Comparative Study of Barrier Coatings on Turbine Blade Cooling

2021 ◽  
Author(s):  
Kishan A. Singh ◽  
Zozimus D. Labana
Keyword(s):  
Heat Transfer ◽  
High Temperature ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Thermal Efficiency ◽  
Considerable Effect ◽  
Inlet Temperature ◽  
Cooling Effectiveness ◽  
Continuous Growth ◽  
Barrier Coatings

Abstract There is a continuous growth in various sectors of engineering and technology resulting in high demand for modern and innovative steps into providing solutions to overcome all barriers. With the increase in demand of modern age technology growing rapidly, the requirement of the advanced system has to keep up with this demand. The power required to drive such technology has one of its sources from the gas turbine propulsion system, which is capable to provide thrust and power to major components. In modern gas turbine applications, high power and thermal efficiency are of essential requirement. The two parameters that play a vital role in increasing the thermal efficiency of the gas turbine are its compression ratio and high turbine inlet temperature. The advanced gas turbine has inlet temperature, which exceeds the material thermal limitation of the blade. This high temperature has an impact on the performance and life of the blade employed for energy extraction in the turbines. The current study analyzes the methods dealing with increasing the cooling effectiveness of the turbine blade, which are working under very high-temperature hot gases that exit from the combustion chamber. This gas expands into the turbine region by which power is extracted. These high-temperature gases can have a considerable effect on the stresses developed that can lead to failure under the cyclic loading of these hot gases. Cooling effectiveness can increase the system working temperature from 800K up to 1000K inlet. The current research compares the means of reducing the heat transfer and improve cooling effectiveness with the help of the latest improved material coatings such as thermal barrier coatings (TBCs) and environmental barrier coatings (EBCs). The study focuses on the effect of these TBCs and EBCs employed on the surface of the blade. Analysis of results obtained from this conjugate heat transfer (CHT) study has shown good agreement with the experimental data. The comparison revealed the use of the SST k-ω model which was efficient and predicted similar trends as that of the experimental for pressure were as with 3% of deviation for temperature.

High-Temperature Potential of Uncooled Radial Turbines

1978 ◽  
Vol 100 (2) ◽  
pp. 294-302 ◽  
Author(s):  
D. J. Arnold ◽  
O. E. Balje
Keyword(s):  
Heat Transfer ◽  
High Temperature ◽  
Gas Turbine ◽  
Thermal Efficiency ◽  
Experimental Investigations ◽  
Inlet Temperature ◽  
High Reaction ◽  
Temperature Potential ◽  
Axial Turbines ◽  
Radial Turbines

Radial turbines are used predominantly for turbo-charges where the geometry is frequently compromised to favor low fabrications costs. Theoretical as well as experimental investigations have shown that the efficiency potential of radial turbines is as high as the efficiency potential of high reaction axial turbines. Structural and heat transfer studies on radial turbines show that the highest stresses in “deep scalloped” radial rotors occur at locations where the metal temperature is considerably lower than at rotor inlet. Thus the maximum allowable gas inlet temperature for radial turbines is several hundred degrees higher than for high-reaction axial turbines. This difference tends to increase with increasing expansion ratios, at least up to expansion ratios of 10:1. Since the thermal efficiency of typical gas turbine cycles increases with increasing gas temperatures and increasing expansion ratios, it results that the application of uncooled radial turbines will yield cycle efficiencies which are not obtainable with uncooled axial turbines.

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.

Improvement of Film Cooling Effectiveness in the Gas Turbine Endwall by Use of Optimization Framework

2019 ◽  
Author(s):  
Daisuke Hata ◽  
Kazuto Kakio ◽  
Yutaka Kawata ◽  
Masahiro Miyabe
Keyword(s):  
Pressure Gradient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Combined Cycle ◽  
Inlet Temperature ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness

Abstract Recently, the number of gas turbine combined cycle plants is rapidly increasing in substitution of nuclear power plants. The turbine inlet temperature (TIT) is constantly being increased in order to achieve higher effectiveness. Therefore, the improvement of the cooling technology for high temperature gas turbine blades is one of the most important issue to be solved. In a gas turbine, the main flow impinging at the leading edge of the turbine blade generates a so called horseshoe vortex by the interaction of its boundary layer and generated pressure gradient at the leading edge. The pressure surface leg of this horseshoe vortex crosses the passage and reaches the blade suction surface, driven by the pressure gradient existing between two consecutive blades. In addition, this pressure gradient generates a cross-flow along the endwall. This all results into a very complex flow field in proximity of the endwall. For this reason, burnouts tend to occur at a specific position in the vicinity of the leading edge. In this research, a methodology to cool the endwall of the turbine blade by means of film cooling jets from the blade surface and the endwall is proposed. The cooling performance is investigated using the transient thermography method. CFD analysis is also conducted to investigate the phenomena occurring at the endwall and calculate the film cooling effectiveness.

Research on Heat-Transfer and Cooling Performance of Gas Turbine Endwall

2017 ◽  
Author(s):  
Kazuto Kakio ◽  
Y. Kawata
Keyword(s):  
Heat Transfer ◽  
Pressure Gradient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Power Plants ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Combined Cycle ◽  
Inlet Temperature ◽  
Cooling Performance

Recently, the number of gas turbine combined cycle plants is rapidly increasing in substitution of nuclear power plants. The turbine inlet temperature (TIT) is being constantly increased in order to achieve higher efficiency. Therefore, the improvement of the cooling technology for high temperature gas turbine blades is one of the most important issue to be solved. In a gas turbine, the main flow impinging at the leading edge of the turbine blade generates a so called horseshoe vortex by the interaction of its boundary layer and generated pressure gradient at the leading edge. The pressure surface leg of this horseshoe vortex crosses the passage and reaches the blade suction surface, driven by the pressure gradient existing between two consecutive blades. In addition, this pressure gradient generates a crossflow along the endwall. This all results into a very complex flow field in proximity of the endwall. For this reason, burnouts tend to occur at a specific position in the vicinity of the leading edge. In this research, a methodology to cool the endwall of the turbine blade by means of film cooling jets from the blade surface is proposed. The cooling performance and heat transfer coefficient distribution is investigated using the transient thermography method. CFD analysis is also conducted to know the phenomena occurring at the end wall and calculate the heat transfer distribution.

