Impact of a Combustor-Turbine Interface on Turbine Vane Aerodynamics and Film Cooling

2019 ◽  
Author(s):  
A. Lerch ◽  
R. Bauer ◽  
J. Krueckels ◽  
M. Henze
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Gas Turbines ◽  
High Speed ◽  
Swirling Flow ◽  
Numerical Study ◽  
Heavy Duty ◽  
Film Cooling Effectiveness ◽  
Transition Piece ◽  
The Impact

Abstract Optimizing the aero-thermal performance of the combustor-turbine interface is an important factor in enhancing the efficiency of heavy-duty gas turbines. Also, it is a key requirement to fulfill the lifetime in this hottest area of the gas turbine. Typically transition pieces of can combustors induce a highly non-uniform swirling flow at the turbine inlet. In order to better understand the impact of the non-uniform combustor flow at the first stage vanes, a combined experimental and numerical study was carried out. The experimental facility consisted of a high speed linear cascade with four vane passages, including an upstream transition piece, which was representative of a heavy duty gas turbine can combustor-turbine interface geometry. The experiments were conducted at engine representative Mach numbers and film cooling effectiveness measurements were performed at three different blowing ratios. CFD RANS simulations were undertaken using a commercial flow solver. The numerical model was first validated with the experimental data, using inlet traverse 5-hole probe measurements, pressure taps along the airfoil perimeter and oil flow visualization results. The investigation shows that the position of the vane relative to the combustor transition piece has a significant impact on the vane aerodynamics and also film cooling behavior. This understanding was key to a robust first vane aerothermal design of the GT36.

Impact of a Combustor–Turbine Interface on Turbine Vane Aerodynamics and Film Cooling

2020 ◽  
Vol 142 (7) ◽  
Author(s):  
Andreas Lerch ◽  
Rainer Bauer ◽  
Joerg Krueckels ◽  
Marc Henze
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Gas Turbines ◽  
High Speed ◽  
Swirling Flow ◽  
Numerical Study ◽  
Heavy Duty ◽  
Film Cooling Effectiveness ◽  
Transition Piece ◽  
The Impact

Abstract Optimizing the aerothermal performance of the combustor–turbine interface is an important factor in enhancing the efficiency of heavy-duty gas turbines. Also, it is a key requirement to fulfill the lifetime in this hottest area of the gas turbine. Typically transition pieces of can combustors induce a highly nonuniform swirling flow at the turbine inlet. In order to better understand the impact of the nonuniform combustor flow at the first stage vanes, a combined experimental and numerical study was carried out. The experimental facility consisted of a high-speed linear cascade with four vane passages, including an upstream transition piece, which was representative of a heavy-duty gas turbine can combustor–turbine interface geometry. The experiments were conducted at engine representative Mach numbers, and film cooling effectiveness measurements were performed at three different blowing ratios. Computational fluid dynamics (CFD) Reynolds-averaged Navier–Stokes simulations were undertaken using a commercial flow solver. The numerical model was first validated with the experimental data, using inlet traverse five-hole probe measurements, pressure taps along the airfoil perimeter, and oil flow visualization results. The investigation shows that the position of the vane relative to the combustor transition piece has a significant impact on the vane aerodynamics and also film cooling behavior. This understanding was important for a robust first vane aerothermal design of the GT36.

Impact of Different Film-Cooling Modes at Trailing Edge on the Aerodynamic Performance of Heavy Duty Gas Turbine Cascade

2011 ◽  
Vol 84-85 ◽  
pp. 259-263
Author(s):  
Xun Liu ◽  
Song Tao Wang ◽  
Xun Zhou ◽  
Guo Tai Feng
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Large Mass ◽  
Turbine Cascade ◽  
Heavy Duty ◽  
Trailing Edge ◽  
Central Difference ◽  
Adiabatic Effectiveness ◽  
Heavy Duty Gas Turbine ◽  
The Impact

In this paper, the trailing edge film cooling flow field of a heavy duty gas turbine cascade has been studied by central difference scheme and multi-block grid technique. The research is based on the three-dimensional N-S equation solver. By way of analysis of the temperature field, the distribution of profile pressure, and the distribution of film-cooling adiabatic effectiveness in the region of trailing edge with different cool air injection mass and different angles, it is found that the impact on the film-cooling adiabatic effectiveness is slightly by changing the injection mass. The distribution of profile pressure dropped intensely at the pressure side near the injection holes line with the large mass cooling air. The cooling effect is good in the region of trailing edge while the injection air is along the direction of stream.

