Pratt-whitney Canada Corp

Eduardo David Hawie

Nigel Caldwell Davenport

Yen-Wen Wang

Stephen Phillips

Bhawan Bhai Patel

Parthasarathy Sampath

Lorin Markarian

Jian-Ming Zhou

Lev Alexander Prociw

Jason Araan Fish

Honza Stastny

Pierre-Yves Legare

Kenneth D Parkman

Dave Menheere

John Greer

Yvan Schraenen

Carol Desjardins

Michael Papple

Robert Sze

Sri Sreekanth

Vincent Paradis

Othmane Leghzaouni

Jasrobin Grewal

Pierre-Yves Légaré

Philippe Boyer

Constantinos Ravanis

Ion Dinu

Reza Ziaei

Marc-Antoine Mailloux-Labrousse

Douglas Maccaul

Hayley Ozem

Nigel Davenport

Gas turbine engines include a compressor section configured to compress a core flow and a combustor section having a combustor that is arranged downstream from the compressor section along a path of the core flow. A turbine section is arranged downstream from the combustor along the path of the core flow, and includes a plurality of first vanes arranged at an outlet of the combustor, with at least one first vane of the plurality of first vanes having an internal vane path. A compressor diffuser is arranged to direct the core flow from the compressor section through the internal vane path. The path of the core flow, in a flow direction, passes through the compressor section, the compressor diffuser, the internal vane path, the combustor, and then between the first vanes of the plurality of first vanes to enter the turbine section.

Integrated axial compressor diffuser and high pressure turbine vane assembly

Gas turbine engines include a compressor section having an impeller configured to compress a core flow. A combustor section having a combustor is arranged downstream from the compressor section along a path of the core flow. A turbine section is arranged downstream from the combustor section along the path of the core flow, with the turbine section having a plurality of first vanes arranged at an outlet of the combustor and at least one first vane of the plurality of first vanes includes an internal vane path. The path of the core flow, in a flow direction, passes through the impeller of the compressor section, the internal vane path internal to the first vane, the combustor, and then between external surfaces of the plurality of first vanes to enter the turbine section.

Integrated centrifugal compressor diffuser and high pressure turbine vane assembly

A combustor for a gas turbine engine comprises an inner annular liner wall and an outer annular liner wall cooperating to form a combustion chamber of the combustor. A first dome wall has a circumferential array of first fuel injection bores. A second dome has a circumferential array of second fuel injection bores. An intermediate wall extends between the first dome wall and the second dome wall. A first combustion stage is defined by the inner liner wall forward end, the first dome wall and the intermediate wall. A second combustion stage is defined at least by the outer liner wall forward end, the second dome wall and the intermediate wall, the first combustion stage communicating with the first fuel injection bores, the second combustion stage communicating with the second fuel injection bores.

Two-stage combustor for gas turbine engine

A turbine blade for a gas turbine engine has an airfoil including leading and trailing edges joined by spaced-apart pressure and suction sides to provide an external airfoil surface extending from a platform in a spanwise direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil defined by a set of Cartesian coordinates set forth in Table 1, the Cartesian coordinates provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by a local axial chord, and a span location.

Turbine blade airfoil profile

A method for delivering fuel to a combustor of a gas turbine engine includes: injecting fuel into a conduit of a fuel manifold; directing the fuel through a continuous first portion of the conduit; directing the fluid from the first portion into an inflexion of the conduit being a single exit of the first portion; directing the fuel from the inflexion into a second portion of the conduit in serial flow communication with the first portion; and carrying the fuel in the second portion in a direction different from that of the fuel in the first portion. As the fuel flows through the second portion of the conduit, the fuel exits from the second portion of the conduit into a plurality of fuel injection nozzles in exclusive fluid flow communication with the second portion for ejection into a combustor of the gas turbine engine.

Fluid manifold for gas turbine engine and method for delivering fuel to a combustor using same

A method for delivering fuel from a fuel nozzle of a combustor of a gas turbine engine includes directing fuel from a fuel source through a flow splitter to provide at least two concentric fuel flows, filming the concentric two fuel flows on concentrically arranged inwardly facing filming surfaces that are disposed downstream of the flow splitter, and atomizing the concentric fuel flows into a core air flow.

