The objective of this work is to examine the flow structure and heat transfer distribution of a squealer tip rotor blade with purge flow cooling and provide a comparison with a basic flat tip rotor blade without purge flow cooling, under transonic conditions and high inlet free stream turbulence intensity. The blade design was provided by Solar Turbines Inc., and consists of a double squealer around the pressure and suction sides, two purge flow blowing holes located downstream of the leading edge and mid-chord, four ribs in the mid-chord region and a trailing edge bleeder exiting on the pressure side. Blade cavity depth is 2.29 mm (0.09 in.) and the total blade turning angle is 107.5°. Tests were performed in a blow-down facility at a turbulence intensity of 12%, in a seven bladed 2-D linear cascade at transonic conditions. Experiments were conducted at isentropic exit Mach numbers of 0.85 and 1.05, corresponding to Reynolds numbers based on axial chord of 9.75x10^5 and 1.15x10^6, respectively, and tip clearance gaps of 1% and 2% of the scaled engine blade span. A blowing ratio of 1.0 was used in the squealer tip experiments. Detailed heat transfer coefficient and film cooling effectiveness distributions were obtained using an infrared thermography technique, while oil flow visualization was used to investigate the flow patterns in the blade tip region. With the addition of a squealer tip, leakage flow was found to decrease, as compared to a flat tip blade. With increasing tip clearance gap, the heat transfer coefficients within the cavity and along the squealer rim were found to decrease and increase, respectively. Film cooling effectiveness decreased with increasing tip clearance gap and was mainly observed within the squealer cavity. The maximum heat transfer coefficient was observed on the leading edge, however, comparatively large values were observed on the mid-chord ribs. The presence of the ribs, greatly affected the flow structure and heat transfer distributions within the cavity and downstream towards the trailing edge. / Master of Science

<p>  </p> <p>In this thesis, two optimizations of squealer tip geometries are completed for first stage turbine blades for use in a supercritical carbon dioxide turbine. First, an optimization is performed on a baseline trapezoidal turbine blade and a set of solution geometries is chosen from along the Pareto front. Next, a second optimization is completed on an advanced blade design and the geometries are grouped by performance characteristics and geometric features. The success of similar geometries across these two optimizations is also analyzed and demonstrates consistency of performance increases from tip geometries over the baseline geometry. An analysis of a flat tip geometry in a stationary condition is also performed to begin validation of annular cascades as a method for testing squealer tip geometries. </p>

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.