General Electric Company

Ge Avio S.r.l.

Ge Aviation Systems Limited

Daniel Alan Niergarth

Brandon Wayne Miller

Randy M Vondrell

Ethan Patrick O'Connor

Darek Tomasz Zatorski

Arthur William Sibbach

Christopher James Kroger

Patrick Michael Marrinan

Jeffrey Douglas Rambo

Ian Francis Prentice

Bugra Han Ertas

Scott Alan Schimmels

Paul Robert Gemin

Richard Alan Wesling

Gert Johannes Van Der Merwe

Nicholas Taylor Moore

Hendrik Pieter Jacobus De Bock

Donald Albert Bradley

Nicholas M Daggett

William Dwight Gerstler

Jeffrey Donald Clements

Christian Xavier Stevenson

Nikolai N Pastouchenko

Thomas Kupiszewski

Brian Lewis Devendorf

Justin Paul Smith

Peter Allen Andrews, Jr

Sridhar Adibhatla

Brian Gene Brzek

Joseph Robert Dickman

Kevin Edward Hinderliter

Lawrence Chih-hui Cheung

David William Crall

Alfred Albert Mancini

Erich Alois Krammer

Arnab Sen

David Justin Brady

Adon Delgado, Jr

Michael Macrorie

Jeffrey S Spruill

Kishanjit Pal

Douglas Downey Turner

John Michael Pyles

Jorge De Luis

Vincenzo Martina

Andrew Michael Tompkins

Keith W Wilkinson

Matthew Kenneth MacDonald

Thomas Lee Becker

Narendra Digamber Joshi

Hao Huang

Christopher Edward Wolfe

Andrew Breeze-Stringfellow

Bhaskar Nanda Mondal

Kurt David Murrow

Alan Roy Stuart

Steven Douglas Johnson

Richard Ludwig Schmidt

Biao Fang

Nathan Evan McCurdy Gibson

Rodrigo Rodriguez Erdmenger

Arthur Vorwerk Radun

Syed Arif Khalid

Kevin Robert Feldmann

John Carl Glessner

David Marion Ostdiek

Jixian Yao

Michele Gravina

Giridhar Jothiprasad

Pranav R Kamat

Raymond Floyd Martell

Matthew Robert Cerny

John Stephen Mears

Adam Joseph Wangler

Kyle Robert Snow

David Vickery Parker

Joel Francis Kirk

Michael Joseph Murray

William James Mailander

Peter David Toot

Robert R Rachedi

Anand Roday

Lorenzo Cipolla

David George Kloos

Samuel Robert Hathaway

Stephanie Glynn Cotten

Stephen William Freund

Michael Robert Bonacum

Martin Manning

Ryan James Matthys

James Laverty

Lawrence Chih-Hui Cheung

A gas turbine includes a turbomachine including a compressor section, a combustion section, and a turbine section arranged in serial flow order. The compressor section includes a high pressure compressor defining a high pressure compressor exit area (A). The turbine section includes a low pressure turbine having at least a first stage of turbine rotor blades and a first stage of turbine stator vanes. The gas turbine engine defines a redline exhaust gas temperature (EGT), a total sea level static thrust output (Fn), an airfoil density (ρ) of one or both of the first stage of turbine rotor blades and the first stage of the turbine stator vanes, and a corrected specific thrust.

Gas turbine engine

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 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 is provided. The gas turbine engine includes: a turbomachine having 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; wherein the gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (Fn) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: Fn×EGT/(A×1000).

A method of detecting an airflow fault condition in a gas turbine engine, the method including: operating the gas turbine engine with a thermal transport bus having an intermediary heat exchange fluid flowing therethrough; determining a performance characteristic of the intermediary heat exchange fluid in the thermal transport bus is outside of a predetermined range, wherein the performance characteristic includes a temperature, a pressure, a flowrate, or a combination thereof; and indicating an airflow fault condition in response to determining the performance characteristic is outside of the predetermined range.

System and method of detecting an airflow fault condition

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.

