Abstract: In an aspect of the present disclosure, an aircraft defining a longitudinal direction and a lateral direction is provided. The aircraft includes: a fuselage; a single unducted rotor engine mounted at a location spaced from the fuselage of the aircraft, the single unducted rotor engine comprising an unducted rotor assembly having a single stage of rotor blades; and a fuselage shield attached to or formed integrally with the fuselage at a location in alignment with the single stage of rotor blades of the unducted rotor assembly along the lateral direction.

Abstract: A fan for a gas turbine engine is provided. The fan includes a plurality of fan blades, a disk, and a trunnion mechanism for attaching the fan blades to the disk. The disk can be formed of a plurality of individual disk segments, with the trunnion mechanism attaching one of the plurality of fan blades to a respective disk segment. A retention member is also provided. The retention member includes a means for catching a portion of the trunnion mechanism should a primary attachment system of the trunnion mechanism fail.

Abstract: A fan for a gas turbine engine is provided. The fan includes a plurality of fan blades, a disk, and a trunnion mechanism for attaching the fan blades to the disk. The disk can be formed of a plurality of individual disk segments, with the trunnion mechanism attaching one of the plurality of fan blades to a respective disk segment. A retention member is also provided. The retention member includes a means for catching a portion of the trunnion mechanism should a primary attachment system of the trunnion mechanism fail.

Abstract: A method and system for distributed power generation is provided. The method includes determining an operational cycle for the system; determining an average energy requirement of the system based on the operational cycle; configuring a plurality of energy sources each corresponding to the load device to produce a peak efficiency corresponding to the average energy requirement of the system; and coupling the energy source to provide energy to the load device.

Abstract: A gas turbine engine includes a turbomachine including a compressor section, a combustion section, a turbine section, and an exhaust section arranged in serial flow order and together defining at least in part a core air flowpath. The gas turbine engine also includes a thermal management system including a flowpath heat exchanger coupled to, or integrated into, one or more components of the compressor section, the combustion section, the turbine section, or the exhaust section such that the flowpath heat exchanger is directly thermally coupled to an airflow through the core air flowpath.

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 thermal management system includes a first heat source assembly including a first heat source exchanger, a first thermal fluid inlet line extending to the first heat source exchanger, and a first thermal fluid outlet line extending from the first heat source exchanger; a second heat source assembly including a second heat source exchanger, a second thermal fluid inlet line extending to the second heat source exchanger, and second a thermal fluid outlet line extending from the second heat source exchanger; a shared assembly including a thermal fluid line and a heat sink exchanger, the shared assembly defining an upstream junction in fluid communication with the first thermal fluid outlet line and second thermal fluid outlet line and a downstream junction in fluid communication with the first thermal fluid inlet line and second thermal fluid inlet line; and a controller configured to selectively fluidly connect the first heat source assembly or the second heat source assembly to the shared assembly.

Abstract: A thermal management system is provided for incorporation into at least one of a gas turbine engine or an aircraft. The thermal management system includes: a thermal transport bus having a heat exchange fluid flowing therethrough, the thermal transport bus further including: a first flow loop comprising at least one first heat source exchanger, at least one first heat sink exchanger, and a first pump to move the heat exchange fluid through the first flow loop; and a second flow loop comprising at least one second heat source exchanger, at least one second heat sink exchanger, and a second pump to move the heat exchange fluid through the second flow loop; wherein the first flow loop is isolated from the second flow loop, and wherein the first heat source exchanger and the second heat source exchanger are configured with a common heat source.

Abstract: A turbomachine includes a turbine section including a turbine. The turbine includes a first plurality of turbine rotor blades and a second plurality of turbine rotor blades, the first plurality of turbine rotor blades and second plurality of turbine rotor blades alternatingly spaced along the axial direction. The turbomachine also includes a gearbox. The first plurality of turbine rotor blades and the second plurality of turbine rotor blades are each coupled to one of a ring gear, a planet gear, or a sun gear of the gearbox such that the first plurality of turbine rotor blades is rotatable with the second plurality of turbine rotor blades through the gearbox. The turbomachine also includes an electric machine assembly including a rotor coupled to one of the ring gear, the planet gear, or the sun gear of the gearbox such that the rotor rotates relative to a stator during operation.

