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1
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.
3
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.
4
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.
5
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.
6
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.
7
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.
9
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.
10
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.
11
Prashanth, Prakash. "Post-combustion emissions control for aero-gas turbine engines". Thesis, Massachusetts Institute of Technology, 2018. https://hdl.handle.net/1721.1/122402.
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Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2018Cataloged from PDF version of thesis.
Includes bibliographical references (pages 47-50).
Aviation NO[subscript x] emissions have an impact on air quality and climate change, where the latter is magnified due to the higher sensitivity of the upper troposphere and lower stratosphere. In the aviation industry, efforts to increase the efficiency of propulsion systems are giving rise to higher overall pressure ratios which results in higher NO[subscript x] emissions due to increased combustion temperatures. This thesis identifies that the trend towards smaller engine cores (gas generators) that are power dense and contribute little to the thrust output presents new opportunities for emissions control that were previously unthinkable when the core exhaust stream contributed significant thrust. This thesis proposes and assesses selective catalytic reduction (SCR), which is a post-combustion emissions control method used in ground-based sources such as power generation and heavy-duty diesel engines, for use in aero-gas turbines.
The SCR system increases aircraft weight and introduces a pressure drop in the core stream. The effects of these are evaluated using representative engine cycle models provided by a major aero-gas turbine manufacturer. This thesis finds that employing an ammonia-based SCR can achieve close to 95% reduction in NO[subscript x] emissions for ~0.4% increase in block fuel burn. The large size of the catalyst needs to be housed in the body of the aircraft and hence would be suitable for future designs where the engine core is also within the fuselage, such as would be possible with turbo-electric or hybrid-electric designs. The performance of the post-combustion emissions control is shown to improve for smaller core engines in new aircraft in the NASA N+3 time-line (2030-2035), suggesting the potential to further decrease the cost of the ~95% NO[subscript x] reduction to below ~0.4% fuel burn.
Using a global chemistry and transport model (GEOS-Chem) this thesis estimates that using ultra-low sulfur (<15 ppm fuel sulfur content) in tandem with post-combustion emissions control results in a ~92% reduction in annual average population exposure to PM₂.₅ and a ~95% reduction in population exposure to ozone. This averts approximately 93% of the air pollution impact of aviation.
by Prakash Prashanth.
S.M.
S.M. Massachusetts Institute of Technology, Department of Aeronautics and Astronautics
12
Allaire, Douglas L. "A physics-based emissions model for aircraft gas turbine combustors". Thesis, Massachusetts Institute of Technology, 2006. http://hdl.handle.net/1721.1/35584.
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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2006.Includes bibliographical references (p. 103-105).
In this thesis, a physics-based model of an aircraft gas turbine combustor is developed for predicting NO. and CO emissions. The objective of the model is to predict the emissions of current and potential future gas turbine engines within quantified uncertainty bounds for the purpose of assessing design tradeoffs and interdependencies in a policy-making setting. The approach taken is to capture the physical relationships among operating conditions, combustor design parameters, and pollutant emissions. The model is developed using only high-level combustor design parameters and ideal reactors. The predictive capability of the model is assessed by comparing model estimates of NO, and CO emissions from five different industry combustors to certification data. The model developed in this work correctly captures the physical relationships between engine operating conditions, combustor design parameters, and NO. and CO emissions. The NO. estimates are as good as, or better than, the NO. estimates from an established empirical model; and the CO estimates are within the uncertainty in the certification data at most of the important low power operating conditions.
by Douglas L. Allaire.
S.M.
13
Jackson, Keith S. (Keith Stuart). "CAD-casting of gas turbine airfoils using three dimensional printing". Thesis, Massachusetts Institute of Technology, 1997. http://hdl.handle.net/1721.1/10518.
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Kluka, James Anthony. "The design of low-leakage modular regenerators for gas-turbine engines". Thesis, Massachusetts Institute of Technology, 1995. http://hdl.handle.net/1721.1/46564.
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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1995.Includes bibliographical references (p. [229]-231).
The design of a modular regenerator concept (patented by Wilson and MIT) for gas-turbine engines is investigated. Mechanical design analysis and theoretical performance calculations were made to show the concept's functionality and performance benefits over current regenerator designs. The modular regenerator concept consists of a ceramic-honeycomb matrix discretized into rectangular blocks, called modules. The modules are exposed to hot (turbine exhaust) and cold (compressor outlet) streams, then are periodically transported through linear passages from one stream to the other. Separating the matrix into modules reduces the transverse sealing lengths substantially. Furthermore, the range of gas-turbine applications increases with the modular concept since larger matrix face areas are possible. Module design is investigated which includes using current research results pertaining to manufacturing technology for rotary regenerators. Mechanical design analysis was made to investigate the possible module-movement schemes. Several regenerator configurations and orientations are introduced. One particular concept balances the pressure forces such that the power requirement for module movement is reduced substantially. Design drawings of a possible modular prototype showing the general configuration and mechanical layout accompany the analysis. A method for determining the regenerator size and measuring its fluid-mechanical and heat-transfer performance is given. An optimization study is made by analyzing the effects when several chosen design parameters are varied. Numerical results of a modular concept for a small gas-turbine engine (120 kW) are given. Seal leakage calculations were made for two modular concepts and compared to the leakage rates for two rotary concepts. The total seal-leakage rates for both modular cases were considerably less than the rotary concepts and can be reduced to well under one percent. In addition, techniques for further leakage reduction are given. Other design issues (to further prove the modular concept's feasibility) not covered in this study have been identified. Guidelines for investigating these issues are given.
by James Anthony Kluka.
S.M.
15
Berg, Rachel A. (Rachel Antonette). "Experimental and analytical assessment of cavity modes in a gas turbine wheelspace". Thesis, Massachusetts Institute of Technology, 2016. http://hdl.handle.net/1721.1/103445.
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Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2016.Cataloged from PDF version of thesis.
Includes bibliographical references (pages 103-106).
High response pressure data from a high-speed 1.5-stage turbine Hot Gas Ingestion Rig shows the existence of cavity modes in the rim-seal-wheelspace cavity for representative turbine engine operating conditions with purge flow. The experimental results and observations are complemented by computational assessments of cavity modes associated with flow in canonical cavity configurations. The cavity modes identified include Shallow Cavity modes and Helmholtz resonance. The response of the cavity modes to variation in design and operating parameters are assessed. These parameters include cavity aspect ratio, purge flow ratio, and flow angle defined by the ratio of primary tangential to axial velocity. Scaling the cavity modal response based on computational results and available experimental data in terms of the appropriate reduced frequencies appears to indicate the potential presence of a Deep Cavity mode as well. Computational assessment of canonical cavity flow suggests that increasing purge flow ratio mitigates Shallow Cavity modal response, in accord with data for the first Shallow Cavity mode but in contrast to data for the second Shallow Cavity mode. Likewise, increasing primary flow angle reduces the Shallow Cavity modal response that vanishes for flow angle beyond 450*. This computational observation is in contrast to the rig data that show the modal response is nevertheless present with a flow angle greater than 45*. An implication from the computational parametric assessments is that increasing purge flow and primary flow angle could provide a stabilizing effect on the response. Experimental requirements to quantify the effects of cavity modes on hot gas ingestion are identified along with inadequacies in the current rig set-up with the associated instrumentation system. As such, the role of cavity modes on hot gas ingestion cannot be clarified based on the current set of data.
by Rachel A. Berg.
S.M.
16
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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Dupuy, Fabien. "Reduced Order Models and Large Eddy Simulation for Combustion Instabilities in aeronautical Gas Turbines". Thesis, Toulouse, INPT, 2020. http://www.theses.fr/2020INPT0046.
