One embodiment includes a multistage infrared suppression exhaust system for an aircraft, including: a stage one including a first exhaust conduit to receive a first exhaust air flow at a first temperature-pressure product T.sub.1P.sub.1, a second exhaust conduit to receive a second exhaust air flow at a second temperature-pressure product T.sub.2P.sub.2, and a flow integrator mechanically configured to mix the first exhaust air flow with the second exhaust air flow in an integration chamber while preventing back flow into the second exhaust conduit; and a stage two including a stage two cooling airflow to cool the mixed first and second exhaust air flows.

Systems and methods for suppressing infrared radiation generated by a turbine engine. A system comprises a primary assembly having a center body, a plurality of vanes extending from the center body, an outer radial duct with the plurality of vanes extending therethrough, a structural baffle, and a mixer. The primary assembly is disposed in the exhaust flow path of a turbine engine and encased in ducting and/or an airfoil. An air flow path defined between the center body and outer radial duct is axially spit by an interface rim and flow segregator. The flow segregator segregates engine core flow from ambient air flow.

One embodiment includes a multistage infrared suppression exhaust system for an aircraft, including: a stage one including a first exhaust conduit to receive a first exhaust air flow at a first temperature-pressure product T.sub.1P.sub.1, a second exhaust conduit to receive a second exhaust air flow at a second temperature-pressure product T.sub.2P.sub.2, and a flow integrator mechanically configured to mix the first exhaust air flow with the second exhaust air flow in an integration chamber while preventing back pressure into the second exhaust conduit; and a stage two including a stage two cooling airflow to cool the mixed first and second exhaust air flows.

The invention concerns an aircraft propulsion control system in which a gas turbine engine has an actuable flow opening for control of flow to or from a portion of the engine. One or more sensor is arranged to sense a condition indicative of vapor trail formation by the exhaust flow from the engine. A controller is arranged to control actuation of the flow opening so as to reduce the efficiency of the engine upon sensing of said condition by the one or more sensor. In one example, the flow opening is a variable area fan nozzle.

A line of sight blocker including a cover defining a duct bounded by the cover and a heated surface when the cover is fastened over the heated surface. The cover blocks transmission of infrared radiation emitted from the heated surface; the cover comprises a material having a lower thermal conductivity than the heated surface; and the duct comprises a vent and a path for latent heat from the heated surface to escape through the vent.

A gas turbine engine includes a main gas flow exhaust nozzle having an annular inner surface which, in use, bounds a flow of exhaust gas. The gas turbine engine further includes cooling passages having respective outlets therefrom to provide a flow of cooling air over a surface of the engine or an adjacent airframe component, thereby protecting the cooled surface from the exhaust gas flow. Adjacent cooling passages of the or each pair of the nested cooling passages are separated from each other by a respective dividing wall. The outlets from the nested cooling passages are staggered in the axial direction of the exhaust nozzle such that cooling air flowing out of an inner one of the adjacent cooling passages of the or each pair of the nested cooling passages flows over the dividing wall separating the adjacent passages.

A line of sight blocker including a cover defining a duct bounded by the cover and a heated surface when the cover is fastened over the heated surface. The cover blocks transmission of infrared radiation emitted from the heated surface; the cover comprises a material having a lower thermal conductivity than the heated surface; and the duct comprises a vent and a path for latent heat from the heated surface to escape through the vent.

Devices, systems, and methods of a casing for a heat suppression system of a gas turbine engine exhaust include a heat shield and an insulation layer for arrangement between the heat shield and an outer skin.

A cooling system for an aircraft has at least one moveable member configured to cover an opening formed within an aircraft outer skin. An actuator moves the at least one moveable member between a fully open position where external atmosphere air can be directed through the opening to an internal passage enclosed by the aircraft outer skin and a fully closed position where the opening is covered. A controller selectively controls the actuator to move the at least one moveable member between the fully open and fully closed positions. An aircraft engine and a method of cooling an aircraft engine in an aircraft are also disclosed.

A seal assembly can accommodate deflection between two components through a cartridge that is slidingly engaged with a slide plate.