General Electric Company

U.S. Department of Energy

General Electric Company Polska Sp. Z O.o.

Jeffrey Douglas Rambo

Brandon Wayne Miller

Arthur William Sibbach

Scott Alan Schimmels

William Dwight Gerstler

Thomas Kupiszewski

Daniel Alan Niergarth

Robert Proctor

Jared Matthew Wolfe

Nicolas Kristopher Sabo

Bhaskar Nanda Mondal

Rajesh Kumar

Arnab Sen

Gregory Michael Petrasko

Scott Gregory Edens

Daniel Jason Erno

Steven Douglas Johnson

Michael Thomas Kenworthy

Kirk Douglas Gallier

Hendrik Pieter Jacobus De Bock

Ramon Martinez

Timothy Richard DePuy

Steven B Morris

Robert R Rachedi

Paul Alexander Intemann

Christopher Charles Glynn

Alan Roy Stuart

Aaron Michael Dziech

Craig Alan Gonyou

David Alan Frey

Kevin Robert Feldmann

William Joseph Bowden

Alexander Martin Sener

Nicholas Taylor Moore

Arvind Namadevan

James Fitzgerald Bonar

Matthew Thomas Beyer

Emily Rosette Clark

Michael John Simonetti

Michael Stephen Popp

Curt Edward Hogan

Ravindra Shankar Ganiger

David Raju Yamarthi

Mohammed El Hacin Sennoun

Donald Michael Corsmeier

John Alan Manteiga

Scott Hunter

Patrick Michael Marrinan

Nathan Evan McCurdy Gibson

Brian Gene Brzek

Erich Alois Krammer

Narendra Anand Hardikar

John Carl Glessner

David Marion Ostdiek

Curtis Walton Stover

Raymond Floyd Martell

Brian Lewis Devendorf

Mark David Durbin

James Michael Kostka

Matthew Robert Cerny

Justin Paul Smith

Neil Richard Garrigan

Vinod Shashikant Chaudhari

Andrew J Martin

Pradeep Hemant Sangli

Michael Joseph Murray

Robert Jon McQuiston

Joel Francis Kirk

David Perveiler

Kevin Edward Hinderliter

Julius John Montgomery

Santosh Kumar Potnuru

John William Moores

Brian Gregg Feie

Tarun Sharma

Bradley Willis Fintel

Erich R Fitzpatrick

Hari Shanker Thakur

Ryan T Roehm

Alex William Ariapad

Victor Moreno Patan

Matthew Ryan Johns

Vaibhav M Deshmukh

Efren Souza Chavez

Tajinder Singh

Bryan Jung Woo Hong

Sushilkumar Shevakari

Nicholas Allan Wilson

Sridhara Rao Venkata Mittapalli

Sarah Anne Edwards

Ryan James Matthys

Yu-Liang Lin

Michal Starzynski

Siddhartha Pal

Eric Barre

Leonardo Aguilar Willys

Rakshit Tirumala

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 propulsion system includes (a) a turbine engine including a compressor section, a combustor, and a turbine section, (b) a ramjet engine, and (c) an inlet section including a turbine engine inlet portion and a ramjet inlet portion, the inlet section providing an inlet airflow to the turbine engine and to the ramjet engine. The turbine engine inlet portion includes an expansion turbine arranged upstream of the compressor section.

Aircraft propulsion system

A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (L), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.

Gas turbine engine having a heat exchanger located in an annular duct

A gas turbine engine is provided, having: a bifurcation; and a turbomachine further including a cooled cooling air (CCA) system, the CCA system having: a cold side bleed assembly defining an inlet positioned to be in fluid communication with an airflow over the bifurcation; a CCA heat exchanger in thermal communication with the cold side bleed assembly downstream of the inlet of the cold side bleed assembly; and a hot side bleed assembly defining an inlet in fluid communication with a working gas flowpath through a compressor section, at a compressor discharge cavity, or both, the hot side bleed assembly in thermal communication with the CCA heat exchanger to cool an airflow through the hot side bleed assembly, the hot side bleed assembly further in thermal communication with a hot component of the turbomachine to cool the hot component of the turbomachine.

Gas turbine engine having cooling systems

A turbine engine includes a cooling air duct for cooling air positioned radially between a core air flow path for core air and a bypass airflow passage for bypass air. A heat exchanger is positioned in the cooling air duct to transfer heat from a heat source from within the turbine engine. The heat exchanger may be a condenser. The turbine engine may further include a steam system that extracts water from the combustion gases, vaporizes the water to generate steam, and injects the steam into the core air flow path, the steam system including the condenser to transfer heat from the combustion gases to the cooling air and to condense the water from the combustion gases. The turbine engine may further include a booster fan to increase the pressure of the cooling air and the core air.

Turbine engine with three air streams

A turbine engine for an aircraft including a combustor, a turbine, and a steam system. The combustor generates combustion gases. The turbine receives the combustion gases and is drivingly coupled to a shaft. The steam system includes a boiler, a superheater, and a steam turbine. The boiler receives water and is fluidly connected to the combustor to receive a first portion of the combustion gases from the combustor and to boil the water to generate steam. The superheater is located downstream of the combustor to receive a second portion of the combustion gases from the combustor and to heat the steam. The steam turbine is fluidly coupled to the superheater to receive the steam from the superheater and to rotate the steam turbine. The steam turbine is drivingly coupled to the shaft.

