United Technologies Corporation

Raytheon Technologies Corporation

Flodesign Sonics, Inc.

Nicholas M LoRicco

Thomas N Slavens

Sasha M Moore

Timothy J Jennings

Dominic J Mongillo, Jr

Clifford J Musto

Christopher Whitfield

John McBrien

San Quach

Jeffrey T Morton

Ricardo Trindade

Richard M Salzillo

Carey J Clum

Steven Bruce Gautschi

Alex J Schneider

James Tilsley Auxier

Brooks E Snyder

Bret M Teller

Shawn M McMahon

Jeffrey J DeGray

Mark A Boeke

David A Niezelski

Justin M Aniello

Parth Jariwala

Thomas Allwood

Vincent E Simms

Jeffery Scott Hembree, Sr

Bart Lipkens

Matthew S Gleiner

William J Brindley

Xuan Liu

Zhong Ouyang

Lane Mikal Thornton

Rafael A Perez

Andrew DeBiccari

Ryan Lundgreen

Mosheshe Camara-Khary Blake

Christopher J Bischof

Kathryn Macauley

Efrain E Vega

Dominic J Mongillo Jr

Sr Jeffery Scott Hembree

Jr Richard M Salzillo

A turbine blade has an attachment root and an airfoil. A cooling passageway system has a plurality of trunks extending from respective inlets along the root inner diameter end from a leading trunk near a first axial end to a trailing trunk near a second axial end; and a plurality of outlets along the airfoil including trailing edge outlets fed by the trailing trunk. Viewed normal to a root end-to-end centerplane: the trailing trunk has a turn passing forward and then rearward; an outside of the turn protrudes forward; and the outside of the turn has a tighter curvature than an inside of the turn.

Turbine blade trailing edge cooling feed

A component for a gas turbine engine includes a mateface with a purge flow interface, the mateface comprises a pocket located in communication with a feather seal slot in a mateface. A vane for a gas turbine engine includes a platform that extends from the airfoil, the platform comprising a mateface with a feather seal slot and a pocket in communication with the feather seal slot, wherein the pocket is of a cross-sectional shape larger than the feather seal slot.

Feather seal mateface cooling pockets

A turbine blade for a gas turbine engine. The turbine blade having a plurality of cooling holes defined therein, wherein the plurality of cooling holes are located in the turbine blade according to the coordinates of Table 1 and at least some of the plurality of cooling holes are located in an airfoil of the turbine blade.

Turbine blade cooling hole arrangement

An airfoil for a gas turbine engine according to an example of the present disclosure includes a first cavity, a second cavity, and a rib extending from a suction sidewall to a pressure sidewall and separating the first cavity from the second cavity. The rib includes a central portion, a first edge portion extending from the pressure sidewall to the central portion, and a second edge portion extending from the suction sidewall to the central portion. The rib defines one or more communication openings from the first cavity to the second cavity in the first or second edge portion.

Airfoil with rib communication

Vane assemblies for turbine engines are described. The vane assemblies include a vane having an internal cavity and a vane platform and a vane rail defining, in part, an outer diameter supply cavity. A blade outer air seal support (BOAS support) is arranged adjacent the vane and engages with a portion of the vane, the blade outer air seal support having BOAS support rail, and a BOAS supported on the BOAS support and engaging with a portion of the vane. The BOAS support includes a first cooling flow aperture configured to enable a cooling flow to cool at least the BOAS and a second cooling flow aperture formed in the BOAS support rail. The vane rail includes a third cooling flow aperture to form a cooling flow path through the second cooling flow aperture and the third cooling flow aperture to fluidly connect to the outer diameter supply cavity.

Turbine vane having dual source cooling

A method is provided that involves rotational equipment that includes a case and a first rotor blade within the case. During the method, a probe is arranged within the case adjacent the first rotor blade. A magnetic characteristic of the first rotor blade is measured using the probe. Presence of internal corrosion within the first rotor blade is determined based on the measured magnetic characteristic.

