Thesis (M. S.)--Mechanical Engineering, Georgia Institute of Technology, 2008. / Committee Chair: Dr. Richard W. Neu; Committee Member: Dr. David L. McDowell; Committee Member: Dr. W. Steven Johnson.

Thesis (Ph. D.)--University of Washington, 2000. / Vita. Includes bibliographical references (leaves 178-184).

Thesis (M.S.)--West Virginia University, 2000. / Title from document title page. Document formatted into pages; contains x, 157 p. : ill. (some col.) Includes abstract. Includes bibliographical references (p. 86-87).

Thesis (M.S. in Management)--Naval Postgraduate School, June 1990. / Thesis Advisor(s): Carrick, Paul M. Second Reader: Doyle, Richard B. "June 1990." Description based on title screen as viewed on October 16, 2009. DTIC Identifier(s): Jet Engines, Engine Components, Cost Analysis, Gas Turbines, Optimizations, Naval Logistics, Aircraft Maintenance, CIP(Component Improvement Program), Benefits, Redesign, Naval Aircraft, Mean Time Between Failure, Data Bases, Theses. Author(s) subject terms: Aircraft Turbine Engine Redesigns Component Improvement Program. Includes bibliographical references (p. 58-60). Also available in print.

Thesis (Ph. D.)--Aerospace Engineering, Georgia Institute of Technology, 2006. / Yedidia Neumeier, Committee Member ; Jerry Seitzman, Committee Member ; Fotis Sotiropoulos, Committee Member ; Tim Lieuwen, Committee Member ; suresh menon, Committee Chair.

Made available in DSpace on 2014-06-11T19:27:11Z (GMT). No. of bitstreams: 0 Previous issue date: 2009-03-27Bitstream added on 2014-06-13T20:35:17Z : No. of bitstreams: 1 silva_rep_me_guara.pdf: 1908689 bytes, checksum: 341019e7e665818b99415ff19ade0d87 (MD5) / Coordenação de Aperfeiçoamento de Pessoal de Nível Superior (CAPES) / Estudos têm sido realizados em relação à utilização de biodiesel em motores alternativos de ciclo Diesel, porém pouco material é encontrado quando se utiliza este combustível em turbinas a gás. Este trabalho analisou os parâmetros de desempenho e emissões de poluentes para várias misturas de biodiesel/querosene de aviação em um turboeixo Rover 1/S60. Os testes mostraram que não houve alterações significativas na operação do turboeixo. O estudo mostrou, também, que houve decréscimo de eficiência térmica e aumento de consumo de combustível com a utilização de misturas mais ricas em biodiesel. As emissões de poluentes também decresceram com o aumento de teor de biodiesel na mistura. / Several tests have been performed comparing diesel and several kinds of esthers (biodiesel), most in reciprocating engines but also in micro-gas-turbines. This work studies a stationary turboshaft Rover 1/S60 performance and pollution emissions in the utilization of several blends of aviation kerosene and biodiesel. The tests also show that no significative changes occurred in turboshaft operation. This study also shows an increase of fuel consumption and a decrease in thermal efficiency and a decrease of pollutant species emission as the higher biodiesel proportion at blends are used at turboshaft operation.

This thesis presents a computational fluid dynamic (CFD) study of high pressure gas turbine blade design with different realistic inlet temperature and velocity boundary conditions. The effects of blade shaping and inlet conditions can only be fully understood by considering the aerodynamics and heat transfer concurrently; this is in contrast to the sequential method of blade design for aerodynamics followed by cooling. The inlet boundary conditions to the NGV simulations are governed by the existence of discrete fuel injectors in the combustion chamber. An appreciation of NGV shaping design under engine realistic inflow conditions will allow for an identification of the correct three dimensional shaping parameters that should be considered for design optimisation. The Rolls-Royce efficient Navier-Stokes solver, HYDRA, was employed in all computational results for a transonic turbine stage. The single passage unsteady method based on the Fourier Shape Correction is adopted. The solver is validated under both rich burn (hot steak only) and the case with swirl inlet profiles for aerothermal characteristics; good agreement is noted with the validation data. Post processing methods were used in order to obtain time-averaged results and blade visualisations. Subsequently, a surrogate design optimisation methodology using machine learning combined with a Genetic Algorithm is implemented and validated. A study of the effect of NGV compound lean on stage performance is carried out and contrasted for uniform and rich burn inlets, and subsequently for lean burn. Compound lean is shown to produce a tip uploading at the rotor inlet, which is beneficial for rich burn, but detrimental for lean burn. It is also found that for rich burn, fluid driving temperature is more dominant than HTC in determining rotor blade heat transfer, the opposite sense to the uniform inlet. Also, for a lean burn inlet, there is another role reversal, with HTC dominating fluid driving temperature in determining heat transfer. A novel NGV design methodology is proposed that seeks to mitigate the combined effects of inlet hot streak and swirling flow. In essence, the concept two NGVs in a pair are shaped independently of each other, thus allowing the inlet flow non uniformity to be suitably accommodated. Finally, two numerical NGV optimisation studies are undertaken for the combined hot streak and swirl inlet for two clocking positions; vane impinging and passage aligned. Due to the prohibitive cost of unsteady CFD simulations for an optimisation strategy, a suitable objective function at the NGV exit plane is used to minimise rotor tip heat flux. The optimised shape for the passage case resulted in the lowest tip heat flux distribution, however the optimum shape for the impinging case led to the highest gain in stage efficiency. This therefore suggests that NGV lean and clocking position should be a consideration for future optimisation and design of the HP stage.

