Pratt-whitney Canada Corp

Jivan, Mihai

Hervé Turcotte

Michel Desjardins

Gabriel Naccache

Guy Lefebvre

Marc Tardif

Adam Logan

Roger Huppe

Oleg Morenko

Remy Synnott

Jean Dubreuil

Ghislain Plante

Kenneth D Parkman

Pierre Gauvin

Paul Weaver

Patrick Valois

Kevin Ng

Gavin Rohiteshwar Kisun

Ryan Miskie

Simon Baker-Ostiguy

Nicholas Guglielmin

Daniel Trottier

John Wilson

Daniel Summers-Lepine

Lubricant pump systems and associated methods for aircraft engines are provided. The method includes receiving an input torque, dividing the input torque between a first load path receiving a first portion of the input torque, and a second load path receiving a second portion of the input torque. A first lubricant pump of the aircraft engine is driven via the first load path using the first portion of the input torque. A second lubricant pump of the aircraft engine is driven via the second load path using the second portion of the input torque. When a malfunction of the second lubricant pump occurs, the method includes ceasing to drive the first lubricant pump and the second lubricant pump using the input torque.

Lubricant pump system and method for aircraft engine

A hybrid multi-spool gas turbine engine, has: a LP spool and a HP spool rotatable about a central axis, the LP spool having an LP compressor and an LP turbine engaged to the LP compressor via an LP shaft, the LP shaft engaged to a rotatable load at a first end thereof, the HP spool having an HP turbine and an HP compressor engaged to the HP turbine via a HP shaft; an accessory gearbox (AGB) engaged to both of the LP shaft and the HP shaft and located proximate a second end thereof, the AGB having at least one accessory output drivingly engageable to at least one accessory and at least one input; and at least one electric motor drivingly engaged to the at least one input of the AGB, the at least one electric motor drivingly engaged to the rotatable load via the AGB and the LP shaft.

Hybrid gas turbine engine

A reduction gearbox is provided that includes a planetary gear arrangement disposed in a casing and a lubrication handling system. The planetary gear arrangement includes a sun gear, planet gear assemblies, and independent fore and aft ring gears. Each planet gear assembly has fore and aft lateral gears respectively disposed on opposite sides of a main gear and rotating therewith. The main gear has a main gear pitch diameter and the fore and aft lateral gears having a lateral gear pitch diameter. The fore ring gear is meshed with the fore lateral gears. The aft ring gear is meshed with the aft lateral gears. The planet gear assemblies are in communication with an output. The lubrication handling system includes a fore ring gear baffle and an aft ring gear baffle.

Reduction gearbox for gas turbine engine with lubrication control system

An aircraft engine has: a compressor section having a compressor rotor rotatable about a central axis; a diffuser downstream of the compressor rotor, the diffuser including a diffuser ring extending circumferentially around the central axis; a bearing housing secured to the diffuser ring, the bearing housing contained within a volume located radially inwardly of the diffuser ring; and an air manifold secured to the diffuser ring, the air manifold defining inlets in fluid flow communication with the compressor section and an outlet in fluid flow communication with the volume.

Diffuser ring with air manifold

An apparatus of an aircraft engine having lubricated moving components is provided. The apparatus comprises a housing having a lubricant cavity for receiving lubricant and a service port cavity, the housing defining a selectively openable service port for accessing the service port cavity. A shaft is mounted for rotation about a rotational axis to the housing. A seal is mounted about the shaft forming a partition between the lubricant cavity and the service port cavity. A vent channel extends between the lubricant cavity and the service port cavity.

Sealing arrangement with vent for an engine component with a service port

Cooling a torque probe involves an elongated sleeve extending along a sleeve axis between a first sleeve end and a second sleeve end. The sleeve is mountable about the torque probe to define a flow passage between an inner surface of the sleeve and an outer surface of the torque probe. The flow passage is in fluid communication with a flow passage inlet of the sleeve and with a flow passage outlet of the sleeve spaced apart along the sleeve axis from the flow passage inlet. A cooling airflow is configured to flow into the flow passage via the flow passage inlet, through the flow passage along the outer surface of the torque probe to cool the torque probe, and out of the flow passage via the flow passage outlet.

