The Experimental and Computational Study of a New Cooling Strategy for Turbomachinery Rotational Components

2005 ◽  
Author(s):  
N. I. Elhabeshi ◽  
S. M. Guo
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Closed Loop ◽  
Imaging System ◽  
Computational Study ◽  
Turbine Blades ◽  
Temperature Fields ◽  
Heat Transfer Fluid ◽  
Rotational Components

Many cooling methods have been developed for gas turbine rotational components. Film cooling is one of the most commonly used technique, which introduces cold air through film cooling holes onto the outer surfaces of gas turbine components to protect them from hot mainstreams. Although film cooling is very effective, it is well known that film cooling could reduce the engine power rating and introduce losses, mainly due to the direct coolant/mainstream interactions. In this paper, a preliminary study of a new closed-loop rotor cooling technique is reported. Filled with suitable heat transfer fluid, this closed loop design is applicable to turbomachinery rotors. Due to the combined buoyancy-centrifugal forces under turbine working conditions, the heat transfer fluid could transport the heat effectively from the outer surface of the blade to the internal cooling air. Instead of relying on the temperature gradient inside the blade walls for heat absorption from the high temperature mainstream side to the low temperature cooling airside, these closed loops act as heat transport superhighways. With proper insulation, the center part of the blade walls could be kept at a lower temperature, comparing to the conventional cooling designs. This would potentially provide vital important extra mechanical strength to the highly loaded turbine blades. Preliminary experimental work and CFD predictions have been conducted to prove this novel design. The tests were conducted using a rotational disk at a speed up to 1500 rpm. A liquid crystal based imaging system was employed to measure the surface temperature under transient conditions. Steady/unsteady 2D/3D CFD predictions were carried out to provide detailed information about the flow and temperature fields inside the proposed cooling configuration. Both experimental and computational work proved that this new design would work, providing the mechanical and manufacturing requirements could be met.

AERODYNAMICS AND HEAT TRANSFER INSIDE A GAS TURBINE MID-PASSAGE GAP

2021 ◽  
pp. 1-13
Author(s):  
Faisal Shaikh ◽  
Budimir Rosic
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Heat Load ◽  
Turbine Blades ◽  
Turbine Stage ◽  
Gas Turbine Blades ◽  
Cooling Flows ◽  
Turbine Efficiency ◽  
Gap Flow

Abstract Gas turbine blades and vanes are typically manufactured with small clearances between adjacent vane and blade platforms, termed the midpassage gap. The midpassage gap reduces turbine efficiency and causes additional heat load into the vane platform, as well as changing the distribution of endwall heat transfer and film cooling. This paper presents a low-order analytical analysis to quantify the effects of the midpassage gap on aerodynamics and heat transfer, verified against an experimental campaign and CFD. Using this model, the effects of the gap can be quantified, for a generic turbine stage, based only on geometric features and the passage static pressure field. It is found that at present there are significant losses and a large proportion of heat load caused by the gap, but that with modified design this could be reduced to negligible levels. Cooling flows into the gap to prevent ingression are investigated analytically and with CFD. Recommendations are given for targets that turbine designers should work toward in reducing the adverse effects of the midpassage gap. A method to estimate the effect of gap flow is presented, so that for any machine the significance of the gap may be assessed.

High-Pressure Gas Turbine Vane Turbulent Flows and Heat Transfer Predicted by RANS/LES/DES

2017 ◽  
Author(s):  
Ping Dong ◽  
R. S. Amano
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Gas Turbine ◽  
Turbulent Flows ◽  
Film Cooling ◽  
Turbine Blades ◽  
Suction Side ◽  
Gas Turbine Blades ◽  
Flow And Heat Transfer ◽  
Turbine Vane

