Bibliographies thématiques / Turbine a Gas Aeronautiche

Littérature scientifique sur le sujet « Turbine a Gas Aeronautiche »

Auteur : Grafiati
Publié le 10 mars 2023

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Articles de revues sur le sujet "Turbine a Gas Aeronautiche"

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Sarti Leme, Alexandre Domingos, Geraldo Creci, Edilson Rosa Barbosa de Jesus, Túlio César Rodrigues et João Carlos Menezes. « Finite Element Analysis to Verify the Structural Integrity of an Aeronautical Gas Turbine Disc Made from Inconel 713LC Superalloy ». Advanced Engineering Forum 32 (avril 2019) : 15–26. http://dx.doi.org/10.4028/www.scientific.net/aef.32.15.

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Gas turbines are very important because they can be used in several areas, such as aeronautics and electric power generation systems. The operation of a gas turbine can be done by less pollutant fuels when compared to traditional kerosene, for example, resulting in less degradation to environment. Gas turbines may fail from a variety of sources, with the possibility of serious damage results. In this work, the structural integrity of the hot disc of an aeronautical gas turbine is addressed. Several numerical analyses have been performed by the finite element method: Temperature Distributions, Thermal Stresses and Dilatations, Structural Stresses and Deformations, Modal Behaviors and Fatigue Analysis. Creep of blades has also been considered. These are the most important failure modes that can happen to the studied hot disc under operating service. All these analysis have been performed considering the boundary conditions at the design point with maximum rotational speed. The mesh of the problem has been strictly evaluated by adaptive refinement of nodes and elements combined with a convergence analysis of results. Then, the material and basic properties of the hot disc have been defined to assure a normal operation free from failures. Therefore, the mechanical drawings of the studied hot turbine disc have been released for manufacturing and the construction of the first prototype of the aeronautical gas turbine is in testing phase showing that the results presented in this work are consistent.
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Dediu, Gabriel, et Daniel Eugeniu Crunteanu. « Automatic Control System for Gas Turbines Test Rig ». Applied Mechanics and Materials 436 (octobre 2013) : 398–405. http://dx.doi.org/10.4028/www.scientific.net/amm.436.398.

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The technical evolution of the industrial and aeronautical groups involving gas turbines, determined by the request of increased efficiency and reliability, imposes the control through modern command and control automation systems. The paper describes a system destined to safely monitor, command and control the working conditions through complete automation of all command functions of a gas turbine. The system is suitable for all series of applications involving gas turbines, also providing a decrease in exploitation and maintenance costs.
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Dunham, J. « 50 years of turbomachinery research at Pyestock — part 2 : turbines ». Aeronautical Journal 104, no 1034 (avril 2000) : 199–207. http://dx.doi.org/10.1017/s0001924000028104.

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Abstract The two parts of this paper summarise the turbomachinery research undertaken at Pyestock during the 50 years since the National Gas Turbine Establishment was formed in 1946. The theoretical and experimental activities are described, and their influence on UK military and civil aero engines is assessed. The way in which NGTE supported non-aeronautical gas turbines is also explained. Part 2 covers turbines.
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Dunham, J. « 50 years of turbomachinery research at Pyestock — part one : compressors ». Aeronautical Journal 104, no 1033 (mars 2000) : 141–51. http://dx.doi.org/10.1017/s0001924000025331.

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Abstract The two parts of this paper summarise the turbomachinery research undertaken at Pyestock during the 50 years since the National Gas Turbine Establishment was formed in 1946. The theoretical and experimental activities are described, and their influence on UK military and civil aero engines is assessed. The way in which NGTE supported non-aeronautical gas turbines is also explained. Part one provides a general introduction and then covers compressors.
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Fatsis, Antonios. « Performance Enhancement of One and Two-Shaft Industrial Turboshaft Engines Topped With Wave Rotors ». International Journal of Turbo & ; Jet-Engines 35, no 2 (25 mai 2018) : 137–47. http://dx.doi.org/10.1515/tjj-2016-0040.

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Abstract Wave rotors are rotating equipment designed to exchange energy between high and low enthalpy fluids by means of unsteady pressure waves. In turbomachinery, they can be used as topping devices to gas turbines aiming to improve performance. The integration of a wave rotor into a ground power unit is far more attractive than into an aeronautical application, since it is not accompanied by any inconvenience concerning the over-weight and extra dimensioning. Two are the most common types of ground industrial gas turbines: The one-shaft and the two-shaft engines. Cycle analysis for both types of gas turbine engines topped with a four-port wave rotor is calculated and their performance is compared to the performance of the baseline engine accordingly. It is concluded that important benefits are obtained in terms of specific work and specific fuel consumption, especially compared to baseline engines with low compressor pressure ratio and low turbine inlet temperature.
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Lino, Vinicius da Silva, Damásio Sacrini, Adilson Vitor Rodrigues, Geraldo Creci et João Carlos Menezes. « Dynamic Characteristics of a Squeeze Film Damper used as Rear Bearing in a Single Spool Aeronautic Gas Turbine ». International Journal of Advanced Engineering Research and Science 10, no 1 (2023) : 019–24. http://dx.doi.org/10.22161/ijaers.101.4.

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Squeeze film dampers are widely used in aeronautic gas turbines because they effectively absorb vibrations and lessen the stresses on the structural components. In this study, we calculated the stiffness and damping dynamic coefficients of a squeeze film damper with open ends and a circumferential oil-feeding groove. This squeeze film damper was used as a rear bearing in an aeronautic gas turbine designed to generate 5-kN of thrust under ISA conditions. Three different radial clearances were investigated to determine the optimal bearing design configuration for the application because the radial clearance of a squeeze film damper is a crucial element in determining its dynamic stiffness and damping coefficients. To provide superior performance and avoid issues, a rotordynamic analysis using the calculated stiffness and damping dynamic coefficients can be conducted to predict the vibratory behavior of the entire rotating assembly.
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Derbal-Habak,, Hassina. « Alternative Materials for Performant TBCs : Short Review ». Journal of Mineral and Material Science (JMMS) 4, no 1 (30 janvier 2023) : 1–2. http://dx.doi.org/10.54026/jmms/1051.

