General Electric Company

Other

Alan Roy Stuart

Thomas Ory Moniz

Jeffrey Donald Clements

Darek Tomasz Zatorski

Richard Ludwig Schmidt

Brandon Wayne Miller

John Robert Fehrmann

Jeffrey Douglas Rambo

Christopher Charles Glynn

Robert Joseph Orlando

Bhaskar Nanda Mondal

Robert Proctor

Rajesh Kumar

Gert Johannes Van Der Merwe

Arnab Sen

Leroy H Smith, Jr

Patrick John Lonneman

Joe E Ellis

James Michael Cosgrove

Daniel Alan Niergarth

Jan Christopher Schilling

Andrew Breeze-Stringfellow

Sridhar Adibhatla

Ambrose A Hauser

Richard Alan Wesling

Tsuguji Nakano

David Marion Ostdiek

Jorge Francisco Seda

Joseph George Rose

Joel Francis Kirk

Mark Edward Linz

Guy C Murphy

William Carl Ruehr

Jack D Wright

James William Simunek

Bruce J Gordon

Stephen William Freund

Kenneth S Scheffel

A propulsion system, the propulsion system comprising a rotating element, a stationary element, and an inlet between the rotating element and the stationary element, wherein the inlet passes radially inward of the stationary element; wherein the inlet passes radially inward of the stationary element; wherein the inlet leads to an inlet duct containing a ducted fan having an axis of rotation and a plurality of blades; and wherein the inlet duct divides into a first duct and a second duct, separate from the first duct. A method of operating a propulsion system, comprising the steps of: operating a first rotating fan assembly to produce a first stream of air; directing a portion of the first stream of air into a second ducted rotating fan assembly; operating the second ducted rotating fan assembly to produce a second stream of air; dividing the second stream of air into a core stream and a fan stream; and directing the core stream into a gas turbine engine core.

Propulsion system architecture

A gas turbine engine is provided including a turbine section including a turbine having a plurality of first speed turbine rotor blades; a compressor section including a compressor having a plurality of first speed compressor rotor blades and a plurality of second speed compressor rotor blades; a gearbox; and a first spool rotatable by the plurality of first speed turbine rotor blades, the first spool coupled to the plurality of first speed compressor rotor blades for driving the plurality of first speed compressor rotor blades in a first direction and to the plurality of second speed compressor rotor blades across the gearbox for driving the plurality of second speed compressor rotor blades in a second direction, opposite the first direction.

Turbomachine with alternatingly spaced rotor blades

An apparatus for a turbine engine comprising an outer casing, an engine core provided within outer casing and having a at least one set of blades, and through which gasses flow in a forward to aft direction, an outer drum located within the outer casing to define an annular cavity. A set of seals extending between the first surface and the second surface to define at least one cooled cavity within the annular cavity.

Turbine engine with annular cavity

A gas turbine engine includes a compressor and a turbine; a plurality of rotor blades positioned in the compressor or the turbine and arranged in a plurality of stages along an axial direction; an outer drum positioned at least partially radially outward of the plurality of rotor blades and rotatable with the plurality of rotor blades; and a casing extending over the outer drum and defining a plurality of air cavities with the outer drum, the plurality of air cavities arranged in series airflow communication along the axial direction.

Turbomachine having an airflow management assembly

The present disclosure is directed to a method of control of a gas turbine engine comprising a fan section coupled to a low turbine together defining a low spool, an intermediate compressor coupled to an intermediate turbine together defining an intermediate spool, and a high compressor coupled to a high turbine together defining a high spool. The method includes providing an intermediate spool speed to low spool speed characteristic curve to a controller; providing a commanded power output to the controller; providing one or more of an environmental condition to the controller; determining, via the controller, a commanded fuel flow rate; determining, via the controller, a commanded intermediate compressor loading; and generating an actual power output of the engine, wherein the actual power output is one or more of an actual low spool speed, an actual intermediate spool speed, an actual high spool speed, and an actual engine pressure ratio.

Method of control of three spool gas turbine engine

An engine includes a static frame member; a spool configured to rotate relative to the static frame member during operation of the engine; a gearbox coupled to the spool; and a gearbox coupling assembly mounting the gearbox to the static frame member using a plurality of moveable connection members, each moveable connection member including a sliding connection allowing for movement of the gearbox relative to the static frame member in a plane perpendicular to an axial direction of the engine.

Gearbox coupling in a turbomachine

A gas turbine engine includes a frame member extending generally along an axial direction and defining a plurality of slots spaced along a circumferential direction; and a gearbox including a ring gear, a plurality of planet gears, and a sun gear, the plurality of planet gears each positioned at least partially in a respective slot of the plurality of slots defined by the frame member such that the frame member extends through the gearbox.

Gearbox mounting in a turbomachine

A method includes installing a first stage assembly including a first ring member and a first stage of rotor blades, the first ring member defining a first end and the first stage of rotor blades defining a second end; installing a second stage assembly including a second ring member and a second stage of rotor blades, the second ring member defining a first end and the second stage of rotor blades defining a second end, wherein installing the second stage assembly includes fitting the first end of the second ring member to the second end of the first stage of rotor blades to form a first attachment interface; and pressing the second stage assembly against the first stage assembly to fix the first attachment interface.

