A gas turbine engine includes an inner annular combustor radially inboard of an outer annular combustor. An outer variable turbine vane array is downstream of the outer annular combustor and an inner variable turbine vane array downstream of the inner annular combustor.

A turbine engine structure includes a turbine engine core having a core cocooning feature, and a fan fore of the turbine engine core, relative to fluid flow through the turbine engine structure. The fan is drivably connected to the turbine engine core via a shaft. A nacelle circumferentially surrounds the turbine engine core, and a bypass flowpath is defined between the turbine engine core and the nacelle. A plurality of augmenter fuel spray bars are disposed in the bypass flowpath.

A turbine engine structure includes a turbine engine core having a core cocooning feature, and a fan fore of the turbine engine core, relative to fluid flow through the turbine engine structure. The fan is drivably connected to the turbine engine core via a shaft. A nacelle circumferentially surrounds the turbine engine core, and a bypass flowpath is defined between the turbine engine core and the nacelle. A plurality of augmenter fuel spray bars are disposed in the bypass flowpath.

A turbine engine structure includes a turbine engine core having a core cocooning feature, and a fan fore of the turbine engine core, relative to fluid flow through the turbine engine structure. The fan is drivably connected to the turbine engine core via a shaft. A nacelle circumferentially surrounds the turbine engine core, and a bypass flowpath is defined between the turbine engine core and the nacelle. A plurality of augmenter fuel spray bars are disposed in the bypass flowpath.

A hypersonic propulsion engine includes: a turbine engine including a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbine engine defining a turbine engine inlet upstream of the compressor section and a turbine engine exhaust downstream of the turbine section; a ducting assembly defining a bypass duct having a substantially annular shape and extending around the turbine engine, an afterburning chamber located downstream of the bypass duct and at least partially aft of the turbine engine exhaust, and an inlet section located at least partially forward of the bypass duct and the turbine engine inlet; and an inlet precooler positioned at least partially within the inlet section of the ducting assembly and upstream of the turbine engine inlet, the bypass duct, or both for cooling an airflow provided through the inlet section of the ducting assembly to the turbine engine inlet, the bypass duct, or both.

A gas turbine engine may include a high pressure compressor coupled to a high pressure turbine by a high pressure shaft, a core combustor located downstream of the high pressure compressor and upstream of the high pressure turbine, and a low pressure compressor provided upstream of the high pressure compressor. The low pressure compressor may be configured to direct core airflow to the high pressure compressor and first bypass airflow which bypasses the high pressure compressor, core combustor and high pressure turbine through a first bypass duct. The engine may further include a mixer downstream of the high pressure turbine and low pressure compressor, the mixer being configured to mix the core and first bypass airflows. The engine also may include a re-heat combustor configured to combust fuel with both core airflow and first bypass airflow. A low pressure turbine may be provided downstream of the re-heat combustor and coupled to the low pressure compressor (14) by a low pressure shaft, the low pressure and high pressure shafts being independently rotatable. A shaft power transfer arrangement may be provided, which is configured to selectively transfer power between the low pressure and high pressure shafts.

A gas turbine engine may include a high pressure compressor coupled to a high pressure turbine by a high pressure shaft, a core combustor located downstream of the high pressure compressor and upstream of the high pressure turbine, and a low pressure compressor provided upstream of the high pressure compressor. The low pressure compressor may be configured to direct core airflow to the high pressure compressor and first bypass airflow which bypasses the high pressure compressor, core combustor and high pressure turbine through a first bypass duct. The engine may further include a mixer downstream of the high pressure turbine and low pressure compressor, the mixer being configured to mix the core and first bypass airflows. The engine also may include a re-heat combustor configured to combust fuel with both core airflow and first bypass airflow. A low pressure turbine may be provided downstream of the re-heat combustor and coupled to the low pressure compressor (14) by a low pressure shaft, the low pressure and high pressure shafts being independently rotatable. A shaft power transfer arrangement may be provided, which is configured to selectively transfer power between the low pressure and high pressure shafts.

An aircraft engine includes an outer casing, an inner casing, a first rotating shaft, a first fan, a second fan, a first combustor, and a second combustor. The outer casing has an outer intake end and an outer exhaust end. The inner casing includes a first section and a second section arranged in an axial direction. The first rotating shaft is rotatably disposed in the inner casing. The first fan includes an inner turbine portion, a first connection portion, and an outer turbine portion. The second fan is connected to the first rotating shaft and disposed on the outer intake end side of the first section wherein an outer diameter of the second fan is larger than that of the first section. The first combustor is disposed between the second fan and the outer turbine portion. The second combustor is disposed between the second fan and the inner turbine portion.

A cooling system for a turbine engine including a heat exchanger in fluid communication with a first fluid inlet stream and disposed upstream and in fluid communication with a core engine. The heat exchanger operative to cool the first fluid inlet stream. The heat exchanger including a heat exchanger inlet for input of a heat exchanging medium for exchange of heat from the first fluid inlet stream to the heat exchanging medium. The heat exchanger further including a heat exchanger outlet for discharge of a heated output stream into one of a turbine of a downstream engine, an augmentor or a combustor of the core engine. The heated output stream provides an additional flow to the downstream engine. A turbine engine including the cooling system is disclosed.

A cooling system for a turbine engine including a heat exchanger in fluid communication with a first fluid inlet stream and disposed upstream and in fluid communication with a core engine. The heat exchanger operative to cool the first fluid inlet stream. The heat exchanger including a heat exchanger inlet for input of a heat exchanging medium for exchange of heat from the first fluid inlet stream to the heat exchanging medium. The heat exchanger further including a heat exchanger outlet for discharge of a heated output stream into one of a turbine of a downstream engine, an augmentor or a combustor of the core engine. The heated output stream provides an additional flow to the downstream engine. A turbine engine including the cooling system is disclosed.