This invention relates to abradable clearance control coatings and, more particularly, to such coatings as are capable of operating up to about 1200.degree. F (650.degree. C).
The efficiency of an axial-flow compressor in a gas turbine engine is at least partially dependent on inhibition of interstage leakage. If a relatively wide clearance exists between a compressor casing and a compressor rotor stage, fluid such as air being compressed can leak from a higher pressure portion to a lower pressure portion of the compressor. Therefore, evolving with gas turbine engine development has been the development of clearance control coatings to minimize interstage leakage.
Solution of such a clearance control problem is made more complex by the fact that during operation, the compressor casing and the rotating compressor blades expand or contract at different rates in an engine cycle. Therefore, one solution has been to allow the blades, rotating, stationary or both, to cut a path into a material mounted in juxtaposition with the blade tips in each stage. Materials so applied have included honeycomb, wire mesh, foamed metals and other porous structures. One such material in wide use in aircraft-type gas turbine engines is described in U.S. Pat. No. 3,342,563 - Butts, issued Sept. 19, 1967, the disclosure of which is incorporated herein by reference.
Many of the existing clearance control systems perform adequately for the type of engine in which they are applied. However, more advanced engines require improved materials having higher temperature capabilities as well as the capability of reducing wear on the tips of blades which rub an abradable coating or material. A more particular problem exists with the wear of blades made from titanium-base alloys.