A gas turbine engine includes a fan, a compressor section, a combustor, and a turbine section where the turbine section is downstream of the combustor section. A shaft connects the turbine section to the compressor section. A bore tube is disposed within the shaft downstream of the compressor section. The bore tube includes an inlet connected to an air source for passing cooling air in an upstream direction of the shaft.

There is provided an axial flow turbine capable of realizing a reduction in gland leakage amount. The axial flow turbine in an embodiment is of a single flow type and includes an upstream-side gland part located on an upstream side of a working medium in an axial direction of a turbine rotor and a downstream-side gland part located on a downstream side of the working medium in the axial direction of the turbine rotor. The axial flow turbine is configured such that a cooling medium lower in temperature and higher in pressure than the working medium is extracted in a middle of flowing from the inside to the outside of the turbine casing in the upstream-side gland part, and the extracted cooling medium is introduced into the stationary blade.

Oil transfer tube for a system to control the regulation of a turbine engine propeller pitch, in particular of an aircraft, the tube being intended to be mounted coaxially inside a tubular shaft of the propeller, the tube having an elongated shape and including a first male longitudinal end part intended to be inserted in a female housing of a stator casing equipped with an oil supply circuit of the tube, and a second male longitudinal end part around which is intended to be mounted a hydrodynamic bearing to guide the tube in the shaft, wherein the first end part includes a free annular end with a convex rounded cross-section intended to bear axially against an annular bottom of said housing, and in that said second end part comprises an outer axial annular bearing surface of an inner ring of the bearing, the annular bearing surface presenting in cross-section a convex curved shape.

A turbocharger includes: a bearing; a bearing wall portion having a bearing hole in which the bearing is arranged; a separation wall portion, which is provided on a radially outer side of the bearing hole with respect to the bearing wall portion, and forms an internal space with the bearing wall portion; an oil discharge port, which is formed in the separation wall portion, and communicates with the internal space; and a guide portion, which is provided to the bearing wall portion facing the internal space, and separates away from the oil discharge port in a direction of a plane perpendicular to a center axis of the bearing as approaching the oil discharge port in a direction of the center axis.

A rotor assembly includes a plurality of rotor disks that each have a rim portion. The plurality of rotor disks includes a first rotor disk and at least one first bleed air passage that extends through a forward rim portion of the first rotor disk. At least one second bleed air passage that extends through an aft rim portion of the first rotor disk.

A gas turbine engine includes a fan, a compressor section, a combustor, and a turbine section where the turbine section is downstream of the combustor section. A shaft connects the turbine section to the compressor section. A bore tube is disposed within the shaft downstream of the compressor section. The bore tube includes an inlet connected to an air source for passing cooling air in an upstream direction of the shaft.

A rotor drum for an aircraft turbomachine includes an annular wall extending around a longitudinal axis (A), the annular wall carrying rotor blades and having at least one bleed device configured to allow at least one liquid to pass through the annular wall. The bleed device includes a series of three adjacent circular orifices, the three orifices being aligned along a line and having a central orifice of larger diameter D1 and two lateral orifices of smaller diameter D2 diametrically opposed with respect to the central orifice.

A gas turbine engine for an aircraft such as a UAV includes a compressor connected to a turbine with a combustor to produce a hot gas stream. The rotor is supported by two radial foil bearings. An axial thrust bearing assembly is positioned between the compressor disk and the turbine disk and includes an axial thrust bearing radial disk extending from a hollow axial tube. Compressed air is bled off from the compressor and passed into an axial thrust balance chamber to provide the axial thrust balance for the rotor. The compressed air from the thrust bearing chamber then flows through both of the radial foil bearings for cooling, is collected in and around the hollow tube, and then discharged into the inlet of the turbine. An orifice can be adjusted to meter and control a pressure occurring in the thrust balance chamber.

A steam turbine system includes: a steam turbine having a casing into which steam is fed from an outside of the casing and a rotating shaft that is provided within the casing so as to be rotatable around a central axis; a deaerator that is connected to the steam turbine and that deaerates leaked steam that has leaked from a gap between the casing and the rotating shaft to the outside of the casing; a vacuum pump that is connected to the deaerator and that lowers a pressure within the deaerator; and a shaft sealing device that is connected to the deaerator and that seals the gap between the casing and the rotating shaft. The shaft sealing device has a seal member having a seal body, a housing, a biasing member, and a negative-pressure introduction part.

Provided is a center vent tube support device that can prevent transmission of abnormal load between members and prevent abrasion of a contact surface. A device includes: an annular sleeve having an inner surface that comes into contact with an outer surface of a center vent tube; a ring formed of a pair of segments and placed between the sleeve and a main shaft; and an annular nut for fixing the ring to the sleeve. An outer surface of the sleeve or the nut includes a pressurizing surface formed as a conical surface. Each of the ring segments includes a supporting surface formed as a cylindrical surface having a diameter equal to an inner diameter of the main shaft in a portion where the device is to be installed, a bearing surface formed as a conical surface having a vertex angle equal to that of the conical surface forming the pressurizing surface, a pair of side surfaces, and a pair of end faces formed as planes each spaced from a plane including the axis. The supporting surface and the bearing surface are formed of an outer surface of a rim portion of the ring segment and an inner surface of a bore portion of the ring segment, respectively. A diameter of a circle circumscribing a projected shape of the device in a direction perpendicular to the end face is smaller than an inner diameter of a smallest-diameter portion of the main shaft.