Abstract: The method is used for making an annular gas turbine engine case from a preform. The method comprises comprising flowforming at least one section of the preform, and then outwardly bending at least one portion of the perform.

Abstract: A gas turbine engine having an electric generator includes a transmission shaft extending along a longitudinal axis of the engine and drivingly interconnecting a turbine shaft of the engine and a rotor shaft of the electric generator. The transmission shaft is engaged by splined mating connections with the turbine shaft and the rotor shaft. The transmission shaft has a shear neck defining a reduced radial wall thickness with respect to a remainder of the transmission shaft such as to provide a weakened region of the transmission shaft. An annular support structure, concentric with and surrounding the transmission shaft, is axially located between the shear neck and a forward end of the transmission shaft engaged to the turbine shaft, and includes a bearing operable to rotationally support the transmission shaft.

Abstract: An airfoil blade of a gas turbine engine includes a root configured for mating attachment with a cooperating rotor hub and an airfoil extending away from the root. The root is composed of a first metal and has a nanocrystalline metal coating, composed of a second metal, over at least a portion thereof. A method of protecting such a blade by applying a nanocrystalline metal coating to a portion of the blade root is also disclosed.

Abstract: An engine casing for a gas turbine engine, such as, but not limited to, a gas turbine engine fan case, is disclosed which includes an annular case shell formed of a substrate material that is at least partially coated by a nanocrystalline metal coating. A method of manufacturing such an engine casing is also provided. The present engine casing provides improved containment capability in the event of a blade release or other failure during operation of the engine.

Abstract: A hybrid vane airfoil for a gas turbine engine, such as a vane of a compressor stator, is disclosed which includes a non-metallic core, and an outer metallic shell at least partially covering the non-metallic core and which defines an outer surface of the airfoil. The non-metallic core is composed for example of a polymer, and the metallic outer shell is composed of a nanocrystalline metallic coating.

Abstract: A method of applying a nanocrystalline coating to a gas turbine engine component is described. The method comprises the steps of applying an intermediate bond coat to at least a portion of the component, and then applying the nanocrystalline coating to at least the portion of the component overtop of the intermediate bond coat. The component may include, for example, a blade of which a dovetail portion of the blade root is protected by applying the intermediate bond coat and the nanocrystalline coating thereto.

Abstract: A fan blade of a gas turbine engine, and a method of producing same, is disclosed which is composed of a core substrate and a nanocrystalline metal coating provided on at least a portion of the airfoil of the fan blade. The method includes the steps of providing a fan blade core having an airfoil and then applying a nanocrystalline metal coating over at least a portion of the fan blade core.

Abstract: The blade dovetails and associated dovetail slots of a gas turbine engine rotor assembly are shaped so that during windmilling the high stress regions of the blade dovetails are shielded from rubbing against the disk and vice versa. Bumper surfaces are provided in low stress regions of the dovetail assembly such that the windmilling contact points between the blade dovetails and the disk are in non-critical areas of the dovetail assembly.

Abstract: A method for repairing damaged fins in a labyrinth seal without the use of welding or metallurgical bonding by removing a section of the shaft containing the damaged fins, and providing a sleeve with replacement fins.

Abstract: A compressor vane having a root, a tip and an airfoil portion extending between the root and the tip is disclosed. A grommet is used in conjunction with the vane and has a lip sealingly surrounding a portion of the vane root and a circumferential portion protruding around a perimeter of the lip in engagement with a surface of a platform of the vane root. The circumferential portion has a profile similar to that of the airfoil portion.

Abstract: A vane for a vane assembly of a gas turbine engine includes a vane root connected to the airfoil portion opposite a tip thereof. The vane root includes a platform and a button portion interconnecting the platform and the airfoil position. The button portion has relatively blunt leading and trailing ends which respectively protrude beyond leading and trailing edges of the airfoil portion.

Abstract: A vane assembly for a gas turbine engine including a shroud with openings defined about a circumference of the shroud and vanes extending radially from the shroud with a vane tip in each opening. Isolating apparatus are disposed in each opening to isolate the vane tip from the shroud ring. Each vane tip is engaged with the shroud by retaining apparatus for restricting movement of the vane tip in an axial direction relative to the shroud.

Abstract: A vane including a root, a tip and an airfoil portion extending between the root and the tip, the vane also including a button portion interconnecting the airfoil portion and the root. The button portion protrudes beyond a profile of the airfoil a distance less than the root. The button portion includes leading and trailing ends free of sharp edges. The leading and trailing ends respectively protrude beyond leading and trailing edges of the airfoil portion.

Abstract: A vane assembly for a gas turbine engine including a shroud ring with openings and isolating arrangement isolating a vane extremity within each opening, the isolating arrangement cooperating to form a continuous gas path surface completely covering at least an axial portion of a surface of the shroud ring.

Abstract: The method is used for making an annular gas turbine engine case from a preform. The method comprises comprising flowforming at least one section of the preform, and then outwardly bending at least one portion of the perform.

Abstract: The method is used for making an annular gas turbine engine case, the method comprises flowforming a first area of the preform to provide a first annular case portion having a first thickness and a second area having a second thickness, the first and the second average thickness being different.

Abstract: A method and apparatus for repairing damaged fins in a labyrinth seal, the seal without the use of welding or metallurgical bonding. The method comprises the steps of removing a section of the shaft containing the damaged fins, and providing a sleeve with replacement fins.

Abstract: A fan case for a gas turbine engine is provided to prevent water from pooling and freezing inside the fan case when the gas turbine is not in operation, and potentially prevent the engine from starting and/or cause engine damage. The fan case is therefore adapted to provide a drainage system allowing water to freely escape the fan case without compromising the structural integrity of the fan case. The fan case comprises: a hollow body; an abradable liner disposed inside the hollow body; a flange on the hollow body defining a surface for mating to a forward part of the engine; and, a first radial slot extending across the mating surface of the hollow body. The abradable liner has an inside surface for circumscribing the fan blades wherein the first radial slot is in a free-draining relation with the inside surface of the abradable liner.

Abstract: A fan case and method of manufacturing a fan case, for a gas turbine engine, where the fan case has a hollow tubular metal shell with: a central axis of symmetry; an inlet; an outlet; and peripheral wall about the axis to encompass the tips of a plurality of rotary fan blades, and the shell wall includes an upstream portion defining an annular abradable material recess that extends axially having an upstream end at the shell inlet and a downstream end located upstream from the shell outlet.

Abstract: An apparatus for balancing a shaft of an aircraft engine preferably includes a nose cone and the use of standard fasteners. The nose cone is adapted to cooperate with at least one standard fastener to retain the standard fastener to the nose cone to rotationally balance the shaft.