Pratt-whitney Canada Corp

Nashed Youssef

Kevin Allan Dooley

Daniel Alecu

Andreas Eleftheriou

Richard Ullyott

Keith Morgan

Lazar Mitrovic

Jean Dubreuil

Joseph Horace Brand

Guy Bouchard

Farid Abrari

Andrew Marshall

Michael J Dowhan

Ilya Medvedev

Stephen Kenny

Darius Jehangir Karanjia

Gary Weir

Richard Harvey

William Savage

Michael Dowhan

David Menheere

The gas turbine engine rotor can have a body having a solid-of-revolution-shaped portion centered around a rotation axis, the body defining an annular cavity centered around the rotation axis, the annular cavity penetrating into the body from an annular opening, the annular cavity extending between two opposite annular wall portions each leading to a corresponding edge of the opening; and at least one structural plate mounted to and extending between the two opposite annular wall portions and forming an interference fit therewith.

Rotor for gas turbine engine

A method for providing an anti-icing airflow, including extracting a compressed airflow from a core flow path of an engine, heating the compressed airflow, mixing the heated compressed airflow with air extracted from a bypass flow path to create the anti-icing airflow having a higher temperature and pressure than that of the air extracted from the bypass flow path, and circulating the anti-icing airflow away from the bypass flow path. Also, an assembly located at least in part inside a turbofan engine and including a heat exchanger, a flow mixing device having a first inlet in the bypass flow path, a second inlet and an outlet, a first conduit providing fluid communication between the heat exchanger and a compressed air portion of the core flow path, a second conduit providing fluid communication between the heat exchanger and the second inlet, and a third conduit in fluid communication with the outlet.

Method and assembly for providing an anti-icing airflow

A gas turbine engine including at least one high pressure turbine rotor and at least one high pressure compressor rotor drivingly engaged to a rotatable high pressure spool, a low pressure spool rotatable independently of the high pressure spool, at least one low pressure turbine rotor drivingly engaged to the low pressure spool, and a rotatable load drivingly engaged to the low pressure spool. A fixed rotational speed ratio is defined between rotational speeds of the at high pressure turbine and compressor rotors. A fixed rotational speed ratio is defined between rotational speeds of the low pressure turbine rotor(s) and of the low pressure spool. A low pressure compressor rotor is in driving engagement with the low pressure spool through a variable transmission which defines a variable rotational speed ratio between the rotational speeds of the low pressure spool and of the low pressure compressor rotor.

Gas turbine engine with transmission and method of adjusting rotational speed

A gas turbine aircraft engine with power variability is provided. The gas turbine aircraft engine comprises a compressor and a turbine mounted on a common shaft, and, a continuously variable transmission coupled to the shaft for transmitting power to a propulsion load.

Gas turbine aircraft engine with power variability

The system is used for taxiing an aircraft and comprises at least one multi-spool gas turbine engine, the engine having an electrical motor in a torque-driving engagement with a low pressure spool of the engine. The low pressure spool has a propulsor connected thereon to generate thrust when rotated. A controller is connected to the electrical motor and an electrical power source to control an amount of electrical power provided from the power source to the electrical motor so as to drive the propulsor and cause at least a major portion of the thrust to be generated by the propulsor for moving the aircraft during taxiing.

Method and system for taxiing an aircraft

An aircraft propulsion system comprises a power plant driving a number of outboard propulsion units to propel a fixed wing aircraft.

Aircraft propulsion system

A two stage compressor for a gas turbine engine with a mixed flow first stage, a centrifugal second stage, and an intermediate duct. The mixed flow stage has a mixed rotor rotatable about the central compressor axis with a circumferential array of mixed flow blades between the mixed flow hub and an associated mixed flow shroud. The downstream centrifugal stage has a centrifugal rotor rotatable about the same compressor axis. The centrifugal rotor has a circumferential array of radially extending centrifugal flow blades between the centrifugal flow hub and an associated centrifugal flow shroud. An intermediate duct has an inner duct wall defining an axially curvilinear transition surface of revolution between an outlet end of the mixed flow hub and an inlet end of the centrifugal flow hub and an outer duct wall defining an axially curvilinear transition surface of revolution between an outlet end of the mixed flow shroud and an inlet end of the centrifugal flow shroud. Preferably, the intermediate duct has a median inlet radius greater than an intermediate duct median outlet radius whereby air flow from the mixed rotor is directed radially inwardly and axially rearwardly toward the centrifugal rotor.

Mixed flow and centrifugal compressor for gas turbine engine

Growing community of inventors

Mississauga, Canada

Nashed Youssef