Literatura académica sobre el tema "Aero-engine combustors"
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Artículos de revistas sobre el tema "Aero-engine combustors"
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Şöhret, Yasin, and T. Hikmet Karakoc. "Exergy indicators of a low-emission aero-engine combustor." Aircraft Engineering and Aerospace Technology 90, no. 2 (2018): 344–50. http://dx.doi.org/10.1108/aeat-03-2016-0045.
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Purpose It is essential to develop more environment-friendly energy systems to prevent climate change and minimize environmental impact. Within this scope, many studies are performed on performance and environmental assessments of many types of energy systems. This paper, different from previous studies, aims to prove exergy performance of a low-emission combustor of an aero-engine. Design/methodology/approach It is a well-known fact that, with respect to previous exergy analysis, highest exergy destruction occurs in the combustor component of the engine. For this reason, it is required to evaluate a low-emission aero-engine combustor thermodynamically to understand the state of the art according to the authors’ best of knowledge. In this framework, combustor has been operated at numerous conditions (variable engine load) and evaluated. Findings As a conclusion of the study, the impact of emission reduction on performance improvement of the aero-engine combustors exergetically is presented. It is stated that exergy efficiency of the low-emission aero-engine combustor is found to be 64.69, 61.95 and 71.97 per cent under various operating conditions. Practical implications Results obtained in this paper may be beneficial for researchers who are interested in combustion and propulsion technology and thermal sciences. Originality/value Different from former studies, the impact of operating conditions on performance of a combustor is examined from the viewpoint of thermodynamics.
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Li, J., X. Sun, Y. Liu, and V. Sethi. "Preliminary aerodynamic design methodology for aero engine lean direct injection combustors." Aeronautical Journal 121, no. 1242 (2017): 1087–108. http://dx.doi.org/10.1017/aer.2017.47.
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ABSTRACTThe Lean Direct Injection (LDI) combustor is one of the low-emissions combustors with great potential in aero-engine applications, especially those with high overall pressure ratio. A preliminary design tool providing basic combustor sizing information and qualitative assessment of performance and emission characteristics of the LDI combustor within a short period of time will be of great value to designers. In this research, the methodology of preliminary aerodynamic design for a second-generation LDI (LDI-2) combustor was explored. A computer code was developed based on this method covering the design of air distribution, combustor sizing, diffuser, dilution holes and swirlers. The NASA correlations for NOxemissions are also embedded in the program in order to estimate the NOx production of the designed LDI combustor. A case study was carried out through the design of an LDI-2 combustor named as CULDI2015 and the comparison with an existing rich-burn, quick-quench, lean-burn combustor operating at identical conditions. It is discovered that the LDI combustor could potentially achieve a reduction in liner length and NOxemissions by 18% and 67%, respectively. A sensitivity study on parameters such as equivalence ratio, dome and passage velocity and fuel staging is performed to investigate the effect of design uncertainties on both preliminary design results and NOxproduction. A summary on the variation of design parameters and their impact is presented. The developed tool is proved to be valuable to preliminarily evaluate the LDI combustor performance and NOxemission at the early design stage.
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Marudhappan, Raja, Chandrasekhar Udayagiri, and Koni Hemachandra Reddy. "Combustion chamber design and reaction modeling for aero turbo-shaft engine." Aircraft Engineering and Aerospace Technology 91, no. 1 (2018): 94–111. http://dx.doi.org/10.1108/aeat-10-2017-0217.
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Purpose The purpose of this paper is to formulate a structured approach to design an annular diffusion flame combustion chamber for use in the development of a 1,400 kW range aero turbo shaft engine. The purpose is extended to perform numerical combustion modeling by solving transient Favre Averaged Navier Stokes equations using realizable two equation k-e turbulence model and Discrete Ordinate radiation model. The presumed shape β-Probability Density Function (β-PDF) is used for turbulence chemistry interaction. The experiments are conducted on the real engine to validate the combustion chamber performance. Design/methodology/approach The combustor geometry is designed using the reference area method and semi-empirical correlations. The three dimensional combustor model is made using a commercial software. The numerical modeling of the combustion process is performed by following Eulerian approach. The functional testing of combustor was conducted to evaluate the performance. Findings The results obtained by the numerical modeling provide a detailed understanding of the combustor internal flow dynamics. The transient flame structures and streamline plots are presented. The velocity profiles obtained at different locations along the combustor by numerical modeling mostly go in-line with the previously published research works. The combustor exit temperature obtained by numerical modeling and experiment are found to be within the acceptable limit. These results form the basis of understanding the design procedure and opens-up avenues for further developments. Research limitations/implications Internal flow and combustion dynamics obtained from numerical simulation are not experimented owing to non-availability of adequate research facilities. Practical implications This study contributes toward the understanding of basic procedures and firsthand experience in the design aspects of combustors for aero-engine applications. This work also highlights one of the efficient, faster and economical aero gas turbine annular diffusion flame combustion chamber design and development. Originality/value The main novelty in this work is the incorporation of scoops in the dilution zone of the numerical model of combustion chamber to augment the effectiveness of cooling of combustion products to obtain the desired combustor exit temperature. The use of polyhedral cells for computational domain discretization in combustion modeling for aero engine application helps in achieving faster convergence and reliable predictions. The methodology and procedures presented in this work provide a basic understanding of the design aspects to the beginners working in the gas turbine combustors particularly meant for turbo shaft engines applications.
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ZIEMANN, J. "Low-NOx combustors for hydrogen fueled aero engine." International Journal of Hydrogen Energy 23, no. 4 (1998): 281–88. http://dx.doi.org/10.1016/s0360-3199(97)00054-2.
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Bake, Friedrich, Ulf Michel, and Ingo Roehle. "Investigation of Entropy Noise in Aero-Engine Combustors." Journal of Engineering for Gas Turbines and Power 129, no. 2 (2006): 370–76. http://dx.doi.org/10.1115/1.2364193.
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Strong evidence is presented that entropy noise is the major source of external noise in aero-engine combustion. Entropy noise is generated in the outlet nozzles of combustors. Low-frequency entropy noise, which was predicted earlier in theory and numerical simulations, was successfully detected in a generic aero-engine combustion chamber. It is shown that entropy noise dominates even in the case of thermo-acoustic resonances. In addition to this, a different noise generating mechanism was discovered that is presumably of even higher relevance to jet engines: There is strong evidence of broad band entropy noise at higher frequencies (1 to 3kHz in the reported tests). This unexpected effect can be explained by the interaction of small scale entropy perturbations (hot spots) with the strong pressure gradient in the outlet nozzle. The direct combustion noise of the flame zone seems to be of minor importance for the noise emission to the ambiance. The combustion experiments were supplemented by experiments with electrical heating. Two different methods for generating entropy waves were used, a pulse excitation and a sinusoidal excitation. In addition, high-frequency entropy noise was generated by steady electrical heating.
