Abstract: A gas turbine engine includes a compressor section and a turbine section together defining a core air flowpath. Additionally, a rotary component is rotatable with at least a portion of the compressor section and at least a portion of the turbine section. An electric machine is mounted coaxially with the rotary component and positioned at least partially inward of the core air flowpath along a radial direction of the gas turbine engine. A cavity wall defines at least in part a buffer cavity surrounding at least a portion of the electric machine to thermally insulate the electric machine, e.g., from the relatively high temperatures within the core air flowpath.

Abstract: A method is provided of controlling a cooled cooling air system for an aeronautical gas turbine engine. The method includes: receiving data indicative of an ambient condition of the aeronautical gas turbine engine, data indicative of a deterioration parameter of the aeronautical gas turbine engine, data indicative of an operating condition of the aeronautical gas turbine engine, or a combination thereof; and modifying a cooling capacity of the cooled cooling air system in response to the received data indicative of the ambient condition of the aeronautical gas turbine engine, data indicative of the deterioration parameter of the aeronautical gas turbine engine, data indicative of an operating condition of the aeronautical gas turbine engine, or the combination thereof.

Abstract: The present disclosure is directed to a gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction. The gas turbine engine includes a fan assembly including a plurality of fan blades rotatably coupled to a fan rotor in which the fan blades define a maximum fan diameter and a fan pressure ratio. The gas turbine engine further includes a low pressure (LP) turbine defining a core flowpath therethrough generally along the longitudinal direction. The core flowpath defines a maximum outer flowpath diameter relative to the axial centerline. The gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter. The fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater.

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: Propulsion systems and methods of operation are provided. An exemplary propulsion system comprises a rotating element; a stationary element; an inlet duct having an inlet between the rotating and stationary elements, the inlet passing radially inward of the stationary element; a ducted fan disposed in the inlet duct downstream of the inlet and having an axis of rotation and a plurality of blades; a gas turbine engine core having a high pressure compressor, a combustor, and a high pressure turbine in serial relationship; and a booster compressor disposed between the ducted fan and the gas turbine engine core. At least one of the ducted fan and the booster compressor is driven by a variable speed power source such that the rotational speed of the ducted fan and/or booster compressor is controllable independently from the rotational speed of any rotor of the propulsion system.

Abstract: A method for energy conversion for a vehicle is provided. The method including extracting a flow of compressed fluid from a compressor section of a propulsion system; flowing the flow of compressed fluid to a turbine operably coupled to a driveshaft, in which the driveshaft is operably coupled to a load device; expanding the flow of compressed fluid through the turbine to generate an output torque at the driveshaft to operate the load device; and flowing the expanded flow of compressed fluid from the turbine to thermal communication with a thermal load.

Abstract: An aeronautical propulsion device including a fan and a plurality of variable guide vanes is provided. The fan includes a plurality of fan blades for providing a flow of air and the plurality of variable guide vanes are configured for directing air to or from the fan in a desired direction. Each of the plurality of guide vanes defines an inner end along the radial direction and is attached to a housing of the portion device at the inner end in a rotatable manner. The propulsion device further includes a pitch change mechanism positioned in the housing and mechanically coupled to at least one of the plurality of guide vanes for changing a pitch of the at least one of the plurality of guide vanes.

Abstract: A turbine engine includes a compressor section, a combustion section, a turbine section, and an exhaust section in serial flow order and together defining a core air flowpath; a fuel delivery system for providing a flow of fuel to the combustion section; and a waste heat recovery system. The waste heat recovery system includes a heat source exchanger in thermal communication with the turbine section, the exhaust section, or both; a heat sink exchanger in thermal communication with the fuel delivery system, the core air flowpath, or both; a thermal transfer bus including a thermal transfer fluid and extending from the heat source exchanger to the heat sink exchanger; and a pump in fluid communication with the thermal transfer bus downstream of the heat source exchanger and upstream of the heat sink exchanger for increasing a temperature and a pressure of the thermal transfer fluid in the thermal transfer bus.