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Des réglementations de plus en plus strictes et un intérêt environnemental grandissant ont poussé les constructeurs de moteurs aéronautiques à développer la génération actuelle de chambres de combustion, affichant des consommations et émissions de polluants plus basses que jamais. Cependant, les phases de conception de chambres modernes ont clairement mis en évidence que celles-ci sont plus susceptibles de développer des instabilités de combustion, où le couplage entre l'acoustique de la chambre et la flamme suscite de larges oscillations de pression ainsi que des vibrations de la structure. Ces instabilités peuvent endommager le moteur, et potentiellement entraîner sa destruction. Dans le même temps, de considérables avancées ont eu lieu dans le domaine de la simulation numérique, et la Mécanique des Fluides Numérique (MFN) a démontré sa capacité à reproduire la dynamique de flammes instationnaires et les instabilités de combustion observées dans les moteurs. Pourtant, même avec le matériel informatique moderne, le temps de calcul reste la contrainte clé de ces simulations haute-fidélité, qui demeurent très coûteuses. Typiquement, couvrir la totalité du domaine de fonctionnement pour un moteur industriel est encore hors de portée. Des modèles dits bas-ordre existent également, et prédire efficacement les instabilités de combustion par leur intermédiaire est envisageable à la condition d'une modélisation appropriée de l'interaction entre l'acoustique et la flamme. La méthode de modélisation la plus commune de cet élément critique est la fonction de transfert de flamme (FTF) qui lie les fluctuations de taux de dégagement de chaleur aux fluctuations de vitesse en un point donné. Cette fonction de transfert peut être obtenue à partir de modèles analytiques, mais très peu existent pour des flammes swirlées turbulentes. Une autre approche consiste à réaliser des mesures expérimentales ou des simulations haute fidélité coûteuses, réduisant à néant la capacité de prédiction rapide recherchée avec les méthodes bas-ordre. Cette thèse vise donc à développer des outils bas ordre à la fois rapides et fiables pour la modélisation des instabilités de combustion, ainsi qu'à améliorer la compréhension des mécanismes inhérents à la réponse acoustique d'une flamme swirlée. A cet effet, une approche hybride nouvelle est proposée, où un nombre réduit de simulations haute fidélité peut être utilisé pour déterminer les paramètres d'entrée d'un modèle analytique représentatif de la fonction de transfert d'une flamme swirlée prémélangée. Le modèle analytique s'appuie sur des travaux antérieurs traitant la flamme comme une interface perturbée, et prend en compte la conversion acoustique-vorticité à travers un swirler. La validité du modèle est mise à l'épreuve en déterminant les divers paramètres nécessaires associés à partir de simulations numériques réactives stationnaires et pulsées d'une flamme prémélangée swirlée académique. Il est également démontré que le modèle peut prendre en compte diverses amplitudes de perturbation. Enfin, des simulations haute-fidélité 3D d'une turbine à gaz industrielle alimentée par un combustible liquide sont réalisées afin de déterminer s'il est possible de prédire numériquement un mode d'instabilité de combustion observé lors des essais. Pour cela, un ensemble de simulations forcées est mené à bien afin de souligner l'importance de l'acquisition de la réponse de la flamme diphasique, en comparant les positions de référence utilisées pour mesurer les vitesses fluctuantes ainsi que l'amplitude et l'origine de la perturbation acoustique. L'applicabilité du modèle analytique à ce cas complexe est aussi étudiée. Les résultats montrent que l'analyse acoustique proposée prédit bien la présence d'un mode instable, mais que le modèle bas ordre nécessite davantage de développements pour étendre son domaine de validité présuméIncreasingly stringent regulations as well as environmental concerns have lead gas turbine powered engine manufacturers to develop the current generation of combustors, which feature lower than ever fuel consumption and pollutant emissions. However, modern combustor designs have been shown to be prone to combustion instabilities, where the coupling between acoustics of the combustor and the flame results in large pressure oscillations and vibrations within the combustion chamber. These instabilities can cause structural damages to the engine or even lead to its destruction. At the same time, considerable developments have been achieved in the numerical simulation domain, and Computational Fluid Dynamics (CFD) has proven capable of capturing unsteady flame dynamics and combustion instabilities for aforementioned engines. Still, even with the current large and fast increasing computing capabilities, time remains the key constraint for these high fidelity yet computationally intensive calculations. Typically, covering the entire range of operating conditions for an industrial engine is still out of reach. In that respect, low order models exist and can be efficient at predicting the occurrence of combustion instabilities, provided an adequate modeling of the flame/acoustics interaction as appearing in the system is available. This essential piece of information is usually recast as the so called Flame Transfer Function (FTF) relating heat release rate fluctuations to velocity fluctuations at a given point. One way to obtain this transfer function is to rely on analytical models, but few exist for turbulent swirling flames. Another way consists in performing costly experiments or numerical simulations, negating the requested fast prediction capabilities. This thesis therefore aims at providing fast, yet reliable methods to allow for low order combustion instabilities modeling. In that context, understanding the underlying mechanisms of swirling flame acoustic response is also targeted. To address this issue, a novel hybrid approach is first proposed based on a reduced set of high fidelity simulations that can be used to determine input parameters of an analytical model used to express the FTF of premixed swirling flames. The analytical model builds on previous works starting with a level-set description of the flame front dynamics while also accounting for the acoustic-vorticity conversion through a swirler. For such a model, validation is obtained using reacting stationary and pulsed numerical simulations of a laboratory scale premixed swirl stabilized flame. The model is also shown to be able to handle various perturbation amplitudes. At last, 3D high fidelity simulations of an industrial gas turbine powered by a swirled spray flame are performed to determine whether a combustion instability observed in experiments can be predicted using numerical analysis. To do so, a series of forced simulations is carried out in en effort to highlight the importance of the two-phase flow flame response evaluation. In that case, sensitivity to reference velocity perturbation probing positions as well as the amplitude and location of the acoustic perturbation source are investigated. The analytical FTF model derived in the context of a laboratory premixed swirled burner is furthermore gauged in this complex case. Results show that the unstable mode is predicted by the acoustic analysis, but that the flame model proposed needs further improvements to extend its applicability range and thus provide data relevant to actual aero-engines
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McNulty, Gregory Scott. "A study of dynamic compressor surge control strategies for a gas turbine engine". Thesis, Massachusetts Institute of Technology, 1993. http://hdl.handle.net/1721.1/47350.
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McCabe, Niall 1971. "A system study on the use of aspirated technology in gas turbine engines". Thesis, Massachusetts Institute of Technology, 2001. http://hdl.handle.net/1721.1/8720.
Texto completoResumen
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2001.Includes bibliographical references (p. 99).
Increasing aircraft engine efficiency and reducing the engines weight have driven innovation in the aircraft engine business since its inception. By simply looking at the Brayton cycle increasing the compressor pressure ratio can bring about an increase in efficiency. To achieve this high pressure ratio, multi-stage axial compressors are used, which tend to be both heavy and expansive. Increasing the number of stages in an axial compressor can increase the pressure ratio and therefore the thermal efficiency; however as the number of stages increases, the engine weight, cost and length also increase, all of which are detrimental to the overall aircraft performance. Recent work by Kerrebrock, Merchant, and Schuler, has led to the possibility of achieving high pressure ratios with a reduction in the number of stages. These compressors use aspiration, or suction on the surface of the blades and endwalls, to keep the boundary layer attached over a greater percentage of the blade chord. Keeping the boundary layer attached longer allows the each blade row to be more highly loaded than the equivalent non-aspirated blade. This higher loading means fewer stages are needed to achieve a given pressure rise. The extracted air is brought inside the blade where it is removed at a convenient location. This bleed air can contain a substantial amount of energy that can be used for numerous purposes on the aircraft or engine. Recovery of the bleed flow and its disposition are important factors in the success of aspirated compressor technology. In this study it is assumed the bleed air can be used for threes purposes: its is returned to the turbine as cooling air, expanded overboard to augment the engine thrust or used to perform "auxiliary work" in a different part of the aircraft. The thermodynamic efficiency (as measured by the specific impulse) and the installed efficiency of the compression system were calculated for different engine/fan configurations and compared with equivalent non-aspirated engines. This allows the effects of aspiration to be quantified and can be used to assess if aspiration is viable for a specific setting.
by Niall McCabe.
