United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

Daniel Bernard Kupratis

Frederick M Schwarz

Paul R Hanrahan

Gabriel L Suciu

William K Ackermann

William G Sheridan

Michael E McCune

Christopher J Hanlon

Francis R Moon

Coy Bruce Wood

Neil J Terwilliger

Jason Husband

Dmitriy B Sidelkovskiy

Zubair Ahmed Baig

Brian D Merry

Gary Collopy

Arthur M Salve, Jr

Michael Winter

David P Houston

Walter A Ledwith

Gary D Roberge

Mark F Zelesky

Anthony R Bifulco

Glenn Levasseur

Michael M Romero

Benjamin T Mylrea

Pitchaiah Vijay Chakka

Arthur W Utay

Michael G McCaffrey

Thomas N Slavens

Tania Bhatia Kashyap

Kevin L Rugg

Nathan L Messersmith

Mark Tu

Howard K Gregory

Sean Danby

Jesse M Chandler

Joseph Brent Staubach

James W Norris

Stephen Karl Kramer

Thomas J Praisner

Mark W Costa

Alan H Epstein

Martin Richard Amari

Geoffrey T Blackwell

Ramesh Rajagopalan

Mary Colby

Andrew P Berryann

William F Schneider

Robert H Bush

Benjamin Bellows

Gary M Stetson

Jams W Norris

Christopher G Hugill

Adam Joseph Suydam

Jacob Shubinski

Flavien Thomas

Michael E Mccune

Gariel L Suciu

Gabriel Suciu

Jr Arthur M Salve

Fredrick M Schwarz

Glen Levasseur

Jr Walter A Ledwith

Paul H Hanrahan

A gas turbine engine includes a compressor section including a first compressor, a turbine section including a first turbine and a second turbine, a first shaft and a second shaft, the first shaft interconnecting the first turbine and the second compressor, and a geared architecture. The first shaft is supported on a first bearing in an overhung manner. A performance ratio is between 0.5 and 1.5.

Gas turbine engine with high speed low pressure turbine section and bearing support features

A system is provided for an aircraft. This aircraft system includes an aircraft and an air system. The airframe includes a body and a vane connected to the body. The vane projects spanwise away from the body to a vane tip. The vane extends longitudinally between a leading edge and a trailing edge. The vane extends laterally between a first vane side and a second vane side. The air system includes a circuit inlet, a circuit outlet and an air circuit. The circuit inlet is arranged with the body and laterally offset from the vane. The circuit outlet is arranged with the vane. The air circuit extends within the airframe from the circuit inlet to the circuit outlet.

Aircraft vertical stabilizer with air system

An assembly is provided for an aircraft. This assembly includes a vertical stabilizer, a first turbine engine, a second turbine engine and an actuation system. The vertical stabilizer extends spanwise from a stabilizer base to a stabilizer tip. The vertical stabilizer extends longitudinally from a stabilizer leading edge to a stabilizer trailing edge. The vertical stabilizer extends laterally between a stabilizer first side and a stabilizer second side. The first turbine engine is movably mounted to the vertical stabilizer at the stabilizer first side. The second turbine engine is moveably mounted to the vertical stabilizer at the stabilizer second side. The actuation system is configured to move the first turbine engine and the second turbine engine relative to the vertical stabilizer.

Aircraft vertical stabilizer with moveable propulsion system(s)

A turbine engine includes integrated electric machines in the compressor section and the turbine section to supplement power produced from fuel with electric power. The example compressor section includes a compressor electric motor that is coupled to a compressor generator. The example turbine section includes a turbine electric motor that is coupled to the geared architecture to supplement power driving the fan section. A turbine generator provides electric power to the turbine electric motor.

Integrated electro-aero-thermal turbine engine

A system for an aircraft includes a vane and an air system. The air system includes a first air circuit and a first door assembly. The first air circuit extends in the vane to a first circuit outlet at a first side of the vane. The first door assembly is arranged at the first circuit outlet. The air system is configured to exhaust a flow of air out of the first air circuit through the first circuit outlet during a first mode and a second mode. The first door assembly is configured to direct the flow of air exhausted from the first circuit outlet longitudinally towards a leading edge of the vane during the first mode. The first door assembly is configured to direct the flow of air exhausted from the first circuit outlet longitudinally towards a trailing edge of the vane during the second mode.

Aircraft airframe component with air system

A gas turbine engine includes a first spool associated with a primary combustor, a second spool associated with a secondary combustor, and a third spool, each spool including a compressor and a turbine mounted to a shaft. A transmission and accessory gearing are enclosed within a housing of an accessory gearbox. The transmission rotationally couples the third spool to the second spool and accessory gearing. A method of operating the gas turbine engine includes supplying a first fuel flow rate to the primary combustor and supplying a second fuel flow rate to the secondary combustor within an intermediate speed range of the gas turbine engine.

Transmission and method for control of boost spool

A powerplant is provided that includes a first gas turbine engine, a second gas turbine engine, a second engine bypass flowpath and a flow control system. The first gas turbine engine includes a first core flowpath fluidly coupled with a first inlet and a first exhaust. The first core flowpath extends sequentially through a first compressor section, a first combustor section and a first turbine section. The second gas turbine engine a second core flowpath fluidly coupled with a second inlet and a second exhaust. The second core flowpath extends sequentially through a second compressor section, a second combustor section and a second turbine section. The flow control system fluidly couples the first inlet and the first exhaust to the second core flowpath during a first mode. The flow control system fluidly couples the first inlet and the first exhaust to the second engine bypass flowpath during a second mode.