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Klose, G., R. Schmehl, R. Meier, et al. "Evaluation of Advanced Two-Phase Flow and Combustion Models for Predicting Low Emission Combustors." Journal of Engineering for Gas Turbines and Power 123, no. 4 (2000): 817–23. http://dx.doi.org/10.1115/1.1377010.
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The development of low-emission aero-engine combustors strongly depends on the availability of accurate and efficient numerical models. The prediction of the interaction between two-phase flow and chemical combustion is one of the major objectives of the simulation of combustor flows. In this paper, predictions of a swirl stabilized model combustor are compared to experimental data. The computational method is based on an Eulerian two-phase model in conjunction with an eddy dissipation (ED) and a presumed-shape-PDF (JPDF) combustion model. The combination of an Eulerian two-phase model with a JPDF combustion model is a novelty. It was found to give good agreement to the experimental data.
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Zhu, M., A. P. Dowling, and K. N. C. Bray. "Self-Excited Oscillations in Combustors With Spray Atomizers." Journal of Engineering for Gas Turbines and Power 123, no. 4 (2000): 779–86. http://dx.doi.org/10.1115/1.1376717.
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Combustors with fuel-spray atomizers are susceptible to a low-frequency oscillation, particularly at idle and sub-idle conditions. For aero-engine combustors, the frequency of this oscillation is typically in the range 50–120 Hz and is commonly called “rumble.” In the current work, computational fluid dynamics (CFD) is used to simulate this self-excited oscillation. The combustion model uses Monte Carlo techniques to give simultaneous solutions of the Williams’ spray equation together with the equations of turbulent reactive flow. The unsteady combustion is calculated by the laminar flamelet presumed pdf method. A quasi-steady description of fuel atomizer behavior is used to couple the inlet flow in the combustor. A choking condition is employed at turbine inlet. The effects of the atomizer and the combustor geometry on the unsteady combustion are studied. The results show that, for some atomizers, with a strong dependence of mean droplet size on air velocity, the coupled system undergoes low-frequency oscillations. The numerical results are analyzed to provide insight into the rumble phenomena. Basically, pressure variations in the combustor alter the inlet air and fuel spray characteristics, thereby changing the rate of combustion. This in turn leads to local “hot spots,” which generate pressure fluctuations as they convect through the downstream nozzle.
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Corsini, A., F. Rispoli, and T. E. Tezduyar. "Stabilized finite element computation of NOx emission in aero-engine combustors." International Journal for Numerical Methods in Fluids 65, no. 1-3 (2010): 254–70. http://dx.doi.org/10.1002/fld.2451.
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Tietz, S., and T. Behrendt. "Development and application of a pre-design tool for aero-engine combustors." CEAS Aeronautical Journal 2, no. 1-4 (2011): 111–23. http://dx.doi.org/10.1007/s13272-011-0012-x.
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Hu, Bin, Yong Huang, Fang Wang, and Fa Xie. "CFD predictions of LBO limits for aero-engine combustors using fuel iterative approximation." Chinese Journal of Aeronautics 26, no. 1 (2013): 74–84. http://dx.doi.org/10.1016/j.cja.2012.12.014.
Texto completoTesis sobre el tema "Aero-engine combustors"
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Vakil, Sachin Suresh. "Flow and Thermal Field Measurements in a Combustor Simulator Relevant to a Gas Turbine Aero-Engine." Thesis, Virginia Tech, 2002. http://hdl.handle.net/10919/36324.
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<p>The highly competitive gas turbine industry has been motivated by consumer demands for higher power-to-weight ratios, increased thermal efficiencies, and reliability while maintaining affordability. In its continual quest, the industry must continually try to raise the turbine inlet temperature, which according to the well-known Brayton cycle is key to higher engine efficiencies. The desire for increased turbine inlet temperatures creates an extremely harsh environment for the combustor liner in addition to the components downstream of the combustor. Shear layers between the dilution jets and the mainstream, as well as combustor liner film-cooling interactions create a complex mean flow field within the combustor, which is not easy to model. A completely uniform temperature and velocity profile at the combustor exit is desirable from the standpoint of reducing the secondary flows in the turbine. However, this seldom occurs due to a lack of thorough mixing within the combustor. Poor mixing results in non-uniformities, such as hot streaks, and allow non-combusted fuel to exit the combustor.
<p>This investigation developed a database documenting the thermal and flow characteristics within a combustor simulator representative of the flowfield within a gas turbine aero-engine. Three- and two-component laser Doppler velocimeter measurements were completed to quantify the flow and turbulence fields, while a thermocouple rake was used to quantify the thermal fields.
<p>The measured results show very high turbulence levels due to the dilution flow injection. Directly downstream of the dilution jets, an increased thickness in the film-cooling was noted with a fairly non-homogeneous temperature field across the combustor width. A highly turbulent shear layer was found at the leading edge of the dilution jets. Measurements also showed that a relatively extensive recirculation region existed downstream of the dilution jets. Despite the lack of film-cooling injection at the trailing edge of the dilution hole, there existed coolant flow indicative of a horse-shoe vortex wrapping around the jet. As a result of the dilution jet interaction with the mainstream flow, kidney-shaped thermal fields and counter-rotating vortices developed. These vortices serve to enhance combustor mixing.<br>Master of Science
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Jaegle, Félix. "LARGE EDDY SIMULATION OF EVAPORATING SPRAYS IN COMPLEX GEOMETRIES USING EULERIAN AND LAGRANGIAN METHODS." Phd thesis, Institut National Polytechnique de Toulouse - INPT, 2009. http://tel.archives-ouvertes.fr/tel-00452501.