The lifetime of the modern gas turbines greatly depends on the durability of hot section components operating at high temperatures. Film cooling is key to air cooling technologies in modern gas turbine and widely used in high-temperature and high-pressure blades as an active cooling scheme. The requirements of accurate prediction of aerodynamic flow and heat transfer in gas turbine blades lay the essential foundation of cooling effectiveness improvement and working life estimation. In recent days, Large Eddy Simulations (LES) is considered as a useful tool to predict turbulent flows and heat transfer around gas turbine blades, but, comparing to the Reynolds-Averaged Navier–Stokes (RANS) methods, the LES method usually needs more computing resource and depends on computational power and mesh quality. In this paper, LES/DES (Detached Eddy Simulation) predictions were compared to RANS prediction with interest in the accuracy and improvement of turbulent flow and heat transfer phenomena around NASA’s C3X high-pressure gas turbine vane with leading edge cooling film. RANS/LES/DES were detailed and further investigated to assess their ability to predict flow and heat transfer in boundary layer around C3X vane. The current predictions showed that the mix between film cooling injections and free stream resulted in complex flow and heat transfer in the boundary layer on the external vane surface. The predictions of the aerodynamic load along the C3X vane with RANS/LES/DES were almost identical and agreed well with the experimental results. However, the heat transfer predictions with RANS/LES/DES were different. The transition prediction showed the best agreement with the experiment data in the most region. The LES prediction only partially agreed with the experimental data before separation point on the suction side and mild pressure gradient region on the pressure side. The DES and RANS predictions agreed with the experiment data after separation point on the suction side and most region on the pressure side.

Computational Modeling of Tip Heat Transfer to a Superscale Model of an Unshrouded Gas Turbine Blade

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Brian M. T. Tang ◽  
Pepe Palafox ◽  
Brian C. Y. Cheong ◽  
Martin L. G. Oldfield ◽  
David R. H. Gillespie
Keyword(s):  
Heat Transfer ◽  
Experimental Data ◽  
Nusselt Number ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Computational Study ◽  
Turbine Blades ◽  
Turbulence Models ◽  
Tip Leakage Flow ◽  
Blade Tip

Control of over-tip leakage flow between turbine blade tips and the stationary shroud is one of the major challenges facing gas turbine designers today. The flow imposes large thermal loads on unshrouded high pressure (HP) turbine blades and is significantly detrimental to turbine blade life. This paper presents results from a computational study performed to investigate the detailed blade tip heat transfer on a sharp-edged, flat tip HP turbine blade. The tip gap is engine representative at 1.5% of the blade chord. Nusselt number distributions on the blade tip surface have been obtained from steady flow simulations and are compared with experimental data carried out in a superscale cascade, which allows detailed flow and heat transfer measurements in stationary and engine representative conditions. Fully structured, multiblock hexahedral meshes were used in the simulations performed in the commercial solver FLUENT. Seven industry-standard turbulence models and a number of different tip gridding strategies are compared, varying in complexity from the one-equation Spalart–Allmaras model to a seven-equation Reynolds stress model. Of the turbulence models examined, the standard k-ω model gave the closest agreement to the experimental data. The discrepancy in Nusselt number observed was just 5%. However, the size of the separation on the pressure side rim was underpredicted, causing the position of reattachment to occur too close to the edge. Other turbulence models tested typically underpredicted Nusselt numbers by around 35%, although locating the position of peak heat flux correctly. The effect of the blade to casing motion was also simulated successfully, qualitatively producing the same changes in secondary flow features as were previously observed experimentally, with associated changes in heat transfer with the blade tip.

scholarly journals Heat Transfer Phenomena in Gas Turbines

1980 ◽  
Author(s):  
J. R. Taylor
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Gas Turbines ◽  
Turbine Blades ◽  
Gas Turbine Engine ◽  
Transfer Coefficients ◽  
Turbine Engine ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients

A discussion of the problems encountered in prediction of heat transfer in the turbine section of a gas turbine engine is presented. Areas of current gas turbine engine is presented. Areas of current concern to designers where knowledge is deficient or lacking are elucidated. Consideration is given to methods and problems associated with determination of heat transfer coefficients, external gas temperatures, and, where applicable, film cooling effectiveness. The paper is divided into parts dealing with turbine airfoil heat transfer, endwall heat transfer, and heat transfer in the internal cavities of cooled turbine blades. Recent literature dealing with these topics is listed.

