Abstract: Methods, apparatus, systems and articles of manufacture are disclosed. An apparatus for mounting a gas turbine engine to a pylon includes: a thrust link coupled to the gas turbine engine and an aft mount, the gas turbine engine coupled to the pylon via a forward mount and the aft mount, a bending restraint having a first end and a second end, the bending restraint including an actuator positioned at the first end to apply a force to a fan section of the gas turbine engine, a first joint positioned at the first end of the bending restraint, the first joint coupled to the first end of the bending restraint and the fan section, and a second joint positioned at the second end of the bending restraint, the second joint coupled to the second end of the bending restraint.

Abstract: A gas turbine engine including a rotating shaft and a bearing assembly. The bearing assembly is configured to support the rotating shaft. The bearing assembly has a first bearing and a second bearing. The first bearing is configured to support an axial load of the rotating shaft when a forward thrust of the rotating shaft is greater than ten percent. The second bearing is configured to support the axial load of the rotating shaft when a thrust of the rotating shaft is between ten percent forward thrust and ten percent aft thrust, inclusive of the end points.

Abstract: A thrust mount assembly includes a thrust link-lever that has a first end region, a second end region, and a fulcrum region disposed between the first end region and the second end region. A thrust link-lever may be coupled or couplable to a fulcrum body of an aft engine-mount at a fulcrum position of the thrust link-lever. When coupled to a turbomachine, the fulcrum position of the thrust link-lever may be laterally offset and/or laterally adjustable relative to an axis of rotation of the turbomachine. A load translated from an engine frame of a turbomachine to an engine-mounting linkage system may be determined and a fulcrum position for the thrust mount assembly may be adjusted laterally relative to an axis of rotation of the turbomachine.

Abstract: A gas turbine engine including a rotating shaft and a bearing assembly. The bearing assembly is configured to support the rotating shaft. The bearing assembly has a first bearing and a second bearing. The first bearing is configured to support an axial load of the rotating shaft when a forward thrust of the rotating shaft is greater than ten percent. The second bearing is configured to support the axial load of the rotating shaft when a thrust of the rotating shaft is between ten percent forward thrust and ten percent aft thrust, inclusive of the end points.

Abstract: Engine-mounting linkage systems may include a plurality of engine mounting links configured to couple an engine frame to an engine support structure of an aircraft. The plurality of engine-mounting links may include a forward link that is connectable to a forward frame portion of the engine frame and the engine support structure, and an aft link that is connectable to an aft frame portion of the engine frame and the engine support structure. The engine mounting linkage system may include an actuator that is connectable to the engine support structure and one of the plurality of engine-mounting links, or to the engine support structure and the engine frame. The actuator, when actuated, changes an inclination angle ? of at least one of the plurality of engine-mounting links.

Abstract: A gas turbine engine includes a compressor rotor shaft assembly, an accessory gearbox, and a bowed-rotor mitigation drive device drivingly coupled with the accessory gearbox. The bowed-rotor mitigation drive device is driven during an engine startup phase so as to induce a mechanical load (mechanical energy) to the bowed-rotor mitigation drive device. The mechanical load (mechanical energy) is retained within the bowed-rotor mitigation drive device during operation of the gas turbine engine. The mechanical load (mechanical energy) retained within the bowed-rotor mitigation drive device is periodically released by the bowed-rotor mitigation drive device in a plurality of periods so as to provide, in each period, a driving force to the accessory gearbox, which provides the driving force to the compressor rotor shaft assembly to periodically rotate the compressor rotor shaft assembly.

Abstract: A turbomachine including a turbine rotor, a compressor rotor, and a shaft, and at least one balance weight assembly connected to the shaft. The shaft drivingly connects the turbine rotor with the compressor rotor to rotate the compressor rotor about a rotational axis when the turbine rotor rotates about the rotational axis. The at least one balance weight assembly including a first chamber, at least one additional chamber, and a balance weight movable between the first chamber and the at least one additional chamber.

Abstract: Example engine apparatus and associated control methods are disclosed. An example engine apparatus includes: a frame; a lug to attach the frame to an aircraft; a mount cover positioned over the lug; and a heating mechanism to regulate a temperature of the lug under the mount cover.

Abstract: A gas turbine engine includes a compressor section with a compressor rotor shaft assembly including a plurality of compressor rotors longitudinally spaced apart from each other via respective ones of a plurality of shaft sections of a rotor shaft, each compressor rotor of the plurality of compressor rotors having a plurality of rotor vanes extending radially outward therefrom and being circumferentially spaced about the compressor rotor. A stator shroud assembly has a stator shroud casing surrounding the compressor rotor shaft assembly, a compressor flow passage being defined between the compressor rotor shaft assembly and the stator shroud casing. A pressurized air source generates a flow of pressurized air to be provided to the compressor section, and a plurality of pressurized airflow nozzles are connected with the pressurized air source and provide the flow of pressurized air into the compressor flow passage to cause the compressor rotor shaft assembly to rotate.

Abstract: Disclosed herein are example variable flowpath casings for blade tip clearance control. An example casing for a turbine engine includes an annular substrate extending along an axial direction, the annular substrate including a first surface and a second surface that is radially inward relative to the first surface; an actuator structure coupled to the second surface of the annular substrate; and a smart structure cantilevered from the second surface of the annular substrate, the smart structure including: a support structure, a first region of the support structure coupled to the second surface of the annular substrate, a second region of the support structure coupled to the actuator structure; and a radially inward surface defining a variable surface, the support structure to move the variable surface in a radial direction.