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Thermal Barrier Coating (TBC) is a thermal insulation, which enables the coated substrate material to work above its melting temperature. The TBCs have been used to enhance the performance of the gas turbine for aeronautic and energy applications. Yttria stabilized zirconia YSZ (ZrO2 +7-8 wt.% Y2 O3 ) is a topcoat ceramic which is applied for more than 40 years to gas turbine components. YSZ has a high toughness and a good temperature capability up to about 1200 °C, higher operating temperatures is required for enhanced efficiency of gas turbine. Alternative materials for TBC application were developed during the last years allowing a higher temperature capability and lower thermal conductivity combined to higher toughness and thermochemical stability a of the TBCs.
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Amato, Giorgio, Matteo Giovannini, Michele Marconcini et Andrea Arnone. « Unsteady Methods Applied to a Transonic Aeronautical Gas Turbine Stage ». Energy Procedia 148 (août 2018) : 74–81. http://dx.doi.org/10.1016/j.egypro.2018.08.032.

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Brookes, Stephen Peter, Hans Joachim Kühn, Birgit Skrotzki, Hellmuth Klingelhöffer, Rainer Sievert, Janine Pfetzing, Dennis Peter et Gunther F. Eggeler. « Multi-Axial Thermo-Mechanical Fatigue of a Near-Gamma TiAl-Alloy ». Advanced Materials Research 59 (décembre 2008) : 283–87. http://dx.doi.org/10.4028/www.scientific.net/amr.59.283.

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A material family to replace the current superalloys in aeronautical gas turbine engines is considered to be that of gamma Titanium Aluminide (-TiAl) alloys. Structural components in aeronautical gas turbine engines typically experience large variations in temperatures and multiaxial states of stress under non-isothermal conditions. The uniaxial, torsional and bi-axial thermo-mechanical fatigue (TMF) behaviour of this -TiAl alloy have been examined at 400 – 800oC with strain amplitudes from 0.15% to 0.7%. The tests were conducted at both in-phase (IP) and out-of-phase (OP). The effects of TMF on the microstructure were also investigated. For the same equivalent mechanical strain amplitude uniaxial IP tests showed significantly longer lifetimes than pure torsional TMF tests. The non-proportional multiaxial OP test showed the lowest lifetimes at the same equivalent mechanical strain amplitude compared to the other types of tests.
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Sarnecki, Jarosław, Tomasz Białecki, Bartosz Gawron, Jadwiga Głąb, Jarosław Kamiński, Andrzej Kulczycki et Katarzyna Romanyk. « Thermal Degradation Process of Semi-Synthetic Fuels for Gas Turbine Engines in Non-Aeronautical Applications ». Polish Maritime Research 26, no 1 (1 mars 2019) : 65–71. http://dx.doi.org/10.2478/pomr-2019-0008.

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Abstract This article concerns the issue of thermal degradation process of fuels, important from the perspective of the operation of turbine engines, especially in the context of new fuels/bio-fuels and their implementation. The studies of the kerosene-based jet fuel (Jet A-1) and its blends with synthetic components manufactured according to HEFA and ATJ technology, were presented. Both technologies are currently approved by ASTM D7566 to produce components to be added to turbine fuels. Test rig investigations were carried out according to specific methodology which reflects the phenomena taking place in fuel systems of turbine engines. The mechanism of thermal degradation process was assessed on the basis of test results for selected properties, IR spectroscopy and calculation of activation energy. The results show that with the increase of the applied temperature there is an increment of the content of solid contaminants, water and acid for all tested fuels. Thermal degradation process is different for conventional jet fuel when compared to blends, but also semi-synthetic fuels distinguished by different thermal stability depending on a given manufacturing technology.
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Thèses sur le sujet "Turbine a Gas Aeronautiche"

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Martinez-Tamayo, Federico. « The impact of evaporatively cooled turbine blades on gas turbine performance ». Thesis, Massachusetts Institute of Technology, 1995. http://hdl.handle.net/1721.1/47385.

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Bradshaw, Sean D. (Sean Darien) 1978. « Probabilistic aerothermal design of gas turbine combustors ». Thesis, Massachusetts Institute of Technology, 2006. http://hdl.handle.net/1721.1/36286.

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Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2006.
Includes bibliographical references (p. 87-89).
This thesis presents a probability-based framework for assessing the impact of manufacturing variability on combustor liner durability. Simplified models are used to link combustor liner life, liner temperature variability, and the effects of manufacturing variability. A probabilistic analysis is then applied to the simplified models to estimate the combustor life distribution. The material property and liner temperature variations accounted for approximately 80 percent and 20 percent, respectively, of the combustor life variability. Furthermore, the typical combustor life was found to be approximately 20 percent less than the life estimated using deterministic methods for these combustors, and the probability that a randomly selected combustor will fail earlier than predicted using deterministic methods is approximately 80 percent. Finally, the application of a sensitivity analysis to a surrogate model for the life identified the leading drivers of the minimum combustor life and the typical combustor life as the material property variability and the variability of the near-wall combustor gas temperature, respectively.
by Sean Darien Bradshaw.
Ph.D.
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Underwood, David Scott. « Primary zone modeling for gas turbine combustors ». Thesis, Massachusetts Institute of Technology, 1999. http://hdl.handle.net/1721.1/32700.