S.M.
20
Goh, Shaun Shiao Sing 1980. "Sustainment of commercial aircraft gas turbine engines : an organizational and cognitive engineering approach". Thesis, Massachusetts Institute of Technology, 2003. http://hdl.handle.net/1721.1/82760.
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Napert, Gary Arthur 1956. "Fatigue performance of electroless nickel coatings on stainless steel gas turbine compressor rotors". Thesis, Massachusetts Institute of Technology, 1989. http://hdl.handle.net/1721.1/41318.
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Monroe, Mark A. (Mark Alan). "A market and engineering study of a 3-kilowatt class gas turbine generator". Thesis, Massachusetts Institute of Technology, 2003. http://hdl.handle.net/1721.1/42200.
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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2003.Includes bibliographical references (p. 147-149).
Market and engineering studies were performed for the world's only commercially available 3 kW class gas turbine generator, the IHI Aerospace Dynajet. The objectives of the market study were to determine the competitive requirements for small generators in various U.S. applications, assess the unit's current suitability for these applications, and recommend ways to modify performance or marketing practices to make it more competitive. Engineering study goals included developing an accurate cycle model and assessing the potential for performance improvement. The market study found that the current high selling price precludes competitiveness in most segments of the U.S. civil market. One potential exception may be the marine market, where price sensitivity is less of an issue and a premium is paid for quiet operation, a distinct advantage of the Dynajet. A gas turbine generator solution has more potential in the military market, where the difference from incumbent prices is smaller than in the civil market. The Dynajet is also an appealing military solution because of its high reliability and quiet operation. The market study concluded that increasing power output and efficiency while reducing purchase price would be the most effective approach to improved competitiveness. Alternatively, the current strengths could be leveraged by adapting it for use with an absorption cooler and by emphasizing its superior emission characteristics to consumers and regulators. The engineering study discovered that cycle performance is degraded by secondary nonidealities including flow leakage, heat leakage, and thermal flow distortion. Although these nonidealities are present to some degree in all gas turbines, their impacts are larger in small-scale engines.
(cont.) The net effect of all nonidealities is a 61 percent reduction in power and 12 point decrease in overall efficiency. Analysis concluded that the best way to enhance Dynajet competitiveness is to reduce or remove those nonidealities that are straightforward to fix while increasing power output to either 3 or 5 kW. Output of 5 kW is most promising in terms of cost and weight competitiveness; however, such an improvement may require turbomachinery redesign. A short-term increase of power output to 3 kW appears practical from an engineering standpoint.
by Mark A. Monroe.
S.M.
23
Martini, Bastien. "Development and assessment of a soot emissions model for aircraft gas turbine engines". Thesis, Massachusetts Institute of Technology, 2008. http://hdl.handle.net/1721.1/45256.
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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2008.Includes bibliographical references.
Assessing candidate policies designed to address the impact of aviation on the environment requires a simplified method to estimate pollutant emissions for current and future aircraft gas turbine engines under different design and operating assumptions. A method for NOx and CO emissions was developed in a previous research effort. This thesis focuses on the addition of a soot mechanism to the existing model. The goal is to estimate soot emissions of existing gas turbine engines within soot measurement uncertainties, and then to use the method to estimate the performance of potential future engines. Soot is non-volatile primary particulate matter. In gas turbine engines the size rarely exceeds l [mu]m. The soot is composed almost exclusively of black carbon, is an aggregate of nearly spherical carbon primary particles, and exhibits fractal behavior. Results of other studies regarding soot nucleation, growth, oxidation, and coagulation rates are integrated within a network of perfectly-stirred reactors and shown to capture the typical evolution of soot inside a gas turbine combustor, with soot formed in the early parts of the combustor and then oxidized. The soot model shows promising results as its emissions estimates are within the measurement uncertainties. Nevertheless, model uncertainties are high. They are the consequence of the large sensitivity to input variables. Therefore, the validity of the model is limited to cases with available engine data. More engine data are needed to develop and assess the soot model.
by Bastien Martini.
S.M.
24
Peck, Jhongwoo 1976. "Development of a catalytic combustion system for the MIT Micro Gas Turbine Engine". Thesis, Massachusetts Institute of Technology, 2003. http://hdl.handle.net/1721.1/28292.
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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2003.Includes bibliographical references (p. 71-72).
As part of the MIT micro-gas turbine engine project, the development of a hydrocarbon-fueled catalytic micro-combustion system is presented. A conventionally-machined catalytic flow reactor was built to simulate the micro-combustor and to better understand the catalytic combustion at micro-scale. In the conventionally-machined catalytic flow reactor, catalytic propane/air combustion was achieved over platinum. A 3-D finite element heat transfer model was also developed to assess the heat transfer characteristics of the catalytic micro-combustor. It has been concluded that catalytic combustion in the micro-combustor is limited by diffusion of fuel into the catalyst surface. To address this issue, a catalytic structure with larger surface area was suggested and tested. It was shown that the larger surface area catalyst increased the chemical efficiency. Design guidelines for the next generation catalytic micro-combustor are presented as well.
by Jhongwoo Peck.
S.M.
25
Lackner, Matthew 1980. "Vibration and crack detection in gas turbine engine compressor blades using Eddy current sensors". Thesis, Massachusetts Institute of Technology, 2003. http://hdl.handle.net/1721.1/28895.
Texto completoResumen
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, September 2004.Includes bibliographical references (p. 97).
(cont.) in the ECS signal, no definitive method for sensing blade vibration using an ECS has yet been developed.
High cycle fatigue (HCF) cracks generated by compressor blade vibrations are a common source of failure in gas turbine engines. Current methods for crack detection are costly, time consuming, and prone to errors. In-situ blade vibration detection would help operators avoid critical engine speeds, and help infer the presence of cracks via a change in the mode of a blade. Blade vibrations can be detected using non-contacting sensors like optical sensors, or contacting sensors like strain gauges. These methods have drawbacks that make them poorly suited for installation in a gas turbine engine. Eddy Current Sensors (ECS) have numerous advantages over other vibration detection methods. This thesis aims to use ECS's for vibration detection. Testing was performed in a spin pit rig in the Gas Turbine Lab at the Massachusetts Institute of Technology. The rig contained a rotor on which three test blades spun, and strain gauge and ECS data were extracted from the rig. Magnet arrays were used to provide an excitation force to the blades, causing them to vibrate as they were spinning. Force hammer testing was used to determine the resonant frequencies and mode shapes of the test blades, as well as transfer functions from the strain gauges to the blade tip acceleration. These transfer functions allowed for independent knowledge of the blade vibration behavior. The case of a cracked blade was also considered. Estimates were performed to determine the proper location and length of a crack in the test blade. A 10 mm edge crack was created in a test blade. The crack was found to lower the resonant frequency of the first torsion mode of the blade by 0.2%, and to alter the transfer function between strain and tip acceleration. While some evidence of the blade vibration appears
by Matthew Lackner.
S.M.
26
Borror, Sean Leander. "Natural and forced response measurement of hydrodynamic stability in an aircraft gas turbine engine". Thesis, Massachusetts Institute of Technology, 1994. http://hdl.handle.net/1721.1/47364.
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Mehra, Amitav. "Development of a high power density combustion system for a silicon micro gas turbine engine". Thesis, Massachusetts Institute of Technology, 2000. http://hdl.handle.net/1721.1/9269.
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Thesis (Ph.D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2000."February 2000."
Includes bibliographical references (p. 203-211).
