Дисертації з теми "Aeronautical turbine"
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Koupper, Charlie. "Unsteady multi-component simulations dedicated to the impact of the combustion chamber on the turbine of aeronautical gas turbines." Phd thesis, Toulouse, INPT, 2015. http://oatao.univ-toulouse.fr/14187/1/koupper_partie_1_sur_2.pdf.
Повний текст джерелаAl-Khudairi, Othman. "Structural performance of horizontal axis wind turbine blade." Thesis, Kingston University, 2014. http://eprints.kingston.ac.uk/32197/.
Повний текст джерелаDupuy, Fabien. "Reduced Order Models and Large Eddy Simulation for Combustion Instabilities in aeronautical Gas Turbines." Thesis, Toulouse, INPT, 2020. http://www.theses.fr/2020INPT0046.
Повний текст джерелаIncreasingly 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
Elfarra, Monier A. K. "Horizontal Axis Wind Turbine Rotor Blade: Winglet And Twist Aerodynamic Design And Optimization Using Cfd." Phd thesis, METU, 2011. http://etd.lib.metu.edu.tr/upload/12612987/index.pdf.
Повний текст джерелаSharma has shown the best agreement with measurements. Launder &ndash
Sharma was chosen for further simulations and for the design process. Before starting the design and optimization, different winglet configurations were studied. The winglets pointing towards the suction side of the blade have yielded higher power output. Genetic algorithm and artificial neural network were implemented in the design and optimization process. The optimized winglet has shown an increase in power of about 9.5 % where the optimized twist has yielded to an increase of 4%. Then the stall regulated blade has been converted into pitch regulated blade to yield more power output. The final design was produced by a combination of the optimized winglet, optimized twist andbest pitch angle for every wind speed. The final design has shown an increase in power output of about 38%.
Effendy, Marwan. "Investigation of turbine blade trailing edge cooling and thermal mixing characteristics." Thesis, Kingston University, 2014. http://eprints.kingston.ac.uk/30604/.
Повний текст джерелаOksuz, Ozhan. "Multiploid Genetic Algorithms For Multi-objective Turbine Blade Aerodynamic Optimization." Phd thesis, METU, 2007. http://etd.lib.metu.edu.tr/upload/12609196/index.pdf.
Повний текст джерелаAkagi, Raymond. "Ram Air-Turbine of Minimum Drag." DigitalCommons@CalPoly, 2021. https://digitalcommons.calpoly.edu/theses/2261.
Повний текст джерелаNotarianni, Gianmarco. "Analysis and modelling of the turbocharger behavior of an internal combustion engine for aeronautical application." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2019.
Знайти повний текст джерелаGorgulu, Ilhan. "Numerical Simulation Of Turbine Internal Cooling And Conjugate Heat Transfer Problems With Rans-based Turbulance Models." Master's thesis, METU, 2012. http://etd.lib.metu.edu.tr/upload/12615000/index.pdf.
Повний текст джерелаmodel, Shear Stress Transport k-&omega
model, Reynolds Stress Model and V2-f model, which became increasingly popular during the last few years, have been used at the numerical simulations. According to conducted analyses, despite a few unreasonable predictions, in the majority of the numerical simulations, V2-f model outperforms other first-order turbulence models (Realizable k-&epsilon
and Shear Stress Transport k-&omega
) in terms of accuracy and Reynolds Stress Model in terms of convergence.
Boiani, Davide. "Finite element structural and thermal analysis of JT9D turbofan engine first stage turbine blade." Bachelor's thesis, Alma Mater Studiorum - Università di Bologna, 2017. http://amslaurea.unibo.it/12566/.
Повний текст джерелаKocer, 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.
Повний текст джерелаerent 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.
Mercan, Bayram. "Experimental Investigation Of The Effects Of Waveform Tip Injection On The Characteristics Of Tip Leakage Vortex In A Lpt Cascade." Master's thesis, METU, 2012. http://etd.lib.metu.edu.tr/upload/12614111/index.pdf.
