A New Method to Calculate the Coolant Requirements of a High-Temperature Gas Turbine Blade

2005 ◽  
Vol 127 (1) ◽  
pp. 191-199 ◽  
Author(s):  
Leonardo Torbidoni ◽  
J. H. Horlock
Keyword(s):  
High Temperature ◽  
Gas Turbine ◽  
Film Cooling ◽  
Guide Vane ◽  
Coolant Flow ◽  
Computational Method ◽  
Inlet Temperature ◽  
Film Cooling Effectiveness ◽  
Cooling Flow ◽  
High Temperature Gas

Earlier papers by the first author have described a computational method of estimating the cooling flow requirements of blade rows in a high-temperature gas turbine, for convective cooling alone and for convective plus film cooling. This method of analysis and computation, when applied to the whole blade chord was compared to a well-known semi-empirical method. In the current paper, a more sophisticated method is developed from the earlier work and is used to calculate the cooling flow required for a nozzle guide vane (the first blade row) of a high-temperature gas turbine, with given inlet gas temperature and coolant inlet temperature. Now the heat flux through an elementary cross-sectional area of the blade, at given spanwise (y) and chordwise (s) locations, is considered, with a guessed value of the elementary coolant flow [as a fraction dΨs of the external gas flow]. At the given s, integration along the blade length gives the blade metal temperatures at the outer and inner walls, Tbgy and Tbcly. If the value of Tbg at the blade tip y=H is assumed to be limited by material considerations to Tbg,max then the elementary coolant flow rate may be obtained by iteration. Summation along the chord then gives the total coolant flow, for the whole blade. Results using the method are then compared to a simpler calculation applied to the whole blade, which assumes chordwise constant temperatures and constant selected values of cooling efficiency and film-cooling effectiveness.

A New Method to Calculate the Coolant Requirements of a High Temperature Gas Turbine Blade

2004 ◽  
Author(s):  
Leonardo Torbidoni ◽  
J. H. Horlock
Keyword(s):  
High Temperature ◽  
Gas Turbine ◽  
Film Cooling ◽  
Guide Vane ◽  
Coolant Flow ◽  
Computational Method ◽  
Inlet Temperature ◽  
Film Cooling Effectiveness ◽  
Cooling Flow ◽  
High Temperature Gas

Earlier papers by the first author have described a computational method of estimating the cooling flow requirements of blade rows in a high temperature gas turbine, for convective cooling alone and for convective plus film cooling; this method of analysis and computation, when applied to the whole blade chord was compared with a well known semi-empirical method. In the current paper, a more sophisticated method is developed from the earlier work and is used to calculate the cooling flow required for a nozzle guide vane [the first blade row] of a high temperature gas turbine, with given inlet gas temperature and coolant inlet temperature. Now the heat flux through an elementary cross-sectional area of the blade, at given spanwise [y] and chordwise [s] locations, is considered, with a guessed value of the elementary coolant flow [as a fraction dΨ(s) of the external gas flow]. At the given s, integration along the blade length gives the blade metal temperatures at the outer and inner walls, Tbg[y] and Tbcl[y]. If the value of Tbg at the blade tip [y = H] is assumed to be limited by material considerations to Tbg,max then the elementary coolant flow rate may be obtained by iteration. Summation along the chord then gives the total coolant flow, for the whole blade. Results using the method are then compared with a simpler calculation applied to the whole blade, which assumes chordwise constant temperatures and constant selected values of cooling efficiency and film cooling effectiveness.

Film Cooling Performance of Tripod Holes on the Endwall Upstream of a First Stage Nozzle Guide Vane

2015 ◽  
Author(s):  
Bharath Viswanath Ravi ◽  
Samruddhi Deshpande ◽  
Sridharan Ramesh ◽  
Prethive Dhilip Dhilipkumar ◽  
Srinath Ekkad
Keyword(s):  
Gas Turbine ◽  
Mass Flow ◽  
Film Cooling ◽  
Guide Vane ◽  
Turbine Engines ◽  
Inlet Temperature ◽  
Gas Turbine Engines ◽  
Film Cooling Effectiveness ◽  
The Difference ◽  
Nozzle Guide

In view of the growing energy demand, there is an increasing need to augment the thermal efficiency of gas turbine engines. The thermal efficiency and power output of gas turbine engines increase with increasing overall pressure ratio which in turn leads to an increase in turbine inlet temperature. The maximum permissible turbine inlet temperature is limited by the material strength of the components of the gas turbine engines. In this regard, it is important to ensure that the endwalls of the first stage nozzle guide vane, which is one of the critical regions, are adequately cooled. The cooling of the endwall is of particular interest because the leading edge region along the endwall of the stator vane experiences high heat transfer rates resulting from formation of horseshoe vortices. In this paper, the performance of upstream purge slot has been compared against discrete film cooling holes. Three different cooling configurations — slot, cylindrical holes and tripod holes have been investigated by comparing the adiabatic film cooling effectiveness. Furthermore, the effect of coolant to mainstream mass flow ratio on the effectiveness of the different cooling schemes has also been studied. The steady-state experiments were conducted in a low speed, linear cascade wind tunnel. Spatially resolved temperature data was captured using infrared thermography technique to compute adiabatic film cooling effectiveness. Amongst the configurations studied, slot ejection offered the best cooling performance at all mass flow ratios. The performance of tripod ejection was comparable to slot ejection at mass flow ratios between 0.5 and 1.5, with the difference in laterally averaged effectiveness being ∼5%. However, at the highest mass flow ratio (MFR=2.5), the difference increased to ∼20%. Low effectiveness values were observed downstream of cylindrical ejection which could be attributed to jet lift-off.