In 1937 the Manchester Engineering Firm Metropolitan Vickers (Metrovick) were awarded a development contract by the Air Ministry to develop a gas turbine for aircraft propulsion in conjunction with the Royal Aircraft Establishment at Farnborough. Over the next decade and a half, the company developed a number of gas turbine designs for a variety of applications in the air, at sea, and on land. This thesis examines the gas turbine work of Metropolitan Vickers, and how the company interacted with a variety of partners across both the military and the civilian realms. These included government research establishments such as the Royal Aircraft Establishment and the Admiralty Engineering Laboratory; commercial partners, such as the aero-engine manufacturer Armstrong Siddeley, Yarrow Shipbuilders, and the Great Western Railway, and state institutions such as the Ministries of Aircraft Production and Fuel and Power. It argues that Metrovick’s technical style was formed by the company’s existing heavy engineering plant business, which privileged design over development and production engineering. Compared to competitors such as Power Jets and Rolls Royce, Metrovick’s progress on aero-engine work was hampered by the lack of a development organisation; though technically advanced, its aircraft engines took a long time to be developed and would not reach production; a factor which was influential in the post-war sale of Metrovick’s aero-engine designs to Armstrong Siddeley. Metrovick did use its gas turbine experience to gain post-war contracts for both naval and civilian gas turbines. The Royal Navy adopted gas turbines for two roles: as lightweight powerplants for short-ranged fast-attack craft, and as part of major warship propulsion systems that were intended to overcome the perceived flaws of the Navy’s interwar steam plants. Metrovick was selected as a development partner because of the company’s existing naval business, as well as its gas turbine expertise. In the civilian realm, the company produced gas turbines for a wide range of applications ranging from railway locomotives to electrical power generation. Most of the customers for these designs were state or quasi-state institutions; this thesis argues that the postwar British state’s support for the civilian gas turbine shows that it was seen as a crucially British technology that could help improve industrial efficiency, as well as utilising indigenous energy resources. However, again Metrovick was content to rely on development contracts rather than commit itself to large-scale production. The company’s gas turbine designs were somewhat marginal to the wider heavy electrical business, and Metrovick never committed the kind of development resources to the gas turbine division that would have been required to produce successful products, nor did it attempt to sell its designs widely to relevant markets.

Within a gas turbine engine, the turbine nozzle guide vanes are subjected to very harsh conditions from the highly turbulent and hot gases exiting the combustor. The temperature and pressure fields exiting combustors are highly nonuniform and dictate the heat transfer and aero losses that occur in the turbine passages. To better understand these effects, the goal of this work was to develop an adjustable combustor exit profile simulator for the Turbine Research Facility (TRF) at the Air Force Research Laboratory. The TRF is a high temperature, high pressure, short duration blow-down test facility that is capable of matching several aerodynamic and thermal nondimensional engine parameters including Reynolds number, Mach number, pressure ratio, corrected mass flow, gas to metal temperature ratio, and corrected speed. The primary research objective was to design, install, and verify a non-reacting simulator device that can provide representative combustor exit total pressure and temperature profiles to the inlet of the TRF turbine test section. This required the upstream section of the facility to be redesigned into multiple concentric annuli that serve the purpose of injecting high momentum dilution jets and low momentum film cooling jets into a central annular chamber, similar to a turbine engine combustor. The design of the section allows for variations in injection levels to generate different pressure profiles with elevated turbulence. The dilution and film cooling temperatures can also be varied to create a variety of exit temperature profiles similar to real combustors. The impact of the generated temperature and pressure profiles on turbine heat transfer and secondary flow development was ultimately investigated. Proposed optimal inlet conditions for the turbine tested in this research effort were determined based on the measured data corresponding to the combustor simulator exit profiles that minimized vane heat transfer and total pressure loss. / Ph. D.

