Abstract: A flowmeter for gaseous fluid includes a conduit composed of non-electrically conductive material for passage of an ionized flow of the gaseous fluid therethrough. The flowmeter further includes an electromagnetic sensor arranged to measure a magnetic field generated about the conduit by the passage of the ionized flow and generate a signal proportional to the magnetic field.

Abstract: A fluid system of an aircraft engine includes a first and second fluid conduits. The first conduit fluid conduit fluidly connects to the aircraft engine. A pump fluidly connects to the first fluid conduit and moves therethrough a fluid imparting an excitation frequency to the fluid. The second fluid conduit fluidly connects to the first fluid conduit and to a pressure sensor. A weight is disposed in the second fluid conduit and movable between a first and second positions. A mass of the weight and a liquid capacitance of the second fluid conduit between the second position and a second end of the second fluid conduit define a resonance frequency that is less than the excitation frequency. Fluid pressure monitoring and pressure noise filter construction methods are also disclosed.

Abstract: An aircraft propulsion system is provided that includes a turbine engine and an exhaust gas condenser. The exhaust gas condenser includes a housing, a nozzle, a plurality of air scoops, and a plurality of exhaust gas conduit banks. The housing has an interior cavity, a bottom side, and first and second lateral sides. The exhaust gas conduit banks are disposed in the interior cavity of the housing and are configured to form an interior bypass air chamber. Each exhaust gas conduit bank includes alternating exhaust gas conduits and bypass air passages. The exhaust gas conduits extend axially between the forward and aft ends, and are configured to conduct exhaust gases to the nozzle. The bypass air passages are configured to receive bypass air from the air scoops and direct the bypass air between the exhaust gas conduits and into the interior bypass air chamber.

Abstract: A journal shaft for a journal bearing assembly of an aircraft engine includes a shaft body extending along a longitudinal axis. The shaft body has a radially outer surface and a radially inner surface radially spaced apart from the longitudinal axis to define an inner cavity. An oil pocket is defined in the radially outer surface. One or more passages extend through the shaft body from the radially inner surface to the oil pocket to provide fluid communication between the inner cavity and the oil pocket. The oil pocket includes a radially inner base surface and interconnecting transition surfaces extending between the radially inner base surface of the oil pocket and the radially outer surface of the shaft body. The interconnecting transition surfaces form a fluid-dynamically smooth and edgeless transition to the radially outer surface of the journal shaft.

Abstract: System and method for detecting an uncommanded high thrust (UHT) event in an aircraft. The method comprises enabling a UHT function associated with an engine when an enabling condition has been met. When the UHT function is enabled, the UHT event is detected when a power lever of the aircraft is at a given position, a parameter indicative of engine speed or power is above a first threshold, and a rate of change of the engine speed is above a second threshold. In response to detecting the UHT event, an alert is output to trigger accommodations to the UHT event for the engine.

Abstract: An engine system is provided that includes an engine core, a fuel containment structure, a fuel containment plenum, a fuel delivery system and a plenum purge circuit. The engine core includes an engine housing structure which houses a compressor section, a combustor section and a turbine section. The fuel containment structure is outboard of and connected to the engine housing structure. The fuel containment plenum is formed by and located between the fuel containment structure and the engine housing structure. The fuel delivery system includes a plurality of fuel injectors, a fuel supply line and a fuel manifold fluidly coupling the fuel supply line to the plurality of fuel injectors. The fuel manifold is arranged within the fuel containment plenum. The plenum purge circuit is configured to purge fuel, when leaked into the fuel containment plenum from the fuel delivery system, out of the fuel containment plenum.

Abstract: An assembly for an aircraft propulsion system includes a propeller, a blade position sensor, and a controller. The propeller is configured for rotation about a rotational axis. The propeller includes a plurality of propeller blades. The plurality of propeller blades have a pitch. Each propeller blade of the plurality of propeller blades is configured for rotation about a respective lengthwise axis to selectively vary the pitch. The blade position sensor is configured to measure the pitch. The blade position sensor is configured to generate a pitch output signal representative of the measured pitch. The controller is in signal communication with the blade position sensor. The controller includes a processor in communication with a non-transitory memory storing instructions, which instructions when executed by the processor, cause the processor to: detect a presence or an absence of an erroneous propeller blade pitch measurement using at least the pitch output signal.