Повний текст джерелаNovikov, 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.
Повний текст джерелаDogan, Eda. "Experimental Investigation Of Boundary Layer Separation Control Using Steady Vortex Generator Jets On Low Pressure Turbines." Master's thesis, METU, 2012. http://etd.lib.metu.edu.tr/upload/12614485/index.pdf.
Повний текст джерелаMartinez-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.
Повний текст джерелаRomanelli, Mirko. "Modellazione del comportamento di un combustore e turbina aeronautica con fogging." Bachelor's thesis, Alma Mater Studiorum - Università di Bologna, 2016. http://amslaurea.unibo.it/12374/.
Повний текст джерелаMazur, Steven (Steven Andrew). "Turbine tip clearance loss mechanisms." Thesis, Massachusetts Institute of Technology, 2013. http://hdl.handle.net/1721.1/82486.
Повний текст джерелаThis electronic version was submitted and approved by the author's academic department as part of an electronic thesis pilot project. The certified thesis is available in the Institute Archives and Special Collections.
Cataloged from department-submitted PDF version of thesis
Includes bibliographical references (p. 97-98).
Three-dimensional numerical simulations (RANS and URANS) were used to assess the impact of two specific design features, and of aspects of the actual turbine environment, on turbine blade tip loss. The calculations were carried out for a subsonic high pressure turbine stage. The loss mechanism examined is that due to tip clearance vortex mixing. The effects examined were three-dimensional blade stacking, downstream transition duct geometry, and unsteadiness due to an upstream nozzle guide vane. Tip leakage loss changes due to three-dimensional blade stacking (bowing or reverse bowing) are verified to be associated with changes in the magnitude of blade tip loading, which create differences in the leakage flow exit velocities. The effect of a downstream diffusing transition duct on tip leakage losses is small; there was a 3.6% increase in tip leakage loss for a 65% increase in duct exit-to-inlet area ratio compared to a constant area duct. For unsteadiness arising from an upstream nozzle guide vane, it is shown that substantial temporal fluctuations in vortex core velocity and loss generation exist. However, the time average tip leakage loss differed less than 5% from the tip leakage loss calculated on a steady flow basis. Based on the computations, the mechanism for tip leakage vortex loss in the three different situations examined appears to be similar to that which is seen for an isolated turbine blade.
by Steven Mazur.
S.M.
Steyn, J. Lodewyk (Jasper Lodewyk) 1976. "A microfabricated ElectroQuasiStatic induction turbine-generator." Thesis, Massachusetts Institute of Technology, 2005. http://hdl.handle.net/1721.1/32463.
Повний текст джерелаIncludes bibliographical references (p. [263]-268).
An ElectroQuasiStatic (EQS) induction machine has been fabricated and has generated net electric power. A maximum power output of 192 [mu]W at 235 krpm has been measured under driven excitation of the six phases. Self excited operation was also demonstrated. Under self-excitation, no external drive electronics are required and sufficient power was produced to dimly light four LED's on two of the six phases. This is believed to be the first demonstration of both power generation and self-excited operation of an EQS induction machine of any scale reported in the open literature. The generator comprises 5 silicon layers, fusion bonded together, and annealed at 700⁰C. The turbine rotor, 4 mm in diameter, is supported on gas bearings. The thrust bearings are formed by a shallow etch of 1.5 [mu]m to define the thrust bearing gap. Thrust bearing pressurization is through 10 [mu]m diameter nozzles, etched 100 [mu]m deep. The journal bearing is a precision, ... wide, 300 [mu]m deep annular trench around the periphery of the turbine disk. The generator airgap is 3 [mu]m. The inner radius of the generator is 1.011 mm, and the outer radius 1.87mm. The machine has ].31 poles for each of the 6 phases, for a total of 786 stator electrodes. Precise microfabrication and aligned, full-wafer fusion bonding enabled turbine generator devices to be operated at rotational speeds as high as 850 krpm. A detailed state-space model of the EQS machine and its external parasitics is presented. The external stray capacitances, and their unbalance, play a critical role in the performance of the device. A method for estimating the strays experimentally is discussed.
