Abstract: An airfoil (10) as may be used in a gas turbine engine includes a ceramic matrix composite (CMC) element (12) that extends to define a leading edge portion (14) and chord portion (16) of the airfoil, and a separately formed but conjoined trailing edge element (18) that defines a desirably thin trailing edge of the airfoil without the need for using an excessively small bend radius for reinforcing fibers in the CMC element. The trailing edge element may include a plurality of interlock elements (26) that extend through the trailing edge attachment wall (22) of the CMC element and provide mechanical attachment there between. Alternatively, the trailing edge element may be adhesively bonded or sinter bonded to the CMC element. A cooling air insert (60) may be disposed within a cooling air cavity (64) of the CMC element and may include cooling tubes (66) that extend into the trailing edge element to deliver cooling air there through.

Abstract: An airfoil (10) as may be used in a gas turbine engine includes a ceramic matrix composite (CMC) element (12) that extends to define a leading edge portion (14) and chord portion (16) of the airfoil, and a separately formed but conjoined trailing edge element (18) that defines a desirably thin trailing edge of the airfoil without the need for using an excessively small bend radius for reinforcing fibers in the CMC element. The trailing edge element may include a plurality of interlock elements (26) that extend through the trailing edge attachment wall (22) of the CMC element and provide mechanical attachment there between. Alternatively, the trailing edge element may be adhesively bonded or sinter bonded to the CMC element. A cooling air insert (60) may be disposed within a cooling air cavity (64) of the CMC element and may include cooling tubes (66) that extend into the trailing edge element to deliver cooling air there through.

Abstract: A turbine vane array (10) for a combustion assembly (8) in a combustion turbine engine. The vane array (10) includes a plurality of stationary vane assemblies (12), each vane assembly (12) including at least one airfoil (24) and inner and outer shroud segments (26, 28) attached to opposing ends of the airfoil (24). The inner and outer shroud segments (26, 28) each include an inner face (34, 36) facing toward a gas path (13) extending through the vane array (10). A removable cover structure may be provided on the inner faces (34, 36) of the inner and outer shroud segments (26, 28). The cover structure may include removably attached insert elements (72, 72?) that are positioned on the inner faces (34, 36) extending between upstream edges (44) and downstream edges (46) of the shroud segments (26, 28) and extending between adjacent airfoils (24).

Abstract: An airfoil (26) for a gas turbine engine (10) having a source of heat for controlling a temperature gradient across the airfoil components. In one embodiment the airfoil includes a CMC outer body (28) defining an airfoil shape and a ceramic inner body core member (36) housed within and bonded to the outer body, and a heating element (54) disposed within the inner body core member. In another embodiment the source of heat may include a conduit (55) for delivering a flow of hot combustion gas from the combustor (14) to an interior of the airfoil. Heat energy may be delivered to the airfoil interior prior to or during startup of the engine in order to reduce the effect of temperature transients, during ongoing operation of the engine to reduce steady state temperature gradients, and/or during shutdown conditions to mitigate differential shrinkage between the core member and the outer body of the airfoil.

Abstract: A method of coating an edge surface (30) of an anisotropic ceramic matrix composite material (10) for use in a high temperature environment is disclosed where the edge surface (30) has exposed reinforced fiber layers (20). A laser beam may be used to melt a portion of the ceramic matrix composite material (10) on the edge surface (30) forming a melt layer. The melt layer is retained proximate the edge surface and the laser beam is controlled to form an isotropic protective coating (32, 34) on a portion of the edge surface (30). A method may be used to form a component for use in a high temperature environment that includes directing a laser beam toward a ceramic matrix composite material (10), controlling the laser beam to melt a portion of the ceramic matrix composite material (10) and forming a homogeneous protective coating (32, 34) from a melt layer that exerts compression on at least a portion of the ceramic matrix composite material (10) when the melt layer is cooled.