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Dû aux efforts apportés à la réduction des émissions de NOx dans des chambres de combustion aéronautiques il y a une tendance récente vers des systèmes à combustion pauvre. Cela résulte dans l'apparition de nouveaux types d'injecteur qui sont caractérisés par une complexité géométrique accrue et par des nouvelles stratégies pour l'injection du carburant liquide, comme des systèmes multi-point. Les deux éléments créent des exigences supplémentaires pour des outils de simulation numériques. La simulation à grandes échelles (SGE ou LES en anglais) est aujourd'hui considérée comme la méthode la plus prometteuse pour capturer des phénomènes d'écoulement complexes qui apparaissent dans une telle application. Dans le présent travail, deux sujets principaux sont abordés: Le premier est le traitement de la paroi ce qui nécessite une modélisation qui reste délicate en SGE, en particulier dans des géométries complexes. Une nouvelle méthode d'implémentation pour des lois de paroi est proposée. Une étude dans une géométrie réaliste démontre que la nouvelle formulation donne de meilleurs résultats comparé à l'implémentation classique. Ensuite, la capacité d'une approche SGE typique (utilisant des lois de paroi) de prédire la perte de charge dans une géométrie représentative est analysée et des sources d'erreur sont identifiées. Le deuxième sujet est la simulation du carburant liquide dans une chambre de combustion. Avec des méthodes Eulériennes et Lagrangiennes, deux approches sont disponibles pour cette tâche. La méthode Eulérienne considère un spray de gouttelettes comme un milieu continu pour lequel on peut écrire des équations de transport. Dans la formulation Lagrangienne, des gouttes individuelles sont suivies ce qui mène à des équations simples. D'autre part, sur le plan numérique, le grand nombre de gouttes à traiter peut s'avérer délicat. La comparaison des deux méthodes sous conditions identiques (solveur gazeux, modèles physiques) est un aspect central du présent travail. Les phénomènes les plus importants dans ce contexte sont l'évaporation ainsi que le problème d'injection d'un jet liquide dans un écoulement gazeux transverse ce qui correspond à une version simplifiée d'un système multi-point. Le cas d'application final est la configuration d'un seul injecteur aéronautique, monté dans un banc d'essai expérimental. Ceci permet d'appliquer de manière simultanée tous les développements préliminaires de ce travail. L'écoulement considéré est non-réactif mais à part cela il correspond au régime ralenti d'un moteur d'avion. Dû aux conditions préchauffées, le spray issu du sstème d'injection multi-point s'évapore dans la chambre. Cet écoulement est simulé, utilisant les approaches Eulériennes et Lagrangiennes et les résultats sont comparés aux données expérimentales.
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Elmi, Carlo Alberto. "Design system integration for multi-objective optimization of aero engine combustors." Doctoral thesis, 2022. http://hdl.handle.net/2158/1276939.
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The transformation towards a climate-neutral civil aviation is providing significant business opportunities to the aero engine market players. To meet this target and keep competitiveness, however, groundbreaking solutions must be introduced at the product’s level in the shortest possible time. Industry lead-ers are increasingly embracing lean and digital approaches for this purpose, by applying these concepts at all company’s levels. Considerable room for im-provements can be identified in the development of complex components as, for instance, the combustor. Due to the complexity of phenomena taking place and interacting into it, there are conflicting functional requirements defined over different physical domains. This leads to a design approach that must be both multidisciplinary and multi-objective, in which the need for supporting know-how and product expertise arises with extensive and structured studies of the design space arises. Nowadays, simulation-based methodologies repre-sent a standard in evaluating multiple configurations of the system, although it may lead to heterogeneous models interacting with each other, sharing miscel-laneous information within the process. In this context, taking advantage of in-tegrated design systems has been proven to be beneficial in standardizing the simulation processes while embedding design’s best practices.
The subject matter of this work is the Combustor Design System Integra-tion (DSI), an integrated methodology aimed at easing and streamlining the preliminary design phase of aero engine combustors. Its concept will be de-scribed in the first part, where the automation of low value-added tasks will be introduced together with four custom integrated tools. It is composed of a CAD generation system, a RANS-based CFD suite for reactive flow calculations, a boundary-conditions processor for 3D thermal FEA and a FE structural envi-ronment for stress and displacement estimation. Particular importance is given to the definition of cooling and quenching systems on combustor’s liners, since their prominent impact on aero-thermal and durability performance. Therefore, specific features for a detailed topological management of holes are presented in this work, providing advance patterning and arrangement capabilities which are not addressed in other design systems. Finally, it will be possible to prove the reduction of lead time for analysis, as well as the enhancement of the overall process robustness. The NEWAC combustor, a lean-burn concept developed in the context of the homonymous European research project, will be exploited as a case allowing, moreover, an assessment of the DSI modelling approach. In the second part will be presented a dedicated framework for multi-objective design optimization, comprising the DSI tools for CAD generation and CFD analysis. A fully automated and water-tight process is here implemented in order to ad-dress the combustor’s problem of dilution mixing, aimed at optimizing the temperature profiles and the emission levels at its outlet. This approach will leverage on advanced neural network algorithms for improving the overall de-sign workflow, so to ensure that the optimal combustor configuration is de-fined as a function of the product’s Critical-To-Quality. The results of the opti-mization will be shown for a rich-quench-lean combustor concept intentionally designed to support this activity, referred as to LEM-RQL.
The general intention of this work, in the end, is to demonstrate how in-tegrated design systems embedded in optimization frameworks could repre-sent both a strategic asset for industry players and a relevant topic for academ-ics. Given the pervasive integration-and-automation of the process, the general-ity in processing multiple design layouts and the possibility to accommodate increasingly advanced and sophisticated optimization algorithms, the DSI pro-cedure configure itself as an ideal platform within the technology maturation process, thus enabling not only the improvement of in-service components but also the development of next-generation combustor products.
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PICCHI, ALESSIO. "Experimental Investigations of Effusion Cooling Systems for Lean Burn Aero-Engine Combustors." Doctoral thesis, 2014. http://hdl.handle.net/2158/857503.
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Legislation limits concerning polluting emissions, for civil aircraft engines, are expected to become even more stringent in the future. To meet these targets, especially in terms of NOx, it is required to maintain the temperature in the combustion zone as low as possible. Lean burn swirl stabilized combustors represent the key technology to reduce NOx emissions. 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. In addition, since up to 70% of the overall air mass flow is utilized for fuel preparation and the initiation of lean combustion, the amount of air available for combustor liner cooling has to be strongly reduced with respect to the traditional diffusive combustor architectures. State-of-the-art of liner cooling technology for modern combustors is represented by the effusion cooling. Effusion cooling is a very efficient cooling strategy based on the use of multi-perforated liners, where metal temperature is lowered by the combined protective effect of coolant film and heat removal through forced convection inside each hole. Beyond that, multiperforated liners act also as passive devices to mitigate thermoacoustic phenomena which is one of the main concern regarding lean combustors operability.
