Autoignition and oxidation of two Methane (CH4) and Dimethyl Ether (CH3OCH3 or DME) mixtures in air were studied in shock tubes over a wide range of equivalence ratios at elevated temperatures and pressures. These experiments were conducted in the reflected shock region with pressures ranging from 0.8 to 35.7 atmospheres, temperatures ranging from 913 to 1650 K, and equivalence ratios of 2.0, 1.0, 0.5, and 0.3. Ignition delay times were obtained from shock-tube endwall pressure traces for fuel mixtures of CH4/CH3OCH3 in ratios of 80/20 percent volume and 60/40 percent volume, respectively. Close examination of the data revealed that energy release from the mixture is occurring in the time between the arrival of the incident shock wave and the ignition event. An adjustment scheme for temperature and pressure was devised to account for this energy release and its effect on the ignition of the mixture. Two separate ignition delay correlations were developed for these pressure- and temperature-adjusted data. These correlations estimate ignition delay from known temperature, pressure, and species mole fractions of methane, dimethyl ether, and air (0.21 O2 + 0.79 N2). The first correlation was developed for ignition delay occurring at temperatures greater than or equal to 1175 K and pressures ranging from 0.8 to 35.3 atm. The second correlation was developed for ignition delay occurring at temperatures less than or equal to 1175 K and pressures ranging from 18.5 to 40.0 atm. Overall good agreement was found to exist between the two correlations and the data of these experiments. Findings of these experiments also include that with pressures at or below ten atm, increased concentrations of dimethyl ether will consistently produce faster ignition times. At pressures greater than ten atmospheres it is possible for fuel rich mixtures with lower concentrations of dimethyl ether to give the fastest ignition times. This work represents the most thorough shock tube investigation for oxidation of methane with high concentration levels of dimethyl ether at gas turbine engine relevant temperatures and pressures. The findings of this study should serve as a validation for detailed chemical kinetics mechanisms.

Gas turbine engines are one of the primary sources of electricity currently in the world. The assembly consists of multiple subsystems all of which are placed within an acoustic enclosure. Dilution ventilation is a widely used and recommended means of protection against the risk of explosion within gas turbine acoustic enclosures. Leakage and accumulation of flammable gas and its ignition from the turbine are one of the reasons for this explosion. The ISO regulations provide a safety criterion for the allowable size of flammable gas cloud as a proportion of enclosure volume. The effectiveness of the ventilation system in maintaining the clouds and the optimum placement of gas detectors to detect these clouds is determined through CFD simulations. This thesis deals with the primary step in the process of detecting the problematic regions which might allow the growth of the gas clouds to pass safety levels if the engine is not shut down. Using the same methodology, other possible configurations of the ventilation system and the effect on the flow due to the addition of a platform and movement of junction boxes are also studied. It is found out that the configuration where the inlet and outlets of the standard configuration are interchanged is arguably the best followed by the configuration with a split inlet, standard configuration, and lastly, the one with reduced volumetric flow rate. The platform and the movement of junction boxes across the walkway/platform do not affect the flow in any significant manner. The vertical movement of junction boxes affect the size of the low-speed region behind the inlet plenum. The thesis satisfied its objectives and provides a guideline for the same.

No description available.

This dissertation presents a detailed experimental and numerical analysis of the aerothermal characteristics of the turbine extremity regions i.e. the blade tip and endwall regions. The heat transfer and secondary flow characteristics were analyzed for different engine relevant configurations and exit Mach/Reynolds number conditions. The experiments were conducted in a linear blowdown cascade at transonic high turbulence conditions of Mexit ~ 0.85, 0.60 and 1.0, with an inlet turbulence intensity of 16% and 12% for the vane and blade cascade respectively. Transient infrared (IR) thermography technique and surface pressure measurement were used to map out the surface heat transfer coefficient and aerodynamic characteristics. The experiments were complemented with computational modeling using the commercial RANS equation solver ANSYS Fluent. The CFD results provided further insight into the local flow characteristics in order to elucidate the flow physics which govern the measured heat transfer characteristics. The results reveal that the highest heat transfer exists in regions with local flow reattachment and new-boundary layer formation. Conversely, the lowest heat transfer occurs in regions with boundary layer thickening and separation/lift-off flow. However, boundary layer separation results in additional secondary flow vortices, such as the squealer cavity vortices and endwall auxiliary vortex system, which significantly increase the stage aerodynamic losses. Furthermore, these vortices result in a low film-cooling effectiveness as was observed on a squealer tip cavity with purge flow. Finally, the importance of transonic experiments in analyzing the turbine section heat transfer and flow characteristics was underlined by the significant shock-boundary layer interactions that occur at high exit Mach number conditions. / Ph. D.

