Influence of Opposing Dilution Jets on Effusion Cooling

2021 ◽  
Author(s):  
M. Riley Creer ◽  
Karen A. Thole
Keyword(s):  
Gas Turbine ◽  
Momentum Flux ◽  
Static Pressure ◽  
Combustion Process ◽  
Flux Ratio ◽  
High Momentum ◽  
Cooling Effectiveness ◽  
Static Pressure Distribution ◽  
The Impact ◽  
Combustor Liner

Abstract The gas turbine combustion process reaches gas temperatures that exceed the melting temperature of the combustor liner materials. Cooling the liner is critical to combustor durability and is often accomplished with double-walled liners that contain both impingement and effusion holes. The liner cooling is complicated with the interruption of the effusion cooling by large dilution jets that facilitate the combustion process. Given the presence of the dilution jets, it is important to understand the effect that the dilution jet has on the opposing wall in respect to the effusion film. This research includes measurements of the local static pressure distribution for a range of dilution jet momentum flux ratios to investigate the impact that the opposing dilution jet has on the effusion film. The interactions with the effusion cooling were also evaluated by measuring the overall cooling effectiveness across the panel. Measurements show that the opposing dilution jets did impact the liner at dilution jet momentum flux ratios that were greater than 20. The impacts at high momentum flux ratios were indicated through increased local static pressures measured on the surface of the combustor liner. Furthermore, the dilution touchdown decreased the overall cooling effectiveness of the effusion cooling. Results also indicated that the opposing dilution jets changed position on the liner as the dilution jet momentum flux ratio changes.

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.

Heat Transfer Measurements Downstream of a Two-Dimensional Jet Entering a Crossflow

1987 ◽  
Vol 109 (4) ◽  
pp. 572-578 ◽  
Author(s):  
S. Wittig ◽  
V. Scherer
Keyword(s):  
Heat Transfer ◽  
Flow Field ◽  
Pressure Distribution ◽  
Momentum Flux ◽  
Static Pressure ◽  
Flux Ratio ◽  
Main Flow ◽  
Two Dimensional ◽  
Reattachment Point ◽  
Static Pressure Distribution

Nusselt and Stanton numbers have been evaluated in and behind the recirculating zone produced by a two-dimensional jet entering a crossflow. The momentum flux ratio of the jet to the main flow was varied from 1.44–8.4 and measurements of the static pressure distribution and of the flow field by a five-hole probe were performed. A relation between the location of the reattachment point of the flow and the maximum of heat transfer was observed. Comparisons with available data are made. The experiments are intended for the verification of calculational codes.

scholarly journals Comprehensive Thermodynamic Analysis of the Humphrey Cycle for Gas Turbines with Pressure Gain Combustion

Energies ◽  
2018 ◽  
Vol 11 (12) ◽  
pp. 3521 ◽  
Author(s):  
Panagiotis Stathopoulos
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Combustion Process ◽  
Ideal Gas ◽  
Point Of View ◽  
Pressure Gain Combustion ◽  
Combustion Technology ◽  
Secondary Air ◽  
And Performance ◽  
The Impact

Conventional gas turbines are approaching their efficiency limits and performance gains are becoming increasingly difficult to achieve. Pressure Gain Combustion (PGC) has emerged as a very promising technology in this respect, due to the higher thermal efficiency of the respective ideal gas turbine thermodynamic cycles. Up to date, only very simplified models of open cycle gas turbines with pressure gain combustion have been considered. However, the integration of a fundamentally different combustion technology will be inherently connected with additional losses. Entropy generation in the combustion process, combustor inlet pressure loss (a central issue for pressure gain combustors), and the impact of PGC on the secondary air system (especially blade cooling) are all very important parameters that have been neglected. The current work uses the Humphrey cycle in an attempt to address all these issues in order to provide gas turbine component designers with benchmark efficiency values for individual components of gas turbines with PGC. The analysis concludes with some recommendations for the best strategy to integrate turbine expanders with PGC combustors. This is done from a purely thermodynamic point of view, again with the goal to deliver design benchmark values for a more realistic interpretation of the cycle.