A large part of the activities and the achievements deriving from the Ph.D. course are collected in the present study, that deals with two experimental campaigns on effusion cooling schemes designed for aero-engine combustor liner applications.
In the first part of the current research, an experimental survey has been performed for the evaluation of thermal performance, in terms of overall and adiabatic effectiveness, of seven multi-perforated planar plates representative of a portion of combustor liner, with uniform mainstream conditions. Effusion geometries were tested imposing 6 blowing ratios in the range 0.5-5, two values of density ratio and two level of mainstream turbulence. Concerning the geometrical features, different porosity levels have been considered: such values are obtained both increasing the hole diameter and pattern spacing. Then, the effect of hole inclination and aspect ratio pattern shape have been tested to assess the impact of typical cooling system features. The analysis of the data points out the impact of the main geometrical and fluid dynamics parameters on the thermal performance, proposing a possible thermal optimization strategy that seems to be promising also from the acoustic damping requirements. Results represent a wide experimental database relevant for the design of an high efficiency effusion cooling systems, even though the survey leaves the impact of the swirled gas flow on thermal performance an open issue.
To enhance the TRL (Technology Readiness Level) of experiments, a planar sector test rig equipped with three AVIO Aero PERM (Partially Evaporated and Rapid Mixing) injector systems and working at atmospheric conditions has been considered in the second part of the work. The test rig allowed to reproduce a representative flow field on the gas side and to test the complete liner cooling scheme composed by a slot system, that reproduced the exhaust dome cooling mass flow, and an effusion array. The final aim of the study is the experimental characterization of the flow field and the measurement of cooling performance in terms of heat transfer coefficient and adiabatic effectiveness due to the interaction of the swirling flow coming out from the injectors and the cooling scheme. Tests were carried out imposing several realistic operating conditions, especially in terms of reduced mass flow rate and pressure drop across swirlers and effusion cooling holes.
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Palanti, Lorenzo. "On the modelling of liquid fuel ignition and atomization in aero engine combustors." Doctoral thesis, 2021. http://hdl.handle.net/2158/1234766.
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A fast and reliable relight of aero engines burners is one of the most critical points to ensure aircraft safety. The so-called altitude relight is the process that allows the combustor to be re-ignited after a flame-out during flight. Several expensive tests must be carried out to obtain the required certifications, which makes important to fully understand the problem of the flame onset. To speedup the design process, Computational Fluid Dynamics established as valid alternative to the experiments to investigate the complex phenomena involved in the ignition process. In this work, a fully reactive Large Eddy Simulation of the ignition process is attempted with the aim of validating the setup for future applications within an advanced design process. In this line, a throughout validation is carried out against some detailed experimental results of a lean spray flame. At first, a non-reactive and reactive simulations are carried out to validate the cold flow field and the stabilized flame structure. Then, an ignition simulation is performed, from initial spark deposition up to flame stabilization. The obtained results are extensively compared with the available experimental data, showing that the employed simulation setup can describe quite well the phenomena involved in the rig ignition. However, this procedure does not allow to perform statistical studies, such as the optimization of the igniter position. This issue is critical since it can help to minimize the amount of energy required for ignition and to increase the durability of the hardware. In fact, several spark discharges must be simulated for each position, to account for different realizations due to the turbulent flow field. Therefore, the previous Large Eddy Simulation approach is not feasible in this scenario, due to the excessive computational effort. In scientific literature, specific low-order models were developed to provide an affordable estimation of the local ignition probability. Due to the large amount of assumptions they clearly sacrifice part of the accuracy and of the physical consistency, but the short turnaround time makes them the first choice at low Technology Readiness Level. In this thesis, a low-order design model is implemented and used to investigate the ignition probability of the same test rig simulated in detail, showing that it can provide good results if a careful sensitivity study is firstly conducted. This set of simulations represents a first attempt in the scientific literature to carefully validate a low-order model using a flow-field from LES and very accurate experimental data. Moreover, the LES of the ignition process is used to further validate the model and highlight its main shortcomings. To the best of the author knowledge, such in-depth validation was not attempted so far. The last step of the thesis concerns the development of a simulation and post-processing methodology to investigate the primary breakup of the spray in altitude relight conditions. Such activity is motivated by the fundamental importance of spray initialization in full chamber simulations. In fact, due to the very poor atomization quality and the negligible evaporation process, the droplets ejected from the nozzle can travel all along the combustion chamber and finally reach the combustor walls. Under these conditions, the ignition tools already described might fail if the spray is injected with a wrong distribution. Therefore, a tentative approach to evaluate the spray size distribution is proposed in the last section of this work. It is based on a run-time evaluation of distributed variables such as liquid/gas density of interface or locally defined variables as the curvature of the interface. In the present work, the Eulerian Lagrangian Spray Atomization model is used as a basis to test such technique. An academic planar prefilmer configuration is selected to evaluate the accuracy of the post-processing, thanks to the availability of experimental measurements in the proximity of the injector. The reported results include a brief description of the simulated liquid structures, the predicted Sauter Mean Diameter, the evolution of interface curvature and a final proposal to derive the spray size distribution. The novelty of this last section is mainly represented by the use of the interface curvature to analyze and postprocess the primary breakup. At the best of the author knowledge, this is the first time that such an approach is attempted to an actual atomizing device. Overall, the aim of this work is to validate and to develop a set of tools to improve altitude relight design in aero engines. Although much work is still needed, this thesis represent a first step towards the use of more advanced numerical tools to optimize this process and to more easily meet the required certifications.
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Langone, Leonardo. "Numerical modelling of partially premixed low-swirl flames for aero-engine applications." Doctoral thesis, 2022. http://hdl.handle.net/2158/1277139.
