Impact of Predicted Combustor Outlet Conditions on the Aerothermal Performance of Film-Cooled High Pressure Turbine Vanes

2018 ◽  
Vol 141 (5) ◽  
Author(s):  
S. Cubeda ◽  
L. Mazzei ◽  
T. Bacci ◽  
A. Andreini
Keyword(s):  
Film Cooling ◽  
Cooling System ◽  
Engine Performance ◽  
Thermal Design ◽  
Separate Analysis ◽  
Lean Burn ◽  
Rans Model ◽  
Temperature Distributions ◽  
Inlet Conditions ◽  
The Impact

Turbine inlet conditions in lean-burn aeroengine combustors are highly swirled and present nonuniform temperature distributions. Uncertainty and lack of confidence associated with combustor-turbine interaction affect significantly engine performance and efficiency. It is well known that only Large-eddy and scale-adaptive simulations (SAS) can overcome the limitations of Reynolds-averaged Navier–Stokes (RANS) in predicting the combustor outlet conditions. However, it is worth investigating the impact of such improvements on the predicted aerothermal performance of the nozzle guide vanes (NGVs), usually studied with RANS-generated boundary conditions. Three numerical modelling strategies were used to investigate a combustor-turbine module designed within the EU Project FACTOR: (i) RANS model of the NGVs with RANS-generated inlet conditions; (ii) RANS model of the NGVs with scale-adaptive simulation (SAS)-generated inlet conditions; (iii) SAS model inclusive of both combustor and NGVs. It was shown that estimating the aerodynamics through the NGVs does not demand particularly complex approaches, in contrast to situations where turbulent mixing is key. High-fidelity predictions of the turbine entrance conditions proved very beneficial to reduce the discrepancies in the estimation of adiabatic temperature distributions. However, a further leap forward can be achieved with an integrated simulation, capable of reproducing the transport of unsteady fluctuations generated from the combustor through the turbine, which play a key role in presence of film cooling. This work, therefore, shows how separate analysis of combustor and NGVs can lead to a poor estimation of the thermal loads and ultimately to a wrong thermal design of the cooling system.

Impact of Predicted Combustor Outlet Conditions on the Aerothermal Performance of Film-Cooled HPT Vanes

2018 ◽  
Author(s):  
S. Cubeda ◽  
L. Mazzei ◽  
T. Bacci ◽  
A. Andreini
Keyword(s):  
Cooling System ◽  
Thermal Design ◽  
Temperature Fields ◽  
Thermal Loads ◽  
Rans Model ◽  
Numerical Predictions ◽  
Safety Margins ◽  
Rans Models ◽  
Inlet Conditions ◽  
The Impact

Turbine inlet conditions in modern aero-engines employing lean-burn combustors are characterised by highly swirled flow and non-uniform temperature distributions. As a consequence of the lack of confidence in numerical predictions and the uncertainty of measurement campaigns, the use of wide safety margins is of common practice in the design of turbine cooling systems, thus affecting the engine performance and efficiency. Previous experiences showed how only scale-resolving approaches such as Large-eddy and Scale-adapting simulations are capable of overcoming the limitations of RANS, significantly improving the accuracy in the prediction of flow and temperature fields at the combustor outlet. However it is worth investigating the impact of such differences on the aerothermal performance of the NGVs, as to understand the limitations entailed in the current approach for their thermal design. Industrial applications in fact usually rely on 1D, circumferentially-averaged profiles of pressure, velocity and temperature at the combustor-turbine interface in conjunction with Reynolds-averaged Navier-Stokes (RANS) models. This paper describes the investigation performed on an experimental test case consisting of a combustor simulator equipped with NGVs. Three numerical modelling strategies were compared in terms of flow field and thermal loads on the film-cooled vanes: i) RANS model of the NGVs with inlet conditions obtained from a RANS simulation of the combustor; ii) RANS model of the NGVs with inlet conditions obtained from a Scale-Adaptive Simulation (SAS) of the combustor; iii) SAS model inclusive of both combustor and NGVs. The results of this study show that estimating the aerodynamics at the NGV exit does not demand particularly complex approaches, whereas the limitations of standard RANS models are highlighted again when the turbulent mixing is key. High-fidelity predictions of the conditions at the turbine entrance proved to be very beneficial to reduce discrepancies in the estimation of local adiabatic wall temperature of even 100 K. However, a further leap forward can be achieved with an integrated simulation, capable of reproducing the transport of the unsteady fluctuations generated in the combustor up into the turbine, which plays a key role in presence of film cooling. This work therefore points out how keeping the analysis of combustor and NGVs separate can lead to a significantly misleading estimation of the thermal loads and ultimately to a wrong thermal design of the cooling system.

