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Valenti, Carlo Alberto. "Development of a control system for an aeronautical engine." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2019.
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The goal of this thesis is to develop a control system for an aeronautical engine. In particular, the engine considered is an automotive one which will be adapted for aeronautical purposes. Therefore, the engine control unit is built in Simulink®, starting from a model in LabVIEW®. The purpose of this unit is to control the engine next cycle parameters and its setpoints. Then, in order to test its effectiveness a simplified engine model is built and the whole system is tested, after a first trial of the control system. The results are presented and discussed.
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Moraru, Laurentiu Eugen. "Numerical Predictions and Measurements in the Lubrication of Aeronautical Engine and Transmission Components." University of Toledo / OhioLINK, 2005. http://rave.ohiolink.edu/etdc/view?acc_num=toledo1125769629.
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Exilard, Gorka. "Large-Eddy Simulation of constant volume combustion in a ground-breaking new aeronautical engine." Thesis, Université Paris-Saclay (ComUE), 2018. http://www.theses.fr/2018SACLC082/document.
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Au cours des dernières années, le transport aérien de passagers connaît un développement sans cesse croissant et continue ainsi d’accroire sa contribution aux émissions mondiale de CO2. Par conséquent, un effort commun entre les avionneurs est fait pour diminuer les émissions de CO2 et de polluants. Pour encourager cet effort, les réglementations deviennent de plus en plus drastiques en terme d'émissions et de polluants tels que le CO2, les NOx mais aussi le bruit. Ces nouvelles limitations sont à la fois définies à court et moyen-long termes pour inciter les motoristes à travailler sur les technologies de plus en plus efficientes.Pour concevoir des moteurs toujours plus performants tout en respectant ces réglementations à court terme, les motoristes travaillent sur l'optimisation de leurs technologies conventionnelles, en améliorant des leviers bien identifiés comme l'augmentation du taux de compression. Cependant, cette optimisation des turbomachines actuelles a déjà atteint un niveau de maturité très élevé. Il semble ainsi difficile de continuer indéfiniment leurs optimisations. Par conséquent, pour atteindre les objectifs à moyen-long terme, les motoristes sont dès aujourd'hui en train d'étudier des nouveaux systèmes propulsifs avancés comme les chambres de Combustion à Volume Constant (CVC) qui peuvent accroître le rendement thermique. Contrairement aux chambres de combustion traditionnelles, qui fonctionnent à flux continu, les chambres CVC opèrent de façon cyclique afin de créer un volume constant pendant la phase de combustion et libérer les gaz chauds dans les étages de turbines.Pendant cette thèse, une approche numérique permettant d'évaluer ce type de chambres est développée. Tout l'enjeu est de pouvoir étudier des chambre de combustion intégrant des parties mobiles, qui permettent de créer le volume constant dédié à la combustion tout en évitant les fuites à travers ces systèmes mobiles lors de l'élévation de la pression dans la chambre. Cette modélisation doit aussi prédire correctement les phases transitoires comme l'admission des gaz frais, qui pilote la phase de combustion. Cette étude utilise des objets immergés pour modéliser les parties mobiles. Les objectifs de cette thèse sont de rendre ces objets immergés imperméables et adapter la méthode aux différents modèles utilisés pour étudier les milieux réactifs tels que le modèle de combustion ECFM-LES ou encore l'injection liquide Lagrangienne utilisée pour résoudre l'injection du fuel.Dans cette étude, une nouvelle formulation est développée puis testée sur différents cas tests de plus en plus représentatifs des chambres CVC. Cette approche numérique est ensuite évaluée sur une chambre réel étudiée expérimentalement au laboratoire PPRIRME de Poitiers. Dans cette dernière étude, deux cas non réactifs permettent de comparer les évolutions de pression à deux endroits dans la dispositif expérimental, ainsi que les champs de vitesse au sein de la chambre de combustion, aux simulations réalisées. Pour ces cas complexes, l'utilisation des objets immergés permet de prédire les résultats expérimentaux à un coût attractif.Un des cas non réactif est ensuite carburé et allumé pour confronter l'évolution pression et les champs de vitesse dans la chambre de combustion des résultats numériques obtenus aux mesures expérimentales. L'approche numérique développée a permis d’enrichir les données expérimentales, d'analyser les variabilités cycle-à-cycle rencontrées au banc et d'identifier les leviers qui permettraient d'optimiser ce type d’architectureOver the past few years, aircrafts have become a common means of transport, thus continuously increasing their contribution to global CO2 emissions. Consequently, there is a common effort between aircraft manufacturers to reduce CO2 and pollutant emissions. To encourage this effort, regulations are becoming more and more stringent on the emissions and pollutants like CO2, NOx and noise. These regulations are both defined in the short and medium-long terms to urge aircraft manufacturers to work on more and more efficient technologies.In order to design more efficient engines while respecting the short term objectives, engine manufacturers are working on the improvement of conventional architectures by using well-known levers like the increase of the Overall Pressure Ratio (OPR). However, the optimization of the present turbomachinery has already reached a high level of maturity and it seems difficult to continuously enhance their performances. Consequently, to reach the medium-long term objectives, engine manufacturers are working on new advanced propulsion systems such as the Constant Volume Combustion (CVC) chambers, which can increase the thermal efficiency of the system. Contrary to present turbomachinery which are burning fresh gases continuously, CVC chambers operate cyclically so as to create the constant vessel dedicated to the combustion phase and to expand the burnt gases into turbine stages.In this PhD thesis, a numerical approach is developed to allow the evaluation of these kind of combustors. The challenge is to be able to evaluate CVC chambers by taking into account the moving parts which create the constant volume and avoid mass leakages through these moving parts during the increase of the combustion chamber pressure when the combustion occurs. This approach also has to correctly predict unsteady phases like the intake, which directly controls the combustion process.These moving parts are modeled with a Lagrangian Immersed Boundary (LIB) method .The main goals of this thesis is to make the LIB as airtight as possible and to render this approach compatible with the different models which are adapted to analyse reactive flows such as the ECFM-LES combustion model or Lagrangian liquid injection, used for fuel sprays. In this study, a new formulation is developed and tested on several test cases from very simple ones to cases more representative of CVC chambers.Then, this approach is evaluated on a real chamber experimentally analysed in PPRIME laboratory in Poitiers. Two non-reactive operating points are used to compare the experimental pressure at two positions in the apparatus and the experimental velocity fields in the combustion chamber with the numerical results. In this complex configuration, the LIB method allows the prediction of the experimental results with a low CPU cost. As in the experiment, one non-reactive case is carburized and ignited to compare the measured pressure and the velocity fields in the combustion chamber with the simulations. The proposed numerical approach allows the data enhancement of the experiment and then the analysis of the cycle-to-cycle variability encountered during the experimental measurements. Last but not least, this method enables the identification of the different levers that could decrease the variability and then could improve operability of this type of combustors
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Koukolíček, Ondřej. "Weather and Aeronautical Data on Map for Airplane EFB." Master's thesis, Vysoké učení technické v Brně. Fakulta informačních technologií, 2015. http://www.nusl.cz/ntk/nusl-234990.
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Práce se zabývá webovými knihovnami pro práci s mapou a jejich možným využitím pro implementaci grafického uživatelského rozhraní nativní aplikace Weather Information Service (WIS) společnosti Honeywell. V práci jsou představeny prvky WIS, které je třeba implementovat pro úspěšné převedení aplikace do webového rozhraní. Dále jsou vysvětleny základy mapových knihoven a podrobněji popsány knihovny Leaflet a Altus Map Engine, které byly vybrány pro vytvoření demonstrační aplikace. Jedna kapitola je věnována metodám použitelným pro vykreslování ve webovém prostředí. Práce dále popisuje implementaci demonstrační aplikace, vytvořené za účelem prezentace možností mapových knihoven implementovat prvky WIS. Na závěr jsou diskutovány výhody a nevýhody obou zkoumaných knihoven a jejich použitelnost pro případné využití v aplikaci WIS.
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Notarianni, Gianmarco. "Analysis and modelling of the turbocharger behavior of an internal combustion engine for aeronautical application." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2019.
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The goal of this thesis project is the creation and the validation of an estimation model developed for the prediction of the Wastegate valve behaviour, to define the splitting of mass flow rate between the turbine and the bypass valve in all the operative conditions of the turbocharger system, by an analysis of the experimental data obtained from tests performed in the engine test bench of the University.
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Pavelec, Sterling Michael. "The development of turbojet aircraft in Germany, Britain, and the United States : a multi-national comparison of aeronautical engineering, 1935-1946 /." The Ohio State University, 2004. http://www.ohiolink.edu/etd/send-pdf.cgi/Pavelec%20Sterling%20Michael.pdf?acc_num=osu1082396007.
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Novikov, Yaroslav. "Development Of A High-fidelity Transient Aerothermal Model For A Helicopter Turboshaft Engine For Inlet Distortion And Engine Deterioration Simulations." Master's thesis, METU, 2012. http://etd.lib.metu.edu.tr/upload/12614389/index.pdf.
