Pratt & Whitney Canada Corp.

Other

Hanwha Aerospace Co., Ltd.

Pratt & Withney Canada Corp.

Guy Lefebvre

Remy Synnott

Eric Durocher

John Pietrobon

Christopher Gover

Francois Doyon

Philippe Savard

Rene Paquet

Jocelyn Bisson

Nicolas Grivas

Franco Di Paola

Richard Bouchard

Denis Leclair

François Doyon

Vincent Paradis

Ian Alexander MacFarlane

Herve Turcotte

Gavin Rohiteshwar Kisun

René Paquet

Claude Giardetti

Jacob Biernat

Gabriel Emard-Parent

Patrick L'Esperance

Eugene Gekht

Daniel Lecuyer

Gabriel Naccache

Lorenzo Sanzari

Samuel Panero

Michael Paolucci

Yves Martin

Gerard Theriault

Young Jung Joo

Eric Langevin

David Pyra

Michael Fryer

Jagdeep Singh Jhinger

Ruoxin Huang

John Walter Pietrobon

A gas turbine engine includes a turbine section, a turbine case, an inner case, and a damping seal baffle. The turbine section extends along a rotational axis of the gas turbine engine. The turbine case includes a cantilevered inner wall. The cantilevered inner wall extends circumferentially about the rotational axis. The cantilevered inner wall extends between and to an upstream axial end and a downstream axial end. The upstream axial end is disposed at and downstream of the turbine section. The cantilevered inner wall forms a first sealing surface and a second sealing surface. The inner case extends and the damping seal baffle extend circumferentially about the rotational axis. The damping seal baffle is mounted to the inner case. The damping seal baffle is disposed in contact with one or both of the first sealing surface or the second sealing surface.

Turbine case structure for an aircraft propulsion system engine

An aircraft engine, comprising a thermal engine, an axial turbine having a turbine inlet, and an exhaust assembly fluidly connecting the thermal engine to the axial turbine. The exhaust assembly includes a housing and a deflector removably mounted within the housing. The deflector has circumferentially distributed vanes. The deflector is a first deflector, having a first set of geometric characteristics, that is removable from the housing and replaceable by a second deflector having a second set of geometric characteristics different from the first set of geometric characteristics. The first deflector is one of a first class of deflectors and the second deflector is one of a second class of deflectors, the first and second class of deflectors respectively defining first and second exhaust flow profiles that differ from each other.

Aircraft engine with exhaust having removable deflector

A bearing assembly for a turbomachine includes a bearing inner race, a bearing outer race located radially outwardly from the bearing inner race, and one or more bearing elements located radially between the bearing inner race and the bearing outer race. A bearing sleeve is positioned radially inboard of the bearing inner race and is configured for engagement with a rotating component. The bearing sleeve includes a sleeve body located radially between the bearing inner race and the rotating component, and a puller feature extending from the sleeve body at a first axial end of the bearing sleeve. The puller feature is configured for application of a removal force thereto to remove the bearing from an installed position in the turbomachine.

Bearing inner race with sleeve and integrated puller

An aircraft engine, has: a turbine including a turbine rotor rotatable about a central axis; a scroll case having an inlet connected to a source of combustion gases and an outlet fluidly connected to the turbine, and a conduit extending around the central axis from the inlet to the outlet; a bearing housing; a turbine support case secured to the bearing housing, the scroll case disposed axially between the turbine support case and the bearing housing, the turbine support case having spokes extending through the scroll case and radially supported by the bearing housing; and heat shields distributed around the central axis in circumferential alignment with the spokes and extending axially through the scroll case for thermally shielding the spokes, a heat shield of the heat shields disposed around a spoke of the spokes, the heat shield defining an air gap surrounding the spoke inside the scroll case.

Turbine support case with axial spokes and heat shields

A containment ring for a turbine casing assembly, including: a main body portion, the main body portion having an outer periphery and an inner periphery and a forward end and an aft end each extending between the outer periphery and the inner periphery, the main body portion also including an outer containment zone and an inner containment zone, the inner containment zone is integrally formed with the outer containment zone in order to provide a single unitary structure and the inner containment zone being radially inward from the outer containment zone, the inner containment zone extending radially outward from the inner periphery and having a plurality of ribs extending between the forward end and the aft end.

Containment ring for gas turbine engine

An aircraft engine, has: a turbine; a scroll case having an inlet connected to a source of combustion gases and an outlet connected to the turbine, and a conduit extending from the inlet to the outlet; a bearing housing including a support flange; an exhaust case downstream of the turbine; and a turbine support case secured to the bearing housing and to the exhaust case, the turbine support case having spokes extending along a direction having an axial component, the spokes extending through the scroll case and radially supported by the bearing housing, a spoke having a distal end secured to the support flange via: one or more fasteners, and a retaining member at the distal end, the retaining member defining an abutment face facing an axial direction and circumferentially overlapping the support flange, a portion of the support flange located axially between the distal end and the abutment face.

