scholarly journals Comparison of Blade Cooling Performance Using Alternative Fluids

2002 ◽  
Author(s):  
Carlo Carcasci ◽  
Stefano Zecchi ◽  
Gianpaolo Oteri
Keyword(s):  
Heat Transfer ◽  
Carbon Dioxide ◽  
Gas Turbine ◽  
Film Cooling ◽  
Cooling System ◽  
Internal Heat ◽  
Transfer Coefficients ◽  
Cooling Systems ◽  
Heat Transfer Coefficients ◽  
Blade Cooling

CO2 emissions reduction has become an important topic, especially after Kyoto protocol. There are several ways to reduce the overall amount of CO2 discharged into the atmosphere, for example using alternative fluids such as steam or CO2. It is therefore interesting to analyze the consequences of their usage on overall performances of gas turbine and blade cooling systems. The presence of steam can be associated with combined or STIG cycle, whereas pure carbon dioxide or air-carbon dioxide mixtures are present in innovative cycles, where the exhaust gas is recirculated partially or even totally. In this paper we will analyze a commercial gas turbine, comparing different fluids used as working and cooling fluids. The different nature of the fluids involved determines different external heat transfer coefficients (external blade surface), different internal heat transfer coefficients (cooling cavities) and affects film cooling effectiveness, resulting in a change of the blade temperature distribution. Results show that the presence of steam and CO2 could determine a non negligible effect on blade temperature. This means that cooling systems need a deep investigation. A redesign of the cooling system could be required. In particular, results show that steam is well suited for internal cooling, whereas CO2 is better used in film cooling systems.

Heat Transfer Coefficients and Film Cooling Effectiveness on the Squealer Tip of a Gas Turbine Blade

2003 ◽  
Vol 125 (4) ◽  
pp. 648-657 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Rotor Blade ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio

Experimental investigations were performed to measure the detailed heat transfer coefficients and film cooling effectiveness on the squealer tip of a gas turbine blade in a five-bladed linear cascade. The blade was a two-dimensional model of a first stage gas turbine rotor blade with a profile of the GE-E3 aircraft gas turbine engine rotor blade. The test blade had a squealer (recessed) tip with a 4.22% recess. The blade model was equipped with a single row of film cooling holes on the pressure side near the tip region and the tip surface along the camber line. Hue detection based transient liquid crystals technique was used to measure heat transfer coefficients and film cooling effectiveness. All measurements were done for the three tip gap clearances of 1.0%, 1.5%, and 2.5% of blade span at the two blowing ratios of 1.0 and 2.0. The Reynolds number based on cascade exit velocity and axial chord length was 1.1×106 and the total turning angle of the blade was 97.9 deg. The overall pressure ratio was 1.2 and the inlet and exit Mach numbers were 0.25 and 0.59, respectively. The turbulence intensity level at the cascade inlet was 9.7%. Results showed that the overall heat transfer coefficients increased with increasing tip gap clearance, but decreased with increasing blowing ratio. However, the overall film cooling effectiveness increased with increasing blowing ratio. Results also showed that the overall film cooling effectiveness increased but heat transfer coefficients decreased for the squealer tip when compared to the plane tip at the same tip gap clearance and blowing ratio conditions.

Heat Transfer in a Vane Trailing Edge Passage With Conical Pins and Pin-Turbulator Integrated Configurations

2014 ◽  
Author(s):  
J. Kruekels ◽  
S. Naik ◽  
A. Lerch ◽  
A. Sedlov
Keyword(s):  
Heat Transfer ◽  
Pressure Drop ◽  
Gas Turbine ◽  
Internal Heat ◽  
Transfer Coefficients ◽  
Trailing Edge ◽  
Heat Transfer Coefficients ◽  
Trailing Edges ◽  
Converging Channel ◽  
Internal Heat Transfer

