Dissertations / Theses on the topic 'Aeronautical turbine'
To see the other types of publications on this topic, follow the link: Aeronautical turbine.
Create a spot-on reference in APA, MLA, Chicago, Harvard, and other styles
Consult the top 50 dissertations / theses for your research on the topic 'Aeronautical turbine.'
Next to every source in the list of references, there is an 'Add to bibliography' button. Press on it, and we will generate automatically the bibliographic reference to the chosen work in the citation style you need: APA, MLA, Harvard, Chicago, Vancouver, etc.
You can also download the full text of the academic publication as pdf and read online its abstract whenever available in the metadata.
Browse dissertations / theses on a wide variety of disciplines and organise your bibliography correctly.
1
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.
Full textAbstract:
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.
2
Al-Khudairi, Othman. "Structural performance of horizontal axis wind turbine blade." Thesis, Kingston University, 2014. http://eprints.kingston.ac.uk/32197/.
Full textAbstract:
The power output from a wind turbine is proportional to rotor swept area and as a result in the past 30 years continuous effort has been made to design larger blades. In this period, the blade length has been increased about 10 times since 1980s to present time. With the longest blade currently measuring more than 100m in length, wind turbine blade designers and manufacturers face enormous challenges to encounter the effect of increased weight and other loads on fatigue durability of the blade. Wind turbine blades are mainly made from glass fibre reinforced plastic (GFRP) composite. materials. As a result, in the design of various parts of wind turbine blades such as the shear web, spar cap and the aerofoil the fatigue behaviour of F RP materials is required. The performance of these parts as well as the adhesively bonded joint under fatigue loading is crucial for structural integrity of a long lasting blade. During operation, delamination can initiate and propagate shortening blade life; hence, characterisation of failure envelope of GFRP laminates under different loading mode is necessary. In this regard in this project, quasi-static tests were carried out to find mode 1, mode 11 and mixed mode I/11 delamination fracture toughness using DCB, ENF and MMB tests and the fracture envelope was established for various mode mixity. In the next stage, the stress-lifetime (S-N) diagrams of the GFRP was studied. Fatigue-life experiments on three different types of loading, i.e. tension-tension at R=0.1, 0.5, tension- compression at R=-1 and compression-compression at R=2 and R=10 were performed. From the results of S-N diagrams, the constant life diagrams (CLD) for 90 degree and 0 degree fibre directions were constructed. CLD diagrams are useful for prediction of fatigue lifetime for loading condition that no experimental data available. The analysis of delamination crack propagation under cyclic loading was next area of the research. The onset life and propagation delamination crack grth of 0//0 interface of GR P laminate in mode I loading using DCB specimens was investigated and the Gm. from the onset life test was determined. From the fitted curve to mode I experimental propagation data the Paris’ law coefficient for the laminated GFRP in mode I was determined. The mode II fatigue crack growth in laminated 0//0 GFRP material was also investigated using ENF specimens. The fatigue behaviour in this mode is analysed based on application of Paris’ law as a function of energy release rate for mode II loading. From the fitted curve to experimental data, the Paris’ law coefficient for the laminated GFRP in mode II was determined. The effect of fatigue delamination growth on fracture surface was studied by fractography analysis of SEM images of fracture surfaces. Studying the behaviour of GFRP under cyclic loading and delamination under static and dynamic load led to full-scale testing of wind turbine blade to establish damage tolerance of the blade under cyclic loading. The sensitivity of wind turbine blade to damage has considerable interest for turbine operators and manufacturers. For full-scale fatigue testing, calibration test and modal analysis of a 45.7m blade has been done and moment-strain diagram and natural frequencies of the blade were obtained. Next, the blade sensitivity to damage under fatigue loading was investigated. The blade has been damaged intentionally by initially inserting a crack of 0.2m between the shear web and spar cap and later it was extended to 1m. The effect of these damages on the modal shape, natural frequencies and strains at various locations of the blade were investigated. The damaged blade fatigue tested, the structural integrity and growth of damage were monitored, and the results were discussed. Finally for the improvement of delamination resistance for joints between spar beam and aero-shell stitching method was used. T-beam and box beam joint were chosen as the platform for testing the stitching effect on the delamination. Various pattern of stitching was applied and the optimum pattern was determined.