Assessment of Aero-Thermal Design Methodology for Effusion Cooled Lean Burn Annular Combustors

2014 ◽  
Author(s):  
A. Andreini ◽  
B. Facchini ◽  
L. Mazzei ◽  
L. Bellocci ◽  
F. Turrini
Keyword(s):  
Film Cooling ◽  
Cooling System ◽  
Heat Removal ◽  
Protective Layer ◽  
Aircraft Engine ◽  
Thermal Design ◽  
Heat Transfer Analysis ◽  
Lean Burn ◽  
Annular Combustor ◽  
Combustion Technology

Increasingly stringent limitations imposed on aircraft engine emissions have led many manufacturers toward lean combustion technology, which involves a relevant increase in mass flow rate dedicated to primary combustion, leading as a consequence to a reduction of air dedicated to cooling of liners. One of the most promising cooling techniques in such conditions is represented by effusion cooling, which consists of an array of closely spaced discrete film cooling holes. This cooling method is based on a protective layer of cooling flow on the hot side of the liner, enhancing at the same time the heat removal within the holes. In the latest years many aero engine manufacturers have increased the research and technology investment on this combustion technology. Working in partnership with the University of Florence, specific component design tools and experimental techniques have been improved by Avio Aero for combustor gas turbine investigation. From a design perspective, CFD analysis has become a key tool up to the early stages of novel combustor design process, producing affordable direct 3D optimization of combustor aerodynamics. Nevertheless, a RANS simulation of even only a single sector of an annular combustor still presents a challenge when the cooling system is taken into account. This issue becomes more critical in case of modern effusion cooled combustors, which may contain up to two thousand holes for the single sector. For this reason, many efforts have been devoted to develop methodologies based on film cooling modeling. Among the approaches published in the literature, models based on local sources represent a good compromise between simplicity and accuracy, with the capability to automatically perform a Conjugate Heat Transfer analysis. This type of methodology has been already defined and validated by the authors, with comparison on effusion cooled plates in terms of experimental overall effectiveness measurements as well as the application on a tubular combustor test case. In the context of this work, the proposed approach has been applied to the analysis of a lean annular combustor with the purpose of investigating pressure losses, flow split and metal temperature field. The results obtained have been compared to experimental data and different numerical tools exploited during the preliminary design of these devices.

Thermal Analysis of Cooling System in a Gas Turbine Transition Piece

2011 ◽  
Author(s):  
Jun Su Park ◽  
Namgeon Yun ◽  
Hokyu Moon ◽  
Kyung Min Kim ◽  
Sin-Ho Kang ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Thermal Stresses ◽  
Cooling System ◽  
Structural Information ◽  
Thermal Analyses ◽  
Thermal Design ◽  
Transfer Coefficients ◽  
Transition Piece ◽  
Inlet Conditions

This paper presents thermal analyses of the cooling system of a transition piece, which is one of the primary hot components in a gas turbine engine. The thermal analyses include heat transfer distributions induced by heat and fluid flow, temperature, and thermal stresses. The purpose of this study is to provide basic thermal and structural information on transition piece, to facilitate their maintenance and repair. The study is carried out primarily by numerical methods, using the commercial software, Fluent and ANSYS. First, the combustion field in a combustion liner with nine fuel nozzles is analyzed to determine the inlet conditions of a transition piece. Using the results of this analysis, pressure distributions inside a transition piece are calculated. The outside of the transition piece in a dump diffuser system is also analyzed. Information on the pressure differences is then used to obtain data on cooling channel flow (one of the methods for cooling a transition piece). The cooling channels have exit holes that function as film-cooling holes. Thermal and flow analyses are carried out on the inside of a film-cooled transition piece. The results are used to investigate the adjacent temperatures and wall heat transfer coefficients inside the transition piece. Overall temperature and thermal stress distributions of the transition piece are obtained. These results will provide a direction to improve thermal design of transition piece.