Torque probe cooling for gas turbine engine using external air

An assembly comprising: a nut having a cylindrical body with a threaded external surface, an axially extending interior tool socket, and a radially outwardly extending flange at a proximal end of the cylindrical body; a first component including a hollow cylindrical tube with a connector end having a radially inwardly extending shoulder, the radially inwardly extending shoulder abutting the radially outwardly extending flange of the nut in an assembled position; and a second component including an annular hub with a threaded internal surface matching the threaded external surface of the nut, and an annular mounting surface abutting the connector end of the propeller shaft in the assembled position.

Nut locking feature

A thermal shielding arrangement for a turbine probe comprises a heat shield having first and second mating portions axially engaged in overlapping relationship around a probe extending through an air cavity between an exhaust case and a turbine housing. The first mating portion is provided on a radially outer surface of the turbine housing. The second mating portion projects radially inwardly from a radially inner surface of the exhaust case.

Thermal protection for a gas turbine engine probe

Cooling a torque probe involves an elongated sleeve extending along a sleeve axis between a first sleeve end and a second sleeve end. The sleeve is mountable about the torque probe to define a flow passage between an inner surface of the sleeve and an outer surface of the torque probe. The flow passage is in fluid communication with a flow passage inlet of the sleeve and with a flow passage outlet of the sleeve spaced apart along the sleeve axis from the flow passage inlet. A cooling fluid is configured to flow into the flow passage via the flow passage inlet, through the flow passage along the outer surface of the torque probe to cool the torque probe, and out of the flow passage via the flow passage outlet.

Torque probe cooling for gas turbine engine using internal fluid flow

A thermal shielding arrangement for a turbine probe extending through an exhaust case and a turbine housing of a gas turbine engine comprises a sleeve having a radially inner end fixedly mounted to the turbine housing and a radially outer end floatingly received in a probe boss on the exhaust case. The radially outer end of the sleeve opens to an air cavity between the exhaust case and a surrounding engine nacelle to allow for air circulation around the probe.

Turbine probe heat shielding

A planetary gear train for an aircraft engine, has a sun gear rotatable about a rotation axis, planet gears meshed with the sun gear and circumferentially distributed about the rotation axis, and a ring gear meshed with the planet gears, a planet carrier having first and second disk members axially offset from one another relative to the rotation axis, the planet carrier having planet bearings secured to the first and second disk members and extending axially therebetween, the planet gears rotatably supported by the planet bearings, the planet carrier having connecting members secured to both of the first and second disk members, the connecting members located circumferentially between the planet bearings, a lubrication network extending within the planet carrier and in fluid flow communication with the planet bearing and extending within the connecting members, the connecting members defining outlets of the lubrication network, the outlets oriented toward the sun gear.

Epicyclic gear stage

An air supply system is configured to provide cooling air with reduced heat pickup to a turbine rotor of a gas turbine engine. The system comprises a first cooling passage extending between a hollow airfoil and an internal pipe extending through the airfoil. The airfoil extends through a hot gas path. A second cooling passage extends through the internal pipe. The coolant flowing through the second cooling passage is thermally isolated from the airfoil hot surface by the flow of coolant flowing through the first cooling passage. The first and second cooling passages have a common output flow to a rotor cavity of the turbine rotor where coolant flows from the first and second cooling passages combine according to a predetermined ratio.

Turbine rotor coolant supply system

The epicyclic gear stage has a carrier body receiving the planet gears via corresponding planet bearings, a first bearing plane where a first axial end of the planet bearings are received by the carrier body, a second bearing plane where a second axial end of the planet bearings are received by the carrier body, a torque transfer body made integral to the carrier body externally to the sun gear, at a balanced torque transfer plane located between the first bearing plane and the second bearing plane, and extending away therefrom to a shaft coupler, for transferring torque between the carrier body and a shaft, and an oil network extending within the carrier, the oil network having an inlet in the torque transfer body, extending within the torque transfer body and into the carrier body, and extending within the carrier body and into the planet bearings.

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Sainte-Julie, Canada

Hervé Turcotte