A gas turbine engine includes a core engine defined about an axis, a gear system driven by the core engine, and a pylon variable area flow system. A fan is driven by the gear system. The variable area flow system operates to effect the bypass flow.

A nacelle for a turbomachine is provided and includes an inner wall that is annular about a longitudinal axis of the nacelle, the annular inner wall being designed to surround part of the turbomachine. The nacelle further includes an annular outer wall surrounding the annular inner wall. The annular outer wall includes a first acoustically porous part, and the annular inner wall including a second acoustically porous part. The first and second porous parts are arranged facing one another so as to allow soundwaves, emitted by the turbomachine housed in the nacelle, to pass through the annular inner wall and then the annular outer wall so as to escape from the nacelle.

An aircraft comprising a fuselage and a propulsion assembly, the propulsion assembly comprising at least one fan rotor located at the rear of the fuselage in the extension thereof along a longitudinal axis, and a nacelle forming a fairing of the at least one fan rotor into which a flow of air passes. The aircraft also comprises a plurality of radial stator arms mounted upstream of the at least one fan rotor and extending between the fuselage and the nacelle, the radial arms comprising blowing means configured for blowing, into the environment of a trailing edge of the radial arms, an additional air flow adding to the airflow in the extension of the trailing edge.

A jet engine which includes: a fan with a plurality of stages of rotor blades; a compressor compressing air sent from the fan; a combustor generating combustion gas by using compressed air generated by the compressor; a turbine generating a driving force from the combustion gas; a nozzle discharging the combustion gas; a variable guide vane disposed upstream of the rotor blades of a second and later stage of the rotor blades of the fan and adjusts an inlet angle of air flow against the second and later stage of the rotor blades; a fluid resistance adjusting device adjusting a fluid resistance at the nozzle; and a controller which controlling the variable guide vane such that the inlet angle at the time of cruise flight is smaller than the inlet angle at the time of acceleration.

An aircraft extending between a forward end and an aft end is provided. The aircraft includes an auxiliary power unit positioned proximate the aft end of the aircraft, the auxiliary power unit having an auxiliary power unit inlet duct and an auxiliary power unit exhaust duct, and a boundary layer ingestion fan positioned proximate the aft end of the aircraft, the boundary layer ingestion fan having a support shaft, wherein the auxiliary power unit exhaust duct extends through a portion of the support shaft of the boundary layer ingestion fan.

A planetary gearbox system with a ring gear, a sun gear, a carrier and planetary gears as parts in a geared aircraft turbofan engine, with one static part of the parts being in rest under nominal operation and with a kinematic switch device for decoupling the static part to become a rotatable part or with a kinematic switch device for coupling the rotatable part to become a static part. The invention also relates to a method for operating a planetary gearbox system.

Systems and methods are provided for an air-driven augmentor fan equipped aircraft propulsor. The augmentor fan may increase the effective bypass ratio of the aircraft propulsor and reduce fuel consumption and carbon emissions of the aircraft. The augmentor fan may be driven by air energized by a ducted fan powered by the core engine of the aircraft propulsor. The energized air may be received by an inlet, flowed through a flow path, and exhausted out the outlet to drive the augmentor fan. The exhausted energized air may impart a torque on the augmentor fan or blades of the augmentor fan. One or more of the inlet, flow path, or outlet may be variable in size to control the volume of air flowed through the flow path.

A fan system includes a rotor having plurality of blades and a ring airfoil, the plurality of blades being rotatably joined to a hub and the ring airfoil. The fan system may include a second contra-rotationally disposed rotor having a plurality of blades and a ring airfoil. The first and second ring airfoils having a cambered shape and an angle of attack between about 5 degrees and about 45 degrees. Passive leading edge slats are attached to the ring airfoil and are configured to open and close using springs.

A turbine engine includes a first compression section includes a last blade trailing edge radial tip length that is greater than about 67% of the radial tip length of a leading edge of a first stage of the first compression section. A second compression section includes a last blade trailing edge radial tip length that is greater than about 57% of a radial tip length of a leading edge of a first stage of the first compression section.

A turbofan engine includes a fan section including a fan blade having a leading edge and hub to tip ratio of less than about 0.34 and greater than about 0.020 measured at the leading edge and a speed change mechanism with gear ratio greater than about 2.6 to 1. A first compression section includes a last blade trailing edge radial tip length that is greater than about 67% of the radial tip length of a leading edge of a first stage of the first compression section. A second compression section includes a last blade trailing edge radial tip length that is greater than about 57% of a radial tip length of a leading edge of a first stage of the first compression section.