A gas turbine engine includes a fan assembly having a plurality of fan blades; and a turbomachine. The turbomachine includes a compressor section, a combustion section, and a turbine section in serial flow order. The turbomachine further including a first input power source; a second input power source configured to counter-rotate relative to the first input power source; a power output component operably connected to the fan assembly; and a gear assembly located forward of the combustion section of the turbomachine, the gear assembly configured to receive power from the first input power source and the second input power source and provide power to the power output component, the gear assembly comprising a helical gear.

A cable that includes a conductor defining a hollow interior, a casing surrounding the conductor, an electrical insulator positioned between the conductor and the casing, and a fluid positioned within the hollow interior of the conductor.

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 oil collector for a mechanical reduction gear of a turbomachine, in particular for an aircraft, the reduction gear including a body having two opposite lateral faces configured to extend in part around planet gears of the reduction gear, the collector further including an internal oil circulation cavity connected firstly to oil inlets located on the faces, and on the other hand to at least one oil outlet, characterised in that at least one of the faces comprises includes columns and rows of several inlets each having a recess with a progressively increasing cross-section, each recess being delimited by walls, at least some of the walls having a hydrodynamic profile.

A turbomachine engine can include a fan assembly, a vane assembly, a core engine, a gearbox, and a gearbox efficiency rating. The fan assembly can include a plurality of fan blades. The vane assembly can include a plurality of vanes, and the vanes can, in some instances, be disposed aft of the fan blades. The core engine can include one or more compressor sections and one or more turbine sections. The gearbox includes an input and an output. The input is coupled to the one or more turbine sections of the core engine and comprises a first rotational speed, the output is coupled to the fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8.

A propulsor section includes a propulsor having a plurality of blades rotatable about an engine longitudinal axis. A compressor section includes a low pressure compressor and a high pressure compressor. A turbine section includes a low pressure turbine that drives the propulsor through an epicyclic gear arrangement, and includes a second turbine that drives the high pressure compressor. A power ratio is provided by the combination of a first power input of the low pressure compressor and a second power input of the high pressure compressor. The power ratio is defined by the second power input divided by the first power input. The power ratio is equal to, or greater than, 1.0 and less than, or equal to, 1.4.

A propulsor section includes a propulsor having a plurality of blades rotatable about an engine longitudinal axis. A compressor section includes a low pressure compressor and a high pressure compressor. A turbine section includes a low pressure turbine that drives the propulsor through an epicyclic gear arrangement, and includes a second turbine that drives the high pressure compressor. A power ratio is provided by the combination of a first power input of the low pressure compressor and a second power input of the high pressure compressor. The power ratio is defined by the second power input divided by the first power input. The power ratio is equal to, or greater than, 1.0 and less than, or equal to, 1.4.

Disclosed is an air outlet of a nacelle for an aircraft turbojet engine, the nacelle forming a solid of revolution about a longitudinal axis, the air outlet having a straightening device having a plurality of flaps mounted circumferentially and protruding, each flap being mounted pivoting around a pivot axis, forming, with respect to the longitudinal axis, an angle of convergence in a radial plane, between a closed position, in which each flap extends along the pivot axis in the aerodynamic prolongation of the trailing edge in order to support a thrust phase and a deployed position, in which each flap extends in a deployed plane forming an angle of deployment with respect to the closed position about the pivot axis, so as to support a reverse thrust phase.

A gas turbine engine includes a fan positioned at an engine central longitudinal axis, and a fan drive turbine located at the engine central longitudinal axis and configured to drive rotation of the fan. A gas generator is non-coaxial with the fan drive turbine and operably connected to the fan drive turbine such that exhaust from the gas generator drives rotation of the fan drive turbine. An auxiliary power core is located at the engine central longitudinal axis, and one or more bleed passages connect the gas generator and the auxiliary power core. The one or more bleed passages are configured to selectably combine a bleed airflow from the gas generator and an auxiliary core airflow at the auxiliary power core to direct the combined airflow to the fan drive turbine to increase output of the fan drive turbine.

A gas turbine engine includes a fan section, a compressor section, and a turbine section. The fan section has a plurality of vane assemblies spaced circumferentially about an engine axis. The vane assemblies each include an airfoil extending between a leading edge and a trailing edge, a control rod extending through the airfoil, and a mechanism driven by the control rod to change the shape of the airfoil. A vane system for a gas turbine engine is also disclosed.