Abstract: A de-aerator for a lubrication system, has: a housing defining an air-oil inlet, an oil outlet, and an air outlet of the de-aerator; a rotor received within the housing and rotatable relative to the housing about a central axis, the rotor having blades distributed about the central axis and extending at least partially radially relative to the central axis, flow passages extending between the blades, the rotor having a hub circumferentially extending around the central axis and around the blades, the hub having a peripheral wall oriented radially inwardly and defining a fore opening leading to the flow passages; and a gap between the housing and the hub of the rotor, a portion of the housing received within the fore opening and axially overlapping the peripheral wall of the hub, the gap having a fore gap inlet between the portion of the housing and the peripheral wall of the hub.

Abstract: There is described a method of detecting failure in an array of thermocouples connected in parallel. The method comprising: during an operation mode of the array of thermocouples, measuring a voltage V across the array of thermocouples, the voltage V associated with a temperature T; during a failure detection mode of the array of thermocouples, shunting the array of thermocouples, and measuring a shunt voltage Vs occurring across a resistive element connected in series with the array of thermocouples; comparing the shunt voltage Vs to an expected shunt voltage Vs_exp for the array of thermocouples at the temperature T; and generating a failure signal indicative of a detected failure in the array of thermocouples when the shunt voltage Vs deviates from the expected shunt voltage Vs_exp by more than a deviation threshold.

Abstract: A low pressure compressor section of a gas turbine engine includes low pressure compressor stages spaced apart along an axis of rotation of the low pressure compressor section. The low pressure compressor section includes at least one of a fluid conduit extending between an air inlet of the gas turbine engine and the low pressure compression stages. The fluid conduit having at least one flow diverter displaceable between a first position and a second position to modulate air through the low pressure compressor stages. And/or a compressor stage disabler is engageable with at least one of the low pressure compressor stages and configured to reduce a rotation thereof.

Abstract: A heating system for heating a component in a liquid distribution system of an aircraft engine. The liquid distribution system feeds a liquid to the component. The heating system includes an acoustic generator disposed in communication with the component via a liquid passage of the liquid distribution system. The liquid passage defines a length between the acoustic generator and the component. The acoustic generator generates a resonant frequency selected as a function of the length of the liquid passage to generate a standing wave in the liquid within the liquid passage. The standing wave transmitting energy to the component to heat the component.

Abstract: In accordance with at least one aspect of this disclosure, there is provided a heat exchange system. The heat exchange system includes a first heat exchanger and a second heat exchanger. The first heat exchanger includes an engine fluid conduit fluidly connecting an engine fluid inlet to an engine fluid outlet. A first internal buffer fluid conduit fluidly connects a first buffer fluid inlet to a first buffer fluid outlet where the engine fluid conduit is in fluid isolation from the first internal buffer fluid conduit but is in thermal communication with the first internal buffer fluid conduit for heat exchange between the engine fluid and the buffer fluid.

Abstract: An air intake for an aircraft propulsion system includes an air inlet duct and a flow control device. The air inlet duct includes an interior surface and an intake inlet. The interior surface extends from the intake inlet. The interior surface forms and surrounds an inlet flow passage through the air inlet duct. The intake inlet includes a top side, a bottom side, a first lateral side, and a second lateral side. Each of the first lateral side and the second lateral side extend between and to the top side and the bottom side. The flow control device is disposed inside the air inlet duct at the first lateral side. The flow control device is configured to form an asymmetrical shape of the intake inlet at the first lateral side relative to the second lateral side.

Abstract: Methods and systems for operating an aircraft having two or more engines are described. The method comprises receiving an engine availability confirmation when a set of engine parameters meet engine operating conditions for an asymmetric operating regime, receiving an aircraft availability confirmation when a set of aircraft parameters meet aircraft operating conditions for the asymmetric operating regime, outputting an availability message to a cockpit of the aircraft in response to receiving the engine availability confirmation and the aircraft availability confirmation, receiving a pilot-initiated request to operate the engine in the asymmetric operating regime, and commanding the engines to operate in the asymmetric operating regime in response to the pilot-initiated request.

Abstract: Systems and methods for validating an inspection software are provided. A dataset which defines properties for a three-dimensional component having at least one known defect is obtained. The three-dimensional component is inspected with the inspection software using the dataset to obtain inspection results for the three-dimensional component. The inspection results are compared to reference results for the three-dimensional component. When the inspection results correspond to the reference results within a predetermined threshold, a signal indicative of validation of the inspection software is issued.

Abstract: A method is provided for operating an aircraft system. During this method, an engine is operated using first fuel provided by a first fuel source. A fuel supply for the engine is switched between the first fuel source and a second fuel source, where the switching of the fuel supply includes shutting down the engine during aircraft flight. The engine is operated using second fuel provided by the second fuel source.

Abstract: A method of positioning exits of diffuser pipes of a centrifugal compressor in an aircraft engine includes aligning an alignment tool with a structure supporting the diffuser pipes, and rotating each diffuser pipe about an inlet axis thereof to close a gap between a surface of the diffuser pipe and part of the alignment tool until the diffuser pipe abuts the part of the alignment tool. The alignment tool includes for example a body defining a center axis and having at least one tool datum configured for abutting against the casing, and a plurality of alignment members that are fixed to the body and extend radially from the body relative to the center axis.

