Hydrogen-powered engines have been gaining increasing interest due to the global concerns of the effects of hydrocarbon combustion on climate change. Gas turbines are suitable for operation on hydrogen fuel. This thesis reports the results of investigations of the special requirements of the fuel controller for a hydrogen gas turbine. In this investigation, a digital fuel controller for a hydrogen-fueled modified Pratt and Whitney PT6A-20 turboprop engine was successfully designed and implemented. Included in the design are safety measures to protect the operating personnel and the engine. A redundant fuel control is part of the final design to provide a second method of managing the engine should there be a malfunction in any part of the primary controller. Parallel to this study, an investigation of the existing hydrogen combustor design was performed to analyze the upper stability limits that were restricting the operability of the engine. The upstream propagation of the flame into the premixer, more commonly known as a flashback, routinely occurred at 150 shaft horsepower during engine testing. The procedures for protecting the engine from a flashback were automated within the fuel controller, significantly reducing the response time from the previous (manual) method. Additionally, protection measures were added to ensure the inter-turbine temperature of the engine did not exceed published limits. Automatic engine starting and shutdown procedures were also added to the control logic, minimizing the effort needed by the operator. The tested performance of the engine with each of the control functions demonstrated the capability of the controller. Methods to generate an engine-specific fuel control map were also studied. The control map would not only takes into account the operability limits of the engine, but also the stability limits of the premixing devices. Such a map is integral in the complete design of the engine fuel controller. / Master of Science

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The goal of a gas turbine engine designer is to reduce the amount of coolant used to cool the critical turbine surfaces, while at the same time extracting more benefit from the coolant flow that is used. Fan-shaped holes offer this opportunity, reducing the normal jet momentum and spreading the coolant in the lateral direction providing better surface coverage. The main drawback of fan-shaped cooling holes is the added manufacturing cost from the need for electrical discharge machining instead of the laser drilling used for cylindrical holes. This research focused on examining the performance of fan-shaped holes on two critical turbine surfaces; the vane and endwall. This research was the first to offer a complete characterization of film-cooling on a turbine vane surface, both in single and multiple row configurations. Infrared thermography was used to measure adiabatic wall temperatures, and a unique rigorous image transformation routine was developed to unwrap the surface images. Film-cooling computations were also done comparing the performance of two popular turbulence models, the RNG-kε and the v2-f model, in predicting film-cooling effectiveness. Results showed that the RNG-kε offered the closest prediction in terms of averaged effectiveness along the vane surface. The v2-f model more accurately predicted the separated flow at the leading edge and on the suction side, but did not predict the lateral jet spreading well, which led to an over-prediction in film-cooling effectiveness. The intent for the endwall surface was to directly compare the cooling and aerodynamic performance of cylindrical holes to fan-shaped holes. This was the first direct comparison of the two geometries on the endwall. The effect of upstream injection and elevated inlet freestream turbulence was also investigated for both hole geometries. Results indicated that fan-shaped film-cooling holes provided an increase in film-cooling effectiveness of 75% on average above cylindrical film-cooling holes, while at the same time producing less total pressure losses through the passage. The effect of upstream injection was to saturate the near wall flow with coolant, increasing effectiveness levels in the downstream passage, while high freestream turbulence generally lowered effectiveness levels on the endwall. / Ph. D.