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Presented in this thesis is the development of a high-fidelity aerothermal model for GE T700 turboshaft engine. The model was constructed using thermodynamic relations governing change of flow properties across engine components, and by applying real component maps for the compressor and turbines as well as empirical relations for specific heats. Included in the model were bleed flows, turbine cooling and heat sink effects. Transient dynamics were modeled using inter-component volumes method in which mass imbalance between two engine components was used to calculate the inter-component pressure. This method allowed fast, high-accuracy and iteration-free calculation of engine states. Developed simulation model was successfully validated against previously published simulation results, and was applied in the simulation of inlet distortion and engine deterioration. Former included simulation of steady state and transient hot gas ingestion as well as transient decrease in the inlet total pressure. Engine deterioration simulations were performed for four different cases of component deterioration with parameters defining engine degradation taken from the literature. Real time capability of the model was achieved by applying time scaling of plenum volumes which allowed for larger simulation time steps at very little cost of numerical accuracy. Finally, T700 model was used to develop a generic model by replacing empirical relations for specific heats with temperature and FAR dependent curve fits, and scaling T700 turbine maps. Developed generic aerothermal model was applied to simulate steady state performance of the Lycoming T53 turboshaft engine.
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Paolucci, Lorenzo. "High efficiency low temperature combustion in compression ignition engines for automotive and aeronautical applications." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2019.
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Internal combustion engines are increasingly subject to ever more stringent and severe Euro-legislations about pollutants emissions in terms of nitrogen oxide, carbon monoxide, unburned hydrocarbons and soot. In last years, thanks to advanced after treatment systems and technological innovations, emission have been improved but, due to even higher costs and complexity of such systems and in a view of further emissions restrictions, advanced combustion methods leading to cleaner and improved efficiency combustion are under investigation. A possible path to follow in order to met requirements on lower emissions, is relative to so called low temperature combustion: a group of innovative combustion methods which by exploiting lean and premixed combustion decreases significantly flame temperature which is mainly responsible for nitrogen oxide production. This work of thesis focus on preliminary study, development and experimental testing of a low temperature combustion strategy, namely "gasoline direct compression ignition" also known as "GDCI".
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Chan, KinYip. "Control algorithms for optimisation of engine combustion process with continuously changing fuel composition." Thesis, Kingston University, 2014. http://eprints.kingston.ac.uk/29888/.
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Current research efforts in the areas of automotive control aim at developing new control algorithms so engine can work efficiently with different fuel compositions. Progressing in this direction, this thesis pioneers and researches novel control strategies to reduce the engine emissions gasses, Carbon Dioxide (CO2), Oxygen (O2), Carbon Monoxide (CO) and Nitric Oxide (NOx), while keeping optimum performance with unknown fuel mixtures. A two-zone engine combustion model is developed and thoroughly validated against the computational data from commercial engine simulation packages. The engine model is suited for the development and testing of control systems. The simulation uses the following fuel mixtures: isooctane (C8-H18), methanol (C1-H4-O1) and ethanol (C2-H6-O1). The results obtained provide better understanding of the control parameters, including fuel-to-air ratio, ignition timing, exhaust-valve timing and intake-valve timing. Moreover, the model facilitates control design. The novel engine controller is studied on the fuel composition as the additional parameter where such parameter has not been widely considered in engine control research. Efficient methodologies to estimate the original fuel composition by using the exhaust gas composition obtained from the engine are proposed and investigated. Two novel approaches based on feed-forward neural network and Adaptive Neuro-Fuzzy Interface System (ANFIS), respectively, are proposed. The portion of mixture for Isooctane-Methanol and Isooctane-Ethanol are effectively calculated. Moreover, results suggest that the feed-forward neural network outperforms the ANFIS approach and that the performance of the fuel estimator is stable in the continuous time process. Further in this research, a Multi-Input-Multi-Output (MIMO) engine control system is developed. The methodology used is a system identification employing a state-space model. In order to reduce the complexity of the state-space model, the developed AI fuel estimator is used to facilitate on the model reduction by feeding gains to controllers for the individual components. In addition, it uses the linear quadric regulator (LQR) method to find the closed-loop gain in the development of the closed-loop control system. The above techniques have been evaluated and results show that the controller is able to identify the minimum levels of the emission gases in terms of CO2, O2, CO and NOx in a continuously changing engine speed.
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Gursoy, Zeynep Ece. "A Numerical Investigation Of Helicopter Flow Fields Including Thermal Effects Of Exhaust Hot Gases." Master's thesis, METU, 2009. http://etd.lib.metu.edu.tr/upload/12611128/index.pdf.
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This thesis investigates the flow field of a twin-engine, medium lift utility helicopter numerically. The effects of the exhaust hot gases emerging from the engines are accounted for in the numerical study. The commercial computational fluid dynamics (CFD) software ANSYS Fluent is employed for the computations. While the effects of engines are included in the computations through simple inlet and outlet boundary conditions, the main and tail rotors are simulated by the Virtual Blade Model in a time-averaged fashion. Forward flight at four different advance ratios and hover in ground effect are studied. The temperature distribution around the tail boom is compared to available flight test data. Good agreement with the flight test data is observed.
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Boiani, Davide. "Finite element structural and thermal analysis of JT9D turbofan engine first stage turbine blade." Bachelor's thesis, Alma Mater Studiorum - Università di Bologna, 2017. http://amslaurea.unibo.it/12566/.
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The objective of this work was to conduct a preliminary finite element static structural and transient thermal analyses of a first stage turbine blade which was previously assembled on a Pratt & Whitney JT9D-7A turbofan engine. This turbine blade was obtained from a collector of aircraft scrap parts. After an extensive theoretical background on airbreathing jet engines and materials used for such components, the process behind the creation of a 3D model was explained. The laser scanning technique and a piezoeletric digitizer were employed to recreate the blade inside a 3D modelling software. The model was then imported into the finite element analysis software ANSYS; the analyses were performed, and the most interesting results were evaluated. The structural and thermal results were found to be congruous with the literature on similar applications of components with the same material, and appear to be a realistic representation of the blade behaviour inside the first stage turbine environment.
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Berger, Sandrine. "Implementation of a coupled computational chain to the combustion chamber's heat transfer." Phd thesis, Toulouse, INPT, 2016. http://oatao.univ-toulouse.fr/16636/1/Berger_Sandrine.pdf.
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The design of aeronautical engines is subject to many constraints that cover performance gain as well as increasingly sensitive environmental issues. These often contradicting objectives are currently being answered through an increase in the local and global temperature in the hot stages of the engine. As a result, the solid parts encounter very high temperature levels and gradients that are critical for the engine lifespan. Combustion chamber walls in particular are subject to large thermal constraints. It is thus essential for designers to characterize accurately the local thermal state of such devices. Today, wall temperature evaluation is obtained experimentally by complex thermocolor tests. To limit such expensive experiments, efforts are currently performed to provide high fidelity numerical tools able to predict the combustion chamber wall temperature. This specific thermal field however requires the consideration of all the modes of heat transfer (convection, conduction and radiation) and the heat production (through the chemical reaction) within the burner. The resolution of such a multi-physic problem can be done numerically through the use of several dedicated numerical and algorithmic approaches. In this manuscript, the methodology relies on a partitioned coupling approach, based on a Large Eddy Simulation (LES) solver to resolve the flow motion and the chemical reactions, a Discrete Ordinate Method (DOM) radiation solver and an unsteady solid conduction code. The various issues related to computer resources distribution as well as the coupling methodology employed to deal with disparity of time and space scales present in each mode of heat transfer are addressed in this manuscript. Coupled application high performance studies, carried out both on a toy model and an industrial burner configuration evidence parameters of importance as well as potential path of improvements. The thermal coupling approach is then considered from a physical point of view on two distinct configurations. First, one addresses the impact of the methodology and the thermal equilibrium state between a reacting fluid and a solid for a simple flame holder academic case. The effect of the flame holder wall temperature on the flame stabilization pattern is addressed through fluid-only predictions. These simulations highlight interestingly three different theoretical equilibrium states. The physical relevance of these three states is then assessed through the computation of several CHT simulations for different initial solutions and solid conductivities. It is shown that only two equilibrium states are physical and that bifurcation between the two possible physical states depends both on solid conductivity and initial condition.Furthermore, the coupling methodology is shown to have no impact on the solutions within the range of parameters tested. A similar methodology is then applied to a helicopter combustor for which radiative heat transfer is additionally considered. Different computations are presented to assess the role of each heat transfer process on the temperature field: a reference adiabatic fluid-only simulation, Conjugate Heat Transfer, RadiationFluid Thermal Interaction and fully coupled simulations are performed. It is shown that coupling LES with conduction in walls is feasible in an industrial context with acceptable CPU costs and gives good trends of temperature repartition. Then, for the combustor geometry and operating point studied, computations illustrate that radiation plays an important role in the wall temperature distribution. Comparisons with thermocolor tests are globally in a better agreement when the three solvers are coupled.
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Hausen, Roberto Begnis. "SISTEMATIZAÇÃO DE CONHECIMENTO PARA O PROJETO O PROJETO DE MOTOR A ETANOL PARA AERONAVE AGRÍCOLA." Universidade Federal de Santa Maria, 2011. http://repositorio.ufsm.br/handle/1/7546.