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Thesis (Sc.D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1999.
"June 1999."
Includes bibliographical references (p. 107-110).
Gas turbine combustor primary zone flows are typified by swirling flow with heat release in a variable area duct, where a central toroidal recirculation zone is formed. The goal of the research was to develop reduced-order models for these flows in an attempt to gain insight into, and understanding of the behavior of swirling flows with combustion. The specific research objectives were (i) to develop a quantitative understanding and ability to compute the behavior of swirling flows with heat addition at conditions typical of gas turbine combustors, (ii) to assess the relative merits of various reduced-order models, and (iii) to define the applicability of these models in the design process. To this end, several reduced-order models of combustor primary zones were developed and assessed. The models represent different levels of modeling approximations and complexity. The models include a quasi-one-dimensional control volume analysis, a streamline curvature model, a quasi-one- dimensional model with recirculation zone capturing (CFLOW), and an axisymmetric Reynolds averaged Navier-Stokes code (UTNS). The models were evaluated through inter-comparison, and comparison with experiment. Following this evaluation, CFLOW was applied to a lean-premixed combustor for which three-dimensional Navier-Stokes solutions existed. These simplified analyses/models were able to capture the features of swirling flows with heat release across flow regimes of interest in gas turbine combustors, provide insight into the underlying physics, and yield guidelines for design purposes. Cross-comparison of the reduced-order models highlighted the aspects of these flows that need to be described accurately. Specifically, modeling of the mixing on the downstream boundary of a recirculation zone is crucial for accurate computation of these flows, with both Reynolds stresses and bulk transport across the interface being accounted for in order to capture recirculation zone closure. The simplified mixing and heat release models used had limitations arising from the need to input empirically-derived parameters. Calibration of these parameters with higher-fidelity computations and experiments allowed comparison of the models across the flow regimes of interest. Following calibration of the mixing and heat release models, CFLOW was able to compute recirculation zone volumes to within 25% of those given by both the axisymmetric and three-dimensional Navier-Stokes codes for swirl ratios between 0.5 and 1.0 and equivalence ratios between 0.0 and 0.8.
by David Scott Underwood.
Sc.D.
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Evans, Simon William 1977. « Thermal design of a cooled micro gas turbine ». Thesis, Massachusetts Institute of Technology, 2001. http://hdl.handle.net/1721.1/8093.

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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2001.
Includes bibliographical references (p. 169-170).
One of the major challenges associated with designing a micro gas turbine engine is the problem of heat transfer. The demonstration version of the engine deals with this problem by transferring excess heat from the turbine, to the compressor wall, through the rotor shaft. This is necessary to keep the turbine wall within its temperature constraints. The resulting heat transfer into the compressor flow however reduces the compressor performance to the point that the cycle will no longer close. A film cooled turbine has thus been pursued as a means of keeping the turbine within its temperature constraints and at the same time reducing heat transfer to the compressor. The thermal design of this cooled micro gas turbine has involved the design of the thermodynamic cycle, a secondary flow system to carry compressor discharge air to the turbine for cooling, and conceptual design of a turbine and rotor shaft to match the compressor. The analysis leading to this design identified turbine wall temperature, turbine exit radius and shaft area as three tools for increasing the power of the turbine, required to close the cycle. The design converged upon revealed that a very high cooling effectiveness is required to close the cycle, if the turbine wall is to be limited to 950K. This high effectiveness is calculated according to an empirical model established with data from full size engines, and thus represents an extrapolation of data with its attendant risks. A comparative model was developed as a regression of CFD results produced for the engine geometry. This model predicts adiabatic cooling effectiveness values too low to close the cycle. From the cycles studied, the recommended cycle configuration includes a 10mm diameter turbine with 1600K at rotor inlet. 41% of compressor inlet air is required to cool the turbine wall to 950K, and shaft area required to be 0.1% of a solid 6mm diameter shaft, i.e. 0.079mm2. The resulting cycle breaks even with a compressor pressure ratio of 2.46 and efficiency of 43%. Turbine efficiency is 63%. This solution shows that closure of the cycle is possible. It however suggests that further design study and technology development is needed to generate useful levels of engine performance.
by Simon William Evans.
S.M.
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Koupper, Charlie. « Unsteady multi-component simulations dedicated to the impact of the combustion chamber on the turbine of aeronautical gas turbines ». Phd thesis, Toulouse, INPT, 2015. http://oatao.univ-toulouse.fr/14187/1/koupper_partie_1_sur_2.pdf.

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De nos jours, seules les turbines à gaz sont à même de propulser les larges aéronefs (avions ou hélicoptères). Depuis les premiers prototypes construits dans les années 40, l’efficacité et la puissance de ces moteurs n’ont cessé de s’améliorer. Chaque composant atteint de tels niveaux de performance que seules une rupture technologique ou un investissement conséquent peuvent permettre de repousser les limites d’efficacité d’une turbine à gaz. Une solution alternative peut être trouvée en constatant qu’un moteur est un système intégré complexe dans lequel tous les composants interagissent entre eux, affectant les performances de chaque module en comparaison de leur fonctionnement isolé. Avec la compacité croissante des turbines à gaz, ces interactions entre modules du moteur sont clairement renforcées et leur étude constitue une potentielle source de gain en termes de performance globale du moteur. Dans ce contexte, l’interface du moteur la plus critique est aujourd’hui la connexion entre la chambre de combustion et la turbine, qui présente les niveaux de pression, température et contraintes les plus élevés du moteur. L’objectif de cette thèse est d’améliorer la caractérisation actuelle de l’interface chambre- turbine afin de juger les méthodes de développement de cette interface et de concourir à l’amélioration des performances de la turbine et sa durée de vie. Pour ainsi faire, un nouveau simulateur de chambre non réactif, représentatif des architectures de chambres pauvres récentes, est développé dans le contexte du projet européen FACTOR (FP7). L’écoulement dans le module est analysé d’une part via le recours massif aux Simulations aux Grandes Echelles (LES), et d’autre part par une caractérisation expérimentale sur une version trisecteur du module, installée à l’Université de Florence (Italie). En tirant profit des complémentarités entre approche numérique et expérimentale, une base de données exhaustive est construite pour qualifier les simulations avancées et caractériser les quantités physiques à l’interface entre la chambre et la turbine. Des diagnostics avancés et des procédures de validation s’appuyant sur les riches données temporelles sont proposés dans l’objectif d’améliorer les processus de design de l’interface chambre-turbine. Par exemple, il est montré qu’il est parfois possible et nécessaire d’aller au-delà d’une simple analyse des moyennes et variances pour qualifier les prédictions à cette interface. Pour approfondir l’étude de l’interaction chambre-turbine, des simulations LES comprenant à la fois le simulateur de chambre et une paire de stators de la turbine haute pression sont réalisées. Ces prédictions purement numériques mettent en évidence l’effet potentiel induit par la présence des stators ainsi que l’influence du calage angulaire par rapport aux injecteurs. Ce dernier ensemble de simulations souligne la difficulté de proprement appréhender l’interface chambre-turbine, mais confirme qu’il peut être simulé par une approche LES à l’avenir.
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Zhang, K. « Turbulent combustion simulation in realistic gas-turbine combustors ». Thesis, City, University of London, 2017. http://openaccess.city.ac.uk/17689/.