A gas turbine engine includes a steam generating system and a turbo-engine. The turbo-engine includes a first spool having a fan and a low-pressure turbine drivingly connected together, a second spool having an intermediate-pressure compressor and an intermediate-pressure turbine drivingly connected together, and a third spool having a high-pressure compressor and a high-pressure turbine drivingly connected together. The turbo-engine also includes a first steam turbine arranged to receive a flow of steam from the steam generating system, and arranged to be connected to, and to drive the first spool, and a second steam turbine arranged to receive a flow of steam from the steam generating system, and arranged to be connected to, and to drive one of the second spool or the third spool.

Gas turbine engine having multiple steam turbines

An aircraft gas turbine includes a combustor that combusts compressed air and at least one fuel flow from at least one fuel source to generate combustion gases. A fuel supply system is arranged to provide at least one of a first flow of fuel to the combustor via a first fuel supply line and a second flow of a heated fuel to the combustor via a second fuel supply line. A fuel heat exchanger is arranged within the fuel supply system to generate the heated fuel that is heated above an autoignition temperature of the fuel, and a heat source communicates with the fuel heat exchanger for generating the heated fuel.

Aircraft gas turbine engine with heated fuel to assist combustion

A turbine engine for an aircraft. The turbine engine includes a fan, a combustor in a core air flowpath that generates combustion gases, a core shaft, a turbine that receives the combustion gases to rotate the turbine, and a steam system. The steam system extracts water from the combustion gases, vaporizes the water to generate steam, and injects the steam into the core air flowpath. The steam system includes one or more water storage devices that store the water therein. The one or more water storage devices include a first state in which the one or more water storage devices increase or maintain a level of the water as the water flows through the one or more water storage devices, and a second state in which the one or more water storage devices decrease the level of the water in the one or more water storage devices.

Turbine engine including a steam system

A fan assembly for a gas turbine engine includes a fan disk, a trunnion, an actuation device, a fan blade, and a counterweight assembly. The trunnion is mounted to the fan disk. The actuation device is operably coupled to the trunnion. The fan blade is rotatably attached to the fan disk. The counterweight assembly includes a link arm, a lever arm, a hinge, and a counterweight. The link arm is connected to the trunnion, to the actuation device, or to both. The link arm is configured to drive rotation of the trunnion relative to the fan disk. The hinge is pivotably connected to the lever arm. The lever arm is connected to the link arm and is disposed to rotate about a connection point of the lever arm and the hinge. The counterweight is mounted to the lever arm at a location spaced from the hinge.

Levered counterweight feathering system

A gas turbine has a core turbine engine, and a fan having a plurality of fan blades. A nacelle surrounds the fan and at least a portion of the core turbine engine. The nacelle defines an inlet arranged upstream of the fan, and a bypass flow passage downstream of the fan defined between the nacelle and the core turbine engine. The gas turbine also includes a bypass flow passage water injection system that includes (a) at least one water injection nozzle assembly arranged to inject water into at least one of the inlet of the nacelle, or into the bypass flow passage, and (b) a water injection supply system arranged to supply water from a storage tank to the at least one water injection nozzle assembly. Water is provided by the bypass flow passage water injection system during a high power operating states of the gas turbine to augment thrust.

Turbine engine with fan bypass water injection to augment thrust

A system and method for using a fuel with an engine, an airframe having an aircraft heat load, a fuel tank, and a fuel oxygen reduction unit are provided. The method includes receiving an inlet fuel flow in the fuel oxygen reduction unit for reducing an amount of oxygen in the inlet fuel flow; separating a fuel/gas mixture within the fuel oxygen reduction unit into an outlet gas flow and an outlet fuel flow exiting the fuel oxygen reduction unit; controlling a first portion of the outlet fuel flow to the engine; and controlling a second portion of the outlet fuel flow to the airframe to transfer heat between the second portion of the outlet fuel flow and the aircraft heat load.

Aircraft thermal transport system and method

A fan for a gas turbine engine is provided. The fan includes a plurality of fan blades, a disk, and a trunnion mechanism for attaching the fan blades to the disk. The disk can be formed of a plurality of individual disk segments, with the trunnion mechanism attaching one of the plurality of fan blades to a respective disk segment. A retention member is also provided. The retention member includes a means for catching a portion of the trunnion mechanism should a primary attachment system of the trunnion mechanism fail.

Trunnion retention for a turbine engine

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Cincinnati, OH, United States of America

Daniel Alan Niergarth