General Electric Company

Ge Avio S.r.l.

General Electric Company Polska Sp. Z O.o.

Ge Aviation Systems Limited

General Electric Deutschland Holding Gmbh

Arthur William Sibbach

Brandon Wayne Miller

Nicholas Joseph Kray

Jeffrey Douglas Rambo

Daniel Alan Niergarth

Bugra Han Ertas

Andrew Hudecki

Andrea Piazza

John C Schilling

Scott Alan Schimmels

Tsuguji Nakano

Robert R Rachedi

Xiaohua Zhang

Elzbieta Kryj-Kos

Brian Lewis Devendorf

Gary Willard Bryant, Jr

Miriam Manzoni

Sean Christopher Binion

Flavia Turi

Ravindra Shankar Ganiger

Michael Anthony Benjamin

Jeffrey Donald Clements

Ethan Patrick O'Connor

Erich Alois Krammer

Frank Worthoff

James Ryan Reepmeyer

Paul Burchell Glaser

Egbert Geertsema

Adam Tomasz Pazinski

Benjamin Thomas Van Oflen

Ryan T Roehm

Vincenzo Martina

Timothy Richard Depuy

Bhaskar Nanda Mondal

Ian Francis Prentice

Wendy Wenling Lin

Aaron Michael Dziech

Scott Hunter

Christopher James Kroger

John Carl Glessner

Gregory Carl Gemeinhardt

Michael John Franks

Timothy Richard DePuy

Michael Macrorie

Vaibhav M Deshmukh

Lawrence William Nurre

Douglas Downey Turner

Sushilkumar Shevakari

Rakshit Tirumala

Allan George Van De Wall

Jorge De Luis

Keith W Wilkinson

Hao Huang

Mohamed Osama

Randy M Vondrell

Kirk Douglas Gallier

Joseph Alfred Iannotti

Ming Xie

Constantinos Minas

Daniel Edward Mollmann

Patrick Michael Marrinan

Rajesh Kumar

Richard Alan Wesling

Brian Gene Brzek

Stefan Joseph Cafaro

Eric Richard Westervelt

David Marion Ostdiek

Andrew J Martin

William Joseph Solomon

Donald Scott Yeager

Vishnu Vardhan Venkata Tatiparthi

Kevin Edward Hinderliter

Adam Tomasz Paziński

Aaron Todd Sellinger

Tomasz Iglewski

Douglas Lorrimer Armstrong

Nagashiresha G

Koji Asari

Douglas Scott Jacobs

Kedar S Vaidya

Ahmet Dindar

Randy W Vondrell

James Laverty

Michał Tomasz Kuropatwa

Robert Eugene Dyson

Brent Michael Push

Avinash Natarajan

Prateek Mathur

A gas turbine engine comprises a fan, a core turbine engine coupled to the fan, a fan case housing the fan and the core turbine engine, a plurality of outlet guide vanes extending between the core turbine engine and the fan case, and an acoustic spacing. The fan comprises a plurality of fan blades that define a fan diameter and a BEAL. The fan case comprises an inlet and an inlet length between the inlet and the fan. The acoustic spacing comprises a distance between the fan and the plurality of outlet guide vanes, and in combination with the BEAL determines an acoustic spacing ratio of the gas turbine engine. The combination of acoustic spacing and corrected specific thrust provide enhanced propulsive efficiency.

Gas turbine engine with acoustic spacing of the fan blades and outlet guide vanes

A gas turbine engine comprises a fan, a core turbine engine coupled to the fan, a fan case housing the fan and the core turbine engine, a plurality of outlet guide vanes extending between the core turbine engine and the fan case, and an acoustic spacing. Relationships between acoustic spacing and a high-speed shaft rating allow for a gas turbine engine that reduces noise emissions while maintaining high performance.

Aircraft engines and high temperature anti-ice systems for aircraft engines are disclosed herein. An example aircraft engine includes: a fan including a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine including a compressor section, a combustion section, and a turbine section; a supply duct to accept bleed air from the compressor section; and a heat exchange system to capture waste heat from the turbine section and convey the waste heat to the bleed air, the bleed air with the waste heat to be conveyed to at least one of an environmental control system or a wing of an aircraft.

High temperature anti-ice systems for aircraft engines

A lubrication system for a turbine engine that includes one or more rotating components. The lubrication system includes one or more tanks that store lubricant, a primary lubrication system, and an auxiliary lubrication system. The primary lubrication system supplies the lubricant from the one or more tanks to the one or more rotating components during stable operating conditions of the lubrication system. The auxiliary lubrication system includes an auxiliary feed line and an auxiliary supply line. The auxiliary lubrication system receives the lubricant from the one or more tanks through the auxiliary feed line. The auxiliary lubrication system supplies the lubricant to the one or more rotating components through the auxiliary supply line when there is a potential lubricant interruption in the lubrication system.

Lubrication system for a turbine engine

A gas turbine engine comprises a fan, a core turbine engine coupled to the fan, a fan case housing the fan and the core turbine engine, a plurality of outlet guide vanes extending between the core turbine engine and the fan case, and an acoustic spacing. Acoustic spacing comprises a distance between the fan and the plurality of outlet guide vanes, and in combination with the use of airfoils with leading edge protectors for both various stages of blades and vanes, the engine effectively reduces noise emissions, enhances aerodynamic efficiency, and improves structural durability.

A gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio is defined by Vis a gutter volume of the gutter and Vis a gearbox volume of the gearbox. A gas turbine engine includes the gearbox assembly and a lubrication system. The lubrication system includes a tank that stores a lubricant therein, one or more lubricant pumps for generating a flow of lubricant from the tank to the gearbox assembly, and a flow rate control valve. The flow rate control valve has a flow inlet and a flow outlet and defines a variable throughput from the flow inlet to the flow outlet for controlling a flowrate of the lubricant to the gearbox assembly.

Gearbox assembly with lubricant extraction volume ratio

A mounting assembly for a gearbox assembly of a gas turbine engine includes at least one mounting member configured to mount a gear of the gearbox assembly to a component of the gas turbine engine, the at least one mounting member characterized by a lateral impedance parameter, a bending impedance parameter, and a torsional impedance parameter. A gas turbine engine includes the mounting assembly. The at least one mounting member may be a flex mount, a fan frame, or a flex coupling. The gas turbine engine includes an electric power system including at least one electric machine. The electric power system includes a plurality of power converters and a plurality of power distribution management units. At least two of the plurality of power converters or the plurality of power distribution management units are integrated together in a single housing.

Mounting assembly for a gearbox assembly

A propulsion and electrical system for an aircraft including a turbine engine and a fuel cell. The turbine engine includes a turbo-engine, a fan having a fan shaft coupled to the turbo-engine, and a steam system. The steam system is fluidly coupled to a water source to receive water. The steam system includes a boiler to vaporize the water and to generate steam. The steam system can be fluidly coupled to the core air flow path to inject the steam into the core air flow path. The fuel cell is fluidly coupled to a hydrogen source and an oxygen source to receive hydrogen and oxygen, respectively, and, when receiving the hydrogen and the oxygen, to generate electricity and water. The fuel cell is fluidly coupled to the steam system as the water source.

Propulsion and electrical system for an aircraft including a fuel cell and turbine engine with a steam system

A gas turbine engine for an aircraft includes a fan and a turbo-engine having a combustor that combusts compressed air and fuel to generate combustion gases and a power turbine coupled to a shaft for driving the fan. The gas turbine engine includes a direct contact condenser for condensing steam in a first flow into liquid water droplets and to generate a second flow including the liquid water droplets, a water tank for collecting the liquid water droplets of the second flow, and a water pump for introducing the liquid water droplets from the water tank into the direct contact condenser to act as condensation nuclei for the steam in the first flow. The first flow is an exhaust-vapor flow extracted from the combustion gases and the second flow is an exhaust-condensate flow from the direct contact condenser.

Turbine engine including a condenser system

An aircraft engine assembly includes a gas turbine engine having an intake channel configured to receive an incoming flow of air and thereby form an intake flow of air, the intake channel configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine, the incoming flow direction reverse of the first axial direction, and an electric machine coupled with the low pressure shaft and located at the aft end of the gas turbine engine proximate the intake channel, the electric machine in heat exchange communication with the intake flow of air such that the electric machine transfers heat to the incoming flow of air within the intake channel when the electric machine is operated.

Reverse flow gas turbine engine having electric machine

A turbine engine for an aircraft including a turbo-engine, a fan having a fan shaft coupled to the turbo-engine, and a steam system. The steam system is fluidly coupled to a core air flow path at a plurality of steam injection locations to selectively inject steam into the core air flow path at each of the plurality of steam injection locations, using a steam flow control valve. The plurality of steam injection locations include (i) an upstream steam injection location located to inject steam into the core air flow path at a primary combustion zone of a combustor or upstream thereof and (ii) a downstream steam injection location located to inject steam into the core air flow path downstream of the primary combustion zone.

Turbine engine including a steam system

An engine component for a turbine engine. The engine component has a composite structure and a cover structure. The composite structure has a composite structure outer wall and a composite structure edge. The cover structure encases at least a portion of the composite structure outer wall. The cover structure has a main body. The main body extends along the at least a portion of the composite structure outer wall.

Engine component

A turbine engine with an engine core defining an engine centerline and comprising a rotor and a stator. The turbine engine including a set of composite airfoils circumferentially arranged about the engine centerline and defining at least a portion of the rotor. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion.

Turbine engine with composite airfoils

A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (A) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (A) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (T) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows: wherein CSP is greater than 0.0001194×EGT−0.103×EGT+22.14 and less than 0.0003294×EGT−0.306×EGT+77.91; and wherein EGT is greater than 525 degrees Celsius and less than 1250 degrees Celsius.

Gas turbine engine

A turbofan engine defining an axial direction and a longitudinal centerline along the axial direction is provided. The turbofan engine includes: a fan section having a fan, the fan comprising a plurality of fan blades; a turbomachine drivingly coupled to the fan, the turbomachine comprising a compressor section with a low pressure compressor, a turbine section with a low pressure turbine, a reduction gearbox, and an outer casing, the low pressure turbine drivingly coupled to the low pressure compressor across the reduction gearbox; an outer nacelle surrounding the fan and at least a portion of the turbomachine; an outlet guide vane extending between the turbomachine and the outer nacelle at a location downstream of the plurality of fan blades, the outlet guide vane defining a base and a tip and being forward swept from the base to the tip.

Gas turbine engine with forward swept outlet guide vanes

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Boxford, MA, United States of America

Arthur William Sibbach