A thrust reverser of a nacelle of a turbojet engine of an aircraft extending about a longitudinal axis and including a fixed structure and at least one door which is rotatable between a closed position and an open position and including a downstream frame spaced axially from a downstream end edge of the door by an axial clearance parallel to the longitudinal axis. The thrust reverser includes a system for diverting fluids to the exterior including at least one projection fixed onto the door or the other fixed structure and extending into the axial clearance and towards one of the fixed structure or the door, while leaving an axial gap between a free end of said projection and one of the fixed structure or the door.

A thrust reverser of a nacelle of a turbojet engine of an aircraft extending about a longitudinal axis and including a fixed structure and at least one door which is rotatable between a closed position and an open position and including a downstream frame spaced axially from a downstream end edge of the door by an axial clearance parallel to the longitudinal axis. The thrust reverser includes a system for diverting fluids to the exterior including at least one projection fixed onto the door or the other fixed structure and extending into the axial clearance and towards one of the fixed structure or the door, while leaving an axial gap between a free end of said projection and one of the fixed structure or the door.

A variable area nozzle assembly for a gas turbine engine includes a nozzle disposed about a nozzle centerline and a fixed ring radially surrounding the nozzle. The nozzle includes a radially outer surface and a radially inner surface. The radially inner surface defines an outlet cross-sectional area of the nozzle. The nozzle is movable relative to the nozzle centerline between a first position of the radially inner surface defining a maximum area of the outlet cross-sectional area and a second position of the radially inner surface defining a minimum area of the outlet cross-sectional area. With the nozzle in the first position, the radially outer surface contacts the fixed ring. With the nozzle in the second position, the radially outer surface is spaced from the fixed ring.

A sealing system may include a bulb seal coupled to a first component and a spring coupled to a second component. The bulb seal may contact the spring to form a seal between the first component and the second component. The bulb seal may be relatively stiff when new and cause the spring to compress. The bulb seal may be relatively flexible as it ages, and the spring may cause the bulb seal to compress. The bulb seal and the spring may maintain a seal throughout the life of the sealing system. The sealing system may maintain a constant separation between the first component and the second component throughout the life of the sealing system.

A process for using an air input of a turboreactor nacelle of an aircraft, comprising an air input lip which comprises at least one fixed portion and at least one portion which can be moved between a first position, in which the air input lip has an aerodynamic profile so as to guide the internal air flow over the internal wall in order to promote a thrust phase, and a second position, in which the portion is displaced in relation to the fixed portion so that the air input lip has a second radial thickness in the second position which is less than the first radial thickness in the first position so as to promote a reverse thrust phase.

A propulsion unit includes a gas turbine engine arranged along an axis and an exhaust system coupled to the gas turbine engine. The gas turbine engine includes an engine core configured to discharge a core flow and a fan configured to be driven by the engine core to discharge a bypass flow. The exhaust system receives the mixed bypass and core flows from the gas turbine engine.

A propulsion unit includes a gas turbine engine arranged along an axis and an exhaust system coupled to the gas turbine engine. The gas turbine engine includes an engine core configured to discharge a core flow and a fan configured to be driven by the engine core to discharge a bypass flow. The exhaust system receives the mixed bypass and core flows from the gas turbine engine.

A propulsion unit includes a gas turbine engine arranged along an axis and an exhaust system coupled to the gas turbine engine. The gas turbine engine includes an engine core configured to discharge a core flow and a fan configured to be driven by the engine core to discharge a bypass flow. The exhaust system receives the mixed bypass and core flows from the gas turbine engine.

A propulsion unit includes a gas turbine engine arranged along an axis and an exhaust system coupled to the gas turbine engine. The gas turbine engine includes an engine core configured to discharge a core flow and a fan configured to be driven by the engine core to discharge a bypass flow. The exhaust system receives the mixed bypass and core flows from the gas turbine engine.

There is provided a load distribution panel assembly for a gas turbine engine. The assembly has a panel structure having at least one circumferential structural panel having a first end, and a second end coupled to a fixed structure of the gas turbine engine. The circumferential structural panel has a first compliant portion extending away from the first end, and has a second stiffened portion angled with respect to and extending radially away from the first compliant portion, and terminating at the second end. The second stiffened portion has a closed stiffened cavity portion integral with a perimeter flange portion. The panel assembly converts fore/aft point load(s) applied to it by load applying apparatus(es), to hoop tension and compression loads, and reacts a load offset in in-plane load(s), to provide uniform load distribution of the fore/aft point load(s) to the fixed structure.