HP Vane Aerodynamics and Heat Transfer in the Presence of Aggressive Inlet Swirl

2012 ◽  
Vol 135 (2) ◽  
Author(s):  
Imran Qureshi ◽  
Andy D. Smith ◽  
Thomas Povey
Keyword(s):  
Heat Transfer ◽  
Cooling System ◽  
Three Dimensional ◽  
Leading Edge ◽  
Good Deal ◽  
Vortex Center ◽  
Research Facility ◽  
Lean Burn ◽  
Numerical Predictions ◽  
The Impact

Modern lean burn combustors now employ aggressive swirlers to enhance fuel-air mixing and improve flame stability. The flow at combustor exit can therefore have high residual swirl. A good deal of research concerning the flow within the combustor is available in open literature. The impact of swirl on the aerodynamic and heat transfer characteristics of an HP turbine stage is not well understood, however. A combustor swirl simulator has been designed and commissioned in the Oxford Turbine Research Facility (OTRF), previously located at QinetiQ, Farnborough UK. The swirl simulator is capable of generating an engine-representative combustor exit swirl pattern. At the turbine inlet plane, yaw and pitch angles of over ±40 deg have been simulated. The turbine research facility used for the study is an engine scale, short duration, rotating transonic turbine, in which the nondimensional parameters for aerodynamics and heat transfer are matched to engine conditions. The research turbine was the unshrouded MT1 design. By design, the center of the vortex from the swirl simulator can be clocked to any circumferential position with respect to HP vane, and the vortex-to-vane count ratio is 1:2. For the current investigation, the clocking position was such that the vortex center was aligned with the vane leading edge (every second vane). Both the aligned vane and the adjacent vane were characterized. This paper presents measurements of HP vane surface and end wall heat transfer for the two vane positions. The results are compared with measurements conducted without swirl. The vane surface pressure distributions are also presented. The experimental measurements are compared with full-stage three-dimensional unsteady numerical predictions obtained using the Rolls Royce in-house code Hydra. The aerodynamic and heat transfer characterization presented in this paper is the first of its kind, and it is hoped to give some insight into the significant changes in the vane flow and heat transfer that occur in the current generation of low NOx combustors. The findings not only have implications for the vane aerodynamic design, but also for the cooling system design.

Assessment of RANS Against LES for the Aero-Thermal Behavior of High Pressure Turbine Stages Under Realistic Inlet Conditions

2016 ◽  
Author(s):  
Stefano Vagnoli ◽  
Tom Verstraete ◽  
Charlie Koupper ◽  
Guillaume Bonneau
Keyword(s):  
High Pressure ◽  
Thermal Behavior ◽  
Turbulence Modeling ◽  
Engine Performance ◽  
High Pressure Turbine ◽  
Lean Burn ◽  
Field Development ◽  
Safety Margins ◽  
Inlet Conditions ◽  
Hot Streak

Modern Lean Burn combustors generate a complex field at the High Pressure turbine (HPT) inlet, characterized by non-uniform velocity and temperature distributions, together with very high turbulence levels (up to 25%). For these extreme conditions, classical numerical methods employed for the HPT design, such as Reynolds Averaged Navier Stokes (RANS) simulation, suffer from a lack of validation. This leads to a reduced confidence in predicting the combustor-turbine interactions, which requires to use extra safety margins, to the detriment of the overall engine performance. Within the European FACTOR project, a 360° non reactive combustor simulator and a 1.5 HPT stage are designed to get more insight into the mutual interaction of these two components. A first experimental and numerical campaign has demonstrated the potential of Large Eddy Simulations (LES) to accurately reproduce the turbulent flow field development at the combustor outlet. The aim of the present paper is to exploit the accuracy of LES to validate less time-consuming RANS models in predicting the hot streak migration in the turbine stage. In this sense, LES results are used as a reference to discriminate the different RANS simulations in terms of turbulence modeling and aerothermal predictions. The current investigations clearly indicate that turbulence and hot streak diffusion within the HPT are strongly linked. In this sense, the choice of the RANS turbulence model and the inlet turbulent conditions plays a major role in modeling the thermal behavior for the stator and rotor blades.

The Impact of Compressor Exit Conditions on Fuel Injector Flows

2012 ◽  
Vol 134 (11) ◽  
Author(s):  
C. L. Ford ◽  
J. F. Carrotte ◽  
A. D. Walker
Keyword(s):  
Flow Field ◽  
Large Scale ◽  
Numerical Data ◽  
Flow Fields ◽  
Main Reaction ◽  
Fuel Injector ◽  
Lean Burn ◽  
Field Development ◽  
Inlet Conditions ◽  
The Impact

