Abstract: An engine includes a gearbox coupled to a first element and a second element. The engine may include a first speed sensor positioned on an input side of the gearbox, the first speed sensor operable to detect a first speed of the first element. The engine may further include a second speed sensor positioned on an output side of the gearbox, the second speed sensor operable to detect a second speed of the second element. The engine includes a controller comprising a processor and a memory coupled to the processor. The processor may be configured to receive data from the first speed sensor and the second speed sensor. The processor is also configured to detect a trigger event based on the data from the first speed sensor and the data from the second speed sensor. Upon detecting the trigger event, the processor communicates a command to an engine operational system.

Abstract: A gearbox assembly includes a gearbox having a gear assembly and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio is defined by V G V GB . VG is a gutter volume of the gutter and VGB is a gearbox volume. A gas turbine engine includes the gearbox assembly and a lubrication system. The lubrication system includes a sump that is a primary reservoir having a first lubricant level and a secondary reservoir in the gearbox assembly. The secondary reservoir has a second lubricant level. The lubrication system fills the secondary reservoir with a lubricant between the first lubricant level and the second lubricant level. The gear assembly collects the lubricant in the secondary reservoir to supply the lubricant to the gear assembly.

Abstract: Fluid-filled thrust link apparatus and associated method are disclosed. A thrust link for an aircraft engine includes a first wall having a forward portion and an aft portion at opposite ends of the thrust link, the forward portion coupled to the aircraft engine, the aft portion coupled to the aircraft engine, a pylon, or an aircraft associated with the aircraft engine, and a second wall within an interior area surrounded by the first wall, the second wall spaced apart from the first wall, a space between the first wall and the second wall defining a channel within the interior area, the channel including a fluid, the fluid pressurized based on a damping ratio to withstand a resonant vibration frequency generated by the aircraft engine.

Abstract: A gas turbine engine includes a gearbox assembly that includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by: V G V G ? B · V G is a gutter volume of the gutter and VGB is a gearbox volume. The gas turbine engine includes a lubricant flow control system that includes a variable flow lubricant pump that generates a pump variable flow of lubricant to the gearbox assembly. The gearbox assembly has a variable consumption demand for delivery of lubricant. A lubricant flow controller is configured to generate a pump control command for the variable flow lubricant pump to produce the pump variable flow of lubricant based on the variable consumption demand.

Abstract: A gas turbine engine includes a fan, a combustor positioned in a core air flowpath that generates combustion gases, a steam system that extracts water from the combustion gases and generates steam, and a gearbox assembly. The steam system includes water storage devices that store the water therein. The water storage devices include a first state in which a level of the water increases or is maintained and a second state in which the level of the water decreases as the water flows through the water storage devices. The gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by: V G V G ? B · V G is a gutter volume of the gutter and VGB is a gearbox volume.

Abstract: A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows: ( T OUT A DTExit ) 2 * EGT A HPCExit * 1 ? 0 - 1 ? 1 ; wherein CSP is greater than 0.0001194×EGT2?0.103×EGT+22.14 and less than 0.0003294×EGT2?0.306×EGT+77.

Abstract: Variable diameter thrust link apparatus are disclosed. A thrust link for an aircraft engine includes a forward end coupled to the aircraft engine, the forward end having a first diameter, and an aft end coupled to the aircraft engine or a pylon, the aft end having the first diameter, wherein the thrust link defines a thrust link span that extends from the forward end to the aft end, wherein a second diameter greater than the first diameter is defined between the forward end and the aft end, wherein the second diameter spans a center diameter span that is equivalent to or between 60% and 90% of the thrust link span.

Abstract: Viscous damper apparatus and associated methods to control a response to a resonant vibration frequency are disclosed. An apparatus to support an aircraft engine includes a thrust link including a forward end and an aft end, the forward end of the thrust link coupled to the aircraft engine, and a damper including a piston rod coupled to the aft end of the thrust link, the piston rod including a piston, and a chamber including a fluid, the piston to move within the chamber.

Abstract: A turbine engine for an aircraft. The turbine engine includes a core turbine engine and a steam system. The steam system extracts water from combustion gases, vaporizes the water to generate steam, and injects the steam into a core air flow path of the core engine to add mass flow to core air. A core hot gas path component may be fluidly connected to a hot gas path that routes combustion gasses from the combustor. The core turbine engine may also include a turbine fluidly connected to the hot gas path to receive the combustion gases. The turbine may include a turbine airfoil. Each of the core hot gas path component and the turbine airfoil includes a combustion-gas-facing surface facing the hot gas path. A hydrophobic coating is formed on the combustion-gas-facing surface, reducing wetting of water vapor within the combustion gases on the core hot gas path component.

Abstract: A turbine engine for an aircraft. The turbine engine includes a combustor and a steam system. Fuel and steam are injected into the combustor to mix with compressed air to generate a fuel and air mixture. The fuel and air mixture is combusted in the combustor to generate combustion gases. The steam system is fluidly coupled to the combustor as the steam source to provide steam to the combustor. The steam system includes a hot gas path and a steam hot gas path component. The hot gas path is fluidly coupled to the combustor to receive the combustion gases and to route the combustion gases through the steam system. The steam hot gas path component includes a wall having a combustion-gas-facing surface facing the hot gas path and a hydrophobic coating formed on the combustion-gas-facing surface.

