Abstract: An air duct interface is disclosed herein. The air duct interface includes a first air duct including a first opening and a first coupling flange disposed circumferentially around the first opening, a second air duct including a second opening and a second coupling flange circumferentially surrounding the second opening, a first clamping block coupled to the first coupling flange by a first bolt, wherein the second coupling flange is disposed between a portion of the first clamping block and the first coupling flange, and a first adjusting block coupled to the first coupling flange by a second bolt, the first adjusting block engaging the first coupling flange and the second coupling flange.

Abstract: A propulsion system for an aircraft includes a gas turbine engine and an engine power extraction system. The gas turbine engine includes a first rotational assembly. The first rotational assembly includes a propeller and a first shaft coupled to the propeller. The first shaft is configured for rotation about a rotational axis of the propulsion system. The engine power extraction system includes a fluid pump, a fluid motor, and a generator. The fluid pump is operably coupled to the first shaft. The fluid pump is in fluid communication with the fluid motor. The fluid pump is configured to pressurize and direct a hydraulic fluid to the fluid motor causing the fluid motor to rotatably drive the generator to generate an electrical power output.

Abstract: A method of controlling an aircraft propulsion system in an aircraft susceptible to ground resonance is provided. The method includes: providing an aircraft having an aircraft propulsion system and a rotary blade propulsion unit, the aircraft propulsion system having an electric motor and a gearbox, wherein the electric motor is in drive communication with the gearbox and the gearbox is in drive communication with the rotary blade propulsion unit; determining the presence or absence of a parameter indicative of a ground resonance condition; and controlling the electric motor to deliver an amount of torque to the gearbox to drive the rotary blade propulsion unit to produce enough lift to make the aircraft airborne when the presence of the parameter indicative of the ground resonance condition is determined.

Abstract: An aircraft system is provided that includes a first propulsor rotor, a first transmission, a second propulsor rotor, a second transmission and an intermittent combustion engine. The first propulsor rotor is rotatable about a first propulsor axis. The first transmission is coupled to the first propulsor rotor. The second propulsor rotor is rotatable about a second propulsor axis. The second transmission is coupled to the second propulsor rotor. The intermittent combustion engine is configured to drive rotation of the first propulsor rotor through the first transmission. The intermittent combustion engine is configured to drive rotation of the second propulsor rotor through the second transmission.

Abstract: An apparatus comprises a combustion chamber of a gas turbine engine, a primary manifold for injecting fuel into the combustion chamber via a first plurality of nozzles and a secondary manifold for injection fuel into the combustion chamber via a second plurality of nozzles. A fuel control system provides fuel flow to the primary manifold and to the secondary manifold. The fuel control system determines at least one control parameter responsive to a first point at which the fuel flow level causes the secondary manifold to start draining and a second point at which an acceleration request is made and selects one of a plurality of secondary manifold filling profiles responsive to the determined at least one control parameter.

Abstract: An airfoil is provided that includes a first end, a second end, a leading edge, a trailing edge, a pressure side and a suction side. The leading edge and the trailing edge are joined by the pressure side and the suction side to provide an exterior airfoil surface extending in a spanwise direction from the first end of the airfoil to the second end of the airfoil. The exterior airfoil surface are formed in conformance with a plurality of cross-section profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by the local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading edge and the trailing edge at the span location.

Abstract: A system is provided for an aircraft. This aircraft system includes a powertrain, a lubrication system and a casing. The powertrain includes a powerplant, a power transmission, a clutch system and a bladed rotor. The power transmission operatively couples the powerplant to the clutch system. The clutch system operatively couples the power transmission to the bladed rotor. The powerplant is configured to drive rotation of the bladed rotor through the power transmission and the clutch system when the clutch system is engaged. The lubrication system includes a lubricant reservoir and a lubricant circuit. The lubrication system is configured to direct lubricant from the lubricant reservoir to the powerplant and the power transmission through the lubricant circuit. The casing houses the power transmission and the lubricant reservoir.

Abstract: An assembly for an aircraft includes an aircraft housing, an auxiliary power unit (APU), an oil system, a door actuator, and a controller. The aircraft housing forms a compartment. The aircraft housing includes an intake door. The intake door is movable between a closed position, an open position, and intermediate positions between the closed position and the open position. The APU is disposed within the compartment. The APU includes an engine. The engine includes an air inlet and a rotational assembly. The oil pump is operatively connected to the rotational assembly. The door actuator is operatively connected to the intake door. The door actuator is operable to position the intake door in the closed position, the open position, and the intermediate positions to control an ambient air flow to the air inlet. The controller is operatively connected to the door actuator.

Abstract: A containment case includes an annular body bound by an inner surface and an outer surface and an outer surface feature extending outward from the outer surface and an inner surface feature extending inwards from the inner surface. The annular body includes a containment section and an inner reinforcement section radially inward from the inner surface subtending a sector of the containment section. The reinforcement section includes a reinforcement thickness that is greater than a casing thickness of the containment section at least partially coinciding with the outer surface feature.

Abstract: A powerplant is provided for an aircraft. This aircraft powerplant includes an aircraft engine, and the aircraft ending includes a shaft and an anode. The shaft is configured to rotate about an axis. The shaft is configured from or otherwise includes a shaft material with a first anodic potential. The anode is mounted to and electrically coupled with the shaft. The anode is configured from or otherwise includes an anode material with a second anodic potential that is greater than the first anodic potential.

Abstract: A variable guide vane assembly for a gas turbine engine stator is provided. The variable guide vane assembly includes a plurality of vanes and a plurality of RT mechanisms. The vanes extend between a shroud and hub. The vanes are circumferentially disposed and spaced apart from one another. Each vane includes inner and outer radial ends, and inner and outer radial posts. Each vane is pivotally mounted to rotate about its rotational axis. Each RT mechanism is in communication with the inner or outer radial post of a respective vane. The RT mechanism includes a pin connected to the vane that is disposed in a ramp slot non-rotational relative to the pivotable vane. The ramp slot extends between first and second lengthwise ends. Rotation of the vane relative to the ramp slot causes the pin to travel within the ramp slot and the vane to translate linearly.

Abstract: A method is provided during which first braze powder is deposited with a substrate. The first braze powder is sintered to the substrate during the depositing of the first braze powder to provide the substrate with sintered first braze material. Second braze powder is deposited with the substrate. The second braze powder is different than the first braze powder. The second braze powder is sintered to the substrate during the depositing of the second braze powder to provide the substrate with sintered second braze material. The sintered first braze material and the sintered second braze material are heated to melt the sintered first braze material and the sintered second braze material and to diffusion bond the sintered first braze material and the sintered second braze material to the substrate.

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.