Aircraft gas turbine engine including a steam system

A turbine engine for an aircraft. The turbine engine includes a combustor fluidly coupled to a fuel delivery assembly to receive fuel from the fuel delivery assembly. The fuel is injected into the combustor and combusted in the combustor to generate combustion gases. A condenser is located downstream of a turbine to receive the combustion gases and to condense water. The fuel heat exchanger is thermally coupled to the condenser to receive heat from the water condensed by the condenser. The fuel heat exchanger is located in the fuel delivery assembly to receive the fuel and to transfer the heat received from the water to the fuel. The boiler is located downstream of the fuel heat exchanger. The boiler receives the water and is fluidly connected to the combustor to receive the combustion gases and to boil the water to generate steam.

Turbine engine including a steam system

A vehicle is provided including a structure including a skin defining an outside surface exposed to ambient cooling flow and an inside surface. The structure includes a first structural member extending from the inside surface of the skin and a second structural member extending from the inside surface of the skin; and a thermal management system including a heat exchanger assembly positioned adjacent to, and in thermal communication with, the inside surface of the skin, the heat exchanger assembly positioned at least partially between the first and second structural members of the structure.

Vehicle heat exchanger system including an inflatable operable with a cooling tube

A propulsion system includes a turbine engine and a ramjet engine. A turbine engine inlet section restriction device permits an inlet airflow to a turbine engine inlet in a first (low speed) operating state of the propulsion system, and restricts the inlet airflow to the turbine engine inlet in a second (high speed) operating state of the propulsion system. A combustor inlet flow control device, in the first operating state, provides an airflow from a compressor to the turbine engine combustor, and restricts an airflow from a ramjet bleed air passage, and, in the second operating state, provides a ramjet inlet bleed airflow to the turbine engine combustor, while restricting airflow from the compressor. In the second operating state, the turbine engine drives at least one accessory gearbox to drive at least one accessory component.

Propulsion system using ramjet bleed air for gas turbine combustion

A turbine engine having a longitudinal centerline axis. The turbine engine including a fan comprising a plurality of fan blades that rotate about the longitudinal centerline axis and a rotational component coupled to the fan. The turbine engine including a fluid circuit for supplying fuel or lubricant to the turbine engine and a hydraulic fan brake coupled to the fluid circuit to prevent rotation of the rotational component, thus preventing rotation of the fan. The hydraulic fan brake including a hydraulic cylinder fluidly coupled to the fluid circuit and a valve coupled to the hydraulic cylinder and having a first valve position that disengages the hydraulic fan brake to allow rotation of the rotational component and a second valve position that engages the hydraulic fan brake to prevent rotation of the rotational component.

Gas turbine engine having a hydraulic fan brake

Flow paths and boundary layer restart features are provided. For example, a flow path comprises a flow path wall defining an inner flow path surface and an asymmetric notch defined in the flow path wall. The asymmetric notch comprises a first surface and a second surface and is asymmetric about a first line extending through an intersection of the first and second surfaces. Further, a flow boundary layer restart feature comprises a first surface extending inward with respect to a flow path surface of a flow path and a second surface extending inward with respect to the flow path surface. The second surface is asymmetric with respect to the first surface such that the first and second surfaces define an asymmetric notch. Additionally, a flow path wall may comprise an asymmetric notch that includes a flow expansion angle and a flow contraction angle that are unequal.

Asymmetric flow path topology

A method for energy conversion for a vehicle is provided. The method including extracting a flow of compressed fluid from a compressor section of a propulsion system; flowing the flow of compressed fluid to a turbine operably coupled to a driveshaft, in which the driveshaft is operably coupled to a load device; expanding the flow of compressed fluid through the turbine to generate an output torque at the driveshaft to operate the load device; and flowing the expanded flow of compressed fluid from the turbine to thermal communication with a thermal load.

Vehicle with energy conversion system

A thermal management system includes a thermal circuit having a thermal transport bus with a thermal fluid flowing therethrough. The thermal transport bus includes first and second bus segments. First and second thermal loads are on the first and second bus segments, respectively, and have different first and second inlet temperature requirements. The thermal management system includes a plurality of heat exchangers in thermal communication with the first and second thermal loads. The heat exchangers include a first heat exchanger and one or more second heat exchangers. The first heat exchanger is on one of the first or second bus segments for rejecting heat to engine fuel. The thermal fluid is split between each of the first and second bus segments such that only a portion of the thermal fluid flows through the first heat exchanger to accommodate the different first and second inlet temperature requirements.

Thermal management system for an engine

An aircraft propulsion system in includes a compressor section, a heat addition system, and an expansion section in serial flow arrangement, wherein the heat addition system is configured to receive a flow of gaseous fuel from a fuel system, and wherein the flow of gaseous fuel is provided to the heat addition system to generate combustion gases. A lubricant system is configured to provide a flow of lubricant through the propulsion system. A hydraulic system is configured to provide a flow of hydraulic fluid to an actuator at the propulsion system. A thermal management system is configured to provide a flow of fluid in thermal communication with the lubricant system and the hydraulic system. The thermal management system comprises a turbine configured to receive the flow of fluid via a fluid circuit extended in fluid communication from the compressor section.

Vehicle and propulsion system with hydrogen fuel system

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Mason, OH, United States of America

Jeffrey Douglas Rambo