A fan case for use in a turbofan engine is provided. The fan case includes an aft portion having a substantially cylindrical cross-sectional shape, and a forward portion extending from the aft portion. A cross-sectional shape of the forward portion progressively decreases in radial size as the forward portion extends from the aft portion.

Fan case for use in a turbofan engine, and method of assembling a turbofan engine

The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction, and a circumferential direction, and an upstream end and a downstream end along the longitudinal direction. The gas turbine engine includes a turbine section, a gearbox proximate to the turbine section, and a driveshaft. The turbine section includes a first rotating component interdigitated with a second rotating component along the longitudinal direction. The first rotating component includes an outer shroud defining a plurality of outer shroud airfoils extended inward of the outer shroud along the radial direction and one or more connecting airfoils coupling the outer shroud to a radially extended rotor. The second rotating component includes an inner shroud defining a plurality of inner shroud airfoils extended outward of the inner shroud along the radial direction. The second rotating component is coupled to an input shaft connected to an input gear of the gearbox. The driveshaft is extended in the longitudinal direction and is connected to an output gear of the gearbox. The first rotating component is coupled to the driveshaft.

Counter rotating turbine with reversing reduction gearbox

The present disclosure is directed to a gas turbine engine including a turbine section including a first rotating component interdigitated along a longitudinal direction with a second rotating component. The first rotating component and the second rotating component are each coupled to a speed reduction assembly in counter-rotating arrangement. The first rotating component comprising an outer shroud and a plurality of outer shroud airfoils extended inward along a radial direction from the outer shroud. A connecting member couples the outer shroud to a radially extended first rotor. The second rotating component comprising an inner shroud and a plurality of inner shroud airfoils extended outward along the radial direction from the inner shroud, the plurality of inner shroud airfoils in alternating arrangement along the longitudinal direction with the plurality of outer shroud airfoils. The gas turbine engine defines a radius per unit thrust defined by a maximum radius at the turbine section over a maximum thrust output between approximately 0.0004 to approximately 0.0010 inches per pound thrust.

Counter rotating turbine with reversing speed reduction assembly

The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction, and a circumferential direction, and an upstream end and a downstream end along the longitudinal direction. The gas turbine engine includes a turbine section and a gear assembly within or downstream of the turbine section. The turbine section includes a first rotating component and a second rotating component along the longitudinal direction. The first rotating component includes one or more connecting airfoils coupled to a radially extended rotor, and the second rotating component includes an inner shroud defining a plurality of inner shroud airfoils extended outward of the inner shroud along the radial direction. The second rotating component is coupled to a second shaft connected to an input accessory of the gear assembly, and the first rotating component is coupled to an output accessory of the gear assembly. The output accessory rotates the first rotating component about the axial centerline at a first speed and wherein the second rotating component rotates about the axial centerline at a second speed.

Counter rotating turbine with reversing reduction gear assembly

The present disclosure is directed to a method of turbine section thermal management for a gas turbine engine. The engine includes a first turbine bearing defining an outer air bearing disposed radially adjacent to a low speed turbine rotor hub of a low speed turbine rotor and an inner bearing disposed radially adjacent to a high pressure (HP) shaft coupled to a high speed turbine rotor, wherein a first manifold is in fluid communication from a pressure plenum of a combustion section to the first turbine bearing, and wherein a second manifold is in fluid communication from the first turbine bearing to a pressure regulating valve and an outer diameter secondary flowpath of the turbine section, and wherein a third manifold is in fluid communication from the pressure plenum of the combustion section to the pressure regulating valve. The method includes flowing air from the pressure plenum to the first turbine bearing; flowing air between the outer air bearing and the low speed turbine rotor hub; flowing air between the inner air bearing and the HP shaft; flowing air from the first turbine bearing to the pressure regulating valve via the second manifold; and flowing air from the first turbine bearing to the outer diameter secondary flowpath via the second manifold.

Method and structure of interdigitated turbine engine thermal management

A system for mounting an engine to an aircraft includes an engine forward mount angled toward the forward end of the engine at a first angle. At least two thrust links extend between an engine aft mount to a link support connection at a second angle. The engine aft mount is angled toward the forward end of the engine at a third angle. A projection of a load vector of the engine forward mount onto a vertical plane extending through the axis of rotation of the engine and a projection of a load vector of each of the at least two thrust links onto the vertical plane intersect the axis of rotation of the engine within a first vertical plane segment extending between a forward end of a nose of a fan assembly and forward of a forward mount interface.

Method and system for mounting an aircraft engine

The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction extended from an axial centerline, and a circumferential direction. The gas turbine engine includes a compressor section, a combustion section, and a turbine section in serial flow arrangement along the longitudinal direction. The gas turbine engine includes a low speed turbine rotor comprising a hub extended along the longitudinal direction and radially within the combustion section; a high speed turbine rotor comprising a high pressure (HP) shaft coupling the high speed turbine rotor to a HP compressor in the compressor section; a first turbine bearing disposed radially between the hub of the low speed turbine rotor and the HP shaft. The HP shaft extends along the longitudinal direction and radially within the hub of the low speed turbine rotor. The first turbine bearing defines an outer air bearing along an outer diameter of the first turbine bearing and adjacent to the hub of the low speed turbine rotor. The first turbine bearing defines an inner air bearing along an inner diameter of the first turbine bearing and adjacent to the HP shaft.

Growing community of inventors

Cincinnati, OH, United States of America

Alan Roy Stuart