As part of an effort to develop a microfabricated gas turbine engine capable of providing 10-50 Watts of electrical power in a package less than one cubic centimeter in volume, this thesis presents the design, fabrication, packaging and testing of the first combustion system for a silicon micro heat engine. The design and operation of a microcombustor is fundamentally limited by the chemical reaction times of the fuel, by silicon material and fabrication constraints, and by the inherently non-adiabatic nature of the operating space. This differs from the design of a modern macro combustion system that is typically driven by emissions, stability, durability and pattern factor requirements. The combustor developed herein is shown to operate at a power density level that is at least an order of magnitude higher than that of any other power-MEMS device (2000 MW/m 3), and establishes the viability of using high power density, silicon-based combustion systems for heat engine applications at the micro-scale. This thesis presents the development of two specific devices - the first device is a 3-wafer level microcombustor that established the viability of non-premixed hydrogen-air combustion in a volume as small as 0.066 cm 3, and within the structural constraints of silicon; the second device is known as the engine "static-structure", and integrated the 3-stack microcombustor with the other non-rotating components of the engine. Fabricated by aligned fusion bonding of 6 silicon wafers, the static structure measures 2.1 cm x 2.1 cm x 0.38 cm, and was largely fabricated by deep reactive ion etching through a total thickness of 3,800 pm. Packaged with a set of fuel plenums, pressure ports, fuel injectors, igniters, fluidic interconnects, and compressor and turbine static airfoils, this structure is the first demonstration of the complete hot flow path of a multi-level microengine. The 0.195 cm 3 combustion chamber has been tested for several tens of hours and is shown to sustain stable hydrogen combustion with exit gas temperatures above 1600K and combustor efficiencies as high as 95%. The structure also serves as the first experimental demonstration of hydrocarbon microcombustion with power density levels of 500 MW/m 3 and 140 MW/m 3 for ethylene-air and propane-air combustion, respectively. In addition to the development of the two combustion devices, this thesis also presents simple analytical models to identify and explain the primary drivers of combustion phenomena at the micro-scale. The chemical efficiency of the combustor is shown to have a strong correlation with the Damkohler number in the chamber, and asymptotes to unity for sufficiently large values of Da. The maximum power density of the combustor is also shown to be primarily limited by the structural and fabrication constraints of the material. Overall, this thesis synthesizes experimental and computational results to propose a simple design methodology for microcombustion devices, and to present design recommendations for future microcombustor development. Combined with parallel efforts to develop thin-film igniters and temperature sensors for the engine, it serves as the first experimental demonstration of the design, fabrication, packaging and operation of a silicon-based combustion system for power generation applications at the micro-scale.
by Amitav Mehra.
Ph.D.
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Rock, Peter Joseph Jr. "Computational assessment of turbine rim seal system parametric variation on hot gas ingestion and flow pattern". Thesis, Massachusetts Institute of Technology, 2018. http://hdl.handle.net/1721.1/119310.
Texto completoResumen
Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2018.Cataloged from PDF version of thesis.
Includes bibliographical references (pages 125-127).
A design of experiments (DOE) is carried out to assess the turbine rim cavity system parametric variation on hot gas ingestion, flow pattern, and turbine stage efficiency. The parameters focused on are purge to main mass flow rate ratio, axial gap to rim radius ratio, radial gap to rim radius ratio, normalized axial position of the blade leading edge, and internal purge cavity radius ratios. The results were used to formulate a non-dimensional sealing parameter, [psi], that has a threshold value of [psi] = 2.3 - 10¹⁵, beyond which there is only a marginal variation in ingestion penetration depth. This non-dimensional sealing parameter is given as a function of rim seal geometry, purge mass flow rate ratio, Rotational Reynolds number, purge flow Reynolds number, rim seal Reynolds number, and Rossby number. The non-dimensional sealing parameter reflects the physical effects associated with rim seal geometry, flow characteristics, and operating parameters. The computed flow field demonstrates the dominant role of vortical structures in the rim cavity flow on effective flow area distribution and hence the ingestion penetration depth. Quantitative attributes of the vortex, such as non-dimensional circulation, maximum vorticity, height to width ratio, and normalized vortex center position, scale with the non-dimensional sealing parameter. As a result, the vortex attributes scale with ingestion penetration depth. The implication is that the sealing parameter potentially provides a guideline for selecting rim seal configurations and operating space to yield marginal levels of hot gas ingestion. The variation in turbine stage efficiency is approximately linear with purge mass flow rate ratio, where a decrease of 0.7% in efficiency is observed for every 1% increase in purge mass flow rate ratio. This result is in accord with published results to-date.
by Peter Joseph Rock Jr.
S.M.
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Kocer, Gulru. "Aerothermodynamic Modeling And Simulation Of Gas Turbines For Transient Operating Conditions". Master's thesis, METU, 2008. http://etd.lib.metu.edu.tr/upload/12609642/index.pdf.
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In this thesis, development of a generic transient aero-thermal gas turbine model
is presented. A simulation code, gtSIM is developed based on an algorithm which
is composed of a set of differential equations and a set of non-linear algebraic
equations representing each gas turbine engine component. These equations are
the governing equations which represents the aero-thermodynamic process of
the each engine component and they are solved according to a specific solving
sequence which is defined in the simulation code algorithm. At each time step,
ordinary differential equations are integrated by a first-order Euler scheme and
a set of algebraic equations are solved by forward substitution. The numerical
solution process lasts until the end of pre-defined simulation time. The objective
of the work is to simulate the critical transient scenarios for different types of
gas turbine engines at off-design conditions. Different critical transient scenarios
are simulated for two di®erent types of gas turbine engine. As a first simulation, a sample critical transient scenario is simulated for a small turbojet engine. As a second simulation, a hot gas ingestion scenario is simulated for a turbo shaft engine. A simple proportional control algorithm is also incorporated into the simulation code, which acts as a simple speed governor in turboshaft simulations. For both cases, the responses of relevant engine parameters are plotted and results are presented. Simulation results show that the code has the potential to correctly capture the transient response of a gas turbine engine under different operating conditions. The code can also be used for developing engine control algorithms as well as health monitoring systems and it can be integrated to various flight vehicle dynamic simulation codes.
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Everitt, Stewart. "Developments in advanced high temperature disc and blade materials for aero-engine gas turbine applications". Thesis, University of Southampton, 2012. https://eprints.soton.ac.uk/348897/.
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The research carried out as part of this EngD is aimed at understanding the high temperature materials used in modern gas turbine applications and providing QinetiQ with the information required to assess component performance in new propulsion systems. Performance gains are achieved through increased turbine gas temperatures which lead to hotter turbine disc rims and blades. The work has focussed on two key areas: (1) Disc Alloy Assessment of High Temperature Properties; and (2) Thermal Barrier Coating Life Assessment; which are drawn together by the overarching theme of the EngD: Lifing of Critical Components in Gas Turbine Engines. Performance of sub-solvus heat treated N18 alloy in the temperature range of 650°C to 725°C has been examined via monotonic and cyclically stabilised tensile, creep and strain controlled low cycle fatigue (LCF) tests including LCF behaviour in the presence of a stress concentration under load-control. Crack propagation studies have been undertaken on N18 and a particular super-solvus heat treatment variant of the alloy LSHR at the same temperatures, in air and vacuum with 1s and 20s dwell times. Comparisons between the results of this testing and microstructural characterisation with RR1000, UDIMET® 720 Low Interstitial (U720Li) and a large grain variant of U720Li have been carried out. In all alloys, strength is linked to a combination of γ' content and grain size as well as slow diffusing atoms in solid solution. High temperature strength improves creep performance which is also dependent on grain size and grain boundary character. Fatigue testing revealed that N18 had the most transgranular crack propagation with a good resistance to intergranular failure modes, with U720Li the most intergranular. Under vacuum conditions transgranular failure modes are evident to higher temperature and ΔK, with LSHR failing almost completely by intergranular crack propagation in air. For N18 significant cyclic softening occurs at 725°C with LCF initiation occurring at pores and oxidised particles. An apparent activation energy technique was used to provide further insights into the failure modes of these alloys, this indicating that, for N18 with 1s dwell, changes in fatigue crack growth rates were attributed to static properties and for LSHR, with 20s dwell in air, that changes were attributed to the detrimental synergistic combination of creep and oxidation at 725°C. Microchemistry at grain boundaries, especially M23C6 carbides, plays an important role in these alloys. Failure mechanisms within a thermal barrier coating (TBC) system consisting of a CMSX4 substrate, PtAl bond coat, thermally grown oxide (TGO) layer and a top coat applied using electron beam physical vapour deposition have been considered. TGO growth has been quantified under isothermal, two stage temperature and thermal cyclic exposures. An Arrhenius relation was used to describe the TGO growth and produce an isothermal TGO growth model. The output from this was used in the QinetiQ TBC Lifing Model. Thermo-mechanical fatigue test methods were also developed including a novel thermocouple placement permitting substrate temperature to be monitored without disturbing the top coat such that the QinetiQ TBC Lifing Model could be validated. The importance of material, system specific knowledge and performance data with respect to a particular design space for critical components in gas turbine engines has been highlighted. Data and knowledge regarding N18, LSHR and TBC systems has been added to the QinetiQ’s databank enhancing their capability for providing independent advice regarding high temperature materials particularly in new gas turbine engines.