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Coordenação de Aperfeiçoamento de Pessoal de Nível SuperiorActually in Brasil there there is a fleet around 1500 agricultural aircraft working on different aplications. The spray aplication by na airplane of crop defenses is increasing on the last years, generating a demand on spraying airplanes, pilots, spare parts for maintenance, fuel and, first of all, operational cost reduction. Base on this, the present master thesis is working as knowledge generator ans presenting results for, in the future, to project an ethanol aeronautical engine that will use on agricultural/spraying airplanes. Operational costs for agricultural airplanes is based on maintenance and in the fuel, that the major part uses aviation gasoline AvGas as fuel. Ethanol comes with a great chance to reduces such costs, due to the price of this fuel is only around 25% of AvGas. The aeronautical engines that flies nowadays with ethanol as fuel are a conversions from AvGas engines, wich generates some problems, the main is consumption increasing. The possibility to reduces operational costs with using a fuel from renewable source and challenge to get an aeronautical engine dedicated to run with ethanol as fuel is what motivates and derecting this research. Development process of a technical system, as an ethanol engine, is complex and requires human and material resources administration, information colect, research and a strenght management of whole development process. Methodology for this work development is based on exploratory research of existing bibliography about development product process, technical systems, aeronautical engines, agricultural aviation and homologation standards for aeronautical engines and your parts and components. Information acquired are presented to provide a data base to meet the need of companies, engineers and designers on a future development. After organization and threatment the data supply the base of influence factors for a project of aeronautical ethanol engine for agricultural airplane. The knowledge about brasilian agricultural airplane fleet and engines used on such airplanes provide detailed information about size and capacity of each airplane and their aplications, including information about their technical and operating characteristics. With this, was able to make entire influence factors for a project of aeronautical ethanol engine for agricultural airplane, such influence factors are divided in four groups that is: Project Scope, Benchmarking, Operating Characteristics and Standards & Homologation. Technical data and information of aeronautical engines and their characteristics was evaluated and compared to observ the tendency of engines with power in a range beetween 151 and 300kW, where more than 80% of those engines in whole brasilian fleet is inside of this range, the major part using AvGas as fuel. Engines running with ethanol generates more power if compared the same engine running with AvGas, for other side, there is a consumption increasing, aorund 40% more and the emissions are decreased, what is a positive point with ethanol use as fuel.
No Brasil, hoje, há uma frota aproximada de 1500 aeronaves agrícolas efetuando operações diversas. A aplicação aérea de defensivos nas culturas agrícolas tem sido crescente nos últimos anos, gerando uma necessidade maior de aeronaves, pilotos agrícolas, peças de reposição e combustível, bem como, e principal quesito, a necessidade de redução dos custos operacionais. É neste ponto que este trabalho entra como gerador de conhecimento apresentando dados e resultados para a obtenção, no futuro, de um motor projetado para o uso de etanol equipar aeronaves de aplicação agrícola.Os custos operacionais das aeronaves agrícolas está baseado na manutenção e no combustível que ela utiliza, na grande maioria é a AvGas gasolina de aviação. O etanol vem com grande chance de reduzir significativamente estes custos, pois seu preço é inferior ao da AvGas, podendo chegar apenas 25% do preço desta. Os motores de aeronaves que voam atualmente com etanol são conversões, o que gera alguns problemas, dentre eles o mais significativo que é o aumento de consumo de combustível. A possibilidade de redução dos custos operacionais com a utilização de combustível produzido de fonte renovável e o desafio de se ter um motor projetado especificamente para o uso de etanol como combustível é o que motiva o estudo e direciona a pesquisa. O processo de desenvolvimento do projeto de sistema técnico, como o de um motor a etanol, é complexo e exige administração dos recursos humanos e materiais, coleta de informações, pesquisa e um rígido gerenciamento de todo o processo para tal desenvolvimento. A metodologia para desenvolvimento do estudo está baseada na pesquisa exploratória de bibliografia existente acerca de desenvolvimento de produtos e sistemas técnicos, motores aeronáuticos, aviação agrícola e normas para homologação de produtos e peças aeronáuticas para aplicação em aeronaves agrícolas. As informações obtidas estão apresentadas de maneira a fornecer uma base de dados capaz de suprir a necessidade de empresas, engenheiros e projetistas num desenvolvimento futuro. Os dados compilados e organizados formam a base para a apresentação dos fatores de influência no projeto de um motor a etanol para aplicação em aeronaves agrícolas. O conhecimento sobre a frota brasileira de aeronaves e os motores utilizados fornecem informações detalhadas sobre o tamanho e capacidade de cada aeronave e suas aplicações, bem como os motores necessários em cada uma destas aeronaves e suas características técnicas e operacionais. Obtêve-se a elaboração completa de todos os fatores de influência para o projeto de um motor aeronáutico a etanol direcionado para aplicação agrícola, tais fatores de influência estão subdivididos em quatro grupos que são: Escopo do projeto, análise comparativa dos motores, características da operação e normas e homologação. Informações e dados técnicos de motores e suas características foram avaliadas e comparadas entre si, de maneira a observar a tandência de utilização de motores na faixa de 151 à 300kW, onde mais de 80% dos motores empregados estão dentro desta faixa e na grande maioria consumindo AvGas. Motores a etanol geram maior potência em detrimento de aumento de consumo, da ordem de 40%, porém, as emissões de gases tóxicos são reduzidas, o que é fator positivo no uso do etanol como combustível.
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De, Paoli Matteo. "Validazione preliminare del modello dinamico di un turboalbero aeronautico." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2014. http://amslaurea.unibo.it/6998/.
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Il bisogno di creare dei metodi per l’identificazione delle performance e strumenti di diagnosi è sempre maggiore nelle turbomacchine. Le case costruttrici di motori, in generale, sono guidate dalle richieste del mercato e da normative sempre più severe, e ricercano quindi da un lato, sempre migliori prestazioni come ad esempio la diminuzione dei consumi e, dall’altro, l’abbattimento degli agenti inquinanti. In ambito industriale si ha l’esigenza di rendere più rapidi i tempi di progettazione e sviluppo dei motori per l’abbattimento dei costi di progettazione e per minimizzare il tempo di commercializzazione. Ecco perché entra in gioco, ed assume importanza, l’uso della simulazione numerica che fornisce informazioni utili sia in fase di verifica, che in fase di progetto di tutti i componenti del motore, anche quelli che sarebbero difficilmente accessibili per misure sperimentali. Al contempo i calcolatori moderni diventano facilmente e rapidamente sempre più potenti riducendo così i tempi di calcolo, e ciò ha consentito l’uso di tecniche numeriche che prima sarebbero risultate impensabili da utilizzare. Per mezzo dell’uso di codici di tipo bi o tri-dimensionale è possibile conoscere e valutare i valori delle grandezze termodinamiche del fluido che attraversa il motore punto per punto. Questi metodi presentano un elevata accuratezza ma hanno anche lo svantaggio di richiedere un elevato costo computazionale con conseguente aumento dei tempi di calcolo. In più hanno bisogno di molti dati di input per il modello matematico, e dipendono fortemente dalle condizioni iniziali e dalle condizioni al contorno. Nasce quindi l'esigenza di un ambiente di sviluppo che consenta una modellazione semplificata degli elementi costituenti il motore, ed un rapido interfacciamento con modelli locali più sofisticati. Inoltre, se si vogliono ottimizzare dei parametri relativi all’impianto globale oppure se si desidera avere un quadro generale delle prestazioni del motore è sufficiente usare modelli 0-D e 1-D per i vari componenti del motore. Sono stati svolti molti studi concentrati sullo sviluppo di strumenti capaci di supportare operazioni di pianificazione e/o per validare analisi di diagnosi.
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York, Stephen P. (Stephen Patrick). "Engine placement for manned descent at Mars considering single engine failures." Thesis, Massachusetts Institute of Technology, 2006. http://hdl.handle.net/1721.1/37949.
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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2006.Includes bibliographical references (p. 93-94).
Previous missions to Mars have landed masses of approximately I metric ton on the surface. Vehicles large enough to support humans on the flight to Mars and land them safely on the surface are closer to 100 metric tons, a two order of magnitude increase. This large mass causes many changes in the EDL of a manned vehicle compared to proven unmanned landers. One critical change is the potential choice of a propulsive descent to replace parachute systems that do not scale to large masses. The placement of these engines on a lander is subject to many concerns such as heat shield packing, cargo handling, and engine out mitigation. Engine out mitigation is of considerable interest because configurations that improve failure mitigation tend to be poorer for the other considerations. This thesis presents the development of a simulation of the descent phase of a manned landing at Mars, an overview of the effects of the various requirements on manned lander engine configuration and the results of a 6 DOF analysis of engine failure scenarios.
by Stephen P. York.
S.M.
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Rochette, Bastien. "Modeling and simulation of two-phase flow turbulent combustion in aeronautical engines." Thesis, Toulouse, INPT, 2019. http://www.theses.fr/2019INPT0059.
