scholarly journals The Effect of Heat Transfer on Turbine Performance

2020 ◽  
Vol 142 (09) ◽  
pp. 56-57
Author(s):  
Lachlan J. Jardine ◽  
Robert J. Miller
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Gas Turbine ◽  
Turbine Blades ◽  
Large Temperature ◽  
Gas Turbine Blades ◽  
Turbine Performance ◽  
Very High

Abstract For over 50 years, high-pressure gas turbine blades have been cooled using air bled from the compressor. This cooling results in very high rates of heat transfer, both within the fluid and within the blade, shown in figure 1. The heat transfer often occurs across large temperature differences and is thus highly irreversible. It is therefore surprising that little is understood about the effect of this heat transfer on turbine performance.

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.

Experimental Study of Heat Transfer in Gas Turbine Blades Using a Transient Inverse Technique

2009 ◽  
Vol 30 (13) ◽  
pp. 1077-1086 ◽  
Author(s):  
Peter Heidrich ◽  
Jens V. Wolfersdorf ◽  
Martin Schnieder
Keyword(s):  
Heat Transfer ◽  
Experimental Study ◽  
Gas Turbine ◽  
Turbine Blades ◽  
Inverse Technique ◽  
Gas Turbine Blades

Laminar and Transitional Boundary Layer Structures in Accelerating Flow With Heat Transfer

1986 ◽  
Vol 108 (1) ◽  
pp. 116-123 ◽  
Author(s):  
K. Rued ◽  
S. Wittig
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Gas Turbine ◽  
Turbine Blades ◽  
Pressure Gradients ◽  
Transfer Coefficients ◽  
Gas Turbine Blades ◽  
Heat Transfer Coefficients ◽  
Layer Structures ◽  
Flow Heat

The accurate prediction of heat transfer coefficients on cooled gas turbine blades requires consideration of various influence parameters. The present study continues previous work with special efforts to determine the separate effects of each of several parameters important in turbine flow. Heat transfer and boundary layer measurements were performed along a cooled flat plate with various freestream turbulence levels (Tu = 1.6−11 percent), pressure gradients (k = 0−6 × 10−6), and cooling intensities (Tw/T∞ = 1.0−0.53). Whereas the majority of previously available results were obtained from adiabatic or only slightly heated surfaces, the present study is directed mainly toward application on highly cooled surfaces as found in gas turbine engines.

Film Cooling of Gas Turbine Blades: A Study of the Effect of Large Temperature Differences on Film Cooling Effectiveness

1977 ◽  
Vol 99 (1) ◽  
pp. 11-20 ◽  
Author(s):  
M. A. Paradis
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Turbine Blades ◽  
Large Temperature ◽  
Constant Property ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Gas Turbine Blades ◽  
Combustion Gases ◽  
Simple Equations

Experiments have been performed on the film cooling of gas turbine blades in order to study the influence of large temperature differences on the effectiveness of film cooling. A two-dimensional flat plate model was tested in a stream of 1000 K combustion gases flowing at between 110 and 170 m/s. The model was cooled on both sides by jets of air coming from flush angled slots. The range of velocity ratios Uc/Ug covered was from 0.3 to 1.7 and the range of blowing rates was between 0.5 and 5. Film cooling effectiveness was measured and boundary layer traverses were performed. It has been found that once radiation and conduction effects are taken into account, the simple equations proposed by previous workers for the constant property case could be used with little error.

Experimental and Numerical Study of Jet Impingement Cooling for Improved Gas Turbine Blade Internal Cooling With In-Line and Staggered Nozzle Arrays

2020 ◽  
Vol 143 (1) ◽  
Author(s):  
Abdel Rahman Salem ◽  
Farah Nazifa Nourin ◽  
Mohammed Abousabae ◽  
Ryoichi S. Amano
Keyword(s):  
Heat Transfer ◽  
Nusselt Number ◽  
Gas Turbine ◽  
Numerical Study ◽  
Turbine Blades ◽  
Jet Impingement ◽  
Target Surface ◽  
Internal Cooling ◽  
Impingement Cooling ◽  
Gas Turbine Blades

Abstract Internal cooling of gas turbine blades is performed with the combination of impingement cooling and serpentine channels. Besides gas turbine blades, the other turbine components such as turbine guide vanes, rotor disks, and combustor wall can be cooled using jet impingement cooling. This study is focused on jet impingement cooling, in order to optimize the coolant flow, and provide the maximum amount of cooling using the minimum amount of coolant. The study compares between different nozzle configurations (in-line and staggered), two different Reynold's numbers (1500 and 2000), and different stand-off distances (Z/D) both experimentally and numerically. The Z/D considered are 3, 5, and 8. In jet impingement cooling, the jet of fluid strikes perpendicular to the target surface to be cooled with high velocity to dissipate the heat. The target surface is heated up by a direct current (DC) power source. The experimental results are obtained by means of thermal image processing of the captured infra-red (IR) thermal images of the target surface. Computational fluid dynamics (CFD) analysis were employed to predict the complex heat transfer and flow phenomena, primarily the line-averaged and area-averaged Nusselt number and the cross-flow effects. In the current investigation, the flow is confined along with the nozzle plate and two parallel surfaces forming a bi-directional channel (bi-directional exit). The results show a comparison between heat transfer enhancement with in-line and staggered nozzle arrays. It is observed that the peaks of the line-averaged Nusselt number (Nu) become less as the stand-off distance (Z/D) increases. It is also observed that the fluctuations in the stagnation heat transfer are caused by the impingement of the primary vortices originating from the jet nozzle exit.

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.

Conjugate Heat Transfer Simulation and Entropy Generation Analysis of Gas Turbine Blades

2021 ◽  
Author(s):  
Yaping Ju ◽  
Yi Feng ◽  
Chuhua Zhang
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Entropy Generation ◽  
Conjugate Heat Transfer ◽  
Turbine Blades ◽  
Measurement Data ◽  
Turbulence Models ◽  
Internal Cooling ◽  
Gas Turbine Blades ◽  
Internally Cooled

Abstract Reynolds averaged Navier-Stokes model-based conjugate heat transfer method is popularly used in simulations and designs of internally cooled gas turbine blades. One of the important factors influencing its prediction accuracy is the choice of turbulence models for different fluid regions because the blade passage flow and internal cooling have considerably different flow features. However, most studies adopted the same turbulence models in passage flow and internal cooling. Another important issue is the comprehensive evaluation of the losses caused by flow and heat transfer for both fluid and solid regions. In this study, a RANS-based CHT solver for subsonic/transonic flows was developed based on OpenFOAM and validated and used to explore suitable RANS turbulence model combinations for internally cooled gas turbine blades. Entropy generation, able to weigh the losses caused by flow friction and heat transfer, was used in the analyses of two internally cooled vanes to reveal the loss mechanisms. Findings indicate that the combination of the k-? SST-?-Re? transition model for passage flow and the standard k-e model for internal cooling agreed best with measurement data. The relative error of vane dimensionless temperature was less than 3%. The variations of entropy generation with different internal cooling inlet velocities and temperatures indicate that reducing entropy generation was contradictory with enhancing heat transfer performance. This study, providing a reliable computing tool and a comprehensive performance parameter, has an important application value for the design of internally cooled gas turbine blades.

Sign in / Sign up

Export Citation Format

Share Document