(cont.) This estimated, updated model made it possible to use computer optimization techniques to find the optimal drive conditions for the device to generate maximum power. Carrier depletion in the moderately doped polysilicon rotor conductor film prevented the generator from producing power at higher voltages, and limited the maximum machine terminal voltage under self-excitation to approximately 30 Vp-p.
by Jasper Lodewyk Steyn.
Ph.D.
D'Hoop, Emmanuel Michel. "Flowfield measurements in a transonic turbine." Thesis, Massachusetts Institute of Technology, 1995. http://hdl.handle.net/1721.1/46442.
Повний текст джерелаKountras, Apostolos 1970. "Probabilistic analysis of turbine blade durability." Thesis, Massachusetts Institute of Technology, 2004. http://hdl.handle.net/1721.1/28893.
Повний текст джерелаIncludes bibliographical references (leaves 71-72).
The effect of variability on turbine blade durability was assessed for seven design/operating parameters in three blade designs. The parameters included gas path and cooling convective parameters, metal and coating thermal conductivity and coating thickness. The durability life was modelled as limited by thermo-mechanical low cycle fatigue and creep. A nominal blade design as well as two additional variants were examined using deterministic and probabilistic approaches. External thermal and pressure boundary conditions were generated by three-dimensional CFD calculations. The location of expected failure was the bottom of the trailing edge cooling slot and was the same for all three designs examined. The nominal design had higher life and less variability for the ranges of design parameters examined. For the temperature range studied fatigue was the primary damage mechanism. The variation in cooling air bulk temperature was most important in setting the variation in blade durability life. This life variation was also affected by main gas bulk temperature and heat transfer coefficient, and cooling heat transfer coefficient, but to a lesser extent.
by Apostolos Kountras.
S.M.
Barry, Pamela S. (Pamela Sue). "Rotational effects on turbine blade cooling." Thesis, Massachusetts Institute of Technology, 1994. http://hdl.handle.net/1721.1/12114.
Повний текст джерелаTitle as it appears in the June 1994 MIT Graduate List: Rotational effects of turbine cooling.
Includes bibliographical references (leaves 102-103).
by Pamela S. Barry.
M.S.
Malbois, Pierre. "Analyse expérimentale par diagnostics lasers du mélange kérosène/air et de la combustion swirlée pauvre prémélangée, haute-pression issue d’un injecteur Low-NOx." Thesis, Normandie, 2017. http://www.theses.fr/2017NORMIR25/document.
Повний текст джерелаAeronautical engine manufacturers are banking on the development of innovative fuel injection systems to reduce fuel consumption and pollutant emissions. The aim of the thesis is to contribute to the experimental investigation of a "Lean Premixed" injector by developing laser diagnostics coupling approaches based on Mie scattering and fluorescent emission of tracers. Measurements are performed at high pressure on the HERON combustion test bench. An innovative approach with fluorescence imaging of kerosene has resulted in the quantification of the kerosene/air mixture. The flame structure was analyzed simultaneously by OH-PLIF and velocity PIV measurements were performed to complete this analysis. A preliminary development of CO-PLIF was also conducted. The numerous measurements provided a detailed analysis of the mechanisms of flame/spray/aerodynamic interactions during a swirl-stabilized kerosene/air combustion at high pressure
Jedamski, Devon (Devon James). "Turbine inlet non-uniformities and unsteady mechanisms." Thesis, Massachusetts Institute of Technology, 2015. http://hdl.handle.net/1721.1/98685.
Повний текст джерелаCataloged from PDF version of thesis.
Includes bibliographical references (pages 87-88).
