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Enhanced Multi-wall High Pressure Turbine Blade Architecture


OBJECTIVE: To enhance the design of small turbo fan engine high-pressure turbine (HPT) blades for improved turbine life and validate design improvements through design analysis and testing. DESCRIPTION: Next generation military unmanned aircraft whose propulsion systems are in the 15,000 lb thrust class, have a need for lower thrust specific fuel consumption (TSFC) and higher power-to-weight ratio that exceeds the capability of the state of the art in this thrust class. Improvements to engines in this thrust class support the warfighter capabilities and help reach the Versatile Affordable Advanced Turbine Engine (VAATE) III Program goals of reduction in TSFC by 35 percent, and increased power-to-weight ratio by 80 percent, and reduced cost by 55 percent of the entire propulsion system. There is a significant body of knowledge on enhancing HPT technology to improve TSFC and power-to-weight ratio for large and very small turbofan engines. The challenge in enhancing performance is to increase HPT inlet temperature and reduce tip clearances. Both of these increase power density, and reduce SFC. Increased HPT inlet temperature, though, creates challenges for HPT durability. Scaling is the key issue for this class of turbines. Heat transfer in the HPT is a function of Reynolds number and momentum ratio. As the Reynolds number decreases, so does the ability to effectively cool the HPT. It is difficult to simply geometrically scale turbine cooling strategies in the same fashion that the engine is scaled for thrust and fuel consumption because as cooling hole size decreases discharge coefficient (Cd) decreases nonlinearly. Decreased Cd requires increased pressure drop to cool the HPT, affecting efficiency. Further, the primary metric for cooling is blowing ratio. Direct scaling holes may increase film penetration, leaving the metal surface vulnerable to hot gases from the combustor. Skin thickness of the blades also decreases making it difficult to achieve film coverage from angled holes due to decreasing aspect ratio of the holes. Finally as the diameter of the HPT decreases tip clearances do not linearly scale, thus more care is required when designing the tip of the blade to ensure that the performance and health of the tip region is good. New technologies are desired that provide enhancements in HPT inlet temperature and tip clearance capability while also maintaining/enhancing durability. Close collaboration with an original equipment manufacturer (OEM) of small gas turbine engines in the 15,000 pound thrust class is highly desirable to increase the probability of transition of the developed cooling and clearance technologies. PHASE I: Demonstrate the feasibility of the innovative technology to improve HPT durability and SFC. Identify a baseline and develop concepts for improved durability of HPT blades with enhanced performance. Conduct CFD, heat transfer analyses and thermo-mechanical-fatigue analysis of candidate technologies and compare their performance to the baseline system. PHASE II: Develop, fabricate, and test concepts analyzed in the Phase I effort. Assess the performance of the proposed HPT technologies and their impact on SFC and power to weight ratio. Assess the durability of the proposed technologies as compared to the baseline. PHASE III: Transition developed technologies to commercial engine manufacturers and consider opportunities within the VAATE Phase II Program. Apply improved HPT technologies to commercial core technologies for regional business jet class of engines. REFERENCES: 1. Vitt, P., Iverson, C., Malak, F.M., and Liu, J.S., 2010,"Impact of Flowfield Unsteadiness on Film Cooling of a High Pressure Turbine Blade,"ASME Turbo Expo, Paper No. GT2010-22773. 2. Martin, E., Wright, L., and Crites, D., 2012,"Impingement Heat Transfer on a Cylindrical, Concave Surface with Varying Jet Geometries,"ASME Turbo Expo, Paper No. GT2012-68818. 3. Ho, K.S., Urwiller, C., Konan, S.M., Liu, J.S., and Aguilar, B., 2012"Conjugate Heat Transfer Analysis for Gas Turbine Cooled Stator,"ASME Turbo Expo, Paper No. GT2012-68196.
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