A process for coating a gas turbine engine disk comprises placing the disk having an outer surface into a chamber, the chamber configured to perform atomic layer deposition; injecting a first reactant into the chamber; forming a first monolayer gas thin film on the outer surface; removing the first reactant from the chamber; injecting a second reactant into the chamber; reacting second reactant with the first monolayer gas thin film; removing the second reactant from the chamber; and forming a protective barrier coating on the outer surface.

A process for coating a gas turbine engine disk comprises placing the disk having an outer surface into a chamber, the chamber configured to perform atomic layer deposition; injecting a first reactant into the chamber; forming a first monolayer gas thin film on the outer surface; removing the first reactant from the chamber; injecting a second reactant into the chamber; reacting second reactant with the first monolayer gas thin film; removing the second reactant from the chamber; and forming a protective barrier coating on the outer surface.

A method of manufacturing an impeller includes attaching blades to a hub. The impeller includes the blades, the hub, and a shroud. The blades each include a tip, and the shroud includes an inner surface and at least one ring extending from the inner surface. The method also includes applying a brazing compound to the tips of the blades and to the inner surface of the shroud. The method includes inserting the blades and the hub into the shroud such that the tips of the blades press against the rings. The rings and the tips of the blades form an interference fit between the rings and the tips of the blades that maintains a consistent gap between the shroud and the blades during manufacture of the impeller. A compressor including the impeller is also disclosed.

An integrally bladed rotor for a turbomachine, in particular a compressor or turbine stage of a gas turbine, to which at least one separately formed impulse element housing (40; 40′) is fastened by at least one fastening element (30; 30′) which engages for this purpose into an opening (41) of the impulse element housing and into an opening (11) of the rotor, the impulse element housing having at least one cavity (44) in which at least one impulse element (5) is accommodated with play.

A vibration mitigation coating for an integrally bladed rotor including a disk including an interior radius proximate an axis and an exterior radius distal from the axis, the disk including a substrate with an external surface, said external surface extends from the interior radius to the exterior radius; and a damping material disposed directly onto the external surface of the substrate.

A turbine wheel consists of a first shroud component and a second bladed disc component. The shroud component comprises a shroud structure, a hub structure and a spoke formed integrally therewith and extending between the shroud structure and the hub structure. The bladed disc component comprises a hub member having inner and outer rims, turbine blades disposed on the outer rim, and at least one receiving zone for receiving the spoke, said at least one receiving zone extending radially between the inner and outer rims. The shroud component and the bladed disc component are connected and thus provide the turbine wheel with a shrouded portion. A shrouded turbine wheel can therefore be conveniently assembled starting from at least two components. Further, these components have simplified geometries for easy manufacture, for example using a casting technique, while the overall mechanical performance of the turbine is preserved or improved.

A turbine wheel consists of a first shroud component and a second bladed disc component. The shroud component comprises a shroud structure, a hub structure and a spoke formed integrally therewith and extending between the shroud structure and the hub structure. The bladed disc component comprises a hub member having inner and outer rims, turbine blades disposed on the outer rim, and at least one receiving zone for receiving the spoke, said at least one receiving zone extending radially between the inner and outer rims. The shroud component and the bladed disc component are connected and thus provide the turbine wheel with a shrouded portion. A shrouded turbine wheel can therefore be conveniently assembled starting from at least two components. Further, these components have simplified geometries for easy manufacture, for example using a casting technique, while the overall mechanical performance of the turbine is preserved or improved.

A method of joining a first component to a second component at respective connection surfaces, comprising, in order, applying a local surface treatment to the connection surface of at least one of the first and second components in order to locally alter the microstructure to a depth of between 60 μm and 10 mm below the connection surface; and joining the first component to the second component using a welding process.

Described is a rotor disk (40) for a compressor (29, 32) of a gas turbine, in particular an aircraft gas turbine (10), the rotor disk having a main body (42), at least one rotor arm (44) projecting from the main body (42) in the axial direction (AR), the rotor arm (44) having, in a sectional view taken in a sectional plane defined by the axial direction (AR) and the radial direction (RR) a beginning portion (44a) merging into the main body (42); an end (44e) portion remote from the main body (42) and forming a kind of free end in the axial direction (AR), the beginning portion (44a) and the end portion (44e) being interconnected by an intermediate portion (44z), characterized in that the intermediate portion (44z) is curved with at least one radius of curvature (Ri, Ra).

A tandem rotor disk apparatus may include a rotor disk body concentric about an axis. The tandem rotor disk apparatus may also include a first blade extending radially outward of the rotor disk body and a second blade extending radially outward of the rotor disk body. The first blade may be offset from the second blade in a direction parallel to the axis. The tandem rotor disk apparatus may be implemented in a gas turbine engine with no intervening stator vane stages disposed between the first blade and the second blade.