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This paper summarizes UTRC’s first-year effort in designing vane ring and integrally bladed rotor made from monolithic ceramics. AERODYNAMIC DESIGN A meanline optimization was performed that resulted in a first-stage design consisting of 15 vanes and 27 blades. The blade count was reduced from that of the baseline design to improve ceramic manufacturability and to allow for increased blade thickness (See Figure 2). Blade trailing edge thickness was increased for improved FOD (foreign object damage) resistance (to be discussed later). The flow path was unchanged from the baseline to reduce the number of dissimilar parts between the two designs. Axial chords were held to those of the baseline to limit airfoil aspect ratios and the associated profile losses, while reaction was limited to 45% to limit blade exit flow angle and Mach number and the associated transition duct losses. This reaction is a substantial increase over that of the baseline, which was held to 25% to reduce the effects of combustion gas temperature on the uncooled single crystal superalloy vanes. To match the compressor performance map, turbine power was varied until the turbine exit flow parameter matched that predicted by the cycle analysis. The meanline optimization resulted in a design with an increase in aerodynamic stage efficiency of between 1 and 2% over that of the baseline. ST5 ST5+ Figure 2. Airfoil has been thickened to improve FOD resistance The meanline and flow path data were converted to an initial three-dimensional design, which was optimized for pressure distribution using a steady, inviscid flow solver. Correlations between pressure distribution characteristics and losses in viscous flows have been derived from experience. The relative Mach number distribution at mid-span of this design can be seen in Figure 3. The optimized aerodynamic design was then converted to CAD format and integrated with the vane ring and integrally bladed rotor (IBR) disk designs to be analyzed for structural integrity and ceramic manufacturability. Another objective of the aerodynamic analysis was to provide thermal boundary conditions for the structural analysis of the ceramic vane ring and IBR. To obtain these boundary conditions, two Navier-Stokes computations on dense meshes were performed. The first case was run with adiabatic wall thermal boundary conditions, while the second was run with a constant wall temperature. The wall temperature was set to be 56oC less than the minimum gas temperature at the airfoil surface found in the adiabatic case. This ensures that the heat transfer will be from the fluid to the airfoil. In both cases, the y+ value of the first grid point off the wall was less than 1. These computations provide convective heat transfer coefficients (Figure 4) and driving temperatures for the airfoil, and inner and outer platforms. The approach used to determine the convective heat transfer coefficients has been extensively validated against experimental data at P&W-C. Thermal boundary conditions were calculated for both average-inlet and hot-streak conditions. The hot-streak inlet profile was obtained from a three- dimensional Navier-Stokes simulation of the combustor and scroll [2]. Figure 3. Relative Mach number distribution for the first stage Figure 4. Heat transfer coefficient determined by CFD CERAMIC VANE DESIGN AND ANALYSIS The primary source of stress in a turbine nozzle vane is thermal in nature. To minimize these stresses, ceramic vanes 6000 5000 4000 3000 2000 1000 0 100 Copyright © 2002 by ASMEPDF Image | ADVANCED MICROTURBINE SYSTEMS Final Report for Tasks 1 Through 4 and Task 6
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