Hydrogen Blending Into Ansaldo Energia AE94.3A Gas Turbine: High Pressure Tests, Field Experience and Modelling Considerations

2021 ◽  
Author(s):  
A. Ciani ◽  
L. Tay-Wo-Chong ◽  
A. Amato ◽  
E. Bertolotto ◽  
G. Spataro
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Power Plants ◽  
Flame Structure ◽  
Flux Ratio ◽  
Fuel Blends ◽  
Pressure Tests ◽  
Fuel Staging ◽  
Combustion Tests ◽  
The Impact

Abstract Fuel flexibility in gas turbine development has become increasingly important and modern engines need to cope with a broad variety of fuels. The target to operate power plants with hydrogen-based fuels and low emissions will be of paramount importance in a future focusing on electric power decarbonization. Ansaldo Energia AE94.3A engine acquired broad experience with operation of various natural gas and hydrogen fuel blends, starting in 2006 in the Brindisi (Italy) power plant. Based on the exhaustive experience acquired in the field, this paper describes the latest advancements characterizing the operation of the AE94.3A burner with high pressure combustion tests adding hydrogen blends ranging from 0 to 40% in volume. The interpretation of the test results is supported by reactive and non-reactive simulations describing the effects of varying fuel reactivity on the flame structure as well as the impact of fuel / air momentum flux ratio on the fuel / air interaction and fuel distribution in the combustion chamber. As expected, increasing amounts of hydrogen in the fuel are also associated with higher amounts of NOx production, however this effect could be countered by optimization of the fuel staging strategy, based on the mentioned CFD considerations and feedback from high pressure tests.

Numerical Investigation of Coolant-to-Mainstream Scaling Parameters With Film Cooling on Pressure and Suction Side of a Gas Turbine Blade

2017 ◽  
Author(s):  
Lingyu Zeng ◽  
Xueying Li ◽  
Jing Ren ◽  
Hongde Jiang
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Momentum Flux ◽  
Velocity Ratio ◽  
Density Ratio ◽  
Flux Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Blowing Ratio ◽  
Adiabatic Effectiveness

Most experiments of blade film cooling are conducted with density ratio lower than that of turbine conditions. In order to accurately model the performance of film cooling under a high density ratio, choosing an appropriate coolant to mainstream scaling parameter is necessary. The effect of density ratio on film cooling effectiveness on the surface of a gas turbine twisted blade is investigated from a numerical point of view. One row of film holes are arranged in the pressure side and two rows in the suction side. All the film holes are cylindrical holes with a pitch to diameter ratio P/d = 8.4. The inclined angle is 30°on the pressure side and 34° on the suction side. The steady solutions are obtained by solving Reynolds-Averaged-Navier-Stokes equations with a finite volume method. The SST turbulence model coupled with γ-θ transition model is applied for the present simulations. A film cooling experiment of a turbine vane was done to validate the turbulence model. Four different density ratios (DR) from 0.97 to 2.5 are studied. To independently vary the blowing ratio (M), momentum flux ratio (I) and velocity ratio (VR) of the coolant to the mainstream, seven conditions (M varying from 0.25 to 1.6 on the pressure side and from 0.25 to 1.4 on the suction side) are simulated for each density ratio. The results indicate that the adiabatic effectiveness increases with the increase of density ratio for a certain blowing ratio or a certain momentum flux ratio. Both on the pressure side and suction side, none of the three parameters listed above can serve as a scaling parameter independent of density ratio in the full range. The velocity ratio provides a relative better collapse of the adiabatic effectiveness than M and I for larger VRs. A new parameter describing the performance of film cooling is introduced. The new parameter is found to be scaled with VR for nearly the whole range.

scholarly journals Drop Size Characteristics of Forward Angled Injectors in Subsonic Cross Flows

2013 ◽  
Author(s):  
Venkat S. Iyengar ◽  
Sathiyamoorthy Kumarasamy ◽  
Srinivas Jangam ◽  
Manjunath Pulumathi
Keyword(s):  
Gas Turbine ◽  
Test Section ◽  
Momentum Flux ◽  
Drop Size ◽  
Cross Flow ◽  
Flux Ratio ◽  
Liquid Jet ◽  
Axial Distance ◽  
Momentum Flux Ratio ◽  
Spray Plume

Cross flow fuel injection is a widely used approach for injecting liquid fuel in gas turbine combustors and afterburners due to the higher penetration and rapid mixing of fuel and the cross flowing airstream. Because of the very limited residence time available in these combustors it is essential to ensure that smaller drop sizes are generated within a short axial distance from the injector in order to promote effective mixing. This requirement calls for detailed investigations into spray characteristics of different injector configurations in a cross-flow environment for identifying promising configurations. The drop size characteristics of a liquid jet issuing from a forward angled injector into a cross-flow of air were investigated experimentally at conditions relevant to gas turbine afterburners. A rig was designed and fabricated to investigate the injection of liquid jet in subsonic cross-flow with a rectangular test section of cross section measuring 50 mm by 70 mm. Experiments were done with a 10 degree forward angled 0.8 mm diameter plain orifice nozzle which was flush mounted on the bottom plate of test section. Laser diffraction using Malvern Spraytec particle analyzer was used to measure drops size and distributions in the near field of the spray. Measurements were performed at a distance of 70 mm from the injector at various locations along the height of the spray plume for a reasonable range of liquid flow rates as in practical devices. The sprays were characterized using the non dimensional parameters such as the Weber number and the momentum flux ratio and drop sizes were measured at three locations along the height of the spray from the bottom wall. The momentum flux ratio was varied from 5 to 25. Results indicate that with increase in momentum flux ratio the SMD reduced at the specific locations and an higher overall SMD was observed as one goes from the bottom to the top of the spray plume. This was accompanied by a narrowing of the drop size distribution.