scholarly journals Adiabatic Effectiveness and Heat Transfer Coefficient of Shaped Film Cooling Holes on a Scaled Guide Vane Pressure Side Model

2004 ◽  
Vol 10 (5) ◽  
pp. 345-354 ◽  
Author(s):  
Jan Dittmar ◽  
Achmed Schulz ◽  
Sigmar Wittig
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Guide Vane ◽  
Inlet Temperature ◽  
Test Model ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness

The demand of improved thermal efficiency and high power output of modern gas turbine engines leads to extremely high turbine inlet temperature and pressure ratios. Sophisticated cooling schemes including film cooling are widely used to protect the vanes and blades of the first stages from failure and to achieve high component lifetimes. In film cooling applications, injection from discrete holes is commonly used to generate a coolant film on the blade's surface.In the present experimental study, the film cooling performance in terms of the adiabatic film cooling effectiveness and the heat transfer coefficient of two different injection configurations are investigated. Measurements have been made using a single row of fanshaped holes and a double row of cylindrical holes in staggered arrangement. A scaled test model was designed in order to simulate a realistic distribution of Reynolds number and acceleration parameter along the pressure side surface of an actual turbine guide vane. An infrared thermography measurement system is used to determine highly resolved distribution of the models surface temperature. Anin-situcalibration procedure is applied using single embedded thermocouples inside the measuring plate in order to acquire accurate local temperature data.All holes are inclined 35° with respect to the model's surface and are oriented in a streamwise direction with no compound angle applied. During the measurements, the influence of blowing ratio and mainstream turbulence level on the adiabatic film cooling effectiveness and heat transfer coefficient is investigated for both of the injection configurations.

Numerical Investigation of the Cooling Performance on the Endwall With and Without Fillet

2018 ◽  
Author(s):  
Qingzong Xu ◽  
Qiang Du ◽  
Pei Wang ◽  
Jun Liu ◽  
Guang Liu
Keyword(s):  
Secondary Flow ◽  
Turbulence Model ◽  
Film Cooling ◽  
Computational Method ◽  
Inlet Temperature ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Numerical Predictions ◽  
Turbine Vane

High inlet temperature of turbine vane increases the demand of high film cooling effectiveness. Vane endwall region was extensively cooled due to the high and flat exit temperature distribution of combustor. Leakage flow from the combustor-turbine gap was used to cool the endwall region except for preventing hot gas ingestion. Numerical predictions were conducted to investigate the flow structure and adiabatic film cooling effectiveness of endwall region in a linear cascade with vane-endwall junction fillet. The simulations were completed by solving the three-dimensional Reynolds-Averaged Navier-Stokes(RANS) equations with shear stress transport(SST) k-ω turbulence model, meanwhile, the computational method and turbulence model were validated by comparing computational result with the experiment. Three types of linear fillet with the length-to-height ratio of 0.5, 1 and 2, named fillet A, fillet B and fillet C respectively, were studied. In addition, circular fillet with radius of 2mm was compared with linear fillet B. The interrupted slot, produced by changing the way of junction of combustor and turbine vane endwall, is introduced at X/Cax = −0.2 upstream of the vane leading edge. Results showed that fillet can significantly affect the cooling performance on the endwall due to suppressing the strength of the secondary flow. Fillet C presented the best cooling performance comparing to fillet A and fillet B because a portion of the coolant which climbs to the fillet was barely affected by secondary flow. Results also showed the effect of fillet on the total pressure loss. The result indicated that only fillet A slightly decreases endwall loss.

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.

Numerical Simulation on Gas Turbine Film Cooling of Curved Surface With Backward Injection

2012 ◽  
Author(s):  
Shashank Shetty ◽  
Xianchang Li ◽  
Ganesh Subbuswamy
Keyword(s):  
Numerical Simulation ◽  
Gas Turbine ◽  
Film Cooling ◽  
Gas Turbines ◽  
Turbine Blades ◽  
Operating Conditions ◽  
Strong Mixing ◽  
Inlet Temperature ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss

Due to the unique role of gas turbine engines in power generation and aircraft propulsion, significant effort has been made to improve the gas turbine performance. As a result, the turbine inlet temperature is usually elevated to be higher than the metal melting point. Therefore, effective cooling of gas turbines is a critical task for engines’ efficiency as well as safety and lifetime. Film cooling has been used to cool the turbine blades for many years. The main issues related to film cooling are its poor coverage, aerodynamic loss, and increase of heat transfer coefficient due to strong mixing. To overcome these problems, film cooling with backward injection has been found to produce a more uniform cooling coverage under low pressure and temperature conditions and with simple cylindrical holes. Therefore, the focus of this paper is on the performance of film cooling with backward injection at gas turbine operating conditions. By applying numerical simulation, it is observed that along the centerline on both concave and convex surfaces, the film cooling effectiveness decreases with backward injection. However, cooling along the span is improved, resulting in more uniform cooling.

