General Electric Company

Safran Aircraft Engines

General Electric Company Polska Sp. Z O.o.

Generale Electric Company Polska Sp. Z O.o

Daniel Edward Mollmann

Nicholas Joseph Kray

David William Crall

Thomas Richard Henning

Nitesh Jain

Max Robert Farson

John Douglas Mickol

Andrew Breeze-Stringfellow

Jan Christopher Schilling

Stefan Joseph Cafaro

Steven Alan Ross

Donald Lee Gardner

James Ryan Reepmeyer

Barry Lynn Allmon

Arjan Johannes Hegeman

Marcia Boyle Johnson

Inenhe Mohammed Khalid

Amit Zutshi

Gert Ven Der Merwe

Arthur William Sibbach

Christopher Charles Glynn

Randy M Vondrell

Trevor Howard Wood

Aspi Rustom Wadia

Michael Harvey Schneider

Howard Paul Weaver

Francois Gallet

Richard Alan Wesling

Gert Johannes Van Der Merwe

Craig William Higgins

Frank Worthoff

Tod Robert Steen

Elzbieta Kryj-Kos

Kenneth Lee Fisher

Charles Stanley Orkiszewski

Mark Edward Stegemiller

Kishore Budumuru

Dong-Jin Shim

Kenneth O Johnson

Jonathan Edward Coleman

Guillaume Glemarec

Bala Corattiyil

Mark Edward Linz

William Carl Ruehr

Todd M Dahling

Thomas Edward Agin

Morris Green Penn

Christopher Ryan Johnson

Rafal Sarba

Valerio Capasso

Jean-Baptiste Manuel Nicolas Vignes

Anthony John Durchholz

Jeffrey Scott Allen

Stephen L Carmichael

Nicolas Maurice Hervé Aussedat

Michael Joseph Gambone

Timothy James Horejs

Michael J Stallone

Michael P Swift

Pamela M Boehm

Brent Michael Push

Stephen Wade Faw

Enzo Luigi Morabito

Matthew A Oakland

Lawrence J Bach

Eugene F Gaffney

Michael Cline

Lyle Douglas Dailey

Anthony Durchholz

Randy Marinus Vondrell

An open rotor engine includes a core engine, a plurality of guide vanes positioned within or extending from the core engine; and a pitch change assembly operably coupled to the plurality of guide vanes. The pitch change assembly includes one or more actuators configured to change a pitch angle of respective ones of the plurality of guide vanes, and a plurality of linkage arms that are respectively movable by actuation of at least one of the one or more actuators. The plurality of linkage arms are directly or indirectly coupled to a corresponding one of the plurality of guide vanes. The plurality of linkage arms may have a length that differs from one another, and such length may orient a displacement or a range of motion of the respective linkage arm to an envelope of rotation about a guide vane axis that differs as between the plurality of guide vanes.

Controlling excitation loads associated with open rotor aeronautical engines

Methods, apparatus, systems and articles of manufacture for mounting a gas turbine engine are described. An apparatus for mounting a gas turbine engine to a pylon, the gas turbine including an upstream section and a downstream section, the gas turbine defining a roll axis, a yaw axis, and a pitch axis, the apparatus including: a first mount to couple the upstream section of the gas turbine engine to the pylon; a second mount to couple the upstream section of the gas turbine engine to the pylon, the second mount downstream of the first mount; a thrust linkage to couple the upstream section to the pylon, wherein the downstream section is decouplable from the upstream section without decoupling the first mount, the second mount, and the thrust linkage.

Methods and apparatus for mounting a gas turbine engine

An airfoil assembly for a turbine engine, the airfoil assembly including a platform defining an inner surface and an outer surface, a variable pitch airfoil extending radially from the outer surface of the platform from a root to a tip to define a span length and a mounting structure connected to the platform.

Variable pitch airfoil

A method of controlling an aeronautical gas turbine engine may be performed with an electronic controller. The method may include determining an airfoil pitch control command for at least one of a plurality of airfoils of the aeronautical gas turbine engine based at least in part on an excitation load acting upon the aeronautical gas turbine engine, and outputting the airfoil pitch control command to one or more actuators actuatable to change a pitch angle of the at least one of the plurality of airfoils. The airfoil pitch control command may be configured to augment and/or compensate for the excitation load acting upon the aeronautical gas turbine engine. The method may be embodied by a non-transitory computer-readable medium that includes computer-executable instructions, which when executed by a processor associated with the electronic controller, cause the electronic controller to perform the method.

Composite airfoils and methods for forming composite airfoils are provided. For example, a composite airfoil of a gas turbine engine comprises opposite pressure and suction sides extending radially along a span from a root to a tip, which define opposite radial extremities of the airfoil. The composite airfoil further comprises a body section and a tip section, which includes the tip, that each extend radially along the span. The composite airfoil is formed from a composite material comprising fibers disposed in a matrix material. The tip section has a tip fiber volume, and the body section has a body fiber volume that is greater than the tip fiber volume. Another composite airfoil comprises a tip cap applied over the tip that tapers from a first end to a second end such that each of the pressure and suction side walls of the tip cap narrows from a first thickness to a second thickness.