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Stanley, Felix. "Dimensional reduction and design optimization of gas turbine engine casings for tip clearance studies". Thesis, University of Southampton, 2010. https://eprints.soton.ac.uk/342789/.
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The objective of this research is to develop a design process that can optimize an engine casing assembly to reduce tip clearance losses. Performing design optimization on the casings that form a gas turbine engine's external structure is a very tedious and cumbersome process. The design process involves the conceptual, the preliminary and the detailed design stages. The redesign costs involved are high when changes are made to the design of a part in the detailed design stage. Normally a 2D configuration is envisaged by the design team in the conceptual design stage. Engine thrust, mass flow, operating temperature, materials and manufacturing processes available at the time of design, mass of the engine, loads and assembly conditions are a few of the many important variables that are taken into consideration when designing an aerospace component. The linking together of this information into the design process to achieve an optimal design using a quick robust method is still a daunting task. In this thesis, we present techniques to extract midsurfaces of complex 3D axisymmetric and non-axisymmetric geometries based on medial axis transforms. We use the proposed FE modeling technique for optimizing the geometry by designing a sequential workflow consisting of CAD, FE analysis and optimization algorithms within an integrated system. An existing commercial code was first used to create a midsurface shell model and the results showed that such models could replace 3D models for defection studies. These softwares being black box codes could not be customized. Such limitations restrict their use in batch mode and development for research purposes. We recognized an immediate need to develop a bespoke code that could be used to extract midsurfaces for FE modeling. Two codes, Mantle-2D and Mantle-3D have been developed using Matlab to handle 3D axisymmetric and non-axisymmetric geometries respectively. Mantle-2D is designed to work with 2D cross-section geometry as an input while Mantle-3D deals with complex 3D geometries. The Pareto front (PF) of 2000 designs of the shell based optimization problem when superimposed on the PF of the solid based optimization, has provided promising results. A DoE study consisting of 200 designs was also conducted and results showed that the shell model differs in mass and defection by <1% and <5.0% respectively. The time taken to build/solve a solid model varied between 45-75 minutes while the equivalent midsurface based shell model built using Mantle-2D required only 3-4 minutes. The Mantle-3D based dimensional reduction process for a complex non-axisymmetric solid model has also been demonstrated with encouraging results. This code has been used to extract and mesh the midsurface of a non-axisymmetric geometry with shell elements for use in finite element analysis. 101 design points were studied and the results compared with the corresponding solid model. The first 10 natural frequencies of the resulting shell model deviates from the solid model by <4.0% for the baseline design, while the mass and defection errors were <3.5% and <9.0% for all 101 design points.
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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.
Texto completoResumen
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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Giles, M. (Michael). "Newton solution of steady two-dimensional transonic flow". Thesis, Massachusetts Institute of Technology, 1985. http://hdl.handle.net/1721.1/15250.
Texto completoResumen
Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1985.MICROFICHE COPY AVAILABLE IN ARCHIVES AND AERONAUTICS.
Bibliography: leaves 167-169.
by Michael Bryce Giles.
Ph.D.
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Tanaka, Shinji S. M. Massachusetts Institute of Technology. "Acoustic and thermal packaging of small gas turbines for portable power". Thesis, Massachusetts Institute of Technology, 2009. http://hdl.handle.net/1721.1/51648.
Texto completoResumen
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2009.Includes bibliographical references (p. 201-203).
To meet the increasing demand for advanced portable power units, for example for use in personal electronics and robotics, a number of studies have focused on portable small gas turbines. This research is concerned with gas turbine generator units in the 1 kW range. The compact and small-scale architecture of the portable gas turbine engine poses major challenges in the acoustic treatment that is required to attenuate the broadband and tonal noise of the high-speed turbomachinery. The challenge in the thermal management is the relatively large required cooling mass flow and the short flow mixing length, constrained by package size considerations. The objective is to conceive a proof-of-concept engine package with exhaust temperatures of 60 °C and a noise signature below 50 dBA at a distance of 7 m. Various liner materials and configurations were investigated in an anechoic chamber using a modular silencer test rig. Acoustic liners based on porous fiber material were developed for both cold intake and hot exhaust gas silencers to reduce the broadband noise. The source noise simulations combined with the measured silencer noise reduction show noise levels below 50 dBA in all directions. A parametric silencer configuration study was carried out to determine the trade-off between liner volume, surface area, and noise reduction. The liner material was demonstrated to withstand hot gas conditions at 700 °C.
(cont.) A mixer/ejector based cooling scheme was proposed and experimentally investigated using vortex generator rings and multi-walled ejectors to enhance the mixing. Although the augmentations achieved a satisfactory mass flow ratio of 16.8:1, hot spots still exist at the exit of the relatively long mixer duct due to the high area-ratio of the ejector configuration. It was concluded that implementation of the scheme into the package is not practical. To overcome this mixing challenge, an alternative cooling scheme was conceived. An inverted dilution liner mixes hot core gas flowing radially through a perforated cylinder with cold fan air. The mixing length is reduced due to jet induced streamwise vortices. The performance of the device was investigated using three-dimensional computational fluid dynamics simulations, which demonstrated improved mixing and uniform, low temperatures of less than 70 °C at the mixer exit. Noise reduction and flow mixing guidelines are established together with a concept package configuration, generally applicable to small scale gas turbine devices.
by Shinji Tanaka.
S.M.
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Leung, Kai Yuen Eric. "3D turbine tip clearance flow redistribution due to gap variation". Thesis, Massachusetts Institute of Technology, 1991. http://hdl.handle.net/1721.1/42513.
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Ozmen, Teoman. "Gas Turbine Monitoring System". Master's thesis, METU, 2006. http://etd.lib.metu.edu.tr/upload/12607957/index.pdf.
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In this study, a new gas turbine monitoring system being able to carry out appropriate run process is set up for a gas turbine with 250 kW power rating and its accessories. The system with the mechanical and electrical connections of the required sub-parts is transformed to a kind of the test stand. Performance test result calculation method is described. In addition that, performance evaluation software being able to apply with the completion of the preliminary performance tests is developed for this gas turbine.
This system has infrastructure for the gas turbine sub-components performance and aerothermodynamics research. This system is also designed for aviation training facility as a training material for the gas turbine start and run demonstration. This system provides the preliminary gas turbine performance research requirements in the laboratory environment.
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Flesland, Synnøve Mangerud. "Gas Turbine Optimum Operation". Thesis, Norges teknisk-naturvitenskapelige universitet, Institutt for energi- og prosessteknikk, 2010. http://urn.kb.se/resolve?urn=urn:nbn:no:ntnu:diva-12409.
