General Electric Company

Ge Avio S.r.l.

Thomas Ory Moniz

Robert Joseph Orlando

Jeffrey Donald Clements

Alan Roy Stuart

Joseph George Rose

Bhaskar Nanda Mondal

Jorge Francisco Seda

Brandon Wayne Miller

Ching-Pang Lee

Richard Ludwig Schmidt

Kattalaicheri Srinivasan Venkataramani

Justin P Stephenson

Joel Francis Kirk

John Christopher Brauer

Steven Alan Ross

Patrick Sean Sage

Mark John Laricchiuta

Randy M Vondrell

Praveen Sharma

Scott Hunter

Atanu Saha

David Glenn Cherry

Vinod Shashikant Chaudhari

John Joseph Rahaim

Daniel Fusinato

Paul Daniel Kemper

Raymond Felix Patt

Ravindra Shankar Ganiger

Darek Tomasz Zatorski

Jan Christopher Schilling

Andrew Breeze-Stringfellow

Ian Francis Prentice

Monty Lee Shelton

Erich Alois Krammer

Anant Pal Singh

Jeffrey Glover

William Andrew Bailey

Scott Michael Carson

Ambika Shivamurthy

John Lewis Baughman

Bhujabal Prashant Mahadeo

Satish Kumar Yadav

Robert John Beacock

John Leslie Henry

Kishanjit Pal

Daniel Robert Dwyer

Hemanth Gudibande Sathyakumar Kumar

William Larson Clifford

James Patrick Dolan

Ann Marie Yanosik

Donald Ray Bond

August H Kramer

Nicholas Joseph Kray

Joshua Tyler Mook

Trevor Howard Wood

Kurt David Murrow

Kishore Ramakrishnan

Byron Andrew Pritchard, Jr

Nitesh Jain

Andrei Tristan Evulet

Sridhar Adibhatla

James Edward Johnson

Rajesh Kumar

Ambrose A Hauser

Richard Alan Wesling

Tsuguji Nakano

Gert Johannes Van Der Merwe

Peter Nicholas Szucs

Thomas Lee Becker, Jr

Robert Alan Frederick

Kenneth Edward Seitzer

Peeyush Pankaj

Arun Kumar

Curtis Walton Stover

Lawrence Wayne Dunbar

Philip Roger Gliebe

Vaishnav Raghuvaran

Nicholas Damlis

Antonio Giuseppe D'Ettole

Robert Harold Weisgerber

Gregory T Steinmetz

Randy Thomas Antelo

Sanjeev Sai Kumar Manepalli

Jordan Tesorero

Vaibhav M Deshmukh

Raymond Gust Holm

Jack Willard Smith, Jr

Paul Cooker

Mohamed Musthafa Thoppil

James William Simunek

Shishir Paresh Shah

Sushilkumar Shevakari

Lawrence Nmn Butler

Jagata Laxmi Narasimharao

Kenneth S Scheffel

David Cory Kirk

Daniel M Allen

Joseph Henry Schleue

Anil Soni

Andrew Mark Del Donno

A passive clearance control limits thermal expansion between stator components relative to rotor components. A control ring controls clearance in a passive manner and is located on or adjacent to stationary components which thermally expand during engine operation. The control ring is formed of material having low coefficient of thermal expansion such as CMCs (Ceramic Matrix Composites) and therefore limits, inhibits or restrains expansion of the adjacent stator components as temperatures increase. Limiting expansion of the stator component reduces rotor/stator clearances and limits parasitic leakage of fluid along the flow path through the engine core.

Passive control of gas turbine clearances using ceramic matrix composites inserts

A gas turbine engine includes a fan located at a forward portion of the gas turbine engine, and a compressor section and a turbine section arranged in serial flow order. The compressor section and the turbine section together define a core airflow path. A rotary member is rotatable with the fan and with a low pressure turbine of the turbine section. The low pressure turbine includes a rotating drum to which a first airfoil structure is connected and extends radially inward toward the rotary member. A torque frame connects the rotating drum to the rotary member and transfers torque from the first airfoil structure mounted to the rotating drum to the rotary member. The torque frame includes an inner disk mounted to the rotary member, an outer ring and a second airfoil structure formed separately from the outer ring and connected thereto by a releasable connecting structure. The second airfoil structure extends radially inward from the outer ring toward the inner disk.