The effect of axial turbine stage inlet non-uniformities are examined through two model problems: wake attenuation and hot streak processing. In the first, twodimensional calculations (RANS and URANS) are used to identify the mechanisms contributing to upstream stator wake attenuation through a turbine blade row. For a representative turbine rotor, pitch and time-averaged wake attenuation by 74% percent is demonstrated at one quarter chord downstream of the trailing edge. Near the pressure surface, the wake stagnation pressure increases by up to 42% above the freestream stagnation pressure. The mechanisms identified are a localized reduction in flow-through time for wake fluid near the pressure surface, compared to the freestream, and an unsteady pressure field (Op/&t) in the rotor reference frame that increases work extraction in the freestream relative to wake fluid. For the second model problem, three-dimensional calculations (RANS and URANS) identify a difference in turbine efficiency sensitivity to thermal distortion between a geometry with no tip gap and a geometry with a finite tip gap. The turbine with a tip clearance is 2.5 times less sensitive, in terms of efficiency decrease, to an inlet hot streak. For the tip gap and no tip gap geometries, the efficiency drops by 0.75% and 1.86% respectively for a peak temperature non-uniformity equal to 0.6 times the combustor temperature rise. The difference in efficiency decrease, due to hot streak, between the two geometries is linked to a reduction in tip leakage mixing losses caused by changes in relative rotor inlet flow angle with and without hot streak.
by Devon Jedamski.
S.M.
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.
Повний текст джерелаIncludes 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.
Underwood, David Scott. "Primary zone modeling for gas turbine combustors." Thesis, Massachusetts Institute of Technology, 1999. http://hdl.handle.net/1721.1/32700.
Повний текст джерела"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.
Guenette, Gerald Roger. "A fully scaled short duration turbine experiment." Thesis, Massachusetts Institute of Technology, 1985. http://hdl.handle.net/1721.1/15249.
Повний текст джерелаMICROFICHE COPY AVAILABLE IN ARCHIVES AND AERO.
Includes bibliographical references.
by Gerald Roger Guenette, Jr.
Sc.D.
Huang, Arthur (Arthur Chan-wei). "Loss mechanisms in turbine tip clearance flows." Thesis, Massachusetts Institute of Technology, 2011. http://hdl.handle.net/1721.1/67067.
Повний текст джерелаThis electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections.
Cataloged from student submitted PDF version of thesis.
Includes bibliographical references (p. 109-110).
Numerical simulations of tip clearance ow have been carried out to dene the loss generation mechanisms associated with tip leakage in unshrouded axial turbines. Mix- ing loss between the leakage, which takes the form of a strong embedded streamwise vortex (u=ux 1 in the vortex core), and the mainstream ow is found to be the main source of loss. Vortex line contraction, and consequent vortex core expansion, and also vortex breakdown, are identied as the two important mechanisms that determine mixing loss. Because of these vortex dynamic features, the behavior is dierent from the conventional view of the effect of pressure level on mixing of non- uniform flows. More specifically, it is shown, through control volume arguments and axisymmetric computations, that as a strongly swirling ow passes through a pres- sure rise, the mixed-out loss can either decrease or increase, the latter occurring if the deceleration becomes large enough to initiate vortex breakdown. It is further shown that tip vortices in turbines experience pressure rises large enough to cause vortex breakdown. The effect of pressure distribution on tip leakage losses is illustrated through examination of two turbine blades, one designed with a forward loaded tip and one with an aft loaded tip. The computations show a 16% difference in tip clearance loss between the two, due to the lower pressure rise encountered by the clearance vortex, and hence lower vortex breakdown losses, with the forward loaded blade. Other computational experiments, on the effects of blade loading, incidence, and solidity, are also shown to be consistent with the ideas developed about blade pressure distribution effects on vortex breakdown and hence clearance mixing loss.
by Arthur Huang.
S.M.