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De nos jours, plus de 80% de l'énergie consommée sur Terre provient de la combustion de combustibles fossiles. Des solutions alternatives à la combustion sont en cours de développement mais les contraintes spécifiques liées au transport aérien ne permettent pas actuellement d'alimenter des moteurs sans introduire de rupture technologique. Ces résultats expliquent les activités de recherche visant à améliorer les connaissances et le contrôle des processus de combustion afin de concevoir des moteurs aéronautiques plus propres et plus efficaces. Dans ce cadre, les Simulations aux Grandes Echelles ("Large Eddy Simulation" LES) sont devenues un outil puissant pour mieux comprendre les processus de combustion et les émissions de polluants. Cette thèse s'inscrit dans ce contexte et se focalise sur les modèles et stratégies de calcul afin de simuler avec plus de précision les écoulements réactifs turbulents gazeux et diphasiques dans la chambre de combustion des moteurs aéronautiques. Tout d'abord, une méthode générique et automatique pour la détection et l'épaississement du front de flamme a été développée pour le modèle TFLES, et validée pour plusieurs configurations académiques de complexité croissante. Cette approche générique est ensuite évaluée dans une simulation LES d'un brûleur de laboratoire et comparée à la méthode d'épaississement classique. Les résultats montrent un épaississement plus précis dans les régions post-flammes. Dans un second temps, à partir de l'analyse de flammes laminaires 1D diphasiques homogènes où la phase dispersée a une vitesse relative comparée à la phase porteuse, deux formulations analytiques pour la vitesse de propagation de ces flammes ont été proposées et validées. La concordance entre les vitesses de flammes mesurées et estimées démontre que le modèle et ses paramètres prennent correctement en compte les principaux mécanismes physiques contrôlant ces flammes diphasiques. Enfin, les modèles TFLES les plus récents ont été testés sur des configurations de flamme turbulente gazeuse/diphasique complexes. Les avantages et les inconvénients de ces modèles ont été étudiés afin de contribuer à la compréhension des mécanismes liés à la combustion turbulente et de proposer une stratégie de modélisation par LES pour améliorer la fidélité des simulations réactivesNowadays, more than 80% of the energy consumed on Earth is produced by burning fossil fuels. Alternative solutions to combustion are being developed but the specific constraints related to air transport do not make it possible to currently power engines without introducing a technological breakthrough. These findings explain the research activity to improve the knowledge and the control of combustion processes to design cleaner, and more efficient aeronautical engines. In this framework, Large Eddy Simulations (LES) have become a powerful tool to better understand combustion processes and pollutant emissions. This PhD thesis is part of this context and focuses on the models and numerical strategies to simulate with more accuracy turbulent gaseous and two-phase reacting flows in the combustion chamber of aeronautical engines. First, a generic and self-adapting method for flame front detection and thickening has been developed for the TFLES model, and validated on several academic configurations of increasing complexity. This generic approach is then evaluated in the LES of a laboratory-scale burner and compared to the classical thickening method. Results show a more accurate thickening in post-flame regions. Second, from the analysis of 1-D homogeneous laminar spray flames where the dispersed phase has a relative velocity compared to the carrier phase, two analytical formulations for the spray flame propagation speed have been proposed and validated. The agreement between the overall trend of both the measured/estimated spray flame speeds demonstrates that the model and its parameters correctly take into account the main physical mechanisms controlling laminar spray flames. Finally, the state-of-the-art TFLES models were tested on complex turbulent gaseous and two-phase reacting configurations. The pros and cons of these models were investigated to contribute to the understanding of the mechanisms related to turbulent combustion, and to propose a LES modeling strategy to improve the fidelity of reactive simulations
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Mulchandani, Hiten. "An engine air-brake integration study." Thesis, Massachusetts Institute of Technology, 2011. http://hdl.handle.net/1721.1/62880.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2011.This electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections.
Cataloged from student submitted PDF version of thesis.
Includes bibliographical references (p. 109-112).
The feasibility of operating an engine air-brake (EAB) integrated with a pylon duct bifurcation in a realistic aircraft engine environment has been analyzed. The EAB uses variable exit guide vanes downstream of a high bypass ratio (BPR) fan rotor to produce drag quietly by swirling flow out of the fan nozzle. The swirling motion yields low pressure in the vortex core from simple radial equilibrium, thereby generating pressure drag. The 4-BB internal plug and 5-BB external plug nozzles of BPR 8 are chosen to provide a realistic environment for model-scale tests at the NASA Aero-Acoustic Propulsion Lab (AAPL). The objectives of this study are to quantify the impact of a pylon on the drag and noise of an EAB, and explore means to mitigate the potential loss of swirling flow and associated drag. Analysis is conducted at approach conditions on the 4-BB nozzle, with fan and core nozzle pressure and temperature ratios obtained from an engine cycle analysis. A pylon is designed to represent engine installations typically encountered in short-range jet aircraft. The pylon is a prismatic NACA 0012 airfoil geometry with swept leading, trailing edges and an extended internal fairing to facilitate compatibility with both nozzles in the AAPL facility. The EAB cases analyzed include three types of pylon/vane configurations: (1) the baseline pylon with un-deflected swirl vanes is used in the calculation of the equivalent drag coefficient (CD); (2) the pylon with the trailing edge (TE) flap deflected full-span by 35 degrees is used to set structural load limits for detailed design of the baseline pylon; and (3) configurations with the pylon TE flap deflected partial-span by 20 degrees and asymmetric swirl vanes are used to generate swirling outflow from the fan nozzle exhaust. The partial-span deflection cases are further categorized by the location of the asymmetric vanes: at the nozzle exhaust (aft) and further upstream. Computational results demonstrate the aft vanes generate CD in the range 0.35-0.61 and the upstream vane cases produce CD between 0.09-0.18. The difference in drag is because the flow avoids the majority of the duct bifurcation in the aft vanes cases to produce stronger swirling outflow. A CD value between 0.7-1.0 is required to achieve a 3-4 degree glidescope change and therefore an overall noise benefit of 2.5 dB for a conventional tube-and-wing aircraft on approach. The aft vane configurations show promise in reaching this target while the upstream vane installation concepts require further investigation.
by Hiten Mulchandani.
S.M.
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Springmann, Philip N. "Lunar descent using sequential engine shutdown." Thesis, Massachusetts Institute of Technology, 2006. http://hdl.handle.net/1721.1/35563.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2006.Includes bibliographical references (p. 111-113).
The notion of sequential engine shutdown is introduced and its application to lunar descent is motivated. The concept calls for the utilization of multiple fixed thrust engines in place of a single continuously throttleable engine. Downrange position control is provided by properly timed engine shutdowns. The principle advantage offered is the potential cost savings that would result from the elimination of the development cost of a throttleable rocket engine. Past lunar landing efforts are reviewed and provide the foundation for a baseline vehicle definition. A descent from a lunar parking orbit is assumed. The powered descent is divided into two phases, and a sequential engine shutdown-based guidance scheme is developed for the earlier phase. The guidance scheme consists of a biased ignition point and an algorithm for calculating shutdown times combined with a linear tangent steering law to provide full terminal position control. The performance of the sequential engine shutdown guidance scheme is assessed against two alternative approaches.
(cont.) A statistical picture of the performance of each guidance scheme is obtained via Monte Carlo trials of a lunar descent simulation that captures, to first order, the interaction between the descent propulsion system, the navigation filter, and the guidance function, allowing a direct comparison to be made on the basis of accuracy and fuel consumption. The impact of variations in the number of engines available in the sequential engine shutdown case is analyzed. While the performance observed with sequential engine shutdown does not match that observed with a throttleable engine, the results suggest that it is a viable solution to the lunar descent guidance problem.
by Philip N. Springmann.
S.M.
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Jamonet, Laurent 1978. "Testing of a microrocket engine turbopump." Thesis, Massachusetts Institute of Technology, 2002. http://hdl.handle.net/1721.1/8129.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2002.Includes bibliographical references (p. 191-194).
Advances in microfabrication suggest its application to rocket engines. A MEMS thrust chamber producing 50 N of thrust at design point was previously developed that requires propellants pressurized as high as 300 atmospheres. Hence the need for turbopumps at the MEMS scale. A demonstration microturbopump approximately 20x20x6mm in size was designed and built using silicon microfabrication technology. Nitrogen and deionized water are used as operating fluids in the turbine and in the pump respectively. The design speed is 750,000 RPM, with a 23 atmospheres pump pressure rise, and an overall 30% turbomachinery efficiency. This thesis addresses the key points of the turbopump design, modelling, fabrication, and testing. A 3D CFD simulation of the pump was run and performance predicted. Cavitation risk was shown to be small. A fabrication process flow was set up and continuously improved using the feedback from experiments. Non-destructive fabrication inspection methods were introduced. A test rig and a packaging were built, on which 13 turbopumps have been tested, 8 of them spinning. The maximum speed reached was 100,000 RPM without pump loading, and 65,000 RPM with pump loading. Structural concerns have been addressed. Rotordynamics issues have been investigated. Pumping tests were performed and have paved the way toward an effective pressure rise. The innovative rotor arrangement with coplanar pump and turbine was validated. Dual phase operation involving water and nitrogen as running fluids was achieved successfully.
by Laurent Jamonet.
S.M.
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Chati, Yashovardhan Sushil. "Statistical modeling of aircraft engine fuel burn." Thesis, Massachusetts Institute of Technology, 2018. http://hdl.handle.net/1721.1/115658.
Full textAbstract:
Thesis: Ph. D., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2018.Cataloged from PDF version of thesis.
Includes bibliographical references (pages 169-177).
Fuel burn is a key driver of aircraft performance, and contributes to airline costs and aviation emissions. While the trajectory (ground track) of a flight can be observed using surveillance systems, its fuel consumption is generally not disseminated by the operating airline. Emissions inventories and benefits assessment tools therefore need models that can predict the fuel flow rate profile and fuel burn of a flight, given its trajectory data. Most existing fuel burn estimation tools rely on an architecture that is centered around the Base of Aircraft Data (BADA), an aircraft performance model developed by EUROCONTROL. Operational data (including trajectory data) are generally processed in order to generate the inputs needed by BADA, which then provides an estimate of the fuel flow rate and fuel burn. Although a versatile tool that covers a large number of aircraft types, BADA makes several assumptions that are not representative of real-world operations. Consequently, the reliance on BADA results in errors in the fuel burn estimates. Additionally, existing fuel burn modeling tools provide deterministic predictions, thereby not capturing the operational variability seen in practice. This thesis proposes an alternative model architecture that enables the development of data-driven, statistical models of fuel burn. The parameters of interest are the instantaneous fuel flow rate (that is, the mass of fuel consumed per unit time) and the fuel burn (cumulative mass of fuel consumed over a particular phase or the entire trajectory). The new model architecture uses supervised learning algorithms to directly map aircraft trajectory variables to the fuel flow rate, and subsequently, fuel burn. The models are trained and validated using operational data from flight recorders, and therefore reflect real-world operations. A physical understanding of aircraft and engine performance is leveraged for feature selection. An important characteristic of statistical methods is that they provide both estimates of mean values, as well as predictive distributions reflecting the variability and uncertainty. Locally expert models are developed for each aircraft type and for each of the flight phases. The Bayesian technique of Gaussian Process Regression (GPR) is found to be well-suited for modeling fuel burn. The resulting models are found to be significantly better than state-of-the-art aircraft performance models in predicting the fuel flow rate and fuel burn of a trajectory, giving up to a 63% improvement in total airborne fuel burn prediction over the BADA model. Finally, the Takeoff Weight (TOW) of an aircraft is recognized as an important variable for determining the fuel burn. The thesis therefore develops and evaluates a methodology to estimate the TOW of a flight, using trajectory data from its takeoff ground roll. The proposed statistical models are found to result in up to a 76% smaller error than the Aircraft Noise and Performance (ANP) database, which is used currently for TOW estimation.
by Yashovardhan Sushil Chati.