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Many offshore installations are dependent on power generated by gas turbines and a critical issue is that these experience performance deterioration over time. Performance deterioration causes reduced plant efficiency and power output as well as increased environmental emissions. It is therefore of highest importance to detect and control recoverable losses in order to reduce their effect. This thesis project was therefore initiated to evaluate parameters for detecting performance deterioration in addition to document different aspects of gas turbine degradation and performance recovery. Compressor fouling is the largest contributor to performance deterioration. Investigating fouling was therefore the main focus of this study.In the present study the deterioration rates of four different gas turbines were evaluated. When choosing gas turbines it was emphasised to select gas turbines operating under equal conditions but with different washing procedures. In addition to offline washing two of the gas turbines had daily online washing routines and one of the gas turbines run idle wash every 1000 hour between each offline wash. Data was extracted from the monitoring software, TurboWatch, and loaded into Excel files. MATLAB scripts were created to handle the large amount of data and visualize performance trends. Series of two parameters were plotted against each other and the graphs were evaluated.The evaluation showed that an overall trend was that the gas turbine that had been running with online washing continuously over a long period of time had higher performance than the reference engine. For the second gas turbine a daily online washing procedure has recently started. The advantage with the evaluation of this gas turbine was that a good reference engine was available. The two engines were operating under quite similar conditions at the same location in addition to having equal filter systems. Some deterioration trends were possible to detect. For the first period both engines seemed to have quite equal deterioration trends. During the second period no clear trends were seen in corrected CDP and corrected EGT when evaluated for constant GG speed. The compressor efficiency had decreasing trends for both engines during the second period as well, but the compressor efficiency for machine 1 was overall higher during the period with online washing than the previous period. The borescope pictures taken after the first period with online washing showed good visual results. However, it is too premature to make a final decision regarding the exact performance gain of online washing. At the time the study was performed the engine had only been running online washing for one operating interval, and more investigation over longer time is recommended. For the engine running with idle wash it was not possible to conclude on the basis of the collected data. No clear deterioration trends were detected and investigations over longer time and several operating intervals are recommended. It is also important to be aware of the fact that the performance gain of idle wash needs to be much higher than for online washing in order for idle wash to be economically profitable. There are several uncertainties related to performance trends. These include inaccuracy in instrumentation, monitoring software, calibration etc. Due to the fact that all the gas turbines evaluated in this study only have standard instrumentation it caused additional uncertainty in the performance trends. One suggestion for further study is to initiate a test instrumented gas turbine into operation with sensors for measuring inlet pressure depression
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Spencer, Matthew Richard. "Gas turbine lubricant evaluation". Thesis, University of Birmingham, 2014. http://etheses.bham.ac.uk//id/eprint/5423/.
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This thesis is a study of the chemical and physical changes which can occur to gas turbine lubricants as a result of exposure to operational conditions. The continual evolution toward more efficient gas turbines is accompanied by increasing thermal and mechanical loading which the lubricant must be able to withstand. In this thesis two major degradation issues are studied; thermal oxidative degradation and lubricant deposition. In the area of thermal oxidative degradation, efforts are made to better understand the key parameters which determine the lubricant breakdown mechanism. Through control of these parameters and comparison to service derived gas turbine oil samples a new laboratory methodology is proposed for the assessment of lubricant oxidative degradation. The study of lubricant deposition in this thesis is concentrated on the regions of highest risk, the bearing chamber feed (single phase) and vent (two phase) oil pipes. Development of existing laboratory scale deposition simulators was conducted to increase how engine representative the methods are of gas turbine conditions. These simulators were used to evaluate the rate of deposition with a range of lubricants, simulated engine cycles and pipe surfaces.
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Bartlett, Michael. "Developing Humidified Gas Turbine Cycles". Doctoral thesis, KTH, Chemical Engineering and Technology, 2002. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-3437.
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As a result of their unique heat recovery properties,Humidified Gas Turbine (HGT) cycles have the potential todeliver resource-effective energy to society. The EvaporativeGas Turbine (EvGT) Consortium in Sweden has been studying thesetypes of cycles for nearly a decade, but now stands at acrossroads, with commercial demonstration remaining. Thisthesis binds together several key elements for the developmentof humidified gas turbines: water recovery and air and waterquality in the cycle, cycle selection for near-term, mid-sizedpower generation, and identifying a feasible niche market fordemonstration and market penetration. Moreover, possiblesocio-technical hinders for humidified gas turbine developmentare examined.
Through modelling saltcontaminant flows in the cycle andverifying the results in the pilot plant, it was found thathumidification tower operation need not endanger the hot gaspath. Moreover, sufficient condensate can be condensed to meetfeed water demands. Air filters were found to be essential tolower the base level of contaminant in the cycle. This protectsboth the air and water stream components. By capturing airparticles of a similar size to the air filters, the humidifieractually lowers air stream salt levels. Measures to minimisedroplet entrainment were successful (50 mg droplets/kg air) andmodels predict a 1% blow down from the water circuit issufficient. The condensate is very clean, with less than 1 mg/lalkali salts and easily deionised.
Based on a core engine parameter analysis for three HGTcycle configurations and a subsequent economic study, asteam-cooled steam injected cycle complemented with part-flowhumidification is recommended for the mid-size power market.This cycle was found to be particularly efficient at highpressures and turbine inlet temperatures, conditions eased bysteam cooling and even intercooling. The recommended HGT cyclegives specific investment costs 30- 35% lower than the combinedcycles and cost of electricity levels were 10-18% lower.Full-flow intercooled EvGT cycles give high performances, butseem to be penalised by the recuperator costs, while stillbeing cheaper than the CC. District heating is suggested as asuitable niche market to commercially demonstrate the HGTcycle. Here, the advantages of HGT are especially pronounceddue their very high total efficiencies. Feasibility prices forelectricity were up to 35% lower than competing combinedcycles. HGT cycles were also found to effectively include wasteheat sources.
Keywords:gas turbines, evaporative gas turbines,humidification, power generation, combined heat and powergeneration.
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Pachidis, Vassilios A. "Gas turbine advanced performance simulation". Thesis, Cranfield University, 2006. http://dspace.lib.cranfield.ac.uk/handle/1826/4529.
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Current commercial 'state of the art' engine simulation software is of a low fidelity. Individual component performance characteristics are typically represented via nondimensional maps with empirical adjustments for off-design effects. Component nondimensional characteristics are usually obtained through the averaging of experimental readings from rig test analyses carried out under nominal operating conditions. In those cases where actual component characteristics are not available and default maps are used instead, conventional simulation tools can offer a good prediction of the performance of the whole engine close to design point, but can deviate substantially at of design and transient conditions. On the other hand, even when real component characteristics are available, zero-dimensional engine cycle simulation tools can not predict the performance of the engine at other than nominal conditions satisfactorily. Low-fidelity simulation tools are generally incapable of analyzing the performance of individual engine components in detail, or capturing complex physical phenomena such as inlet flow distortion. Although the available computational power has increased exponentially over the last two decades, a detailed, three-dimensional analysis of an entire propulsion system still seems to be so complex and computationally intensive as to remain cost-prohibitive. For this reason, alternative methods of integrating different types and levels of analysis are necessary. The integration of simulation codes that model at different levels of fidelity into a single simulation provides the opportunity to reduce the overall computing resource needed, while retaining the desired level of analysis in specific engine components. The objective of this work was to investigate different simulation strategies for communicating the performance characteristics of an isolated gas turbine engine component, resolved from a detailed, high-fidelity analysis, to an engine system analysis carried out at a lower level of resolution. This would allow component-level, complex physical processes to be captured and analyzed in the context of the whole engine performance, at an affordable computing resource and time. More specifically, this work identified and thoroughly investigated several advanced simulation strategies in terms of their actual implementation and potential, by looking into relative changes in engine performance after integrating into the basic, nondimensional cycle analysis, the performance characteristics of i) two-dimensional Streamline Curvature (SLC) and ii) three-dimensional Computational Fluid Dynamics (CFD), engine component models. In the context of this work, several case studies were carried out, utilising different two-dimensional and three-dimensional component geometries, under different operating conditions, such as different types and extents of compressor inlet pressure distortion and turbine inlet temperature distortion. More importantly, this research effort established the necessary methodology and technology required for a full, twodimensional engine cycle analysis at an affordable computational resource.
