Enhanced High Pressure Turbine Blade Cooling Architecture

ABSTRACT: Next generation military unmanned aircraft have a need for lower thrust specific fuel consumption (TSFC) and higher power-to-weight ratio, exceeding the capability of current engines in this thrust class. Attainment of these goals can be aided through an increase in turbine inlet temperature. In a modern gas turbine engine, turbine inlet temperatures can exceed the blade and disk material limits by 600°F or more. Micro Cooling Concepts has developed a turbine blade cooling concept that provides enhanced internal impingement cooling effectiveness. The concept uses an impingement scheme with a micro-structured face on the interior wall in the turbine blade stagnation region. Air flow exits the nozzle, impinges in the grooved region, and then exits via the microchannel passages. The flow through the microchannel passages cools the sidewalls as it travels back along the blade contour, exiting out the trailing edge. A laminated foil construction approach will used to form the blade in which thin metal foils are etched with micro passages, then stacked and diffusion bonded. This design approach has several advantages including improved internal heat transfer, low pressure losses through internal heat transfer sections, consistent with thin trailing edge designs, and high strength, lightweight blade designs. BENEFIT: Anticipated benefits of the enhanced internal cooling design include reduction in thrust specific fuel consumption, increased power-to-weight, reduced cost for turbine blades due to the ability to use lower cost superalloys, and reduced specific pollutant formation. The commercial applications are varied including most axial flow aero, marine, and power generation gas turbines.