A gas turbine engine according to the present disclosure includes, among other things, a propulsor section including a propulsor having a plurality of blades, the plurality of blades having a peak tip radius Rt and an inboard leading edge radius Rh at a first inboard boundary of a first flowpath, and a core engine including a first turbine that drives a first compressor and a second turbine that drives the propulsor section. A second inboard boundary of a core flowpath has a radius R1 defined at a first stage of a second compressor and has a radius R2 defined at a splitter rim that guides flow into the core flowpath.

A gas turbine engine according to the present disclosure includes, among other things, a propulsor section including a propulsor having a plurality of blades, the plurality of blades having a peak tip radius Rt and an inboard leading edge radius Rh at a first inboard boundary of a first flowpath, and a core engine including a first turbine that drives a first compressor and a second turbine that drives the propulsor section. A second inboard boundary of a core flowpath has a radius R1 defined at a first stage of a second compressor and has a radius R2 defined at a splitter rim that guides flow into the core flowpath.

An exhaust nozzle assembly for a gas turbine engine. The assembly includes concentrically arranged inner mixer and outer exhaust nozzles, the exhaust nozzle extending axially downstream of said mixer nozzle. A centre-body is axially mounted within and extends axially downstream from the mixer nozzle. A core flow duct is defined by the mixer nozzle and the centre-body, the core flow duct having a core exit area. An exhaust duct is defined at least in part by the exhaust nozzle downstream of the mixer nozzle, the exhaust duct having an exhaust exit area. The mixer nozzle includes a mixer cowl which is axially-translatable along the centre axis and the exhaust nozzle includes an exhaust cowl which is either axially-translatable along or angularly-adjustable relative to the centre axis. The assembly further includes an actuation mechanism and the mixer cowl and exhaust cowl are movable by the actuation mechanism.

An exhaust nozzle assembly for a gas turbine engine. The assembly includes concentrically arranged inner mixer and outer exhaust nozzles, the exhaust nozzle extending axially downstream of said mixer nozzle. A centre-body is axially mounted within and extends axially downstream from the mixer nozzle. A core flow duct is defined by the mixer nozzle and the centre-body, the core flow duct having a core exit area. An exhaust duct is defined at least in part by the exhaust nozzle downstream of the mixer nozzle, the exhaust duct having an exhaust exit area. The mixer nozzle includes a mixer cowl which is axially-translatable along the centre axis and the exhaust nozzle includes an exhaust cowl which is either axially-translatable along or angularly-adjustable relative to the centre axis. The assembly further includes an actuation mechanism and the mixer cowl and exhaust cowl are movable by the actuation mechanism.

A gas turbine engine includes a core engine that has at least a compressor section, a combustor section and a turbine section disposed along a central axis. A fan is coupled to be driven by the turbine section. A fan nozzle is aft of the fan and defines an exit area. The fan nozzle has a body with an airfoil cross-section geometry. The body includes a wall that has a controlled mechanical property distribution that varies in material macro- or micro-structure by location on the wall in accordance with a desired flutter characteristic at the location.

A gas turbine engine includes a gear assembly, a bypass flow passage, a fan located upstream of the bypass flow passage, a first shaft and a second shaft, a first turbine coupled through the gear assembly to the fan, a first compressor coupled with the first shaft, and a second turbine coupled with the second shaft. The fan has a bypass ratio of greater than 8.5. The fan includes a hub and a row of fan blades that extend from the hub. The row includes a number (N) of the fan blades that is from 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to 20.

A gas turbine engine includes a gear assembly, a bypass flow passage, a fan located upstream of the bypass flow passage, a first shaft and a second shaft, a first turbine coupled through the gear assembly to the fan, a first compressor coupled with the first shaft, and a second turbine coupled with the second shaft. The fan has a bypass ratio of greater than 8.5. The fan includes a hub and a row of fan blades that extend from the hub. The row includes a number (N) of the fan blades that is from 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to 20.

A combined cycle propulsion system for a flight vehicle includes a compressor-fed combustion engine, and a multi-mode supersonic engine. The multi-mode supersonic engine includes an adjustable inlet section, a combustion section arranged downstream of the adjustable inlet section and including a first combustor portion having at least one rotating detonation combustor and a second combustor portion having a supersonic combustion type combustor, and an adjustable exhaust nozzle section arranged downstream of the combustion section. The at least one rotating detonation combustor functions as a pilot for the supersonic combustion type combustor.

A combined cycle propulsion system for a flight vehicle includes a compressor-fed combustion engine, and a multi-mode supersonic engine. The multi-mode supersonic engine includes an adjustable inlet section, a combustion section arranged downstream of the adjustable inlet section and including a first combustor portion having at least one rotating detonation combustor and a second combustor portion having a supersonic combustion type combustor, and an adjustable exhaust nozzle section arranged downstream of the combustion section. The at least one rotating detonation combustor functions as a pilot for the supersonic combustion type combustor.

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.