A machine (10) with lifting system (18) includes a method for generating a resultant force (424) from circulating toroidal flow field (410) and poloidal flow field (420) within nearly confined toroid volume (413). The machine (10) architecture has designs for adaptability at assembly level to achieve modularity for ease of fabrication, maintenance, and operations.

A propulsion system for an includes a combustion engine, a propulsor, and an electric machine configured to either be driven by the combustion engine or configured to drive the propulsor. The electric machine defines an axis. The electric machine includes a rotor extending along and rotatable about the axis, and a stator having a plurality of winding assemblies, the plurality of winding assemblies spaced along the axis of the electric machine, each winding assembly operable with the rotor independently of an adjacent winding assembly during operation of the electric machine.

A propulsion unit for an aircraft is provided. The propulsion unit includes a main engine that supplies main thrust during a takeoff operating condition and a top of climb operating condition, and an auxiliary engine, distinct from the main engine, that supplies auxiliary thrust to complete the main thrust of the main engine during the takeoff operating condition. The main engine includes a high-pressure compressor. The main engine is dimensioned taking into account the thrust of the auxiliary engine in the takeoff operating condition, in such a manner that a temperature ratio of the high-pressure compressor, corresponding to the ratio between an outlet temperature of the high-pressure compressor of the main engine in the top of climb operating condition and an outlet temperature of the high-pressure compressor of the main engine in the takeoff operating condition, is between 0.90 and 1.10.

An aircraft propulsion assembly including a fan. It include a first engine and a second engine which are not coaxial and a mechanical energy transmission device configured to enable the fan to be conjointly rotated by the first engine and the second engine. This allows an aircraft propulsion assembly to be produced of which the fan may be positioned so as to ingest the boundary layer formed at the surface of a member of the aircraft equipped with the propulsion assembly, while allowing operating modes in the case of certain failures, and certification for commercial use of an aircraft equipped with such a propulsion assembly, to which the invention also relates.

An integrated propulsion system comprising at least two gas turbine engines, at least one fan, and a transmission assembly coupling the at least two gas turbine engines to the at least one fan wherein the at least two gas turbine engines are disposed within a main body of an airframe comprising the main body and a pair of wings, and wherein the number of gas turbine engines is greater than the number of fans.

An integrated propulsion system comprising at least two gas turbine engines, at least one fan, and a transmission assembly coupling the at least two gas turbine engines to the at least one fan wherein the at least two gas turbine engines are disposed within a main body of an airframe comprising the main body and a pair of wings, and wherein the number of gas turbine engines is greater than the number of fans.

A method of operation is provided during which a heating fluid is directed through a first passage into a combustor section to heat a component within the combustor section when an aircraft engine is in a first standby mode. The aircraft engine includes a flowpath, a compressor section, the combustor section, a turbine section and an exhaust section. The flowpath extends through the compressor section, the combustor section, the turbine section and the exhaust section from an inlet into the flowpath to an exhaust from the flowpath. The heating fluid is directed through a second passage into the exhaust section when the aircraft engine is in a second mode. The heating fluid is directed through the second passage bypasses the turbine section.

A method of operation is provided during which a heating fluid is directed through a first passage into a combustor section to heat a component within the combustor section when an aircraft engine is in a first standby mode. The aircraft engine includes a flowpath, a compressor section, the combustor section, a turbine section and an exhaust section. The flowpath extends through the compressor section, the combustor section, the turbine section and the exhaust section from an inlet into the flowpath to an exhaust from the flowpath. The heating fluid is directed through a second passage into the exhaust section when the aircraft engine is in a second mode. The heating fluid is directed through the second passage bypasses the turbine section.

The invention concerns an aircraft propelled by a turbine engine having contrarotating fans (7, 8), the turbine engine being incorporated at the rear of a fuselage (1) of the aircraft, in the extension of same and comprising at least two gas generators (2a, 2b) that supply, via a shared central stream (4), a power turbine (3), the turbine (3) comprising two contrarotating rotors (5, 6) for driving two fans (7,8) disposed downstream from the gas generators (2a, 2b), said aircraft comprising means (15) arranged for separating the gas flow in the power turbine (3) into at least two concentric streams (16, 17) and a device comprising first means for distributing the gas flow (21-24) between said streams (16, 17) from the central stream (4), the first distribution means being configured to be able to open or close the supply of at least one so-called sealable stream (16) of the streams (16, 17) of the power turbine (3).

The invention concerns an aircraft propelled by a turbine engine having contrarotating fans (7, 8), the turbine engine being incorporated at the rear of a fuselage (1) of the aircraft, in the extension of same and comprising at least two gas generators (2a, 2b) that supply, via a shared central stream (4), a power turbine (3), the turbine (3) comprising two contrarotating rotors (5, 6) for driving two fans (7,8) disposed downstream from the gas generators (2a, 2b), said aircraft comprising means (15) arranged for separating the gas flow in the power turbine (3) into at least two concentric streams (16, 17) and a device comprising first means for distributing the gas flow (21-24) between said streams (16, 17) from the central stream (4), the first distribution means being configured to be able to open or close the supply of at least one so-called sealable stream (16) of the streams (16, 17) of the power turbine (3).