A turbofan engine according to an example of the present disclosure includes, among other things, a fan section including a plurality of fan blades, a gear train, a low pressure turbine driving the fan section through the gear train, a fan nacelle and a core nacelle, the fan nacelle at least partially surrounding the core nacelle, a fan bypass flow path defined between the core nacelle and the fan nacelle, and a fan variable area nozzle in communication with the fan bypass flow path, and defining a fan nozzle exit area between the fan nacelle and the core nacelle. The fan variable area nozzle includes a first fan nacelle section and a second fan nacelle section, the second fan nacelle section movably mounted relative the first fan nacelle section and moveable axially along an engine axis of rotation relative the first fan nacelle section, defining an auxiliary port that extends between the first fan nacelle section and the second fan nacelle section.

A variable area nozzle assembly for a gas turbine engine includes a fixed structure surrounding an exhaust duct extending along a nozzle centerline. The fixed structure includes an upper side and a lower side opposite the upper side. The variable area nozzle assembly further includes a nozzle disposed about the nozzle centerline. The nozzle includes a nozzle throat cross-sectional area and a nozzle outlet cross-sectional area downstream of the nozzle throat cross-sectional area. The nozzle includes an upper panel and a lower panel. The upper panel includes an upper downstream end and the lower panel including a lower downstream end. The upper downstream end and the lower downstream end define a portion of the nozzle throat cross-sectional area. The variable area nozzle assembly further includes a nozzle actuation system including an upper shaft connected to the upper panel and a lower shaft connected to the lower panel.

A variable area nozzle assembly for a gas turbine engine includes a fixed structure surrounding an exhaust duct extending along a nozzle centerline. The fixed structure includes an upper side and a lower side opposite the upper side. The variable area nozzle assembly further includes a nozzle disposed about the nozzle centerline. The nozzle includes a nozzle throat cross-sectional area and a nozzle outlet cross-sectional area downstream of the nozzle throat cross-sectional area. The nozzle includes an upper panel and a lower panel. The upper panel includes an upper downstream end and the lower panel including a lower downstream end. The upper downstream end and the lower downstream end define a portion of the nozzle throat cross-sectional area. The variable area nozzle assembly further includes a nozzle actuation system including an upper shaft connected to the upper panel and a lower shaft connected to the lower panel.

A system for a thrust reverser of an aircraft includes a primary sleeve and a secondary sleeve having cascades. The secondary sleeve is coupled to a set of blocker doors. The sliding motions of the primary sleeve and the secondary sleeve are not directly coupled when each moves between its stowed and deployed positions. The sliding motion of the primary sleeve may begin at a different time and continue at a different rate from the sliding motion of the secondary sleeve.

A turbofan gas turbine engine includes a core engine within a core nacelle, a fan nacelle at least partially surrounding the core nacelle to define a bypass flow path and a variable fan nozzle exit area for bypass flow, and a pylon variable area flow system which operates to effect the bypass flow. A method of operating a turbofan gas turbine engine is also disclosed.

A turbofan engine includes fan section including a plurality of fan blades, a gear train, a low spool including a low pressure turbine and a low pressure compressor, the low pressure turbine driving the plurality of fan blades through the gear train, and a high spool including a high pressure turbine driving a high pressure compressor. A fan nacelle at least partially surrounds a core nacelle to define a fan bypass flow path. A fan variable area nozzle is in communication with the fan bypass flow path and defines a fan nozzle exit area between the fan nacelle and the core nacelle. The fan variable area nozzle varies the fan nozzle exit area.

A turbofan engine includes fan section including a plurality of fan blades, a gear train, a low spool including a low pressure turbine and a low pressure compressor, the low pressure turbine driving the plurality of fan blades through the gear train, and a high spool including a high pressure turbine driving a high pressure compressor. A fan nacelle at least partially surrounds a core nacelle to define a fan bypass flow path. A fan variable area nozzle is in communication with the fan bypass flow path and defines a fan nozzle exit area between the fan nacelle and the core nacelle. The fan variable area nozzle varies the fan nozzle exit area.

A turbofan gas turbine engine includes a core engine within a core nacelle, a fan nacelle at least partially surrounding the core nacelle to define a bypass flow path and a variable fan nozzle exit area for bypass flow, and a pylon variable area flow system which operates to effect the bypass flow. A method of operating a turbofan gas turbine engine is also disclosed.

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 mixing tube with multiple shapes is provided, allowing additional injection of gas in order to keep the flow detached from the second shape in an ascent phase and to bring about, in a descent phase, a controlled detachment as a result of the change of slope between the two shapes. A propulsion nozzle for an engine of a spacecraft or aircraft is provided including such a mixing tube and a method for controlling the speed transition of the propulsion gases in such a nozzle in accordance with the altitude. Also, a method is provided for vectorising the thrust in such a nozzle by radial and asymmetrical injection of gas and a control method which prevents re-attachment of the jet to the second shape of such a propulsion nozzle for an engine of a spacecraft when it is in the take-off or landing phase.