Numerical research on the film-cooling gas turbine blade with the conjugate heat transfer method

2015 ◽  
Vol 19 (sup6) ◽  
pp. S6-175-S6-180
Author(s):  
H. O. Sun ◽  
S. Bu ◽  
Y. G. Luan ◽  
T. Sun ◽  
X. M. Pei
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Gas Turbine Blade ◽  
Numerical Research ◽  
Transfer Method

A new conjugate heat transfer method to analyse a 3D steam cooled gas turbine blade with temperature-dependent material properties

2011 ◽  
Vol 226 (5) ◽  
pp. 1309-1320 ◽  
Author(s):  
M M Jafari ◽  
G Atefi ◽  
J Khalesi ◽  
A Soleymani
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Power Plants ◽  
Conjugate Heat Transfer ◽  
Three Dimensional ◽  
Computer Code ◽  
Gas Turbine Blade ◽  
Temperature Dependent ◽  
Transfer Method

The erosion of the hot regions in a gas turbine is one of the most important challenges encountered by the power plants. Though several numerical simulations of the problem have been reported so far, little is known to give accurate results. In this article, the thermoelastic behaviour of a gas turbine blade with internal steam-cooled channels positioned within a three-dimensional cascade configuration has been studied. A computer code based on the conjugate heat transfer method using the simultaneous solution of Navier–Stokes and heat transfer equations has been developed. From this study, the temperature distribution along with the stress values at high temperatures has been obtained. The blade parameters such as E, α, and K were considered to be a function of the temperature. In the previous works, usually only one or two of these parameters was considered as temperature dependent and the others constant. In this article, all the blade parameters, though making the equations highly non-linear, were considered as a function of temperature. The results have been compared with the available experimental data and a good agreement is observed. According to these findings, taking the temperature dependency of materials into account increases the estimations accuracy and brings the results closer to the reality.

Short-Duration Measurements and Numerical Simulation of Heat Transfer Along the Suction Side of a Film-Cooled Gas Turbine Blade

1985 ◽  
Vol 107 (4) ◽  
pp. 991-997 ◽  
Author(s):  
C. Camci ◽  
T. Arts
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Suction Side ◽  
Gas Turbine Blade ◽  
Transfer Coefficients ◽  
Isentropic Compression ◽  
Curvature Effects ◽  
Convex Wall

This paper deals with an experimental investigation of heat transfer across the suction side of a high-pressure, film-cooled gas turbine blade and with an attempt to numerically predict this quantity both with and without film cooling. The measurements were performed in the VKI isentropic compression tube facility under well-simulated gas turbine conditions. Data measured in a stationary frame, with and without film cooling, are presented. The predictions of convective heat transfer, including streamwise curvature effects, are compared with the measurements. A new approach to determine the augmented mixing lengths near the ejection holes on a highly convex wall is discussed and numerical results agree well with experimentally determined heat transfer coefficients in the presence of film cooling.

Convective Heat Transfer Through Film Cooling Holes of a Gas Turbine Blade Leading Edge

2005 ◽  
Author(s):  
Elon J. Terrell ◽  
Brian D. Mouzon ◽  
David G. Bogard
Keyword(s):  
Heat Transfer ◽  
Convective Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Convective Heat ◽  
Film Cooling ◽  
Leading Edge ◽  
Gas Turbine Blade ◽  
Film Cooling Holes ◽  
Blade Leading Edge

Studies of film cooling performance for a turbine airfoil predominately focus on the reduction of heat transfer to the external surface of the airfoil. However, convective cooling of the airfoil due to coolant flow through the film cooling holes is potentially a major contributor to the overall cooling of the airfoil. This study used experimental and computational methods to examine the convective heat transfer to the coolant as it traveled through the film cooling holes of a gas turbine blade leading edge. Experimental measurements were conducted on a model gas turbine blade leading edge composed of alumina ceramic which approximately matched the Biot number of an engine airfoil leading edge. The temperature rise in the coolant from the entrance to the exit of the film cooling holes was measured using a series of internal thermocouples and an external traversing thermocouple probe. A CFD simulation of the model of the leading edge was also done in order to facilitate the processing of the experimental data and provide a comparison for the experimental coolant hole heat transfer. Without impingement cooling, the coolant hole heat transfer was found to account for 50 to 80 percent of the airfoil internal cooling, i.e. the dominating cooling mechanism.

