A jet engine having a flow duct that is arranged in an engine nacelle and is delimited radially on the inside by a central body and radially on the outside by an outer nozzle wall is described. The free cross section of the flow duct is variable by an elastically embodied, radially variable wall of at least one of these components. The radially variable wall at least regionally has an at least approximately wavelike structure and is connected to at least one adjusting device having an adjusting unit. The variable wall of one of these components is variable in a radial direction at two defined control sections, at least one of which defines an axial wall end of the radially variable wall.

A gas turbine engine includes a fan, a nacelle arranged about the fan, and an engine core at least partially within the nacelle. A fan bypass passage downstream of the fan between the nacelle and the gas turbine engine conveys a bypass airflow from the fan. A nozzle associated with the fan bypass passage is operative to control the bypass airflow. The nozzle includes a shape memory material having a first solid state phase that corresponds to a first nozzle position and a second solid state phase that corresponds to a second nozzle position.

A gas turbine engine includes a first duct, a second duct, and a bladder. The bladder is disposed in the first duct and communicates with the second duct. A first gas flow is capable of passing through the first duct and a second gas flow is capable of passing through the second duct. The bladder is adapted to receive a bleed gas flow from the second duct and inflate within the first duct, thereby decreasing an area in the first duct through which the first gas flow is capable of traveling.

An exhaust section of a gas turbine engine includes an exhaust plug defining a plurality of exhaust plug apertures circumferentially spaced from each other. Also included is an exhaust nozzle radially offset from the exhaust plug defining an exhaust pathway between the exhaust plug and the exhaust nozzle. Further included is an exhaust plug liner having a non-uniform outer surface axially aligned with the exhaust plug apertures. The exhaust plug liner is rotatable relative to exhaust plug between a first position and a second position to selectively change a cross-sectional area of the exhaust pathway during thrust reversal operation to increase an amount of reverse thrust.

A turbine engine having an engine core, an inner cowl radially surrounding the engine core, an outer cowl radially surrounding the inner cowl and spaced from the inner cowl to form an annular passage between the inner and outer cowls that defines a nozzle, at least one control surface provided on the inner cowl and movable between a retracted position, where the nozzle has a first cross-sectional area, and an extended position where the nozzle has a second cross-sectional area that is less than the first cross-sectional area and an actuator operably coupled to the control surface and configured to move the control surface to control the cross-sectional area of the nozzle.

A seal system for sealing a first component with a second component is provided. The seal system includes layers of abutting planar seal segments. A first seal element is in a slot in the first component to slidably seal against a surface of the layers, and a second seal element is on a surface of the second component and engages with the second end of certain layer(s). Members extend from seal segments of one of the layers and a tension cable engages the members. A tensioner is coupled to end(s) of the tension cable to apply a tension force on the tension cable to force the seal segments to slidably seal the second seal element with the surface of the second component. The abutting planar seal segments can move freely perpendicular and parallel to the second component and can thermally contract/expand as necessary.

An aircraft gas turbine engine includes a fan that in operation moves air through both a core airflow path and a bypass airflow path of the gas turbine engine. The core airflow path and the bypass airflow path converge at a jet nozzle of the gas turbine engine. A method of controlling the gas turbine engine includes detecting, at a controller of the gas turbine engine, a cruise operating condition of the gas turbine engine, and in response to detecting the cruise operating condition, operating a mechanism of the gas turbine engine to decrease an effective area of the jet nozzle.

A gas turbine engine for an aircraft includes an outer bypass section wall, and a jet nozzle including at least one inflatable diaphragm. The at least one inflatable diaphragm is disposed along the outer bypass section wall. The gas turbine engine also includes a fluid pressure sensor configured to measure a fluid pressure within the at least one inflatable diaphragm, an inlet valve configured to control a pressurized flow of a fluid into the at least one inflatable diaphragm in response to a command from a controller, and a release valve configured to control a release of the fluid from within the at least one inflatable diaphragm in response to a command from the controller.

A gas turbine engine includes a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order. The turbomachine defines an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct. The primary fan is driven by the turbomachine, and a secondary fan is located downstream of the primary fan within the inlet duct. One or more actuation devices operably associated with the fan duct, the one or more actuation devices actuable to increase or decrease an exit area of the fan duct.

An apparatus is provided for an aircraft engine. This apparatus includes a variable area nozzle. The variable area nozzle includes a nozzle wall, a nozzle sleeve, an actuation system and a flowpath extending axially along an axis through the variable area nozzle and radially between the nozzle wall and the nozzle sleeve. The nozzle sleeve includes a shell, a liner, a cooling cavity and an ejector. The shell extends axially along and circumferentially about the axis. The liner axially overlaps and circumscribes the shell. The cooling cavity is formed by and radially between the shell and the liner. The ejector is arranged at a downstream end of the liner along the flowpath. The ejector fluidly couples the cooling cavity to the flowpath. The actuation system is configured to move the nozzle sleeve axially along the axis relative to the nozzle wall.