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The development of innovative aero-engine combustors has been devoted to drastically reducing pollutants emissions and improving engine performances in recent years. These aspects are not only crucial to meet the severe regulations imposed by ICAO-CAEP, but also to enable potential new engine architectures. Especially considering Nitrogen Oxidizes (NOx) emissions, the most promising concept carried out so far is represented by Lean burn combustors, which however introduce several challenges in terms of ame stability. A possible solution to this problem is the novel burner concept proposed in the EU project CHAiRLIFT (Compact Helical Arranged combustoRs with lean LIFTed ames). The proposal of this project sees the employment of low-swirl ultra-lean spray lifted flames with an inclined disposition of the burners. Both these concepts have been investigated separately at the Karlsruhe Institute of Technology (KIT). In particular, this type of flame has shown superior performances in terms of NOx emissions a good resistance to Lean Blow-Off (LBO) occurrences, while avoiding flashback risks. The inclined arrangement of the burners, instead, establishes a macro-recirculation in the combustion chamber responsible for the transport of vitiated gas among the burners, promoting flame stability. Moreover, it contributes to reduce the need of cooling air and the overall weight. This ambitious project indeed requires proper tools to study flame interaction between adjacent burners in deep. The present research effort is therefore devoted to numerically investigating the CHAiRLIFT concept through Computational Fluid Dynamics (CFD) simulations. With this goal, both Scale Resolved (SR) simulations of the low-swirl burner in single sector con guration and Reynolds Averaged Navier-Stokes (RANS) simulations of the multiburner rig, currently under investigation at KIT, have been employed. The numerical study of the multiburner con figuration had a twofold objective: assess the numerical approach with the available data and support the experimental investigations, especially concerning the sensitivity to the tilt angle. The outcomes have shown that, in non-preheated conditions, the numerical simulation can fairly reproduce the spray lifted flames with a reasonable computational effort. Also, it points out that the best setup in terms of tilt angle for maximizing the exhausts recirculation lays between 20 and 30 degrees, which is a lower value concerning the original experimental investigation with high-swrirl flames. Another point to be addressed is the turbulent combustion modelling in preheated conditions, which has shown to be more challenging from the modelling point of view. To this aim, the investigation is focused on the same burner in single sector con guration, operated with gaseous fuel. State-of-the-art numerical models for Large Eddy Simulations context are employed to understand how the ame is reproduced. The results highlighted that both the Flamelet Generated Manifold (FGM) approach and a modifi ed version taking into account the stretch and heat loss effects are mispredicting the flame lift-off height. Instead, a good reproduction is achieved with the Thickened Flame (TF) model. Despite the good agreement of the latter approach, this suffers of some disadvantages in terms of chemistry description: aiming to overcome these issues, a hybrid TF-FGM approach is introduced and validated in the conclusive part of this work, with very interesting results in terms of lift-off and flame shape. In the nal part, the hybrid TF-FGM model is applied to the same con guration operated with spray together with dedicated spray boundary conditions carried out from the research activity of the COmplexe de Recherche Interprofessionnel en Aerothermochimie (CORIA) research team. The results pointed out that although the ame is still not perfectly predicted, a large improvement is reached concerning the combustion and spray modeling approaches commonly present in the literature.
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GIUSTI, ANDREA. "Development of numerical tools for the analysis of advanced airblast injection systems for lean burn aero-engine combustors." Doctoral thesis, 2014. http://hdl.handle.net/2158/867029.
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The liquid fuel preparation has a strong impact on the combustion process and consequently on pollutant emissions.
However, currently there are no validated and computational affordable methods available to predict the spray breakup process and to reliably compute the spray distribution generated after primary breakup. This research activity, carried out within the framework of the European project FIRST (Fuel Injector Research for Sustainable Transport), is aimed at developing reliable tools to be used in the industrial design process able to describe the processes involved in liquid fuel preparation in advanced injection systems based on prefilming airblast concept.
A multi-coupled solver for prefilming airblast injectors which includes liquid film evolution and primary breakup was developed in the framework of OpenFOAM. The solver is aimed at improving the description of the complex physical phenomena characterizing liquid fuel preparation and spray evolution in advanced airblast injection systems within the context of typical RANS (U-RANS) industrial calculations. In this kind of injectors, gas-phase, droplet and liquid film interact with each other, thus, in order to properly predict spray evolution and fuel distribution inside the combustor, proper tools able to catch the most important interactions among the different phases are necessary. A steady-state Eulerian-Lagrangian approach was introduced in the code together with up-to-date evaporation and secondary breakup models. Particular attention was devoted to the liquid film primary breakup and to the interactions between gas-phase and liquid film. A new primary breakup model for liquid films, basically a phenomenological model which exploits liquid film and gas-phase solutions for the computation of spray characteristics after breakup, was developed and implemented in the code. Different formulations for the computation of droplet diameter after breakup were evaluated and revised on the basis of recent experimental findings. The multi-coupled solver was validated against literature test cases with detailed experimental measurements and eventually applied to the simulation of an advanced prefilming airblast injector based on the PERM concept in a tubular combustor configuration.
The proposed approach allows us to better describe the fuel evolution in the injector region leading to a more comprehensive and physically consistent description of the phenomena regulating liquid fuel preparation compared to standard approaches which neglect the presence of liquid film and its interaction with both droplets and gas-phase.
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MAZZEI, LORENZO. "A 3D coupled approach for the thermal design of aero-engine combustor liners." Doctoral thesis, 2015. http://hdl.handle.net/2158/993808.
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The recent limitations imposed by ICAO-CAEP towards a drastic reduction of NOx emissions is driving the development of modern aeroengines towards the implementation of lean burn concept. The increased amount of air dedicated to the combustion process (up to 70%) involves several technological issues, including a signicant reduction of coolant available for the thermal management of combustor liners. This, from a design perspective, involves the continuous research for effective cooling schemes, such as effusion cooling, and the necessity of more accurate methodologies for the estimation of metal temperature, so as to properly assess the expected duration of hot gas path components.
The flame stabilization through swirler characterized by large effective area leads to extended recirculating zones, which interact considerably with the liner cooling system. As highlighted in the first part of this dissertation, the impact on the near-wall flow field makes any consideration based on a correlative approach untrustworthy, demanding for more reliable evaluations through CFD analysis. Unfortunately, the application of effusion cooling entails a huge computational effort due to the high number of film cooling holes involved, therefore many approaches have been proposed in literature with the aim of modelling the coolant injection through mass sources.
This work presents SAFE (Source based effusion model), a methodology for the CFD simulation of the entire combustor, which is based on the local coolant injection through point sources and a calculation of mass flow rate according to local flow conditions. A further step in reduction in the computational effort is represented by a different methodology, called Therm3D, which involves the simulation of the flame tube, whereas the solution of the remaining part of the combustor is fulfilled through the modelling of an equivalent flow network, which provides for the estimation of flow split and cold side heat loads.
