United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

Paul M Lutjen

Timothy M Davis

Christopher W Strock

Ken Francis Blaney

Susan M Tholen

Dominic J Mongillo, Jr

Anthony B Swift

Jose R Paulino

Richard K Hayford

Christopher R Joe

Mark John Rogers

Dmitriy A Romanov

Thurman Carlo Dabbs

Carson A Roy Thill

Gary L Grogg

James N Knapp

Michael J Bruskotter

Matthew A Devore

John R Farris

Neil L Tatman

Jeremy Drake

Kevin J Ryan

Brian T Hazel

Mario P Bochiechio

John A Sharon

Shawn J Gregg

James F Wiedenhoefer

Brian C McLaughlin

Winston Gregory Smiddy

Russell E Keene

Lawrence A Binek

Eric A Hudson

Joel H Wagner

Daniel S Rogers

Philip Robert Rioux

Jonathan J Earl

Patrick D Couture

Joseph Micucci

Terence P Tyler

Morely S Sherman

Bruce E Chick

Timothy J Harding

Virginia L Ross

Christopher M Jarochym

Tracy A Propheter-Hinckley

Michael G McCaffrey

Mark F Zelesky

Thomas E Clark

Russell J Bergman

Gregory E Reinhardt

Thomas J Praisner

Michael Papple

John R Paulus

Mohammed Ennacer

Michael F Blair

Nicholas R Leslie

Joseph T Caprario

David Richard Griffin

Blake J Luczak

Richard W Hoff

Richard H Page

John D Wiedemer

Roger Gates

Mikayla M Rogers

Philip R Belanger

Brian R Pelletier

Billie Rosalie Maldonado

Christina G Ciamarra

Christopher William Moore

Peter J Milligan

James R Plante

Marshall F Thayer

James E Weisheit

Paul C McClay

Paul S Bean

Hannah M Pier

Dairus D McDowell

Thurman C Dabbs

Mary E M Hiller

Michael F Benamati

Nolan Frederic Quinn

Matthew James Bennison

Nicholas Anderson

Cole Romano

Kevin D Tracy

Jeffrey A Perry

Terence P Tyler, Jr

Jr Dominic J Mongillo

Dominic J Mongillo Jr

Terence P TYLERJr

Phillip R Rioux

David Griffin

Christopher Moore

Susan Tholen

A gas turbine engine includes a turbine section that has a ceramic matrix composite body that defines a core gaspath side that a row of blades, a forward end, an aft end, and circumferential sides. The forward end abuts at least one vane of a first row of vanes, and the aft end including a seal channel. A compliant seal is disposed in the seal channel and seals against at least one vane of a second row of vanes.

Blade outer air seal with compliant seal

A method of assembling a part is provided and includes forming a first section of the part, defining, in the first section, passages with dimensions as small as 0.005 inches (0.127 mm), forming a second section of the part, metallurgically bonding the first and second sections whereby the passages are delimited by the first and second sections and executing the metallurgically bonding without modifying a condition of the passages.

Hybrid bonded configuration for blade outer air seal (BOAS)

An additively manufactured vane cluster has a platform, a shroud, and airfoils joining the platform to the shroud. The platform has openings, each opening between a respective two of the airfoils and forming accommodating differential thermal expansion of a leading end of the platform relative to a trailing end of the platform.

Additively manufactured turbine vane cluster

A component for a gas turbine engine, including: at least one internal cavity extending through the component, the internal cavity having at least one inlet opening and at least one outlet opening each being in fluid communication with the at least one internal cavity; a plurality of cooling features extending from a surface of the at least one internal cavity, the plurality of cooling features are formed in accordance with at least one of the following groups: i) a plurality of airfoil shaped features that extend upwardly from the surface of the at least one internal cavity and a plurality of wedge shaped features each having a triangular base that has an upstream portion and a downstream portion, the upstream portion extending further from the surface than the downstream portion; ii) a plurality features having a curved or 'J' shaped base that extends upwardly from the surface, a plurality features having a double curved or symmetrically “J” shaped base that extends upwardly from the surface, and a plurality features having a base that extends upwardly from the surface with a curved portion that defines an opening therethrough; iii) a first plurality of pins with a plurality of grooves that are formed into a peripheral surface of each of the first plurality of pins and a second plurality of pins with a plurality of grooves that are formed into a peripheral surface of each of the second plurality of pins the plurality of grooves formed in the peripheral surface of each of the second plurality of pins are configured such that “V” shapes or inverted “V” shapes are formed in the peripheral surface of each of the second plurality of pins; and iv) a plurality of chevron shaped trip strips that are located in a channel, the plurality of chevron shaped trip strips are spaced from each other such that a “U” shaped passage is formed therebetween and each chevron shaped trip strip has a top portion that curls inwardly towards the channel and a plurality of pairs of features that each extend from a surface of another channel towards each other where a gap is located between distal ends of the plurality of pairs of features.

Cooling features for a component of a gas turbine engine

An inlet protector for a vane cooling air passage inlet including a vane including a vane cooling air inlet; an inlet protector fluidly coupled to the vane cooling air inlet, the inlet protector comprising a protector body having a discharge portion and a suction portion, an interior flow passage fluidly coupling the suction portion with the discharge portion; the discharge portion fluidly coupled with the vane cooling air inlet; and a suction inlet formed in the suction portion, wherein the suction inlet is located distally from the vane cooling air inlet.