Composite airfoils with frangible tips

An airfoil defining a span extending between a root and a tip. The airfoil includes a frangible airfoil portion extending between a leading edge and a trailing edge and extending between the tip and a fusion line along the span. The frangible airfoil portion includes a first material defining a first modulus of elasticity and further defines an internal cavity. The airfoil includes a residual airfoil portion coupled to the frangible airfoil portion extending from the fusion line to the root along the span and including a second material defining a second modulus of elasticity greater than the first modulus of elasticity. The residual airfoil portion meets the frangible airfoil portion at the fusion line.

Frangible gas turbine engine airfoil including an internal cavity

A method for operating a gas turbine engine includes receiving data indicative of an operational vibration within a section of the gas turbine engine; and providing electrical power to a shaker mechanically coupled to one or more components of the section of the gas turbine engine to generate a canceling vibration to reduce or minimize the operational vibration within the section of the gas turbine engine.

Vibration control for a gas turbine engine

An airfoil defining a span extending between a root and a tip and a chord at each point along the span extending between a leading edge and a trailing edge. The airfoil includes a frangible airfoil portion at the tip extending between the leading edge and the trailing edge and extending between the tip and a frangible line along the span. The frangible airfoil portion includes an exterior surface. The exterior surface at least partially defines at least one fusion cavity at least partially defining the frangible line. The airfoil further includes a residual airfoil portion extending from the frangible line to the root along the span. The residual airfoil portion meets the frangible airfoil portion at the frangible line. As such, the frangible airfoil portion deforms or partially or fully detaches relative to the residual airfoil portion at the frangible line following an event creating imbalance.

Frangible gas turbine engine airfoil with fusion cavities

A blade for a propulsion apparatus. The blade includes a body having opposed pressure and suction sides. The blade extends in span between a root and a tip and the blade extends in chord between a leading edge and a trailing edge. The blade includes a fracture structure that is defined within the body and positioned such that it is spaced-away from the tip. The fracture structure includes at least one chamber. A first fracture wall is disposed between the at least one chamber and the pressure side, and a second fracture wall is disposed between the at least one chamber and the suction side. The fracture structure is configured to fail when a predetermined force is applied to the tip such that the tip separates from the remainder of the body.

Metallic compliant tip fan blade

A containment structure comprising an annular structure configured to provide close tolerance to a rotation structure. The containment structure includes a first abradable structure, a second abradable structure, and a shroud that is defined in part by the first abradable structure and in part by the second abradable structure. An annular surface of the shroud is defined by the first abradable structure and by the second abradable structure. The first abradable structure is harder than the second abradable structure.

Unsymmetrical turbofan abradable grind for reduced rub loads

A gas turbine engine airfoil includes leading and trailing edges, pressure and suction sides extending from airfoil base to airfoil tip, trailing edge cladding made of cladding material bonded to composite core made of composite material, cladding material less brittle than composite material, composite core including central core portion extending downstream from leading edge portion to trailing edge portion of composite core, and trailing edge cladding including wavy wall and trailing edge. Pressure and suction side flanks of trailing edge cladding may be bonded to pressure and suction side surfaces of trailing edge portion. Waves of wavy wall may extend normal to and away from the pressure and suction side surfaces. Trailing edge cladding may include wavy pressure and suction side trailing edge guards including waves of wavy wall. Airfoil may extend outwardly from platform of a blade. Root may include integral dovetail.

Fan blade with composite core and wavy wall trailing edge cladding

A method for reducing tip rub loads in a rotor blade is described having the steps of selecting a location of contact between the rotor blade and a static structure; selecting a location on the blade for incorporating a tip-cutter; selecting a cutter-profile for the tip-cutter; and incorporating the tip-cutter on the selected location on the blade such that tip-cutter is capable of removing a portion of the static structure during a tip rub such that tip rub loads are facilitated to be reduced.

Method for reducing tip rub loading

An airfoil for use in a rotor assembly is disclosed, the airfoil having a tip-cutter located on a sidewall near a tip portion, wherein the tip-cutter is capable of removing a portion of an abradable material during a tip rub. In another embodiment, an airfoil has a tip-grinder located on the tip portion, wherein the tip-grinder capable of removing a portion of an abradable material during a tip rub. In another embodiment, an airfoil has a tip-rake that facilitates reducing loading induced to said airfoil during tip rubs.

Rotor blade with reduced rub loading

Vibration is monitored in a dual rotor engine. A first total vibration of the engine is measured during operation. Relative phase of the two rotors is then clocked, and a second total vibration of the engine is then measured during operation. First and second total unbalances are then resolved at a common response frequency from the measured total vibrations. First and second rotor unbalances are then resolved for the corresponding rotors from the first and second total unbalances. In a further method, corresponding corrective balance weights may be determined for offsetting the resolved rotor unbalances.

Dual rotor vibration monitoring

A processor-implemented method of assembling a rotatable machine is provided. The machine includes a plurality of blades that extend radially outwardly from a rotor. The method includes determining a geometric parameter for each blade in a row of blades that is relative to a ratio, R of an inlet area and an outlet area of a predetermined volume defined between each pair of blades, determining an initial sequence map for the row of blades that facilitates minimizing a difference of R between circumferentially adjacent pairs of blades, and iteratively remapping the sequence of the blades to facilitate reducing a moment weight vector sum of the rotor to a value that is less than a predetermined value.

Methods and apparatus for assembling rotatable machines

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Cincinnati, OH, United States of America

Daniel Edward Mollmann