3
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.
Full textAbstract:
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
4
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.
Full textAbstract:
The main purpose of this study is to aerodynamically design and optimize winglet, twist angle distribution and pitch angle for a wind turbine blade using CFD to produce more power. The RANS solver of Numeca Fine/Turbo was validated by two test cases, the NREL II and NREL VI blades. The results have shown a
considerable agreement with measurements for both cases. Two different preconditioners have been implemented for the low Mach number flow. The results have shown the superiority of Merkle preconditioner over Hakimi one and Merkle
was selected for further simulations. In addition to that, different turbulence models have been compared and the Launder &ndashSharma 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%.
5
Effendy, Marwan. "Investigation of turbine blade trailing edge cooling and thermal mixing characteristics." Thesis, Kingston University, 2014. http://eprints.kingston.ac.uk/30604/.
Full textAbstract:
The present computation investigates a turbine blade with trailing-edge cutback coolant ejection designs, aiming for a comparison study of aerothermal performances such as discharge coefficient and film cooling effectiveness due to the change of trailing-edge geometries and blowing ratios. The shear-stress transport (SST) k-w turbulence model is adopted and numerical studies are carried out by two-stage investigations:- firstly, validation of an existing cutback blade model with staggered circular pin-fins array inside the cooling passage that has been extensively studied by other researchers and predicted internal passage discharge coefficient and film-cooling effectiveness along the cutback surface are compared to experimental measurements. RANS/URANS and DES are applied during this stage; secondly, further investigation of four main cases considering different key design parameters such as the ratio of lip thickness to slot height (t/H = 0.25, 0.5, 1.0 and 1.5), the design of internal features (i.e. circular pin-fin array, elliptic pin-fin array, and empty duct), the coolant ejection angle (alpha = 5 degrees, 10 degrees and 15 degrees). In addition, a trailing-edge cutback model with suction-side (SS) ─ pressure-side (PS) walls and lands is considered to create a more realistic blade design. The results show that both steady and unsteady RANS predictions are able to produce discharge coefficients in fairly good agreement with test data, but not the film-cooling effectiveness on cutback surfaces which over-predicts in far-field wake region. Further prediction improvements can be made by using unsteady DES approach. In terms of film-cooling effectiveness and shedding frequency, computational results indicate a strong dependency on those aforementioned key design parameters. This film-cooling effectiveness is strongly affected by turbulent flow structures along the cutback region, which is representing the dynamic mixing process between the mainstream flow and the ejecting coolant from the slot-exit. The use of elliptic pin-fin inside the cooling passage and thin lip thickness could improve the effectiveness of film-cooling. The increase of ejection angle yields almost near unity cooling effectiveness along the protected wall. Significant improvements on cooling performance are also achieved with higher blowing ratios. Computations of the trailing-edge cutback cooling with pressure-side (PS) and suction-side (SS) wall demonstrates that performance of the case without lands is better than that of the case with lands by discrepancy up to 18% in terms of overall-averaged film-cooling effectiveness. The blade trailing-edge design with lands causes a rapid decay of the averaged film-cooling effectiveness.
6
Oksuz, 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.
Full textAbstract:
To decrease the computational cost of genetic algorithm optimizations,
surrogate models are used during optimization. Online update of surrogate
models and repeated exchange of surrogate models with exact model during
genetic optimization converts static optimization problems to dynamic ones.
However, genetic algorithms fail to converge to the global optimum in
dynamic optimization problems. To address these problems, a multiploid
genetic algorithm optimization method is proposed. Multi-fidelity surrogate
models are assigned to corresponding levels of fitness values to sustain the
static optimization problem. Low fidelity fitness values are used to decrease
the computational cost. The exact/highest-fidelity model fitness value is used for converging to the global optimum. The algorithm is applied to
single and multi-objective turbine blade aerodynamic optimization
problems. The design objectives are selected as maximizing the adiabatic
efficiency and torque so as to reduce the weight, size and the cost of the gas
turbine engine. A 3-D steady Reynolds-Averaged Navier-Stokes solver is
coupled with an automated unstructured grid generation tool. The solver is
validated by using two well known test cases. Blade geometry is modelled
by 37 design variables. Fine and coarse grid solutions are respected as high
and low fidelity surrogate models, respectively. One of the test cases is
selected as the baseline and is modified in the design process. The effects of
input parameters on the performance of the multiploid genetic algorithm are
studied. It is demonstrated that the proposed algorithm accelerates the
optimization cycle while providing convergence to the global optimum for
single and multi-objective problems.