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The work presented in this thesis addresses issues involving the accurate and efficient numerical modelling of turbulence combustion with an emphasis on an industrially representative Tay model combustor. This combustor retained all essential features of a modern aero-engine rich burn combustor and thus the turbulence combustion within this combustor is much more complicated than those observed in the combustor-like burners typically considered in laboratory experiments. A comparative study of two combustion models based on a non-premixed assumption or a partially premixed assumption using the previously proposed models Zimont Turbulent Flame Speed Closure (ZTFSC) and Extended Coherent Flamelet Method (ECFM)) is presented in a first step. Comprehensive chemical reactions containing 244 reactions and 50 species are taken into account using a tabulated detailed chemistry approach and an assumed shape PDF to account for turbulence effects. The purpose of this study is to validate and compare the effectiveness of these models in predicting complex combustion and to improve upon for the defects observed in previous predictions of the same combustor. It is concluded that the use of models invoking the partially premixed combustion assumption can provide much more accurate results than models using a non-premixed combustion assumption especially in the primary zone of the combustor where turbulence combustion interaction is strong. In addition, certain shortcomings of steady RANS type models are identified as a result of strong unsteady effects and their inability to resolve the turbulence spectrum. Following this, two URANS models and the scale resolving simulation (SRS) approach such as a shear stress transport, K-omega, scale adaptive simulation (SSTKWSAS) combined with the partially premixed method identified in the first step are employed in a second step to further improve the accuracy achieved and to provide evidence and guidance in terms of the trade-off between accuracy and computational cost for complex turbulent combustion simulations. The second generation SRS model (SSTKWSAS) is applied to the complicated flow environment of a realistic combustor for the first time. The present work highlights the superiority of the combination of the SSTKWSAS approach and a partially premixed combustion model in terms of both accuracy and efficiency for predicting such combustion problems.
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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.

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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.
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Kleiven, Thomas J. (Thomas John). « Effect of gas path heat transfer on turbine loss ». Thesis, Massachusetts Institute of Technology, 2017. http://hdl.handle.net/1721.1/112466.

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Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2017.
Cataloged from PDF version of thesis.
Includes bibliographical references (pages 117-118).
This thesis presents an assessment of the impact of gas path, i.e., streamtube-to-streamtube, heat transfer on aero engine turbine loss and efficiency. The assessment, based on the concept of mechanical work potential [19], was carried out for two model problems to introduce the ideas. Three-dimensional RANS calculations were also conducted to show the application to realistic configurations. The first model problem, a constant area mixing duct, demonstrates the importance of selecting a fluid component loss metric appropriate to the purpose of the overall system in which the component resides. The phenomenon of thrust increase due to mixing is analyzed to show that system performance can increase even though there is a loss of thermodynamic availability. Gas path heat transfer affects mechanical work potential, and thus turbine loss, through a mechanism called thermal creation [19]. The second model problem, an inviscid heat exchanger, illustrates how thermal creation is due to enthalpy redistribution between flow regions with different local Brayton efficiency. Heat transfer across a static pressure difference, or between gases with different specific heat ratios, can cause turbine efficiency to increase or decrease depending on the direction of the heat flow. Three-dimensional RANS calculations have also been interrogated to define and determine the thermal creation, and thus the losses, in a modern two-stage cooled high pressure turbine. At representative engine operating conditions the effect of thermal creation was a 0.1% decrease in efficiency, with the thermal creation accounting for 1% of the overall lost work. Introducing coolant flow into the main gas path increased the loss from thermal creation in the first stage by 84% and decreased the loss from thermal creation in the second stage by 8%.
by Thomas J. Kleiven.
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.

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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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Livres sur le sujet "Turbine a Gas Aeronautiche"

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Aerothermodynamics of gas turbine and rocket propulsion. 3e éd. Reston, VA : American Institute of Aeronautics and Astronautics, 1997.

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Aerothermodynamics of gas turbine and rocket propulsion. Washington, DC : American Institute of Aeronautics and Astronautics, 1988.

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3

J, Holt Mark, dir. The turbine pilot's flight manual. Ames : Iowa State University Press, 1995.

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J, Holt Mark, dir. The turbine pilot's flight manual. 2e éd. Ames : Iowa State University Press, 2001.

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5

United States. National Aeronautics and Space Administration., dir. PROBABILISTIC ANALYSIS OF GAS TURBINE FIELD PERFORMANCE... NASA/TM--2002-211699... NATIONAL AERONAUTICS AND SPACE ADMINISTRATION... NOVEMBER. [S.l : s.n., 2003.

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Beck, Douglas Stephen. Gas-turbine regenerators. New York : Chapman & Hall, 1996.

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Beck, Douglas Stephen, et David Gordon Wilson. Gas-Turbine Regenerators. Boston, MA : Springer US, 1996. http://dx.doi.org/10.1007/978-1-4613-1209-3.

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Lieuwen, Tim C., et Vigor Yang, dir. Gas Turbine Emissions. Cambridge : Cambridge University Press, 2013. http://dx.doi.org/10.1017/cbo9781139015462.

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Walsh, Philip P. Gas turbine performance. Oxford : Blackwell Science, 1998.

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Beck, Douglas Stephen. Gas-Turbine Regenerators. Boston, MA : Springer US, 1996.

