Abstract: A fuel injector for a gas turbine engine combustor is provided that includes a swirler, a mounting stage, and a distributor. The swirler has a shaft, a collar, a throat section, and first and second axial ends. The throat section includes an inner radial surface that defines a central passage that extends between the swirler inner bore and the collar. The collar includes a plurality of apertures extending therethrough disposed radially outside of the central passage. The mounting stage is disposed in the inner bore, and has an annular flange, a central hub, and at least one strut. The distributor has a stem attached to a head. The stem has a distal end opposite the head portion engaged with the central hub. The head portion has an end surface and a side surface. The distributor is selectively positionable relative to the throat section.

Abstract: An engine housing includes a rotor housing body, a side housing body a side plate and a seal assembly. The rotor housing body extends about an axis to form a rotor cavity. The rotor housing body extends between and to a first axial end and a second axial end. The side housing body is disposed at the first axial end. The side plate includes an inner side, an outer side, and a perimeter edge extending from the inner side to the outer side. The seal assembly includes a support ring and a spacer. The support ring and the spacer extend about the axis. The support ring is mounted to the perimeter edge by a braze joint. The spacer extends between and to a first axial spacer end and a second axial spacer end. The first axial spacer end is disposed at the support ring and the outer side.

Abstract: A method for monitoring a bleed valve of an aircraft engine includes obtaining operational data of the aircraft engine, the operational data indicative of operational performance of the aircraft engine. The method further includes utilizing one or more neural networks to determine a health status of a bleed valve of the aircraft engine based on the operational data, the bleed valve configured to bleed air from a compressor section of the aircraft engine, wherein the one or more neural networks have been trained with root cause data relating to a plurality of potential failure modes of the bleed valve.

Abstract: A method is disclosed for providing a component. During this method, a substrate is scanned using computed tomography to provide substrate scan data. The substrate scan data is compared to substrate reference data to provide additive manufacturing data. Braze powder is deposited with the substrate based on the additive manufacturing data. The braze powder is sintered together during the depositing of the braze powder to provide the substrate with sintered braze material. The sintered braze material is heated to melt the sintered braze material and to diffusion bond the sintered braze material to the substrate.

Abstract: A method of and system for operating a hybrid electric propulsion system for an aircraft in a start sequence is provided. The hybrid electric propulsion system includes a thermal engine, an electric motor, a gearbox, an electric power storage unit, and a propulsion unit having a propeller having propeller blades. The method includes: driving the propeller from a static state to a target rotational speed within a predetermined range of speeds using the electric motor; transitioning the propeller blades from a feathered mode to an unfeathered mode while the propeller is being driven at the target rotational speed solely by the electric motor; starting the thermal engine; and transitioning the driving of the propeller using the electric motor to driving the propeller using the thermal engine.

Abstract: An aircraft engine lubrication system for an aircraft engine having a first component (FC) and a second component (SC) is provided. The system includes a lubricant tank (LT), a first pump (FP), a second pump (SP), and a de-aerator (DA). The LT outlet is in fluid communication with a FP inlet. The FP outlet is in fluid communication with a SP inlet. The SP outlet is in fluid communication with a FC inlet and the SC inlet. The SC outlet is in fluid communication with a LT first inlet. The DA gas outlet is in fluid communication with a LT second inlet. The DA liquid outlet is in fluid communication with a SP inlet.

Abstract: A gas turbine engine includes a turbine for driving a shaft to in turn drive a rotor. A combustor receives compressed air, and combusts the compressed air. The shaft is supported by at least one bearing. An oil delivery system is associated with the at least one bearing including an oil scoop rotating with the at least one shaft and a stationary oil jet provided with a source of lubricant. A mixed fluid chamber is within the shaft. The mixed fluid chamber is connected to an air exit. There further is an oil chamber radially outward of the mixed fluid chamber. The oil chamber is connected to the at least one bearing.

