Abstract: The tooling system can have an extension unit having a pushing member and a pulling member extending along a length, the pushing member and the pulling member in sliding engagement and displaceable relative to one another, the extension unit having an internal passage configured for receiving a portion of the shaft via the first end; a pushing adapter engageable with and disengageable from the pushing member at the first end, and further engageable with and disengageable from one of the shaft and the component; and a pulling adapter engageable with and disengageable from the pulling member at the first end, and further engageable with and disengageable from an other one of the shaft and the component.

Abstract: An apparatus is provided for a gas turbine engine. This gas turbine engine apparatus includes a hot section structure of the gas turbine engine. The hot section structure is configured with a plurality of surfaces respectively forming boundaries of a gas path through the hot section structure. The surfaces include a first surface and a second surface. The hot section structure includes metal and thermal barrier material. The first surface is formed by the metal. The second surface is formed by the thermal barrier material.

Abstract: A rotary internal combustion engine includes a main rotor housing that has a peripheral wall that circumscribes a rotor cavity, a first interface surface and a second interface surface. A rotor is disposed within the rotor cavity. First and second side housings are secured against corresponding first and interface surfaces of the main rotor housing. The main rotor housing, the first side housing and the second side housing are formed from an aluminum alloy and at least one of the first interface surface and the second interface surface include an anti-fretting coating.

Abstract: A method for calibrating a machining system includes providing the machining system which includes a base, a cantilevered arm, and a rotary table positioned at the second arm end of the cantilevered arm. The rotary table is rotatable relative to the cantilevered arm about a first axis. The first axis has a first unloaded position and a first unloaded orientation with the machining system in an unloaded condition. The method further includes installing a measurement artifact on the rotary table, measuring a first position of the measurement artifact, and installing a load on the rotary table. The first axis has a first loaded position and a first loaded orientation with the machining system in a loaded condition. The method further includes measuring a second position of the measurement artifact and determining a positional deviation of the second position from the first position.

Abstract: An engine assembly for an aircraft propulsion system includes an engine, a rotational assembly, a gearbox assembly, and a controller. The engine includes an engine shaft. The rotational assembly includes a shaft, a bladed compressor rotor, and a bladed turbine rotor. The shaft interconnects the bladed compressor rotor and the bladed turbine rotor. The gearbox assembly includes a transmission assembly. The transmission assembly includes a first pump, a second pump, and a conduit circuit connecting the first pump and the second pump in fluid communication. The first pump is coupled with the engine shaft. The first pump is a variable displacement pump. The second pump is coupled with the rotational assembly. The controller is configured to identify a target rotation speed for the rotational assembly with an outer closed loop, identify a control signal with an inner closed loop using the target rotation speed, and controlling a displacement of the first pump using the control signal.

Abstract: A system is provided for an aircraft. This aircraft system includes a propulsion system, and the propulsion system includes a first thermal engine, a second thermal engine and a first electric machine. The propulsion system is configured to operate the first thermal engine and the second thermal engine, without operating the first electric machine, during a first mode of operation to provide aircraft thrust. The propulsion system is configured to operate the first electric machine and the second thermal engine, without operating the first thermal engine, during a second mode of operation to provide the aircraft thrust.

Abstract: An impending bypass switch includes a housing having a fluid inlet for admitting fluid into an interior of the housing. A piston is displaceable in the interior of the housing in response to a pressure variation of the fluid in the interior. A magnetic sensor is configured to receive a current and is fixedly mounted to the housing. A magnet is mounted to the piston to be displaced with the piston. Displacement of the magnet relative to the magnetic sensor causes a change in the current through the magnetic sensor. The change in the current is indicative of an impending bypass of a component.

Abstract: An aircraft system is provided that includes a thrust rotor and a powerplant coupled to and configured to drive rotation of the thrust rotor. The powerplant includes a gas turbine engine, a drivetrain and a fluid motor. The gas turbine engine includes a rotating assembly, a compressor section, a combustor section, a turbine section and a flowpath extending through the compressor section, the combustor section and the turbine section. The rotating assembly includes a turbine rotor in the turbine section. The turbine rotor is configured to convert fluid power of combustion products flowing through the flowpath within the turbine section into mechanical power for driving rotation of the rotating assembly during a mode of engine operation. The drivetrain is configured to rotatably couple the fluid motor to the rotating assembly. The fluid motor is configured to boost the mechanical power for driving the rotation of the rotating assembly during the mode of engine operation.

Abstract: An assembly for an aircraft propulsion system includes an output shaft and a gear assembly. The gear assembly is configured to drive rotation of the output shaft. The gear assembly includes a sun gear, a plurality of planet gears, a first ring gear, and a second ring gear. The sun gear is rotatable about a rotational axis. The sun gear includes a first double helical gear pattern. Each planet gear of the plurality of planet gears includes a main gear, a first lateral gear, and a second lateral gear. The main gear includes a second double helical gear pattern engaged with the first double helical gear pattern. The first lateral gear is engaged with the first ring gear. The second lateral gear is engaged with the second ring gear. The first ring gear is axially spaced from the second ring gear relative to the rotational axis.

