Pratt-whitney Canada Corp

Karan Anand

Alexandre Capron

Hien Duong

Tibor Urac

Krishna Prasad Balike

Tammy Yam

Paul Stone

Peter Townsend

Thomas Veitch

Daniel Fudge

Sylvain Claude

Ali Azmi

Ignatius Theratil

Vijay Kandasamy

Farid Abrari

Hong Yu

Kari J Heikurinen

Paul Aitchison

Ernest Adique

Yau-Wai Li

Prakul Mittal

Anthony Robert Bruni

Rakesh Munlyappa

César Carlos Valbuena Merino

Ignacio Javier Ucin Jarne

Pascal Doran

Charles Mason

Edwin Wong

Juan Ignacio Ruiz-Gopegui Gonzalez

Terry Brownridge

Cesar Valbuena Merino

Kari Heikurinen

A method includes: obtaining a rotor having a hub and a plurality of blades protruding from the hub, the plurality of blades including first blades and second blades disposed in alternation around a central axis of the rotor, natural vibration frequencies of the first blades different from natural vibration frequencies of the second blades; determining that a difference between a first natural vibration frequency of a first blade of the first blades and a second natural vibration frequency of a second blade of the second blades is below a threshold; and modifying a shape of the first blade until the difference between the first natural vibration frequency and the second natural vibration frequency is at or above the threshold.

Method of manufacturing a mistuned rotor

Device and methods for guiding bleed air in a turbofan gas turbine engine are disclosed. The devices provided include louvers and baffles that guide bleed air toward a bypass duct of the turbofan engine. The louvers and baffles have a geometric configuration that promotes desirable flow conditions and reduced energy loss.

Devices and methods for guiding bleed air in a turbofan engine

An aircraft engine, has: an upstream stator having upstream stator vanes distributed about a central axis; and a downstream stator having downstream stator vanes distributed about the central axis, the downstream stator located downstream of the upstream stator, a number of the upstream stator vanes different than a number of the downstream stator vanes, the downstream stator vanes including: a first vane, a major portion of a leading edge of the first vane circumferentially overlapped by one of the upstream stator vanes; and a second vane differing from the first vane by a geometric parameter, the geometric parameter causing the second vane to have one or more of: a stiffness greater than that of the first vane, and a major portion of a leading edge of the second vane circumferentially overlapped by another one of the upstream stator vanes.

Aircraft engine with stator having varying geometry

An aircraft engine, has: an upstream stator having upstream stator vanes circumferentially distributed about a central axis; and a downstream stator having downstream stator vanes circumferentially distributed about the central axis, the downstream stator located downstream of the upstream stator relative to an airflow flowing within a core gaspath of the aircraft engine, a number of the upstream stator vanes being different than a number of the downstream stator vanes, major portions of leading edges of the downstream stator vanes circumferentially overlapped by the upstream stator vanes, the downstream stator vanes including: a first pair of circumferentially adjacent vanes of the downstream stator vanes spaced apart by a first pitch, and a second pair of circumferentially adjacent vanes of the downstream stator vanes spaced apart by a second pitch different than the first pitch.

Aircraft engine with stator having varying pitch

Assemblies and methods for directing a flow of air through a compressor section of a gas turbine engine are disclosed. A method includes receiving the flow of air over a strut extending radially across a substantially annular gas path of the gas turbine engine, and at least partially confining and modifying a strut wake generated in the flow of air by the strut using one or more variable orientation guide vanes.

Assembly for a compressor section of a gas turbine engine

An aircraft engine, has: an upstream stator having upstream stator vanes circumferentially distributed about a central axis; and a downstream stator having downstream stator vanes circumferentially distributed about the central axis, the downstream stator located downstream of the upstream stator relative to an airflow flowing within a core gaspath of the aircraft engine, a number of the upstream stator vanes being different than a number of the downstream stator vanes, the downstream stator vanes including: a first vane made of a first material, a major portion of a leading edge of the first vane circumferentially overlapped by one of the upstream stator vanes, and a second vane made of a second material having a greater stiffness, strength, and/or ductility than that of the first material, a major portion of a leading edge of the second vane exposed via a spacing defined between two of the upstream stator vanes.

Aircraft engine having stator vanes made of different materials

A stator that has vanes extending between a first end and a second end along a span and from a leading edge to a trailing edge along a chord length, the vanes having a first end portion extending from the first end to about 30% of the span to a first location, a chord ratio of the chord length at the first end to the chord length at the first location greater than or equal to 1.1, a throat ratio of a width of a throat between two adjacent vanes at the first location to a width of the throat at the first end is greater than or equal to 1.3, a sweep angle difference between a maximum sweep angle of the leading edge along the first end portion and a minimum sweep angle of the leading edge along the first end portion is at least 15 degrees.

Compressor stator

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

There is disclosed a fan assembly including a fan rotor including a hub and fan blades. The fan blades have a leading edge and a trailing edge. A fan stator downstream of the fan rotor relative to a direction of an airflow through the fan assembly. The fan stator includes vanes extending between radially inner ends and radially outer ends. A flow recirculation circuit has an inlet downstream of radially inner ends of the vanes of the fan stator and an outlet upstream of radially inner ends of the vanes. A recirculation stator has a plurality of stationary guide vanes circumferentially distributed around the axis and located in the flow recirculation circuit between the inlet and the outlet A method of operating the fan assembly is also disclosed.

Fan assembly having flow recirculation circuit with guide vanes

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 bleed of valve comprises a ring axially translatable between a retracted position in which the ring is configured to close an annular bleed off opening defined in a converging portion of a radially outer annular wall of a gas turbine engine gaspath and a deployed position in which the ring protrudes into the gaspath to mechanically scoop out incoming air and water/hail particles.

Translating gaspath bleed valve

Compressor stators () for gas turbine engines () are disclosed. An exemplary compressor stator () comprises a circumferential array of stator vanes () and a shroud () for supporting the stator vanes (). The shroud () has a circumferentially extending inner endwall () exposed to an annular gas path () of the compressor (). The endwall () has a circumferentially uniform axial cross-sectional profile (). The axial cross-sectional profile () comprises at least one deviation () from a nominal axial cross-sectional profile () defining an overall shape of the annular gas path (). The at least one deviation comprises a concave deviation () and a convex deviation ().

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Mississauga, Canada

Karan Anand