Determining presence of internal corrosion within a rotor blade by measuring magnetic characteristic(s)

A turbine blade for a gas turbine engine having a plurality of cooling holes defined therein, the plurality of cooling holes are located in an airfoil of the turbine blade according to the coordinates of Table 1.

A gas turbine engine component includes a wall that provides an exterior surface and an interior flow path surface. The wall has a wall thickness. A protrusion is arranged on the wall that extends a height beyond the wall thickness and provides a portion of the interior flow path surface. A film cooling hole that has an inlet is provided on the protrusion and extends to an exit on the exterior surface.

Gas turbine engine component with film cooling hole

Divided baffles for a gas turbine engines are provided. The divided baffles include a baffle body having a first end and a second end, at least one divider located within the baffle body and extending from the first end to the second end arranged to divide an interior of the baffle body into two or more feed cavities, at least one cap arranged at the first end to bound a first feed cavity of the two or more feed cavities, and at least one cap arranged at the second end to bound a second feed cavity of the two or more feed cavities.

Divided baffle for components of gas turbine engines

One exemplary embodiment of this disclosure relates to a method of forming an engine component. The method includes forming an engine component having an internal passageway, the internal passageway formed with an initial dimension. The method further includes establishing a flow of machining fluid within the internal passageway, the machining fluid changing the initial dimension.

Fluidic machining method and system

Core assemblies for manufacturing airfoils and airfoils for gas turbine engines are described. The core assemblies include a tip flag cavity core having an upstream portion, a tapering portion, and a downstream portion, with the tapering portion located between the upstream portion and the downstream portion and the downstream portion defines an exit in a formed airfoil. The upstream portion has a first radial height H, the downstream portion has a second radial height Hthat is less than the first radial height H, the tapering portion transitions from the first radial height Hat an upstream end to the second radial height Hat a downstream end, and at least one metering pedestal aperture is located within the tapering portion.

Airfoils having tapered tip flag cavity and cores for forming the same

An airfoil includes pressure and suction side walls that extend in a chord-wise direction between leading and trailing edges. The pressure and suction side walls extend in a radial direction to provide an exterior airfoil surface. A main-body core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a platform. A skin core cooling passage is arranged in one of the pressure and suction side walls to form a hot side wall and a cold side wall. The hot side wall defines a portion of the exterior airfoil surface and the cold side wall defines a portion of the core passage. The skin core cooling passage is divided by a wall into two discrete first and second skin core cooling passages each supplied with cooling fluid from opposing sides.

Double wall turbine gas turbine engine vane with discrete opposing skin core cooling configuration

A gas turbine engine component includes an airfoil that includes a trailing edge feed cavity. A baffle is located in the trailing edge feed cavity and includes a plurality of pressure side cooling openings in fluid communication with a corresponding one of a plurality of pressure side cooling passages. A plurality of suction side cooling openings are in fluid communication with a corresponding one of a plurality of suction side cooling passages.

Gas turbine engine airfoil with cooling path

A component for a gas turbine engine, includes a component outer surface exposed to flowpath gases of the gas turbine engine, a cooling channel located in the component, and at least one channel rib located in the cooling channel extending across the cooling channel from a channel inner surface to a channel outer surface. A cooling slot extends from the cooling channel to a slot outlet at the component outer surface. The slot outlet has a radial width greater than an axial length.

Cooling passage with structural rib and film cooling slot

Components for gas turbine engines are provided. The components include a hot external wall that is exposed to hot gaspath air when installed within a gas turbine engine, and an interior impingement wall, wherein the interior impingement wall defines a feed cavity and at least one impingement cavity is defined between the impingement wall and the external wall. The impingement wall includes a plurality of impingement holes that fluidly connect the feed cavity to the at least one impingement cavity, the external wall includes a plurality of film holes that fluidly connect the at least one impingement cavity to an exterior surface of the external wall, and wherein the only source of cooling air within the at least one impingement cavity is the feed cavity.

Dual-wall impingement cavity for components of gas turbine engines

Growing community of inventors

Windsor, CT, United States of America

Nicholas M LoRicco