The second law is still rarely used as a design decision tool. However, information obtained from second law analysis is valuable in the design process for thermodynamic systems. This investigation reviews the past development of second law analysis. A clear, operational vocabulary is established. Then two examples of microcomputer-based design procedures are presented. The first is a second law analysis of refrigeration cycles. As a part of this example, the correlation of physical property data by simple methods is demonstrated. The second example is a second law analysis of gas turbine systems. The salient point is that all of this can be done on a microcomputer, and consequently is readily available to any engineer. / M.S.

This thesis describes research on acoustically excited bluff body flow-fields, motivated by the problem of combustion instabilities in devices utilizing these types of flame-holders. Vortices/convective-structures play a dominant role in perturbing the flame during these combustion instabilities. This thesis addresses a number of issues related to the origin, evolution and the interaction of these structures with the flame. The first part of this thesis reviews the fluid mechanics of non-reacting and reacting bluff body flows. The second part describes the spatio/temporal characteristics of bluff-body flames responding to excitation. The key processes controlling the flame response have been identified as 1) the anchoring of the flame at the bluff body, 2) the excitation of flame-front wrinkles by the oscillating velocity field and 3) flame propagation normal to itself at the local flame speed. The first two processes control the growth of the flame response and the last process controls the decay. The third part of this thesis describes the effect of acoustic excitation on the velocity field of reacting bluff body flows. Acoustic disturbances excite the Kelvin-Helmholtz (KH) instability of the reacting shear layer. This leads to a spatially decaying vorticity field downstream of the bluff body in the shear layers. The length over which the decay occurs was shown to scale with the length of the recirculation zone of the bluff body, i.e. the length over which the velocity profile transitions from shear layer to wake. The flame influences this decay process in two ways. Gas expansion across the flame reduces the extent of shear by reducing the magnitude of negative velocities within the recirculation zone. This combined with the higher product diffusivity reduces the length of the recirculation zone, thereby further augmenting the decay of the vorticity fluctuations. Lastly, these results also revealed phase jitter - a cycle-to-cycle variation in the position of the rolled-up vortices. Close to the bluff-body, phase jitter is very low but increases monotonically in the downstream direction. This leads to significant differences between instantaneous and ensemble averaged flow fields and, in particular, the decay rate of the vorticity in the downstream direction.

