Rtx Corporation

Raytheon Technologies Corporation

United Technologies Corporation

Raytheon Technology Corporation

Jon E Sobanski

Steven D Porter

Howard J Liles

Jacob C Snyder

Andrew E Breault

Bryan P Dube

Jeffrey T Morton

Bryan H Farrar

Neil J Terwilliger

Joseph Brent Staubach

Robert A White, Iii

Tyler G Vincent

Jesse M Chandler

Paul R Hanrahan

Abbas A Alahyari

Kathryn S Read

Alyson T Burdette

Wojciech Lipka

Marc J Muldoon

Raymond C Surace

Ying She

Jonathan Jeffery Eastwood

David Wasserman

James F Wiedenhoefer

Thomas Gregory Tillman

Malcolm P MacDonald

Simon W Evans

Cheng Gao

Stuart S Ochs

Brandon M Evans

Lucas Dvorozniak

Rutva A Patel

Chang Gao

Kenneth Petroski

Jill Klinowski

Robert A White Iii

Iii Robert A White

Malcolm Macdonald

A fan drive turbine drives a drive shaft. A compressor section and a turbine section rotate about a rotational axis, with the turbine section positioned in a downstream direction. A fan has fan blades and a fan hub positioned radially within a fan case, with the fan case defining a bypass duct. A fan inlet guide vane case is upstream of the fan blades and a fan exit guide vane case is mounted downstream. The drive shaft is mounted on at least one input drive shaft bearing inwardly of the fan inlet guide vane case. The drive shaft drives a flexible coupling to in turn drive a gear reduction. The gear reduction drives a fan drive shaft to in turn drive the fan hub. At least two fan drive shaft bearings support the fan drive shaft within the fan exit guide vane case. An aircraft is also disclosed.

Aft mounted pusher fan for gas turbine engine

A turbine engine is provided that includes a first rotating assembly, a bearing, a turbine engine core and a flowpath. The first rotating assembly includes a propulsor rotor, a power turbine rotor and a power turbine shaft coupled with and axially between the propulsor rotor and the power turbine rotor. The bearing rotatably supports the first rotating assembly at an intermediate location along the power turbine shaft. The turbine engine core includes a second rotating assembly and a combustor. The second rotating assembly includes a core compressor rotor and a core turbine rotor. The power turbine rotor is arranged axially between the propulsor rotor and the second rotating assembly. The flowpath extends sequentially across the core compressor rotor, the combustor, the core turbine rotor and the power turbine rotor from an inlet into the flowpath to an exhaust from the flowpath.

Supporting elongated rotating assembly within a turbine engine

A propulsion system for an aircraft includes a propulsor section that includes a fan with a plurality of fan blades that are rotatable about a fan axis, a core engine that is configured to generate a gas flow, a power turbine that is mechanically uncoupled from the core engine and rotatable independent of the core engine, the gas flow from the core engine is expanded through the power turbine to generate shaft power, a support structure where the power turbine is mounted relative the core engine, a fan drive gear system that is driven by the power turbine and at least partially mounted to the support structure, and a power turbine shaft that includes a first coupling to the fan drive gear system and a second coupling to the propulsor section.

AFT gear based engine with heat recovery system

A fuel system for a gas turbine engine includes a manifold assembly with an assembly inlet, an assembly outlet, and a plurality of fuel injectors. The manifold assembly also includes at least one fuel manifold fluidically connecting each fuel injector of the plurality of fuel injectors to the assembly inlet and to the assembly outlet. A fuel return line is fluidically connected to the assembly outlet. A fuel circulator pump includes a pump outlet fluidically connected to the assembly inlet.

Continuous flow fuel system

A grid arrayed microtube heat exchanger with vibration dampening support including an upper portion comprising an upper portion support wall having multiple upper portion receivers; a lower portion comprising a lower portion support wall having multiple lower portion receivers; a grid array comprising multiple rows of the lower portion receivers and the upper portion receivers; multiple microtubes supported by the upper portion receivers and the lower portion receivers; a gap located between each microtube; and a support insertable through the gap between the multiple microtubes, the support including at least one cam contacting the microtube, the at least one cam being rigid.

Grid arrayed microtube heat exchanger with midspan support components

A turbine engine is provided that includes a propulsor rotor and an engine core. The propulsor rotor is rotatable about a propulsor axis. The engine core is configured to power operation of the propulsor rotor. The engine core includes a core compressor section, a core combustor section, a core turbine section, a first rotating structure and a second rotating structure. The first rotating structure includes a first compressor rotor arranged within the core compressor section. The first rotating structure is rotatable about a first structure axis which is offset from the propulsor rotor axis. The second rotating structure includes a second compressor rotor arranged within the core compressor section. The second rotating structure is rotatable about a second structure axis which is offset from the propulsor rotor axis and the first structure axis.

Aircraft propulsion system engine with multiple independent rotating structures

A propulsion system for an aircraft includes a gas generating core engine generates an exhaust gas flow that is expanded through a turbine section, a power turbine driven by the exhaust gas flow, a propulsor coupled to the power turbine, a hydrogen fuel system configured to supply hydrogen fuel to the combustor through a fuel flow path, a condenser arranged along the core flow path and configured to extract water from the exhaust gas flow, and an evaporator arranged along the core flow path receiving a portion of the water extracted by the condenser to generate a steam flow that is injected into the core flow path upstream of the turbine section.