7
Akagi, Raymond. "Ram Air-Turbine of Minimum Drag." DigitalCommons@CalPoly, 2021. https://digitalcommons.calpoly.edu/theses/2261.
Full textAbstract:
The primary motivation for this work was to predict the conditions that would yield minimum drag for a small Ram-Air Turbine used to provide a specified power requirement for a small flight test instrument called the Boundary Layer Data System. Actuator Disk Theory was used to provide an analytical model for this work.
Classic Actuator Disk Theory (CADT) or Froude’s Momentum Theory was initially established for quasi-one-dimensional flows and inviscid fluids to predict the power output, drag, and efficiency of energy-extracting devices as a function of wake and freestream velocities using the laws of Conservations of Mass, Momentum, and Energy. Because swirl and losses due to the effects of viscosity have real and significant impacts on existing turbines, there is a strong motivation to develop models which can provide generalized results about the performance of an energy-extractor, such as a turbine, with the inclusion of these effects. A model with swirl and a model with losses due to the effects of viscosity were incorporated into CADT which yielded equations that predicted the performance of an energy-extractor for both un-ducted and ducted cases. In both of these models, for this application, additional performance parameters were analyzed including the drag, drag coefficient, power output, power coefficient, force coefficient, and relative efficiency.
For the un-ducted CADT, it is well known that the wake-to-freestream velocity ratio of 1/3 will give the maximum power extraction efficiency of 59.3%; this result is called the Betz limit. However, the present analysis shows that reduced drag for a desired power extraction will occur for wake-to-freestream velocity ratios higher than the value of 1/3 which results in maximum power extraction efficiency. This in turn means that a turbine with a larger area than the smallest possible turbine for a specified power extraction will actually experience a lower drag.
The model with the inclusion of swirl made use of the Moment of Momentum Theorem applied to a single-rotor actuator disk with no stators, in addition to the laws of Conservation of Mass, Momentum, and Energy from the CADT. The results from the model w/swirl showed that drag remains unchanged while power extracted decreases with the addition of swirl, with swirl effects becoming more severe for tip speed ratios below about 5. As for CADT, reduced drag for a specified power extraction can be achieved when the wake-to-freestream velocity ratio is higher that than which provides maximum power extraction efficiency. The model w/losses due to viscosity incorporated the losses into the Conservation of Energy relationship. The results from the model w/losses showed that there is a distinct wake-to-freestream velocity ratio at which minimum drag for a specified power output is achieved, and that this velocity ratio is usually—but not always—higher than that for which the power extraction efficiency is a maximum.
It was concluded that a lower drag for a specified power output of an energy-extractor can usually be achieved at a wake-to-freestream velocity ratio higher than that which produces the v maximum power extraction efficiency. The latter condition, known as the Betz limit for CADT, and which defines the minimum size for a turbine to provide a specified power extraction, is therefore not the correct target design condition to achieve lowest drag for a small Ram-Air Turbine to power BLDS.
8
Notarianni, Gianmarco. "Analysis and modelling of the turbocharger behavior of an internal combustion engine for aeronautical application." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2019.
Find full textAbstract:
The goal of this thesis project is the creation and the validation of an estimation model developed for the prediction of the Wastegate valve behaviour, to define the splitting of mass flow rate between the turbine and the bypass valve in all the operative conditions of the turbocharger system, by an analysis of the experimental data obtained from tests performed in the engine test bench of the University.
9
Gorgulu, 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.