Durante as últimas décadas, um dos desafios da indústria aeroespacial é em relação ao aumento da eficiência dos motores de turbinas a gás. A eficiência dos motores é limitada principalmente pela temperatura dos gases de combustão, que não pode ser aumentada devido às limitações intrínsecas relacionadas ao uso das superligas à base de Ni nas partes quentes da turbina, onde as temperaturas podem atingir valores acima de 1000 °C. Este trabalho visa caracterizar novos materiais do sistema Nb-Ti para aplicações aeronáuticas, materiais de baixa massa específica que podem substituir as superligas de Ni. As ligas foram produzidas através de fusão a arco, tratadas termicamente a 1200 °C durante 48 h e expostas a temperaturas semelhantes às encontradas na seção de baixa pressão de um turborreator. Os materiais foram caracterizados em termos de composição química, propriedades mecânicas e microestrutura. Foram utilizadas as seguintes técnicas: difração de raios; microscopia eletrônica de varredura, espectrometria de raios X por energia dispersiva e Microscopia Eletrônica de Transmissão. A caracterização microestrutural revelou que as ligas expostas a 1000 °C durante 168 h apresentam uma microestrutura de duas fases composta principalmente de uma matriz ?0-BCC (Nb/Ti) com precipitados de uma segunda fase rica em titânio. Microestruturas de duas fases também foram observadas para as ligas expostas a 800 °C durante 168 h, na qual uma matriz ?0-BCC (Nb/Ti) com precipitados de uma segunda fase identificada como O2-Ti2NbAl foi observada. As ligas estudadas apresentaram massa específica inferior às superligas à base de Ni normalmente utilizadas na indústria aeronáutica. Em termos de propriedades mecânicas, as amostras expostas e testadas a 1000°C apresentaram valores baixos de resistência à compressão (100 MPa) quando comparado as amostras expostas e testadas a 800 °C (565 MPa). Pelos resultados de oxidação observou-se uma maior resistência a oxidação das ligas testadas a 800 °C, porém tanto a 1000 °C como a 800 °C não houve a formação de um filme protetor. / During the last decades, one of the challenges in the aerospace industry is with respect to increase the efficiency of gas turbine engines. The efficiency of the engines is a function of temperatures of the fluel gas, which cannot be increased because of intrinsic limitations related the use of Ni-based superalloys in the hot parts, where temperatures can reach values above 1000 °C. This work aims to investigate new materials in the Nb-Ti system, with low-density materials that may substitute Ni superalloys. The alloys were processed via arc melting, heat treated at 1200°C for 48h and exposed at temperatures similar to those encountered at the low-pressure section in a turbojet engine. The materials were characterized in terms of chemical composition, mechanical properties and microstructure. The following techniques have been used: X-ray diffraction; Scanning Electron Microscopy; Energy Dispersive X-ray Spectrometry and Transmission Electron Microscopy. The microstructural characterization have revealed that the alloys exposed at 1000 °C for 168 hours present a two-phase microstructure composed mainly of a ?0-BCC (Nb/Ti) matrix with precipitations of a second phase rich in titanium. Two-phase microstructures were also observed for the alloys exposed at 800 °C for 168 hours, where a ?0-BCC (Nb/Ti) matrix is observed with precipitates of a second phase identified as O2-Ti2NbAl. The studies alloys reported a lower density when comparing with the Ni based superalloys normally used in the aeronautical industry. In terms of mechanical properties, specimens exposed and tested at 1000 °C showed lower values of compressive strength (100 MPa) than those exposed and tested at 800 °C (565 MPa). The oxidation results allowed to observe a higher oxidation resistance of the alloys tested at 800 °C, however there was no protective film formation at 1000 °C as at 800 °C.

Pour des raisons de sécurité, les moteurs aéronautiques doivent pouvoir redémarrer en vol sur toute leur plage d'opération. Mais les contraintes sur les émissions polluantes nécessitent le développement de nouvelles chambres de combustion dont la conception peut détériorer les capacités d'allumage du moteur. Afin d'améliorer la compréhension du processus d'allumage et d'aider à l'optimisation de la conception, les recherches actuelles combinent les études expérimentales de plus en plus complexes et les simulation numériques hautes fidélités. Dans ce travail, l'étude numérique du processus d'allumage des chambres de combustion aéronautiques, de l'étincelle à la propagation azimutale de la flamme, est conduite avec plusieurs objectifs: améliorer la robustesse et la confiance de l'outil LES pour l'étude de l'allumage, étudier les mécanismes qui affectent l'allumage dans des conditions représentatives des conditions réelles et enfin améliorer les méthodes bas-ordre pour la prédiction des performances d'allumage. Dans une première partie, la SGE d'un monobruleur installé au CORIA permet de mettre en évidence les bons résultats de la LES et de construire une base de données pour l'analyses des mécanismes d'extinction. Ces données sont aussi utilisées pour développer une méthodologie permettant de prédire les performances d'allumage à bas coût en utilisant les résultats d'une SGE non-réactive. Dans une seconde partie, la propagation inter-secteur est investiguée par l'étude de deux cas expérimentaux et la SGE est capable de reproduire les modes de propagation mais aussi les temps d'allumage avec précision. Sur la bases de ces bons résultats, une analyse plus fine de la simulation permet d'identifier les mécanismes qui contrôlent la propagation de la flamme. / For safety reasons, in-flight relight of the engine must be guaranteed over a wide range of operating conditions but the increasing stringency of pollutant emission constraints requires the development of new aero-engine combustor whose design might be detrimental to the ignition capability. To improve the knowledge of the ignition process in aeronautical gas turbines and better combine conflicting technological solutions, current research relies on both complex experimental investigation and high fidelity numerical simulations. In this work, numerical study of the ignition process in gas turbines from the energy deposit to the light-around is performed with several objectives: increase the level of confidence of Large Eddy Simulations tool for the analysis of the ignition process, investigate the mechanisms controlling ignition in conditions representative of realistic aeronautical gas turbine flows and improve the low-order methodologies for the prediction of ignition performance. In a first part, LES of the single burner installed at CORIA (France) is carried out and allows to highlight the LES accuracy and to build a database on which the main mechanisms controlling the ignition success are identified. Based on these results, a methodology is developed to predict the ignition performance at a low computational cost using the non-reacting flow statistics only. In a second part, the light-around process is studied on two experimental set-ups and the very good agreement of the LES results with experiments is the starting point from an analysis of the mechanisms driving the flame propagation process.