Livera, Filippo. "Dimensionamento ed ottimizzazione di un compressore assiale per turbina aeronautica ad altissima potenza." Bachelor's thesis, Alma Mater Studiorum - Università di Bologna, 2014. http://amslaurea.unibo.it/7564/.
Повний текст джерелаFaggi, Elia. "Validazione di un modello real-time di un turboalbero aeronautico." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2015. http://amslaurea.unibo.it/9207/.
Повний текст джерелаCai, Yi 1970. "Aerodynamic performance measurements in a fully scaled turbine." Thesis, Massachusetts Institute of Technology, 1998. http://hdl.handle.net/1721.1/47406.
Повний текст джерелаEvans, Simon William 1977. "Thermal design of a cooled micro gas turbine." Thesis, Massachusetts Institute of Technology, 2001. http://hdl.handle.net/1721.1/8093.
Повний текст джерелаIncludes 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.
Philippon, Baudoin (Baudoin Henry) 1975. "Design of a film cooled MEMS micro turbine." Thesis, Massachusetts Institute of Technology, 2001. http://hdl.handle.net/1721.1/35488.
Повний текст джерелаIncludes bibliographical references (leaf 120).
As part of an effort to develop a portable power generation system, a fluid dynamics and thermal transfer investigation of a micro radial inflow turbine was carried out. The 3-D numerical performance assessment revealed that the baseline 2-D designed turbine stage was not matched to the baseline compressor, resulting in off design operation. The CFD predicts that the baseline turbine has a total to static efficiency of 29%, and does not provide enough power to drive the compressor at the matched pressure ratio of 1.65. Reasons for this low efficiency are the blockage due to end walls effects and to the exit right angle turn, and 3-D secondary flows in the blade passage leading to boundary layer separation. The turbine was then redesigned. An analytical design procedure, based on a mean line analysis and correlations from 3-D CFD solutions was formulated and validated against numerical results. It was shown that significant performance gains could be achieved by increasing the turbine exit area to reduce the exit viscous loss and by increasing the blade exit angle. Shaping of the exit diffuser turned out not to be viable because of the difficulties in keeping the boundary layer attached. An improved turbine was then designed. Numerical simulations of the improved design predicted a 20% gain in efficiency, at a matched pressure ratio of 2.1. Still, the turbine cannot drive the compressor. The turbine is conduction cooled by the compressor, but the large heat addition to the compressor flow causes a 30% drop in efficiency. Film cooling schemes for the turbine were investigated. An axisymmetric model showed that a coolant layer flowing radially inward may sustain the adverse centrifugal force at design speed. Film cooling schemes were then proposed for disk and blade cooling. Effectiveness drivers are surface coverage, thermal mixing with the main flow, and coolant matching. The overall peak cooling effectiveness of the proposed cooling schemes was approximately 30% for a 30% coolant flow.
by Baudoin Philippon.
S.M.
Kim, Yusik. "Wind turbine aerodynamics in freestream turbulence." Thesis, University of Southampton, 2013. https://eprints.soton.ac.uk/360372/.
Повний текст джерелаDuffner, John D. "The effects of manufacturing variability on turbine vane performance." Thesis, Massachusetts Institute of Technology, 2008. http://hdl.handle.net/1721.1/44933.
Повний текст джерелаThis electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections.
Includes bibliographical references (leaves 71-73).
Gas turbine vanes have airfoil shapes optimized to deliver specific flow conditions to turbine rotors. The limitations of the manufacturing process with regards to accuracy and precision mean that no vane will exactly match the design intent. This research effort is an investigation of the effects of manufacturing-induced geometry variability on the performance of a transonic turbine vane. Variability is characterized by performing Principal Components Analysis (PCA) on a set of measured vanes and then applied to a different vane design. The performance scatter of that design is estimated through Monte Carlo analysis. The effect of a single PCA mode on performance is estimated and it is found that some modes with lower geometric variability can have greater impact on performance metrics. Linear sensitivity analysis, both viscous and inviscid, is carried out to survey performance sensitivity to localized surface perturbations, and tolerances are evaluated using these results. The flow field is seen to be practically insensitive to shape changes upstream of the throat. Especially sensitive locations like the throat and trailing edge are investigated further through nonlinear sensitivity analysis.
by John D. Duffner.