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Pachidis, Vassilios. "Gas Turbine Advanced Performance Simulation". Thesis, Cranfield University, 2006. http://hdl.handle.net/1826/4529.
Texto completoResumen
Current commercial `state of the art' engine simulation software is of a low fidelity.
Individual component performance characteristics are typically represented via nondimensional
maps with empirical adjustments for off-design effects. Component nondimensional
characteristics are usually obtained through the averaging of
experimental readings from rig test analyses carried out under nominal operating
conditions.
In those cases where actual component characteristics are not available and default
maps are used instead, conventional simulation tools can offer a good prediction of
the performance of the whole engine close to design point, but can deviate
substantially at of design and transient conditions. On the other hand, even when real
component characteristics are available, zero-dimensional engine cycle simulation
tools can not predict the performance of the engine at other than nominal conditions
satisfactorily. Low-fidelity simulation tools are generally incapable of analyzing the
performance of individual engine components in detail, or capturing complex physical
phenomena such as inlet flow distortion.
Although the available computational power has increased exponentially over the last
two decades, a detailed, three-dimensional analysis of an entire propulsion system still
seems to be so complex and computationally intensive as to remain cost-prohibitive.
For this reason, alternative methods of integrating different types and levels of
analysis are necessary. The integration of simulation codes that model at different
levels of fidelity into a single simulation provides the opportunity to reduce the
overall computing resource needed, while retaining the desired level of analysis in
specific engine components.
The objective of this work was to investigate different simulation strategies for
communicating the performance characteristics of an isolated gas turbine engine
component, resolved from a detailed, high-fidelity analysis, to an engine system
analysis carried out at a lower level of resolution. This would allow component-level,
complex physical processes to be captured and analyzed in the context of the whole
engine performance, at an affordable computing resource and time.
More specifically, this work identified and thoroughly investigated several advanced
simulation strategies in terms of their actual implementation and potential, by looking
into relative changes in engine performance after integrating into the basic, nondimensional
cycle analysis, the performance characteristics of i) two-dimensional
Streamline Curvature (SLC) and ii) three-dimensional Computational Fluid Dynamics
(CFD), engine component models.
In the context of this work, several case studies were carried out, utilising different
two-dimensional and three-dimensional component geometries, under different
operating conditions, such as different types and extents of compressor inlet pressure
distortion and turbine inlet temperature distortion. More importantly, this research
effort established the necessary methodology and technology required for a full, twodimensional
engine cycle analysis at an affordable computational resource.
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Spencer, A. "Gas turbine combustor port flows". Thesis, Loughborough University, 1998. https://dspace.lboro.ac.uk/2134/6883.
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Competitive pressure and stringent emissions legislation have placed an urgent demand on research to improve our understanding of the gas turbine combustor flow field. Flow through the air admission ports of a combustor plays an essential role in determining the internal flow patterns on which many features of combustor performance depend. This thesis explains how a combination of experimental and computational research has helped improve our understanding, and ability to predict, the flow characteristics of jets entering a combustor. The experiments focused on a simplified generic geometry of a combustor port system. Two concentric tubes, with ports introduced into the inner tube's wall, allowed a set of radially impinging jets to be formed within the inner tube. By investigating the flow with LDA instrumentation and flow visualisation methods a quantitative and qualitative picture of the mean and turbulent flow fields has been constructed. Data were collected from the annulus, port and core regions. These data provide suitable validation information for computational models, allow improved understanding of the detailed flow physics and provide the global performance parameters used traditionally by combustor designers. Computational work focused on improving the port representation within CFD models. This work looked at the effect of increasing the grid refinement, and improving the geometrical representation of the port. The desire to model realistic port features led to the development of a stand-alone port modelling module. Comparing calculations of plain-circular ports to those for more realistic chuted port geometry, for example, showed that isothermal modelling methods were able to predict the expected changes to the global parameters measured. Moreover, these effects are seen to have significant consequences on the predicted combustor core flow field.
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Ahmad, N. T. "Swirl stabilised gas turbine combustion". Thesis, University of Leeds, 1986. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.356423.
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Pishva, S. M. R. (S Mohammed Reza) Carleton University Dissertation Engineering Mechanical. "Rejuvenation of gas turbine discs". Ottawa, 1988.
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Asere, Abraham Awolola. "Gas turbine combustor wall cooling". Thesis, University of Leeds, 1986. http://etheses.whiterose.ac.uk/2590/.
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The need for better methods of cooling gas turbine combustors and a review of current cooling techniques have been presented. Three cooling methods are investigated: (a) Full Coverage Discrete Hole Film Cooling (Effusion), (b) Impingement/Effusion Hybrid Cooling Systems, and (c) Transpiration Cooling. The aim of these cooling techniques is to effectively and efficiently cool gas turbine combustors with a significant reduction in current cooling air requirements. The range of test conditions were coolant temperature, Tc, of 289 < Tc 710 K and combustion gases temperature, Tg, of 500 Tg N< 1900 K. The discharge coefficients of the effusion and the impingement/effusion systemshave also been studied. A detailed analysis has been made of the heat transfer of the cooling systems, jet penetration into the cross-stream, prediction of the cooling jet temperatures at various stages in the cooling process and the cooling film heat transfer coefficient. The results of the discharge coefficient (Cd) indicate a decreasing C with increasing wall thickness to diameter ratio, t/D, and a weak effect of cross-stream flow. The results of both the effusion and the impingement/effusion hybrid systems indicate a high cooling performance of similar magnitude to that of the transpiration system. Graphical design correlations for the cooling wall have been made. The optimum hole geometries for both cooling configurations have been developed. The influence of the coolant to hot gas density ratio has been studied over the range 1.4-3.4. In the design of effusion and impingement/effusion cooling systems, wall thickness, hole density, hole diameter and wall design pressure loss are significant parameters for cooling performance maximisation.
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Løver, Kristian Aase. "Biomass gasification integration in recuperative gas turbine cycles and recuperative fuel cell integrated gas turbine cycles : -". Thesis, Norwegian University of Science and Technology, Department of Energy and Process Engineering, 2007. http://urn.kb.se/resolve?urn=urn:nbn:no:ntnu:diva-9658.
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A multi-reactor, multi-temperature, waste-heat driven biomass thermochemical converter is proposed and simulated in the process simulation tool Aspen Plus. The thermochemical converter is in Aspen Plus integrated with a gas turbine power cycle and a combined fuel cell/gas turbine power cycle. Both power cycles are recuperative, and supply the thermochemical converter with waste heat. For result comparison, the power cycles are also integrated with a reference conventional single-reactor thermochemical converter, utilizing partial oxidation to drive the conversion process. Exergy analysis is used for assessment of the simulation results. In stand-alone simulation, the proposed thermochemical shows high performance. Cold gas efficiency is 108.0% and syngas HHV is 14.5 MJ/kg on dry basis. When integrated with the gas turbine power cycle, the proposed converter fails to improve thermal efficiency of the integrated cycle significantly, compared to reference converter. Thermal efficiency is 41.8% and 40.7%, on a biomass HHV basis, with the proposed and the reference converter respectively. This is despite superior cold gas efficiency for the proposed converter, and the gas turbine cycle is found not to be able to properly take advantage of the high chemical energy in the syngas of the proposed converter. When integrated with the combined fuel cell/gas turbine power cycle, the proposed converter significantly improves the thermal efficiency of the integrated cycle, compared to the reference converter. Thermal efficiency is 56.0% and 51.2%, on a biomass HHV basis, with the proposed and the reference converter respectively. The fuel cell is found to be able to take advantage of the high chemical energy in the syngas of the proposed converter, which is the main cause of increase in thermal efficiency. Operation of the proposed thermochemical converter is found to be feasible at a wide range of operating conditions, although low operating temperatures in the converter may cause problems at very high carbon conversion ratios.
