Turbogenerator (1) for an aircraft (2) comprising:—a turboshaft engine (3); —an electric generator (4) comprising a rotor (5) driven mechanically by the turboshaft engine (3) and a stator (6) supported by a housing (7) of the electric generator (4); characterized in that the turbogenerator (1) comprises a static separator (8) for separating an air/oil mixture coming from the turboshaft engine (3), the static separator (8) being positioned around the housing (7) of the electric generator (4).

Turbogenerator (1) for an aircraft (2) comprising:—a turboshaft engine (3); —an electric generator (4) comprising a rotor (5) driven mechanically by the turboshaft engine (3) and a stator (6) supported by a housing (7) of the electric generator (4); characterized in that the turbogenerator (1) comprises a static separator (8) for separating an air/oil mixture coming from the turboshaft engine (3), the static separator (8) being positioned around the housing (7) of the electric generator (4).

An internal combustion engine, ICE, system, includes a turbocharger having a turbine and a compressor for compressing intake air and feeding the intake air to the ICE. A turbo turbine unit is disposed in an exhaust gas path downstream the turbocharger to receive exhaust gas from the turbocharger. The turbo turbine unit having a turbine wheel, a bearing housing defining an inside volume for containing lubrication oil, and a sealing arrangement positioned in the vicinity of the turbine wheel for preventing lubrication oil from escaping from the inside volume of the bearing housing to an exhaust duct of the turbo turbine unit. A buffer air conduit extends between a selected position at the compressor and a buffer air inlet of the turbo turbine unit. The buffer air inlet is in fluid communication with a buffer air channel inside the turbo turbine unit to direct buffer air to a position in-between axially opposite arranged annular sealing members to provide a counter-pressure against at least one of the annular sealing members.

A bearing assembly for a charging apparatus. The bearing assembly comprises a bearing housing and a shaft. The bearing assembly further comprises a compressor-side bearing bushing and a turbine-side bearing bushing which together support the shaft inside a bearing bore of the bearing housing. The compressor-side bearing bushing is configured differently than the turbine-side bearing bushing.

A propulsion system includes a drive shaft, a fan, a fan shaft, and a reduction device coupling the drive and fan shafts. The reduction device includes a first reduction stage and a second reduction state, and include a sun gear, centered on an axis of rotation of the drive and fan shafts and driven in rotation by the drive shaft, a ring gear, coaxial with the sun gear and that drives the fan shaft in rotation about the axis, and planet gears distributed circumferentially about the axis between the sun and ring gears, each planet gear including a first portion which is meshed with the sun gear and a second portion which is meshed with the ring gear, a diameter of the first portion being different from a diameter of the second portion, and an oil transfer bearing positioned between the fan and the reduction device.

A service tube assembly for an aircraft engine, comprising: a service tube having a threaded end portion, an opposed end portion and an annular tube surface proximate to the threaded end portion; a housing having an outer surface defining a tube socket extending in the outer surface, and a ramp extending toward the tube socket so as to define an engagement direction, the tube socket engaged with the threaded end portion of the service tube; a locking member having a bottom surface disposed against the ramp and an engagement surface facing toward the service tube, the locking member slidable along the ramp in the engagement direction between a first member position in which the engagement surface is spaced from the annular tube surface and a second member position in which the engagement surface contacts the annular tube surface; and a fastener releasably holding the locking member against the ramp.

A service tube assembly for an aircraft engine, comprising: a service tube having a threaded end portion, an opposed end portion and an annular tube surface proximate to the threaded end portion; a housing having an outer surface defining a tube socket extending in the outer surface, and a ramp extending toward the tube socket so as to define an engagement direction, the tube socket engaged with the threaded end portion of the service tube; a locking member having a bottom surface disposed against the ramp and an engagement surface facing toward the service tube, the locking member slidable along the ramp in the engagement direction between a first member position in which the engagement surface is spaced from the annular tube surface and a second member position in which the engagement surface contacts the annular tube surface; and a fastener releasably holding the locking member against the ramp.

A split labyrinth seal assembly having split stationary and rotary sealing elements, and a split clamping mechanism. The clamping mechanism is configured to secure the rotary element to a rotating shaft. The rotary element can also include an integrally formed valve element that moves between a contacting and a non-contacting position in response to rotation of the shaft. The rotary element and the clamping mechanism can have surface features associated therewith to promote nesting between the components.

A method of regulating pressure in a thermal transport bus of a gas turbine engine, the method including: operating the gas turbine engine with the thermal transport bus having an intermediary heat exchange fluid flowing therethrough, the thermal transport bus including one or more heat source heat exchangers and one or more heat sink heat exchangers in thermal communication through the intermediary heat exchanger fluid; and adjusting a flow volume of the thermal transport bus using a variable volume device in fluid communication with the thermal transport bus in response to a pressure change associated with the thermal transport bus.

A lubrication system is provided for a turbine engine. A lubricant source includes a source outlet. Feed circuits are fluidly coupled with the source outlet in parallel. The feed circuits include a first feed circuit and a second feed circuit. The second feed circuit includes a pump with a pump inlet and a pump outlet. A bypass circuit is fluidly coupled with the pump inlet and the pump outlet. A bleed circuit is fluidly coupled with the first feed circuit. A flow regulator is configured to regulate flow through the bypass circuit during a first mode of operation and a second mode of operation. The flow regulator is configured to close the bleed circuit during the second mode of operation. A sensor system is configured to monitor fluid flow directed to the first feed circuit and/or the second feed circuit.