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

Christopher James Kroger

Arthur William Sibbach

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

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

Lawrence Chih-hui Cheung

Kevin Edward Hinderliter

Jeffrey Donald Clements

David William Crall

Alfred Albert Mancini

Erich Alois Krammer

Arnab Sen

David Justin Brady

Adon Delgado, Jr

Michael Macrorie

Douglas Downey Turner

Jeffrey S Spruill

Kishanjit Pal

Jorge De Luis

John Michael Pyles

Vincenzo Martina

Andrew Michael Tompkins

Thomas Lee Becker

Matthew Kenneth MacDonald

Keith W Wilkinson

Lawrence Chih-Hui Cheung

Narendra Digamber Joshi

Christopher Edward Wolfe

Hao Huang

Andrew Breeze-Stringfellow

Kurt David Murrow

Bhaskar Nanda Mondal

Alan Roy Stuart

Steven Douglas Johnson

Richard Ludwig Schmidt

Biao Fang

Rodrigo Rodriguez Erdmenger

Nathan Evan McCurdy Gibson

Arthur Vorwerk Radun

Syed Arif Khalid

Kevin Robert Feldmann

John Carl Glessner

David Marion Ostdiek

Michele Gravina

Jixian Yao

Giridhar Jothiprasad

Raymond Floyd Martell

Matthew Robert Cerny

Pranav R Kamat

John Stephen Mears

Adam Joseph Wangler

David Vickery Parker

Kyle Robert Snow

Michael Joseph Murray

Joel Francis Kirk

William James Mailander

Peter David Toot

Anand Roday

Robert R Rachedi

Lorenzo Cipolla

Samuel Robert Hathaway

David George Kloos

Stephanie Glynn Cotten

Stephen William Freund

Michael Robert Bonacum

Martin Manning

Ryan James Matthys

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).

Gas turbine engine

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

A three-stream engine is provided. The three-stream engine includes a fan section, a core engine disposed downstream of the fan section, and a core cowl annularly encasing the core engine and at least partially defining a core duct. A fan cowl is disposed radially outward from the core cowl and annularly encasing at least a portion of the core cowl. The fan cowl at least partially defining an inlet duct and a fan duct. The fan duct and the core duct at least partially co-extending axially on opposite sides of the core cowl. A heat exchanger disposed within the fan duct. The heat exchanger provides for thermal communication between a fluid flowing through fan duct and a motive fluid flowing through the heat exchanger.

Three-stream engine having a heat exchanger

A method of regulating pressure in a thermal transport bus of a gas turbine engine, the method including: operating the gas turbine engine with the thermal transport bus having an intermediary heat exchange fluid flowing therethrough, the thermal transport bus including one or more heat source heat exchangers and one or more heat sink heat exchangers in thermal communication through the intermediary heat exchanger fluid; and adjusting a flow volume of the thermal transport bus using a variable volume device in fluid communication with the thermal transport bus in response to a pressure change associated with the thermal transport bus.

System and method of regulating thermal transport bus pressure

Example split valves for regulating a first flowrate and a second flowrate of a fluid within a closed loop systems are disclosed herein. An example split valve includes an electrohydraulic servo valve coupled to a first piston via a first hydraulic flowline and a second hydraulic flowline, the first piston to include a piston shaft, a first head, and a second head; one or more bellows fixed to at least one of the first head or the second head, the one or more bellows to hermetically seal the fluid from a hydraulic fluid; and a control system connected to the electrohydraulic servo valve, the control system to adjust the first flowrate and the second flowrate of the fluid through a first fluid chamber, the first piston to be located in the first fluid chamber.

Split valves for regulating fluid flow in closed loop systems

Growing community of inventors

Cincinnati, OH, United States of America

Daniel Alan Niergarth