Velocity-coupled combustion instability is a major issue facing lean combustor design in modern gas turbine applications. In this study, we analyze the complex acoustic field excited by a transverse acoustic mode in an annular combustor. This work is motivated by the need to understand the various velocity disturbance mechanisms present in the flame region during a transverse instability event. Recent simulation and experimental studies have shown that much of the flame response during these transverse instabilities may be due to the longitudinal motion induced by the fluctuating pressure field above the nozzles. This transverse to longitudinal coupling has been discussed in previous work, but in this work it is given a robust acoustic treatment via computational methods in order to verify the mechanisms by which these two motions couple. We will provide an in-depth discussion of this coupling mechanism and propose a parameter, Rz, also referred to as the Impedance Ratio, in order to compare the pressure/velocity relationship at the nozzle outlet to quasi one-dimensional theoretical acoustic approximations. A three-dimensional inviscid simulation was developed to simulate transversely propagating acoustic pressure waves, based on an earlier experiment designed to measure these effects. Modifications to this geometry have been made to account for lack of viscosity in the pure acoustic simulation and are discussed. Results from this study show that transverse acoustic pressure excites significant axial motion in and around the nozzle over a large range of frequencies. Furthermore, the development of Rz offers a defined physical parameter through which to reference this important velocity-coupled instability mechanism. Therefore, this study offers an in-depth and quantifiable understanding of the instability mechanism caused by transversely propagating acoustic waves across a combustor inlet, which can be applied to greatly improve annular combustor design in future low-emissions gas turbine engines.

High-pressure experiments and chemical kinetics modeling were performed for laminar spherically expanding flames for methane/air, ethane/air, methane/ethane/air and propane/air mixtures at pressures between 1 and 10 atm and equivalence ratios ranging from 0.7 to 1.3. All experiments were performed in a new flame speed facility capable of withstanding initial pressures up to 15 atm. The facility consists of a cylindrical pressure vessel rated up to 2200 psi. Vacuums down to 30 mTorr were produced before each experiment, and mixtures were created using the partial pressure method. Ignition was obtained by an automotive coil and a constant current power supply capable of reducing the spark energy close to the minimum ignition energy. Optical cine-photography was provided via a Z-type schlieren set up and a high-speed camera (2000 fps). A full description of the facility is given including a pressure rating and a computational conjugate heat transfer analysis predicting temperature rises at the walls. Additionally, a detailed uncertainty analysis revealed total uncertainty in measured flame speed of approximately +-0.7 cm/s. This study includes first-ever measurements of methane/ethane flame speeds at elevated pressures as well as unique high pressure ethane flame speed measurements. Three chemical kinetic models were used and compared against measured flame velocities. GRI 3.0 performed remarkably well even for high-pressure ethane flames. The C5 mechanism performed acceptably at low pressure conditions and under-predicted the experimental data at elevated pressures. Measured Markstein lengths of atmospheric methane/air flames were compared against values found in the literature. In this study, Markstein lengths increased for methane/air flames from fuel lean to fuel rich. A reverse trend was observed for ethane/air mixtures with the Markstein length decreasing from fuel lean to fuel rich conditions. Flame cellularity was observed for mixtures at elevated pressures. For both methane and ethane, hydrodynamic instabilities dominated at stoichiometric conditions. Flame acceleration was clearly visible and used to determine the onset of cellular instabilities. The onset of flame acceleration for each high-pressure experiment was recorded.