Abstract: A method and system for correlating engine thrust and engine thrust lever angle (TLA) during operation of a gas turbine engine powered aircraft is provided. The method includes: a) providing current flight conditions including a Mach number of the aircraft, altitude of the aircraft, and a TLA; b) determining a corrected fan speed value based on a sensed fan speed and flight conditions; c) determining a first corrected net thrust value based on the corrected fan speed value and the Mach number; d) determining a compensation factor using the corrected fan speed value, the Mach number, and the corrected net thrust value; e) determining a second corrected net thrust value as a function of the TLA; and f) correlating an engine fan speed to the TLA using the second corrected net thrust value as a function of the TLA and the compensation factor.

Abstract: A method for checking a closing point for a bleed-off valve for a gas turbine engine includes determining a modulation characteristic curve for the bleed-off valve, determining a nominal closing point value for the bleed-off valve on the modulation characteristic curve, operating the gas turbine engine and increasing an engine power of the gas turbine engine until the gas turbine engine parameter reaches a predetermined testing value, and determining a bleed-off valve measured value and a gas turbine engine measured value when the gas turbine engine parameter reaches the predetermined testing value. The gas turbine engine measured value is different than the nominal closing point value. The method further includes determining an extrapolated closing point value and checking the closing point for the bleed-off valve by comparing the bleed-off valve measured value or the gas turbine engine measured value to the extrapolated closing point value.

Abstract: An engine system is provided that includes an engine rotating structure, an electric machine rotating structure and a flex coupler. The flex coupler rotatably connects the electric machine rotating structure to the engine rotating structure. The flex coupler includes a first mount, a second mount and a flex plate. The first mount includes a plurality of first mount fingers arranged circumferentially about an axis. The second mount includes a plurality of second mount fingers arranged circumferentially about the axis. The flex plate connects the first mount to the second mount. The flex plate includes a plurality of first flex plate fingers and a plurality of second flex plate fingers. Each of the first flex plate fingers is attached to a respective one of the first mount fingers. Each of the second flex plate fingers is attached to a respective one of the second mount fingers.

Abstract: An aircraft propulsion system includes a bladed rotor configured for rotation about a rotational axis. A rotor blade of the bladed rotor of rotor blades includes a blade body. The blade body has a blade span extending between and to a base end and a tip end. The blade body has a blade chord extending between and to a leading edge and a trailing edge. The blade body forms a pressure side surface and a suction side surface. The blade body forms a recess at the pressure side surface. The recess includes a recess surface portion forming a portion of the pressure side surface. The recess surface portion has a recess height. The recess height is greater than or equal to a Y-value of 0.025 for a normalized two-dimensional curve of a surface profile of the recess surface portion taken along a Y-Z plane orthogonal to the rotational axis for a recess span.

Abstract: A method of braking a rotational air mover portion of an aircraft propulsion system is provided. The method includes: using an electrical device to apply a torque to a propulsion unit driven by a propulsion system of an aircraft, the propulsion unit including a rotational air mover configured to provide thrust for propelling the aircraft, wherein the electrical device is configured to produce electrical energy during the process of applying the torque to the propulsion unit; controlling the electrical device to apply an amount of the torque that is sufficient to cause the rotational air mover to decelerate or to maintain the rotational air mover non-rotational; and directing the electrical energy produced by the electrical device to an aircraft component.

Abstract: A rotor housing for an aircraft rotary engine includes a side housing body and a rail. The side housing body extends along an axis between and to an inner side and an outer side. The side housing body forms a fluid cooling passage and a plurality of ribs. The fluid cooling passage extends about the axis at the inner side. The plurality of ribs are coincident with and extend into the fluid cooling passage. The plurality of ribs are distributed about the fluid cooling passage as an array of ribs. The rail is disposed at the plurality of ribs. The rail extends about the fluid cooling passage. The side housing body and the rail form a plurality of fluid cooling channels connected in fluid communication with the fluid cooling passage.

Abstract: A rotary engine rotor and a method for supplying rotary engine rotors are provided. The rotary engine rotor has a center axis, a rotor body, a plurality of peripheral surfaces, and a plurality of recess members. The rotor body has a central bore that extends along the center axis. The plurality of peripheral surfaces are disposed around an exterior perimeter of the rotor body. Each of the plurality of recess members are disposed in a respective one of the plurality of peripheral surfaces. Each recess member includes an interior cavity open to an exterior of the rotor.