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Chapitres de livres sur le sujet "Turbine a Gas Aeronautiche"

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Scharnell, Lennart, et Stuart Sabol. « Gas Turbine Combustion ». Dans Practical Dispute Resolution, 2–4. Cham : Springer International Publishing, 2018. http://dx.doi.org/10.1007/978-3-031-01493-2_2.

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Beck, Douglas Stephen, et David Gordon Wilson. « Gas-Turbine Cycles ». Dans Gas-Turbine Regenerators, 37–62. Boston, MA : Springer US, 1996. http://dx.doi.org/10.1007/978-1-4613-1209-3_3.

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El Hefni, Baligh, et Daniel Bouskela. « Gas Turbine Modeling ». Dans Modeling and Simulation of Thermal Power Plants with ThermoSysPro, 297–309. Cham : Springer International Publishing, 2019. http://dx.doi.org/10.1007/978-3-030-05105-1_11.

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Beck, Douglas Stephen, et David Gordon Wilson. « Introduction ». Dans Gas-Turbine Regenerators, 1–26. Boston, MA : Springer US, 1996. http://dx.doi.org/10.1007/978-1-4613-1209-3_1.

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Beck, Douglas Stephen, et David Gordon Wilson. « Background ». Dans Gas-Turbine Regenerators, 27–35. Boston, MA : Springer US, 1996. http://dx.doi.org/10.1007/978-1-4613-1209-3_2.

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Beck, Douglas Stephen, et David Gordon Wilson. « Regenerator Designs ». Dans Gas-Turbine Regenerators, 63–78. Boston, MA : Springer US, 1996. http://dx.doi.org/10.1007/978-1-4613-1209-3_4.

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Beck, Douglas Stephen, et David Gordon Wilson. « Design Procedures and Examples ». Dans Gas-Turbine Regenerators, 79–120. Boston, MA : Springer US, 1996. http://dx.doi.org/10.1007/978-1-4613-1209-3_5.

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Beck, Douglas Stephen, et David Gordon Wilson. « Regenerator Performance ». Dans Gas-Turbine Regenerators, 121–233. Boston, MA : Springer US, 1996. http://dx.doi.org/10.1007/978-1-4613-1209-3_6.

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Zohuri, Bahman, et Patrick McDaniel. « Gas Turbine Working Principals ». Dans Combined Cycle Driven Efficiency for Next Generation Nuclear Power Plants, 149–74. Cham : Springer International Publishing, 2017. http://dx.doi.org/10.1007/978-3-319-70551-4_7.

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Kulikov, Gennady G., et Haydn A. Thompson. « Linear Gas Turbine Modelling ». Dans Advances in Industrial Control, 89–116. London : Springer London, 2004. http://dx.doi.org/10.1007/978-1-4471-3796-2_6.

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Actes de conférences sur le sujet "Turbine a Gas Aeronautiche"

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Rettler, M. W., M. L. Easley et J. R. Smyth. « Ceramic Gas Turbine Technology Development ». Dans ASME 1995 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1995. http://dx.doi.org/10.1115/95-gt-207.

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Under the U.S. Dept. of Energy/National Aeronautics and Space Administration (DOE/NASA) funded Ceramic Turbine Engine Demonstration Project, formerly the Advanced Turbine Technology Applications Project (ATTAP), AlliedSignal Engines is addressing the remaining critical concerns slowing the commercialization of structural ceramics in gas turbine engines. These issues include demonstration of ceramic component reliability, readiness of ceramic suppliers to support ceramic production needs, and development of ceramic design technologies. The AlliedSignal/Garrett Model 331-200[CT] Auxiliary Power Unit (APU) is being used as a ceramics test bed engine. The first-stage turbine blades and nozzles were redesigned using ceramic materials, employing the design methods developed during the earlier DOE/NASA-funded Advanced Gas Turbine (AGT) and ATTAP programs. Ceramic engine components have been fabricated and are now being evaluated in laboratory engine testing. The fabrication processes for these components will provide the framework for a demonstration of manufacturing process scale-up to the minimum level for commercial viability. The laboratory engine testing is helping to refine the component designs and focus the development of ceramic component technologies. Extended engine endurance testing and field testing in commercial aircraft is planned to demonstrate ceramic component reliability. Significant progress has been made during 1994. An engine with ceramic turbine nozzles was successfully operated and engine tests in the laboratory are continuing to gather useful data. An engine equipped with ceramic blades was also tested, but blade fractures occurred, interrupting operation. An extensive investigation has identified possible vibration and contact problems. Investigative evaluation efforts are continuing to identify the problem source and determine go-forward plans for ceramic blade development. Component design technologies have progressed in the areas of modeling particle impact pulverization, development of a ceramic hot corrosion environmental life model, and methods for evaluating ceramic contact damage. The planned ceramic manufacturing scale-up was initiated with two ceramics vendors, Norton Advanced Ceramics (East Granby, CT) and AlliedSignal Ceramic Components (Torrance, CA). The scaleup demonstration program is emphasizing improvement of ceramic processing yields and increased production rates. Work summarized in this paper was funded by the U.S. Dept. of Energy (DOE) Office of Transportation Technologies, as part of the Turbine Engine Technologies Program, and administered by the NASA Lewis Research Center, Cleveland, OH under Contract No. DEN3-335.
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Easley, M. L., et J. R. Smyth. « Ceramic Gas Turbine Technology Development ». Dans ASME 1996 International Gas Turbine and Aeroengine Congress and Exhibition. American Society of Mechanical Engineers, 1996. http://dx.doi.org/10.1115/96-gt-367.