The effect of global warming due to increased level of greenhouse gas emissions from coal fired thermal power plants and crisis of reliable energy resources has profoundly increased the importance of natural gas based power generation as a major alternative in the last few decades. Although gas turbine propulsion system had been primarily developed and technological advancements over the years had focused on application in civil and military aviation industry, use of gas turbine engines for land based power generation has emerged as the most promising candidate due to higher thermal efficiency, abundance of natural gas resources, development in generation of hydrogen rich synthetic fuel (Syngas) using advanced gasification technology for further improved emission levels and strict enforcement in emission regulations on installation of new coal based power plants. The fundamental thermodynamic principle behind gas turbine engines is Brayton cycle and higher thermal efficiency is achieved through maximizing the Turbine Inlet Temperature (TIT). Modern gas turbine engines operate well beyond the melting point of the turbine component materials to meet the enhanced efficiency requirements especially in the initial high pressure stages (HPT) after the combustor exit. Application of thermal barrier coatings (TBC) provides the first line of defense to the hot gas path components against direct exposure to high temperature gases. However, a major portion of the heat load to the airfoil and passage is reduced through injection of secondary air from high pressure compressor at the expense of a penalty on engine performance. External film cooling comprises a significant part of the entire convective cooling scheme. This can be achieved injecting coolant air through film holes on airfoil and endwall passages or utilizing the high pressure air required to seal the gaps and interfaces due to turbine assembly features. The major objective is to maximize heat transfer performance and film coverage on the surface with minimum coolant usage. Endwall contouring on the other hand provides an effective means of minimizing heat load on the platform through efficient control of secondary flow vortices. Complex vortices form due to the interaction between the incoming boundary layer and endwall-airfoil junction at the leading edge which entrain the hot gases towards the endwall, thus increasing surface heat transfer along its trajectory. A properly designed endwall profile can weaken the effects of secondary flow thereby improving the aerodynamic and associated heat transfer performance. This dissertation aims to investigate heat transfer characteristics of a non-axisymmetric contoured endwall design compared to a baseline planar endwall geometry in presence of three major endwall cooling features – upstream purge flow, discrete hole film cooling and mateface gap leakage under transonic operating conditions. The preliminary design objective of the contoured endwall geometry was to minimize stagnation and secondary aerodynamic losses. Upstream purge flow and mateface gap leakage is necessary to prevent ingestion to the turbine core whereas discrete hole cooling is largely necessary to provide film cooling primarily near leading edge region and mid-passage region. Different coolant to mainstream mass flow ratios (MFR) were investigated for all cooling features at design exit isentropic Mach number (0.88) and design incidence angle. The experiments were performed at Virginia Tech's quasi linear transonic blow down cascade facility. The airfoil span increases in the mainstream flow direction in order to match realistic inlet/exit airfoil surface Mach number distribution. A transient Infrared (IR) thermography technique was employed to measure the endwall surface temperature and a novel heat transfer data reduction method was developed for simultaneous calculation of heat transfer coefficient (HTC) and adiabatic cooling effectiveness (ETA), assuming a 1D semi-infinite transient conduction. An experimental study on endwall film cooling with endwall contouring at high exit Mach numbers is not available in literature. Results indicate significant benefits in heat transfer performance using the contoured endwall in presence of individual (upstream slot, discrete hole and mateface gap) and combined (upstream slot with mateface gap) cooling flow features. Major advantages of endwall contouring were observed through reduction in heat transfer coefficient and increase in coolant film coverage by weakening the effects of secondary flow and cross passage pressure differential. Net Heat Flux Reduction (NHFR) analysis was carried out combining the effect of heat transfer coefficient and film cooling effectiveness on both endwall geometries (contoured and baseline) where, the contoured endwall showed major improvement in heat load reduction near the suction side of the platform (upstream leakage only and combined upstream with mateface leakage) as well as further downstream of the film holes (discrete hole film cooling). Detailed interpretation of the heat transfer results along with near endwall flow physics has also been discussed. / Ph. D.

The current demands for higher performance in gas turbine engines can be reached by raising combustion temperatures to increase thermal efficiency. Hot combustion temperatures create a harsh environment which leads to the consideration of the durability of the combustor and turbine sections. Improvements in durability can be achieved through understanding the interactions between the combustor and turbine. The flow field at a combustor exit shows non-uniformities in pressure, temperature, and velocity in the pitch and radial directions. This inlet profile to the turbine can have a considerable effect on the development of the secondary flows through the vane passage. This thesis presents a computational study of the flow field generated in a non-reacting gas turbine combustor and how that flow field convects through the downstream stator vane. Specifically, the effect that the combustor flow field had on the secondary flow pattern in the turbine was studied. Data from a modern gas turbine engine manufacturer was used to design a realistic, low speed, large scale combustor test section. This thesis presents the results of computational simulations done in parallel with experimental simulations of the combustor flow field. In comparisons of computational predictions with experimental data, reasonable agreement of the mean flow and general trends were found for the case without dilution jets. The computational predictions of the combustor flow with dilution jets indicated that the turbulence models under-predicted jet mixing. The combustor exit profiles showed non-uniformities both radially and circumferentially, which were strongly dependent on dilution and cooling slot injection. The development of the secondary flow field in the turbine was highly dependent on the incoming total pressure profile. For a case with a uniform inlet pressure in the near-wall region no leading edge vortex was formed. The endwall heat transfer was found to also depend strongly on the secondary flow field, and therefore on the incoming pressure profile from the combustor. / Master of Science