Abstract: A fan assembly for a gas turbine engine includes a fan disk, a trunnion, an actuation device, a fan blade, and a counterweight assembly. The trunnion is mounted to the fan disk. The actuation device is operably coupled to the trunnion. The fan blade is rotatably attached to the fan disk. The counterweight assembly includes a link arm, a lever arm, a hinge, and a counterweight. The link arm is connected to the trunnion, to the actuation device, or to both. The link arm is configured to drive rotation of the trunnion relative to the fan disk. The hinge is pivotably connected to the lever arm. The lever arm is connected to the link arm and is disposed to rotate about a connection point of the lever arm and the hinge. The counterweight is mounted to the lever arm at a location spaced from the hinge.

Abstract: A heat engine comprising a compressor providing a flow of compressed air from a core flowpath of the heat engine; a cooled cooling air (CCA) heat exchanger system to which the flow of compressed air is provided from the compressor; a coolant supply system providing a flow of coolant to the CCA heat exchanger in thermal communication with the flow of compressed air at the CCA heat exchanger, in which the coolant supply system and CCA heat exchanger together define a CCA circuit through which the compressed air flows in thermal communication with the coolant; and a hot section disposed downstream of the compressor section along the core flowpath through which combustion gases flow, in which the hot section defines a secondary flowpath through which the flow of compressed air from the CCA heat exchanger is provided.

Abstract: A method of regulating pressure in a thermal transport bus of a gas turbine engine, the method including: operating the gas turbine engine with the thermal transport bus having an intermediary heat exchange fluid flowing therethrough, the thermal transport bus including one or more heat source heat exchangers and one or more heat sink heat exchangers in thermal communication through the intermediary heat exchanger fluid; and adjusting a flow volume of the thermal transport bus using a variable volume device in fluid communication with the thermal transport bus in response to a pressure change associated with the thermal transport bus.

Abstract: A method of detecting an airflow fault condition in a gas turbine engine, the method including: operating the gas turbine engine with a thermal transport bus having an intermediary heat exchange fluid flowing therethrough; determining a performance characteristic of the intermediary heat exchange fluid in the thermal transport bus is outside of a predetermined range, wherein the performance characteristic includes a temperature, a pressure, a flowrate, or a combination thereof; and indicating an airflow fault condition in response to determining the performance characteristic is outside of the predetermined range.

Abstract: A thermal management system for transferring heat between fluids includes a thermal transport bus through which a heat exchange fluid flows. Additionally, the system includes a heat source heat exchanger arranged along the bus such that heat is added to the fluid flowing through the heat source heat exchanger. Moreover, the system includes a plurality of heat sink heat exchangers arranged along the bus such that heat is removed from the fluid flowing through the plurality of heat sink heat exchangers. Furthermore, the system includes a bypass conduit fluidly coupled to the bus such that the bypass conduit allows the fluid to bypass one of the heat source heat exchanger or one of the heat sink heat exchangers. In addition, the system includes a valve configured to control a flow of the fluid through the bypass conduit based on a pressure of the fluid within the bus.

Abstract: A fuel oxygen conversion unit for a vehicle or an engine of the vehicle includes a contactor; a mechanically-driven, first fuel gas separator defining a liquid fuel outlet and a stripping gas outlet, the fuel oxygen conversion unit defining a liquid fuel outlet path in fluid communication with the liquid fuel outlet of the first fuel gas separator; and a second fuel gas separator positioned in fluid communication with the liquid fuel outlet path at a location downstream of the first fuel gas separator.

Abstract: A gas turbine engine is provided including a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order; a rotor assembly driven by the turbomachine, the rotor assembly, the turbomachine, or both comprising a substantially annular duct relative to the centerline of the gas turbine engine, the annular duct defining a flowpath; a heat exchanger positioned within the annular duct and extending substantially continuously along the circumferential direction, the heat exchanger comprising a first material defining a heat exchange surface exposed to the flowpath, wherein the first material defines a heat exchange coefficient and wherein the heat exchange surface defines a surface area (A), and wherein the heat exchanger has an effective transmission loss (ETL) of between 5 decibels and 1 decibel for an operating condition.

Abstract: A fan assembly for a gas turbine engine includes a fan disk, a trunnion, a fan blade, and a counterweight assembly. The fan disk is configured to rotate about an axial centerline of the gas turbine engine when installed in the gas turbine engine. The trunnion is mounted to the fan disk and defines a slot extending through a portion of the trunnion. The fan blade defines a pitch axis and is rotatably attached to the fan disk about its pitch axis through the trunnion. The counterweight assembly includes a link arm extending to the trunnion and an engagement device mounted to the link arm that is disposed to move through the slot of the trunnion.