Ph. D.
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Graysmith, J. L. "Using CFD in engine design." Thesis, University of Warwick, 1995. http://wrap.warwick.ac.uk/4252/.
Full textAbstract:
In this thesis the author presents two areas of work; exploring the integration of Computational Fluid Dynamics (CFD) into engine design for Jaguar Cars Ltd and developing a novel 'mesh construction' method for making mesh generation both easy and fast. It is concluded that Jaguar can use CFD in the evaluation stage of the engine design process, although not in the concept stage of design. The CFD predictions are shown to be useful for detecting flow related faults and determining the general flow trends, but they should not be used as an absolute measure of the flow variables. The author has determined an efficient method for obtaining good quality meshes using commercial modelling and mesh generation software which requires a skilled CFD analyst. Steady flow analysis of an engine port and cylinder design could currently be completed in about six weeks using a high-powered workstation. The author recommends dedicated workstations for CFD analysis and training Jaguar's draughtsmen to create CAD models with computer analysis requirements in mind. The author's mesh construction program automatically joins two overlapping meshes or cuts one mesh from another. Whilst the program works well on the test cases considered, it is not at a stage for commercial exploitation. Further development is therefore recommended.
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Seal, David A. (David Allen). "Fault tolerant issues in jet engine compressor control." Thesis, Massachusetts Institute of Technology, 1991. http://hdl.handle.net/1721.1/44277.
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Corn, Brian A. 1971. "Surge dynamics of a helicopter engine gas generator." Thesis, Massachusetts Institute of Technology, 1998. http://hdl.handle.net/1721.1/50328.
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Deux, Antoine 1975. "Design of a silicon microfabricated rocket engine turbopump." Thesis, Massachusetts Institute of Technology, 2001. http://hdl.handle.net/1721.1/8714.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2001.Includes bibliographical references (p. 175-178).
The advances in silicon microfabrication technology suggest the feasibility of high-precision mechanical devices for power conversion. This thesis describes the design of a silicon demonstration turbopump for a micro-rocket engine, and the analysis and experimental investigation of liquid bearings that may be implemented in a future turbopump. Liquid micro-scale lubrication is investigated. Models are developed to predict the performance of hydrostatic liquid thrust bearings, and hydrostatic and hydrodynamic liquid journal bearings. These models suggest that liquid operation of the existing micro-bearing rig is feasible. This device was tested with water to assess the bearings performance. The maximum speed achieved was 21,000 revolutions per minute, and was limited by the drag in this device designed for gas operation. A micro-scale turbopump producing a pressure rise of 30 atm for water was designed, as a demonstration of this concept for fluid pressurization in the rocket engine system. This thesis addresses several of the key design trades and identifies the fundamental engineering issues. This micropump integrates high-speed turbomachinery and micro-gas bearings. An innovative arrangement is proposed with coplanar pump and turbine for ease of fabrication and reduction of imbalance.
by Antoine Deux.
S.M.
25
Robinson, A. J. "Petrol engine development strategy : executive summary." Thesis, University of Warwick, 2000. http://wrap.warwick.ac.uk/66665/.
Full textAbstract:
The automotive industry is becoming more global and cosmopolitan, while the markets are becoming fragmented and differentiated because of the sophistication of the customer. This requires the automotive manufacturers to have a product well suited to these factors more quickly and cheaply, which necessitates the accurate definition of the product. Two European automotive manufacturers NPI processes were studied and this showed that poor performance in the product definition phase was deemed to be a major reason for sub-optimal performance in both companies, even though the approaches were very different. It was therefore decided to develop and apply tools to assist in the development of petrol engines to overcome these deficiencies. When these shortcomings are considered in the context of petrol engine development it can be seen through examination of the literature and industry that: • There is no method for the translation of the values of the company or product into tangible engineering terms and in the context of this project, with focus on vehicle and engine performance. This can result in the poor positioning of a vehicle in the market due to inappropriate characteristics • Having defined the required vehicle performance, it is not possible to demonstrate the feel of the vehicle until the hardware is physically available. This means that often there is a long lead-time between the setting of a target and the concept ratification through driving the new vehicle. This often results in a point of no (or very costly) return very early during the programme. • Fuel economy is becoming an increasingly significant issue with the introduction of fiscal penalties for poor fuel consumption vehicles. There are currently no processes available for the calculation of steady state or drive cycle fuel economy which allow for the accurate modelling to include combustion, pumping and friction losses, and the control of the engine with the engine management system To overcome these limitations three main groups of innovative tools/techniques have been developed and applied on new engine and vehicle programmes. • Marque engineering: a proposal to translate the brand objectives into engineering terms has been developed to a level where the engine torque output can be linked to the product position • Engine and vehicle performance simulation: a vehicle has been built to demonstrate how a new vehicle/engine will feel to drive by controlling the original engine performance to allow the simulation of the concept engine performance in the concept vehicle. This means it is possible to demonstrate and ratify a given engine/vehicle performance based on experience of the simulated product and to conduct sensitivity studies to discreet aspects of the performance feel of the vehicle • Steady state and drive cycle fuel economy: simulation programs have been written that take the basic engine efficiency relationships and through manipulation it is possible to determine the exact operating point of the engine, steady state or transient, and then determine the fuel economy. In order to determine the exact operating point it is necessary to consider detailed component data and key calibration data. Therefore it is possible to understand the effects of small changes to engine geometry, components or calibration on the fuel used. These developed techniques have been compared with traditional methods to determine the benefits in the concept confirmation phase of a new programme. This investigation showed that the reductions in phase duration, resource requirements and cost could be achieved in the order of 49, 27 and 17% respectively. This is coupled with the ability to obtain a more accurate product positioning through the capacity precisely to predict the product attributes. To conclude, the ability to define the product well is paramount to the success of an automotive manufacturer, and in support of this my project has developed tools and techniques that will greatly assist the development of petrol engines.
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Jensen, Jonathan Andrew. "Robust discrete estimation of the space shuttle main engine." Thesis, Massachusetts Institute of Technology, 1996. http://hdl.handle.net/1721.1/49608.
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Greenman, Matthew David. "Design and construction of a miniature internal combustion engine." Thesis, Massachusetts Institute of Technology, 1996. http://hdl.handle.net/1721.1/10829.
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York, Martin A. S. M. Massachusetts Institute of Technology. "Turbofan engine sizing and tradeoff analysis via signomial programming/." Thesis, Massachusetts Institute of Technology, 2017. http://hdl.handle.net/1721.1/112383.
Full textAbstract:
Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2017.This electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections.
Cataloged from student-submitted PDF version of thesis.
Includes bibliographical references (pages 79-80).
This thesis presents a full 1D core+fan flowpath turbofan optimization model, based on first principles, and meant to be used during aircraft conceptual design optimization. The model is formulated as a signomial program, which is a type of optimization problem that can be solved locally using sequential convex optimization. Signomial programs can be solved reliably and eciently, and are straightforward to integrate with other optimization models in an all-at-once manner. To demonstrate this, the turbofan model is integrated with a simple commercial aircraft sizing model. The turbofan model is validated against the Transport Aircraft System OPTimization turbofan model as well as two Georgia Tech Numerical Propulsion System Simulation turbofan models. Four integrated engine/aircraft parametric studies are performed, including a 2,460 variable multi-mission optimization that solves in 28 seconds.
by Martin A. York.
S.M.
29
Groshenry, Christophe. "Preliminary design study of a micro-gas turbine engine." Thesis, Massachusetts Institute of Technology, 1995. http://hdl.handle.net/1721.1/10386.
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Liu, Chunmeni 1970. "Dynamical system modeling of a micro gas turbine engine." Thesis, Massachusetts Institute of Technology, 2000. http://hdl.handle.net/1721.1/9249.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2000.Also available online at the MIT Theses Online homepage
Includes bibliographical references (p. 123).