Fuel nozzle

A gas turbine engine combustor is described which includes outer and inner annular combustor liners formed of sheet metal. The exit duct circumscribes an annular combustor exit and defines a combustion gas path. The exit duct includes a large exit duct having an annular forged metal section which is butt welded at an upstream end to the outer combustor liner to form a first annular joint. The annular forged metal section is fixed at a downstream end to an annular sheet metal wall to form a second annular joint. Methods of forming and of repairing a gas turbine engine combustor are also disclosed.

Gas turbine engine combustor and method of forming same

A fuel nozzle for a combustor of a gas turbine engine includes a body defining an axial direction and a radial direction, an air passageway defined axially in the body, and a fuel passageway defined axially in the body radially outwardly from the air passageway. The fuel passageway has an outer wall including an exit lip at a downstream portion of the outer wall. The lip generally increases in diameter as it extends downstream.

The described gas turbine engine fuel manifold includes one or more fuel conduits defined within a ring-shaped manifold body, each of the fuel conduits being fluidly connected to a respective separate group of fuel injection nozzles. Each of the fuel conduits, which extend from a conduit inlet to a conduit end, include a first portion extending continuously between the conduit inlet and an inflexion of the conduit, and a second portion downstream of the first portion and connected thereto in serial flow communication, the second portion extending between the inflexion and the conduit end. The inflexion being a single exit of the first portion such that fuel flows uninterrupted between the conduit inlet and the inflexion. The respective separate group of fuel injection nozzles of the fuel conduit fluidly communicating exclusively with the second portion of the fuel conduit.

A fuel nozzle for a combustor of a gas turbine engine includes a body defining an axial direction and a radial direction, a primary air passageway centrally defined axially in the body, and a plurality of concentrically-arranged nozzle tip projections disposed at a downstream portion of the body. Each of the plurality of nozzle tip projections has a radially inwardly facing fuel filming surface communicating with respective fuel passages. The fuel filming surfaces are disposed radially outwardly of an outlet of the primary air passageway. A method for delivering fuel from a fuel nozzle of a combustor of a gas turbine engine is also presented.

A fuel nozzle for a combustor of a gas turbine engine includes a body defining an axial direction and a radial direction, an air passageway defined axially in the body, and a fuel passageway defined axially in the body radially outwardly from the air passageway. The fuel passageway has an outer wall including an exit lip at a downstream portion of the outer wall. The exit lip has a surface treatment including a swirl-inducing relief. A gas turbine engine and a method of inducing swirl in at least one of pressurised fuel and air exiting a fuel nozzle of a gas turbine engine are also presented.

A combustor heat shield assembly comprises a circumferential array of heat shield panels individually mounted to an inner surface of a combustor shell. Each heat shield panel has opposed front and back faces, the back face facing the inner surface of the combustor shell and being spaced therefrom to define an air gap. The front and back faces have a perimeter including opposed lateral edges extending between opposed circumferentially extending edges. The lateral edges of adjacent heat shield panels have complementary non-linear profiles defining an asymmetric heat shield panel interface.

Asymmetric combustor heat shield panels

The combustor has a laser ignitor mounted to the casing, remotely from the liner of the combustion chamber. The laser ignitor has an igniter beam path for igniting the fuel and air mixture in the combustion chamber, the igniter beam path extending at least partially across the air plenum surrounding the liner and into the combustion chamber through a corresponding beam path aperture provided in the liner.

Laser-ignition combustor for gas turbine engine

A combustor for a gas turbine engine comprises an inner annular liner and an outer annular liner. A first and a second combustion stages are defined between the liners, each said combustion stage having a plurality of fuel injection bores distributed in a liner wall defining the respective stage. Valves at the fuel injection bores of one of the combustion stages, the valves each defining an air passage from an exterior to an interior of the combustion stage, the valves each having an actuatable member for adjusting a size of a respective air passage for air staging the combustor.

Growing community of inventors

Woodbridge, Canada

Eduardo David Hawie