Fiber-Reinforced Metal Matrix Composites for High-Pressure Turbines

For a selected vehicle and jet engine configuration, fuel consumption is to a significant degree dictated by the thermal efficiency of the engine. Achieving optimal thermal efficiency and minimizing fuel consumption requires high temperature operation with minimum cooling of hot gas path components, including combustors, high-pressure turbine blades, and inlet nozzle (stator) vanes. In parallel, cost reduction and reliability enhancements can be achieved by minimizing system complexity, in part by eliminating or minimizing cooling and by eliminating the need for thermal barrier coatings. Currently, cooled blades and vanes are constrained to a peak operating temperature of ~3000 degrees F. Successful implementation of cost-effective and reliable materials and processes with operational capability to 3500 degrees F would result in dramatic gains in jet engine efficiency. Gains to date have in large part been achieved by evolutionary improvements in turbine blade alloys, thermal barriers, and cooling gas path designs. Anticipated gains expected from the use of ceramics and ceramic matrix composites (CMC) have been slow to accrue. This project will pursue revolutionary improvements in materials and processing capabilities via the application of innovative fiber interfaces and melt infiltration processing and proven oxidation protection coatings to produce a durable and cost-effective carbon fiber-reinforced metal matrix composite (Cf/MMC) with mechanical properties and environmental resistance suited to cyclic and long-duration operation at turbine inlet temperatures to 3500 degrees F within the jet engine environment. In Phase I, conceptual design will be performed of representative composite hardware; preliminary goals will be identified for thermal, chemical, and mechanical properties of these key turbine subcomponents; demonstrator subelements will be designed; test articles will be fabricated and tested; and demonstrator subelements will be fabricated as a preliminary demonstration of feasibility. In the Phase I option, the demonstrator subelements will be tested in a laboratory environment. A cost-effective and representative engine demonstration will be performed in Phase II to provide an early path for commercialization leading to opportunities for future implementation of the technology. Project emphasis will be on demonstration of turbine components, but the developed technology would also be applicable to applications for other hot section components including combustors.