This paper examines the effect of compressor generated inlet conditions on the air flow uniformity through lean burn fuel injectors. Any resulting nonuniformity in the injector flow field can impact on local fuel air ratios and hence emissions performance. The geometry considered is typical of the lean burn systems currently being proposed for future, low emission aero engines. Initially, Reynolds-averaged Navier-Stokes (RANS) computational fluid dynamics (CFD) predictions were used to examine the flow field development between compressor exit and the inlet to the fuel injector. This enabled the main flow field features in this region to be characterized along with identification of the various stream-tubes captured by the fuel injector passages. The predictions indicate the resulting flow fields entering the injector passages are not uniform. This is particularly evident in the annular passages furthest away from the injector centerline which pass the majority of the flow which subsequently forms the main reaction zone within the flame tube. Detailed experimental measurements were also undertaken on a fully annular facility incorporating an axial compressor and lean burn combustion system. The measurements were obtained at near atmospheric pressure/temperatures and under nonreacting conditions. Time-resolved and time-averaged data were obtained at various locations and included measurements of the flow field issuing from the various fuel injector passages. In this way any nonuniformity in these flow fields could be quantified. In conjunction with the numerical data, the sources of nonuniformities in the injector exit plane were identified. For example, a large scale bulk variation (+/−10%) of the injector flow field was attributed to the development of the flow field upstream of the injector, compared with localized variations (+/−5%) that were generated by the injector swirl vane wakes. Using this data the potential effects on fuel injector emissions performance can be assessed.

Analysis of Swirl number effects on effusion flow behaviour using time resolved PIV

2022 ◽  
pp. 1-33
Author(s):  
Tommaso Lenzi ◽  
Alessio Picchi ◽  
Antonio Andreini ◽  
Bruno Facchini
Keyword(s):  
Film Cooling ◽  
Cooling System ◽  
Tangential Velocity ◽  
Particle Image Velocimetry System ◽  
Swirl Number ◽  
Time Resolved ◽  
Experimental Apparatus ◽  
Turbulence Length ◽  
The Impact ◽  
Near Wall

Abstract The analysis of the interaction between the swirling and liner film-cooling flows is a fundamental task for the design of turbine combustion chambers since it influences different aspects such as emissions and cooling capability. Particularly, high turbulence values, flow instabilities, and tangential velocity components induced by the swirlers deeply affect the behavior of effusion cooling jets, demanding for dedicated time-resolved near-wall analysis. The experimental setup of this work consists of a non-reactive single-sector linear combustor test rig scaled up with respect to engine dimensions; the test section was equipped with an effusion plate with standard inclined cylindrical holes to simulate the liner cooling system. The rig was instrumented with a 2D Time-Resolved Particle Image Velocimetry system, focused on different field of views. The degree of swirl is usually characterized by the swirl number, Sn, defined as the ratio of the tangential momentum to axial momentum flux. To assess the impact of such parameter on the near-wall effusion behavior, a set of three axial swirlers with swirl number equal to Sn = 0.6 − 0.8 − 1.0 were designed and tested in the experimental apparatus. An analysis of the main flow by varying the Sn was first performed in terms of average velocity, RMS, and Tu values, providing kinetic energy spectra and turbulence length scale information. Following, the analysis was focused on the near-wall regions: the effects of Sn on the coolant jets was quantified in terms of vorticity analysis and jet oscillation.

The Impact of Compressor Exit Conditions on Fuel Injector Flows

2012 ◽  
Author(s):  
C. L. Ford ◽  
J. F. Carrotte ◽  
A. D. Walker
Keyword(s):  
Flow Field ◽  
Large Scale ◽  
Numerical Data ◽  
Flow Fields ◽  
Main Reaction ◽  
Fuel Injector ◽  
Lean Burn ◽  
Field Development ◽  
Inlet Conditions ◽  
The Impact

This paper examines the effect of compressor generated inlet conditions on the air flow uniformity through lean burn fuel injectors. Any resulting non-uniformity in the injector flow field can impact on local fuel air ratios and hence emissions performance. The geometry considered is typical of the lean burn systems currently being proposed for future, low emission aero engines. Initially, RANS CFD predictions were used to examine the flow field development between compressor exit and the inlet to the fuel injector. This enabled the main flow field features in this region to be characterized along with identification of the various stream-tubes captured by the fuel injector passages. The predictions indicate the resulting flow fields entering the injector passages are not uniform. This is particularly evident in the annular passages furthest away from the injector center-line which pass the majority of the flow which subsequently forms the main reaction zone within the flame tube. Detailed experimental measurements were also undertaken on a fully annular facility incorporating an axial compressor and lean burn combustion system. The measurements were obtained at near atmospheric pressure/temperatures and under non-reacting conditions. Time-resolved and time-averaged data were obtained at various locations and included measurements of the flow field issuing from the various fuel injector passages. In this way any non-uniformity in these flow fields could be quantified. In conjunction with the numerical data, the sources of non-uniformities in the injector exit plane were identified. For example, a large scale bulk variation (+/−10%) of the injector flow field was attributed to the development of the flow field upstream of the injector, compared with localized variations (+/−5%) that were generated by the injector swirl vane wakes. Using this data the potential effects on fuel injector emissions performance can be assessed.