Abstract: A system for controlling blade clearances within a gas turbine engine includes a rotor disk and a rotor blade coupled to the rotor disk. Additionally, the system includes an outer turbine component positioned outward of the rotor blade such that a clearance is defined between the rotor blade and the outer turbine component. Furthermore, the system includes a heat exchanger configured to receive a flow of cooling air bled from the gas turbine engine and transfer heat from the received flow of the cooling air to a flow of coolant to generate cooled cooling air. Moreover, the system includes a valve configured to control the flow of the coolant to the heat exchanger. In this respect, the cooled cooling air is supplied to at least one of the rotor disk or the rotor blade to adjust the clearance between the rotor blade and the outer turbine component.

Abstract: Systems and methods of operating systems are provided. For example, a system comprises a fuel cooling loop including a cold fuel flowpath having a fuel flowing therethrough, a fuel cooler heat exchanger for cooling the fuel in fluid communication with the cold fuel flowpath, and a cold fuel tank disposed along the cold fuel flowpath for accumulating at least a portion of the cooled fuel. The system further comprises a fuel heating loop including a hot fuel flowpath for a flow of the fuel, a fuel heater heat exchanger for heating the fuel in fluid communication with the hot fuel flowpath, and a hot fuel tank disposed along the hot fuel flowpath for accumulating at least a portion of the heated fuel. The fuel cooling loop is coupled to the fuel heating loop such that the fuel circulates through both the fuel cooling loop and the fuel heating loop.

Abstract: In one exemplary embodiment of the present disclosure, a method of operating a fuel system for an aeronautical gas turbine engine is provided. The method includes: providing a flow of fuel to a fuel nozzle of the aeronautical gas turbine engine during a wind down condition; operating a fuel oxygen reduction unit to reduce an oxygen content of the flow of fuel provided to the fuel nozzle of the aeronautical gas turbine engine during the wind down condition; and ceasing providing the flow of fuel to the fuel nozzle of the aeronautical gas turbine engine, the fuel nozzle comprising a volume of fuel after ceasing providing the flow of fuel to the fuel nozzle; wherein operating the fuel oxygen reduction unit comprises operating the fuel oxygen reduction unit to reduce an oxygen content of the volume of fuel in the fuel nozzle to less than 20 parts per million.

Abstract: A fuel oxygen conversion unit includes a contactor; a fuel gas separator, the fuel oxygen conversion unit defining a circulation gas flowpath from the fuel gas separator to the contactor; and an isolation valve in airflow communication with the circulation gas flowpath for modulating a gas flow through the circulation gas flowpath to the contactor.

Abstract: A system for heat exchange and leak detection is generally provided, the system including a heat exchanger including a first wall defining a first passage containing a first fluid. A leak detection enclosure containing a leak detection medium is defined between the first wall and a second wall surrounding the first wall.

Abstract: A system for heat exchange and leak detection is generally provided, the system including a heat exchanger including a first wall defining a first passage containing a first fluid. A leak detection enclosure containing a leak detection medium is defined between the first wall and a second wall surrounding the first wall.

Abstract: A turbine engine that includes an engine core, an inner cowl, an outer cowl and an accessory gearbox. The engine core includes at least a compressor section, a combustion section, and a turbine section in axial flow arrangement. The accessory gearbox is operably coupled to the engine core and includes a first portion and a second portion.

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 a plurality of heat source exchangers and at least one heat sink exchanger. The plurality of heat source exchangers and the at least one heat sink exchanger are in thermal communication with the heat exchange fluid in the thermal transport bus. The plurality of heat source exchangers are arranged along the thermal transport bus and configured to transfer heat from one or more accessory systems to the heat exchange fluid, and the at least one heat sink exchanger is located downstream of the plurality of heat source exchangers and configured to remove heat from the heat exchange fluid.

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: Provided are systems and methods for configuring a robotic simulator that is used to train a robot via machine learning. In one example, a method may include storing a domain description which comprises information about an operating environment of a robot, generating, via an automated planner, a plurality configuration files for a robotic simulator based on the domain description, where each configuration file comprises different configurations of a simulation environment for training a machine learning algorithm of the robot, and storing the plurality of configuration files in a memory device.

Abstract: Provided are systems and methods for decomposing learned robotic activities into smaller sub-activities that can be used independently. In one example, a method may include storing simulation data comprising an activity of a robot during a training simulation performed via a robotic simulator, decompose the activity into a plurality of sub-activities that are performed by the robot during the training simulation based on changes in behavior of the robot identified within the simulation data, and generating and storing a plurality of programs for executing the plurality of sub-activities, respectively, in the storage.