Counter-rotating gas turbine engines including turbine sections with separable torque frames

Structures for reducing forward loads in compressors are described. A compressor includes inner and outer circumferential support structures positioned concentrically around a central axis, and an aft-most stage including a vane extending radially inward from the outer circumferential support structure. An axial length of the aft-most stage is defined by a spacer arm of the inner circumferential support structure. The vane includes a root, a tip, and a trailing edge extending between the root and the tip. A ratio of a first radial distance between a first point located at an intersection of the tip of the vane and the trailing edge to a radially inner wall of the spacer arm of the inner circumferential support structure to a second radial distance between the first point and a second point located at an intersection of the root of the vane and the trailing edge is between 0.95 and 4.5.

Forward load reduction structures for aft-most stages of high pressure compressors

A rotor blade assembly for a turbine engine, including an airfoil blade including an inner diameter end and an outer diameter end, a lower blade carrier coupled to the inner diameter end of the airfoil blade and rigidly coupled to a disk via a pin, an upper blade carrier coupled to the outer diameter end of the airfoil blade, and an outer drum coupled to the upper blade carrier via a radial joint. The radial joint supports radial motion of the upper blade carrier relative to an axis extending through a center of the rotor blade assembly.

Rotor blade assemblies for turbine engines

The present disclosure is directed to a gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction. The gas turbine engine includes a fan assembly including a plurality of fan blades rotatably coupled to a fan rotor in which the fan blades define a maximum fan diameter and a fan pressure ratio. The gas turbine engine further includes a low pressure (LP) turbine defining a core flowpath therethrough generally along the longitudinal direction. The core flowpath defines a maximum outer flowpath diameter relative to the axial centerline. The gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter. The fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater.

Aircraft and direct drive engine under wing installation

An unducted single fan engine includes a housing having one or more fan blades coupled to the housing and configured to rotate circumferentially. The engine has one or more outlet guide vanes coupled to the housing. Each of the one or more guide vanes has a leading edge portion having a variable leading edge. The engine has one or more actuation devices coupled to each of the one or more outlet guide vanes. The one or more actuation devices are configured to control the variable leading edge of the respective outlet guide vane. The variable leading edge is controllable to vary the pitch, camber, lean angle, or sweep of the respective outlet guide vane.

Outlet guide vane

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

A passive clearance control limits thermal expansion between stator components relative to rotor components. A control ring controls clearance in a passive manner and is located on or adjacent to stationary components which thermally expand during engine operation. The control ring is formed of material having low coefficient of thermal expansion such as CMCs (Ceramic Matrix Composites) and therefore limits, inhibits or restrains expansion of the adjacent stator components as temperatures increase. Limiting expansion of the stator component reduces rotor/stator clearances and limits parasitic leakage of fluid along the flowpath through the engine core.

A turbine assembly including a first rotor assembly with a rotatable outer drum from which one or more stages of a plurality of outer drum airfoils is extended radially inward is provided. An outer casing surrounds the outer drum of the first rotor assembly. A seal assembly is coupled to the outer casing and positioned radially outward from an upstream-most stage of the plurality of outer drum airfoils. The seal assembly is positioned in axial alignment with the upstream-most stage of the plurality of outer drum airfoils. The seal assembly separates a first plenum from a second plenum. The second plenum is formed axially aft of the first plenum and is formed by the seal assembly, the outer casing, and the outer drum of the first rotor assembly. The first plenum is positioned radially outward from the upstream-most stage of the plurality of outer drum airfoils.

Flow structure for turbine engine

A gas turbine engine includes a compressor section having a high pressure compressor and a core casing surrounding the compressor section and defining an inlet. The gas turbine engine also includes a cooling system for cooling air in or to the compressor section. The cooling system includes a fluid tank for storing a volume of cooling fluid and a fluid line assembly in fluid communication with the fluid tank. The fluid line assembly includes an outlet positioned upstream of the high pressure compressor and downstream of the inlet defined by the core casing for injecting cooling fluid into an airflow upstream of the high pressure compressor.

Cooling system for a turbine engine

A support assembly for a bearing of a gas turbine engine including a spring finger ring positioned radially exterior to an outer race of the bearing. The spring finger ring includes an outer ring positioned radially exterior to the outer ring, an inner ring positioned radially interior to the outer ring, and a plurality of spring fingers extending between the inner and outer rings. One or more spring fingers configured as two-sided spring fingers including a first ligament coupled to the outer ring and extending at a first circumferential angle to a first radial bumper proximate to the inner ring and a second ligament coupled to the inner ring and extending at a different second circumferential angle to a second radial bumper proximate to the outer ring. The first and second radial bumpers define first and second radial gaps between the bumpers and the inner and outer rings, respectively.

Spring finger ring support assembly for a bearing

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

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Loveland, OH, United States of America

Thomas Ory Moniz