Turbine support case with axial spokes and retaining members

A containment ring assembly for a turbine casing assembly, including: a first inner containment ring; a second inner containment ring; an outer containment ring, the first inner containment ring and the second inner containment ring being radially inward from the outer containment ring, wherein a radial space is located between the outer containment ring and the first inner containment ring and the second inner containment ring, the radial space allowing deformation of the first inner containment ring and the second inner containment ring.

A centrifugal compressor is provided that includes an impeller, a diffuser, a compressor scroll, a bearing housing structure, and a coupling ring. The impeller is configured to rotate about a rotational axis. The diffuser is disposed radially outside of the impeller. The coupling ring is connected to the compressor scroll by a plurality of first fasteners and connected to the bearing housing structure by a plurality of second fasteners. The compressor is configured such that the compressor scroll will decouple from the bearing housing structure as a result of an impeller failure.

Compressor frangible containment structure

A containment ring assembly for a turbine casing assembly, including: a first containment ring; a second containment ring, the first containment ring being radially inward from the second containment ring, wherein a radial space is located between the first containment ring and the second containment ring, the radial space allowing deformation of the first containment ring prior to the first containment ring contacting the second containment ring.

A rotor assembly comprises an internal snap ring seated in a groove defined in an inner diameter surface of a rotor component. The snap ring has centrifugal weights to centrifugally expand enough to compensate for a difference in thermal growth between the rotor component and the snap ring.

Internal retaining ring arrangement for a rotor assembly

An aircraft fuel system comprises a fuel manifold adapter interchangeably connectable between respective fuel manifolds and flow divider valves of different aircraft engine platforms. The adapter has a body having an input interface and an output interface. The input interface is rigidly connectable via a rigid supply line to the flow divider valve of a given one of the different engine platforms so as to locate the output interface in axial alignment with a nozzle-input interface of the fuel manifold of the given one of the engine platforms. The output interface is axially slidably engaged with an upstream-tube end of a linear transfer tube having a downstream-tube end slidably engaged with the nozzle-input interface. The output interface of the adapter and the nozzle-input interface are slidably engaged and thermally decoupled via the linear transfer tube.

Fuel manifold adapter

An engine carcass stiffener for a turboprop engine is provided. The engine carcass stiffener includes a first longitudinal reinforcement member, a proximal end of the first longitudinal reinforcement member is coupled to a first end of a proximal flange; a second longitudinal reinforcement member, a proximal end of the second longitudinal reinforcement member is coupled to a second end of the proximal flange; a first bushing holder, the first bushing holder is coupled to a first end of a distal flange; a second bushing holder, the second bushing holder is coupled to a second end of the distal flange; a first bolt, the first bolt is inserted through the first bushing holder and couples to a distal end of the first longitudinal reinforcement member; and a second bolt, the second bolt is inserted through the second bushing holder and couples to a distal end of the second longitudinal reinforcement member.

Engine carcass stiffener for high maneuver loads

An aircraft engine, has: a turbine including a turbine rotor rotatable about a central axis; a scroll case having an inlet fluidly connected to a source of combustion gases and an outlet fluidly connected to the turbine, and a conduit extending around the central axis from the inlet to the outlet, the conduit spiraling towards the central axis; a bearing housing extending around the central axis; an exhaust case disposed downstream of the turbine; and a turbine support case secured to the bearing housing and to the exhaust case, the turbine support case having spokes distributed around the central axis and extending along a direction having an axial component relative to the central axis, the spokes extending through the conduit of the scroll case and radially supported by the bearing housing.

Turbine support case having axial spokes

The damper segment can be assembled between adjacent segments of a pressurized gas pipe of an aircraft engine. The damper segment can have a proximal end, a distal end, a rigid tube at the proximal end, a damper tube extending between the rigid tube and the distal end, the damper tube being made of a metal mesh, a proximal catch structurally connecting a proximal end of the damper tube to the rigid tube, and a distal catch structurally connected between a distal end of the damper tube and the distal end, the damper tube having an unsupported length extending between the distal catch and the proximal catch, the rigid tube having a liner portion projecting into the distal segment, the liner portion extending internally relative the damper tube.

Damper segment for pressurized gas pipe of aircraft engine

An integrated blade rotor (IBR) is provided and includes a disc, blades integrally formed with the disc and radially outwardly facing surfaces of the disc. Each radially outwardly facing surface is disposed adjacent to a corresponding blade and includes a curved profile.

Growing community of inventors

St-Bruno-de-Montarville, Canada

Guy Lefebvre