Numerical Study of Mist Film Cooling in Combustor at Operating Conditions

2009 ◽  
Author(s):  
Srinivasa Rao Para ◽  
Xianchang Li ◽  
Ganesh Subbuswamy
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Numerical Study ◽  
Operating Conditions ◽  
Inlet Temperature ◽  
Gas Temperature ◽  
Cooling Effectiveness ◽  
Wall Functions ◽  
Film Cooling Effectiveness ◽  
Combustor Liner

To improve the gas turbine thermal performance, apart from using a high compression ratio, the turbine inlet temperature must be increased. Therefore, the gas temperature inside the combustion chamber needs to be maintained at a very high level. Hence, cooling of the combustor liner becomes critical. Among all the cooling techniques, film cooling has been successfully applied to cool the combustor liner. In film cooling, coolant air is introduced through discrete holes and forms a thin film between the hot gases and the inner surface of the liner, so that the inner wall can be protected from overheating. The film will be destroyed in the downstream flow because of mixing of hot and cold gases. The present work focuses on numerical study of film cooling under operating conditions, i.e., high temperature and pressure. The effect of coolant injection angles and blowing ratios on film cooling effectiveness is studied. A promising technology, cooling with mist injection, is studied under operating conditions. The effect of droplet size and mist concentration is also analyzed. The results of this study indicate that the film cooling effectiveness can increase ∼11% at gas turbine operating conditions with mist injection of 2% coolant air when droplets of 10μm and a blowing ratio of 1.0 are applied. The cooling performance can be further improved by higher mist concentration. The commercial CFD software, Fluent 6.3.26, is used in this study and the standard k-ε model with enhanced wall functions is adopted as the turbulence model.

Film Cooling Effectiveness in a Gas Turbine Engine: A Review

2014 ◽  
Vol 71 (2) ◽  
Author(s):  
Ehsan Kianpour ◽  
Nor Azwadi Che Sidik ◽  
Iman Golshokouh
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Future Research ◽  
Inlet Temperature ◽  
Turbine Engine ◽  
Film Cooling Effectiveness ◽  
Engine Efficiency ◽  
External Cooling ◽  
Critical Components ◽  
Film Cooling Holes

This study was carried out to extend database knowledge about the function of film cooling holes at the end of combustor and the inlet of turbine. Using the well-known Brayton cycle, raising the turbine inlet temperature is the key to obtain higher engine efficiency in gas turbine engines. However, high temperature of the combustor exit flow causes non-uniformities. These non-uniformities lead to the reduction of expected life of critical components. Therefore, an appropriate cooling technique should be designed to protect these parts. Film cooling is one of the most effective external cooling methods. Various film cooling techniques presented in the literature have been investigated. Moreover, challenges and future directions of film cooling techniques have been reviewed and presented in this paper. The aim of this review is to summarize recent development in research on film cooling techniques and attempt to identify some challenging issues that need to be solved for future research.

Radiative Effectiveness on the Aero- and Thermodynamics in a Highly Thermally Loaded Film Cooling System

2011 ◽  
Author(s):  
Wenping Wang ◽  
Peng Sun ◽  
Jing Ren ◽  
Hongde Jiang
Keyword(s):  
Heat Transfer ◽  
High Temperature ◽  
Gas Turbine ◽  
Film Cooling ◽  
Cooling System ◽  
Numerical Study ◽  
Discrete Ordinates ◽  
Spanwise Direction ◽  
Inlet Temperature ◽  
Gas Channel

With the increasing of the gas turbine inlet temperature, the radiative heat transfer plays a more important role in the total heat transfer. In this paper, a high temperature test rig has been built to research the radiative effect in high temperature film cooling. The test section is made up of a high temperature hot gas channel and a middle temperature coolant air channel which are separated by a flat plate with a row of film cooling holes. The goal is to analyze the effects of radiation and its interaction between conduction and convection in the internal and film cooling which consider the heat transfer in both gas and solid. Meanwhile, the numerical study on the test cases are also carried out by combining conjugate heat transfer with radiative models. The fluid and solid regions were solved simultaneously. The Discrete Ordinates (DO) model and the Weighted Sum of Gray Gases Model (WSGGM) has been used to solve the radiative transfer equation for the radiation modeling. The results show that the temperature of the plate increase greatly when the radiation is taken into account and the temperature gradient through the plate becomes much larger. The temperature distribution has been changed and become smoother in spanwise direction. The results also indicate that the internal emissivity of the inlet has an influence mainly on the whole temperature of the plate, which suggests that the control of inlet emissivity is a good way for prevent over-high temperature on the first stage gas turbine vane.

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