The trailing edge sections of gas turbine vanes and blades are generally subjected to extremely high heat loads due to the combined effects of high external accelerating Mach numbers and gas temperatures. In order to maintain the metal temperatures of these trailing edges to a level, which fulfills the mechanical integrity of the parts, highly efficient cooling of the trailing edges is required without increasing the coolant consumption, as the latter has a detrimental effect on the overall gas turbine performance. In this paper the characteristics of the heat transfer and pressure drop of two novel integrated pin bank configurations were investigated. These include a pin bank with conical pins and a pin bank consisting of cylindrical pins and intersecting broken turbulators. As baseline case, a pin bank with cylindrical pins was studied as well. All investigations were done in a converging channel in order to be consistent with the real part. The heat transfer and pressure drop of all the pin banks were investigated initially with the use of numerical predictions and subsequently in a scaled experimental wind tunnel. The experimental study was conducted for a range of operational Reynolds numbers. The TLC (thermochromic liquid crystal) method was used to measure the detailed heat transfer coefficients in scaled Perspex models representing the various pin bank configurations. Pressure taps were located at several positions within the test sections. Both local and average heat transfer coefficients and pressure loss coefficients were determined. The measured and predicted results showed that the local internal heat transfer coefficient increases in the flow direction. This was due to the flow acceleration in the converging channel. Furthermore, both the broken ribs and the conical pin banks resulted in higher heat transfer coefficients compared with the baseline cylindrical pins. The conical pins produced the highest average internal heat transfer coefficients in contrast to the pins with the broken ribs, though this was also associated with a higher pressure drop.

Transient Heat Transfer Experiments in Complex Passages

2005 ◽  
Author(s):  
Rico Poser ◽  
Jens von Wolfersdorf ◽  
Klaus Semmler
Keyword(s):  
Heat Transfer ◽  
Internal Heat ◽  
Direct Consequence ◽  
Data Evaluation ◽  
Transient Heat Transfer ◽  
Inlet Temperature ◽  
Transfer Coefficients ◽  
Turbine Blade Cooling ◽  
Heat Transfer Coefficients ◽  
Blade Cooling

Transient heat transfer experiments were performed in a model of a multi-pass gas turbine blade cooling circuit. The inner surface of the Plexiglas model was coated with thermochromic liquid crystals in order to determine the internal heat transfer coefficients. A change in inlet temperature is applied using a pre-cooled heat exchanger. As for simple geometries the analytical solution of Fourier’s equation can often be directly used for data evaluation, one ought to pay attention to complex passages. The reason has to be seen that the flow in complex passages has to be characterized by local and time dependent fluid temperatures. As a direct consequence data evaluation might be limited to small evaluation areas especially far downstream. Otherwise the uncertainties in the heat transfer results will increase substantially. In the present study the sensitivity of the transient method for complex passages has been analyzed theoretically and applied experimentally.

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.

An Experimental and Numerical Investigation of Near Cooling Hole Heat Fluxes on a Film-Cooled Turbine Blade

1989 ◽  
Vol 111 (1) ◽  
pp. 63-70 ◽  
Author(s):  
C. Camci
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Internal Heat ◽  
External Heat ◽  
Heat Fluxes ◽  
Temperature Gradients ◽  
Transfer Coefficients ◽  
External Heat Transfer ◽  
Heat Transfer Coefficients ◽  
Cooling Hole

Discrete hole film cooling on highly curved surfaces of a gas turbine blade produces very significant wall temperature gradients and wall heat flux variations near downstream and upstream of rows of circular cooling holes. In this study a set of well-defined external heat transfer coefficient distributions in the presence of discrete hole film cooling is presented. Heat transfer coefficients are measured on the suction side of an HP rotor blade profile in a short-duration facility under well-simulated gas turbine flow conditions. The main emphasis of the study is to evaluate the internal heat flux distributions in a detailed way near the cooling holes by using a computational technique. The method uses the measured external heat transfer coefficients as boundary conditions in addition to available internal heat transfer correlations for the internal passages. The study shows the details of the near hole temperature gradients and heat fluxes. The convective heat transfer inside the circular film cooling holes is shown to be very significant even with their relatively small diameter and lengths compared to the chord length. The study also indicates a nonnegligible wall temperature reduction at near upstream of discrete cooling holes. This is explained with the elliptic nature of the internal conduction field of the blade and relatively low coolant temperature levels at the exit of a film cooling hole compared to the mean blade temperature.

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.

