General Electric Company

David William Crall

Nicholas Joseph Kray

Daniel Edward Mollmann

Brandon Wayne Miller

Bhaskar Nanda Mondal

Ian Francis Prentice

Frank Worthoff

Christopher Charles Glynn

Tsuguji Nakano

Douglas Duane Ward

Jeffrey Donald Clements

Christopher James Kroger

Bruce Clark Busbey

Max Robert Farson

Trevor Wayne Goerig

Daniel Alan Niergarth

Nitesh Jain

Scott Roger Finn

Erich Alois Krammer

Craig William Higgins

Peter Nicholas Szucs

Donald Lee Gardner

Richard David Cedar

Gregory Carl Gemeinhardt

Mustafa Dindar

Vinod Shashikant Chaudhari

Howard Paul Weaver

Courtney James Tudor

Peter John Linko, Iii

Marcia Boyle Johnson

Seth Alexander McDonald

Gregory Scott McNulty

Joseph Donofrio

Darek Tomasz Zatorski

Aspi Rustom Wadia

Sridhar Adibhatla

Ming Xie

Wendy Wenling Lin

James Edward Johnson

Syed Arif Khalid

Stephen Craig Mitchell

Donald George LaChapelle

David Lawrence Bedel

William Howard Hasting

Rollin G Giffin, Iii

Michael John Franks

Tod Winton Davis

Lee Alan Blanton

John Robert Kelley

Kevin Lee Kirkeng

Jack W Baldwin

Govindarajan Rengarajan

Thomas Richard Henning

Richard Clay Haubert

Richard Thomas Price

Jason Francis Pepi

Todd M Dahling

John W Salvage

Carl C Koch

Donald Ray Bond

Robert Clayton Von Der Esch

Christopher Lee McAfee

Charles Daniel Califf

Steven Roy Manwaring

Michael D Toye

David G Glista

Anthony L DiPietro

Benjamin Vincent Rawe

Matthew A Oakland

Michael De Wayne Toye

David Andrew Goshorn

Michael Dewayne Toye

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 aircraft defining a longitudinal direction and a lateral direction is provided. The aircraft includes a fuselage; an engine mounted at a location spaced from the fuselage of the aircraft, the engine comprising rotor blades; and at least one fuselage shield removably coupled to the fuselage at a location in alignment with the rotor blades along the lateral direction.

Removeable fuselage shield for an aircraft

A fan for a turbine engine having a core air flowpath and a bypass air flowpath. The fan includes a disk, a plurality of fan blades, and a plurality of splitter blades. The disk rotates about a centerline axis. The plurality of fan blades are coupled to the disk. The plurality of splitter blades are coupled to the disk and positioned between the plurality of fan blades. The plurality of splitter blades are axially aligned on the disk at a splitter blade axial distance from an annular inlet of the core air flowpath such that the plurality of splitter blades direct debris that passes through the fan into the bypass air flowpath.

Fan for a turbine engine

An aircraft defining a longitudinal direction and a lateral direction is provided. The aircraft includes: a fuselage; an engine mounted at a location spaced from the fuselage of the aircraft, the engine comprising a plurality of rotor blades; and a fuselage shield attached to or formed integrally with the fuselage at a location in alignment with the plurality of rotor blades along the lateral direction, the fuselage shield comprising a first layer defining a first density and a second layer defining a second density, the first density being different than the second density.

Layered fuselage shield

Structures, such as compressor casings, for reducing a thermal gradient are provided. For example, a compressor case is split such that it includes first and second case segments. The first case segment extends over a first portion of the compressor case circumference and comprises a first inner structure, a first outer structure, and a first porous structure integrally formed as a monolithic component. The first porous structure is defined between the first inner structure and the first outer structure. The second case segment extends over a second portion of the compressor case circumference and comprises a second inner structure, a second outer structure, and a second porous structure integrally formed as a monolithic component. The second porous structure is defined between the second inner structure and the second outer structure. Methods of cooling structures such as compressor casings also are provided.