Ultimately, this work introduces innovative approaches for the CFD investigation of effusion cooled combustor, with a special focus on the metal temperature prediction. A model for the film cooling injection is proposed to overcome the issues related to the necessity of meshing the perforation, nevertheless several improvable aspects have been highlighted, pointing the way for further enhancements.
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INSINNA, MASSIMILIANO. "Investigation of the Aero-Thermal Aspects of Combustor/Turbine Interaction in Gas Turbines." Doctoral thesis, 2015. http://hdl.handle.net/2158/986426.
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Bacci, Tommaso. "Experimental investigation on a high pressure NGV cascade in the presence of a representative lean burn aero-engine combustor outflow." Doctoral thesis, 2018. http://hdl.handle.net/2158/1128260.
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Experimental Investigation of the effects of a modern lean burn combustor outflow on the performance of a film-cooled NGV cascade. Evaluation of chamber flow field, NGV inlet/outlet aerothermal field, turbulence decay and adiabatic effectiveness on the NGV profiles
Libros sobre el tema "Aero-engine combustors"
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Panigrahi, Shashi Kanta, and Niranjan Sarangi. Aero Engine Combustor Casing. Taylor & Francis Group, 2020.
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Sarangi, Niranjan, and Sashi Kanta Panigrahi. Aero Engine Combustor Casing: Experimental Design and Fatigue Studies. Taylor & Francis Group, 2017.
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Sarangi, Niranjan, and Sashi Kanta Panigrahi. Aero Engine Combustor Casing: Experimental Design and Fatigue Studies. Taylor & Francis Group, 2017.
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Sarangi, Niranjan, and Sashi Kanta Panigrahi. Aero Engine Combustor Casing: Experimental Design and Fatigue Studies. Taylor & Francis Group, 2017.
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Sarangi, Niranjan, and Sashi Kanta Panigrahi. Aero Engine Combustor Casing: Experimental Design and Fatigue Studies. Taylor & Francis Group, 2017.
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Aero Engine Combustor Casing: Experimental Design and Fatigue Studies. Taylor & Francis Group, 2017.
Buscar texto completoCapítulos de libros sobre el tema "Aero-engine combustors"
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"Introduction." In Aero Engine Combustor Casing. CRC Press, 2017. http://dx.doi.org/10.1201/9781315116754-1.
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"References." In Aero Engine Combustor Casing. CRC Press, 2017. http://dx.doi.org/10.1201/9781315116754-10.
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"Index." In Aero Engine Combustor Casing. CRC Press, 2017. http://dx.doi.org/10.1201/9781315116754-11.
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"Fatigue Design Philosophy of an Aero Engine Combustor Casing." In Aero Engine Combustor Casing. CRC Press, 2017. http://dx.doi.org/10.1201/9781315116754-2.
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"Development of Test Facility and Test Setup." In Aero Engine Combustor Casing. CRC Press, 2017. http://dx.doi.org/10.1201/9781315116754-3.
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"Manufacturing of an Aero Engine Combustor Casing, the Experimental Evaluation of Its Fatigue Life, and Correlation with Numerical Results." In Aero Engine Combustor Casing. CRC Press, 2017. http://dx.doi.org/10.1201/9781315116754-4.
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"Reassessment of Fatigue Life of the Modified Combustor Casing." In Aero Engine Combustor Casing. CRC Press, 2017. http://dx.doi.org/10.1201/9781315116754-5.
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"Safety Test on Modified Combustor Casing." In Aero Engine Combustor Casing. CRC Press, 2017. http://dx.doi.org/10.1201/9781315116754-6.
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"Effect of Fatigue on the Proof Strength of an Aero Engine Combustor Casing*." In Aero Engine Combustor Casing. CRC Press, 2017. http://dx.doi.org/10.1201/9781315116754-7.
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"Conclusions." In Aero Engine Combustor Casing. CRC Press, 2017. http://dx.doi.org/10.1201/9781315116754-8.
Texto completoActas de conferencias sobre el tema "Aero-engine combustors"
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Bake, Friedrich, Ulf Michel, and Ingo Roehle. "Investigation of Entropy Noise in Aero-Engine Combustors." In ASME Turbo Expo 2006: Power for Land, Sea, and Air. ASMEDC, 2006. http://dx.doi.org/10.1115/gt2006-90093.
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Strong evidence is presented that entropy noise is the major source of external noise in aero-engine combustion. Entropy noise is generated in the outlet nozzles of combustors. Low frequency entropy noise — which was predicted earlier in theory and numerical simulations — was successfully detected in a generic aero-engine combustion chamber. It is shown that entropy noise dominates even in the case of thermo-acoustic resonances. In addition to this, a different noise generating mechanism was discovered that is presumably of even higher relevance to jet engines: There is strong evidence of broad band entropy noise at higher frequencies (1 kHz to 3 kHz in the reported tests). This unexpected effect can be explained by the interaction of small scale entropy perturbations (hot spots) with the strong pressure gradient in the outlet nozzle. The direct combustion noise of the flame zone seems to be of minor importance for the noise emission to the ambiance. The combustion experiments were supplemented by experiments with electrical heating. Two different methods for generating entropy waves were used, a pulse excitation and a sinusoidal excitation. In addition, high-frequency entropy noise was generated by steady electrical heating.
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Zelina, J., D. T. Shouse, J. S. Stutrud, G. J. Sturgess, and W. M. Roquemore. "Exploration of Compact Combustors for Reheat Cycle Aero Engine Applications." In ASME Turbo Expo 2006: Power for Land, Sea, and Air. ASMEDC, 2006. http://dx.doi.org/10.1115/gt2006-90179.
