A gas turbine engine for an aircraft comprising an engine core and including a bypass channel which radially surrounds the engine core at least in part is described. A core shaft is operatively connected to an engine accessory gearbox, which is arranged between the engine core and the bypass channel, by means of a radial shaft of a drive train. An electric machine is provided which is designed to start the gas turbine engine during motor operation and to generate electrical energy during alternator operation. The electric machine is arranged coaxially with the core shaft and connected thereto for conjoint rotation. Alternatively, the electric machine can be arranged radially outside the bypass channel and can be operatively connected to the core shaft by means of the radial shaft, wherein a rotor of the electric machine is arranged coaxially with the radial shaft and connected thereto for conjoint rotation.

A gas turbine engine including an engine core and including a bypass channel of an aircraft is described. The bypass channel radially surrounds the engine core at least in part. At least one core shaft extending in the axial direction is provided, which shaft is operatively connected, by means of a drive train, to a core accessory gearbox arranged between the engine core and the bypass channel, and to an aircraft accessory gearbox. The aircraft accessory gearbox is arranged radially outside the bypass channel. The drive train extends substantially in the radial direction between the core shaft and the aircraft accessory gearbox. The drive train has an angle drive between a radial shaft and the core shaft. Furthermore, the drive train comprises another angle drive between the radial shaft and the core accessory gearbox and an additional angle drive between the radial shaft and the aircraft accessory gearbox.

A gas turbine engine including an engine core and including a bypass channel of an aircraft is described. The bypass channel radially surrounds the engine core at least in part. At least one core shaft extending in the axial direction is provided, which shaft is operatively connected, by means of a drive train, to a core accessory gearbox arranged between the engine core and the bypass channel, and to an aircraft accessory gearbox. The aircraft accessory gearbox is arranged radially outside the bypass channel. The drive train extends substantially in the radial direction between the core shaft and the aircraft accessory gearbox. The drive train has an angle drive between a radial shaft and the core shaft. Furthermore, the drive train comprises another angle drive between the radial shaft and the core accessory gearbox and an additional angle drive between the radial shaft and the aircraft accessory gearbox.

A propulsion device with double-layer flow guiding assembly and a flight vehicle using the same are provided. The propulsion device includes a propulsion body, a first-layer flow guiding assembly and a second-layer flow guiding assembly. The propulsion body includes a housing, an airflow suction port and an airflow discharge port. The first-layer flow guiding assembly includes a front flow guiding ring and at least one first-layer flow guiding plate. The front flow guiding ring is disposed outside the airflow discharge port and has a first axis. The front flow guiding ring swings relative to the airflow discharge port along a first rotation axis. The first rotation axis intersects the first axis. The first-layer flow guiding plate is fixed in the front flow guiding ring and extends along the first rotation axis. The second-layer flow guiding assembly has a structure similar to the first-layer flow guiding assembly.

A propeller-less unmanned aerial vehicle having a body having a plurality of channels, an inlet formed in the body and configured to allow air flow to enter the plurality of channels from an exterior of the body, an anechoic chamber formed in the body and coupled to the plurality of channels, a rotor comprising a plurality of angled fins located in the anechoic chamber, a control system configured to direct air flow within the plurality of channels, and one or more circular tubes coupled to the exterior of the body and in communication with the plurality of channels. The air flows into the body through the inlet, into the plurality of channels and the anechoic chamber, and exits through the one or more circular tubes to provide lift and directional control to the propeller-less unmanned aerial vehicle.

A propeller-less unmanned aerial vehicle having a body having a plurality of channels, an inlet formed in the body and configured to allow air flow to enter the plurality of channels from an exterior of the body, an anechoic chamber formed in the body and coupled to the plurality of channels, a rotor comprising a plurality of angled fins located in the anechoic chamber, a control system configured to direct air flow within the plurality of channels, and one or more circular tubes coupled to the exterior of the body and in communication with the plurality of channels. The air flows into the body through the inlet, into the plurality of channels and the anechoic chamber, and exits through the one or more circular tubes to provide lift and directional control to the propeller-less unmanned aerial vehicle.

Systems and methods include providing an aircraft with a fuselage and a convertible engine disposed within the fuselage. The convertible engine is operable as a turbofan engine in a thrust mode and a turboshaft engine in a shaft power mode. The convertible engine includes a housing, an engine core having a low pressure turbine shaft, and a bypass fan system. The bypass fan system includes a bypass fan having a fan clutch. The fan clutch selectively couples at least a portion of the bypass fan to the low pressure turbine shaft when the convertible engine is operated in the thrust mode and decouples at least a portion of the bypass fan from the low pressure turbine shaft when the convertible engine is operated in the shaft power mode.

Systems and methods include providing an aircraft with a fuselage and a convertible engine disposed within the fuselage. The convertible engine is operable as a turbofan engine in a thrust mode and a turboshaft engine in a shaft power mode. The convertible engine includes a housing, an engine core having a low pressure turbine shaft, and a bypass fan system. The bypass fan system includes a bypass fan having a fan clutch. The fan clutch selectively couples at least a portion of the bypass fan to the low pressure turbine shaft when the convertible engine is operated in the thrust mode and decouples at least a portion of the bypass fan from the low pressure turbine shaft when the convertible engine is operated in the shaft power mode.

A motor-integrated fan having an intake port and a blow-out port comprises a rotary part 12 that is rotatably supported by a shaft, and a motor 14 that rotates the rotary part 12. The motor 14 is an outer periphery drive motor that rotates the rotary part 12 by supplying motive power from a duct provided on the outer peripheral side of the shaft. The motor 14 includes: a permanent magnet 45 provided on a rotary support ring 33 connected to the outer peripheral side of blades 32 of the rotary part 12; and a coil 46 provided opposite the permanent magnet 45 in the axial direction of the axis of rotation. Among the plurality of blades 32, a first blade 32a and a second blade 32b are located at different positions in the axial direction.

An aerospace vehicle including: a body, wherein the body is configured to generate heat during operation; a coating disposed over at least a portion of the body, the coating being configured to shift a frequency of at least one wavelength of the heat generated by the body from a first frequency to a second frequency having higher transmissivity relative to a neighboring medium surrounding the body as compared to the first frequency.