Abstract: The present disclosure is directed to a gas turbine engine defining a radial direction, a circumferential direction, an axial centerline along a longitudinal direction, and an upstream end and a downstream end along the longitudinal direction. The gas turbine engine defines a core flowpath extended generally along the longitudinal direction. The gas turbine engine includes a first turbine rotor. The first turbine rotor includes an annular outer band disposed outward of the core flowpath along the radial direction. The first turbine rotor further includes a plurality of airfoils coupled to an inner diameter of the outer band in which the plurality of airfoils are extended generally inward along the radial direction. The outer band defines a plurality of airfoil cooling passages in which the plurality of airfoil cooling passages are extended at least partially in the radial direction in fluid communication with the plurality of airfoils.

Abstract: 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.

Abstract: An electric machine is provided which includes a rotor disk extending along a radial direction and having a rotor flange attached to or formed integrally with the rotor disk and extending substantially along the axial direction. A plurality of rotor magnets are mounted on the rotor disk and supported by the rotor flange such that the rotor flange absorbs centrifugal loads exerted on the magnets and enables high speeds of operation.

Abstract: A thermal management system for a gas turbine engine and/or an aircraft is provided including a thermal transport bus having a heat exchange fluid flowing therethrough. The thermal management system also includes one or more heat source exchangers and a deicing module. The one or more heat source exchangers and the deicing module are each in thermal communication with the heat exchange fluid in the thermal transport bus. The one or more heat source exchangers are configured to transfer heat from one or more accessory systems to the heat exchange fluid, and the deicing module is located downstream of the one or more heat source exchangers for transferring heat from the thermal transfer fluid to a surface of one or more components of the gas turbine engine and/or the aircraft.

Abstract: Systems and methods for adjusting airflow distortion in a gas turbine engine using a translating inlet assembly are provided. In one embodiment, a core engine of a gas turbine engine can include a compressor section, a combustion section, and a turbine section in series flow and defining at least in part an engine airflow path. The compressor section can include an inner flowpath surface. A core casing can enclose the core engine. A forward end of the core casing can include a translating inlet assembly moveable between a first position and a second position. The translating inlet assembly and the inner flowpath surface can together define an inlet to an engine airflow path. A translating inlet assembly can define a first inlet area in the first position and a second inlet area in the second position, the first inlet area being greater than the second inlet area.

Abstract: A planet gear train for an engine casing includes a sun gear rotatable by a shaft, a ring gear, a plurality of planet gears rotatably mounted in a planet carrier and meshing with the sun gear and the ring gear, and a damping system. The damping system includes a soft mount disposed about the shaft and coupling the planet carrier the engine casing. The soft mount includes a flexible inner sleeve and a flexible outer sleeve. The flexible inner sleeve and the flexible outer sleeve are fixedly coupled together at respective central portions thereof near the shaft. The damping system further includes a plurality of damping couplers disposed about an outer flange of the planet carrier. The plurality of damping couplers is configured to flexibly couple the flexible inner sleeve to the flexible outer sleeve.

Abstract: A gas turbine engine includes a compressor section and a turbine section. The turbine section includes a drive turbine and is located downstream of the compressor section. The gas turbine engine also includes a fan mechanically coupled to and rotatable with the drive turbine such that the fan is rotatable by the drive turbine at the same rotational speed as the drive turbine, the fan defining a fan pressure ratio and including a plurality of fan blades, each fan blade defining a fan tip speed. During operation of the gas turbine engine at a rated speed, the fan pressure ratio of the fan is less than 1.5 and the fan tip speed of each of the fan blades is greater than 1,250 feet per second.

Abstract: A propulsion system for an aircraft having an aft end is provided herein. The propulsion system can include an electric propulsion engine defining a central axis. The electric propulsion engine can include an electric motor, a fan rotatable about the central axis of the electric propulsion engine by the electric motor, a bearing supporting rotation of the fan, and a thermal management system. The thermal management system can include a lubrication oil circulation assembly for providing the bearing with lubrication oil. The lubrication oil circulation assembly can be driven independently of a shaft of the electric propulsion engine.

Abstract: A propulsion system for an aircraft includes a gas turbine engine and an electric propulsion engine defining a central axis. The electric propulsion engine includes an electric motor and a fan rotatable about the central axis of the electric propulsion engine by the electric motor. The electric propulsion engine additionally includes a bearing supporting rotation of the fan and a thermal management system including a thermal fluid circulation assembly. The thermal fluid circulation assembly is in thermal communication with at least one of the electric motor or the bearing and is further in thermal communication with a heat exchanger of a thermal management system of the gas turbine engine.

Abstract: Instrumented airflow path components configured to determine airflow path distortion in an airflow path of a gas turbine engine (e.g., using for propulsion of an aircraft) are provided. In one embodiment, a gas turbine engine for an aircraft can include a compressor section, a combustion section, and turbine section in series flow. The compressor section, combustion section, and turbine section define at a portion of an engine airflow path for the gas turbine engine. The gas turbine engine further includes one or more members extending at least partially into the engine airflow path of the gas turbine engine and one or more pressure sensor devices at least partially integrated into the one or more members extending at least partially into the engine airflow path. The one or more pressure sensor devices are configured to obtain measurements for determining a distortion condition for the gas turbine engine.