Split casings and methods of forming and cooling casings

The disclosure is towards an inlet cowl for a turbine engine including a surface defining an inlet with a flow path and a method towards controlling the airflow in the flow path. The inlet cowl further includes an inlet lip and inner and outer barrels. The inlet lip confronts the inner barrel at a junction defining a gap.

Inlet cowl for a turbine engine

Methods, apparatus, systems and articles of manufacture for compact compressors are disclosed including a gas turbine engine defining an axial direction and a radial direction, the gas turbine engine including an axial flow compressor and a radial flow compressor, wherein the axial flow compressor is located axially forward of the radial flow compressor, a blade assembly including a splitter shroud to divide incoming air into axial air flow for the axial flow compressor and radial air flow for the radial flow compressor, the blade assembly rotating relative to the axial flow compressor and counter-rotating relative to the radial flow compressor, and wherein the blade assembly is located axially aft of the radial flow compressor.

Compact low-pressure compressor

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

An airfoil that includes an airfoil body having a root and a tip, and convex and concave sides that extend between a leading edge and a trailing edge. The airfoil also includes at least a first cladding element that is attached to the airfoil body. The first cladding element includes a first portion and a second portion. The second portion is configured to separate from the first portion when the first cladding element encounters a force of at least a predetermined amount.

Multi-material leading edge protector

A combustion engine includes a combustion section; a fuel delivery system for providing a fuel flow to the combustion section, the fuel delivery system including an oxygen reduction unit for reducing an oxygen content of the fuel flow; a thermal management system including a heat sink heat exchanger, the heat sink heat exchanger in thermal communication with the fuel delivery system at a location downstream of the oxygen reduction unit; and a control system including a sensor operable with the fuel delivery system for sensing data indicative of an operability of the oxygen reduction unit and a controller operable with the sensor, the controller configured to initiate a corrective action based on the data sensed by the sensor indicative of the operability of the oxygen reduction unit.

Thermal management system

Containment assemblies and methods for forming containment assemblies of gas turbine engines are provided. For example, a containment assembly comprises a containment case including a plurality of coated fibers. Each coated fiber comprises a fiber surrounded by a ceramic material such that the ceramic material coats the fiber. As another example, a containment assembly comprises an inner case and a containment case comprising a plurality of coated fibers. Each coated fiber comprises a fiber surrounded by a ceramic material such that the ceramic material coats the fiber. The containment case includes a greater proportion of the coated fibers at an inner surface of a layer of the containment case than at a location within the containment case that is radially outward from the inner surface. Methods for forming a containment assembly of a gas turbine engine are provided.

Containment case having ceramic coated fibers

A method and icing effects mitigation system are provided. The icing effects mitigation system includes a fluid duct configured to channel a first flow of fluid through the fluid duct from a duct opening to a rotatable member at least partially positioned within the fluid duct. The rotatable member includes a radially inner rotatable portion and a radially outer rotatable portion. The icing effects mitigation system also includes a duct member extending through the fluid duct in a direction approximately orthogonal to a direction of the first flow of fluid. The duct member is configured to channel a second flow of a second fluid therethrough that causes ice accreted on the duct member to shed on a trajectory that impacts the rotatable member at the radially inner portion.

Method and system for ice tolerant bleed takeoff

A gas turbine engine includes a compressor section and a turbine section. The turbine section includes a drive turbine and is located downstream of the compressor section. The gas turbine engine also includes a fan mechanically coupled to and rotatable with the drive turbine such that the fan is rotatable by the drive turbine at the same rotational speed as the drive turbine, the fan defining a fan pressure ratio and including a plurality of fan blades, each fan blade defining a fan tip speed. During operation of the gas turbine engine at a rated speed, the fan pressure ratio of the fan is less than 1.5 and the fan tip speed of each of the fan blades is greater than 1,250 feet per second.

High tip speed gas turbine engine

A blade for the fan section of a turbine engine comprising a composite core defining a pressure side and a suction side extending axially between a core leading edge and a core trailing edge defining a chord-wise direction and extending radially between a core root and a core tip defining a span-wise direction and a leading edge strip made of materials with different elasticity and mounted to the core leading edge.

Turbine engine blade with leading edge strip

Growing community of inventors

Loveland, OH, United States of America

David William Crall