Smooth outer coating for combustor components and coating method therefor

Abstract

A coating and method for overcoating a TBC on a component used in a high-temperature environment, such as the combustor section of an industrial gas turbine. The coating defines the outermost surface of the component and is formed of at least two layers having different compositions. An inner layer of the coating contains alumina in a first silica-containing matrix material that is free of zinc titanate. An outer layer of the coating contains alumina, a glass material, and zinc titanate in a second silica-containing matrix material. The outer layer of the coating has a surface roughness of not greater than three micrometers Ra and forms the outermost surface of the component. The coating reduces the component temperature by reducing the convective and radiant heat transfer thereto.

Description

[0001] This application is a Division of U.S. patent application Ser. No. 10 / 710,110, filed Jun. 18, 2004.BACKGROUND OF THE INVENTION [0002] The present invention generally relates to coatings for components exposed to high temperatures, such as the hostile thermal environment of a gas turbine. More particularly, this invention relates to a smooth outer coating for combustor components of a gas turbine component, in which the coating reduces the component temperature by reducing the convective and radiant heat transfer to the component in the combustor section of the turbine. [0003] Hot section components of aircraft and industrial (power generation) gas turbine engines are often protected by a thermal barrier coating (TBC), which reduces the temperature of the underlying component substrate and thereby prolongs the service life of the component. Ceramic materials and particularly yttria-stabilized zirconia (YSZ) are widely used as TBC materials because of their high temperature cap...

AI Technical Summary

Benefits of technology

[0007] The present invention generally provides a coating and method for overcoating a TBC on a component used in a high-temperature environment, such as the combustor section of a gas turbine. The invention is particularly directed to a coating that reduces the component temperature by reducing the convective and radiant heat transfer to the component in the combustor section of an industrial gas turbine.

[0010] As noted above, the coating of this invention reduces the component temperature by reducing the convective and radiant heat transfer to the component. In particular, the fine particle size distribution of the outer layer enables the outermost surface defined by the outer layer to be sufficiently smooth to significantly reduce convective heat transfer to the component, and the zinc titanate contained in the outer layer serves to reduce the IR transmissivity of the coating. The bimodal particle size distribution of the inner layer promotes the chemical inertness and stability of the inner layer. Furthermore, the inner layer is chemically compatible with the outer layer and the absence of zinc titanate in the inner layer promotes the adhesion of the outer layer to the component. In addition to the above benefits, the coating of this invention improves the spallation and erosion resistance of the TBC, and is therefore capable of significantly extending the life of the gas turbine component protected by the thermal barrier coating.

Problems solved by technology

Hot section components of aircraft and industrial (power generation) gas turbine engines are often protected by a thermal barrier coating (TBC), which reduces the temperature of the underlying component substrate and thereby prolongs the service life of the component.

As a result, the thermal protection provided by APSTBC is compromised in environments that have high thermal radiation loads, such as within the combustor section of a gas turbine.

The latter requirement is particularly demanding due to the different coefficients of thermal expansion (CTE) between ceramic materials and the substrates they protect, which are typically superalloys though ceramic matrix composite (CMC) materials are also used.

The service life of a TBC system is typically limited by a spallation event brought on by thermal fatigue.

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