Reverse flow hydrogen steam injected turbine engine

A heat exchanger for heat transfer between an external first flow along a first flowpath and a second flow along an internal second flowpath, has: a first manifold; a second manifold; and a plurality of tubes extending from the first manifold to the second manifold and having respective interiors bounding respective legs of the second flowpath. The plurality of tubes comprises a plurality groups of tubes. For each of the groups of the tubes: the tubes of the group have first ends mounted to the first manifold at respective first locations; and the tubes of the group have second ends mounted to the second manifold at respective second locations. From the first manifold to the second manifold, each tube has: a upstream concave first turn; an upstream convex second turn; and an upstream concave third turn; and the second locations are offset downstream along the first flowpath from the respective first locations.

Aircraft heat exchanger

A propulsion system for an aircraft includes an open propulsor rotor and a turbine engine. The turbine engine includes a fan section, an engine core, a first geartrain, a core flowpath and a bypass flowpath. The fan section includes a fan rotor. The engine core includes a first rotating assembly, a low pressure compressor section, a high pressure compressor section, a combustor section, a high pressure turbine section and a low pressure turbine section. The first rotating assembly drives rotation of the open propulsor rotor and the fan rotor through the first geartrain. The core flowpath extends through the low pressure compressor section, the high pressure compressor section, the combustor section, the high pressure turbine section and the low pressure turbine section. The bypass flowpath extends outside of and along the engine core.

Open rotor aircraft propulsion system with bypass flowpath

A propulsion system for an aircraft includes an open propulsor rotor and a turbine engine configured to drive rotation of the open propulsor rotor about an axis. The turbine engine includes a core flowpath, a bypass flowpath, a fan section and an engine core. The engine core includes a compressor section, a combustor section and a turbine section. The core flowpath extends through the compressor section, the combustor section and the turbine section from an inlet into the core flowpath to an exhaust from the core flowpath. The inlet is next to and downstream of the fan section. The bypass flowpath extends outside of and along the engine core from an inlet into the bypass flowpath to an exhaust from the bypass flowpath. The inlet is next to and downstream of the fan section. The engine core is arranged axially between the fan section and the open propulsor rotor.

A tooling assembly suitable for use in infiltrating a fibrous preform comprises an outer tooling fixture and an inner tooling fixture. The outer tooling fixture comprises a first portion removably attached to a second portion, and a first pair of oppositely disposed windows defined by the first portion and the second portion. Each of the first portion and the second portion comprise a base with a recess. The inner tooling fixture comprises a body defined by at least one portion, the body configured to at least partially surround the fibrous preform, a plurality of holes extending through a thickness of the at least one portion, each hole of the plurality of holes having a length (L) and a diameter (D), and a first neck region engageable with a recess of the first portion or the second portion.

Floating tooling assembly for chemical vapor infiltration

A gas turbine engine support system includes a support arc segment piece that defines radially inner and outer sides, axially forward and aft sides, and first and second circumferential mate face sides. The radially inner side has a radially open socket that is configured to engage a mating component and transmit loads from the mating component. The radially open socket is elongated in an axial direction with respect to the axially forward and aft sides The support arc segment piece has radial and axial reaction surfaces that are configured, respectively, to transmit radial and axial components of the loads from the mating component.

Vane support system

In one aspect, a panel for lining a combustor shell of a combustor of a gas turbine engine includes a front surface configured to face a combustion chamber and a back surface configured to face the combustor shell, a plurality of cooling holes extending from the back surface to the front surface, a rail extending from the back surface and defining a perimeter of the panel, and a plurality of studs disposed adjacent to and connected to the rail. The rail is configured to contact the combustor shell. The plurality of studs are configured to be received in a plurality of holes of the combustor shell. The plurality of studs is asymmetrically arranged around the perimeter of the panel.

Combustor panel with rail integrated perimeter studs

An apparatus comprises a variable pitch blade configured to connect with an endwall of an open rotor engine. The variable pitch blade defines a chamber therein on a bottom edge thereof. A retractable edge member pivotally connects within the chamber to move between a first position wherein at least a portion of the retractable edge member is located within the chamber and a second position wherein a portion of the retractable edge member extends downward from the bottom edge of the variable pitch blade to block a gap between the bottom edge of the variable pitch blade and the endwall.

Open rotor variable pitch blade with retracting inboard trailing edge

An aircraft propulsion includes a core engine section that defines a core flow path where an inlet airflow is compressed, mixed with fuel, and ignited to generate an exhaust gas flow, an inner nacelle assembly that surrounds the core engine section, an outer nacelle assembly that is spaced radially apart from the inner nacelle assembly, and a strut heat exchanger that extends radially between the inner nacelle assembly and the outer nacelle assembly, within the strut heat exchanger a portion of the exhaust gas flow is placed in thermal communication with a second flow for transferring thermal energy.

Growing community of inventors

Glastonbury, CT, United States of America

Jon E Sobanski