Thesis (Ph. D.)--Ohio State University, 2008. / Title from first page of PDF file. Includes bibliographical references (p. 144-152).

Thesis (Ph. D.)--Florida State University, 2005. / Advisor: Dr. Anjaneyulu Krothapalli, Florida State University, College of Engineering, Dept. of Mechanical Engineering. Title and description from dissertation home page (viewed June 14, 2005). Document formatted into pages; contains xx, 224 pages. Includes bibliographical references.

Thesis (M.S.)--Rochester Institute of Technology, 2004. / Typescript. Includes bibliographical references (leaves 65-66).

A facility for the investigation of stall flutter in aircraft engine compressors and fans was designed. Stall flutter was achieved in the test fan and verified through sonic and photographic methods. The frequency components of the sonic output during flutter were determined using a real-time analyzer. This frequency analysis indicated a dominant peak within 7 percent of the theoretical torsional natural frequency of the blades. Photographs taken during stall flutter indicated the presence of an interblade phase angle. The effect of blade stagger angle, flow incidence angle and solidity on flutter speed was determined. / Master of Science

The aerodynamic and thermodynamic performance of an aircraft gas turbine directly affects the fuel consumption of the engine and the life of the turbine components. Hence, it is important to be able to understand and predict the fluid flow and heat transfer in turbine blades to enable the modifications and improvements in the design process. The use of numerical experiments for the above purposes is becoming increasingly common. The present thesis is involved with the development of a flow solver for turbine flow and heat transfer computations. A 3-D Navier-Stokes code, the Moore Elliptic Flow Program (MEFP) is used to calculate steady flow and heat transfer in turbine rotor cascades. Successful calculations were performed on two different rotor profiles using a one-equation q-L transitional turbulence model. A series of programs was developed for the post-processing of the output from the flow solver. The calculations revealed details of the flow including boundary layer development, trailing edge shocks, flow transition and stagnation and peak heat transfer rates. The calculated pressure distributions, losses, transition ranges, boundary layer parameters and peak heat transfer rates to the blade are compared with the available experimental data. The comparisons indicate that the q-L transitional turbulence model is successful in predicting flows in transonic turbine blade rows. The results also indicate that the calculated loss levels are independent of the gridding used while the heat transfer rate predictions improve with finer grids. / Master of Science

The persistent thrust for a cleaner, greener environment has prompted air pollution regulations to be enforced with increased stringency by environmental protection bodies all over the world. This has prompted gas turbine manufacturers to move from non-premixed combustion to lean, premixed combustion. These lean premixed combustors operate quite fuel-lean compared to the stochiometric, in order to minimize CO and NOx productions, and are very susceptible to oscillations in any of the upstream flow variables. These oscillations cause the heat release rate of the flame to oscillate, which can engage one or more acoustic modes of the combustor or gas turbine components, and under certain conditions, lead to limit cycle oscillations. This phenomenon, called thermoacoustic instabilities, is characterized by very high pressure oscillations and increased heat fluxes at system walls, and can cause significant problems in the routine operability of these combustors, not to mention the occasional hardware damages that could occur, all of which cumulatively cost several millions of dollars. In a bid towards understanding this flow-flame interaction, this research works studies the heat release response of premixed flames to oscillations in reactant equivalence ratio, reactant velocity and pressure, under conditions where the flame preheat zone is convectively compact to these disturbances, using the G-equation. The heat release response is quantified by means of the flame transfer function and together with combustor acoustics, forms a critical component of the analytical models that can predict combustor dynamics. To this end, low excitation amplitude (linear) and high excitation amplitude (nonlinear) responses of the flame are studied in this work. The linear heat release response of lean, premixed flames are seen to be dominated by responses to velocity and equivalence ratio fluctuations at low frequencies, and to pressure fluctuations at high frequencies which are in the vicinity of typical screech frequencies in gas turbine combustors. The nonlinear response problem is exclusively studied in the case of equivalence ratio coupling. Various nonlinearity mechanisms are identified, amongst which the crossover mechanisms, viz., stoichiometric and flammability crossovers, are seen to be responsible in causing saturation in the overall heat release magnitude of the flame. The response physics remain the same across various preheat temperatures and reactant pressures. Finally, comparisons between the chemiluminescence transfer function obtained experimentally and the heat release transfer functions obtained from the reduced order model (ROM) are performed for lean, CH4/Air swirl-stabilized, axisymmetric V-flames. While the comparison between the phases of the experimental and theoretical transfer functions are encouraging, their magnitudes show disagreement at lower Strouhal number gains show disagreement.