S.M.
Groshenry, Christophe. "Preliminary design study of a micro-gas turbine engine." Thesis, Massachusetts Institute of Technology, 1995. http://hdl.handle.net/1721.1/10386.
Повний текст джерелаLiu, Chunmeni 1970. "Dynamical system modeling of a micro gas turbine engine." Thesis, Massachusetts Institute of Technology, 2000. http://hdl.handle.net/1721.1/9249.
Повний текст джерелаAlso 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.
Steptoe, William James. "Integral boundary layer heat transfer prediction on turbine blades." Thesis, Massachusetts Institute of Technology, 1989. http://hdl.handle.net/1721.1/42188.
Повний текст джерелаCrawford, Curran A. (Curran Alexander) 1978. "An integrated CAD methodology applied to wind turbine optimization." Thesis, Massachusetts Institute of Technology, 2003. http://hdl.handle.net/1721.1/17047.
Повний текст джерелаIncludes bibliographical references (p. 169-172).
This electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections.
Modern engineering practice for designing physical products requires the creation of a CAD model of the design for documentation and manufacturing. As the design evolves from concept through to production, it is analyzed a number of times, in some cases using general parameters and in others requiring fine details of the product's form. The setup of each analysis is typically disjoint from the previous steps, inhibiting design changes and optimization. This thesis addresses these bottlenecks by proposing a methodology to use the CAD model of the system as the central element. The model is created at the earliest possible stage of the process following a strict synthesis procedure; it then forms a common base for all of the follow-on analyses and development. Computational tools are developed to aid in using the geometric and parametric information in the CAD model to setup simulations, as well as to dynamically drive the CAD model itself for design studies and optimization. The methods and tools are then applied to the design of a wind turbine for power production. Using a common CAD model, various analysis codes and optimization algorithms are applied to the design of the system, to lower the cost of delivered energy. Multidisciplinary aspects of wind turbine design including aerodynamics, structures, and economics are presented together with the employed modeling techniques. The demonstrated improvements achieved over the baseline design lend credence to the methodology, and demonstrate its effectiveness in enhancing the systems performance. The insights gleaned from the present work intimate promising directions for continued development both at the level of software tools and also effective methods for approaching multi-disciplinary design.
by Curran A. Crawford.
S.M.
Keogh, Rory (Rory Colm) 1968. "Aerodynamic performance measurements of a film-cooled turbine stage." Thesis, Massachusetts Institute of Technology, 2001. http://hdl.handle.net/1721.1/8869.
Повний текст джерелаIncludes bibliographical references (p. 167-168).
Thesis (Ph.D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2001.
The goal of this research is to measure the aerodynamic performance of a film-cooled turbine stage and to quantify the loss caused by film-cooling. A secondary goal of the research is to provide a detailed breakdown of the losses associated with film-cooling for the turbine stage being tested. The experimental work was carried out at the MIT Blowdown Turbine Facility using a highly loaded turbine stage. The Blowdown Turbine Facility is a short duration test facility capable of testing turbine stages under fully scaled conditions for a test duration of 0.5 seconds. The facility was modified to enable the measurement of the turbine mass flow and shaft torque. These newly developed measurement techniques, along with previously developed total pressure and temperature instruments, have enabled the measure- ment of the stage isentropic efficiency. A highly loaded turbine stage (without film-cooling) was designed, fabricated, and tested using the newly developed measurement techniques. The turbine stage was then modified to incorporate vane, blade and rotor casing coolant manifolds using precision electrical discharge machining. The film-cooling hole geometry was created using a laser drilling process to produce the required 43,000 cooling holes. The film-cooled stage was then tested over a range of operating conditions (pressure ratios and corrected speeds) and over a range of coolant-to-mainstream mass flow and temperature ratios.