Abstract: An aircraft propulsion system is provided that includes a turbine engine and an exhaust gas condenser. The exhaust gas condenser includes a housing, a nozzle, air scoops, a bank of exhaust gas conduits, and a center collection channel. The housing includes an interior cavity, top and bottom sides, and first and second lateral sides. The bank of exhaust gas conduits is disposed in the interior cavity of the housing and includes exhaust gas conduits and bypass air passages. The exhaust gas conduits and bypass air passages are disposed in an alternating configuration. The center collection channel is disposed radially inside of the bank of exhaust conduits. The exhaust gas conduits and bypass air passages extend spirally along the central axis between the forward and aft ends. Each bypass air passage is in fluid communication with an air scoop and the center collection channel.

Abstract: An aircraft propulsion system is provided that includes a turbine engine and an exhaust gas condenser. The exhaust gas condenser includes a housing, a nozzle, inlet and outlet air scoops, and an exhaust gas conduit bank. The housing has an interior cavity, top and bottom sides, and first and second lateral sides. The exhaust gas conduit bank is disposed in the interior cavity of the housing. The exhaust gas conduit bank includes a plurality of exhaust gas conduits and bypass air passages that are disposed in an alternating configuration. The exhaust gas conduits extend axially between the forward and aft ends, are open at the forward end and at the nozzle. The bypass air passages receive bypass air from the inlet air scoop and direct the bypass air laterally across the condenser to the outlet air scoop.

Abstract: A tool assembly for disassembling a crankshaft assembly with a main shaft received within a hollow extension shaft along a shaft axis and coupled thereto via an interference fit, the extension shaft having a fastener slot. The tool assembly includes a main tool body having a central bore, a main tool body inner surface circumscribing the central bore, and a main tool body outer surface having a fastener bore, the central bore dimensioned for axially receiving the extension shaft, with the fastener bore axially aligned with the extension shaft fastener slot. A fastener is concurrently fastenable to the main tool body via the fastener bore and the hollow extension shaft via the fastener slot. An actuator assembly with a linearly-displacing actuating element is coupled to the main tool body and applies a force, via the actuating element, against an end of the main shaft in a direction along the axis.

Abstract: A gas turbine engine apparatus includes an engine flowpath, a protuberance and a variable vane. The protuberance projects into the engine flowpath. The variable vane extends across the engine flowpath. 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. A recess extends spanwise into the airfoil from the first end. The airfoil, at the first end, is spaced from the protuberance when the variable vane is in the first position. The airfoil, at the first end, is aligned with the protuberance and the protuberance projects into the recess when the variable vane is in the second position.

Abstract: A method of operating an aircraft power plant having an interburner fluidly connected between a combustion engine and a turbine, includes: operating the combustion engine at an air-fuel ratio corresponding to a nominal air-fuel ratio to generate combustion gases; receiving the combustion gases generated by the combustion engine into the interburner; increasing a temperature of the combustion gases generated by a combustion chamber of the combustion engine by lowering the air-fuel ratio to an ignition air-fuel ratio lower than the nominal air-fuel ratio; injecting fuel into the interburner to cause ignition of a mixture of the combustion gases and the fuel received into the interburner; and after ignition of the interburner, increasing the air-fuel ratio above the ignition air-fuel ratio.

Abstract: A rotor assembly has: blades having airfoils and roots protruding from platform segments; a rotor disc having a peripheral face defining recesses, and slots, a recess located between two adjacent ones of the slots and bounded by a step; feather seals located radially between the peripheral face and the platform segments, a feather seal having a core extending from a trailing end to a leading end and overlapping a gap defined between two platform segments and tabs protruding from the core, the tabs including: trailing tabs positioned axially outside the recess; and leading tabs, a leading tab extending from a root to a tip and having one or more of: the tip axially positioned outside of the recess; and a fillet at an intersection between the tip and an edge of the leading tab, the edge extending between the tip and the core, and facing the step.