Inlet protector for vane coupling hole

An accessible debris separator for a gas turbine engine is provided. The accessible debris separator including a chamber formed by a entrance wall, an exit wall, an inner wall, and an outer wall; a plurality of inlet openings in the entrance wall; a plurality of outlet openings in the exit wall; and a component within the chamber, the component forcing cooling air with debris particulates entering the chamber via an inlet opening of the plurality of inlet openings to take a circuitous path thereby separating the debris particulates from the cooling air prior to the cooling air exiting the chamber via an outlet opening of the plurality of outlet openings.

Accessible debris separator for high pressure turbine outside diameter fed static components

A component of a gas turbine engine includes a continuous hoop portion extending at least partially around an engine central longitudinal axis of the gas turbine engine and a bolting feature extending from the hoop portion, configured for securing the component to an adjacent component. The bolting feature includes a circumferential arm extending radially and circumferentially from the hoop portion. The circumferential arm defines a radial gap between the circumferential arm and the hoop portion. The circumferential arm includes a bolt opening for installation of a bolt therethrough. A connecting leg extends radially from the hoop portion across the radial gap toward the connecting arm. An arm interlocking feature is located at the circumferential arm and is configured to radially interlock with a complementary leg interlocking feature of the connecting leg. The arm interlocking feature and the leg interlocking feature have a radial interconnect gap therebetween in a free state.

Bolted joint of gas turbine engine

An assembly is provided for a turbine engine. This assembly includes a nozzle structure and a combustor wall. The nozzle structure includes a first platform, a second platform and a plurality of nozzle vanes arranged circumferentially about an axis. The nozzle vanes extends radially between and are connected to the first platform and the second platform. The combustor wall includes a plurality of apertures. An upstream portion of the combustor wall is radially between and borders a plenum and a combustion chamber. A downstream portion of the combustor wall is radially between and borders the plenum and a gap. The downstream portion of the combustor wall axially overlaps the nozzle structure with the gap formed by and extending between the combustor wall and the first platform. The apertures extends through the downstream portion of the combustor wall and are aligned with the nozzle vanes.

Cooling nozzle vanes of a turbine engine

An article comprises a first portion comprising a first alloy and a second portion comprising a second alloy that is metallurgically bonded to the first portion to form a monolithic article. The metallurgical bonding involves the application of an electrical current across the bond line and results in a retention of a metallurgical structure of the first portion and of a metallurgical structure of the second portion immediately adjacent to a bond line. The first portion has a first dominant property and the second portion has a second dominant property. The first dominant property is different from the second dominant property. The first dominant property is selected to handle operating conditions at a first position of the article where the first portion is located and the second dominant property is selected to handle operating conditions at a second position of the article where the second portion is located.

Hybrid superalloy article and method of manufacture thereof

An advanced passive clearance control (APCC) control ring is provided. The APCC control ring includes first and second cover sections, first and second wall sections and a control ring. At least one of the first and second cover sections is bonded to corresponding edges of the first and second wall sections by field assisted sintering technology (FAST) processing along a bond surface to form an enclosure for the control ring.

Advanced passive clearance control (APCC) control ring produced by field assisted sintering technology (FAST)

A blade outer air seal segment including a radially outward surface, a radially inward surface oriented away from the radially outward surface, and a cooling channel located between the radially outward surface and the radially inward surface. The blade outer air seal segment also including a stress-relief boss extending into the cooling channel and an inlet orifice fluidly coupled to the cooling channel through the stress-relief boss. The blade outer air seal segment further including a stress-relief recess. The stress-relief boss being located within the stress relief recess.

Fatigue resistant blade outer air seal

A novel method of depositing grit particles onto a fan blade tip coating is provided. The method enhances grit capture by presenting a softened coating surface for the impinging particles. The softened surface is achieved without high substrate temperatures that could degrade the base metal properties in the fan blade. An auxiliary heat source is used to establish a locally heated and softened surface where the grit deposition takes place. The softened surface greatly increases the probability of grit capture.

Method of forming an abrasive coating on a fan blade tip

There is a blade outer air seal positioned to be radially outward of blades in the rotating blade row relative to the rotational axis. There is at least one brush seal having bristles extending from a mount location radially inward to contact a sealing surface on the blade outer air seal. An angle is defined between the bristles and relative to the sealing surface having a radially inward extending component, and with an axial component in a direction toward a leading edge attachment feature and a trailing edge attachment feature, and the angle is less than 90° on a side of the brush seal opposite the leading edge and trailing edge mount hooks. A gas turbine engine component is also disclosed.

Angled brush seal and gas turbine engine component combination

A process for manufacturing a preformed sheet having geometric surface features for a geometrically segmented abradable ceramic thermal barrier coating on a turbine engine component, the process comprising the steps of providing a preformed sheet material. The process includes forming a partially of geometric surface features in the sheet material. The process includes joining the sheet material to a substrate of the turbine engine component. The process includes disposing a thermally insulating topcoat over the geometric surface features and forming segmented portions that are separated by faults extending through the thermally insulating topcoat from the geometric surface features.

Cost effective manufacturing method for GSAC incorporating a stamped preform

Growing community of inventors

Kennebunkport, ME, United States of America

Paul M Lutjen