No description available.

No description available.

No description available.

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 unsteady interaction of shock waves emerging from the trailing edge of modern turbine nozzle guide vanes and impinging on downstream rotor blades is modeled in a linear cascade. The Reynolds number based on blade chord and exit conditions (5*10^6) and the exit Mach number (1.2) are representative of modern engine operating conditions. The relative motion of shocks and blades is simulated by sending a shock wave along the leading edges of the linear cascade instead of moving the blades through an array of stationary shock waves. The blade geometry is a generic version of a modern high turning rotor blade with transonic exit conditions. The blade is equipped with a showerhead film cooling scheme. Heat flux, surface pressure and surface temperature are measured at six locations on the suction side of the central blade. Pressure measurements are taken with Kulite XCQ-062-50a high frequency pressure transducers. Heat flux data is obtained with Vatell HFM-7/L high speed heat flux sensors. High speed heat flux and pressure data are recorded during the time of the shock impact with and without film cooling. The data is analyzed in detail to find the relative magnitudes of the shock effect on the heat transfer coefficient and the recovery temperature or adiabatic wall temperature (in the presence of film cooling). It is shown that the variations of the heat transfer coefficient and the film effectiveness are less significant than the variations of recovery temperature. The effect of the shock is found to be similar in the cases with and without film cooling. In both cases the variation of recovery temperature induced by the shock is shown to be the main contribution to the overall unsteady heat flux. The unsteady heat flux is compared to results from different prediction models published in the literature. The best agreement of data and prediction is found for a model that assumes a constant heat transfer coefficient and a temperature difference calculated from the unsteady surface pressure assuming an isentropic compression. / Ph. D.

This work presents a comparison of thin film gauge (TFG) and infrared (IR) thermography measurement techniques to simultaneously determine heat transfer coefficient and film cooling effectiveness. The first comparison was with an uncooled vane where heat transfer coefficient was measured at Mex=0.77 and Tu=16%. Relatively good agreement was found between the results of the two methods and the effect of recovery temperature and data reduction time was analyzed. Improvements were made to the experimental set up for the next comparison, a showerhead film cooled vane. This geometry was tested at BR=0, 2.0, Mex=0.76 and Tu=16%. The TFG and IR results did not compare well for heat transfer coefficient or film cooling effectiveness. The effects of measured and calculated recovery temperature were analyzed as well as the respective data reduction methods, though the analysis could not account for the effectiveness trend seen on the suction surface. Finally, a vane with showerhead and shaped film cooling holes were presented at BR=0, 1.7, 2.0, 2.8, Mex=0.85, and Tu=13% to assess a new film cooling geometry measured with the IR technique. Similarities on the suction surface trend between the different film cooled geometries tested with IR indicate a flaw in the experiment that will require further analysis, changes and testing to complete the comparison with TFG. / Master of Science

In the drive to increase cycle efficiency, gas turbine designers have increased turbine inlet temperatures well beyond the metallurgical limits of engine components. In order to prevent failure and meet life requirements, turbine components must be cooled well below these hot gas temperatures. Film cooling is a widely employed cooling technique whereby air is extracted from the compressor and ejected through holes on the surfaces of hot gas path components. The cool air forms a protective film around the surface of the part. Accurate numerical prediction of film cooling performance is extremely difficult so experiments are required to validate designs and CFD tools. In this study, a first stage turbine vane with five rows of showerhead cooling was instrumented with platinum thin-film gauges to experimentally characterize film cooling performance. The vane was tested in a transonic vane cascade in Virginia Tech's heated, blow-down wind tunnel. Two freestream exit Mach numbers of 0.76 and 1.0—corresponding to exit Reynolds numbers based on vane chord of 1.1x106 and 1.5x106, respectively—were tested at an inlet freestream turbulence intensity of two percent and an integral length scale normalized by vane pitch of 0.05. The showerhead cooling scheme was tested at blowing ratios of 0 (no cooling), 1.5, and 2.0 and a density ratio of 1.35. Midspan Nusselt number and film cooling effectiveness distributions over the surface of the vane are presented. Film cooling was found to augment heat transfer and reduce adiabatic wall temperature downstream of injection. In general, an increase in blowing ratio was shown to increase augmentation and film cooling effectiveness. Increasing Reynolds number was shown to increase heat transfer and reduce effectiveness. Finally, comparing low turbulence measurements (Tu = 2%) to measurements performed at high freestream turbulence (Tu = 16%) by Nasir et al. [13] showed that large-scale high freestream turbulence can reduce heat transfer coefficient downstream of injection. / Master of Science

The single-plate interferometer is a powerful flow visualization and aerodynamic measurement tool. It can provide full-field data for the density distribution in a non-intrusive manner, and it can be used for highly unsteady flows. While the device itself represents a large decrease in complexity over other forms of interferometry, the data reduction procedure has traditionally been laborious and difficult. To remove these difficulties and to improve the accuracy of the Virginia Tech interferometer setup, the software has been revamped into a black box design removing the need to handle the code directly. Furthermore, the software has been made to be platform independent by implementing the algorithms using the Java programming language. New hardware has also been added which further simplifies the setup procedure. The improved setup and the new software is used to study the flow around a film cooled turbine blade in the Virginia Tech cascade wind tunnel. The study of this flowfield is used as a validation for the new algorithms and to illustrate the ease of use of the system. Through this analysis, the density distribution for the entire flowfield is acquired. Furthermore the use of Plexiglas as window material was tried. This proved to work, however the manufacturing processing of these windows proved relatively difficult. Studying the film layer close to the surface proved difficult because of inherent limitations with the single-plate interferometer. / Master of Science

Increased efficiency and durability of gas turbine components is driven by demands for reduced fuel consumption and increased reliability in aircraft and power generation applications. The complex flow near the endwall of an axial gas turbine has been identified as a significant contributing factor to aerodynamic loss and increased part temperatures. Three-dimensional (non-axisymmetric) contouring of the endwall surface has been shown to reduce aerodynamic losses, but the effect of the contouring on endwall heat transfer is not well understood. This research focused on understanding the general flow physics of contouring and the sensitivity of the contouring to perturbations arising from leakage features present in an engine. Two scaled low-speed cascades were designed for spatially-resolved measurements of endwall heat transfer and film cooling. One cascade was intended for flat and contoured endwall studies without considering typical engine leakage features. The other cascade modeled the gaps present between a stator and rotor and between adjacent blades on a wheel, in addition to the non-axisymmetric endwall contouring. Comparisons between a flat and contoured endwall showed that the contour increased endwall heat transfer and increased turbulence in the forward portion of the passage due to displacement of the horseshoe vortex. However, the contour decreased heat transfer further into the passage, particularly in regions of high heat transfer, due to delayed development of the passage vortex and reduced boundary layer skew. Realistic leakage features such as the stator-rotor rim seal had a significant effect on the endwall heat transfer, although leakage flow from the rim seal only affected the horseshoe vortex. The contours studied were not effective at reducing the impact of secondary flows on endwall heat transfer and loss when realistic leakage features were also considered. The most significant factor in loss generation and high levels of endwall heat transfer was the presence of a platform gap between adjacent airfoils. / Ph. D.