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Under the U.S. Department Of Energy/National Aeronautics and Space Administration (DOE/NASA) funded Ceramic Turbine Engine Demonstration Program, AlliedSignal Engines is addressing the remaining critical concerns slowing the commercialization of structural ceramics in gas turbine engines. These issues include demonstration of ceramic component reliability, readiness of ceramic suppliers to support ceramic production needs, and enhancement of ceramic design methodologies. The AlliedSignal/Garrett Model 331-200[CT] Auxiliary Power Unit (APU) is being used as a ceramics test bed engine. For this program, the APU First-stage turbine blades and nozzles were redesigned using ceramic materials, employing the design methods developed during the earlier DOE/NASA funded Advanced Gas Turbine (AGT) and Advanced Turbine Technologies Application Project (ATTAP) programs. The present program includes ceramic component design, fabrication, and testing, including component bench tests and extended engine endurance testing and field testing. These activities will demonstrate commercial viability of the ceramic turbine application. In addition, manufacturing process scaleup for ceramic components to the minimum level for commercial viability will be demonstrated. Significant progress has been made during the past year. Engine testing evaluating performance with ceramic turbine nozzles has accumulated over 910 hours operation. Ceramic blade component tests were performed to evaluate the effectiveness of vibration dampers and high-temperature strain gages, and ceramic blade strength and impact resistance. Component design technologies produced impact-resistance design guidelines for inserted ceramic axial blades, and advanced the application of thin-film thermocouples and strain gages on ceramic components. Ceramic manufacturing scaleup activities were conducted by two ceramics vendors, Norton Advanced Ceramics (East Granby, CT) and AlliedSignal Ceramic Components (Torrance, CA). Following the decision of Norton Advanced Ceramics to leave the program, a subcontract was initiated with the Kyocera Industrial Ceramics Company Advanced Ceramics Technology Center (Vancouver, WA). The manufacturing scaleup program emphasizes improvement of process yields and increased production rates. Work summarized in this paper was funded by the U.S. Dept. Of Energy (DOE) Office of Transportation Technologies, part of the Turbine Engine Technologies Program, and administered by the NASA Lewis Research Center, Cleveland, OH under Contract No. DEN3-335.
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Kinney, Troy W., et Michael L. Easley. « Ceramic Gas Turbine Technology Development ». Dans ASME 1997 International Gas Turbine and Aeroengine Congress and Exhibition. American Society of Mechanical Engineers, 1997. http://dx.doi.org/10.1115/97-gt-465.

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Under the U.S. Dept. of Energy (DoE) funded Ceramic Turbine Engine Demonstration Project (CTEDP), AlliedSignal Engines is addressing remaining critical concerns slowing the commercialization of structural ceramics in gas turbine engines. These issues include demonstration of ceramic component reliability, readiness of ceramic suppliers to support ceramic production needs, and development of ceramic design technologies. The AlliedSignal/Garrett Model 331-200[CT] auxiliary power unit (APU) is being used as a ceramics test bed engine. The first-stage turbine blades and nozzles were redesigned for ceramic materials, employing design methods developed during the earlier Dept. of Energy/National Aeronautics and Space Administration (DoE/NASA)-funded Advanced Gas Turbine (ACT) and Advanced Turbine Technology Applications (ATTAP) programs. The fabrication processes for these components provide the framework for demonstration of ceramic manufacturing process scale-up to the minimum level for commercial viability. Ceramic engine components have been fabricated and are now being evaluated in laboratory engine testing. This testing is helping to refine the component designs and focus the development of ceramic component technologies. Extended engine endurance testing and field testing in commercial aircraft is planned, to demonstrate ceramic component reliability. Significant progress was made during 1996 in the ceramic component manufacturing scale-up activities. The CTEDP ceramics subcontractors, AlliedSignal Ceramic Components (Torrance, CA) and Kyocera Industrial Ceramics Corporation (Vancouver, WA) demonstrated increased capacity and improved yields of silicon nitride materials. Planned ceramic turbine nozzle manufacturing demonstrations were initiated by both companies. Ceramic design technology was further refined in several areas. Work continued in defining boundary conditions for impact modeling of ceramic turbine engines, including completion of a three-dimensional trajectory analysis for combustor carbon particles in the engine flowpath. Contact rig tests and supporting analyses helped define the effectiveness of compliant layers in reducing ceramic turbine blade attachment contact stresses, and the results are aiding the evolution of more effective compliant layer configurations. This work supported evaluation of various ceramic turbine blade attachment designs in subelement and engine tests. Thin-film strain gage technology for measuring vibratory levels at high temperatures was successfully applied on ceramic turbine blades. Ceramic materials were screened for susceptibility to cyclic hot corrosion fatigue at the conditions affecting turbine blades. Stress rupture testing in support of the proof test methodology development was completed. Engine endurance tests with ceramic turbine nozzles accumulated over 482 additional hours of successful operation. Ceramic turbine blades were successfully demonstrated in over 190 hours of engine operation. This work brought the combined ceramic component engine test experience to over 1500 operating hours. Work summarized in this paper was funded by the DoE Office of Transportation Technologies, as part of the Turbine Engine Technologies Program, and administered through Fiscal Year 1996 by the NASA Lewis Research Center, Cleveland, OH under Contract No. OEN3-335.
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Zhang, Chenkai, Jun Hu, Zhiqiang Wang et Xiang Gao. « Design Work of a Compressor Stage Through High-to-Low Speed Compressor Transformation ». Dans ASME 2013 Gas Turbine India Conference. American Society of Mechanical Engineers, 2013. http://dx.doi.org/10.1115/gtindia2013-3506.

