Pratt-whitney Canada Corp

Pratt & Whitney Canada Cororation

Richard Ivakitch

Andreas Eleftheriou

David Harold Menheere

Paul Stone

David Denis

David Beamish

Peter Townsend

Ignatius Theratil

Krishna Prasad Balike

Aldo Abate

Czeslaw Wojtyczka

Daniel Fudge

Ronald Dutton

Milica Kojovic

Sean Kelly

Robert Venditti

Philippe Beauchesne-Martel

Mark Cunningham

Enzo Macchia

Daniel Poick

Philippe Bonniere

Alain Carrieres

Oleg Iskra

Mateusz Ożóg

Philip Ridyard

Guorong Yan

Damien Le Pape

Patrick Chiu

James O Brien

David Menheere

David H Menheere

Richard Chiu

Sean Michael Kelly

A variable pitch stator vane stage of a gas turbine engine includes a plurality of stator vanes extending across a flowpath of the gas turbine engine, and an actuation assembly operably connected to the plurality of start vanes. The actuation assembly includes an actuation ring extending circumferentially around an engine central longitudinal axis of the gas turbine engine, and a plurality of vane links. Each vane link operably connects a corresponding stator vane of the plurality of stator vanes to the actuation ring. A vane link of the plurality of vane links includes a link body, and a first link end connected the link body. The first link end includes a first spherical connection to the actuation ring. Circumferential movement of the actuation ring about the engine central longitudinal axis urges rotation of the plurality of vanes about their respective vane axes.

Joint for connection to variable guide vane of gas turbine engine

An assembly is provided for an aircraft motor. This assembly includes a bearing, a stationary structure and a flexible support. The bearing extends circumferentially around an axis. The stationary structure circumscribes the bearing. The flexible support is arranged radially between and radially engages the bearing and the stationary structure. The flexible support includes a first ring, a second ring, a plurality of bridges and a plurality of first ribs. The bridges are arranged circumferentially about the axis. Each of the bridges extends radially between and is connected to the first ring and the second ring. The first ribs are arranged circumferentially about the axis and interspersed with the bridges. Each of the first ribs projects radially out from the first ring towards the second ring. Each of the first ribs is connected to the first ring and is disengaged from the second ring.

Flexible support for aircraft motor bearing

A bleed arrangement of a gas turbine engine of an aircraft includes a core casing having one or more bleed openings define therein. The core case separates a core flowpath from a bypass flowpath of the gas turbine engine. A slider is positioned radially inboard of a case inner surface of the core casing relative to an engine central longitudinal axis. The slider is configured to be moved circumferentially between a closed position blocking the one or more bleed openings and an open position in which at least a portion of a core airflow is diverted from the core flowpath into the bypass flowpath as bleed airflow.

Gas turbine bleed off valve

An assembly for an aircraft powerplant includes a first component, a second component, an annular seal element and a third component. The first component includes a first land surface which extends axially along and circumferentially about an axis. The second component includes a second land surface radially opposite the first land surface. The second land surface extends axially along and circumferentially about the axis. The annular seal element extends radially between a first side and a second side. The annular seal element includes a base, a plurality of ribs and a groove. The base includes a seal element surface disposed at the first side that radially engages the first land surface. The ribs are disposed at the second side. Each rib projects out from the base and radially engages the second land surface. The groove projects partially axially into the base. The third component projects axially into the groove.

Ribbed seal element for aircraft powerplant

An aircraft engine comprises an air inlet duct extending from an inlet through which air is configured to enter the aircraft engine, to an outlet within the aircraft engine. At least one strut has a leading edge and a trailing edge and extends across at least part of the air inlet duct. The at least one strut has a strut passage and a plurality of static pressure measurement taps spaced apart on the trailing edge and in fluid communication with the strut passage.

Air inlet strut for aircraft engine

A gas turbine engine, has: an annular gaspath extending around a central axis, the annular gaspath defined between a first casing and a second casing, the first casing defining pockets; and a variable guide vane assembly having: variable guide vanes circumferentially distributed around the central axis, the variable guide vanes having airfoils extending into the annular gaspath and extending between first and second pivot members at respective first and second ends of the airfoils, the variable guide vanes rotatable about respective spanwise axes, bushings received within the pockets of the first casing, the first pivot members of the variable guide vanes rollingly engaged to the bushings, and resilient members disposed radially between surfaces of the first casing and the bushings relative to the spanwise axes, the resilient members in abutment against both of the surfaces of the first casing and the bushings.

Variable guide vane assembly and bushing therefor

A fan blade anti-icing system comprises a fan hub and a fan blade extending radially outwardly from the fan hub. The fan blade has a base and an airfoil extending radially outwardly from the fan base. The airfoil having a leading edge, a trailing edge, a convex side surface between the leading and trailing edge and a concave side surface between the leading and trailing edge. The fan blade further has a radial passage extending from a blade air inlet in the blade base in communication with a source of heated air, and a rearwardly directed passage in communication with the radial passage and having a blade air outlet forward of the trailing edge and oriented tangentially to the convex side surface or concave side surface of the airfoil.