Since 1995, MIT has been developing the technology for a micro gas turbine engine capable of producing tens of watts of power in a package less than one cubic centimeter in volume. The demo engine developed for this research has low and diabtic component performance and severe heat transfer from the turbine side to the compressor side. The goals of this thesis are developing a dynamical model and providing a simulation platform for predicting the microengine performance and control design, as well as giving an estimate of the microengine behavior under current design. The thesis first analyzes and models the dynamical components of the microengine. Then a nonlinear model, a linearized model, and corresponding simulators are derived, which are valid for estimating both the steady state and transient behavior. Simulations are also performed to estimate the microengine performance, which include steady states, linear properties, transient behavior, and sensor options. A parameter study and investigation of the startup process are also performed. Analysis and simulations show that there is the possibility of increasing turbine inlet temperature with decreasing fuel flow rate in some regions. Because of the severe heat transfer and this turbine inlet temperature trend, the microengine system behaves like a second-order system with low damping and poor linear properties. This increases the possibility of surge, over-temperature and over-speed. This also implies a potentially complex control system. The surge margin at the design point is large, but accelerating directly from minimum speed to 100% speed still causes surge. Investigation of the sensor options shows that temperature sensors have relatively fast response time but give multiple estimates of the engine state. Pressure sensors have relatively slow response time but they change monotonically with the engine state. So the future choice of sensors may be some combinations of the two. For the purpose of feedback control, the system is observable from speed, temperature, or pressure measurements. Parameter studies show that the engine performance doesn't change significantly with changes in either nozzle area or the coefficient relating heat flux to compressor efficiency. It does depend strongly on the coefficient relating heat flux to compressor pressure ratio. The value of the compressor peak efficiency affects the engine operation only when it is inside the range of the engine operation. Finally, parameter studies indicate that, to obtain improved transient behavior with less possibility of surge, over-temperature and over-speed, and to simplify the system analysis and design as well as the design and implementation of control laws, it is desirable to reduce the ratio of rotor mechanical inertia to thermal inertia, e.g. by slowing the thermal dynamics. This can in some cases decouple the dynamics of rotor acceleration and heat transfer. Several methods were shown to improve the startup process: higher start speed, higher start spool temperature, and higher start fuel flow input. Simulations also show that the efficiency gradient affects the transient behavior of the engine significantly, thereby effecting the startup process. Finally, the analysis and modeling methodologies presented in this thesis can be applied to other engines with severe heat transfer. The estimates of the engine performance can serve as a reference of similar engines as well.
by Chunmei Liu.
S.M.
31
London, Adam Pollok. "Development and test of a microfabricated bipropellant rocket engine." Thesis, Massachusetts Institute of Technology, 2000. http://hdl.handle.net/1721.1/9293.
Full textAbstract:
Thesis (Ph.D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2000.Includes bibliographical references (p. 253-255).
The development of high aspect-ratio, high precision micromachining in silicon or silicon carbide suggests the feasibility of rnicrofabricated, high chamber pressure chemical rocket engines. Such an engine, approximately 20x15x3 mm in size, would produce about three pounds of thrust using 300 sec I sp propellants. As part of the present work, the feasibility of these engines has been investigated and a liquid-cooled, pressure-fed thrust chamber has been designed, fabricated, and tested to evaluate the feasibility of the concept. The results of the tests to date using oxygen and methane as propellants support the feasibility of the concept, producing a maximum thrust of 1 N at a chamber pressure of 12 atm. Given the 1.2 gram mass of the thrust chamber, this corresponds to a thrust-to-weight ratio of 85:1. The characteristic exhaust velocity, c*, a measure of combustion effectiveness, appears to be nearly independent of chamber pressure, indicating that chemical reaction rates are not limiting the combustion. Additionally, when effects of chamber heat loss are included, c* appears to approach its predicted ideal value, indicating that the transport and mixing of propellants in the combustion chamber is of the right order to provide for complete combustion. The thrust chamber was fabricated by etching the required patterns into each side of six 0.5 mm thick silicon wafers, and then diffusion bonding the six wafers together to create the one-piece thrust chamber. A packaging technique is presented to interface high pressure and high temperature fluids to the silicon rocket engine chip. Additionally, initial modelling work has lead to the development of a methodology for mapping the feasible design space of microrocket engines, and for optimizing the performance of such systems given current limitations in microfabrication technology.
by Adam Pollok London.
Ph.D.
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Wang, Vincent S. M. Massachusetts Institute of Technology. "Characterization of cavitation instabilities in rocket engine turbopump inducers." Thesis, Massachusetts Institute of Technology, 2016. http://hdl.handle.net/1721.1/107057.
Full textAbstract:
Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2016.Cataloged from PDF version of thesis.
Includes bibliographical references (pages 135-138).
Characterized by super-synchronous rotation of cavities around the periphery of rocket engine turbopump inducers, rotating cavitation is the primary cavitation instability considered in this thesis. A recently developed hypothesis for rotating cavitation onset is assessed through novel experimental analysis and a previously developed body force modeling approach using the MIT inducer, representative of the design of the Space Shuttle main engine low-pressure oxidizer pump inducer. A previously developed temporal and spatial Fourier decomposition, known as Traveling Wave Energy (TWE) analysis, of experimental unsteady inlet pressure measurements of the cavitating MIT inducer is demonstrated. TWE analysis offers several advantages over the current experimental analysis methods, resolving frequency, spatial mode shapes, and rotation direction of cavitation phenomena. Cut-on/cut-off behavior between rotating cavitation and alternate blade cavitation is observed, supporting the hypothesis that alternate blade cavitation is a necessary precursor to rotating cavitation onset. TWE is adapted for use on high speed borescope video data taken in the same experimental campaign. The frequency content extracted is qualitatively correlated with the results from the pressure data, establishing TWE as a viable tool for quantitative analysis of optical data. The video TWE results indicate that cavitation instability signatures are uniform in the radial direction, suggesting that a pressure transducer array can be established as the primary detection method for rotating cavitation and thereby simplifying test setups. A body force based modeling approach typically used for aero-engine compressor stability prediction is assessed for use in predicting rotating cavitation. A previously developed inducer-specific body force model formulation is validated in a representative compressor geometry, capturing global performance across the characteristic within 7%. However, the model exhibits convergence issues when applied to the inducer, hypothesized to be due to sensitivity in the inducer's loss characteristics. The investigation suggests the low flow coefficient design of the inducer drives the loss sensitivity and is the root cause behind the model's convergence issues. The results indicate the body force model is valid for the higher flow coefficient designs and lower stagger angles typically found in aero-engine compressors and fans. Suggestions for desensitizing the model for the inducer as well as further diagnostics defining the limiting geometry case for body force modeling are made.
by Vincent Wang.
S.M.
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Savoulides, Nicholas 1978. "Development of a MEMS turbocharger and gas turbine engine." Thesis, Massachusetts Institute of Technology, 2004. http://hdl.handle.net/1721.1/17815.
Full textAbstract:
Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2004.Includes bibliographical references.
As portable electronic devices proliferate (laptops, GPS, radios etc.), the demand for compact energy sources to power them increases. Primary (non-rechargeable) batteries now provide energy densities upwards of 180 W-hr/kg, secondary (rechargeable) batteries offer about 1/2 that level. Hydrocarbon fuels have a chemical energy density of 13,000-14,000 W-hr/kg. A power source using hydrocarbon fuels with an electric power conversion efficiency of order 10% would be revolutionary. This promise has driven the development of the MIT micro gas turbine generator concept. The first engine design measures 23 x 23 x 0.3 mm and is fabricated from single crystal silicon using MEMS micro-fabrication techniques so as to offer the promise of low cost in large production. This thesis describes the development and testing of a MEMS turbocharger. This is a version of a simple cycle, single spool gas turbine engine with compressor and turbine flow paths separated for diagnostic purposes, intended for turbomachinery and rotordynamic development. The turbocharger design described herein was evolved from an earlier, unsuccessful design (Protz 2000) to satisfy rotordynamic and fabrication constraints. The turbochargers consist of a back-to-back centrifugal compressor and radial inflow turbine supported on gas bearings with a design rotating speed of 1.2 Mrpm. This design speed is many times the natural frequency of the radial bearing system. Primarily due to the exacting requirements of the micron scale bearings, these devices have proven very difficult to manufacture to design, with only six near specification units produced over the course of three years. Six proved to be a small number for this development program since these silicon devices are brittle
(cont.) and do not survive bearing crashes at speeds much above a few tens of thousands of rpm. The primary focus of this thesis has been the theoretical and empirical determination of strategies for the starting and acceleration of the turbocharger and engine and evolution of the design to that end. Experiments identified phenomena governing rotordynamics, which were compared to model predictions. During these tests, the turbocharger reached 40% design speed (480,000 rpm). Rotordynamics were the limiting factor. The turbomachinery performance was characterized during these experiments. At 40% design speed, the compressor developed a pressure ratio of 1.21 at a flow rate of 0.13 g/s, values in agreement with CFD predictions. At this operating point the turbine pressure ratio was 1.7 with a flow rate of 0.26 g/s resulting in an overall spool efficiency of 19%. To assess ignition strategies for the gas turbine, a lumped parameter model was developed to examine the transient behavior of the engine as dictated by the turbomachinery fluid mechanics, heat transfer, structural deformations from centrifugal and thermal loading and rotordynamics. The model shows that transients are dominated by three time constants - rotor inertial (10⁻¹ sec), rotor thermal (lsec), and static structure thermal (10sec). The model suggests that the engine requires modified bearing dimensions relative to the turbocharger and that it might be necessary to pre-heat the structure prior to ignition ...
by Nicholas Savoulides.
Ph.D.
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Protz, Christopher S. (Christopher Stephen) 1977. "Systems analysis of a microfabricated storable bipropellant rocket engine." Thesis, Massachusetts Institute of Technology, 2000. http://hdl.handle.net/1721.1/9297.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2000.Includes bibliographical references (leaves 95-97).
