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.

Design Improvement Survey for NOx Emissions Reduction of a Heavy-Duty Gas Turbine Partially Premixed Fuel Nozzle Operating With Natural Gas: Experimental Campaign

2015 ◽  
Author(s):  
Matteo Cerutti ◽  
Roberto Modi ◽  
Danielle Kalitan ◽  
Kapil K. Singh
Keyword(s):  
Pressure Drop ◽  
Natural Gas ◽  
Gas Turbine ◽  
Pollutant Emissions ◽  
Nox Emissions ◽  
Heavy Duty ◽  
Fuel Nozzle ◽  
Partially Premixed ◽  
Heavy Duty Gas Turbine ◽  
The Impact

As government regulations become increasingly strict with regards to combustion pollutant emissions, new gas turbine combustor designs must produce lower NOx while also maintaining acceptable combustor operability. The design and implementation of an efficient fuel/air premixer is paramount to achieving low emissions. Options for improving the design of a natural gas fired heavy-duty gas turbine partially premixed fuel nozzle have been considered in the current study. In particular, the study focused on fuel injection and pilot/main interaction at high pressure and high inlet temperature. NOx emissions results have been reported and analyzed for a baseline nozzle first. Available experience is shared in this paper in the form of a NOx correlative model, giving evidence of the consistency of current results with past campaigns. Subsequently, new fuel nozzle premixer designs have been investigated and compared, mainly in terms of NOx emissions performance. The operating range of investigation has been preliminarily checked by means of a flame stability assessment. Adequate margin to lean blow out and thermo-acoustic instabilities onset has been found while also maintaining acceptable CO emissions. NOx emission data were collected over a variety of fuel/air ratios and pilot/main splits for all the fuel nozzle configurations. Results clearly indicated the most effective design option in reducing NOx. In addition, the impact of each design modification has been quantified and the baseline correlative NOx emissions model calibrated to describe the new fuel nozzles behavior. Effect of inlet air pressure has been evaluated and included in the models, allowing the extensive use of less costly reduced pressure test campaigns hereafter. Although the observed effect of combustor pressure drop on NOx is not dominant for this particular fuel nozzle, sensitivity has been performed to consolidate gathered experience and to make the model able to evaluate even small design changes affecting pressure drop.

scholarly journals Quasi-3D Numerical Computations on a Film-Cooled Gas Turbine Nozzle

1998 ◽  
Author(s):  
F. Bassi ◽  
S. Rebay ◽  
M. Savini
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Complex Geometry ◽  
Computational Effort ◽  
Navier Stokes ◽  
Gas Turbine Blade ◽  
Trailing Edge ◽  
Density Ratios ◽  
The Impact

The aim of this work is to assess the accuracy of a “quasi-3d” Navier-Stokes solver equipped with the k-ω turbulence model in the computation of a film-cooled gas turbine blade under a variety of flow conditions. The “quasi-3d” formulation was chosen as a cheap approach to investigate a large number of test conditions for a nozzle of complex geometry (around 400 cooling holes) which would require a large computational effort for a truly 3d simulation. The developed code has been used to investigate the influence of various cooling geometries and blowing conditions (mass flow rate and/or density ratios) on the aerodynamic behaviour of the cascade (in terms of loading, losses and flow angles) and their impact on the mixing process downstream of the trailing edge. The investigated nozzle is an advanced design turbine vane working in high subsonic regime. It is characterized by a marked endwall contouring at the casing and by the presence of 12 rows of holes (including a trailing edge row of slots) so as to obtain full-coverage film-cooling of the solid surfaces. This vane has been extensively tested in the Politecnico di Milano Fluid Dynamics Laboratory (formerly C.N.P.M.) blowdown transonic wind tunnel and a great amount of data are therefore available for validation purposes. The uselfulness of the proposed approach is fully analyzed and discussed throughout the paper and it is shown that the relation between the cascade performance and the variation of the investigated parameters is correctly described. In addition we address and discuss which ejection boundary conditions and which loss definitions are best suited for a meaningful comparison with the experimental measurements. In conclusion, in the case considered the developed code seems to be a valuable tool to determine the impact of film-cooling on the aerodynamic performance of a gas turbine blade.

Effect of Natural Gas Composition on Low NOx Burners Operation in Heavy Duty Gas Turbine

2019 ◽  
Author(s):  
Serena Romano ◽  
Matteo Cerutti ◽  
Giovanni Riccio ◽  
Antonio Andreini ◽  
Christian Romano
Keyword(s):  
Natural Gas ◽  
Gas Turbine ◽  
Operating Conditions ◽  
Fuel Composition ◽  
Gas Composition ◽  
Heavy Duty ◽  
Wide Range ◽  
Annular Combustor ◽  
Heavy Duty Gas Turbine ◽  
The Impact

Abstract Development of lean-premixed combustion technology with low emissions and stable operation in an increasingly wide range of operating conditions requires a deep understanding of the mechanisms that affect the combustion performance or even the operability of the entire gas turbine. Due to the relative wide range of natural gas composition supplies and the increased demand from Oil&Gas customers to burn unprocessed gas as well as LNG with notable higher hydrocarbons (C2+) content; the impact on gas turbine operability and combustion related aspects has been matter of several studies. In this paper, results of experimental test campaign of an annular combustor for heavy-duty gas turbine are presented with focus on the effect of fuel composition on both emissions and flame stability. Test campaign involved two different facilities, a full annular combustor rig and a full-scale prototype engine fed with different fuel mixtures of natural gas with small to moderate C2H6 content. Emissions trends and blowout for several operating conditions and burner configurations have been analyzed. Modifications to the burner geometry and fuel injection optimization have shown to be able to reach a good trade-off while keeping low NOx emissions in stable operating conditions for varying fuel composition.

Effect of Natural Gas Composition on Low Nox Burners Operation in Heavy Duty Gas Turbine

2019 ◽  
Vol 141 (11) ◽  
Author(s):  
Serena Romano ◽  
Matteo Cerutti ◽  
Giovanni Riccio ◽  
Antonio Andreini ◽  
Christian Romano
Keyword(s):  
Natural Gas ◽  
Gas Turbine ◽  
Operating Conditions ◽  
Fuel Composition ◽  
Gas Composition ◽  
Heavy Duty ◽  
Wide Range ◽  
Annular Combustor ◽  
Heavy Duty Gas Turbine ◽  
The Impact

Abstract Development of lean-premixed combustion technology with low emissions and stable operation in an increasingly wide range of operating conditions requires a deep understanding of the mechanisms that affect the combustion performance or even the operability of the entire gas turbine. Due to the relative wide range of natural gas composition supplies and the increased demand from Oil&Gas customers to burn unprocessed gas as well as liquified natural gas (LNG) with notable higher hydrocarbons (C2+) content, the impact on gas turbine operability and combustion related aspects has been matter of several studies. In this paper, results of experimental test campaign of an annular combustor for heavy-duty gas turbine are presented with focus on the effect of fuel composition on both emissions and flame stability. Test campaign involved two different facilities, a full annular combustor rig and a full-scale prototype engine fed with different fuel mixtures of natural gas with small to moderate C2H6 content. Emission trends and blowout for several operating conditions and burner configurations have been analyzed. Modifications to the burner geometry and fuel injection optimization have shown to be able to reach a good tradeoff while keeping low NOx emissions in stable operating conditions for varying fuel composition.

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 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.

Sign in / Sign up

Export Citation Format

Share Document