(cont.) The loss due to film-cooling is defined as the difference in performance between the film-cooled turbine and an ideal turbine with the same velocity triangles and airfoil Mach number distributions. However, there is no uncooled turbine geometry that will produce the same flow conditions as the film-cooled turbine stage, and consequently, there is no experimental baseline that can be tested to determine the loss due to film- cooling. A meanline velocity triangle model of the turbine stage was developed using published correlations and loss models to estimate the performance of this ideal stage. The model was calibrated against the baseline test results without coolant and it was then used to estimate the loss due to film-cooling. The estimated loss due to film-cooling was 3.0% at the design point, which corresponds to 0.3% per percent of coolant. The estimated repeatability (U95) for the efficiency measurement of the uncooled tur- bine geometry is ± 0.14%. Based on this measurement repeatability, the net effect of a design change can be determined with an uncertainty of just ± 0.1% if four measurements are repeated for each design configuration. The estimated measurement uncertainty for the film-cooled stage efficiency is 0.55% and for back-to-back measurements the uncertainty is 0.45%.
by Rory Keogh.
Ph.D.
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.
Повний текст джерелаCataloged 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.
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.
Повний текст джерелаIncludes 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.
Shang, Tonghuo. "Influence of inlet temperature distortion on turbine heat transfer." Thesis, Massachusetts Institute of Technology, 1995. http://hdl.handle.net/1721.1/47370.
Повний текст джерелаPrashanth, Prakash. "Post-combustion emissions control for aero-gas turbine engines." Thesis, Massachusetts Institute of Technology, 2018. https://hdl.handle.net/1721.1/122402.
Повний текст джерелаCataloged 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
Catalfamo, Peter T. "Characterization of turbine rim seal flow and its sealing effectiveness." Thesis, Massachusetts Institute of Technology, 2013. http://hdl.handle.net/1721.1/82504.
Повний текст джерелаThis thesis was scanned as part of an electronic thesis pilot project.
Cataloged from PDF version of thesis
Includes bibliographical references (p. 83-84).
In a gas turbine engine, ingestion of hot gas from the flowpath into the gaps between the turbine rotor and stator can lead to elevated metal temperatures and a deterioration of component life. To prevent ingestion, bleed air from the compressor is used to "purge" the rim seal cavities. Establishing a quantitative understanding of the wheelspace and rim cavity flow processes driving ingestion is critical to optimizing seal design and minimizing the associated performance penalty. A computational model of the wheelspace that does not limit the spatial or temporal scales of flow processes is formulated. This allows the assessment of the response of the wheelspace to external stimuli set up by the turbine main flow path, and the development of causal links between flow processes and their drivers. Varying the axisymmetric turbine flowpath pressure on a quasi-steady basis when the purge flow supply seal is choked has no impact on ingestion; the pressure field in the wheelspace merely scales with the flowpath pressure, leaving the flow structure unchanged. Introducing circumferential variation in the external pressure field can, however, lead to ingestion with the ratio of disturbance wavelength to the trench depth emerging as a key parameter. Varying rotational speed alone does not drive ingestion as a stagnation point is formed on the outer shroud that is ingestion resistant. It is shown that excitation at frequencies corresponding to the natural modes of the wheelspace system can lead to large responses in pressure and seal flow rate, with the seal reduced frequency appearing as a characterizing parameter. The existence and parametric dependence of these modes is further assessed through a small disturbance flow analysis. A generalized small disturbance flow analysis is formulated that provides a direct enumeration of the key characterizing parameters.
by Peter T. Catalfamo.
S.M.
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.
Повний текст джерелаIncludes 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.
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.
Повний текст джерелаShannon, Kevin R. (Kevin Robert). "Loss mechanisms in a highly loaded transonic axial turbine stage." Thesis, Massachusetts Institute of Technology, 2018. http://hdl.handle.net/1721.1/120440.