scholarly journals Detailed Film Cooling Measurements over a Gas Turbine Blade Using a Transient Liquid Crystal Image Technique

2001 ◽  
Vol 7 (6) ◽  
pp. 415-424 ◽  
Author(s):  
Hui Du ◽  
Srinath V. Ekkad ◽  
Je-Chin Han ◽  
C. Pang Lee
Keyword(s):  
Heat Transfer ◽  
Liquid Crystal ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbine Blade ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio ◽  
Image Technique

Detailed heat transfer coefficient and film effectiveness distributions over a gas turbine blade with film cooling are obtained using a transient liquid crystal image technique. The test blade has three rows of film holes on the leading edge and two rows each on the pressure and suction surfaces. A transient liquid crystal technique maps the entire blade midspan region, and helps provide detailed measurements, particularly near the film hole. Tests were performed on a five-blade linear cascade in a low-speed wind tunnel. The mainstream Reynolds number based on cascade exit velocity is5.3×105. Two different coolants (air andCo2) were used to simulate coolant density effect. Coolant blowing ratio was varied between 0.8 and 1.2 for air injection and 0.4–1.2 forCo2injection. Results show that film injection promotes earlier laminar-turbulent boundary layer transition on the suction surface and also enhances local heat transfer coefficients (up to 80%) downstream of injection. An increase in coolant blowing ratio produces higher heat transfer coefficients for both coolants. This effect is stronger immediately downstream of injection holes. Film effectiveness is highest at a blowing ratio of 0.8 for air injection and at a blowing ratio of 1.2 forCo2injection. Such detailed results will help provide insight into the film cooling phenomena on a gas turbine blade.

scholarly journals Experimental and Numerical Investigation of Effect of Blowing Ratio on Film Cooling Effectiveness and Heat Transfer Coefficient Over a Gas Turbine Blade Leading Edge Film Cooling Configurations

2013 ◽  
Author(s):  
Yepuri Giridhara Babu ◽  
Gururaj Lalgi ◽  
Ashok Babu Talanki Puttarangasetty ◽  
Jesuraj Felix ◽  
Sreenivas Rao V. Kenkere ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbine Blade ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio

Film cooling is one of the cooling techniques to cool the hot section components of a gas turbine engines. The gas turbine blade leading edges are the vital parts in the turbines as they are directly hit by the hot gases, hence the optimized cooling of gas turbine blade surfaces is essential. This study aims at investigating the film cooling effectiveness and heat transfer coefficient experimentally and numerically for the three different gas turbine blade leading edge models each having the one row of film cooling holes at 15, 30 and 45 degrees hole orientation angle respectively from stagnation line. Each row has the five holes with the hole diameter of 3mm, pitch of 20mm and has the hole inclination angle of 20deg. in spanwise direction. Experiments are carried out using the subsonic cascade tunnel facility of National Aerospace Laboratories, Bangalore at a nominal flow Reynolds number of 1,00,000 based on the leading edge diameter, varying the blowing ratios of 1.2, 1.50, 1.75 and 2.0. In addition, an attempt has been made for the film cooling effectiveness using CFD simulation, using k-€ realizable turbulence model to solve the flow field. Among the considered 15, 30 and 45 deg. models, both the cooling effectiveness and heat transfer coefficient shown the increase with the increase in hole orientation angle from stagnation line. The film cooling effectiveness increases with the increase in blowing ratio upto 1.5 for the 15 and 30 deg. models, whereas on the 45 deg. model the increase in effectiveness shown upto the blowing ratio of 1.75. The heat transfer coefficient values showed the increase with the increase in blowing ratio for all the considered three models. The CFD results in the form of temperature, velocity contours and film cooling effectiveness values have shown the meaningful results with the experimental values.

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