This thesis discusses the systems analysis of a storable bipropellant micro-rocket engine. Micro-rockets are built using MEMS technology and are projected to deliver a thrust to weight ratio up to two orders of magnitude greater than conventional rocket motors at small thrust levels making them very attractive for satellite propulsion applications and propulsion of very small launch vehicles. Several propellant combinations and engine cycles have been analyzed. Propellant combinations have been evaluated for a 125 atm combustion chamber on the basis of their performance, handling, and cooling properties. A non-toxic hydrogen peroxide/ethanol combination (302 s Isp) is chosen over a nitrogen tetroxide/hydrazine combination (322 s Isp) and a hydrogen peroxode/JP-7 combination (315 s Isp) on the basis of handling and cooling properties. Studies indicate that nitrogen tetroxide/hydrazine expander and decomposition topping cycles and hydrogen peroxide/ethanol and hydrogen peroxide/JP-7 decomposition topping cycles are feasible. It is shown that a 300 atm turbopump feed system is possible while further investigations on bearings and pump design are required to fully validate the concept. The analysis suggests that the storable bipropellant micro-rocket engine concept is feasible and identifies the engineering challenges ahead.
by Christopher S. Protz.
S.M.
35
Al-Midani, Omar M. (Omar Mouaffak) 1974. "Preliminary design of a liquid bipropellant microfabricated rocket engine." Thesis, Massachusetts Institute of Technology, 1998. http://hdl.handle.net/1721.1/9620.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1998."June 1998."
Includes bibliographical references (p. 135-137).
This thesis discusses the design of a microfabricated rocket engine. Micro-rockets feature a thrust to weight ratio up to two orders of magnitude greater than conventional rocket motors at small thrust levels and hence are very attractive for satellite propulsion applications and micro-satellite development. All major rocket components have been characterized and evaluated for micro-scale operation. These include a 300 atm pumping system, a 3000 K and 125 atm combustion chamber and a Mach 3.5 thrusting nozzle. Studies indicate that a turbopump system is feasible while further investigations on bearings are required to fully validate the concept. The viability of the combustion chamber is believed to be dependent on the mixing performance of an innovative injection scheme which features inter-digitated fuel/oxidizer jets impinging at a 180° angle. The nozzle is projected to perform satisfactorily, incurring a mere 2% loss in thrust according to 2D CFD calculations. Modeling of the system transients has indicated an acceleration time on the order of 0.1 sec. as well as notable sensitivities to the injector diameter and turbine blade turning angles. The analysis suggests that the micro-rocket engine concept is feasible and identifies the engineering challenges ahead.
by Omar M. Al-Midani.
S.M.
36
CATY, Fabien. "DESIGN OF SECONDARY AIR SYSTEM AND THERMAL MODELS FOR TRIPLE SPOOL JET ENGINES." Thesis, KTH, Kraft- och värmeteknologi, 2012. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-98875.
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This master thesis deals with the understanding of the secondary air system of athree spool turbofan. The main purpose is the creation of secondary air systemand thermal models to evaluate the behavior of this kind of engine architectureand estimate the pros and cons in comparison with a typical two spool turbofan. Afinite element model of the secondary air system of the engine has been designedbased on the experience of typical jet engines manufactured by Snecma. Theinner thermodynamic pattern and mass flow rates of the engine were obtained.Some local improvements were then made by making analogies with the enginesmanufactured by Snecma. After having communicated the results to theperformance unit to get updates thermodynamic cycles, a quite reliable model wasobtained and
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Joppin, Carole 1979. "Cooling performance of storable propellants for a micro rocket engine." Thesis, Massachusetts Institute of Technology, 2002. http://hdl.handle.net/1721.1/8130.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2002.Includes bibliographical references (p. 211-213).
This thesis studies the selection of propellants for a liquid regeneratively cooled micro rocket engine focusing on the characterization of their cooling performance. Propellants will be at high pressures and under high heat fluxes in the cooling passages and will be supercritical. A summary of the propellant combination selection process and a brief evaluation of potential propellants are presented. A series of heat transfer tests in electrically heated stainless steel micro tubes 95 microns inner diameter has been conducted with two hydrocarbons JP7 and JP10 at subcritical, critical and supercritical conditions and under high heat fluxes. JP7 and JP10 have been evaluated on the basis of their heat transfer capabilities, their stability and the formation of deposits in micro channels. JP7 offers a high heat capacity. An increase in the heat transfer coefficient at the end of the tube, combined with an increase in the Stanton number, seems to indicate that JP7 undergoes an endothermic decomposition which causes a significant enhancement in heat transfer capacity. JP10 offers lower heat transfer coefficients. Both hydrocarbons show a good stability and no evidence of deposits has been seen. Previous results with supercritical ethanol were compared to the results with JP7 and JP10. JP7 seems to provide the highest heat transfer coefficients at high pressures and seems to be the most promising coolant for the regeneratively cooled rocket engine. Compatibility issues associated with the use of hydrogen peroxide as oxidizer for the liquid rocket engine have been addressed. Materials used in MEMS devices show good compatibility with 98 % hydrogen peroxide after passivation in 30 % hydrogen peroxide except for platinum.
by Carole Joppin.
S.M.
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Tsay, Michael Meng-Tsuan. "Two-dimensional numerical modeling of Radio-Frequency ion engine discharge." Thesis, Massachusetts Institute of Technology, 2010. http://hdl.handle.net/1721.1/62713.
Full textAbstract:
Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2010.Cataloged from PDF version of thesis.
Includes bibliographical references (p. 106-109).
Small satellites are gaining popularity in the space industry and reduction in spacecraft size requires scaling down its propulsion system. Low-power electric propulsion poses a unique challenge due to various scaling penalties. Of high-performance plasma thrusters, the radio-frequency ion engine is most likely to succeed in scaling as it does not require an externally applied magnetic field and is structurally simple to construct. As part of a design package an original two-dimensional simulation code for radio-frequency ion engine discharge is developed. The code models the inductive plasma with fluid assumption and resolves the electromagnetic wave in the time domain. Major physical effects considered include magnetic field diffusion and coupling, plasma current induction and ambipolar plasma diffusion. The discharge simulation is benchmarked with data from an experimental thruster. It shows excellent performance in predicting the load power and the internal power loss of the plasma. Predictability of anode current depends on the operating power but is generally adequate. Optimum skin depth on the order of half of chamber radius is suggested by the simulation. The code also demonstrates excellent scaling ability as it successfully predicts the performance of a smaller thruster with errors less than 10%. Using the code a brief optimization study was conducted and the results suggest the maximum thrust efficiency does not necessarily occur at the same frequency that maximizes the power coupling efficiency of the matching circuit.
by Michael Meng-Tsuan Tsay.
Ph.D.
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Powell, Ricardo F. (Ricardo Fernandez) 1973. "Impact of parametric aerothermal variability on aircraft engine operating cost." Thesis, Massachusetts Institute of Technology, 2002. http://hdl.handle.net/1721.1/82257.
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Cali, Philip Mark. "Implicit Euler calculation of supersonic vortex wake/engine plume interaction." Thesis, Massachusetts Institute of Technology, 1992. http://hdl.handle.net/1721.1/43266.
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Kaiser, Sascha [Verfasser], Mirko [Akademischer Betreuer] Hornung, Mirko [Gutachter] Hornung, and Anders [Gutachter] Lundbladh. "Multidisciplinary Design of Aeronautical Composite Cycle Engines / Sascha Kaiser ; Gutachter: Mirko Hornung, Anders Lundbladh ; Betreuer: Mirko Hornung." München : Universitätsbibliothek der TU München, 2020. http://d-nb.info/1204199965/34.
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Cellier, Antony Hermann Guy. "Detection and Identification of Instability and Blow-off/Flashback Precursors in Aeronautical Engines using Deep Learning techniques." Thesis, KTH, Kraft- och värmeteknologi, 2020. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-272079.
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The evolution of injection processes toward more fuel efficient and less polluting combustion systems tend to make them more prone to critical events such as Thermo-Acoustic Instabilities, Blow-Off and Flash-Back. Moreover, the addition of Di-Hydrogen as a secondary or as the main fuel is in discussion by aeronautical engines manufacturers. It drastically modifies the stability of the system and thus raise several interrogations concerning the multiplicity of its use. Being able to predict critical phenomena becomes a necessity in order to efficiently operate a system without having to pre-test every configuration and without sacrificing the safety of the user. Based on Deep Learning techniques and more specifically Speech Recognition, the following study presents the steps to develop a tool able to successfully detect and translate precursors of instability of an aeronautical grade swirled injector confined in a tubular combustion chamber. The promising results obtained lead to proposals for future transpositions to real-size systems.Utvecklingen av injektionsprocesser mot mer bränsleeffektiva och mindre förorenande förbränningssystem, tenderar att göra dem mer benägna att utsättas för kritiska händelser som Thermo-Acoustic Instabilities, Blow-Off och Flash-Back. Dessutom diskuterar flygmotorkonstruktörer möjligheten att använda Dihydrogen som sekundärt eller som huvudbränsle. Det modifierar drastiskt systemets stabilitet och det väcker frågan hur man kan använda det effektivt. Att kunna förutsäga kritiska fenomen blir en nödvändighet för att använda ett system utan att behöva att på förhand testa varje konfiguration och utan att reducera användarens säkerhet. Baserat på Deep-Learning-tekniker och Speech-Recognition-tekniker, presenterar följande studie stegen för att utveckla ett verktyg som kan upptäcka och översätta föregångare till instabilitet hos en swirled flygmotorerinsprutningspump som är innesluten i en förbränningskammare. De lovande resultaten leder till idéer om hur man kan anpassa det här verktyg till ett system i verklig storlek.
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Wong, J. S. "Whole life cost methods for aero-engine design." Thesis, University of Southampton, 2012. https://eprints.soton.ac.uk/360369/.