Abstract: A locking tab washer for rotatably locking a bolt fastened to a component in an aircraft engine has a body with an opening and at least four locking tabs positioned about the opening and circumferentially spaced apart from one another at integer multiples of a common angle relative to the center of the opening. The locking tabs are bendable towards to rotatably lock the bolt. The locking tab washer also has a retaining feature positioned about the opening and operable to rotatably retain the locking tab washer to the component.

Abstract: An electric machine system described herein comprises a first electric machine rotor, a first gear and a second gear. The first gear is connected to the first electric machine rotor and to a first input/output shaft. The second gear is connected to the first electric machine rotor and to a second input/output shaft. The first electric machine rotor is disposed between the first gear and the second gear.

Abstract: A component configured for use in a gas turbine engine, including: a wall portion formed from sheet metal and extending between two end portions, the wall portion being stamped with a plurality of instances of a pre-defined three dimensional shape structured to increase a stiffness of the wall portion in a direction normal to the wall portion, each instance of the plurality of instances of the pre-defined three dimensional shape being orientated at 90 degrees with respect to each other adjacent instance of the plurality of instances, wherein the three dimensional shape has a pair of curved distal peripheral end portions that are at opposite ends of the three dimensional shape and each pair of curved distal peripheral end portions are connected to each other by a pair of opposing concave peripheral portions that curve towards each other at a mid-section.

Abstract: A heat exchanger, has: a heat exchanger core having an inlet, an outlet, at least one first conduit connecting the inlet to the outlet, and at least one second conduit in heat exchange relationship with the at least one first conduit; a bypass conduit connecting the inlet to the outlet while bypassing the at least one first conduit; and a valve in communication with the bypass conduit and having: a valve member movable between a closed position in which fluid communication through the valve and via the bypass conduit is blocked and an open position in which fluid communication through the bypass conduit and through the valve is permitted, the valve member defining an auxiliary passage, and a thermal fuse overlapping the auxiliary passage and blocking fluid communication through the valve member via the auxiliary passage, the thermal fuse having a melting temperature above a predetermined temperature of the fluid.

Abstract: A hybrid-electric powertrain system for aircraft includes a gearbox having a first rotary shaft for output to drive an air mover for aircraft thrust. The system includes a first prime mover connected by a second rotary shaft to the gearbox for power input to the gearbox. Further, the system includes a second prime mover connected by a third rotary shaft to the gearbox. The second prime mover can have a hollow core, and at least one of the first and second rotary shafts passes through the hollow core and the third rotary shaft.

Abstract: The method can include analyzing the composition of the particulate polymer debris, including ascertaining a presence of at least one main polymer element and a presence or absence of at least one signature polymer element in the particulate debris; establishing a correlation between the particulate polymer debris and a set of components, including matching the ascertained presence of the at least one main polymer element to a family of polymers from which the components of the set, including the source component, are made; and determining the source component amongst the components of the set, including matching the ascertained presence or absence of the at least one signature polymer element to a polymer composition of the source component.

Abstract: A propeller control unit (PCU) for controlling pitch angles of blades of a propeller, has: a pitch angle actuator; a servo valve hydraulically connected to the pitch angle actuator and to a first hydraulic fluid source; and a feather valve having a body movable within a cavity, the feather valve having a first actuation port and a second actuation port both in fluid communication with the cavity, the body between the first actuation port and the second actuation port, the body being movable to selectively hydraulically connect the pitch angle actuator to the servo valve through the feather valve or to hydraulically connect the pitch angle actuator to a drain line through the feather valve, the first actuation port and the second actuation port hydraulically connected to a second hydraulic fluid source independent from the first hydraulic fluid source.

Abstract: A gas turbine engine according to a first embodiment, having a case; a core within the case; an exhaust nozzle that is fluidly coupled to the case; a duct having a duct outlet that is fluidly coupled to the exhaust nozzle and a duct body extending from the duct outlet to a duct inlet; an air/oil cooler having an airflow outlet disposed at the duct inlet and a cooler body extending from the airflow outlet to an airflow inlet; and an airflow control baffle disposed at one of: the duct inlet; the duct outlet; and the airflow inlet of the cooler, wherein the airflow control baffle is configured to open when oil temperature in the engine is above a first threshold and close when the oil temperature is below a second threshold that is lower than the first threshold.

Abstract: A rotor for an aircraft rotary engine includes a rotor body. The rotor body extends about an axial centerline. The rotor body includes an outer body portion, an annular inner body portion, and ribs. The outer body portion forms a plurality of sides and a plurality of apex portions of the rotor. The annular inner body portion is disposed radially inward of the outer body portion. The ribs extend radially between and connect the outer body portion and the annular inner body portion. The rotor body forms at least a first lubrication passage within the inner body portion and within a first rib of the ribs. The first lubrication passage includes at least one passage inlet and at least one passage outlet. The at least one passage inlet is disposed at the annular inner body portion.

Abstract: A gas turbine engine apparatus includes a variable vane. The variable vane includes a pivot axis and an airfoil. The variable vane is configured to pivot about the pivot axis between a first position and a second position. The airfoil extends spanwise along a span line between a first end and a second end. The airfoil extends chordwise along a chord line between a leading edge and a trailing edge. The chord line is angularly offset from a reference plane containing the pivot axis by a twist angle. A first section of the airfoil is disposed at the first end. The twist angle varies as the first section extends spanwise along the span line. A second section of the airfoil is disposed spanwise between the first section and the second end. The twist angle is uniform as the second section extends spanwise along the span line.