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An aero gas turbine engine has been proposed that uses a near-constant-temperature (NCT) cycle and an Inter-Turbine Burner (ITB) to provide large amounts of power extraction from the low-pressure turbine. This level of energy is achieved with a modest temperature rise across the ITB. The additional energy can be used to power a large geared fan for an ultra-high bypass ratio transport aircraft, or to drive an alternator for large amounts of electrical power extraction. Conventional gas turbines engines cannot drive ultra-large diameter fans without causing excessively high turbine temperatures, and cannot meet high power extraction demands without a loss of engine thrust. Reducing the size of the combustion system is key to make use of a NCT gas turbine cycle. Ultra-compact combustor (UCC) concepts are being explored experimentally. These systems use high swirl in a circumferential cavity about the engine centerline to enhance reaction rates via high cavity g-loading on the order of 3000 g’s. Any increase in reaction rate can be exploited to reduce combustor volume. The UCC design integrates compressor and turbine features which will enable a shorter and potentially less complex gas turbine engine. This paper will present experimental data of the Ultra-Compact Combustor (UCC) performance in vitiated flow. Vitiation levels were varied from 12–20% oxygen levels to simulate exhaust from the high pressure turbine (HPT). Experimental results from the ITB at atmospheric pressure indicate that the combustion system operates at 97–99% combustion efficiency over a wide range of operating conditions burning JP-8 +100 fuel. Flame lengths were extremely short, at about 50% of those seen in conventional systems. A wide range of operation is possible with lean blowout fuel-air ratio limits at 25–50% below the value of current systems. These results are significant because the ITB only requires a small (300°F) temperature rise for optimal power extraction, leading to operation of the ITB at near-lean-blowout limits of conventional combustor designs. This data lays the foundation for the design space required for future engine designs.
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James, S., M. S. Anand, and B. Sekar. "Towards Improved Prediction of Aero-Engine Combustor Performance Using Large Eddy Simulations." In ASME Turbo Expo 2008: Power for Land, Sea, and Air. ASMEDC, 2008. http://dx.doi.org/10.1115/gt2008-50199.
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The paper presents an assessment of large eddy simulation (LES) and conventional Reynolds averaged methods (RANS) for predicting aero-engine gas turbine combustor performance. The performance characteristic that is examined in detail is the radial burner outlet temperature (BOT) or fuel-air ratio profile. Several different combustor configurations, with variations in airflows, geometries, hole patterns and operating conditions are analyzed with both LES and RANS methods. It is seen that LES consistently produces a better match to radial profile as compared to RANS. To assess the predictive capability of LES as a design tool, pretest predictions of radial profile for a combustor configuration are also presented. Overall, the work presented indicates that LES is a more accurate tool and can be used with confidence to guide combustor design. This work is the first systematic assessment of LES versus RANS on industry-relevant aero-engine gas turbine combustors.
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Hu, Bin, Yong Huang, and Jianzhong Xu. "A Hybrid Semi-Empirical Model for Lean Blow-Out Limit Predictions of Aero-Engine Combustors." In ASME Turbo Expo 2014: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2014. http://dx.doi.org/10.1115/gt2014-26271.
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Lean blow-out (LBO) is critical to operational performance of combustion systems in propulsion and power generation. Current predictive tools for LBO are based on decades-old empirical correlations that have limited applicability for modern combustor designs. Based on Lefebvre’s model for LBO and flame volume concept, an FV (Flame Volume) model was proposed by Authors in early study. The FV model adds two key parameters of α and β that represent the fraction of dome air and dimensionless flame volume defined as the ratio of flame volume and combustor volume. Due to the flame volume is obtained from the experimental image, FV model could only be used in LBO analysis instead of predictions. In the present study, a hybrid FV model is proposed that combines the FV model with numerical simulation for LBO predictions. In the hybrid FV model, α and β are estimated from the numerical simulation result of the non-reacting flow in the combustor. Comparing with the experimental data for 11 combustors, the LBO fuel/air ratio obtained by hybrid FV model shows better agreement than that obtained by Lefebvre’s model. The maximum prediction uncertainties of hybrid FV model and Lefebvre’s model are about ±16% and ±48%, respectively. Moreover, the time cost of the LBO prediction using hybrid FV model for each case is about 6 hours with the computer equipment of CPU×12 and 24G memory, showing that the hybrid FV model is reliable and efficient to be used for the performance evaluation of the combustor, even the so called “paper combustors” in the primary design stage.
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Rolt, Andrew, Victor Martínez Bueno, Mirko Romanelli, et al. "Numerical Studies of Novel Aero Engine Secondary Combustors for Low-NOx Emissions." In ASME Turbo Expo 2020: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2020. http://dx.doi.org/10.1115/gt2020-16081.
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Abstract Gas turbine thermal efficiency and fuel burn are very dependent on turbine entry temperature and overall pressure ratio (OPR). Unfortunately, increases in these two parameters compromise other key aspects of engine operation and tend to increase emissions of nitrogen oxides (NOx). The European Horizon 2020 ULTIMATE project researched advanced-cycle aero engines with synergistic combinations of novel technologies to increase thermal efficiency without increasing emissions. One candidate technology was the addition of secondary combustion to increase the mean temperature of heat addition to improve thermal efficiency while limiting the primary combustor flame temperatures and NOx formation. However, an overall reduction in NOx also requires the secondary combustor to be a low-NOx design. This paper describes numerical studies carried out on novel aero engine secondary combustor concepts developed in two MSc-thesis research projects. The studies have explored the potential of oxy-poor-flame combustion concepts. These annular combustor designs featured two distinct regions: (i) the vortex zone, which promotes recirculation of combustion products, a prerequisite for low-oxygen combustion, and (ii) a through-flow region where part of the incoming flow bypasses the vortex before the flows mix again. These studies have demonstrated the advantages and some limitations of the proposed designs and emissions assessments in comparison with previous secondary combustor studies. They suggest very low NOx is achievable with oxy-poor combustion, but will be more difficult if the incoming oxygen levels are above 10%. More-accurate assessments will require LES modelling and inclusion of the primary combustor in the simulations. However, if the low overall NOx emissions would include relatively higher levels of nitrous oxide (N2O) then this might raise concerns with respect to global warming.
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Leung, Ho Yin, Efstathios Karlis, Yannis Hardalupas, and Andrea Giusti. "Evaluation of Blow-Off Dynamics in Aero-Engine Combustors Using Recurrence Quantification Analysis." In ASME Turbo Expo 2021: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2021. http://dx.doi.org/10.1115/gt2021-59484.