Full textAbstract:
The present study considers the numerical simulation of the different flow characteristics involved in the conjugate heat transfer analysis of an internally cooled gas turbine blade. Conjugate simulations require full coupling of convective heat transfer in fluid regions to the heat diffusion in solid regions. Therefore, accurate prediction of heat transfer quantities on both external and internal surfaces has the uppermost importance and highly connected with the performance of the employed turbulence models. The complex flow on both surfaces of the internally cooled turbine blades is caused from the boundary layer laminar-to-turbulence transition, shock wave interaction with boundary layer, high streamline curvature and sequential flow separation. In order to discover the performances of different turbulence models on these flow types, analyses have been conducted on five different experimental studies each concerned with different flow and heat transfer characteristics. Each experimental study has been examined with four different turbulence models available in the commercial software (ANSYS FLUENT13.0) to decide most suitable RANS-based turbulence model. The Realizable k-&epsilonmodel, 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.
10
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/.
Full textAbstract:
The objective of this work was to conduct a preliminary finite element static structural and transient thermal analyses of a first stage turbine blade which was previously assembled on a Pratt & Whitney JT9D-7A turbofan engine. This turbine blade was obtained from a collector of aircraft scrap parts. After an extensive theoretical background on airbreathing jet engines and materials used for such components, the process behind the creation of a 3D model was explained. The laser scanning technique and a piezoeletric digitizer were employed to recreate the blade inside a 3D modelling software. The model was then imported into the finite element analysis software ANSYS; the analyses were performed, and the most interesting results were evaluated. The structural and thermal results were found to be congruous with the literature on similar applications of components with the same material, and appear to be a realistic representation of the blade behaviour inside the first stage turbine environment.
11
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.
Full textAbstract:
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.
12
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.
Full textAbstract:
This study presents the results of an experimental study that investigates the effects of uniform/waveform tip injection along the camberline on the total pressure loss characteristics downstream of a row of Low Pressure Turbine (LPT) blades. The experiments are performed in a low speed cascade facility. This injection technique involves spanwise jets at the tip that are issued from a series of holes along the camber line normal to the freestream flow direction. The injection mass flow rate from each hole is individually controlled using computer driven solenoid valves and therefore the flow injection geometrical pattern at the tip can be adjusted to any desired waveform shape, and can be uniform as well as waveform along the camber. Measurements involve Kiel probe traverses for different injection scenarios 0.5 axial chords downstream of the blades. Results show that, instead of performing uniform mass injection along the camberline, by selecting an appropriate waveform injection pattern one can reduce the total loss levels of the blade, including the tip leakage loss as well as the wake losses.
13
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.
Full textAbstract:
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.
14
Dogan, 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.
Full textAbstract:
This thesis presents the results of an experimental study that investigates the effects of steady vortex generator jets (VGJs) integrated to a low pressure turbine blade to control the laminar separation bubble occurring on the suction surface of the blade at low Reynolds numbers. The injection technique involves jets issued from the holes located near the suction peak of the test blade which is in the middle of a five-blade low speed linear cascade facility. Three injection cases are tested with different blowing ratio values ranging from low to high. Surface pressure and particle image velocimetry (PIV) measurements are performed. The results show that steady VGJ is effective in eliminating the laminar separation bubble. Also it is observed that to have fully developed attached boundary layer, blowing ratio should be chosen accordingly since a very thin separation zone still exists at low blowing ratios.
15
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.
Full text
16
Romanelli, 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/.
Full textAbstract:
Tra i vari metodi di produzione di energia meccanica le turbomacchine rappresentano sicuramente uno dei metodi più sviluppati e di maggiore interesse per via del loro utilizzo in ambito aeronautico, per la produzione di energia elettrica e la movimentazione di grosse macchine
motrici. A causa della loro diffusione è diventato di fondamentale importanza trovare delle metodologie per aumentarne la potenza, l'ecosostenibilità e i rendimenti. A tale scopo la tecnica del fogging rappresenta una valida opzione per cercare di raggiungere tali obiettivi.
In questo elaborato è stato prima effettuato uno studio in ambiente MatLab del comportamento di un turboalbero Allison 250 C18 variando la portata di acqua iniettata all'interno della motorizzazione cercando di determinare analiticamente, mediante le conoscenze teoriche acquisite,
l'andamento di alcune grandezze fondamentali quali temperatura e pressione per poi confrontare i risultati ottenuti con i dati sperimentali raccolti precedentemente presso il "Laboratorio di macchine e propulsione" della Scuola di Ingegneria e Architettura, sede di Forlì. Successivamente si è passati alla creazione tramite Simulink, un tool di MatLab, di un modello dinamico che ci permettesse di estendere i risultati ottenuti dallo studio di alcuni punti di funzionamento a un range più elevato in modo da avere un quadro completo del comportamento del turboalbero.