Gas turbine engines are prevalent in the today’s aviation and power generation industries. The majority of commercial aircraft use a turbofan gas turbine engines. Gas turbines used for power generation can achieve thermodynamic efficiencies as high as 60% when coupled with a steam turbine as part of a combined cycle. The success of gas turbines is a direct result of a half century’s development of the technology necessary to create such efficient, powerful, and reliable machines. One key area of technical advancement is the turbine cooling system. In short, increasing the turbine inlet temperature leads to a rise in cycle efficiency. Before the development of modern turbine cooling schemes, this temperature was limited by the softening temperature of the metallic turbine components. The evolution of component cooling systems – in conjunction with metallurgical advancements and the introduction of Thermal Barrier Coatings (TBC) – allowed for gradual increases in power output and efficiency. Today, the walls of gas turbine combustors are protected by a cool film that bypassed combustion; the 1st (and often 2nd) stage turbine blades and vanes are cooled via internal convection, a combination of turbulent channel flow, pin fin arrays, and impingement cooling; and some coolant air is bled onto the external surface of the blade and the blade endwall to establish a protective film on the exposed geometry. Modern research continues to focus on the optimization of these cooling designs, and a better understanding of the physics behind fluid behavior. The current study focuses on one particular cooling design: an impingement-effusion cooling system. While a single entity, the cooling schemes used in this system can be separated into impingement cooling on the backside iv of the cooled component and full coverage film cooling on the exposed surface. The result of this combination is a very high level of cooling effectiveness. The goal of this study is to explore a wide range of geometrical parameters and their effect on the overall cooling performance. Several parameters are taken outside the ranges normally investigated by the available literature. New methods of data comparison and normalization are offered in order to create an objective comparison of different configurations. Particular attention is given to the total coolant spent per unit surface area cooled. This study is also unique as it is a multi-modal heat transfer study, unlike the majority of impingement-effusion investigations, which only evaluate impingement heat transfer. Through determination of impingement heat transfer, film cooling effectiveness, and film cooling heat transfer on the target wall, a simplified heat transfer model of the cooled component is developed to show the relative impact of each parameter on the overall cooling effectiveness. The use of Temperature Sensitive Paint (TSP) for data acquisition allows for high resolution local heat transfer and effectiveness results. This has a quantitative benefit, giving the ability to average as desired and/or compare local data, for example the lateral distribution of film cooling effectiveness. However, the qualitative benefit of viewing the contours of heat transfer coefficient under an impinging jet array or downstream of a film cooling jet is instrumental in drawing conclusions about the behavior of the flow. The local data provides, in essence, a flow visualization on the test surface and adds (quite literally) another dimension to the heat transfer results. Impingement arrays with local extraction of coolant via effusion are able to produce higher overall heat transfer, as no significant cross flow is present to deflect the impinging jets. Low jet-to-target-plate spacing produces the highest yet most non-uniform heat transfer v distribution; at high spacing the heat transfer rate is much less sensitive to impingement height. Arrays with high hole-to-hole spacing and high jet Reynold’s number are more effective (per mass of coolant used) than tightly spaced holes at low jet Reyonld’s number. On the effusion side, staggered hole arrangements provide significantly higher film cooling effectiveness than their in-line counterparts as the staggered arrangement minimizes jet interactions and promotes a more even lateral distribution of coolant. These full coverage film cooling geometries typically show increases in effectiveness with each row of injection. Some additional cases were show with 15 film cooling rows, and generally the adiabatic wall temperature was decreasing through the last row. In the recovery region, results were highly dependant on blowing ratio; injection of excess coolant into the boundary layer at high blowing ratio allowed for cooling effectiveness to penetrate well downstream of the end of the array. From a heat transfer standpoint, compound angle injection resulted in higher enhancement than purely inclined injection, but this negative effect was outweighed by the substantial increase in film cooling effectiveness with the compounded geometry. Overall, the additive film superposition method under-predicted full coverage film cooling effectiveness trends for staggered hole arrangements; however, with more accurate estimation (or measurement) of recovery region trends for a single row of holes, this method may produce an acceptable result.