A device for balancing a shaft in a turbomachine engine includes an annular insert configured to be positioned inside the shaft at any axial position along the shaft, and a weight that is attached to the annular insert at any angular position inside the shaft.

In accordance with at least one aspect of this disclosure, there is provided a drive shaft system. In embodiments, the drive shaft system includes a generator shaft extending along a longitudinal axis with a longitudinal bore defined through the generator shaft. A thru shaft extends through the longitudinal bore of the generator shaft. In embodiments, the thru shaft includes, a front coupler mount at a first end operative to receive torque input, and a rear coupler mount at a second end opposite the first end.

An assembly is provided for a turbine engine. This turbine engine assembly includes an impeller rotor, a seal land and a lip seal. The impeller rotor is configured to rotate about a rotational axis. The impeller rotor is configured from or otherwise includes impeller rotor material. The seal land extends axially along and circumferentially about the rotational axis. The seal land is mechanically attached to and rotatable with the impeller rotor. The seal land is configured from or otherwise includes seal land material that is different than the impeller rotor material. The lip seal radially engages the seal land.

A planetary gear box, which has a sun gear, a plurality of planet gears, a ring gear, a plurality of planet pins an axially front carrier plate and an axially rear carrier plate, wherein the planet pins are connected to the carrier plates. Provision is made whereby the planet pins, at their axial ends, have in each case one platform, radially oriented slots are formed on the carrier plates, and the platforms are pushed via guide grooves into the radially oriented slots. The invention furthermore relates to a method for the assembly of a planetary gear box.

A fuel pump arrangement including a housing and an impeller located within the housing configured for receiving cold fuel through an inlet of the housing, wherein the impeller includes a series of vanes for drawing fluid and a nose cone defining an axis, wherein the nose cone includes a channel therethrough for passing warm fuel to an exterior surface of the nose cone for mixing with the cold fuel.

A gas turbine engine for an aircraft. The gas turbine engine comprises a rotatable shaft defining a rotational axis extending between rearward and forward ends of the gas turbine engine and a compressor drum surrounding, and coupled to, the shaft so as to define an annular gap therebetween. The gas turbine engine further comprises an intermediate case arranged axially rearward of the compressor drum and comprising a bearing support element extending into the annular gap, and a forward bearing mounted between the bearing support element and the shaft proximate a forward end of the compressor drum.

A centrifugal compressor and a refrigerating device. The compressor includes: a shell, which has a fluid inlet at a first position of the shell, and a fluid outlet at a second position of the shell, a motor assembly, which is arranged in the shell and includes a stator and a rotor, the rotor including a vertically arranged rotor shaft, and the rotor shaft including a lower end and an upper end; a centrifugal compression mechanism, an impeller of which is connected with the rotor shaft so as to be driven by the motor assembly; and a guide member, which is located above the centrifugal compression mechanism, and which defines a flow passage alone or together with a top part of the shell.

An air starter for an engine, such as a turbine engine, that includes a turbine for rotatably extracting mechanical power from a flow of fluid. The turbine includes a drive shaft that can be coupled to an output shaft to provide a rotational output. The interface between the drive shaft and the output shaft can be offset.

The invention relates to a shaft for supplying air to a fuel cell, comprising a tubular body (12), at least sections of which are hollow, and which has a first component (14) and a second component (16) that are connected to one another at respective neighbouring axial end sections (14a, 16a), and wherein air can flow through a hollow space (12a) of the tubular body (12) for cooling components arranged next to the shaft (10). The invention also relates to a radial compressor (1). The invention further relates to a method for producing a radial compressor (1).

An assembly comprising a vent tube intended to be mounted inside a turbine shaft of a turbomachine. The vent tube has a radially inner annular part from which at least one annular centering part extends radially outwards. The centering part has a groove and a sealing part of a sealing member is mounted in the groove. The sealing part is designed to come into sealing abutment against an inner surface of the turbine shaft. The sealing member has a protective part covering at least one area of the centering part situated upstream and/or downstream of the groove.