Two-Phase Flow Simulation of Mist Film Cooling on Turbine Blades With Conjugate Internal Cooling

2008 ◽  
Vol 130 (10) ◽  
Author(s):  
Xianchang Li ◽  
Ting Wang
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Turbine Blades ◽  
Flow Simulation ◽  
Solid Base ◽  
Internal Cooling ◽  
Adiabatic Wall ◽  
Two Phase ◽  
Air Film

Effective cooling of gas turbine combustor liners, combustor transition pieces, turbine vanes (nozzles), and blades (buckets) is a critical task to protect these components from the flue gas at extremely high temperature. Air film cooling has been successfully used to cool these hot sections for the past half century. However, the net benefits from the traditional methods seem to be incremental, but the temperature of working gas is continuously increasing to achieve high thermal efficiency. Therefore, new cooling techniques need to be developed. One of the promising techniques is to enhance film cooling with mist injection. While the previous study reported the effect of mist on the cooling effectiveness with an adiabatic wall, this paper focuses on the effect of mist injection on heat transfer of film cooling with a nonadiabatic flat wall, using the commercial computational fluid dynamics software package FLUENT. Both 2D and 3D cases are considered with a 2D slot and diffusive compound-angle holes. Modeling of the interaction of a droplet with a uniformly cooled wall as well as conjugate heat conduction inside the solid base are conducted. Different mist droplet sizes and mist concentrations are adopted. Conditions both in a gas turbine operating environment (15 atm and 1561 K) and in a laboratory environment (1 atm and 450 K) are considered. Results show that injecting 2–10% mist reduces the heat transfer coefficient and the wall temperature. Especially, mist has the prolonged effect of cooling the region downstream for 15 jet hole diameters, where conventional air film cooling is not effective.

scholarly journals Turbulent Flow and Heat Transfer in Gas Turbine Blade Cooling Passages

1992 ◽  
Author(s):  
F. J. Cunha
Keyword(s):  
Heat Transfer ◽  
Turbulent Flow ◽  
Gas Turbine ◽  
Turbine Blades ◽  
Operating Conditions ◽  
Temperature Fields ◽  
Air Cooling ◽  
Turbine Blade Cooling ◽  
Gas Turbine Blades ◽  
Analysis Methodology

A numerical analysis methodology has been created to predict the heat transfer within the air cooling passages of gas turbine blades. In this paper, the turbulent flow heat convection with developed velocity and temperature fields is studied for cavities with turbulators. The influence of Coriolis forces and rotational buoyancy effects were also included. The k-equation turbulence model was employed over most of the cross section while a modified Van Driest’s version of the mixing length hypothesis is used in the near-wall sublayer. This methodology was successfully benchmarked against experimental results for air cooling passages of turbine blades. Analytical results are presented in terms of the Reynolds, Rossby and rotational Rayleigh numbers for realistic operating conditions.

Aerodynamics and Heat Transfer Inside a Gas Turbine Mid-Passage Gap

2020 ◽  
Author(s):  
Faisal Shaikh ◽  
Budimir Rosic
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Heat Load ◽  
Turbine Blades ◽  
Turbine Stage ◽  
Gas Turbine Blades ◽  
Cooling Flows ◽  
Turbine Efficiency ◽  
Gap Flow

Abstract Gas turbine blades and vanes are typically manufactured with small clearances between adjacent vane and blade platforms, termed the midpassage gap. The midpassage gap reduces turbine efficiency and causes additional heat load into the vane platform, as well as changing the distribution of endwall heat transfer and film cooling. This paper presents a low-order analytical analysis to quantify the effects of the midpassage gap on aerodynamics and heat transfer, verified against an experimental campaign and CFD. Using this model, the effects of the gap can be quantified, for a generic turbine stage, based only on geometric features and the passage static pressure field. It is found that at present there are significant losses and a large proportion of heat load caused by the gap, but that with modified design this could be reduced to negligible levels. Cooling flows into the gap to prevent ingression are investigated analytically and with CFD. Recommendations are given for targets that turbine designers should work toward in reducing the adverse effects of the midpassage gap. A method to estimate the effect of gap flow is presented, so that for any machine the significance of the gap may be assessed.

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