Abstract: A casing for use in a turbofan engine is provided. The casing includes an inner hub, an intermediate casing positioned radially outward from the inner hub, an outer casing positioned radially outward from the intermediate casing, and a plurality of struts spaced circumferentially about the inner hub and extending between the inner hub and the outer casing. At least one strut of the plurality of struts includes a flow channel extending therethrough. The casing also includes a fluid scoop defined within the inner hub, and a collector that couples the fluid scoop in flow communication with the flow channel.

Abstract: Waterproof apparatus for cables and cable interfaces are provided herein. An exemplary apparatus includes a coupler body that includes a first end configured to releaseably couple with a connector bulkhead and a second end having an opening that is sized to receive a sealing gland, a cavity for receiving the sealing gland, the sealing gland comprising an outer peripheral surface configured to sealingly engage with an inner surface of the cavity, the sealing gland comprising an aperture that is configured to receive a cable.

Abstract: A computer-implemented method of assessing a health of a gas turbine engine of a hybrid-electric propulsion system for an aircraft includes receiving, by one or more computing devices, data indicative of an amount of electrical power provided to, or extracted from, an electric machine. The method also includes receiving, by the one or more computing devices, data indicative of an operating parameter of the hybrid-electric propulsion system. The method also includes assessing, by the one or more computing devices, a health of the gas turbine engine based at least in part on the received data indicative of the amount of electrical power provided to, or extracted from, the electric machine and the received data indicative of the operating parameter of the hybrid electric propulsion system. The method also includes providing, by the one or more computing devices, information to a user indicative of the health of the gas turbine engine.

Abstract: A fuel system is provided for an aircraft having a fuel source. The fuel system includes a fuel oxygen reduction unit defining a liquid fuel supply path, a stripping gas supply path, a liquid fuel outlet path, and a stripping gas return path, wherein the stripping gas return path is in airflow communication with the fuel source.

Abstract: Systems and methods for adjusting airflow distortion in a gas turbine engine using a translating inlet assembly are provided. In one embodiment, a core engine of a gas turbine engine can include a compressor section, a combustion section, and a turbine section in series flow and defining at least in part an engine airflow path. The compressor section can include an inner flowpath surface. A core casing can enclose the core engine. A forward end of the core casing can include a translating inlet assembly moveable between a first position and a second position. The translating inlet assembly and the inner flowpath surface can together define an inlet to an engine airflow path. A translating inlet assembly can define a first inlet area in the first position and a second inlet area in the second position, the first inlet area being greater than the second inlet area.

Abstract: A gas turbine engine includes a turbomachine and a fan rotatable by the turbomachine. The fan includes a plurality of fan blades. The gas turbine engine also includes an outer nacelle surrounding the plurality of fan blades and a plurality of part-span inlet guide vanes cantilevered from the outer nacelle at a location forward of the plurality of fan blades along the axial direction. Each of the plurality of inlet guide vanes defines an inner end along the radial direction and it is unconnected with an adjacent part-span inlet guide vane at the inner end.

Abstract: A gas turbine engine includes a turbomachine and a fan rotatable by the turbomachine. The fan includes a plurality of fan blades, each of the plurality of fan blades defining a fan blade span along a radial direction. The gas turbine engine further includes an outer nacelle surrounding the plurality of fan blades and a plurality of part-span inlet guide varies attached to the outer nacelle at a location forward of the plurality of fan blades along an axial direction. Each of the plurality of inlet guide vanes defines an IGV span along the radial direction, the IGV span being at least about five percent of the fan blade span and up to about fifty-five percent of the fan blade span.

Abstract: A turbine engine has a core with a compressor, combustor, and turbine sections in axial flow arrangement and with corresponding rotating elements mounted to a shaft defining engine rotor elements. The turbine engine has a rotary driver operably coupled to the engine rotor elements. The turbine engine has at least one thermoelectric generator in thermal communication with the core and in electrical communication with the rotary driver to provide power to the rotary driver to turn the engine rotor elements.

Abstract: An aspect of the present disclosure is directed to a rotor assembly for a turbine engine. The rotor assembly includes an airfoil assembly and a hub to which the airfoil assembly is attached. A wall assembly defines a first cavity and a second cavity between the airfoil assembly and the hub. The first cavity and the second cavity are at least partially fluidly separated by the wall assembly. The first cavity is in fluid communication with a flow of first cooling fluid and the second cavity is in fluid communication with a flow of second cooling fluid different from the first cooling fluid.

Abstract: A propulsion system for an aircraft includes an electric propulsion engine. The electric propulsion engine includes an electric motor and a fan rotatable about a central axis of the electric propulsion engine by the electric motor. The electric propulsion engine also includes a bearing supporting rotation of the fan and a thermal management system. The thermal management system includes a lubrication oil circulation assembly and a heat exchanger thermally connected to the lubrication oil circulation assembly. The lubrication oil circulation assembly is configured for providing the bearing with lubrication oil. Such an electric propulsion engine may be a relatively self-sufficient engine.