Full textAbstract:
This research was motivated by the move of aero-engine manufacturers to provide services as well as products. With leasing arrangements such as TotalCare©, the aero-engine manufacturers are responsible for the operation, management and maintenance of their products while airlines pay a contracted rate for their use. As a result, aero-engine manufacturers need to minimise the cost their products incur over their lifecycle to increase profits. It is widely accepted that the greatest scope to reduce costs is in the design stage. Hence, the aim of this thesis is to create tools and methodologies for designers which will allow them to monitor the impact of design decisions on whole life cost. Two different approaches to designing for whole life cost were presented. The Life Cycle Cost (LCC) approach and the comparatively novel Value Driven Design (VDD) approach. It was observed from research literature that models for both LCC and VDD need to be tailored to specific objectives in order to keep the scope of the model manageable. This makes generic LCC or VDD models unfeasible and consequently the reuse of these models is limited. With this in mind, a methodology was developed for creating integrated analyses models which were customisable, modular and transparent so as to facilitate future modification and reuse. It used a commercial software integration package called Isight, and modular analyses modules. Model Based Systems Engineering (MBSE) was also used in the development of the integrated model architecture. Case studies were performed for both LCC and VDD approaches to allow comparison of their respective merits and flaws. Finally several avenues of future work in VDD and MBSE were discussed.
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Sorensen, William Alarik. "A body force model for cavitating inducers in rocket engine turbopumps." Thesis, Massachusetts Institute of Technology, 2014. http://hdl.handle.net/1721.1/93771.
Full textAbstract:
Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2014.This electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections.
Cataloged from student-submitted PDF version of thesis.
Includes bibliographical references (pages 111-113).
Modern rocket engine turbopumps utilize cavitating inducers to meet mass and volume requirements. Rotating cavitation and higher order cavitation instabilities have frequently been observed during inducer testing and operation and can cause severe asymmetric loading on the inducer blades and shaft, potentially leading to failure of the inducer. To date no broadly applicable design method exists to characterize and suppress the onset of cavitation instabilities. This thesis presents the development of a body force model for cavitating inducers with the goal of enabling interrogation of the onset of rotating cavitation and higher order cavitation instabilities and characterization of the governing uid dynamic mechanisms. Building on body force models of gas turbine compressors for compressor stability, the model introduces an additional force component, the binormal force, to capture the strong radial flows observed in inducer ow fields. The body forces were defined and the methodology was successfully validated for two test inducers, a helical inducer and a more advanced design resembling the Space Shuttle Main Engine Low Pressure Oxidizer Pump. The head rise characteristic of each test inducer was captured with less than 4% error across the operating range and the extent of the upstream backflow region was predicted to within 18% at every operating condition. Several challenges with the blade passage model were encountered during the course of the research and the diagnostics performed to investigate them are detailed. An extension of the body force model to two-phase flows was formulated and preliminary calculations with the extended model are presented. The preliminary two-phase results are encouraging and pave the way for future assessment of rotating cavitation instabilities.
by William Alarik Sorensen.
S.M.
45
Bell, Jabin Todd. "Measurements of forced and unforced aerodynamic disturbances in a turbojet engine." Thesis, Massachusetts Institute of Technology, 1993. http://hdl.handle.net/1721.1/46423.
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Pantalone, Giulia Bissinger. "Development of an engine model for an integrated aircraft design tool." Thesis, Massachusetts Institute of Technology, 2015. http://hdl.handle.net/1721.1/98811.
Full textAbstract:
Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2015.Cataloged from PDF version of thesis.
Includes bibliographical references (pages 101-103).
This thesis describes the development of a new engine weight surrogate model and High Pressure Compressor (HPC) polytropic efficiency correction for the propulsion module in the Transport Aircraft OPTtimization (TASOPT) code. The goal of this work is to improve the accuracy and applicability of TASOPT in conceptual design of advanced technology, high bypass ratio, small-core, geared and direct-drive turbofan engines. The engine weight surrogate model was built as separate engine component weight surrogate models using least squares and Gaussian Process regression techniques on data data generated from NPSS/WATE++ and then combined to estimate a "bare" engine weight-including only the fan, compressor, turbine, and combustor-and a total engine weight, which also includes the nacelle, nozzle, and pylon. The new model estimates bare engine weight within +/-10% of published values for seven existing engines, and improves TASOPT's accuracy in predicting the geometry, weight, and performance of the Boeing 737-800. The effects of existing TASOPT engine weight models on optimization od D8-series aircraft concepts are also discussed. The HPC polytropic efficiency correction correlation, which reduces user-input HPC polytropic efficiency based on compressor exit corrected mass flow, was implemented based on data from Computational Fluid Dynamics (CFD). When applied to TASOPT optimization studies of three D8-series aircraft, the efficiency correction drives the optimizer to increase engine core size.
by Giulia Bissinger Pantalone.
S.M.
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Chan, Nicholas Y. S. "Scaling considerations for small aircraft engines." Thesis, Massachusetts Institute of Technology, 2008. http://hdl.handle.net/1721.1/45236.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2008.Includes bibliographical references (p. 81-84).
Small aircraft engines traditionally have poorer performance compared to larger engines, which until recently, has been a factor that outweighed the aerodynamic benefits of commoditized and distributed propulsion. Improvements in the performance of small engines have, however, prompted another look at this old concept. This thesis examines aspects of aircraft engines that may have application to commodity thrust or distributed propulsion applications. Trends of engine performance with size and time are investigated. These trends are further extended to justify parameter choices for conceptual engines of the current, mid-term (10 years) and far-term (20 years). Uninstalled and installed performances are evaluated for these engines, and parametric studies are performed to determine the most influential and limiting factors. It is found that scaling down of engines is detrimental to SFC and fuel burn, mainly due to the Reynolds number effect. The more scaling done, the more prominent the effect. It is determined that new technology such as higher TIT, OPR and turbomachinery [eta]poly's for small aircraft engines enable the operation of larger bypass ratios, which is the most influential parameter to SFC and fuel bum. The increase of bypass ratio up to a value of 8 is found to be effective for such improvement. SFC decrease from the current to mid-term model is found to be ~20% and ~9% from mid-term to far-term. Range and endurance improvements are found to be ~30% and ~10% respectively for the mission examined. Finally, the mid-term engine model has performance comparable to that of a current, larger state-of-the-art engine, thus suggesting that improvement in small gas turbine technology in the next 10 years will make the application of commodity thrust or distributed propulsion an attractive option for future aircraft.
by Nicholas Y.S. Chan.
S.M.
48
Bae, Jinwoo W. "An experimental study of surge control in a helicopter gas turbine engine." Thesis, Massachusetts Institute of Technology, 1998. http://hdl.handle.net/1721.1/50319.
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Sheehan, Jerry R. (Jerrard Robert). "Commericalization and transfer of technology in the U.S. jet aircraft engine industry." Thesis, Massachusetts Institute of Technology, 1991. http://hdl.handle.net/1721.1/42508.
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Protz, Christopher S. (Christopher Stephen) 1977. "Experimental investigation of microfabricated bipropellant rocket engines." Thesis, Massachusetts Institute of Technology, 2004. http://hdl.handle.net/1721.1/17794.
Full textAbstract:
Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2004.MIT Institute Archives copy: p. 301-328 bound in reverse order.
Includes bibliographical references (p. 325-328).
As satellite missions begin to require smaller satellites, launch systems and attitude control thrusters of reduced mass will be required. Microrocket engines could provide a low mass, high specific impulse, modular answer to these needs. These small rocket engines would produce thrust of order of 10's of Newtons at a thrust-to-weight of over 1000, over 10 times the thrust-to-weight of conventional chemical liquid bipropellant engines. The first microrocket engine thrust chamber and nozzle design measures 18 x 14 x 3 mm and is fabricated from single crystal silicon using MEMS microfabrication techniques offering the promise of low cost in production. This thesis describes an experimental investigation of bi-propellant microrocket engines and encompasses the fields of materials, microfabrication, combustion and chemical kinetics, instrumentation, packaging, and fluid dynamics. It builds on London's earlier gaseous propellant work, expanding the operating envelope of his motors to higher thrust levels and using these results to design liquid bi-propellant regeneratively cooled engines. Failure analysis of the original devices indicated failures were primarily caused by structural design flaws. Second generation gaseous propellant devices were built and tested. Providing reliable packaging interfaces between the macro test setup and the device proved very difficult. Two packaging methods involving modified geometries and glass seals were developed and allowed higher performance tests. Combustion experiments spanned a range of oxidizer-to-fuel ratios by mass of 1.6 to 2.5 and reached a maximum chamber pressure of 30 bar with a maximum thrust of 3 N at a thrust coefficient of 1.12. A maximum c* of 1650 m/s has been recorded.
(cont.) Experimental results were compared with CFD predictions which suggest that the low thrust coefficient of these devices is due to the overexpansion of gases in the nozzle at the test pressures in combination with the planar extruded nozzle geometry. CFD suggests that at higher chamber pressures the thrust coefficient will approach values up to 90-95 percent of the 1-D ideal case. Experimental values of characteristic exhaust velocity are in agreement with non- adiabatic predictions indicating that combustion is nearly complete. The chamber pressure and thrust limits in the current devices are due to localized failures at bond interfaces in the coolant passages. The potential of the current design is limited to approximately 60 bar by the coolant passage pressure limit, chamber structural limit, and injector manifold pressure limit. Potential liquid propellants for a regeneratively cooled storable bipropellant microrocket engine are examined. The design space for devices using these propellants is explored based on the thermal, structural, and fabrication constraints, and a design for a regeneratively cooled microrocket engine utilizing liquid nitrogen tetroxide and liquid JP-7 as propellants at a vacuum specific impulse of 267 s is presented. Directions for improved specific impulse engines include increasing the engine size by a factor of 2 to 4 and continuing research on hydrogen peroxide as a coolant.
by Christopher S. Protz.
Ph.D.