Abstract: A turbine engine for an aircraft. The turbine engine includes a combustor and a steam system. Fuel and steam are injected into the combustor to mix with compressed air to generate a fuel and air mixture. The fuel and air mixture is combusted in the combustor to generate combustion gases. The steam system is fluidly coupled to the combustor as the steam source to provide steam to the combustor. The steam system includes a hot gas path and a steam hot gas path component. The hot gas path is fluidly coupled to the combustor to receive the combustion gases and to route the combustion gases through the steam system. The steam hot gas path component includes a wall having a combustion-gas-facing surface facing the hot gas path and a hydrophobic coating formed on the combustion-gas-facing surface.

Abstract: A turbine engine for an aircraft. The turbine engine includes a combustor fluidly coupled to a fuel delivery assembly to receive fuel from the fuel delivery assembly. The fuel is injected into the combustor and combusted in the combustor to generate combustion gases. A condenser is located downstream of the combustor to receive the combustion gases and to condense water. A fuel heat exchanger is thermally coupled to the condenser by at least one heat transfer loop to receive heat from the condenser.

Abstract: A turbine engine for an aircraft. The turbine engine includes a combustor fluidly coupled to a fuel delivery assembly to receive fuel from the fuel delivery assembly. The fuel is injected into the combustor and combusted in the combustor to generate combustion gases. A condenser is located downstream of a turbine to receive the combustion gases and to condense water. The fuel heat exchanger is thermally coupled to the condenser to receive heat from the water condensed by the condenser. The fuel heat exchanger is located in the fuel delivery assembly to receive the fuel and to transfer the heat received from the water to the fuel. The boiler is located downstream of the fuel heat exchanger. The boiler receives the water and is fluidly connected to the combustor to receive the combustion gases and to boil the water to generate steam.

Abstract: An aircraft gas turbine includes a combustor that combusts compressed air and at least one fuel flow from at least one fuel source to generate combustion gases. A fuel supply system is arranged to provide at least one of a first flow of fuel to the combustor via a first fuel supply line and a second flow of a heated fuel to the combustor via a second fuel supply line. A fuel heat exchanger is arranged within the fuel supply system to generate the heated fuel that is heated above an autoignition temperature of the fuel, and a heat source communicates with the fuel heat exchanger for generating the heated fuel.

Abstract: An aircraft gas turbine includes a combustor that combusts compressed air and at least one fuel flow from at least one fuel source to generate combustion gases. A fuel supply system is arranged to provide at least one of a first flow of fuel to the combustor via a first fuel supply line and a second flow of a heated fuel to the combustor via a second fuel supply line. A fuel heat exchanger is arranged within the fuel supply system to generate the heated fuel that is heated above an autoignition temperature of the fuel, and a heat source communicates with the fuel heat exchanger for generating the heated fuel.

Abstract: A fuel leak detection system for hydrogen fuel system including a monitored component. The fuel leak detection system including a sensor and controller communicatively coupled to the sensor. The sensor is positioned to monitor at least a portion of the monitored component. The sensor is configured (i) to sense a parameter corresponding to a hydrogen fuel leak of the monitored component and (ii) to generate an output. The controller is configured (i) to receive the output of the sensor, (ii) to determine, based on the output of the sensor, if a leak has occurred in the monitored component, and (iii) to generate an output indicating a fuel system leak when the controller determines that the leak has occurred in the monitored component. The monitored component may be a component of one of a fuel tank, a power generator, and a fuel delivery assembly.

Abstract: A turbine engine for an aircraft including a turbo-engine, a fan having a fan shaft coupled to the turbo-engine, and a steam system. The steam system is fluidly coupled to a core air flow path at a plurality of steam injection locations to selectively inject steam into the core air flow path at each of the plurality of steam injection locations, using a steam flow control valve. The plurality of steam injection locations include (i) an upstream steam injection location located to inject steam into the core air flow path at a primary combustion zone of a combustor or upstream thereof and (ii) a downstream steam injection location located to inject steam into the core air flow path downstream of the primary combustion zone.

Abstract: A turbine engine for an aircraft. The turbine engine includes a combustor, a core shaft, a turbine, and a steam system providing steam to a core air flow path. The steam system includes a boiler that receives combustion gases to boil water to generate steam. A steam turbine is fluidly coupled to the boiler to receive the steam from the boiler and to cause the steam turbine to rotate. The steam turbine is coupled to the core shaft to rotate the core shaft when the steam turbine rotates. The steam system may include a bypass flow path selectively operable to redirect at least one of the steam or water, and to bypass the core air flow path. The turbine engine may also include at least one steam control valve located downstream of the boiler and upstream of the core air flow path to control the flow of steam.

Abstract: A lubrication system for a turbine engine having a longitudinal centerline axis and one or more rotating components. The lubrication system includes a primary lubrication system that supplies lubricant to the one or more rotating components during normal operation of the turbine engine. The lubrication system also includes an auxiliary lubrication system. The auxiliary lubrication system includes an auxiliary reservoir and a lubricant dispersion device. The auxiliary reservoir stores the lubricant therein. The lubricant dispersion device rotates about the longitudinal centerline axis. The lubricant dispersion device collects the lubricant in the auxiliary reservoir and disperses the lubricant to the one or more rotating components as the lubricant dispersion device rotates.

Abstract: A turbine engine for an aircraft. The turbine engine includes a fan, a combustor in a core air flowpath that generates combustion gases, a core shaft, a turbine that receives the combustion gases to rotate the turbine, and a steam system. The steam system extracts water from the combustion gases, vaporizes the water to generate steam, and injects the steam into the core air flowpath. The steam system includes one or more water storage devices that store the water therein. The one or more water storage devices include a first state in which the one or more water storage devices increase or maintain a level of the water as the water flows through the one or more water storage devices, and a second state in which the one or more water storage devices decrease the level of the water in the one or more water storage devices.