Abstract: There are described methods and systems for operating an aircraft having two or more engines. One method comprises operating the two or more engines of the aircraft in an asymmetric operating regime, wherein a first of the engines is in an active mode to provide motive power to the aircraft and a second of the engines is in a standby mode to provide substantially no motive power to the aircraft; governing the first engine in the active mode using a first governing logic; and governing the second engine in the standby mode using a second governing logic, the second governing logic based on a target compressor speed and variable geometry mechanism (VGM) settings that are adjusted using trim values dependent on at least one parameter of the second engine in the standby mode.

Abstract: A filter housing for containing a cartridge assembly having a filter cartridge secured to a filter cover, the filter housing has a peripheral wall extending around a longitudinal axis, and end wall secured to the peripheral wall, the filter housing defining a cavity partially enclosed by the peripheral wall and the end wall, the peripheral wall defining a cartridge-receiving section extending from the end wall along the longitudinal axis and sized for receiving the filter cartridge and a cover-receiving section extending along the longitudinal axis from the cartridge-receiving section to an open end of the filter housing opposed to the end wall and sized for receiving the filter cover, a portion of the cover-receiving section recessed radially outwardly from the cartridge-receiving section relative to the longitudinal axis, a lubricant outlet of the filter housing extending through the portion of the cover-receiving section of the peripheral wall.

Abstract: A vapor-to-air heat exchanger for an aircraft powerplant that includes: a pressurized vapor source supplying vapor; and a condenser including a condenser inlet in fluid communication with the pressurized vapor source to receive the vapor, a condenser outlet, and at least one pneumatic vessel defining a cavity in fluid communication between the condenser inlet and the condenser outlet. The pneumatic vessel is reversibly inflatable to be configurable between a collapsed vessel configuration and an inflated vessel configuration. A volume of the cavity is greater in the inflated vessel configuration than in the collapsed vessel configuration. The pneumatic vessel is inflatable from the collapsed vessel configuration to the inflated vessel configuration when the cavity is pressurized by the vapor.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a turbine engine structure and a sensor system. The sensor system includes a probe and an optical sensor. The probe is connected to the turbine engine structure. The probe projects into a gas path of the turbine engine. The sensor system is configured to measure a temperature of the probe using the optical sensor.

Abstract: A valve for an air system in an aircraft engine, comprising: a housing defining a chamber having a valve axis; a body within the chamber about a piston axis collinear with the valve axis, extending from a first surface to a second surface, defining a bore extending from the first to the second surface, having a mating connector defined by the second surface and located radially outward of the bore relative to the piston axis; and a sleeve extending from a first end matingly engaged with the mating connector to a second end along a sleeve axis collinear with the valve axis, the first end axially stacked on the body via the first surface to define a first distance between the first end and the first surface, and via the second surface to define a second distance between the first end and the second surface greater than the first distance.

Abstract: There is described a method of operating a multi-engine system of an helicopter. The multi-engine system has a first turboshaft engine having a first shaft, a second turboshaft engine having a second shaft, a gearbox having a clutch system, and a range of rotation speeds defined as a placarded zone. The method generally has rotating the first shaft at a flight rotation speed when clutched and rotating the second shaft at a first idle rotation speed when unclutched, the first idle rotation speed above the placarded zone; decreasing a rotation speed of the first shaft from the flight rotation speed to a given rotation speed within the placarded zone; decreasing a rotation speed of the second shaft to the given rotation speed; clutching the second shaft; and decreasing the rotation speeds of the first and second shafts to a second idle rotation speed below the placarded zone.

Abstract: An aircraft engine, has: an upstream stator having upstream stator vanes circumferentially distributed about a central axis; and a downstream stator having downstream stator vanes circumferentially distributed about the central axis, the downstream stator located downstream of the upstream stator relative to an airflow flowing within a core gaspath of the aircraft engine, a number of the upstream stator vanes being different than a number of the downstream stator vanes, the downstream stator vanes including: a first vane made of a first material, a major portion of a leading edge of the first vane circumferentially overlapped by one of the upstream stator vanes, and a second vane made of a second material having a greater stiffness, strength, and/or ductility than that of the first material, a major portion of a leading edge of the second vane exposed via a spacing defined between two of the upstream stator vanes.

Abstract: An integrally bladed rotor (IBR) for a gas turbine engine and method is provided. The IBR is configured for use in blade off testing and includes a hub, a plurality of rotor blades, a central passage, and first and second lateral cavities. The hub has forward and aft ends and a circumferentially extending exterior surface. The central passage is disposed in the hub radially below a test rotor blade, extending along a path between an inlet at or forward of the test blade leading edge and an outlet at or aft of the test blade trailing edge. The first and second lateral cavities are disposed in the hub, extending generally parallel to the central passage path, on opposite circumferential sides. The first lateral cavity is disposed a distance (MSD1) from the central passage and the second lateral cavity is disposed a distance (MSD2) from the central passage.

Abstract: A fuel assembly for a gas turbine engine includes a fuel supply tube, a fuel port, a fuel manifold, and a fuel manifold adapter. The fuel supply tube is configured to convey a fuel. The fuel port is fluidly coupled to the fuel supply tube and configured to receive the fuel from the fuel supply tube. The fuel manifold includes a fuel inlet and a plurality of fuel outlets. The fuel inlet is fluidly coupled to the fuel port and configured to receive the fuel from the fuel port. The fuel manifold adapter includes a first mount portion and a second mount portion. The first mount portion is connected to the fuel port. The first mount portion is moveable relative to the second mount portion.