Abstract: There is provided a method and a system for detecting and mitigating a propeller failure condition. An actual value of a rotational speed of the propeller and/or of a pitch angle of blades of the propeller is obtained. In response to determining that the speed is below a reference rotational speed for the propeller and/or determining that the pitch angle is above a pitch angle threshold, an actuator operatively connected to the blades is commanded to decrease the pitch angle to increase the speed towards the reference speed. After commanding of the actuator to decrease the pitch angle, a subsequent value of the speed and/or a subsequent value of the pitch angle is obtained. The actuator is commanded to hold the pitch angle in response to determining that the speed has failed to increase towards the reference speed and/or determining that the pitch angle has failed to decrease.

Abstract: A fuel delivery apparatus is provided for a gas turbine engine. This fuel delivery apparatus includes a fuel injector. The fuel injector includes a plurality of fuel passages, a guide passage, a plurality of air passages, a mixing cavity and a fuel injector outlet. The fuel passages extend along an axis to the guide passage. The fuel passages converge radially inwards towards the axis as the fuel passages spiral about the axis towards the guide passage. The guide passage turns radially outwards away from the axis as the guide passage extends from the fuel passages to the mixing cavity. The air passages converge radially inwards towards the axis as the air passages extend axially to the mixing cavity. The mixing cavity fluidly couples the guide passage and the air passages to the fuel injector outlet.

Abstract: The aircraft can include a rotary airfoil device having an output shaft rotatable around an airfoil rotation axis; an electric engine having a source shaft drivingly coupled to the output shaft; a controller in communication with the electric engine; a low voltage electric system operable to power the controller; a high voltage electric system having a high voltage battery operable to power the electric engine, the high voltage battery further operable to power the low voltage electric system via a voltage converter; and an electromagnetic generator having a generator shaft coupled to the output shaft, the electromagnetic generator operable to power the low voltage electric system.

Abstract: A compressor for an aircraft engine, has: a gaspath extending around a central axis between a radially-inner wall and a radially-outer wall; a rotor having blades circumferentially distributed around the central axis, the blades having airfoils extending from roots to tips along a span and extending from leading edges to trailing edges along a chord, the airfoils extending across the gaspath; grooves defined in an inner face of the radially-outer wall, circumferentially distributed around the central axis, and having groove lengths extending in a first direction having a circumferential component relative to the central axis, the grooves located between the leading edges and the trailing edges of the airfoils; and slots defined in the inner face of the radially-outer wall, circumferentially distributed around the central axis, and having slot lengths extending in a second direction having an axial component relative to the central axis, the slots projecting from the grooves.

Abstract: A hybrid propulsion system includes a heat engine configured to drive a heat engine shaft. An electric motor is configured to drive an electric motor shaft. A transmission system includes at least one gearbox. The transmission system is configured to receive rotational input power from each of the heat engine shaft and the electric motor shaft and to convert the rotation input power to output power.

Abstract: A method of determining an inlet total air pressure includes determining a first parameter indicative of a first inlet total air pressure. The method includes executing a sequence that includes: determining a mass air flow passing through the air inlet based on the first parameter, determining a Mach number of air passing through the air inlet based on the mass air flow, determining a static air pressure at the air inlet, determining an air pressure ratio based on the Mach number, generating a subsequent parameter indicative of the revised inlet total air pressure based on the air pressure ratio and the static air pressure, and substituting the subsequent parameter for the first parameter. The method includes executing at least one additional instance of the sequence with the subsequent parameter, and outputting the subsequent parameter as the inlet total air pressure.

Abstract: A blade for mounting on a rotor disc of a gas turbine engine includes platform segments between a airfoil and blade root and extending laterally from the blade, the platform segment having a curved profile defining a platform recess on a root side of the platform segment, the recess configured to cooperate with recesses of adjacent blades to define a blade pocket. A recessed region is defined in a rear surface of the blade root, and is configured to receive a complementary projection of the rotor disc upon mounting the blade on the rotor disc. The recessed region has a surface configured to abut a corresponding surface on the projection of the rotor disc to form a face seal. Methods of manufacturing such a rotor disc and of sealing a circumferential joint between such blades and the rotor disc are also disclosed.