Abstract: A method of making a ceramic matrix composite blade includes laying up at least one ply of the plurality of fibrous ceramic plies in a preselected arrangement to form an airfoil and dovetail preform, laying up at least one additional ply of the plurality of fibrous ceramic plies on the airfoil and dovetail preform in a second preselected arrangement to form an integral platform as part of the non-rigidized blade preform, rigidizing the blade preform to form a rigidized blade preform, and densifying the rigidized blade preform to form a ceramic matrix composite blade having an integral platform structure. A tool for making the ceramic matrix composite blade and a ceramic matrix composite blade are also disclosed.

Abstract: A method of making a ceramic matrix composite blade includes laying up at least one ply of the plurality of fibrous ceramic plies in a preselected arrangement to form an airfoil and dovetail preform, laying up at least one additional ply of the plurality of fibrous ceramic plies on the airfoil and dovetail preform in a second preselected arrangement to form an integral platform as part of the non-rigidized blade preform, rigidizing the blade preform to form a rigidized blade preform, and densifying the rigidized blade preform to form a ceramic matrix composite blade having an integral platform structure. A tool for making the ceramic matrix composite blade and a ceramic matrix composite blade are also disclosed.

Abstract: A turbine airfoil includes at least one spar arrangement having a length less than an associated turbine airfoil length. During operation, the turbine airfoil has an outer skin surface which operates at a substantially higher temperature than that of an internal supporting parted spar arrangement. The parted spar arrangement permits the turbine airfoil outer skin surface to thermally expand between spar arrangements, thus preventing self-constraining thermal stresses from forming within the spar arrangement or the airfoil skin surfaces.

Abstract: A gas turbine engine hollow turbine airfoil, such as a first stage vane of a high pressure turbine of a gas turbine engine includes an outer wall surrounding a hollow interior. The outer wall extends radially outwardly in a spanwise direction from an airfoil base to an airfoil tip and has chordwise spaced apart leading and trailing edges and widthwise spaced apart pressure and suction side portions extending chordwise between the leading edge and a trailing edge block which terminates at the trailing edge. A plurality of trailing edge cooling air ducts extend chordwise through the trailing edge block aftwardly from the hollow interior and a plurality of trailing edge film cooling holes extend from the ducts through the trailing edge block. The cooling air ducts are preferably centered near a neutral axis of chordwise and spanwise bending of the trailing edge block.

Abstract: A method and apparatus for film cooling of an aerodynamically shaped airfoil uses a plurality of curved slots extending through the airfoil in an area upstream of the high curvature region of the airfoil, i.e., in an area of low Mach number of the gas stream passing over the airfoil surface. The curved slots are configured to inject cooling air at an angle of about 16.5 degrees. The cooling air is injected at a blowing ratio in excess of 1.0 and yet is effective to form a film on the vane surface.