Conceptual Design and Cooling Blade Development of 1700°C Class High-Temperature Gas Turbine

2005 ◽  
Vol 127 (2) ◽  
pp. 358-368 ◽  
Author(s):  
Shoko Ito ◽  
Hiroshi Saeki ◽  
Asako Inomata ◽  
Fumio Ootomo ◽  
Katsuya Yamashita ◽  
...  
Keyword(s):  
Heat Transfer ◽  
High Temperature ◽  
Gas Turbine ◽  
Conceptual Design ◽  
Thermal Efficiency ◽  
Internal Cooling ◽  
Closed Cycle ◽  
Steam Consumption ◽  
Flow Phenomena ◽  
High Temperature Gas

In this paper we describe the conceptual design and cooling blade development of a 1700°C-class high-temperature gas turbine in the ACRO-GT-2000 (Advanced Carbon Dioxide Recovery System of Closed-Cycle Gas Turbine Aiming 2000 K) project. In the ACRO-GT closed cycle power plant system, the thermal efficiency aimed at is more than 60% of the higher heating value of fuel (HHV). Because of the high thermal efficiency requirement, the 1700°C-class high-temperature gas turbine must be designed with the minimum amount of cooling and seal steam consumption. The hybrid cooling scheme, which is a combination of closed loop internal cooling and film ejection cooling, was chosen from among several cooling schemes. The elemental experiments and numerical studies, such as those on blade surface heat transfer, internal cooling channel heat transfer, and pressure loss and rotor coolant passage distribution flow phenomena, were conducted and the results were applied to the conceptual design advancement. As a result, the cooling steam consumption in the first stage nozzle and blade was reduced by about 40% compared with the previous design that was performed in the WE-NET (World Energy Network) Phase-I.

Development and Evaluation of Ceramic Components for Gas Turbine

2002 ◽  
Author(s):  
Takero Fukudome ◽  
Sazo Tsuruzono ◽  
Wataru Karasawa ◽  
Yoshihiro Ichikawa
Keyword(s):  
High Temperature ◽  
Silicon Nitride ◽  
Gas Turbine ◽  
Thermal Efficiency ◽  
Technology Development ◽  
Inlet Temperature ◽  
Combustion Gas ◽  
Development Organization ◽  
Nitride Material ◽  
Ceramic Components

An 8000 kW class Hybrid Gas Turbine (HGT) project, administered by the New Energy and Industrial Technology Development Organization (NEDO), has been ongoing since July of 1999 in Japan. Targets of this project are improvement in thermal efficiency and output power by using ceramic components, and early commercialization of the gas turbine system. The ceramic components are used for stationary parts subjected to high temperature, such as combustor liners, transition ducts, and first stage turbine nozzles. Development of the gas turbine is conducted by Kawasaki Heavy Industries, Ltd. (KHI), to achieve the Turbine Inlet Temperature (TIT) of 1250°C, thermal efficiency of 34%, NOx emission less than standard regulation values, and 4,000 h engine durability. Kyocera is in charge of the development and evaluation of the ceramic components. Recently, recession of the Si based ceramic materials under the combustion gas is the focus of attention to improve the reliability of ceramic components for gas turbine. For the HGT project, the silicon nitride material (SN282 : silicon nitride material produced by Kyocera Corporation) is used for the components subjected to high temperature. The SN282 was evaluated under the combustion gas, and clear recession was observed. Our technology of the Environmental Barrier Coating (EBC) is under development to obtain reliable heat resistive SN282 components, against the recession by combustion gas. Reliability of the SN282 with EBC has been evaluated by exposure and hydrothermal corrosion test. Ceramic components made of SN282 with EBC will be also evaluated by a proof engine test of 4,000 h, which starts in the spring of 2002.

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.

Thermodynamic and Energy Study of a Regenerator in Gas Turbine Cycle and Optimization of Performances

2016 ◽  
Vol 5 (2) ◽  
pp. 25-44
Author(s):  
Saria Abed ◽  
Taher Khir ◽  
Ammar Ben Brahim
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Thermal Efficiency ◽  
Parametric Study ◽  
Pressure Ratio ◽  
Operating Conditions ◽  
Inlet Temperature ◽  
Compressor Pressure Ratio ◽  
Compressor Inlet ◽  
Gas Turbine Cycle

In this paper, thermodynamic study of simple and regenerative gas turbine cycles is exhibited. Firstly, thermodynamic models for both cycles are defined; thermal efficiencies of both cycles are determined, the overall heat transfer coefficient through the heat exchanger is calculated in order to determinate its performances and parametric study is carried out to investigate the effects of compressor inlet temperature, turbine inlet temperature and compressor pressure ratio on the parameters that measure cycles' performance. Subsequently, numerical optimization is established through EES software to determinate operating conditions. The results of parametric study have shown a significant impact of operating parameters on the performance of the cycle. According to this study, the regeneration technique improves the thermal efficiency by 10%. The studied regenerator has an important effectiveness (˜ 82%) which improves the heat transfer exchange; also a high compressor pressure ratio and an important combustion temperature can increase thermal efficiency.

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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