Account of Film Turbulence for Predicting Film Cooling Effectiveness in Gas Turbine Combustors

1979 ◽  
Author(s):  
G. J. Sturgess
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Cooling Effectiveness ◽  
Turbine Engine ◽  
Film Cooling Effectiveness ◽  
Wide Range ◽  
Engine Combustion ◽  
Turbulence Effects ◽  
The Impact ◽  
Combustor Liner

The paper deals with a small but important part of the overall gas turbine engine combustion system and continues earlier published work on turbulence effects in film cooling to cover the case of film turbulence. Film cooling of the gas turbine combustor liner imposes certain geometric limitations on the coolant injection device. The impact of practical film injection geometry on the cooling is one of increased rates of film decay when compared to the performance from idealized injection geometries at similar injection conditions. It is important to combustor durability and life estimation to be able to predict accurately the performance obtainable from a given practical slot. The coolant film is modeled as three distinct regions, and the effects of injection slot geometry on the development of each region are described in terms of film turbulence intensity and initial circumferential non-uniformity of the injected coolant. The concept of the well-designed slot is introduced and film effectiveness is shown to be dependent on it. Only slots which can be described as well-designed are of interest in practical equipment design. A prediction procedure is provided for well-designed slots which describes growth of the film downstream of the first of the three film regions. Comparisons of predictions with measured data are made for several very different well-designed slots over a relatively wide range of injection conditions, and good agreement is shown.

Account of Film Turbulence for Predicting Film Cooling Effectiveness in Gas Turbine Combustors

1980 ◽  
Vol 102 (3) ◽  
pp. 524-534 ◽  
Author(s):  
G. J. Sturgess
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Cooling Effectiveness ◽  
Turbine Engine ◽  
Film Cooling Effectiveness ◽  
Wide Range ◽  
Engine Combustion ◽  
Turbulence Effects ◽  
The Impact ◽  
Combustor Liner

The paper deals with a small but important part of the overall gas turbine engine combustion system and continues earlier published work on turbulence effects in film cooling to cover the case of film turbulence. Film cooling of the gas turbine combustor liner imposes certain geometric limitations on the coolant injection device. The impact of practical film injection geometry on the cooling is one of increased rates of film decay when compared to the performance from idealized injection geometries at similar injection conditions. It is important to combustor durability and life estimation to be able to predict accurately the performance obtainable from a given practical slot. The coolant film is modeled as three distinct regions, and the effects of injection slot geometry on the development of each region are described in terms of film turbulence intensity and initial circumferential non-uniformity of the injected coolant. The concept of the well-designed slot is introduced and film effectiveness is shown to be dependent on it. Only slots which can be described as well-designed are of interest in practical equipment design. A prediction procedure is provided for well-designed slots which describes growth of the film downstream of the first of the three film regions. Comparisons of predictions with measured data are made for several very different well-designed slots over a relatively wide range of injection conditions, and good agreement is shown.

Axial Compressor Degradation Effects on Heavy Duty Gas Turbines Overall Performances

2013 ◽  
Author(s):  
Pio Astrua ◽  
Stefano Cecchi ◽  
Stefano Piola ◽  
Andrea Silingardi ◽  
Federico Bonzani
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Axial Compressor ◽  
Performance Parameters ◽  
Heavy Duty ◽  
Compressor Blades ◽  
Modular Model ◽  
Secondary Air ◽  
The Impact ◽  
Insight Into

The operation of a gas turbine is the result of the aero-thermodynamic matching of several components which necessarily experience aging and degradation over time. An approach to treat degradation phenomena of the axial compressor is provided, with an insight into the impact they have on compressor operation and on overall GT performances. The analysis is focused on the surface fouling of compressor blades and on rotor tip clearances variation. A modular model is used to simulate the gas turbine operation in design and off-design conditions and the aerodynamic impact of fouling and rotor tip clearances increase is assessed by means of dedicated loss and deviation correlations implemented in the 1D mid-streamline code of the compressor modules. The two different degradation sources are individually considered and besides the overall GT performance parameters, the analysis includes an evaluation of the compressor degradation impact on the secondary air system.

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.

Advances in Gas Turbine Technology and the Digital Computer

1996 ◽  
Author(s):  
Melvin Platt
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
High Speed ◽  
Historical Context ◽  
Interactive Operation ◽  
Technology Advancement ◽  
Coherent Image ◽  
Key Technologies ◽  
The Impact ◽  
Digital Computers

Recent advances in gas turbine technology from computational fluid dynamics (CFD) to concurrent engineering are identified and discussed. These advances are placed in a historical context. Parallels are drawn between advances in gas turbine technology and digital computers. The advent of high-speed digital computers in the late 50s, followed by interactive operation and display graphics in the 70s. and orders-of-magnitude increases in affordable processing speed more recently, have each enabled key technologies for the gas turbine. Examples of such key technologies are given and their evolution is considered. Further, the impact of that technology on gas turbine development is discussed. Those perspectives provide a more coherent image of technology advancement, and thus a greater ability to identify new and future advances that will effect the evolution of gas turbines. The paper concludes with a look into the near future of gas turbine technology.