Powerplant with multiple integrated gas turbine engines

An aircraft system is provided that includes a first gas turbine engine, a propulsor rotor, a second gas turbine engine, an electric generator and an electric component. The first gas turbine engine includes a first compressor section, a first combustor section, a first turbine section and a first flowpath. The propulsor rotor is rotatably driven by the first gas turbine engine. The second gas turbine engine includes a second compressor section, a second combustor section, a second turbine section and a second flowpath between a second inlet and a second exhaust. The second inlet and the second exhaust are each fluidly coupled with the first flowpath upstream of a combustor of the first combustor section. The electric generator is rotatably driven by the second gas turbine engine. The electric component receives electricity generated by the electric generator. The electric component is discrete from the first gas turbine engine.

Aircraft powerplant with boosted gas turbine engine

A section for a gas turbine engine according to an example of the present disclosure includes, among other things, a rotor including a row of blades extending in a radial direction outwardly from a hub. The row of blades deliver flow to a bypass flow path, an intermediate flow path, and a core flow path. A first case surrounds the row of blades to establish the bypass flow path. A first flow splitter divides flow between the bypass flow path and a second duct. An aftmost row of guide vanes extends in the radial direction across the bypass flow path. A second flow splitter radially inboard of the first flow splitter divides flow from the second duct between the intermediate flow path and the core flow path. A bypass port interconnects the intermediate and bypass flow paths.

Splitter and guide vane arrangement for gas turbine engines

A gas turbine engine according to an example of the present disclosure includes, among other things, propulsor means, first compression means, second compression means, reduction means for reducing a rotational speed of an output that drives the propulsor means relative to an input, first expansion means, and second expansion means. The reduction means includes an epicyclic gear system with a gear reduction. The gear system is straddle-mounted by first and second bearings on opposite sides of the gear reduction relative to an engine longitudinal axis.

A gas turbine engine includes a low speed spool mechanically interconnecting a low pressure turbine and at least one of a fan and a prop, a high speed spool mechanically interconnecting a high pressure turbine and a high pressure compressor, and an epicyclic gear system mechanically coupled to the high speed spool. The gas turbine engine also includes a low pressure compressor mechanically coupled to the high speed spool via the epicyclic gear system. The low pressure compressor may be mechanically independent of the low speed spool. The gas turbine engine may include a plurality of motor-generators for transferring power between the high speed spool and the low pressure compressor.

Low pressure compressor control for a gas turbine engine

A fan has a rotating bulge defining a radially outermost point of the rotating cone relative to a rotational axis. The rotating bulge is defined by a radially outwardly extending forward portion leading to the radially outermost point on the bulge and then a radially inwardly extending portion aft of the radially outermost point of the bulge leading into the core engine path within the splitter housing. The core splitter housing has a leading edge, and the bulge radially outermost point being upstream of the leading edge of the splitter housing. The bulge radially outermost point is a first radial distance from a rotational axis, and the leading edge of the splitter housing being at a second radial distance of the rotational axis, and a ratio of the first radial distance to the second radial distance being at least 0.95.

Gas turbine engine with core debris protection

An assembly is provided for an aircraft propulsion system. This assembly includes an inner wall structure, an outer wall structure, a splitter and an actuation system. The inner wall structure extends axially along a first axis to an inner wall upstream end. The outer wall structure extends axially along the first axis to an outer wall upstream end. The splitter extends axially along the first axis to a leading edge. An inner section of the splitter is connected to the inner wall structure at the inner wall upstream end. An outer section of the splitter is connected to the outer wall structure at the outer wall upstream end and meets the inner section of the splitter at the leading edge. The actuation system is mechanically coupled to the splitter and is configured to move the leading edge radially relative to the first axis.

Actuatable airflow splitter for an aircraft powerplant

A gas turbine engine with hybrid fan exit guide vanes including a fan duct supporting a circumferential pattern of split variable fan exit guide vanes, the split variable fan exit guide vanes comprising an upper section and a lower section, the upper section and the lower section each being adjustable about an axis extending along a span of each of the split variable fan exit guide vanes; an upper actuator in operative communication with the upper section, the upper actuator configured to independently adjust an incidence angle of the upper section responsive to predetermined gas turbine operating conditions; and a lower actuator in operative communication with the lower section, the lower actuator configured to independently adjust an incidence angle of the lower section responsive to the predetermined gas turbine operating conditions.

Gas turbine engine with hybrid fan exit guide vanes

An aircraft propulsion system is disclosed and includes a first gas turbine engine including a first input shaft driving a first gear system, a first fan driven by the first gear system, a first generator supported on the first input shaft and a fan drive electric motor providing a drive input to the first fan, a second gas turbine engine including a second input shaft driving a second gear system, a second fan driven by the second gear system, a second generator supported on the second input shaft and a second fan drive electric motor providing a drive input to the second fan and a controller controlling power output from each of the first and second generators and directing the power output between each of the first and second fan drive electric motors.

Aircraft hybrid propulsion fan drive gear system dc motors and generators

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Wallingford, CT, United States of America

Daniel Bernard Kupratis