Heat Transfer Coefficient and Film-Cooling Effectiveness on the Squealer Tip of a Gas Turbine Blade

2002 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Rotor Blade ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio

Experimental investigations were performed to measure the detailed heat transfer coefficients and film-cooling effectiveness on the squealer tip of a gas turbine blade in a five-bladed linear cascade. The blade was a 2-dimensional model of a first stage gas turbine rotor blade with a profile of the GE-E3 aircraft gas turbine engine rotor blade. The test blade had a squealer (recessed) tip with a 4.22% recess. The blade model was equipped with a single row of film-cooling holes on the pressure-side near the tip region and the tip surface along the camber line. A hue detection based transient liquid crystal technique was used to measure heat transfer coefficients and film-cooling effectiveness. All measurements were done for the tip gap clearances of 1.0%,1.5%, and 2.5% of blade span at the two blowing ratios of 1.0 and 2.0. The Reynolds number based on cascade exit velocity and axial chord length was 1.1 × 106 and the overall pressure ratio was 1.32. The turbulence intensity level at the cascade inlet was 9.7%. Results showed that the overall heat transfer coefficients increased with increasing tip gap clearance, but decreased with increasing blowing ratio. However, the overall film-cooling effectiveness increased with increasing blowing ratio. Results also showed that the overall film-cooling effectiveness increased but heat transfer coefficients decreased for the squealer tip when compared to the plane tip at the same tip gap clearance and blowing ratio conditions.

scholarly journals Gas Turbine Heat Transfer With Alternate Cooling Fluids

1988 ◽  
Author(s):  
N. Abuaf ◽  
A. Cohn
Keyword(s):  
Heat Transfer ◽  
Carbon Dioxide ◽  
Film Cooling ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
Fully Developed Turbulent Flow ◽  
Gas Turbine Heat Transfer ◽  
Property Changes

Steam or gases other than air have been proposed as substitutes for cooling industrial turbine components positioned along the hot gas path. To check the effect of coolant gas physical properties on the internal passage heat transfer coefficients, tests were conducted with air, helium, carbon dioxide, and refrigerant-12 in a serpentine passage with turbulence promoters. For the fully developed turbulent flow regions, the correlation proposed by Han et al. (1985) properly accounts for the cooling gas property changes. With respect to the external heat transfer aspects with film cooling, tests were run in a wind tunnel with a two-dimensional slit injection geometry. The film cooling effectiveness results measured with the same gases injected as a secondary gas into the main flow of air show that the correlations proposed by Kutateladze and Leont’ev (1963) and Librizzi and Cresci (1964) agree with the present data for the air/air, air/helium, and air/carbon dioxide combinations. The air/refrigerant-12 results are lower than all the correlations proposed in the literature.

Heat Transfer and Flow Studies of Different Cooling Configurations for Gas Turbine Rotor Blade

2014 ◽  
Author(s):  
Batchu Suresh ◽  
Ainapur Brijesh ◽  
V. Kesavan ◽  
S. Kishore Kumar
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Metal Temperature ◽  
Coolant Temperature ◽  
Fe Model ◽  
Transfer Coefficients ◽  
Flow Network ◽  
Blade Cooling ◽  
Film Cooling Holes

Military gas turbine engine operates at turbine entry temperatures (TET) of the order of 2000K. Increase in TET improves thermal efficiency and power output. The gas temperature is far above the allowable metal temperature of turbine components. Hence, there is a need to cool the components such as blades and vanes for safe operation. The blades are cooled by combination of internal convective cooling and external film-cooling. Rib tabulators are widely used in blade cooling passages to enhance heat transfer. In the present study, different rib tabulator configurations have been studied. 1D flow network model of blade cooling passages have been modeled using Flowmaster software. Flowmaster software estimates pressure losses, rotational effects and heat transfer of the coolant flow in the blade passages. Cooling passages are modeled as ducts while film cooling holes, impingement holes, tip holes and ejection holes are modeled as orifices. Experimentally measured heat transfer and pressure loss correlations are used in the analysis. The coolant pressure at inlet and sink pressure at exit of film cooling holes are given as input. The heat load coming on to the blade is also given as input for predicting the coolant temperature rise and blade metal temperature. The thermal analysis is carried out with different shaped rib turbulators such as V and W ribs with broken and continuous pattern. It is observed that thermal performance factor for a broken V rib configuration is better than other configurations. The metal temperature for broken V ribbed configuration is 25°C less compared other configurations. The effect of rotation on the blade temperature is also studied. The convective bulk temperatures and convective heat transfer coefficients obtained from 1D flow network are applied on 2D Finite Element (FE) model to obtain nodal temperature distribution.

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