The invention relates to an aircraft comprising a fuselage (1), which is propelled by a turbine engine with two coaxial fans, namely an upstream fan (7) and a downstream fan (8), driven by two contra-rotating rotors (5, 6) of a power turbine (3). The two fans (7, 8) and the turbine (3) are integrated into a nacelle (14) which projects downstream from the fuselage (1) and through which air flows. According to the invention, at least one of the fans (7, 8) of the aircraft and, in particular, the downstream fan (8) comprises variable-spacing blades, and at least one stator-forming variable-spacing blade ring (25) in the aircraft is placed upstream of the upstream fan (7). The variable-spacing stator blades (25) and the variable-spacing blades of the downstream fan (8) are mutually configured to direct the air flow in a first mode in which the air flows through the nacelle (14) from upstream to downstream and in a second mode in which the air is pushed back upstream through the nacelle (14).

An aircraft including a fuselage, a lateral support wing (1) and a propulsion assembly (100) mounted under the wing. The wing includes at least two structural spars (11ba, 11bf) extending from the fuselage toward the tip of the wing, one of these (11ba) being upstream and the other (11bf) downstream. The propulsion assembly includes a gas generator (106) and at least two offset fans (102, 104) arranged on either side of the axis of the gas generator. The offset fans (102, 104) are attached directly to one of the spars (11ba, 11bf) and the gas generator (106) is attached directly to the two spars. The leading edge of the wing forms a given sweep of angle (α) with the axis of the fuselage. The two offset fans (102, 104) are axially offset from one another.

An aircraft including a fuselage, a lateral support wing (1) and a propulsion assembly (100) mounted under the wing. The wing includes at least two structural spars (11ba, 11bf) extending from the fuselage toward the tip of the wing, one of these (11ba) being upstream and the other (11bf) downstream. The propulsion assembly includes a gas generator (106) and at least two offset fans (102, 104) arranged on either side of the axis of the gas generator. The offset fans (102, 104) are attached directly to one of the spars (11ba, 11bf) and the gas generator (106) is attached directly to the two spars. The leading edge of the wing forms a given sweep of angle (α) with the axis of the fuselage. The two offset fans (102, 104) are axially offset from one another.

Various embodiments may include a drive system for an aircraft comprising: an electrical generator; and an internal combustion engine with a gas expansion turbine. The generator includes a rotor rotating on a rotor shaft. The gas expansion turbine rotates on a turbine shaft. The rotor shaft and the turbine shaft are couplable to transmit torque such that the internal combustion engine can deliver mechanical energy to the generator. A supply air stream impinges at least one component of the generator.

A exhaust nozzle assembly comprises a nozzle that extends about an axial centerline and includes an exhaust nozzle flange; an radially inner surface that comprises an axially forward inner surface; a noise attenuating structure; a through hole inlet formed in the axially forward inner surface; a through hole outlet formed in the axially rear inner surface. The nozzle assembly also includes a radially outer surface that is radially separated from the radially inner surface by a nozzle cavity, where engine core air enters the nozzle cavity from the through hole inlet and exits the nozzle cavity axially downstream of the hole inlet via the through hole outlet.

An exhaust nozzle of a gas turbine engine includes: a nozzle wall, a centerbody arranged in a flow channel, and two struts connecting the centerbody to the wall. One of the struts is connected to the wall by a coupling arrangement that includes two first brackets and a third bracket, the brackets being spaced in an axial direction and being connected either directly to the wall or to a sliding element that is arranged in a displaceable manner in the wall. The brackets each have a first, highest stiffness in a first direction and smaller stiffnesses in a second and third direction. The brackets are oriented such that with the two first brackets, the first direction is aligned with a circumferential direction of the nozzle and that with the third bracket, the first direction is aligned with the axial direction of the nozzle.

An exhaust nozzle of a gas turbine engine includes: a nozzle wall, a centerbody arranged in a flow channel, and two struts connecting the centerbody to the wall. One of the struts is connected to the wall by a coupling arrangement that includes two first brackets and a third bracket, the brackets being spaced in an axial direction and being connected either directly to the wall or to a sliding element that is arranged in a displaceable manner in the wall. The brackets each have a first, highest stiffness in a first direction and smaller stiffnesses in a second and third direction. The brackets are oriented such that with the two first brackets, the first direction is aligned with a circumferential direction of the nozzle and that with the third bracket, the first direction is aligned with the axial direction of the nozzle.

An exhaust nozzle for use with a gas turbine engine includes an outer shroud and a nozzle-plug assembly coupled to the outer shroud. The nozzle-plug assembly includes an inner plug and at least one support vane that is coupled to the outer shroud to support the inner plug in an exhaust nozzle flow path.

An exhaust nozzle for use with a gas turbine engine includes an outer shroud and a nozzle-plug assembly coupled to the outer shroud. The nozzle-plug assembly includes an inner plug and at least one support vane that is coupled to the outer shroud to support the inner plug in an exhaust nozzle flow path.

A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan defining a fan tip radius at the fan face; wherein a downstream blockage ratio is defined as: $\frac{\begin{array}{c}\mathrm{the}\mathrm{turbine}\mathrm{diameter}\mathrm{at}\mathrm{an}\mathrm{axial}\mathrm{location}\\ \mathrm{of}\mathrm{the}\mathrm{lowest}\mathrm{pressure}\mathrm{rotor}\mathrm{stage}\end{array}}{a\mathrm{distance}f\mathrm{rom}}$