A turbine for a power plant unit for a hybrid power plant is arranged in a turbine housing in which a flow channel for a compressible medium is arranged. A drive shaft and at least one output impeller are arranged in the flow channel, the output impeller containing an output shaft for operating a generator, wherein the output impeller is connected in a rotationally fixed manner to the output shaft. The drive shaft is not connected to the output shaft.

A seal includes a seal body disposed about a seal axis and configured to be mounted to a first component at a second axial end of the seal body. The seal body includes an interior surface defining a seal gland circumferentially extending about the seal axis. A packing is disposed within the seal gland. A retaining ring is in communication with a second radial side of the packing and is disposed within the seal gland. A plurality of exciter springs are mounted to a second radial side of the retaining ring. The plurality of exciter springs are biased against the interior surface of the seal body and are configured to center the packing within the seal body. The packing is configured to receive a second component and form a seal interface between the packing and the second component. The seal body is configured to form a portion of a passage.

The present disclosure provides a method of removing a bearing compartment from a gas turbine engine comprising inserting a tool between an aft portion of a low pressure turbine and a forward portion of a turbine exhaust case, the tool comprising one of a single wedge-shaped ring or a plurality of wedge-shaped segments, removing a plurality of fasteners coupling a flange of a bearing compartment housing to the turbine exhaust case, the bearing compartment housing containing the bearing compartment, and removing the bearing compartment housing from an aft portion of the gas turbine engine.

An apparatus and method for a component for a turbine engine, which generates a hot gas flow, and provides a cooling fluid flow, comprising a wall separating the hot gas flow from the cooling fluid flow and having a heated surface along which the hot gas flows and a cooled surface facing the cooling fluid flow and at least one cooling hole comprising a connecting passage extending between an inlet at the cooled surface and an outlet located at the heated surface, with the connecting passage comprising a diffusing section.

A vaned ring for a turbomachine stator includes an outer shell, a vane having a vane head, and two platforms arranged circumferentially on either side of the vane head, and fixed on a radially inner face of the outer shell by detachable fasteners. The vane head has first conical seatings bearing on respective second conical seatings of the two platforms so as to prevent a displacement of the vane radially inwards and along the circumferential direction. The vaned ring includes a frangible pin including a base mounted on the outer shell and a head projecting radially inwards with respect to the outer shell and having a head surface forming, vis--vis a top surface of the vane head, a stop opposing a displacement of the vane radially outwards.

A shaft assembly for a gas turbine engine is provided. The shaft assembly comprises: a first shaft having an outer surface; a first coupling ring disposed around the outer surface of the first shaft, an inner surface of the first coupling ring being coupled to the outer surface of the first shaft; a second shaft having an inner surface; and a second coupling ring disposed around the inner surface of the second shaft, an outer surface of the second coupling ring being coupled to the inner surface of the second shaft, wherein an outer surface of the first coupling ring is configured to mate with an inner surface of the second coupling ring, such that concentricity of the first and second shafts is maintained at the shaft assembly by virtue of the mating of the first and second coupling rings. Methods of assembling and re-assembling a shaft assembly are also provided.

A gas turbine engine with a lubrication system includes a ball bearing assembly and a rotor circumscribing a rotational axis and journaled within the ball bearing assembly. The gas turbine engine also includes a lubrication system located radially outward from a rotational axis and radially outward and adjacent to the ball bearing assembly, which includes a lubrication channel having an inlet and an outlet and a dispersion cone adjacent to the outlet of the lubrication channel.

A turbomachine assembly and, in particular, a low-pressure compressor of an aircraft turbojet engine includes an annular row of upstream vanes with trailing edges extending radially from an upstream support; an annular row of downstream vanes with leading edges axially facing the trailing edges and extending radially from a downstream support; an annular passageway delimited by the upstream support and the downstream support. The downstream support has a profile with: an upstream portion delimiting the annular passageway forming an annular slide, a downstream portion axially at the level of downstream vanes, and a connecting arc connecting the upstream portion to the downstream portion. The connecting arc is arranged downstream of the leading edges.

It is described a containment casing, in particular a containment casing of a gas turbine engine of an aircraft engine, which at least regionally has a structure selected from at least one of the structure types beam structure, cylinder structure, strips cage structure, foam structure, honeycomb structure, corrugated structure or net structure, each of which is formed from SMA. Furthermore, a gas turbine engine including an impactor containment casing is proposed.

A single-crystal turbine blade and a method of making such single-crystal turbine blade are disclosed. During manufacturing, a secondary crystallographic orientation of the material of the single-crystal turbine blade is controlled based on a parameter of a root fillet between an airfoil of the single-crystal turbine blade and a platform of the single-crystal turbine blade. The parameter can be a location of peak stress in the root fillet expected during use of the turbine blade.