Abstract: A cooling system for a turbine engine. The turbine engine includes a compressor, a turbine, and a shaft that drivingly couples the compressor and the turbine. The cooling system includes one or more cavities of the compressor. The cooling system includes a shaft flowpath defined in the shaft. The shaft includes one or more shaft apertures that provide fluid communication between the shaft flowpath and the one or more cavities. Air passes through the one or more shaft apertures and the shaft flowpath during a shutdown of the turbine engine to cool the one or more cavities.

Abstract: A gas turbine engine includes a casing, a rotor assembly rotatably mounted within the casing, the rotor assembly comprising a forward rotor portion defining a forward rotor cavity and an aft rotor portion defining an aft rotor cavity, the aft rotor cavity being fluidly isolated from the forward rotor cavity, and a cooling system. The cooling system includes a forward cooling fluid supply, a forward supply line providing fluid communication between the forward cooling fluid supply and the forward rotor cavity, an aft cooling fluid supply, an aft supply line providing fluid communication between the aft cooling fluid supply and the aft rotor cavity, and a flow control system fluidly coupled to the forward supply line and the aft supply line for regulating a flow of forward cooling fluid and a flow of aft cooling fluid.

Abstract: A method for cooling a compressor section of a gas turbine engine includes sensing, via at least one first temperature sensor, a first temperature at the first location on a stationary assembly of the compressor section. The method also includes sensing, via at least one second temperature sensor, a second temperature at a second location on the stationary assembly of the compressor section. The second location is spaced apart from the first location. The method also includes determining, via a controller, a delta between the first temperature and the second temperature. Further, the method includes operating, via the controller, at least one cooling element when the delta exceeds a predetermined threshold, the at least one cooling element provided at the first location of the stationary assembly of the compressor section.

Abstract: A gas turbine engine includes a compressor rotor shaft assembly, an accessory gearbox, and a bowed-rotor mitigation drive device drivingly coupled with the accessory gearbox. The bowed-rotor mitigation drive device is driven during an engine startup phase so as to induce a mechanical load (mechanical energy) to the bowed-rotor mitigation drive device. The mechanical load (mechanical energy) is retained within the bowed-rotor mitigation drive device during operation of the gas turbine engine. The mechanical load (mechanical energy) retained within the bowed-rotor mitigation drive device is periodically released by the bowed-rotor mitigation drive device in a plurality of periods so as to provide, in each period, a driving force to the accessory gearbox, which provides the driving force to the compressor rotor shaft assembly to periodically rotate the compressor rotor shaft assembly.

Abstract: Disclosed herein are example variable flowpath casings for blade tip clearance control. An example casing for a turbine engine includes an annular substrate extending along an axial direction, the annular substrate defining a cavity at a radially inward surface of the annular substrate, and a smart structure coupled to the annular substrate, the smart structure including a support structure; an actuator structure to at least one of expand or contract in response to a change in temperature of the actuator structure, and a variable surface coupled to the support structure, the support structure to move the variable surface in a radial direction.

Abstract: A gas turbine engine includes a casing, a rotor assembly rotatably mounted within the casing, the rotor assembly comprising a forward rotor portion defining a forward rotor cavity and an aft rotor portion defining an aft rotor cavity, the aft rotor cavity being fluidly isolated from the forward rotor cavity, and a cooling system. The cooling system includes a forward cooling fluid supply, a forward supply line providing fluid communication between the forward cooling fluid supply and the forward rotor cavity, an aft cooling fluid supply, an aft supply line providing fluid communication between the aft cooling fluid supply and the aft rotor cavity, and a flow control system fluidly coupled to the forward supply line and the aft supply line for regulating a flow of forward cooling fluid and a flow of aft cooling fluid.

Abstract: A gas turbine engine includes a compressor section with a compressor rotor shaft assembly including a plurality of compressor rotors longitudinally spaced apart from each other via respective ones of a plurality of shaft sections of a rotor shaft, each compressor rotor of the plurality of compressor rotors having a plurality of rotor vanes extending radially outward therefrom and being circumferentially spaced about the compressor rotor. A stator shroud assembly has a stator shroud casing surrounding the compressor rotor shaft assembly, a compressor flow passage being defined between the compressor rotor shaft assembly and the stator shroud casing. A pressurized air source generates a flow of pressurized air to be provided to the compressor section, and a plurality of pressurized airflow nozzles are connected with the pressurized air source and provide the flow of pressurized air into the compressor flow passage to cause the compressor rotor shaft assembly to rotate.

Abstract: A gas turbine engine including a first rotor assembly comprising a first drive shaft extended along a longitudinal direction; a housing coupled to the first rotor assembly to provide rotation of the first rotor assembly around an axial centerline; a first accessory assembly, wherein the first accessory assembly sends and/or extracts energy to and from the first rotor assembly; and a first clutch assembly disposed between the first rotor assembly and the first accessory assembly. The first clutch assembly engages and disengages the first rotor assembly to and from the first accessory assembly.

Abstract: A system for modulating airflow into a bore defined by a rotor of a gas turbine engine defining an axial direction, a circumferential direction, and a radial direction is provided. The system includes a movable member positioned forward of a first stage of rotor blades of the rotor. The movable member is movable between at least a first position and a second position to modulate airflow into the bore via a plurality of opening in fluid communication with the bore.

Abstract: Systems and methods for optimizing clearances within an engine include an adjustable coupling configured to couple a thrust link to the aircraft engine, an actuator coupled to the adjustable coupling, where motion produced by the actuator adjusts a hinge point of the adjustable coupling, sensors configured to capture real time flight data, and an electronic control unit. The electronic control unit receives flight data from the sensors, implements a machine learning model trained to predict clearance values within the engine based on the received flight data, predicts, with the machine learning model, the clearance values within the engine based on the received flight data, determines an actuator position based on the clearance values, and causes the actuator to adjust to the determined actuator position.