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Abstract The lean blow-out performance of an engine and the ability to re-ignite the flame, especially at high-altitude conditions, are important aspects for the safe operability of airplanes. The operability margins of the engine could be extended if it was possible to predict the occurrence of flame blowout from in-flight measurements and take actions to dynamically control the flame behaviour before complete extinction. In this work, the use of Re-currence Quantification Analysis (RQA), an established tool for the analysis of non-linear dynamical systems, is explored to reconstruct and study the blow-off dynamics starting from pressure measurements taken from blow-off experiments of an engine rig. It is shown that the dynamics of the combustor exhibit chaotic characteristics far away from blow-off and that the dynamics become more coherent as the blow-off condition is approached. The degree of determinism and recurrence rate are studied during the entire combustor’s dynamics, from stable flame to flame extinction. It is shown that the flame extinction is anticipated by an increase of the degree of determinism and recurrence rate at all investigated conditions, which indicates intermittent behavior of the combustor before the blow-off condition is reached. Therefore, in the configuration investigated here, the determinism and the recurrence rate of the system could be good predictors of blow-off occurrence and could potentially enable control actions to avoid flame extinction. This study opens up new possibilities for engine control and operability. The development of real-time RQA should be addressed in future research.
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Hu, Bin, Yong Huang, Fang Wang, and Fa Xie. "Numerical Simulation of Cold Flow Field of Aero-Engine Combustors for Lean Blow Off Analysis." In ASME 2011 Turbo Expo: Turbine Technical Conference and Exposition. ASMEDC, 2011. http://dx.doi.org/10.1115/gt2011-45467.
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Lean blow off (LBO) performance is critical to the operational performance of combustion system in propulsion and power generation. Current predictive tools for LBO are based on decades-old empirical correlations that have limited applicability for modern combustor designs. Recent advances in computational fluids dynamics (CFD) have provided new insight into the fundamental processes that occur in these flows. In this paper, it is envisaged a new methodology for the LBO predictions that is predicting the LBO fuel/air ratio based on the cold flow field of the combustor. Comparing to the traditional tools, this methodology has the lower prediction cost, especially in the designing stage of the combustor. The study presented here is the preliminary study of this method. According to the Lefebvre’s LBO model, a new load parameter (mr·Vf) extracted from the cold flow field is obtained for LBO analysis. Commercial software FLUENT is used to simulate the velocity and concentration field without combustion in different combustors. LBO fuel/air ratios are obtained from the model combustor experiments. Flammable zone volume (Vf) is used instead of Vc (as defined in Lefebvre’s model: combustor volume ahead of the dilution holes) in this LBO analysis. Vf is defined according to the lean/rich limits and increased with the increase of φLBO. In addition, the mass flow rate of back-flow air which enters the flammable zone (mr) is used to account for the combustion air. φLBO is increased in a parabolic way with the increase of mr. The load parameter (mr·Vf) could represent the actual combustion load of the combustor near LBO and relates φLBO to the cold flow field of the combustor. It will be encouraging and beneficial to the study of LBO prediction in the future.
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Franzelli, B., E. Riber, B. Cuenot, and M. Ihme. "Numerical Modeling of Soot Production in Aero-Engine Combustors Using Large Eddy Simulations." In ASME Turbo Expo 2015: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2015. http://dx.doi.org/10.1115/gt2015-43630.
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Numerical simulations are regarded as an essential tool for improving the design of combustion systems since they can provide information that is complementary to experiments. However, although numerical simulations have already been successfully applied to the prediction of temperature and species concentration in turbulent flames, the production of soot is far from being conclusive due to the complexity of the processes involved in soot production. In this context, first Large Eddy Simulations (LES) of soot production in turbulent flames are reported in the literature in laboratory-scale configurations, thereby confirming the feasibility of the approach. However numerous modeling and numerical issues have not been completely solved. Moreover, validation of the models through comparisons with measurements in realistic complex flows typical of aero-engines is still rare. This work therefore proposes to evaluate the LES approach for the prediction of soot production in an experimental swirl-stabilized non-premixed ethylene/air aero-engine combustor, for which soot and flame data are available. Two simulations are carried out using a two-equation soot model to compare the performance of a hybrid chemical description (reduced chemistry for the flame structure/tabulated chemistry for soot precursor chemistry) to a classical full tabulation method. Discrepancies of soot concentration between the two LES calculations will be analyzed and the sensitivity to the chemical models will be investigated.
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Dauch, Thilo Ferdinand, Samuel Braun, Lars Wieth, et al. "Computational Prediction of Primary Breakup in Fuel Spray Nozzles for Aero-Engine Combustors." In ILASS2017 - 28th European Conference on Liquid Atomization and Spray Systems. Universitat Politècnica València, 2017. http://dx.doi.org/10.4995/ilass2017.2017.4693.
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Primary breakup of liquid fuel in the vicinity of fuel spray nozzles as utilized in aero-engine combustors is numericallyinvestigated. As grid based methods exhibit a variety of disadvantages when it comes to the prediction of multi- phase flows, the ”Smoothed Particle Hydrodynamics“ (SPH)-method is employed. The eligibility of the method to analyze breakup of fuel has been demonstrated in recent publications by Braun et al, Dauch et al and Koch et al [1, 2, 3, 4]. In the current paper a methodology for the investigation of the two-phase flow in the vicinity of fuel spray nozzles at typical operating conditions is proposed. Due to lower costs in terms of computing time, 2D predictions are desired. However, atomization of fluids is inherently three dimensional. Hence, differences between 2D and 3D predictions are to be expected. In course of this study, predictions in 2D and based on a 3D sector are presented.Differences in terms of gaseous flow, ligament shape and mixing are assessed.DOI: http://dx.doi.org/10.4995/ILASS2017.2017.4693
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Venkatesan, Krishna, Arin Cross, and Fei Han. "Acoustic Flame Transfer Function Measurements in a Liquid Fueled High Pressure Aero-Engine Combustor." In ASME Turbo Expo 2022: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2022. http://dx.doi.org/10.1115/gt2022-81769.
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Abstract This paper describes an experimental approach and study of thermo-acoustic flame transfer functions in a high-pressure liquid-fueled rich burn combustor. The presence of high background flame luminosity in high-pressure sooty flame combustors precludes the application of any direct optical flame transfer function method. Instead, an acoustic method based on multiple microphones was employed to characterize the combustor acoustic pressure and velocity responses to acoustic forcing. A high-pressure siren device was employed to acoustically excite the combustor air flow over a broad range of frequencies from 150–1000Hz and modulate the combustor inlet dynamic pressure amplitudes. The acoustic pressures measured from the microphones located upstream and downstream of the flame were processed to obtain swirler impedances and flame transfer functions. Nonlinear behavior of the liquid fuel flame transfer function was studied by systematically varying the siren excitation pressure amplitudes. A parametric study of varying inlet air pressure, inlet air temperature, and thermal power was performed to study the impact of operating conditions on the measured liquid flame transfer function.