Abstract: A method includes using an electric motor to start a thermal engine. The electric motor and thermal engine are connected to one another as the electric motor and the thermal engine of a hybrid-electric power plant in an aircraft. The electric motor can be connected to a combining gear box. The thermal engine can be connected to the combining gear box so that the electric motor and the thermal engine can provide torque to the combining gear box in a parallel hybrid-electric configuration. The combining gearbox can output torque to an air mover for providing thrust to the aircraft.

Abstract: A system is provided for an aircraft. This aircraft system includes a gas turbine engine and a sensor system. The gas turbine engine includes an inlet and a compressor section. A flowpath projects radially inward into the gas turbine engine from the inlet and extends through the compressor section. The sensor system includes a plurality of static pressure sensors at least partially within the flowpath. The sensor system is configured to determine a total pressure characteristic within the flowpath using the plurality of static pressure sensors.

Abstract: A labyrinth seal assembly comprising: an outer seal having a radially inner cavity surface, and an axial cavity surface inward of the radially inner cavity surface; a floating seal received by the outer seal so as to be movable in a radial direction relative to the outer seal, the floating seal having a radially outer seal surface spaced inwardly from the radially inner cavity surface, a radially inner surface inward of the radially outer seal surface, and an axial seal surface bearing against the axial cavity surface; and an inner seal received by the floating seal, the inner seal having a radially outer surface spaced inwardly from the radially inner surface, a first one of the radially outer surface and the radially inner surface defining one tooth projecting toward a second one of the radially outer surface and the radially inner surface to inward of the axial seal surface.

Abstract: A method of operating a reverse thrust system of an aircraft engine, the method comprising: receiving a status signal indicative that the aircraft is on-ground or in-flight; and upon detecting that the aircraft is on-ground, overriding a protection module such that the reverse thrust system is operable regardless of the protection module being in an active state or in a disabled state, the protection module causing, absent the overriding, the reverse thrust system to be inoperable when in the active state. An aircraft comprising: an engine including a reverse thrust system; a thrust control input device configured for generating an input signal indicative of a reverse thrust demand; at least one sensor configured for generating at least one status signal indicative of the aircraft being in-flight or on-ground; and a control system electronically connected with the at least one sensor, the thrust control input device and the reverse thrust system.

Abstract: There is described a multi-channel, multi-range transducer for measuring a parameter, the transducer comprising N channels and M sensing elements, the M sensing elements centered on distinct calibration points of distinct measuring ranges, the M sensing elements distributed across the N channels of the transducer, wherein sensing elements having adjacent measuring ranges are provided in different ones of the N channels.

Abstract: A method of manufacturing a shroud segment for gas turbine engine includes providing an insert having a plurality of pins that extend into a platform cavity portion of a mold cavity. A powder injection molding feedstock is injected. When the green part is disengaged from the mold, each elongated feature is slid out of the green part to define a respective elongated cooling passage in the platform. The method may include, after debinding and sintering, projecting a coating material while defining an obstruction between source of coating material and the open end of each elongated feature with a shoulder of the element to prevent the coating material from reaching the open end, followed by machining to remove at least a part of the shoulder.

Abstract: During an inspection method, an inspection scope is inserted into an interior of an apparatus of a gas turbine engine. An image of the interior of the apparatus is captured to provide/output image data. A location of the inspection scope within the interior of the apparatus when the image is captured is determined to provide location data, where the determining of the location includes comparing the image data to model data from a model of the apparatus. During the course of inspection, movement of the inspection scope is tracked within the interior of the apparatus using the location data. The location data may be derived by feature recognition and/or motion tracking of live feed from a scope video output to features available in a field of view within duplicate CAD model.

Abstract: An aircraft engine oil filler apparatus includes a filler tube configured to be connected to an oil tank such that a bottom portion of the filler tube is disposed inside the oil tank, a valve received in the bottom portion of the filler tube movable between an open position in which the valve hydraulically connects the filler tube to the oil tank, and a closed position in which the valve hydraulically disconnects the filler tube from the oil tank, and a float disposed above the valve and operatively connected to the valve to move the valve from the open position to the closed position when oil inside the oil tank rises to a threshold level. The valve is movable independently from the float when pressure in the oil tank is greater than pressure in the filler tube. A method of operation of an oil filler apparatus is also described.

Abstract: An air intake for an aircraft propulsion system includes an air inlet duct, a core flow duct, a bypass flow duct, a splitter, and a flow control device. The air inlet duct includes an intake inlet and a gas path floor. The core flow duct includes a core flow outlet. The bypass flow duct includes a bypass flow outlet. The bypass flow duct includes the gas path floor. The splitter separates the core flow duct and the bypass flow duct. The flow control device is disposed on a portion of the gas path floor. The flow control device is configured to be selectively positioned to control an air flow path for air flowing through the air inlet duct, the core flow duct, and the bypass flow duct.

Abstract: A bearing housing cover for a gas turbine engine having an impeller. The bearing housing comprises an annular collar having a flange mountable to an exterior surface of a bearing housing. The bearing housing further comprises an impeller baffle integrated with the annular collar, the impeller baffle having an annular body with a front baffle face positionable adjacent a rear face of the impeller, a rear baffle face having a plurality of stiffening elements extending between the rear baffle face and the annular collar, and a central baffle opening.

Abstract: A seal device includes a first sidewall, a second sidewall and a carbon seal element. The first sidewall is arranged with a first tool die which centers the first sidewall about an axis. The first sidewall axially contacts the first tool die along the axis. The carbon seal element is arranged with a centering element which centers the carbon seal element about the axis. The carbon seal element circumscribes and radially contacts the centering element. The second sidewall is arranged with a second tool die which centers the second sidewall about the axis. The second sidewall axially contacts the second tool die along the axis. The first tool die and the second tool die are moved axially together along the axis to press fit the second sidewall into a bore of the first sidewall and axially capture the carbon seal element between the first sidewall and the second sidewall.