Fan blade anti-icing concept

A gas turbine engine having an engine core defining a core flow path delimited by an engine casing, a washing device having a spraying member defining a conduit in fluid communication with a source of washing fluid, the spraying member having a spraying section for outputting the washing fluid from the conduit, the spraying member movable relative to the engine casing from a retracted position in which the spraying section is flush with a surface of the engine casing defining said core flow path, and an expanded position in which the spraying section protrudes beyond the surface of the engine casing into the core flow path

Retractable washing device

A method of heating a fan blade of a gas turbine engine for anti-icing includes emitting jets of heated air from a radial fin disposed upstream from a radially inward portion of a fan blade airfoil, the jets of heated air being directed by outlet orifices in a downstream direction substantially parallel to a flow of incoming air over the fan blade airfoil.

Method for fan blade heating using coanda effect

A multi-film oil damper for accommodating radial movement of a rotary shaft bearing relative to a bearing housing, the multi-film oil damper comprising: an annular damper cavity defined within the bearing housing between a radially outward wall, a first radially extending side wall and a second radially extending side wall, the annular damper cavity having an oil inlet in the radially outward wall, the oil inlet being in communication with a source of pressurized oil; an inner damper ring having axial ends abutting the first and second radially extending side walls of the annular damper cavity; a plurality of outer damper rings coaxially nested between the inner damper ring and the radially outward wall, each outer damper ring having axial ends adjacent the first and second radially outward walls of the annular damper cavity, each outer damper ring having a first cylindrical surface and a second cylindrical surface; a spacer ring disposed at each axial end of the plurality of outer damper rings, a contact surface of the spacer ring extending radially beyond the first surface of an associated outer damper ring and engaging the second surface of an adjacent outer damper ring defining an oil filled gap there between; and a radial oil channel through a central portion of the plurality of outer damper rings in fluid communication with the oil inlet.

Method of separating and sealing multi film damper rings

A multi-film oil damper has a housing defining an annular damper cavity having an oil inlet in communication with a source of pressurized oil. A plurality of nested damper rings is disposed within the annular damper cavity, the plurality of nested damper rings defining a plurality of squeeze film annuli. Spacer rings are disposed adjacent opposed ends of the damper rings. A contact surface of the spacer rings extends radially beyond a first cylindrical surface of an associated damper ring for engaging a second cylindrical surface of an adjacent damper ring. Recesses are defined in the second cylindrical surface of the damper rings, the recesses fluidly communicating between the squeeze film annuli and the oil inlet.

Method to supply oil to a multi-film damper

A multi-film oil damper for accommodating radial movement of a rotary shaft bearing relative to a bearing housing, the multi-film oil damper comprising: an annular damper cavity defined within the bearing housing between a radially outward wall, a first radially extending side wall and a second radially extending side wall, the annular damper cavity having an oil inlet in the radially outward wall, the oil inlet being in communication with a source of pressurized oil; an inner damper ring having axial ends abutting the first and second radially extending side walls of the annular damper cavity; a plurality of outer damper rings coaxially nested between the inner damper ring and the radially outward wall, each outer damper ring having axial ends adjacent the first and second radially outward walls of the annular damper cavity, each outer damper ring having a first cylindrical surface and a second cylindrical surface; a plurality of spacer bosses circumferentially spaced apart at each axial end of the plurality of outer damper rings, a contact surface of each spacer boss extending radially beyond the first surface of an associated outer damper ring and engaging the second surface of an adjacent outer damper ring defining an oil filled gap there between; and a radial oil channel through a central portion of the plurality of outer damper rings in fluid communication with the oil inlet.

Multi-film damper rings with spacers

A compressor rotor for a gas turbine engine is described which includes sets of blades having different airfoil thickness distributions, each including a frequency modifier forming a thickness differential relative to a baseline blade thickness. The frequency modifiers provide different natural vibration frequencies for each of the blades, and facilitate modifying natural vibration frequency separation between adjacent blades of the compressor rotor.

Mistuned fan

A gas turbine engine rotor including a disc adapted to be mounted for rotation about an axis, the disc including an annular balancing flange integrally connected thereto, the balancing flange having a first radial dimension around a first arc angle and a second radial dimension greater than the first radial dimension around a second arc angle, the second arc angle corresponding to 360 minus the first arc angle, the second arc angle being less than 180 degrees, the balancing flange around the second arc angle defining a protuberance, the protuberance being defined through machining of the balancing flange, the protuberance compensating for an unbalance of the rotor.

Growing community of inventors

Toronto, Canada

Richard Ivakitch