Abstract: A propulsion system according to aspects of the present disclosure is provided, the propulsion system including a rotor assembly with a plurality of blades extended radially relative to the engine centerline axis, and a vane assembly positioned in aerodynamic relationship with the rotor assembly. The vane assembly includes a plurality of vanes extended radially relative to the engine centerline axis, and the propulsion system includes a ratio of a quantity of blades to a quantity of vanes between 2:5 and 2:1.

Abstract: A propulsion system, the propulsion system comprising a rotating element, a stationary element, and an inlet between the rotating element and the stationary element, wherein the inlet passes radially inward of the stationary element; wherein the inlet passes radially inward of the stationary element; wherein the inlet leads to an inlet duct containing a ducted fan having an axis of rotation and a plurality of blades; and wherein the inlet duct divides into a first duct and a second duct, separate from the first duct. A method of operating a propulsion system, comprising the steps of: operating a first rotating fan assembly to produce a first stream of air; directing a portion of the first stream of air into a second ducted rotating fan assembly; operating the second ducted rotating fan assembly to produce a second stream of air; dividing the second stream of air into a core stream and a fan stream; and directing the core stream into a gas turbine engine core.

Abstract: A gas turbine engine includes a gas turbine engine core having a high pressure compressor, a combustor, and a high pressure turbine in serial relationship; and a low pressure compressor upstream of the gas turbine engine core; wherein the low pressure compressor driven by a variable speed power source such that the rotational speed of the low pressure compressor is controllable independently from the rotational speed of any turbine of the gas turbine engine.

Abstract: A gas turbine engine includes a fan section having a fan rotatable with a fan shaft and a turbomachinery section having a turbine and a turbomachine shaft rotatable with the turbine. A power gearbox is also provided mechanically coupled to both the fan shaft and the turbomachine shaft such that the fan shaft is rotatable by the turbomachine shaft across the power gearbox. A torque monitoring system includes a gearbox sensor operable with a gear of the power gearbox and a shaft sensor operable with at least one of the turbomachine shaft or the fan shaft. The torque monitoring system determines an angular position of the gear of the gearbox relative to at least one of the fan shaft or the turbomachine shaft using the gearbox sensor and the shaft sensor to determine a torque within the gas turbine engine.

Abstract: In one an exemplary aspect of the present disclosure, an engine includes a drive shaft; an electric machine including a stator assembly and a rotor assembly, the rotor assembly rotatable relative to the stator assembly; and an electrical break, the drive shaft coupled to the rotor assembly through the electrical break.

Abstract: An aeronautical gas turbine engine includes a turbomachine including a compressor section, a combustion section, a turbine section, and an exhaust section in serial flow order. The aeronautical gas turbine engine additionally includes a closed cycle heat engine including a compressor configured to compress a working fluid; a primary heat exchanger in thermal communication with the turbomachine and the working fluid, the primary heat exchanger configured to transfer heat from the turbomachine to the working fluid; an expander coupled to the compressor for expanding the working fluid; and an output shaft driven by the expander.

Abstract: A gas turbine engine including core engine is provided. Air may enter the core engine through an inlet and travel through and engine air flowpath extending through the core engine, e.g., generally along an axial direction of the gas turbine engine. The gas turbine engine additionally includes a cooling air flowpath extending outwardly generally along the radial direction of the gas turbine engine. The cooling air flowpath extends between an inlet in flow communication with engine air flowpath and an outlet defined by an opening in an outer casing of the core engine. Moreover, the gas turbine engine includes a heat exchanger positioned at least partially within the outer casing the core engine with the cooling air flowpath extending over or through the heat exchanger.

Abstract: A gas turbine engine includes a core turbine engine and a fan mechanically coupled to the core turbine engine. The fan includes a plurality of fan blades, each fan blade defining a base and an inner end along a radial direction of the gas turbine engine. The fan also includes a hub covering the base of each of the plurality of fan blades. Further, the fan includes one or more bearings for supporting rotation of the plurality of fan blades. The one or more bearings define a fan bearing radius along a radial direction of the gas turbine engine. Similarly, the hub defines a hub radius along the radial direction of the gas turbine engine. The ratio of the hub radius to the fan bearing radius is less than about three, providing for desired packaging of the various components within the fan of the gas turbine engine.

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: 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.