Abstract: Methods and systems for operating an aircraft engine. A health parameter for the aircraft engine is monitored by a health evaluation device, the health parameter received from a first instrument. the health parameter is compared, by the health evaluation device, to a predetermined threshold. When the health parameter reaches the predetermined threshold, the health evaluation device wirelessly transmits a fault signal to a controller associated with the aircraft engine to elicit a health response from the controller, the fault signal containing at least two mutually-exclusive fault codes associated with an operating condition of the aircraft engine monitored by a second instrument.

Abstract: In accordance with at least one aspect of this disclosure, there is provided a fuel control system for gaseous fuel in an aircraft. The system includes a control module operatively connected to a metering device in a fuel flow conduit, the control module operable to control the flow of fuel through the fuel flow conduit. The control module includes an input line operable to receive a command input indicative of a requested engine state. In embodiments, the control module includes a compressibility logic and machine readable instructions. The machine readable instruction can be configured to cause the control module to control the metering device to achieve the requested engine state based on a compressibility factor input from the compressibility logic.

Abstract: A method of controlling an aircraft engine with an engine control system for limiting acceleration in response to reslam operation. The method includes triggering a reslam logic, detecting an acceleration that exceeds an acceleration threshold, and determining an elapsed time debit based on time elapsed from triggering of the reslam logic threshold. The elapsed time debit is inversely proportional to the elapsed time from triggering of the reslam logic and is utilized to set an engine acceleration reference value limit and maximum fuel rate limit.

Abstract: Methods and system for operating a gas turbine engine are described. The method comprises determining an actual engine output power based on a torque of the gas turbine engine, comparing the actual engine output power to an estimated engine output power to obtain an error, obtaining an actual engine speed and biasing the actual engine speed using the error to produce a biased engine speed, determining the estimated engine output power using a model-based estimator having the biased engine speed as input, detecting a torque-related fault based on the error and a first threshold, and accommodating the torque-related fault in response to detecting the torque-related fault.

Abstract: A gas turbine engine includes a propulsor, a compressor and a turbine section having a first turbine rotor connected to the compressor through a first shaft. A second turbine rotor is connected to at least the propulsor through a second shaft. The first turbine rotor drives the compressor at a first speed, and the second turbine rotor drives the propulsor at a second speed. There is a control for the gas turbine engine. The control is programmed to compare one of the first speed and the second speed to a commanded speed to determine a difference. The control also is programmed to identify a flame out condition, and identify a locked in stall condition when the difference between the two speeds exceeds a threshold for at least a threshold period of time at least in part concurrently with the flame out condition. The control is programmed to take a corrective action should the locked in stall condition be identified. A method is also disclosed.

Abstract: Gas turbine engines include a compressor section having an impeller configured to compress a core flow. A combustor section having a combustor is arranged downstream from the compressor section along a path of the core flow. A turbine section is arranged downstream from the combustor section along the path of the core flow, with the turbine section having a plurality of first vanes arranged at an outlet of the combustor and at least one first vane of the plurality of first vanes includes an internal vane path. The path of the core flow, in a flow direction, passes through the impeller of the compressor section, the internal vane path internal to the first vane, the combustor, and then between external surfaces of the plurality of first vanes to enter the turbine section.

Abstract: Lubricant pump systems and associated methods for aircraft engines are provided. The method includes receiving an input torque, dividing the input torque between a first load path receiving a first portion of the input torque, and a second load path receiving a second portion of the input torque. A first lubricant pump of the aircraft engine is driven via the first load path using the first portion of the input torque. A second lubricant pump of the aircraft engine is driven via the second load path using the second portion of the input torque. When a malfunction of the second lubricant pump occurs, the method includes ceasing to drive the first lubricant pump and the second lubricant pump using the input torque.

Abstract: An aircraft propulsion system for an aircraft having a nacelle that includes a pylon structure is provided. The system includes compressor and turbine sections, an IC engine, a fan, and an IC engine cooling system. The compressor section is powered by a first electric motor. The turbine section is configured to power a first electric generator configured to produce electrical power. The first fan is rotationally driven by a second electric motor. The fan has a hub and a plurality of fan blades extending radially outward from the hub. The hub is disposed in the pylon interior region and the fan blades are configured to extend outside of the pylon structure. The fan is positioned downstream of the compressor section. The IC engine cooling system has a heat exchanger and a pump configured to provide coolant communication between the IC engine and the heat exchanger.

Abstract: A method is provided for operating a system of an aircraft. During this method, rotation of a propulsor rotor is driven using mechanical power output by a powerplant. The powerplant includes a first drive device and a second drive device. The first drive device generates a first portion of the mechanical power. The second drive device generates a second portion of the mechanical power. A ratio between the first portion of the mechanical power and the second portion of the mechanical power is adjusted to control vibrations of the powerplant.

Abstract: Sandwich-structured noise-attenuating and/or structural panels and methods of manufacturing such panels using ultrasonic welding are described. The method includes: receiving a backing member, a sheet and a cellular structure; assembling the cellular structure between the backing member and the sheet; and ultrasonically welding the backing member and the sheet together.

Abstract: A test blade for a blade release test, including: an airfoil extending radially from a platform; and a root portion secured to the platform, the root portion having an elongated opening extending longitudinally from a leading edge of the root portion towards a trailing edge of the root portion, wherein the elongated opening has a width extending along the leading edge of the root portion that is greater than or equal to 3 times a radial height of the elongated opening.