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US 2006/0140771 A1Published application
Ceramic composite with integrated compliance/wear layer
AssigneeGeneral Electric Company
InventorsDouglas Carper, Suresh Subramanian
Life of this patent
Estimating the term…
Abstract
The integral layer provides a ductile interface for attachment locations of a turbine engine component where a metallic surface is adjacent the attachment location. The ductile layer provides a favorable load distribution through the composite at the attachment location, and eliminates the need for a metallic shim.
Claims23
A component for use in a heated environment comprising:
- a body composed of a ceramic matrix material, the body including an attachment location, the attachment location defining at least a portion of a surface of the component, and the attachment location being configured to be adjacent a second component during exposure in a heated environment;
- and
- an interface layer defining at least a portion of the surface of the attachment location and having a matrix side and an outer side, wherein the interface layer is interposed between the body and the second component during exposure in the heated environment, the matrix side of the interface layer being metallurgically bonded to the body.
A method of manufacturing a turbine engine component comprising the steps of:
- providing a plurality of ceramic fibers;
- coating a preselected portion of the fibers with a layer of boron nitride to form a plurality of coated fibers, wherein the coating includes using chemical vapor infiltration;
- laying up the plurality of fibers in a preselected arrangement to form a component preform;
- partially infiltrating the component preform using a carbon-containing slurry;
- further infiltrating the component preform with at least silicon to form a ceramic matrix composite aircraft engine component;
- and
- bonding an interface layer to an attachment location defined by a surface of the ceramic matrix composite aircraft engine component, wherein the interface layer contains at least silicon.
4 more elements
A method of manufacturing a ceramic matrix composite aircraft engine component, the component the method comprising the steps of:
- providing a plurality of unidirectional prepreg ceramic fiber plies, the plies comprising coated prepreg ceramic fiber tows;
- laying up the plurality of prepreg ceramic fiber plies in a preselected arrangement to form a component shape;
- overlying at least a portion of the outside surface of the component shape with at least one matrix ply;
- heating the component shape to form a preform;
- infiltrating the preform with at least silicon to form a ceramic matrix composite component, wherein the component includes an attachment area defining a surface of the component;
- and
- forming an interface layer on the attachment area, wherein the interface layer includes at least silicon, and the interface layer and the attachment area are integrated.
4 more elements
Dependencies are inferred from the claim text.