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Low-speed model testing has advantages such as great accuracy, low cost and risk, so it’s widely used in the design procedure of HPC exit stage. The low-speed model testing project is conducted in Nanjing University of Aeronautics and Astronautics, to represent aerodynamic load and flow field structure of the seventh stage of a high-performance 10-stage high-pressure compressor. This paper outlines the design work of the low speed four-repeating-stage axial compressor, the third stage of which is the testing stage. The first two stages and the last stage provide the compressor with entrance and exit conditions respectively. The high-to-low speed transformation process involves both geometric and aerodynamic considerations. Accurate similarities demand the same Mach number and Reynolds number, which will not be maintained due to motor power/size and its low-speed feature. Compromises of constraints are obvious. Modeling principles are presented in high-to-low speed transformation. Design work is carried out based on these principles. Four main procedures are proceeded subsequently in the general design, including establishment of low-speed modeling target, global parameter design of modeling stage, throughflow aerodynamic design and blading design. In global parameter design procedure, rotational speed, shroud diameter, hub-tip ratio, mid-span chord and axial spacing between stages are determined by geometrical modeling principles. During throughflow design process, radial distributions of aerodynamic parameters such as D-Factor, pressure-rise coefficient, loss coefficients, stage reaction and other parameters are obtained by determined aerodynamic modeling principles. Finally, rotor and stator blade profiles of LSRC at seven span locations are adjusted, to make sure that blade surface pressure coefficients agree well with that of the HPC. 3D flow calculations are performed on low-speed four-stage axial compressor, and the resultant flow field structures agree well with that of the HPC. It’s worth noting that a large separation zone appears in both suction surfaces of LSRC and HPC. How to diminish it through 3D blading design in LSRC test rig is our further work.
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Rouser, Kurt P., Caitlin R. Thorn, Aaron R. Byerley, Charles F. Wisniewski, Scott R. Nowlin et Kenneth W. Van Treuren. « Integration of a Turbine Cascade Facility Into an Undergraduate Thermo-Propulsion Sequence ». Dans ASME Turbo Expo 2013 : Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2013. http://dx.doi.org/10.1115/gt2013-94744.

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The Department of Aeronautics at the United States Air Force Academy utilizes a closed-loop, two-dimensional turbine cascade wind tunnel to reinforce a learning-focused undergraduate thermo-propulsion sequence. While previous work presented in the literature outlined the Academy thermo-propulsion sequence and the contextual framework for instruction, this current paper addresses how the Academy turbine cascade facility is integrated into the aeronautical engineering course sequence. Cadets who concentrate in propulsion are to some extent prepared for each successive course through their contact with the cascade, and ultimately they graduate with an exposure to experimental research that enhances their grasp of gas turbine engine fundamentals. Initially, the cascade is used to reinforce airfoil theory to all cadets in the Fundamentals of Aeronautics course. Aeronautical engineering majors take this course during the first semester of their sophomore year. The next semester all aeronautical engineering majors take Introduction to Aero-thermodynamics. In this course, the closed-loop aspect of the cascade facility is used to reinforce concepts of work addition to the flow. Heat transfer is also discussed, using the heat exchanger that regulates test section temperature. Exposure to the cascade also prepares cadets for the ensuing Introduction to Propulsion and Aeronautics Laboratory courses, taken in the junior and senior year, respectively. In the propulsion course, cadets connect thermodynamic principles to component analysis. In the laboratory course, cadets work in pairs on propulsion projects sponsored by the Air Force Research Laboratory, including projects in the cascade wind tunnel. Individual cadets are selected from the cascade research teams for summer internships, working at an Air Force Research Laboratory turbine cascade tunnel. Ultimately, cadet experiences with the Academy turbine cascade help lay the foundation for a two-part senior propulsion capstone sequence in which cadets design a gas turbine engine starting with the overall cycle selection leading to component-level design. The turbine cascade also serves to integrate propulsion principles and fluid mechanics through a senior elective Computational Fluid Dynamics course. In this course, cadets may select a computational project related to the cascade. Cadets who complete the thermo-propulsion sequence graduate with a thorough understanding of turbine engine fundamentals from both conceptual and applied perspectives. Their exposure to the cascade facility is an important part of the process. An assessment of cadet learning is presented to validate the effectiveness of this integrated research-classroom approach.
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Anbarasan, Selwyn, S. Esakki Muthu, Hardik Roy, P. Udayanan et Girish K. Degaonkar. « Residual Life Estimation of Axial Compressor Blade of a Turbo-Shaft Engine ». Dans ASME 2014 Gas Turbine India Conference. American Society of Mechanical Engineers, 2014. http://dx.doi.org/10.1115/gtindia2014-8241.

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In aeronautical industry, flight safety is the first and foremost concern. Structural failure in aero engine aids to high risk in flight safety and human lives. High Cycle Fatigue (HCF) failure accounts for forty percent of blade structural failure. All critical components in the aero engine are life limited components and are replaced when its prescribed life is reached. Earlier components are designed as per safe-life design philosophy. Ninety percent of the critical components are retired utilizing less than fifty percent of its safe-life capability. Extending the life of component reduces the operating cost and lowers downtime for the fleet operator. This paper describes the life extension methodology used for a first stage axial compressor blade of turbo-shaft engine. The blade which has completed its safe life in the aircraft in service is identified. The studies are carried out to determine the residual life of the blade to extend the life. First stage axial compressor blade is made from titanium alloy and is fixed with the disc by pins. The force on the blade due to aerodynamic excitation which is required for life estimation is predicted using CFD tools followed by the estimation of alternate stress on the blade using FEM tools. Stress based life estimation methodology is used to estimate HCF life of the blade under engine operating condition. The estimated life of the blade is to be confirmed in rig testing. Life completed blades are identified from different engines for testing of the residual life. Metallurgical studies are carried out on the specimens from the blade to check the residual properties. The blades which qualify after the dimension inspection and crack detection are used for testing residual life. The blades are tested in vibration test rig facility in which incremental fatigue testing method is employed. The blades are excited by the electro-dynamic shaker at its first natural frequency (1F mode) and the amplitude is increased in steps in a steady interval of time. The HCF life of the blade is calculated by Miner’s hypothesis. The residual life of the blade is estimated by vibration rig testing and is compared with the numerical estimation. It is found that the residual life of the life completed component has got potential life of one more overhaul.
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Li, Jibao, Arthur H. Lefebvre et James R. Rollbuhler. « Effervescent Atomizers for Small Gas Turbines ». Dans ASME 1994 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1994. http://dx.doi.org/10.1115/94-gt-495.

