This paper reports the numerical study performed during the setup of a tubular laboratory combustor developed by Avio Group. The main purpose of this study is the assessment of some numerical tools which might be employed for the design of aero-engines combustors. Adopted methodologies pertains with heat transfer and exhaust emissions issues of combustor design, and they refer to both detailed three dimensional and simplified zero/one dimensional formulations. An exhaustive discussion about the obtained result and about the adopted modelling criteria is reported in the paper. The rig has a tubular geometry that allows to study the single burner of a typical annular combustor of jet engines. Considered operating condition for the tests refers to Max Take-Off (ICAO 100%) in the cycle of a typical jet engine. Three different numerical tools were considered, representing the fundamental steps during the preliminary and detailed design of combustors. The first one is a 1D procedure capable to analyze the cooling flow network of the combustor and to predict liner wall temperature and heat loads. The second is a classical chemical reactor network code for flame temperature and exhaust emissions prediction. Such preliminary design tools are usually supported by reactive CFD computations. In this work both two dimensional and three dimensional CFD models are considered with different goals. A 2D CFD model of flame tube with prescribed mass flow rates allows to quickly test various turbulence, combustion and emissions models, while a 3D model of the complete hardware (liner, burner and cooling network), permits to predict air flow splits and wall radiative and convective heat loads. Furthermore a final 3D reactive CFD computation with radiation conjugated with the thermal conduction solution across the liner, permits to estimate the wall temperature distribution. The whole set of numerical results has pointed out an appreciable agreement among the various tools representing a valid assessment of the validity and robustness of selected design methodologies. A more significant validation of codes accuracy will be possible as soon as the scheduled experimental results will be available.
Lean burn combustion is increasing its popularity in the aeronautical framework due to its potential in reducing drastically pollutant emissions (NOx and soot in particular). Its implementation, however, involves significant issues related to the increased amount of air dedicated to the combustion process, demanding the redesign of injection and cooling systems. Also, the conditions at the combustor exit are a concern, as high turbulence, residual swirl, and the impossibility to adjust the temperature profile with dilution holes determine a harsher environment for nozzle guide vanes. This work describes the final stages of the design of an aeronautical effusion-cooled lean burn combustor. Full annular tests were carried out to measure temperature profiles and emissions (CO and NOx) at the combustor exit. Different operating conditions of the ICAO cycle were tested, considering Idle, Cruise, Approach, and Take-off. Scale-adaptive simulations with the flamelet generated manifold (FGM) combustion model were performed to extend the validation of the employed computational fluid dynamics (CFD) methodology and to reproduce the experimental data in terms of radial temperature distribution factor (RTDF)/overall temperature distribution factor (OTDF) profiles as well as emission indexes (EIs). The satisfactory agreement paved the way to an exploitation of the methodology to provide a deeper understanding of the flow physics within the combustion chamber, highlighting the impact of the different operating conditions on flame, spray evolution, and pollutant formation.
Lean burn combustion is increasing its popularity in the aeronautical framework due to its potential in reducing drastically pollutant emissions (NOx and soot in particular). Its implementation however involves significant issues related to the increased amount of air dedicated to the combustion process, demanding the redesign of injection and cooling systems. Also the conditions at the combustor exit are a concern, as high turbulence, residual swirl and the impossibility to adjust the temperature profile with dilution holes determine a harsher environment for nozzle guide vanes. This work describes the final stages of the design of an aeronautical effusion-cooled lean burn combustor. Full annular tests were carried out to measure temperature profiles and emissions (CO and NOx) at the combustor exit. Different operating conditions of the ICAO cycle were tested, considering Idle, Cruise, Approach and Take-Off. Scale-adaptive simulations with the Flamelet Generated Manifold combustion model were performed to extend the validation of the employed CFD methodology and to reproduce the experimental data in terms of RTDF/OTDF profiles as well as emission indexes. The satisfactory agreement paved the way to an exploitation of the methodology to provide a deeper understanding of the flow physics within the combustion chamber, highlighting the impact of the different operating conditions on flame, spray evolution and pollutant formation.
Effusion cooled liners, commonly used in gas turbine combustion chambers to reduce wall temperature, may also help reducing the propagation of pressure fluctuations due to thermoacoustic instabilities.Large Eddy Simulations were conducted to accurately model the flow field and the acoustic response of effusion plates subject to a mean bias flow under external sinusoidal forcing. Even though existing lower order computational models showed good predicting capabilities, it is interesting to verify directly the influence of those parameters such as the staggered arrangement, the hole inclination, the presence of a grazing flow and the level of bias flow, which are not fully included in those models.A first bi-periodic single hole configuration with normal acoustic forcing was selected to investigate the acousting behavior with varying inclination angle, bias and grazing flow. 90 • and 30 • perforations were simulated for bias flow Mach number in the range 0.05-0.1 and grazing flow between 0 and 0.08. Those conditions were chosen to expand the knowledge of acoustic properties towards actual liners working conditions. A second more computationally expensive set-up, including 4 inclined holes at 30 • , focused on the damping of parallel to the plate waves.Details of the computational methods implemented in the general purpose open-source unstructured CFD code OpenFOAM R exploited to conduct this analysis are reported together with an analysis of the results obtained from the acoustic computations both regarding the flow field generated and the absorption and energy dissipation coefficient.
In order to reduce NOx emissions, modern gas turbines are often equipped with lean burn combustion systems, where the engine operates near the lean blow-out limits. One of the most critical issues of lean combustion technology is the onset of combustion instabilities related to a coupling between pressure oscillations and thermal fluctuations excited by the unsteady heat release. In this work a thermoacoustic analysis of a full annular combustor developed by AVIO is discussed. The system is equipped with an advanced PERM (Partially Evaporating and Rapid Mixing) injection system based on a piloted lean burn spray flame generated by a pre-filming atomizer. Combustor walls are based on multi-perforated liners to control metal temperature: these devices are also recognized as very effective sound absorbers, thus in innovative lean combustors they could represent a good means both for wall cooling and damping combustion instabilities. The performed analysis is based on the resolution of the eigenvalue problem related to an inhomogeneous wave equation which includes a source term representing heat release fluctuations (the so called Flame Transfer Function, FTF) in the flame region using a three-dimensional FEM code. A model representing the entire combustor was assembled including all the acoustically relevant geometrical features. In particular, the acoustic effect of multi-perforated liners was introduced by modeling the corresponding surfaces with an equivalent internal impedance. Different simulations with and without the presence of the flame were performed analyzing the influence of the multi-perforated liners. Furthermore, different modeling approaches for the FTF were examined and compared with each other. Comparisons with available experimental data showed a good agreement in terms of resonant frequencies in the case of passive simulations. On the other hand, when the presence of the flame is considered, comparisons with experiments showed the inadequacy of FTFs commonly used for premixed combustion and thus the necessity of an improved FTF, more suitable for liquid fueled gas turbines where the evaporation process could play an important role in the flame heat release fluctuations.
scite is a Brooklyn-based organization that helps researchers better discover and understand research articles through Smart Citations–citations that display the context of the citation and describe whether the article provides supporting or contrasting evidence. scite is used by students and researchers from around the world and is funded in part by the National Science Foundation and the National Institute on Drug Abuse of the National Institutes of Health.
customersupport@researchsolutions.com
10624 S. Eastern Ave., Ste. A-614
Henderson, NV 89052, USA
Copyright © 2024 scite LLC. All rights reserved.
Made with 💙 for researchers
Part of the Research Solutions Family.