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Eskner, Mats. "Mechanical Behaviour of Gas Turbine Coatings". Doctoral thesis, KTH, Materials Science and Engineering, 2004. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-3776.
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Coatings are frequently applied on gas turbine components inorder to restrict surface degradation such as corrosion andoxidation of the structural material or to thermally insulatethe structural material against the hot environment, therebyincreasing the efficiency of the turbine. However, in order toobtain accurate lifetime expectancies and performance of thecoatings system it is necessary to have a reliableunderstanding of the mechanical properties and failuremechanisms of the coatings.
In this thesis, mechanical and fracture behaviour have beenstudied for a NiAl coating applied by a pack cementationprocess, an air-plasma sprayed NiCoCrAlY bondcoat, a vacuumplasma-sprayed NiCrAlY bondcoat and an air plasma-sprayed ZrO2+ 6-8 % Y2O3topcoat. The mechanical tests were carried out ata temperature interval between room temperature and 860oC.Small punch tests and spherical indentation were the testmethods applied for this purpose, in which existing bending andindentation theory were adopted for interpretation of the testresults. Efforts were made to validate the test methods toensure their relevance for coating property measurements. Itwas found that the combination of these two methods givescapability to predict the temperature dependence of severalrelevant mechanical properties of gas turbine coatings, forexample the hardness, elastic modulus, yield strength, fracturestrength, flow stress-strain behaviour and ductility.Furthermore, the plasma-sprayed coatings were tested in bothas-coated and heat-treated condition, which revealedsignificant difference in properties. Microstructuralexamination of the bondcoats showed that oxidation with loss ofaluminium plays an important role in the coating degradationand for the property changes in the coatings.
Keywords:small punch test, miniaturised disc bendingtests, spherical indentation, coatings, NiAl, APS-NiCoCrAlY,VPS-NiCrAlY, mechanical properties
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Cavaliere, Davide Egidio. "Blow-off in gas turbine combustors". Thesis, University of Cambridge, 2014. https://www.repository.cam.ac.uk/handle/1810/265575.
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This thesis describes an experimental investigation of the flame structure close to the extinction and the blow-off events of non-premixed and spray flames stabilized on an axisymmetric bluff body in a confined swirl configuration. The comparison of flames of different canonical types in the same basic aerodynamic field allows insights on the relative blow-off behaviour. The first part of the thesis describes several velocity measurements in non-reacting and reacting flows. The main usefulness of this data is to provide the aerodynamic flow pattern and some discussion on the velocity field and the related recirculation zones. The velocity and turbulence information obtained are particularly useful for providing data, which is crucial for validation of computational models. The second part describes an experimental investigation of non-premixed stable flames very close to the blow-off condition. The measurements included visualisation of the blow-off transient with 5 kHz OH* chemiluminescence, which allowed a quantification of the average duration of the blow-off transient. OH-PLIF images at 5 kHz for flames far from and close to extinction showed that the non-premixed flame intermittently lifts-off the bluff body, with increasing probability as the fuel velocity increases. The flame sheet shows evidence of localised extinctions, which are more pronounced as approaching blow-off. The measurements include blow-off limits and their attempted correlation. It was found that a correlation based on a Damkohler number does a reasonable job at collapsing the dataset. The final part examines the blow-off behaviour of swirling spray flames for two different fuels: n-heptane and n-decane. The measurements include blow-off limits and their att~mpted correlation, visualisation of the blow-off transient with 5 kHz OH* chemiluminescence, and the quantification of the average duration of the blow-off transient. It was found that the average duration of the blow-off event is in order of the tens of ms for both spray flames (10-16 ms). The blow-off event is therefore a relatively slow process for the spray ~ames using n-heptane and decane fuels. This suggests that control measures, such as fast fuel injection, coupled with appropriate detection, such as with chemiluminescence monitoring, may have a reasonable chance of success in keeping the flame alight very close to the blow-off limit. These results, together with those obtained for the non-premixed gaseous case form a wide body of experimental data available for the validation of turbulent flame models. The quantification of some properties during the blow-off transient can assist studies of extinction based on large-eddy simulation that have a promise of capturing combustion transients.
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Bengtsson, Karl. "ThermoacousticInstabilities in a Gas Turbine Combustor". Thesis, KTH, MWL Marcus Wallenberg Laboratoriet, 2017. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-226530.
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Stationary gas turbines are widely used today for power generation and mechanical drive applications. The introduction of new regulations on emissions in the last decades have led to extensive development and new technologies used within modern gas turbines. The majority of the gas turbines sold today have a so called DLE (Dry Low Emission) combustion system that mainly operates in the leanpremixed combustion regime. The lean-premixed regime is characterized by low emission capabilities but are more likely to exhibit stability issues compared to traditional non-premixed combustion systems. Thermoacoustic instabilities are a highly unwanted phenomena characterized by an interaction between an acoustic eld and a combustion process. This interaction may lead to self-sustained large amplitude oscillations which can cause severe structural damage to the gas turbine if it couples with a structural mode. However, since a coupled phenomena, prediction of thermoacoustic stability is a complex topic still under research. In this work, the mechanisms responsible for thermoacoustic instabilities are described and a 1- dimensional stability modelling approach is applied to the Siemens SGT-750 combustion system. The complete combustor is modelled by so called acoustic two-port elements in which a 1-dimensional ame model is incorporated. The simulations is done using a generalized network code developed by Siemens. The SGT-750 shows today excellent stability and combustion performance but a deeper knowledge in the thermoacoustic behaviour is highly valued for future development. In addition, measurement data from an engine test is evaluated, post-processed and compared with the results from the 1-dimensional network model. The results are found to be in good agreement and the thermoacoustic response of the SGT-750 is found to be dominated by both global modes including all cans as well as local modes within the individual cans.
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Murthy, J. N. "Gas turbine combustor modelling for design". Thesis, Cranfield University, 1988. http://hdl.handle.net/1826/2626.
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The design and development of gas turbine combustors is a crucial but uncertain part of an engine development process. Combustion within a gas turbine is a complex interaction of, among other things, fluid dynamics, heat and mass transfer and chemical kinetics. At present, the design process relies upon a wealth of experimental data and correlations. The proper use of this information requires experienced combustion engineers and even for them the design process is very time consuming.
Some major engine manufacturers have attempted to address the above problem by developing one dimensional computer programs based on the above test and empirical data to assist combustor designers. Such programs are usually proprietary. The present work, based on this approach has yielded DEPTH, a combustor design program. DEPTH ( Design and Evaluation of Pressure, Temperature and Heat transfer in combustors) is developed in Fortran-77 to assist in preliminary design and evaluation of conventional gas turbine combustion chambers.
DEPTH can be used to carry out a preliminary design along with prediction of the cooling slots for a given metal temperature limit or to evaluate heat transfer and temperatures for an existing combustion chamber. Analysis of performance parameters such as efficiency, stability and NOx based on stirred reactor theories is also coupled. DEPTH is made sufficiently interactive/user-friendly such that no prior expertise is required as far as computer operation is concerned. The range of variables such as operating conditions, geometry, hardware, fuel type can all be effectively examined and their contribution towards the combustor performance studied. Such comprehensive study should provide ample opportunity for the designer to make the right decisions. It should also be an effective study aid. Returns in terms of higher thermal efficiencies is an incentive to go for combined cycles and cogeneration. In such cases, opting for higher cycle pressures together with a second or reheat combustor promise higher thermal efficiencies and exhaust temperatures and hence such designs are likely to be of interest. The concepts that are needed for understanding a double or reheat combustor are also addressed using the programme. A specific application of the programme is demonstrated through the design of a double combustor.