Gas turbine reheat thrust augmenters known as afterburners are used to provide additional thrust during emergencies, take off, combat, and in supersonic flight of high-performance aircrafts. During the course of reheat development, the most persistent trouble has been the onset of high frequency combustion instability, also known as screech, invariably followed by rapid mechanical failure. The coupling of acoustic pressure upstream of the flame stabilizer with in-phase heat-release downstream, results in combustion instability by which the amplitude at various resonant modes — longitudinal (buzz — low frequency), tangential or radial (screech — high frequency) – amplifies leading to deterioration of the afterburner components. Various researchers in early 1950s have performed extensive testing on straight jet afterburners, to identify screech frequencies. Theoretical and experimental work at test rig level has been reported in the case of buzz to validate the heat release combustion models. In this work, focus is given to study the high frequency tangential combustion instability by vibro-acoustic software and the tests are conducted on the scaled bypass flow afterburner for confirmation of predicted screech frequencies. The wave equation for the afterburner is solved taking the appropriate geometry of the afterburner and taking into account the factors affecting the stability. Nozzle of the afterburner is taken into account by using the nozzle admittance condition derived for a choked nozzle. Screech liner admittance boundary condition is imposed and the effect on acoustic attenuation is studied. A new combustion model has been proposed for obtaining the heat release rate response function to acoustic oscillations. Acoustic wave – flame interactions involve unsteady kinetic, fluid mechanic and acoustic processes over a large range of time scales. Three types of flow disturbances exist such as : vortical, entropy, and acoustic. In a homogeneous, uniform flow, these three disturbance modes propagate independently in the linear approximation. Unsteady heat release also generates entropy and vorticity disturbances. Since flow is not accelerated in the region of uniform area duct, vortical and entropy disturbances are treated as in significant, as these disturbances are convected out into atmosphere like an open-ended tube, but these are considered in deriving the nozzle admittance condition. Heat release fluctuations that arise due to fluctuating pressure and temperature are taken into consideration. The aim is to provide results on how flames respond to pressure disturbances of different amplitudes and characterised by different length scales. The development of the theory is based on large activation energy asymptotics. One-dimensional conservation equations are used for obtaining the response function for the heat release rate assuming the laminar flamelet model to be valid. The estimates are compared with the published data and deviations are discussed. The normalized acoustic pressure variation in the afterburner is predicted using the models discussed earlier to provide an indication of the resonant modes of the pressure oscillations and the amplification and attenuation of oscillations caused by the various processes. Similar frequency spectrum is also obtained experimentally using a test rig for a range of inlet mean pressures and temperatures with combustion and core and bypass flows simulated, for confirmation of predicted results. Without the heat source only longitudinal acoustic modes are found to be excited in the afterburner test section. With heat release, three additional tangential modes are excited. By the use of eight probes in the circumferential cross section of afterburner it was possible to identify the tangential modes by their respective phase shift in the experiments. Comparison of normalized acoustic pressure and phase with and without the incorporation of perforate liner is made to study the effectiveness of the screech liner in attenuating the amplitude of screech modes. By the analysis, conclusion is drawn about modes that get effectively attenuated with the presence of perforate liner. Parametric study of screech liner porosity factor of 1.5 % has not shown appreciable attenuation. Whereas with 2.5 % porosity significant attenuation is noticed, but with 4 % porosity, the gain is very minimal. Hence, the perforate screech liner with the porosity of 2.5 % is finalized. From the rig runs, first pure screech tangential mode and second screech coupled tangential modes are captured. The theoretical frequencies for first and second tangential modes with their phases are comparable with experimental results. Though third tangential mode is predicted, it was not excited in the experiments. There was certain level of deviation in the prediction of these frequencies, when compared to the experimentally obtained values. For this test section of length to diameter ratio of 5, no radial modes are encountered both in the analysis and experiments in the frequency range of interest. In summary, an acoustic model has been developed for the afterburner combustor, taking into account the combustion response, the screech liner and the nozzle to study the acoustic instability of the afterburner. The model has been validated experimentally for screech frequencies using a model test rig and the results have given sufficient confidence to apply the model for full scale afterburners as a predictive design tool.

