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 turbocharger includes a shaft extending between a first shaft end coupled to a turbine wheel, and a second shaft end coupled to a compressor wheel. The turbocharger also includes a bearing housing assembly including a bearing housing presenting an interior surface having a first diameter. The interior surface defines a bearing housing interior. A bushing is disposed in the bearing housing interior, extends between first and second bushing ends, is coupled to the shaft, and includes a first anti-rotation component. A retention shoulder extends away from the interior surface and radially inward proximate to one of the first and second bushing ends, and presents an inner surface having a second diameter less than the first diameter to limit axial movement of the bushing. The retention shoulder includes a second anti-rotation component engagable with the first anti-rotation component to limit rotational movement between the bushing and the bearing housing.

A face sheet of an acoustic liner for a gas turbine engine, having: a first surface, a second surface, and a sheet body extending from the first surface to the second surface; a forward end, an aft end, and the sheet body extending from the forward end to the aft end; and a hole formed in the sheet body, extending from the first surface to the second surface, the hole defining a hole opening at the first surface of the face sheet, and an elongated edge that extends aft at the hole opening, wherein the elongated edge defines a flow ramp surface that tapers toward the first surface and is configured to direct flow toward the first surface, wherein the flow ramp surface is a streamlining feature.

An apparatus of an aircraft engine having lubricated moving components is provided. The apparatus comprises a housing having a lubricant cavity for receiving lubricant and a service port cavity, the housing defining a selectively openable service port for accessing the service port cavity. A shaft is mounted for rotation about a rotational axis to the housing. A seal is mounted about the shaft forming a partition between the lubricant cavity and the service port cavity. A vent channel extends between the lubricant cavity and the service port cavity.

The invention relates to a positioning element for a guide vane arrangement of a guide vane stage of a gas turbine, with at least one base section curved in the peripheral direction; a plurality of uptake openings arranged next to each other in the peripheral direction on the base section, whose aperture axis runs substantially in the radial direction and which are designed to take up a respective radially inner guide vane section; a coupling section provided on the base section, which is or can be coupled to a seal carrier of a seal arrangement. According to the invention, it is proposed that at least one recess is provided in the base section, which is arranged between two neighboring uptake openings and runs from inside to outside at least in the radial direction.

A device for removing and/or installing an annular component, which is arranged at a position within a turbine housing, which position is accessible via an access point of the turbine housing. A method for removing and for installing a component of this type by a device of this type.

A turbine blade having a gas film cooling structure with a composite irregular groove. The turbine blade has a hollow structure, and a plurality of first grooves which are recessed grooves are provided on an outer surface thereof. A plurality of discrete holes A extending to an inner surface of the turbine blade are provided at the groove bottom of each first groove. The first groove is an irregular groove, and includes at least two portions in a depth direction. A portion having a depth H.sub.1 from the groove bottom of the first groove is a first portion, and the rest thereof is a second portion. At least one side surface of the second portion is formed by expanding in lateral direction from a corresponding side surface of the first portion.

The invention relates to a method for introducing a balancing mark into the compressor wheel of a turbocharger. According to the method, a milling tool is firstly moved in a first direction in order to introduce a recess into the compressor wheel, and the milling tool located in the recess then runs out in a second direction in order to convert the recess into a pear segment-shaped balancing mark. The invention furthermore relates to a turbocharger which comprises a compressor wheel that has one or more pear segment-shaped balancing marks.

A grommet may define a dilution hole in a combustor panel. The grommet may comprise a ridge having a stepped geometry formed about an inner diameter of the grommet, the ridge comprising a passage. The passage may comprise an outlet. The ridge may further comprise a fillet about the inner diameter of the grommet, wherein the outlet is configured to direct a cooling flow circumferentially along the fillet and fill the ridge with the cooling flow.

A modular casing manifold for cooling fluids of a gas turbine engine is presented. The modular casing manifold has an annular shape including an axial inner plate, an axial outer plate, a radial forward plate and a radial aft plate. The forward plate is attached to the inner and outer plates at forward end. At least a portion of the aft plate is attachable to and removable from the inner and outer plates at aft end for enabling cooling fluid to cool turbine blades of the gas turbine engine. The modular casing manifold includes preswirler segments. At least a number of the preswirler segments are attachable to and removable from the forward plate for enabling cooling fluid to cool turbine blades of the gas turbine engine. The modular casing manifold enables alternative cooling fluids to cool turbine blades of the gas turbine engine with minimal cost and assembly flexibility.