Pratt-whitney Canada Corp

Thomas Veitch

Tibor Urac

Hien Duong

Paul Stone

Farid Abrari

Kari J Heikurinen

Daniel Fudge

Ernest Adique

Peter Townsend

Krishna Prasad Balike

Raman Warikoo

Keegan Lobo

Ignatius Theratil

Paul Aitchison

David Beamish

Karan Anand

Tyler Richardson

Jason Nichols

Nathan Tomes

Philippe Bonniere

Bruce Fielding

Guiherme Watson

Terry Brownridge

Phillipe Bonniere

Guilherme Watson

Kari Heikurinen

An aircraft power plant, comprising: a fan for providing power to an aircraft; a gas turbine engine located on a first axial side of the fan, the gas turbine engine having a compressor drivingly engaged by a turbine via an engine shaft; an electric motor powered by a power source, the electric motor driving a motor shaft; and a gearbox located on a second axial side of the fan opposite the first axial side, the gearbox in driving engagement with the engine shaft, the motor shaft, and the fan, the gearbox defining a first load path between the gas turbine engine and the fan and a second load path between the electric motor to the fan, the fan in continuous driving engagement with both of the gas turbine engine and the electric motor along the first load path and the second load path.

Hybrid aircraft power plant

A multi-film oil damper has an annular damper cavity defined within a housing between a radially outward wall, a first radially extending side wall and a second radially extending side wall. Nested damper rings are disposed within the annular damper cavity for defining squeeze film annuli therebetween. The squeeze film annuli are fluidly connected in parallel to an inlet gallery.

Method and system to supply oil to a multi-film oil damper

A multi-film oil damper has a housing defining an annular damper cavity between a radially outward wall and radially extending side walls. The annular damper cavity has an oil inlet configured for connection to a source of pressurized oil. A closure ring defines a radially inward boundary of the annular damper cavity. First and second damper rings are nested together coaxially within the annular damper cavity. At least one of the first damper ring and the second damper ring has an axial end radial thickness less than an intermediate-portion radial thickness.

Multi-film oil damper with tapered damper rings

A gas turbine engine casing apparatus includes annular first and second annular cases connected by a spigot joint. The spigot joint includes an annular projection of the first annular case fitted into an annular recess of the second annular case. A plurality of circumferentially spaced apart scallops are formed on one of surfaces of the annular projection or of the annular recess, and are located in selective circumferential locations adjacent respective enhanced stiff areas of the first and second annular cases.

Intermittent spigot joint for gas turbine engine casing connection

A compressor rotor for a gas turbine engine includes a hub disposed about an axis of rotation and an outer surface forming a radially inner gaspath boundary, the outer surface defining a nominal hub diameter. A circumferential array of blades extends radially outwardly from the hub. A first inter-blade passage is defined between a first set of adjacent blades and has a first throat area. A second inter-blade passage is defined between a second set of adjacent blades and has a second throat area that is smaller than the first throat area. At least one scoop is disposed in the second inter-blade passage, the scoop defining a cavity extending radially into the outer surface of the hub relative to the nominal hub diameter.

Mistuned compressor rotor with hub scoops

A rotor for a gas turbine engine. The rotor includes blades circumferentially distributed around a hub. The blades have airfoils with a span defined between a root and tip, a chord defined between a leading edge and a trailing edge, and a thickness defined between a pressure side surface and suction side surface. The blades include first blades and second blades. The airfoil of the first blades has a first thickness distribution defining a first natural vibration frequency of the airfoils of the first blades. The airfoil of the second blades has a second thickness distribution defining a second natural vibration frequency different than the first natural vibration frequency. The first thickness distribution is different than the second thickness distribution along a radially-inner half of the span, and the first thickness distribution matches the second thickness distribution along a radially-outer half of the span.

Mistuned rotor for gas turbine engine

A fan for a gas turbine engine comprises blades distributed around a hub. The blades include first and second blades, having geometric parameters and/or material properties that differ from each other to frequency mistune the fan. The blades are distributed about the hub with at least one second blade between adjacent first blades. The leading edge of the airfoil of the second blades is disposed axially aft of the corresponding leading edge of the first blades in at least a portion of the blade span.

Mistuned fan for gas turbine engine

A multi-film oil damper suited 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 and radially extending side walls; an inner damper ring; an outer damper ring disposed between the inner damper ring and the radially outward wall of the annular damper cavity; wherein the annular damper cavity has an oil inlet in communication with a source of pressurized oil; and wherein a radial thickness of the inner damper ring exceeds a radial thickness of the outer damper ring.

Method to seal damper cavity of multi-film oil damper

An inter-compressor case for a gas turbine engine is disclosed. The inter-compressor case comprises: an outer casing and an inner casing radially spaced apart relative to a longitudinal axis; a gas path extending from a plurality of radial inlets arranged in a circumferentially spaced apart array around the outer casing to an annular outlet defined by an axially extending downstream portion of the outer casing and an axially extending downstream portion of the inner casing; a plurality of struts having a gas path surface extending across the gas path between the outer casing and the inner casing; and a plurality of flow separators extending from the adjacent radial inlets. The flow separators have trailing edges disposed upstream of the annular outlet and include a plurality of full length flow separators and a plurality of truncated flow separators.

Inter-compressor flow divider profiling

A compressor rotor airfoil in a gas turbine engine is presented. Opposed pressure and suction sides are joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. A leading edge dihedral angle is defined at a point on the leading edge between a tangent to the airfoil and a vertical. The leading edge dihedral angle has a span-wise distribution. The distribution has at least one inflection point. A method of reducing a rub angle between a compressor rotor blade and a casing surrounding the blade is also presented.

Compressor rotor airfoil

An airfoil for a gas turbine engine comprises at least three successive sections defined between a root and a tip. A first decrease of maximum thickness-to-chord ratios defined by a difference between a maximum thickness-to-chord ratio at the root and a maximum thickness-to-chord ratio at a first spanwise position, a second decrease of the maximum thickness-to-chord ratios defined by a difference between the maximum thickness-to-chord ratio at the first spanwise position and a maximum thickness-to-chord ratio at a second spanwise position, a third decrease of the maximum thickness-to-chord ratios defined by a difference between the maximum thickness-to-chord ratio at the second spanwise position and a maximum thickness-to-chord ratio at the tip, the second decrease being greater than the corresponding first and third decreases.

Airfoil with stepped spanwise thickness distribution

A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades.

Compressor rotor with coated blades

A gas turbine engine is described having a rotor with a plurality of rotor blades each extending radially between a blade root to a blade tip and axially between a leading edge and a trailing edge. An annular casing body housing the rotor blades includes an abradable segment of the inner surface facing the blade tips and having an abradable member. The abradable segment having one or more annular grooves, extending into the casing body from an inner surface thereof, and including an edge groove axially aligned with and facing the leading edge or the trailing edge of the rotor blades. The edge groove extends axially between a first position on the inner surface upstream of the leading edge or the trailing edge, and a second position on the inner surface downstream of the leading edge or the trailing edge.

Gas turbine engine blade casing

A rotary airfoil in a gas turbine engine is provided. The airfoil includes opposed pressure and suction sides joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. An axial component of a center of gravity of a cross-section taken chordally toward the tip of the airfoil being upstream relative to an axial component of a center of gravity of a cross-section taken chordally toward the root of the airfoil. A method for forming such blade is also presented.

Growing community of inventors

Scarborough, Canada

Thomas Veitch