A hybrid jet engine, including a front cowling having a tubular shape, a rear cowling having a tubular shape and connected to the front cowling, a central core disposed within the front cowling and the rear cowling, a shaft disposed longitudinally within the central core, a plurality of main fan blades disposed at a first end of the shaft at the front cowling, a plurality of auxiliary fan blades disposed at a second end of the shaft at the rear cowling, at least one aerodynamic stabilizer disposed on a surface of the central core to extend therefrom, and at least one Tesla one way valve disposed on the surface of the central core to receive the ram air and to provide the ram air to the plurality of auxiliary fan blades.

A hybrid jet engine, including a front cowling having a tubular shape, a rear cowling having a tubular shape and connected to the front cowling, a central core disposed within the front cowling and the rear cowling, a shaft disposed longitudinally within the central core, a plurality of main fan blades disposed at a first end of the shaft at the front cowling, a plurality of auxiliary fan blades disposed at a second end of the shaft at the rear cowling, at least one aerodynamic stabilizer disposed on a surface of the central core to extend therefrom, and at least one Tesla one way valve disposed on the surface of the central core to receive the ram air and to provide the ram air to the plurality of auxiliary fan blades.

The present invention relates to an aircraft including a fuselage and a thruster downstream of the fuselage. The thruster includes a power turbine, located inside a main flow jet, and at least one fan, located inside a secondary flow jet and mechanically driven by the power turbine. The main flow jet of the power turbine and the secondary flow jet of the fan are concentric. The power turbine is supplied with gases from two gas turbine gas generators via two supply channels. The gas turbine gas generators have axes parallel to that of the fuselage. The air inlet sleeve is spaced apart from the fuselage, and the supply channels each have a hatch for controlling the flow between a position for guiding the gas flow to the power turbine and a position for ejecting the gases into the atmosphere while bypassing the power turbine.

A gas turbine engine includes an engine core including: a compressor system including first, lower pressure, compressor, and a second, higher pressure, compressor; and an outer core casing surrounding the compressor system and including a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection, wherein the first flange connection is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on the trailing edge of the most downstream aerofoil of the first compressor and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor; a nacelle surrounding the engine core and defining a bypass duct between the engine core and the nacelle; wherein an axial midpoint of the radially outer edge is defined as the fan OGV tip centrepoint.

An aircraft propulsion assembly, comprising a gas generator and two fans rotated by the gas generator and offset on either side of a vertical plane passing through the axis of said gas generator. The propulsion assembly comprises an air inlet sleeve comprising an inlet pipe oriented along a first axis that is substantially parallel and offset with respect to a longitudinal axis of the gas generator, the inlet pipe dividing into a supply pipe that is connected to an inlet opening of the gas generator and a discharge pipe configured such that particles ingested by the inlet pipe are discharged without entering the gas generator.

An aircraft propulsion assembly, comprising a gas generator and two fans rotated by the gas generator and offset on either side of a vertical plane passing through the axis of said gas generator. The propulsion assembly comprises an air inlet sleeve comprising an inlet pipe oriented along a first axis that is substantially parallel and offset with respect to a longitudinal axis of the gas generator, the inlet pipe dividing into a supply pipe that is connected to an inlet opening of the gas generator and a discharge pipe configured such that particles ingested by the inlet pipe are discharged without entering the gas generator.

A hybrid gas turbine engine comprises a turbine, a compressor, and a gear system that is driven by the turbine where the gear system includes a sun gear, a ring gear and a plurality of planet gears, where the ring gear has a ring gear radius of R.sub.Ring gear and the sun gear has a sun gear radius of R.sub.Sun gear such that the gear system provides a reduction ratio of R.sub.Ring gear/R.sub.Sun gear. A fan is driven by the ring gear, and a generator is driven by the ring gear and provides a generator output signal. A power circuit receives the generator output signal and provides a gain R.sub.Ring gear/R.sub.Sun gear to the generator output signal and provides a gain signal indicative thereof to an electrical load.

A hybrid gas turbine engine comprises a turbine, a compressor, and a gear system that is driven by the turbine where the gear system includes a sun gear, a ring gear and a plurality of planet gears, where the ring gear has a ring gear radius of R.sub.Ring gear and the sun gear has a sun gear radius of R.sub.Sun gear such that the gear system provides a reduction ratio of R.sub.Ring gear/R.sub.Sun gear. A fan is driven by the ring gear, and a generator is driven by the ring gear and provides a generator output signal. A power circuit receives the generator output signal and provides a gain R.sub.Ring gear/R.sub.Sun gear to the generator output signal and provides a gain signal indicative thereof to an electrical load.

The present invention relates to an aircraft comprising a fuselage (1) and a propulsion unit at the rear of the fuselage, the propulsion unit comprising at least one fan rotor (7, 8), a nacelle (14) fairing the fan and at least one connection means (15) connecting the nacelle to the fuselage, the fan being rotated by the energy supplied by at least one gas-turbine gas generator (2a, 2b) housed in the fuselage, said gas generator comprising auxiliary equipment cooled by a cooling circuit. The aircraft is characterised in that said cooling circuit comprises at least one heat exchanger exchanging heat with the ambient air housed in one of said connection means (15) and/or in said nacelle (14). The cooling circuit optionally comprises also a heat exchanger exchanging heat with the ambient air, housed in the tail unit.

Disclosed is an air cycle machine (ACM) having: a turbine; a compressor; a compressor shaft connected to the compressor and configured to receive rotational energy from a gearbox; and a turbine shaft connected to the turbine and configured to provide rotational energy to the gearbox; wherein the turbine shaft and the compressor shaft operate at different rotational speeds.