17
Mazur, Steven (Steven Andrew). "Turbine tip clearance loss mechanisms." Thesis, Massachusetts Institute of Technology, 2013. http://hdl.handle.net/1721.1/82486.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2013.This 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.
18
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.
Full textAbstract:
Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2005.Includes 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.
19
D'Hoop, Emmanuel Michel. "Flowfield measurements in a transonic turbine." Thesis, Massachusetts Institute of Technology, 1995. http://hdl.handle.net/1721.1/46442.
Full text
20
Kountras, Apostolos 1970. "Probabilistic analysis of turbine blade durability." Thesis, Massachusetts Institute of Technology, 2004. http://hdl.handle.net/1721.1/28893.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2004.Includes 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.
21
Barry, Pamela S. (Pamela Sue). "Rotational effects on turbine blade cooling." Thesis, Massachusetts Institute of Technology, 1994. http://hdl.handle.net/1721.1/12114.
Full textAbstract:
Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1994.Title 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.
22
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.
Full textAbstract:
Les motoristes aéronautiques misent sur le développement de systèmes d’injection de carburant innovants pour réduire la consommation de carburant et les émissions de polluants. L’objectif de la thèse est de contribuer à l’étude expérimentale d’un injecteur « Lean Premixed » par le développement de diagnostics lasers couplant des approches basées sur la diffusion de Mie et l’émission fluorescente de traceurs. Les mesures ont été réalisées sur le banc de combustion haute pression HERON. Une approche novatrice avec l’imagerie de fluorescence du kérosène a permis d’obtenir une quantification du mélange kérosène/air. La structure de flamme a été mesurée simultanément par PLIF-OH et des mesures PIV de vitesse ont complété cette analyse. Un développement préliminaire de la PLIF-CO a également été mené. Les nombreuses mesures permettent de fournir une analyse détaillée des interactions flamme/spray/aérodynamique lors d’une combustion swirlée stabilisée kérosène/air à haute pressionAeronautical 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
23
Jedamski, Devon (Devon James). "Turbine inlet non-uniformities and unsteady mechanisms." Thesis, Massachusetts Institute of Technology, 2015. http://hdl.handle.net/1721.1/98685.
Full textAbstract:
Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2015.Cataloged from PDF version of thesis.
Includes bibliographical references (pages 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.
24
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.
Full textAbstract:
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.
25
Underwood, David Scott. "Primary zone modeling for gas turbine combustors." Thesis, Massachusetts Institute of Technology, 1999. http://hdl.handle.net/1721.1/32700.
Full textAbstract:
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.
26
Guenette, Gerald Roger. "A fully scaled short duration turbine experiment." Thesis, Massachusetts Institute of Technology, 1985. http://hdl.handle.net/1721.1/15249.
Full textAbstract:
Thesis (Sc. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1985.MICROFICHE COPY AVAILABLE IN ARCHIVES AND AERO.
Includes bibliographical references.
by Gerald Roger Guenette, Jr.
Sc.D.
27
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.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2011.This electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections.
Cataloged from student submitted PDF version of thesis.
Includes bibliographical references (p. 109-110).
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.
28
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/.
Full text
29
Faggi, 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/.
Full textAbstract:
L'obiettivo di questo lavoro di tesi è realizzare e validare un modello del propulsore Allison 250 c18 realizzato in Simulink. In particolare utilizzando i dati provenienti da prove sperimentali effettuate al banco prova del laboratorio di Propulsione e macchine della Scuola di Ingegneria e Archittetura vengono dapprima create le mappe prestazionali dei singoli componenti. Esse sono poi implementate all'interno del modello in Simulink, realizzato con struttura modulare a blocchi permettendo il modellamento di ogni singolo componente del motore. Successivamente i dati provenienti dalle simulazioni sono confrontati con i dati sperimentali, riscontrando risultati soddisfacenti.
In ultimo si riporta brevemente una descrizione della compilazione del modello da utilizzare in una interfaccia LabVIEW per verificarne il funzionamento del modello in real-time.