Numerical Investigation of Gas Turbine Combustor Liner Film Cooling Slots

2018 ◽  
Author(s):  
Firat Kiyici ◽  
Ahmet Topal ◽  
Ender Hepkaya ◽  
Sinan Inanli
Keyword(s):  
Heat Transfer ◽  
Prandtl Number ◽  
Surface Temperature ◽  
Gas Turbine ◽  
Film Cooling ◽  
Turbulent Prandtl Number ◽  
Surface Cooling ◽  
Gas Turbine Combustor ◽  
Cooling Effectiveness ◽  
Combustor Liner

A numerical study, based on experimental work of Inanli et al. [1] is conducted to understand the heat transfer characteristics of film cooled test plates that represent the gas turbine combustor liner cooling system. Film cooling tests are conducted by six different slot geometries and they are scaled-up model of real combustor liner. Three different blowing ratios are applied to six different geometries and surface cooling effectiveness is determined for each test condition by measuring the surface temperature distribution. Effects of geometrical and flow parameters on cooling effectiveness are investigated. In this study, Conjugate Heat Transfer (CHT) simulations are performed with different turbulence models. Effect of the turbulent Prandtl Number is also investigated in terms of heat transfer distribution along the measurement surface. For this purpose, turbulent Prandtl number is calculated with a correlation as a function of local surface temperature gradient and its effect also compared with the constant turbulent Prandtl numbers. Good agreement is obtained with two-layered k–ϵ with modified Turbulent Prandtl number.

Effect of Cross-Flow Swirl on the Trajectory of Spray in an Annular Passage

2020 ◽  
Author(s):  
Tushar Sikroria ◽  
Abhijit Kushari
Keyword(s):  
Momentum Flux ◽  
High Speed ◽  
Air Flow ◽  
Cross Flow ◽  
Flux Ratio ◽  
Liquid Jet ◽  
Swirl Number ◽  
Flow Swirl ◽  
Air Stream ◽  
The Impact

Abstract This paper presents the experimental analysis of the impact of swirl number of cross-flowing air stream on liquid jet spray trajectory at a fixed air flow velocity of 42 m/s with the corresponding Mach number of 0.12. The experiments were conducted for 4 different swirl numbers (0, 0.2, 0.42 and 0.73) using swirl vanes at air inlet having angles of 0°, 15°, 30° and 45° respectively. Liquid to air momentum flux ratio (q) was varied from 5 to 25. High speed (@ 500 fps) images of the spray were captured and those images were processed using MATLAB to obtain the path of the spray at various momentum flux ratios. The results show interesting trends for the spray trajectory and the jet spread in swirling air flow. High swirling flows not only lead to spray with lower radial penetration due to sharp bending and disintegration of liquid jet, but also result in spray with high jet spread and spray area. Based on the results, correlations for the spray path have been proposed which incorporates the effects of the swirl number of the air flow.

Assessment of the Operational Performance of Fischer-Tropsch Synthetic-Paraffinic Kerosene in a T63 Gas Turbine Compared to Conventional Jet A-1 Fuel

2009 ◽  
Author(s):  
Nigel Bester ◽  
Andy Yates
Keyword(s):  
Gas Turbine ◽  
Thermal Loading ◽  
Flame Temperature ◽  
Combustion Products ◽  
Combustion Efficiency ◽  
Efficiency Gain ◽  
Engine Efficiency ◽  
Exit Temperature ◽  
The Impact ◽  
Combustor Liner

The performance implications of operating on Synthetic-Paraffinic Kerosene (SPK) were investigated using a RR-Allison T63-A-700 Model 250-C18 B gas turbine and compared to conventional Jet A-1. The SPK was aromatic–free and possessed a greater hydrogen/carbon ratio than petroleum derived Jet A-1. The variation in aromatic content had several implications with respect to soot and NOx emissions. Reduced aromatics also implied a reduction in the radiative heat transfer to the combustor liner. A simple model was used to explore the effect of H/C ratio on the adiabatic flame temperature, the combustor exit temperature and the engine efficiency via the impact on the gas properties and these were compared to the experimental data. It was found that operation with SPK changed directionally toward improving energy extraction via a turbine and an overall efficiency gain of about 1.2% was attained with operation on SPK through increased combustion efficiency, a reduction in liner pressure loss and an improvement in the combustion products properties. A modified combustion liner was fitted to enable the thermal loading on the combustor liner to be investigated and the expected trend with the SPK fuel was confirmed and quantified.

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