An example method of assembling a turbomachine airfoil array includes, among other things, securing a partial airfoil array within a fixture, the partial airfoil array having at least one existing airfoil extending radially between an inner and an outer fairing and an open area where at least one existing airfoil has been removed. The method includes mounting a positioning saddle relative to a base of the fixture, the positioning saddle aligned with the open area, holding a replacement airfoil using the positioning saddle, applying a curable material at an interface between the replacement airfoil and the inner and outer fairing, and curing the curable material while maintaining a relative position between the replacement airfoil and the inner and outer fairing.

A vane assembly (10) having: an airfoil (12) and a shroud (14) held together without metallurgical bonding there between; a channel (22) disposed circumferentially about the airfoil (12), between the airfoil (12) and the shroud (14); and a seal (20) disposed in the channel (22), wherein during operation of a turbine engine having the vane assembly (10) the seal (20) has a sufficient ductility such that a force generated on the seal (20) resulting from relative movement of the airfoil (12) and the shroud (14) is sufficient to plastically deform the seal (20).

Embodiments of an axially-split radial turbine, as are embodiments of a method for manufacturing an axially-split radial turbine. In one embodiment, the method includes the steps of joining a forward bladed ring to a forward disk to produce a forward turbine rotor, fabricating an aft turbine rotor, and disposing the forward turbine rotor and the aft turbine rotor in an axially-abutting, rotationally-fixed relationship to produce the axially-split radial turbine.

Dual alloy turbine rotors and methods for manufacturing the same are provided. The dual alloy turbine rotor comprises an assembled blade ring and a hub bonded to the assembled blade ring. The assembled blade ring comprises a first alloy selected from the group consisting of a single crystal alloy, a directionally solidified alloy, or an equi-axed alloy. The hub comprises a second alloy. The method comprises positioning a hub within a blade ring to define an interface between the hub and the blade ring. The interface is a non-contacting interface or a contacting interface. The interface is enclosed by a pair of diaphragms. The interface is vacuum sealed. The blade ring is bonded to the hub after the vacuum sealing step.

A weldment member for a gas turbine engine including a forward welding member and an aft welding member. The forward welding member has an annular shape with a forward welding face formed at one end. The forward welding face has a forward radiographic marking hole formed therein. The aft welding member has an annular shape with an aft welding face formed at one end. The aft welding face has an aft radiographic marking hole formed therein. The forward welding face is aligned with the aft welding face and the forward radiographic marking hole is angularly offset from the aft radiographic marking hole.

A method of joining a first work piece and a second workpiece. The first and second workpieces may be rotor wheels of a rotor for a turbomachine. At least one of the workpieces includes an oxide dispersion strengthened alloy material and the first and second work pieces may be joined by welding a cladding on at least one of the workpieces to the other of the workpieces, without welding a substrate of the at least one workpiece which includes an oxide dispersion strengthened alloy material.

The invention relates to a flow guiding system for a turbo engine, in particular an aircraft engine, characterized by an air intake device taking a thermal management air stream from a free flowing air stream during the operation of the turbo engine and at least one flow guiding means for guiding the thermal management air stream to at least one target region to be thermally managed in the turbo engine. The invention also relates to a turbo engine with a flow guiding system and a method for manufacturing a flow guiding system.

An exhaust centerbody for a turbine engine is provided. The centerbody includes a truncated downstream part, which is connected to an upstream part by an annular ridge marking a discontinuity between the outer surfaces of the upstream and downstream parts. The outer surface of the downstream part has a substantially conical general shape, of which the tip is oriented downstream and is positioned in the region of the axis A, the axial half-section of this outer surface defining a line of which the upstream end part is substantially tangential to a straight line passing through the ridge and forming a non-zero angle α with a tangent to the outer surface of the upstream part, in the region of the ridge, and of which the downstream end part is substantially tangential to a straight line passing through the tip and forming a non-zero angle β with the axis A.

An epicyclic gear train component includes spaced apart walls with circumferentially spaced mounts that interconnect the walls. The mounts provide circumferentially spaced apart apertures between the mounts at an outer circumference of the walls. Baffles are arranged between the walls near the mounts. Gear pockets are provided between the baffles and the baffles include a lubrication passage that terminates at least one of the gear pockets. One of the walls includes a hole which has a tube that extends through the hole and is received in an opening in the baffle. The tube is in communication with the lubrication passage.

An assembly adapted for use in a gas turbine engine includes a carrier and a blade track segment. The carrier extends at least partway about an axis. The blade track segment is supported by the carrier radially relative to the axis to define a portion of a gas path of the assembly.