| dc.description.abstract | The development of gas turbine engines for advanced subsonic and supersonic aircraft as well as for potential utilization of these high performance engines for stationary and marine applications requires, as a key element, the ability to operate at turbine inlet temperatures above the actual melting temperatures of the turbine materials. A limit on gas temperature levels is imposed by the fact that current alloys available for use in turbines, exhibit inadequate strength and oxidation characteristics above 1600–1800 deg F. However, the performance gains offered by operating engines at a high turbine inlet temperature may be realized through the application of an efficient method of cooling the highly stressed turbine components. As a step toward demonstrating that transpiration cooling of turbine blading is an effective means for achieving reliable and efficient gas turbine operation in a high gas temperature environment, a full-scale engine was tested at average gas temperatures of 2750–2800 deg F with a transpiration cooled turbine fabricated from normally used turbine alloys which are limited to metal temperatures of 1600–1800 deg F. The authors discuss the design of the transpiration air-cooled turbine, the technique used in fabricating the porous turbine blading, and the experimental test results obtained from operating the high-temperature engine. Furthermore, correlation of the test results on blade cooling with analytical predictions is presented. | |