Повний текст джерелаCataloged from PDF version of thesis.
Includes bibliographical references (pages 129-130).
Flow in a one-and-a-half stage highly loaded transonic axial turbine representative of future generation turbine technology is assessed for its role in loss generation. Steady and unsteady two-dimensional and three-dimensional flow computations, complemented by simplistic control volume analyses as well as test data, provided results for establishing the quantitative level of loss from various sources. The test data has been acquired in a cascade and blowdown turbine research rig. Specifically, the overall loss determined from unsteady three-dimensional flow computations of a cooled one-and-a-half stage turbine is within 6% of that inferred from the blowdown turbine rig test data. The computed flows with different levels of flow and configuration complexities are post-processed and interrogated to allow an estimation of blade profile loss, trailing edge loss, shock loss, endwall loss, secondary flow loss, tip leakage loss, cooling injection loss, and unsteady flow loss. The dominant sources of loss are determined to be the trailing edge loss, profile loss, and tip leakage loss. The computed flows show that the flow deviation in a highly loaded transonic turbine airfoil with trailing edge shocks is negative (-2° to -4°); estimating the trailing edge loss by assuming zero flow deviation in a simple control volume approach would yield a significantly higher value. Loss arising from flow unsteadiness contributes an additional loss of about 1/6 of that in steady flow approximation; 3/4 of the flow unsteadiness induced loss occurs in the downstream vane where the flow is threaded with propagating shocks from the upstream blade and downstream shock reflections; and the remaining 1/4 is from unsteadiness in NGV wakes and shock oscillations from influence of the adjacent airfoil row. 1/5 of the overall loss in the one-and-a-half stage turbine is from the cooling and purge flows. A preliminary assessment of loss variation with turbine stage pressure ratio shows a non-monotonic trend.
by Kevin R. Shannon.
S.M.
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.
Повний текст джерелаLee, Jinwook S. M. Massachusetts Institute of Technology. "Aerothermodynamics and operation of turbine system under unsteady pulsating flow." Thesis, Massachusetts Institute of Technology, 2015. http://hdl.handle.net/1721.1/98692.
Повний текст джерелаCataloged from PDF version of thesis.
Includes bibliographical references (pages 133-135).
An assessment of a turbine system operating under highly pulsating flow environment typically found in vehicular turbochargers is made to: identify the key operating parameters, enable the formulation of a reduced order model, delineate the sources of loss and suggest strategies for performance improvement. The turbine system consists of a scroll-volute followed by a turbine wheel and then a diffuser. The assessment includes calculating unsteady three-dimensional flow in the turbine system followed by in-depth interrogation complemented with flow modeling. The key findings are (1) The flow mechanisms behind the turbine wheel performance, the diffuser loss and the wastegate port loss appear locally quasi-steady such that we can characterize the performance of the components based on a series of steady calculations subjected to varying inlet conditions reflecting the inlet flow pulsation; (2) the operation of scroll-volute and the diffuser pressure recovery can be adequately determined using a quasi-one-dimensional unsteady flow model; (3) A significant fraction of the loss that is not from skin frictions occurs downstream of turbine wheel exit (18%pts out of 34%pts in Peak Torque and 20%pts out of 56%pts in Turbo Initial Transient based on cycle loss debit); (4) The condition of maximum power extraction on unsteady pulsating environment can be approximated with a simple modeling of volute storage effect. A physically consistent definition of ideal power that elucidates the role of unsteadiness in an unsteady turbine system is derived; it informs one on what the extractable power is compared to what it could be for an ideal system. Finally the findings are used to define the required attributes of methodology for estimating efficiency with a specified uncertainty bandwidth.
by Jinwook Lee.
S.M.
Zhang, K. "Turbulent combustion simulation in realistic gas-turbine combustors." Thesis, City, University of London, 2017. http://openaccess.city.ac.uk/17689/.
Повний текст джерела