30
Cai, Yi 1970. "Aerodynamic performance measurements in a fully scaled turbine." Thesis, Massachusetts Institute of Technology, 1998. http://hdl.handle.net/1721.1/47406.
Full text
31
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.
Full textAbstract:
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.
32
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.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2001.Includes 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.
33
Kim, Yusik. "Wind turbine aerodynamics in freestream turbulence." Thesis, University of Southampton, 2013. https://eprints.soton.ac.uk/360372/.
Full textAbstract:
Topics in wind turbine aerodynamics are reviewed. These include the effect of freestrearn turbulence all the flows over wind turbine blades; dynamic stall phenomenon; and rotational augmentation. The advantages of numerical studies on these topics are highlighted and large-eddy simulation (LES) is selected to overcome the defects for other numerical approaches, e.g. Reynolds Average Navier-Stokes (RANS) , all such applications
34
Duffner, John D. "The effects of manufacturing variability on turbine vane performance." Thesis, Massachusetts Institute of Technology, 2008. http://hdl.handle.net/1721.1/44933.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2008.This 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.
35
Groshenry, Christophe. "Preliminary design study of a micro-gas turbine engine." Thesis, Massachusetts Institute of Technology, 1995. http://hdl.handle.net/1721.1/10386.
Full text
36
Liu, Chunmeni 1970. "Dynamical system modeling of a micro gas turbine engine." Thesis, Massachusetts Institute of Technology, 2000. http://hdl.handle.net/1721.1/9249.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2000.Also available online at the MIT Theses Online homepage
Includes bibliographical references (p. 123).
Since 1995, MIT has been developing the technology for a micro gas turbine engine capable of producing tens of watts of power in a package less than one cubic centimeter in volume. The demo engine developed for this research has low and diabtic component performance and severe heat transfer from the turbine side to the compressor side. The goals of this thesis are developing a dynamical model and providing a simulation platform for predicting the microengine performance and control design, as well as giving an estimate of the microengine behavior under current design. The thesis first analyzes and models the dynamical components of the microengine. Then a nonlinear model, a linearized model, and corresponding simulators are derived, which are valid for estimating both the steady state and transient behavior. Simulations are also performed to estimate the microengine performance, which include steady states, linear properties, transient behavior, and sensor options. A parameter study and investigation of the startup process are also performed. Analysis and simulations show that there is the possibility of increasing turbine inlet temperature with decreasing fuel flow rate in some regions. Because of the severe heat transfer and this turbine inlet temperature trend, the microengine system behaves like a second-order system with low damping and poor linear properties. This increases the possibility of surge, over-temperature and over-speed. This also implies a potentially complex control system. The surge margin at the design point is large, but accelerating directly from minimum speed to 100% speed still causes surge. Investigation of the sensor options shows that temperature sensors have relatively fast response time but give multiple estimates of the engine state. Pressure sensors have relatively slow response time but they change monotonically with the engine state. So the future choice of sensors may be some combinations of the two. For the purpose of feedback control, the system is observable from speed, temperature, or pressure measurements. Parameter studies show that the engine performance doesn't change significantly with changes in either nozzle area or the coefficient relating heat flux to compressor efficiency. It does depend strongly on the coefficient relating heat flux to compressor pressure ratio. The value of the compressor peak efficiency affects the engine operation only when it is inside the range of the engine operation. Finally, parameter studies indicate that, to obtain improved transient behavior with less possibility of surge, over-temperature and over-speed, and to simplify the system analysis and design as well as the design and implementation of control laws, it is desirable to reduce the ratio of rotor mechanical inertia to thermal inertia, e.g. by slowing the thermal dynamics. This can in some cases decouple the dynamics of rotor acceleration and heat transfer. Several methods were shown to improve the startup process: higher start speed, higher start spool temperature, and higher start fuel flow input. Simulations also show that the efficiency gradient affects the transient behavior of the engine significantly, thereby effecting the startup process. Finally, the analysis and modeling methodologies presented in this thesis can be applied to other engines with severe heat transfer. The estimates of the engine performance can serve as a reference of similar engines as well.
by Chunmei Liu.
S.M.