High Coverage Blade Tip Film Cooling

2004 ◽  
Author(s):  
M. Kuwabara ◽  
Keizo Tsukagoshi ◽  
T. Arts
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbines ◽  
Separation Bubble ◽  
Density Ratio ◽  
Tip Clearance ◽  
Experimental Program ◽  
Film Cooling Effectiveness ◽  
Blade Tip

More sophisticated cooling schemes are required for the turbine blade due to the demand of increased turbine temperature for improved performance. Although the tip portion of a turbine blade is one of the most critical portions in a gas turbine, there are few studies on cooling this portion compared to those for airfoil, especially film cooling strategies. Industrial gas turbines have a more uniform gas temperature profile than aero engines. For these applications, it is more important to understand the characteristics of tip film cooling to improve the blade durability and gas turbine performance by reducing cooling air. A numerical and experimental program was initiated to study film cooling effectiveness on a flat blade tip as a function of tip gap and mass flux ratios. Flow visualization tests were conducted with and without film cooling to verify the numerical CFD findings. The predictions and visualization results showed that a separation bubble forms at the pressure side edge that increases with tip gap. Film effectiveness measurements were carried out on a 1.3X scale blade model in a low speed test while simulating the normalized pressure distribution typical of an engine design. The engine density ratio of the coolant to mainstream was replicated in the film cooling tests to provide the best simulation of the engine. Two rows of holes were placed near the tip of the blade to provide high film coverage prior to the flowing over the tip. The data shows that film effectiveness increases with decreasing tip clearance. Blowing ratio provides an improvement due to the added mass flow, which was shown by a non-dimensionalized correlation.

Numerical Study of Film Cooling Enhancement in Gas Turbine Combustor Liner

2008 ◽  
Author(s):  
Ganesh Subbuswamy ◽  
Xianchang Li
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Structural Integrity ◽  
Numerical Study ◽  
Current Work ◽  
Working Fluid ◽  
Large Temperature ◽  
Film Cooling Effectiveness ◽  
Hot Air ◽  
Combustor Liner

Combustion chamber or combustor is one of the hottest parts of a gas turbine. Liner is where the actual flame occurs in a combustor and thus, the hottest part of the combustor. The temperature of working fluid inside a liner is about 1200 to 2000K. Because of the hot fluid, the liner is heated up to a temperature almost impossible for the material to withstand. Although the mechanical stresses experienced by the combustor liner are within acceptable limits, high temperatures and large temperature gradients affect the structural integrity of its components, which makes the liner a very critical component of a gas turbine in structural and thermal designs. Film cooling is a traditional method of cooling the inner surface of liner. In film cooling for a combustor, axial holes are drilled along the surface of the liner at discrete locations, through which cold air is injected axially into the liner to provide a film of cool air that prevents direct contact of hot air, and thus, protects the inner wall surface. The film is destroyed in the downstream to the flow because of mixing of cool and hot air. Though this method provides an acceptable cooling, there is a compromise with the increased net benefits of the gas turbine. Therefore, there is a need for new cooling techniques or enhancing the techniques available. The current work is a numerical simulation of film cooling in a model combustor. The effect of coolant injection angles and blowing ratios on film cooling effectiveness is studied. One innovative method, cooling with mist injection, is explored to enhance the performance of film cooling. The effect of droplet size and mist concentration, which can affect the performance of the mist injection, is also analyzed. Fluent, a commercial CFD software, is used in the current work for numerical simulations.

Double-Jet Film-Cooling for Highly Efficient Film-Cooling With Low Blowing Ratios

2008 ◽  
Author(s):  
Karsten Kusterer ◽  
Anas Elyas ◽  
Dieter Bohn ◽  
Takao Sugimoto ◽  
Ryozo Tanaka
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Gas Turbines ◽  
Original Design ◽  
Long Distance ◽  
Mixing Processes ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Air ◽  
Very High

Further improvement of the thermal efficiency of modern gas turbines can be achieved by a further reduction of the cooling air amount. Therefore, it is necessary to increase the cooling effectiveness so that the available cooling air fulfils the cooling task even if the amount has been reduced. In particular, the cooling effort for the vanes and blades of the first stage in a modern gas turbine is very high. The task of the film-cooling is to protect the blade material from the hot gas attack to the surface. Unfortunately, aerodynamic mixing processes are enhanced by secondary vortices in the cooling jets and, thus, the film-cooling effectiveness is reduced shortly behind the cooling air ejection through the holes. By improvement of the hole positioning the negative interaction effects can be reduced. The Double-jet Film-cooling (DJFC) Technology invented by the authors is one method to reach a significant increase in film-cooling effectiveness by establishing an anti-kidney vortex pair in a combined jet from the two jets starting from cylindrical ejection holes. This has been shown by numerical investigations and application to an industrial gas turbine as reported in recent publications. Whereas the original design application has been for moderate and high blowing ratios, the present numerical investigation shows that the DJFC is also applicable for lower blowing ratios (0.5<M<1.0) with only slight modification of the geometry of the configuration. The anti-kidney vortex concept can also be established for the lower blowing ratios and, as a result, a very high film-cooling effectiveness is reached not only behind the ejection holes but also for a very long distance downstream (> 30 hole diameters).

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