Description: Gas turbine engine technology is constantly challenged to operate at higher combustor outlet temperatures. In a modern gas turbine engine, these temperatures can exceed the blade and disk material limits by 600 ¿F or more, necessitating both internal and film cooling schemes in addition to the use of thermal barrier coatings. Internal convective cooling is inadequate in many blade locations, and both internal and film cooling approaches can lead to significant performance penalties in the engine. Micro Cooling Concepts has developed a turbine blade cooling concept that provides enhanced internal impingement cooling effectiveness via the use of micro-structured impingement surfaces. These surfaces significantly increase the cooling capability of the impinging flow, as compared to a conventional untextured surface. This approach can be combined with microchannel cooling and external film cooling to tailor the cooling capability per the external heating profile. The cooling system can then be optimized to minimize impact on engine performance.

Benefits: The proposed Phase I effort directly supports core NASA research efforts in turbine engine development, as well as the multi-agency Verstaile Affordable Advanced Turbine Engine (VAATE) initiative. It is also applicable to two-state to orbit designs.

Military and commercial aircraft can both benefit from this technology, which permits higher combustion temperatures with lower engine penalties than state-of-the-art turbine blade cooling technologies. The blade cooling concept could be also applied to power generation plants, which are also seeking means of operating at higher temperatures and efficiencies. More generally, the enhanced impingement cooling techniques proposed here could be applied to a variety of cooling problems in the electronics, industrial processes, automotive, and laser industries.