An engine system is provided that includes an engine rotating structure, an electric machine rotating structure and a flex coupler. The flex coupler rotatably connects the electric machine rotating structure to the engine rotating structure. The flex coupler includes a first mount, a second mount and a flex plate. The first mount includes a plurality of first mount fingers arranged circumferentially about an axis. The second mount includes a plurality of second mount fingers arranged circumferentially about the axis. The flex plate connects the first mount to the second mount. The flex plate includes a plurality of first flex plate fingers and a plurality of second flex plate fingers. Each of the first flex plate fingers is attached to a respective one of the first mount fingers. Each of the second flex plate fingers is attached to a respective one of the second mount fingers.

A hybrid electric gas turbine engine is provided. The hybrid electric gas turbine engine includes: a turbomachine having a compressor section and a turbine section arranged in serial flow order, the compressor section and turbine section together defining a core air flowpath, the turbomachine defining a core air flowpath exhaust; and an electric machine assembly having an electric machine disposed aft of the core air flowpath exhaust and mechanically connected to the turbine section.

A fan drive turbine drives a drive shaft. A compressor section and a turbine section rotate about a rotational axis, with the turbine section positioned in a downstream direction. A fan has fan blades and a fan hub positioned radially within a fan case, with the fan case defining a bypass duct. A fan inlet guide vane case is upstream of the fan blades and a fan exit guide vane case is mounted downstream. The drive shaft is mounted on at least one input drive shaft bearing inwardly of the fan inlet guide vane case. The drive shaft drives a flexible coupling to in turn drive a gear reduction. The gear reduction drives a fan drive shaft to in turn drive the fan hub. At least two fan drive shaft bearings support the fan drive shaft within the fan exit guide vane case. An aircraft is also disclosed.

A fan drive turbine drives a drive shaft. A compressor section and a turbine section rotate about a rotational axis, with the turbine section positioned in a downstream direction. A fan has fan blades and a fan hub positioned radially within a fan case, with the fan case defining a bypass duct. A fan inlet guide vane case is upstream of the fan blades and a fan exit guide vane case is mounted downstream. The drive shaft is mounted on at least one input drive shaft bearing inwardly of the fan inlet guide vane case. The drive shaft drives a flexible coupling to in turn drive a gear reduction. The gear reduction drives a fan drive shaft to in turn drive the fan hub. At least two fan drive shaft bearings support the fan drive shaft within the fan exit guide vane case. An aircraft is also disclosed.

The invention relates to a system for rotating a fan (2) of a turbojet engine (1) about a first rotation axis , including a reduction gear (3) made up of a planetary gearset placed at the centre of the fan (2), which includes: a central sun gear (31); an outer planet gear (33) attached to the fan (2); at least one satellite gear (32) arranged between the central sun gear (31) and the outer planet gear (33) in order to transmit a rotary movement between the central sun gear (31) and the outer planet gear (33); characterised in that the outer planet gear (33) includes an inner portion (33a) which meshes with the satellite gear (32), as well as an outer portion (33b) to which blades (21) of the fan (2) are directly attached.

Disclosed is a gas turbine engine comprising: a low pressure spool having a low pressure compressor and a low pressure turbine connected by a low pressure shaft; a reduction gear train having a sun gear, a carrier having a plurality of planet gears attached thereto, and a ring gear, wherein the sun gear is driveably connected to the low pressure shaft, and either of carrier and ring gear provides an output drive connected to a propulsive fan; wherein the connection between the low pressure shaft and the sun gear includes a flexible drive shaft having serially connected concentrically nested first and second input shafts between the low pressure shaft and sun gear.

What is described is an aircraft engine with a compressor device and with a fan device. In the area of a compressor shaft, the compressor device is connected via an epicyclic gear to a fan shaft. A planetary web is operatively connected via bearing devices to the planetary gears. A ring gear is coupled with the fan shaft and a sun gear of the epicyclic gear is coupled with the compressor shaft, while the planetary carrier is held at the housing side in a torque-proof manner. The ring gear is connected via a flexible connection device to the fan shaft and/or the planetary carrier is connected via a flexible connection device to the housing. In the area of the connection device, movements between the ring gear and the planetary gears and/or between the planetary carrier and the housing can be at least approximately compensated in the radial and the axial direction.

A gas turbine engine comprises a fan drive turbine driving a shaft. The shaft engages a gear reduction. The gear reduction drives a fan rotor at a speed that is less than the speed of the fan drive turbine. The gear reduction is a non-epicyclic gear reduction.

Embodiments of a compliant shaft for engines are provided herein. In some embodiments, a compliant shaft for a turbine engine may include: a body having a first end configured to be coupled to a shaft of a turbine and a second end coupled to a gear of a gear box, wherein at least a portion of the body is flexible.

An aircraft gas turbine engine comprises a high pressure compressor coupled to a high pressure turbine by a high pressure shaft, and a low pressure compressor coupled to a low pressure turbine by a low pressure shaft. The engine includes a first shaft bearing configured to support a forward end of the low pressure shaft, the first shaft bearing being mounted to a static structure forward of the high pressure turbine. The engine also includes a second shaft bearing configured to support a rearward end of the low pressure shaft, the second shaft bearing being mounted to a static structure rearward of the high pressure turbine, and forward of the low pressure turbine. The engine further includes a third shaft bearing configured to support the low pressure shaft, the third shaft bearing being mounted to a static structure rearward of the low pressure turbine.