Abstract: A fan assembly for a gas turbine engine includes a fan disk, a trunnion, an actuation device, a fan blade, and a counterweight assembly. The trunnion is mounted to the fan disk. The actuation device is operably coupled to the trunnion. The fan blade is rotatably attached to the fan disk. The counterweight assembly includes a link arm, a lever arm, a hinge, and a counterweight. The link arm is connected to the trunnion, to the actuation device, or to both. The link arm is configured to drive rotation of the trunnion relative to the fan disk. The hinge is pivotably connected to the lever arm. The lever arm is connected to the link arm and is disposed to rotate about a connection point of the lever arm and the hinge. The counterweight is mounted to the lever arm at a location spaced from the hinge.

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: 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 system and method for using a fuel with an engine, an airframe having an aircraft heat load, a fuel tank, and a fuel oxygen reduction unit are provided. The method includes receiving an inlet fuel flow in the fuel oxygen reduction unit for reducing an amount of oxygen in the inlet fuel flow; separating a fuel/gas mixture within the fuel oxygen reduction unit into an outlet gas flow and an outlet fuel flow exiting the fuel oxygen reduction unit; controlling a first portion of the outlet fuel flow to the engine; and controlling a second portion of the outlet fuel flow to the airframe to transfer heat between the second portion of the outlet fuel flow and the aircraft heat load.

Abstract: A gas turbine engine includes a compressor section and a turbine section. A first spool is rotatable with a first turbine of the turbine section and a first compressor of the compressor section. Additionally, a second spool is rotatable with a second turbine of the turbine section and a second compressor of the compressor section. An electric machine is positioned at least partially inward of the core air flowpath of the gas turbine engine along a radial direction of the gas turbine engine. Additionally, the exemplary gas turbine engine includes a gear assembly mechanically coupled to both the first spool and the second spool, such that the first and second spools may each drive the electric machine.

Abstract: Turbine engine systems equipped with a fuel deoxygenation system and turboelectric power system are provided. In one aspect, a turbine engine system includes a turboelectric power system having an electric generator operatively coupled with a gas turbine engine. A fuel line provides fuel to the gas turbine engine. A fuel deoxygenation system is positioned along the fuel line for reducing an amount of oxygen in the fuel. A main fuel pump is positioned downstream of the fuel deoxygenation system along the fuel line. An electric motor drives the main fuel pump. An electric generator of a turboelectric power system is operatively coupled with the gas turbine engine and generates electrical power. The electrical power is provided to the electric motor and the fuel deoxygenation system. Heat is recovered from the turboelectric power system and the electric motor and is imparted to fuel provided to the gas turbine engine.

Abstract: A thermal management system for transferring heat between fluids includes a thermal transport bus through which a heat exchange fluid flows. Additionally, the system includes a heat source heat exchanger arranged along the bus such that heat is added to the fluid flowing through the heat source heat exchanger. Moreover, the system includes a plurality of heat sink heat exchangers arranged along the bus such that heat is removed from the fluid flowing through the plurality of heat sink heat exchangers. Furthermore, the system includes a bypass conduit fluidly coupled to the bus such that the bypass conduit allows the fluid to bypass one of the heat source heat exchanger or one of the heat sink heat exchangers. In addition, the system includes a valve configured to control a flow of the fluid through the bypass conduit based on a pressure of the fluid within the bus.

Abstract: A method of operating a fuel oxygen reduction unit for a vehicle or a gas turbine engine of the vehicle is provided. The fuel oxygen reduction unit including a contactor and a fuel gas separator, and further defining a stripping gas flowpath in flow communication with a stripping gas inlet of the contactor and a stripping gas outlet of the fuel gas separator. The method includes receiving data indicative of a parameter of a stripping gas flow through the stripping gas flowpath or of a component in flow communication with the stripping gas flow through the stripping gas flowpath; and determining an operability condition of the fuel oxygen reduction unit, or a component operable with the fuel oxygen reduction unit, based on the data received indicative of the parameter of the stripping gas flow or of the component in flow communication with the stripping gas flow.

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 three-stream engine is provided. The three-stream engine includes a fan section, a core engine disposed downstream of the fan section, and a core cowl annularly encasing the core engine and at least partially defining a core duct. A fan cowl is disposed radially outward from the core cowl and annularly encasing at least a portion of the core cowl. The fan cowl at least partially defining an inlet duct and a fan duct. The fan duct and the core duct at least partially co-extending axially on opposite sides of the core cowl. A heat exchanger disposed within the fan duct. The heat exchanger provides for thermal communication between a fluid flowing through fan duct and a motive fluid flowing through the heat exchanger.

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: 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.