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An experimental investigation is conducted into the potential of effervescent atomizers as fuel injectors for gas turbine engines. The designs studied include three different configurations of multihole effervescent atomizers and an effervescent/airblast hybrid atomizer. In all tests the liquid employed is water. The spray characteristics investigated include drop size distributions and liquid flux distributions within the spray. The results obtained show that multi hole effervesent atomizers combine good atomization with uniform liquid flux distribution. This makes them especially suitable for application to annular combustors because they allow appreciable reductions to be made 1n the number of fuel injectors needed to achieve uniform circumferential fuel distribution. The hybrid atomizer also combines good atomization with the capability of wide cone angles. The only drawback exhibited by these atomizers is the need for a separate supply of atomizing air. This drawback could restrict their applications to non-aeronautical gas turbine engines.
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Andriani, Roberto, Fausto Gamma et Umberto Ghezzi. « Main Effects of Intercooling and Regeneration on Aeronautical Gas Turbine Engines ». Dans 46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit. Reston, Virigina : American Institute of Aeronautics and Astronautics, 2010. http://dx.doi.org/10.2514/6.2010-6539.

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Cirtwill, Joseph D., Sina Kheirkhah, Pankaj Saini, Krishna Venkatesan et Adam M. Steinberg. « Analysis of intermittent thermoacoustic oscillations in an aeronautical gas turbine combustor ». Dans 55th AIAA Aerospace Sciences Meeting. Reston, Virginia : American Institute of Aeronautics and Astronautics, 2017. http://dx.doi.org/10.2514/6.2017-0824.

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Popescu, Jeni A., Valeriu A. Vilag, Romulus Petcu, Valentin Silivestru et Virgil Stanciu. « Researches Concerning Kerosene-to-Landfill Gas Conversion for an Aero-Derivative Gas Turbine ». Dans ASME Turbo Expo 2010 : Power for Land, Sea, and Air. ASMEDC, 2010. http://dx.doi.org/10.1115/gt2010-23436.

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The aero-derivative gas turbine represents an advanced solution for technologic transfer from aeronautics to industrial applications, including high efficiency, reduced dimensions and high reliability. The paper, as result of a research project, is focused on an application using an aero-derivative gas turbine as an installation for CO2 rich landfill gas valorization. The paper also presents the potential for landfill gas production in Romania, in the context of the requirements imposed by the environmental laws. A calculation is realized based on demographic statistics, showing the most suitable areas in the country for obtaining the landfill gas. The first part is dedicated to a comparative examination of classical liquid fuel, kerosene, and two gaseous fuels, methane and landfill gas with equal volume ratio of methane and carbon dioxide, analyzed from the point of view of their combustion performances in the gas turbine, with the help of CEA program developed by NASA. Considering the nowadays utilization of CFD simulations for design purpose in many activity fields from the engineering domain, the results provided by the CEA program, along with the ones provided by the gas turbine’s producer, were considered input data for the numerical approaches of the combustion process of methane and landfill gas in the known combustion chamber using a commercial CFD code. The main goal of the CFD applications is to determine the optimum geometric configuration of the new injection system in order to obtain a stabilized process and high performances in safety conditions, for low working regimes and nominal regime, as defined by experimental data and producer’s recommendations. Previous successful experimentations on test bench following the combustion simulation of methane gas and the encouraging results from the CFD simulations lead to new experimentations of the gas turbine working on landfill gas in order to validate the numerical approaches, activity described in the third part of the paper. A technological fueling scheme was designed, the geometrical adjustments were made according to previous simulations and the landfill gas was simulated using a homogenization device installed on the fuel line for a forced mixing of the two non-reactive substances, methane and carbon dioxide. The gas turbine was prepared and instrumented for bench testing and stable working was obtained for speeds of 27–63% of the nominal one. The conclusions are related to the execution of an installation allowing experimentation of gas turbines working on landfill gas and future researches focusing on tests for higher working regimes.
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Rapports d'organisations sur le sujet "Turbine a Gas Aeronautiche"

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Epstein, A. H., K. S. Breuer, J. H. Lang, M. A. Schmidt et S. D. Senturia. Micro Gas Turbine Generators. Fort Belvoir, VA : Defense Technical Information Center, décembre 2000. http://dx.doi.org/10.21236/ada391343.

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Pint, Bruce A., Michael M. Kirka, Gary S. Marlow, Charles S. Hawkins, Jim Kesseli et Jim Nash. Internally Cooled Turbine Rotor for Small Gas Turbine. Office of Scientific and Technical Information (OSTI), novembre 2017. http://dx.doi.org/10.2172/1427664.

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Unknown. ADVANCED GAS TURBINE SYSTEMS RESEARCH. Office of Scientific and Technical Information (OSTI), janvier 2002. http://dx.doi.org/10.2172/791987.

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Unknown. ADVANCED GAS TURBINE SYSTEMS RESEARCH. Office of Scientific and Technical Information (OSTI), février 2002. http://dx.doi.org/10.2172/793004.

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Unknown. ADVANCED GAS TURBINE SYSTEMS RESEARCH. Office of Scientific and Technical Information (OSTI), avril 2002. http://dx.doi.org/10.2172/794939.

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Unknown. ADVANCED GAS TURBINE SYSTEMS RESEARCH. Office of Scientific and Technical Information (OSTI), janvier 2000. http://dx.doi.org/10.2172/766242.

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Unknown. ADVANCED GAS TURBINE SYSTEMS RESEARCH. Office of Scientific and Technical Information (OSTI), juillet 1999. http://dx.doi.org/10.2172/769312.

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Unknown. ADVANCED GAS TURBINE SYSTEMS RESEARCH. Office of Scientific and Technical Information (OSTI), octobre 1999. http://dx.doi.org/10.2172/769313.

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Metz, Stephen D., et David L. Smith. Survey of Gas Turbine Control for Application to Marine Gas Turbine Propulsion System Control. Fort Belvoir, VA : Defense Technical Information Center, janvier 1989. http://dx.doi.org/10.21236/ada204713.

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Korjack, T. A. A Twisted Turbine Blade Analysis for a Gas Turbine Engine. Fort Belvoir, VA : Defense Technical Information Center, août 1997. http://dx.doi.org/10.21236/ada329581.

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