Abstract: A heat exchanger assembly for an aircraft includes a housing, a heat exchanger, and a nozzle. The housing includes an inner flow path wall, an outer flow path wall, an inner inflatable diaphragm, and an outer inflatable diaphragm. The inner flow path wall and the outer flow path wall form an air inlet and an air outlet. The inner inflatable diaphragm and the outer inflatable diaphragm form a venturi channel between the air inlet and the air outlet. The heat exchanger is disposed downstream of the venturi channel within the air flow path. The nozzle includes a nozzle housing disposed within the air flow path. The nozzle housing forms a nozzle inlet and a nozzle outlet. The nozzle outlet is disposed at the venturi channel.

Abstract: A reduction gearbox is provided that includes a sun gear, planet gear assemblies, first and second ring gears, and a ring gear positional alignment system. Each planet gear assembly includes a main gear and first and second lateral gears coupled to one another. The ring gear positional alignment system includes a system controller, a ring gear actuator, and a sensor. The ring gear actuator is engaged with the first ring gear and is configured to rotate the first ring gear relative to the second ring gear. The system controller is in communication with the sensor, the ring gear actuator, and a non-transitory memory storing instructions. The instructions when executed cause the system controller to receive and process signals from the sensor relating to the position of the first ring gear, and control the ring gear actuator to selectively position the first ring gear using the signals from the sensor.

Abstract: A test rotor blade includes a blade body and a plug. The blade body extends between and to a leading edge and a trailing edge. The blade body forms a first side surface and a second side surface. The blade body forms a hole. The blade body includes a first side wall and a second side wall. The first side wall extends between and to the hole and the first side surface. The second side wall extends between and to the hole and the second side surface. The plug is disposed within the hole. The plug forms a perimeter surface extending about the axial centerline. The perimeter surface includes a sealing surface portion and a diverging surface portion. The diverging surface portion forms a gap between the plug and the blade body. The gap is disposed between the first side wall and the diverging surface portion.

Abstract: An engine assembly includes a propulsor, a gearbox, an engine, a gas turbine engine, and an electric control assembly. The gearbox includes a gear assembly coupled with the propulsor. The engine includes an air inlet, an exhaust outlet, and an engine output shaft. The engine output shaft is coupled with the gear assembly and configured to drive rotation of the propulsor through the gear assembly. The gas turbine engine includes a first rotational assembly, a compressor section, a turbine section, and a combustor. The compressor section is connected to the air inlet. The combustor is connected to the exhaust outlet. The combustor is configured to direct a combustion gas through the turbine section to drive rotation of the first rotational assembly. The electric control assembly includes a first electric motor. The first electric motor is coupled with the gear assembly. The first electric motor is configured to selectively apply a rotational force to the gear assembly to further drive rotation of the propulsor.

Abstract: A planetary gearbox is provided that includes a sun gear, a plurality of planet gear assemblies, each planet gear assembly having a main gear meshed with the sun gear, a fore lateral gear and an aft lateral gear disposed on opposite sides of the main gear and rotating therewith, a planet carrier rotatably supporting at least some of the planet gear assemblies, and at least one fore ring gear meshed with the fore lateral gears, at least one aft ring gear meshed with the aft lateral gears, wherein one of the sun gear, the planet carrier, and the ring gears is configured to be operatively connected to an input, one is configured to be operatively connected to an output, and rotation of a remaining one is limited.

Abstract: An assembly for an aircraft powerplant includes a first component, a second component, an annular seal element and a third component. The first component includes a first land surface which extends axially along and circumferentially about an axis. The second component includes a second land surface radially opposite the first land surface. The second land surface extends axially along and circumferentially about the axis. The annular seal element extends radially between a first side and a second side. The annular seal element includes a base, a plurality of ribs and a groove. The base includes a seal element surface disposed at the first side that radially engages the first land surface. The ribs are disposed at the second side. Each rib projects out from the base and radially engages the second land surface. The groove projects partially axially into the base. The third component projects axially into the groove.