37
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.
Full text
38
Crawford, 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.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2003.Includes 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.
39
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.
Full textAbstract:
"February 2001."Includes 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.
40
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.
Full textAbstract:
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.
41
Savoulides, Nicholas 1978. "Development of a MEMS turbocharger and gas turbine engine." Thesis, Massachusetts Institute of Technology, 2004. http://hdl.handle.net/1721.1/17815.
Full textAbstract:
Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2004.Includes bibliographical references.
As portable electronic devices proliferate (laptops, GPS, radios etc.), the demand for compact energy sources to power them increases. Primary (non-rechargeable) batteries now provide energy densities upwards of 180 W-hr/kg, secondary (rechargeable) batteries offer about 1/2 that level. Hydrocarbon fuels have a chemical energy density of 13,000-14,000 W-hr/kg. A power source using hydrocarbon fuels with an electric power conversion efficiency of order 10% would be revolutionary. This promise has driven the development of the MIT micro gas turbine generator concept. The first engine design measures 23 x 23 x 0.3 mm and is fabricated from single crystal silicon using MEMS micro-fabrication techniques so as to offer the promise of low cost in large production. This thesis describes the development and testing of a MEMS turbocharger. This is a version of a simple cycle, single spool gas turbine engine with compressor and turbine flow paths separated for diagnostic purposes, intended for turbomachinery and rotordynamic development. The turbocharger design described herein was evolved from an earlier, unsuccessful design (Protz 2000) to satisfy rotordynamic and fabrication constraints. The turbochargers consist of a back-to-back centrifugal compressor and radial inflow turbine supported on gas bearings with a design rotating speed of 1.2 Mrpm. This design speed is many times the natural frequency of the radial bearing system. Primarily due to the exacting requirements of the micron scale bearings, these devices have proven very difficult to manufacture to design, with only six near specification units produced over the course of three years. Six proved to be a small number for this development program since these silicon devices are brittle
(cont.) and do not survive bearing crashes at speeds much above a few tens of thousands of rpm. The primary focus of this thesis has been the theoretical and empirical determination of strategies for the starting and acceleration of the turbocharger and engine and evolution of the design to that end. Experiments identified phenomena governing rotordynamics, which were compared to model predictions. During these tests, the turbocharger reached 40% design speed (480,000 rpm). Rotordynamics were the limiting factor. The turbomachinery performance was characterized during these experiments. At 40% design speed, the compressor developed a pressure ratio of 1.21 at a flow rate of 0.13 g/s, values in agreement with CFD predictions. At this operating point the turbine pressure ratio was 1.7 with a flow rate of 0.26 g/s resulting in an overall spool efficiency of 19%. To assess ignition strategies for the gas turbine, a lumped parameter model was developed to examine the transient behavior of the engine as dictated by the turbomachinery fluid mechanics, heat transfer, structural deformations from centrifugal and thermal loading and rotordynamics. The model shows that transients are dominated by three time constants - rotor inertial (10⁻¹ sec), rotor thermal (lsec), and static structure thermal (10sec). The model suggests that the engine requires modified bearing dimensions relative to the turbocharger and that it might be necessary to pre-heat the structure prior to ignition ...
by Nicholas Savoulides.
Ph.D.
42
Shang, Tonghuo. "Influence of inlet temperature distortion on turbine heat transfer." Thesis, Massachusetts Institute of Technology, 1995. http://hdl.handle.net/1721.1/47370.
Full text
43
Prashanth, Prakash. "Post-combustion emissions control for aero-gas turbine engines." Thesis, Massachusetts Institute of Technology, 2018. https://hdl.handle.net/1721.1/122402.
Full textAbstract:
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
44
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.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2013.This 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.
45
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.
Full textAbstract:
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.
46
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.
Full text
47
Shannon, 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.
Full textAbstract:
Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2018.Cataloged 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.
48
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.
Full text
49
Lee, 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.
Full textAbstract:
Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2015.Cataloged from PDF version of thesis.
Includes bibliographical references (pages 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.
50
Zhang, K. "Turbulent combustion simulation in realistic gas-turbine combustors." Thesis, City, University of London, 2017. http://openaccess.city.ac.uk/17689/.
Full textAbstract:
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.