Experimental Investigation of the Flow Field and the Heat Transfer on a Scaled Cooled Combustor Liner With Realistic Swirling Flow Generated by a Lean-Burn Injection System

2014 ◽  
Author(s):  
Antonio Andreini ◽  
Gianluca Caciolli ◽  
Bruno Facchini ◽  
Alessio Picchi ◽  
Fabio Turrini
Keyword(s):  
Heat Transfer ◽  
Flow Field ◽  
Film Cooling ◽  
Swirling Flow ◽  
Injection System ◽  
Lean Burn ◽  
Real Geometry ◽  
Slot Film Cooling ◽  
The Impact ◽  
Combustor Liner

Lean burn swirl stabilized combustors represent the key technology to reduce NOx emissions in modern aircraft engines. The high amount of air admitted through a lean-burn injection system is characterized by very complex flow structures such as recirculations, vortex breakdown and processing vortex core, that may deeply interact in the near wall region of the combustor liner. This interaction and its effects on the local cooling performance make the design of the cooling systems very challenging, accounting for the design and commission of new test rigs for detailed analysis. The main purpose of the present work is the characterization of the flow field and the wall heat transfer due to the interaction of a swirling flow coming out from real geometry injectors and a slot cooling system which generates film cooling in the first part of the combustor liner. The experimental setup consists of a non-reactive three sector planar rig in an open loop wind tunnel; the rig, developed within the EU project LEMCOTEC, includes three swirlers, whose scaled geometry reproduces the real geometry of an Avio Aero PERM (Partially Evaporated and Rapid Mixing) injector technology, and a simple cooling scheme made up of a slot injection, reproducing the exhaust dome cooling mass flow. Test were carried out imposing realistic combustor operating conditions, especially in terms of reduced mass flow rate and pressure drop across the swirlers. The flow field is investigated by means of PIV, while the measurement of the heat transfer coefficient is performed through Thermochromic Liquid Crystals steady state technique. PIV results show the behavior of flow field generated by the injectors, their mutual interaction and the impact of the swirled main flow on the stability of the slot film cooling. TLC measurements, reported in terms of detailed 2D heat transfer coefficient maps, highlight the impact of the swirled flow and slot film cooling on wall heat transfer.

Effusion Cooling System Optimization for Modern Lean Burn Combustor

2016 ◽  
Author(s):  
Antonio Andreini ◽  
Riccardo Becchi ◽  
Bruno Facchini ◽  
Lorenzo Mazzei ◽  
Alessio Picchi ◽  
...  
Keyword(s):  
Film Cooling ◽  
Cooling System ◽  
Heat Removal ◽  
System Optimization ◽  
Field Investigation ◽  
Metal Temperature ◽  
Back Side ◽  
Lean Burn ◽  
Standard Configuration ◽  
Cooling Air

Stricter legislation limits concerning NOx emissions are leading main aero-engine manufacturers to update the architecture of the combustors towards the implementation of lean burn combustion concept. Cooling air availability for the thermal management of combustor liners is significantly reduced, demanding even more effective liner cooling schemes. The state-of-the-art of liner cooling technology is represented by effusion cooling, consisting in a very efficient cooling strategy based on multi-perforated liners, where metal temperature is lowered by the combined protective effect of coolant film and heat removal inside the holes. The present research study aims at deepening the knowledge of effusion systems, exploiting the results of a thorough experimental campaign carried out in two different planar test rigs, equipped with a complete liner cooling scheme composed by slot injection and effusion array. The film cooling protection was analysed using PSP (Pressure Sensitive Paint) technique, while the effect of cooling injection and extraction from the annulus on heat transfer distribution were studied by means of TLC (Thermochromic Liquid Crystals) thermography. Thermal measurements were supported by flow field investigation with standard 2D PIV (Particle Image Velocimetry) in order to highlight the typical velocity distributions generated by a realistic lean injector. These detailed experimental data were exploited in a 1D thermal flow-network solver that allows to better assess the main cooling mechanisms characterising the proposed cooling system. Moreover, an optimized cooling configuration with enhanced back-side convective cooling was proposed and compared with the standard configuration in terms of metal temperature and cooling consumption.

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