A novel combustor design, referred to as the Stagnation Point Reverse-Flow (SPRF) combustor, was recently developed that is able to operate stably at very lean fuel-air mixtures and with low NOx emissions even when the fuel and air are not premixed before entering the combustor. The primary objective of this work is to elucidate the underlying physics behind the excellent stability and emissions performance of the SPRF combustor. The approach is to experimentally characterize velocities, species mixing, heat release and flame structure in an atmospheric pressure SPRF combustor with the help of various optical diagnostic techniques: OH PLIF, chemiluminescence imaging, PIV and Spontaneous Raman Scattering. Results indicate that the combustor is primarily stabilized in a region downstream of the injector that is characterized by low average velocities and high turbulence levels; this is also the region where most of the heat release occurs. High turbulence levels in the shear layer lead to increased product entrainment levels, elevating the reaction rates and thereby enhancing the combustor stability. The effect of product entrainment on chemical timescales and the flame structure is illustrated with simple reactor models. Although reactants are found to burn in a highly preheated (1300 K) and turbulent environment due to mixing with hot product gases, the residence times are sufficiently long compared to the ignition timescales such that the reactants do not autoignite. Turbulent flame structure analysis indicates that the flame is primarily in the thin reaction zones regime throughout the combustor, and it tends to become more flamelet like with increasing distance from the injector. Fuel-air mixing measurements in case of non-premixed operation indicate that the fuel is shielded from hot products until it is fully mixed with air, providing nearly premixed performance without the safety issues associated with premixing. The reduction in NOx emissions in the SPRF combustor are primarily due to its ability to stably operate under ultra lean (and nearly premixed) condition within the combustor. Further, to extend the usefulness of this combustor configuration to various applications, combustor geometry scaling rules were developed with the help of simplified coaxial and opposed jet models.

Thesis. 1978. Ocean E.--Massachusetts Institute of Technology. Dept. of Ocean Engineering. / MICROFICHE COPY AVAILABLE IN ARCHIVES AND ENGINEERING. / Includes bibliographical references. / Ocean E.

Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Mechanical Engineering, 1981. / MICROFICHE COPY AVAILABLE IN ARCHIVES AND ENGINEERING. / Includes bibliographical references. / A theoretical study is presented of the fluid mechanics of the inlet vortex (or ground vortex) phenomenon. The vorticity field associated with the vortex is investigated using a secondary flow approach. In this approach the flow is assumed to be composed of an irrotational primary flow and a weak shear flow, with the vortex filaments associated with the latter being regarded as convected by the former. The potential flow field induced by the inlet-ground plane combination is computed using the panel method developed by · Hess, Mack and Stockman. Using the analysis, material lines (which coincide with vortex lines) can be tracked between a far upstream location, where this vorticity can be taken as known, and the engine face location. The deformation of the material lines thus shows directly the generation and amplification of the streamwise component of vorticity, which is responsible for the velocity distortion at the compressor face. Two representative flow configurations are considered, one with headwind only and one with the flow at forty-five degrees to the inlet axis of symmetry. Although the results so far yield only qualitative information, they appear to provide some insight into one mechanism associated with the inlet vortex formation. / by Henri Charles Viguier. / M.S.