Cooling of Two Pass Gas Turbine Blades Channels Using Broken V Ribs

2014 ◽  
Author(s):  
Sourabh Kumar ◽  
R. S. Amano
Keyword(s):  
Gas Turbine ◽  
Thermal Stresses ◽  
Turbine Blades ◽  
Research Work ◽  
Local Heat Transfer ◽  
Local Heat ◽  
Inlet Temperature ◽  
Internal Cooling ◽  
Cooling Channel ◽  
Square Duct

Improvements in the thermal efficiency of a gas turbine can be obtained by operating it at high inlet temperatures. This high inlet temperature develops high thermal stresses on the turbine blades in addition to improving the performance. Cooling methodologies are implemented inside the blades to withstand those high temperatures. Four different combinations of broken 60° V ribs in cooling channel are considered. The research work investigates and compares numerically and experimentally, internal cooling of channels with broken V ribs. Local heat transfer in a square duct roughened with 60° V broken ribs is investigated for three different Reynolds numbers. Aspect ratio of the channel is taken to be 1:1. The pitch of the rib is considered to be 10 times the width of the rib, which is 0.0635 m. The square cross section of the channel is 0.508 × 0.508 m2 with 0.6096 m length. This study provides information about the best configuration of a broken V rib in a cooling channel.

scholarly journals Numerical Prediction of Film Cooling Effectiveness and the Associated Aerodynamic Losses With a Three-Dimensional Calculation Procedure

1990 ◽  
Author(s):  
G. H. Dibelius ◽  
R. Pitt ◽  
B. Wen
Keyword(s):  
Experimental Data ◽  
Film Cooling ◽  
Reverse Flow ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Temperature Pattern ◽  
Main Stream ◽  
Cooling Efficiency ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Losses

Film cooling of turbine blades by injecting air through holes or slots affects the main stream flow. A numerical model has been developed to predict the resulting three-dimensional flow and the temperature pattern under steady flow conditions. An elliptic procedure is used in the near injection area to include reverse flow situations, while in the upstream area as well as far downstream a partial-parabolic procedure is applied. As first step an adiabatic wall has been assumed as boundary condition, since for this case experimental data are readily available for comparison. At elevated momentum blowing rates, zones of reverse flow occur downstream of the injection holes resulting in a decrease of cooling efficiency. A variation of the relevant parameters momentum blowing rate m, injection angle α and ratio of hole spacing to diameter s/d revealed the combination of m ≈ 1, α ≈ 30° and s/d ≈ 2 to be the optimum with respect to the averaged cooling efficiency and to the aerodynamic losses. Cooling is more efficient with slots than with a row of holes not considering the related problems of manufacture and service life. The calculated temperature patterns compare well with the experimental data available.

Influence of Film Cooling Hole Angles and Geometries on Aerodynamic Loss and Net Heat Flux Reduction

2011 ◽  
Author(s):  
Chia Hui Lim ◽  
Graham Pullan ◽  
Peter Ireland
Keyword(s):  
Heat Flux ◽  
Film Cooling ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Design Parameters ◽  
Aerodynamic Loss ◽  
Design Engineers ◽  
Net Heat Flux ◽  
Cooling Hole ◽  
Shaped Hole

Turbine design engineers have to ensure that film cooling can provide sufficient protection to turbine blades from the hot mainstream gas, while keeping the losses low. Film cooling hole design parameters include inclination angle (α), compound angle (β), hole inlet geometry and hole exit geometry. The influence of these parameters on aerodynamic loss and net heat flux reduction is investigated, with loss being the primary focus. Low-speed flat plate experiments have been conducted at momentum flux ratios of IR = 0.16, 0.64 and 1.44. The film cooling aerodynamic mixing loss, generated by the mixing of mainstream and coolant, can be quantified using a three-dimensional analytical model that has been previously reported by the authors. The model suggests that for the same flow conditions, the aerodynamic mixing loss is the same for holes with different α and β but with the same angle between the mainstream and coolant flow directions (angle κ). This relationship is assessed through experiments by testing two sets of cylindrical holes with different α and β: one set with κ = 35°, another set with κ = 60°. The data confirm the stated relationship between α, β, κ and the aerodynamic mixing loss. The results show that the designer should minimise κ to obtain the lowest loss, but maximise β to achieve the best heat transfer performance. A suggestion on improving the loss model is also given. Five different hole geometries (α = 35.0°, β = 0°) were also tested: cylindrical hole, trenched hole, fan-shaped hole, D-Fan and SD-Fan. The D-Fan and the SD-Fan have similar hole exits to the fan-shaped hole but their hole inlets are laterally expanded. The external mixing loss and the loss generated inside the hole are compared. It was found that the D-Fan and the SD-Fan have the lowest loss. This is attributed to their laterally expanded hole inlets, which lead to significant reduction in the loss generated inside the holes. As a result, the loss of these geometries is ≈ 50% of the loss of the fan-shaped hole at IR = 0.64 and 1.44.

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