United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

Anthony B Swift

Paul M Lutjen

Ken Francis Blaney

Neil L Tatman

Dominic J Mongillo, Jr

Timothy M Davis

Mark John Rogers

Richard K Hayford

Carson A Roy Thill

Mark Broomer

Brian C McLaughlin

Craig R McGarrah

Russell E Keene

Christopher M Jarochym

Thomas E Clark

Matthew A Devore

Jose R Paulino

Thomas J Praisner

Daniel S Rogers

Terence P Tyler

Christina G Ciamarra

Timothy A Dunbar

Jr Dominic J Mongillo

Jr Terence P Tyler

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 blade outer air seal for a gas turbine engine includes a platform having a leading edge and a trailing edge. A pair of circumferential edges connect the leading edge and the trailing edge. An end wall protrudes radially outward from the platform at the trailing edge. A first support rib connects one of the circumferential edges to the end wall and structurally supports the end wall. A first boss portion extends axially forward from the end wall and is disposed radially outward of the first support rib.

Axial retention geometry for a turbine engine blade outer air seal

A gas turbine engine component has a body extending between two circumferential sides, and between a leading edge and a trailing edge. A refractory metal core within the body forms at least one cooling circuit to utilize fluid to cool the body. When the refractory metal core is removed from the body, the at least one cooling circuit includes an inlet, an outlet, and a passage that varies in cross-sectional area between the inlet and outlet. A method of manufacturing a gas turbine engine, a method of manufacturing a core, and a refractory metal core are also disclosed.

Variable thickness core for gas turbine engine component

A gas turbine engine component includes a wall portion that includes a first side and a second opposite side. A plurality of passages extends between the first side of the wall portion and the second side of the wall portion and includes a plurality of inlets located on the first side of the wall portion. A plurality of outlets are located on a second side of the wall portion. The plurality of outlets include a first plurality of outlets located on a first side of the plurality of inlets and a second plurality of outlets located on a second side of the plurality of inlets.

Cooling passages for gas turbine engine component

A seal assembly includes a blade outer air seal, a downstream vane, and a pressure wall, according to various embodiments. The blade outer air seal may include a radially outer surface and the downstream vane may be coupled to the blade outer air seal via a fluid sealing engagement. The pressure wall may be coupled to the blade outer air seal and may define a metering orifice. In various embodiments, the metering orifice of the pressure wall is configured to meter air flow from a first plenum upstream of the pressure wall to a second plenum downstream of the pressure wall. In various embodiments, at least a preponderance of the radially outer surface of the blade outer air seal at least partially defines the first plenum and the fluid sealing engagement at least partially defines the second plenum.

Air metering for blade outer air seals

A casting core for a Blade Outer Air Seal includes a heat exchange cavity core section in communication with a first plenum section and a second plenum section, the first plenum and the second plenum section are of a thickness greater than the heat exchange cavity section.

Reversible flow blade outer air seal

A cooling fluid system for a gas turbine engine includes a structure that provides a fluid passageway. The structure has a wall with an aperture that is in fluid communication with the fluid passageway. The aperture is configured to provide a fluid in a flow direction. Fingers are arranged in the fluid passageway facing into flow direction. The fluid passageway includes a cooling cavity immediately downstream from the fingers and it is configured to receive fluid having passed over or through the fingers.

Air-driven particle pulverizer for gas turbine engine cooling fluid system

A seal assembly for a gas turbine engine includes an engine static structure. First and second members fluidly separate cavities from one another. A seal assembly is captured by the engine structure. The seal assembly includes a carrier and a seal that engages the first member. The second member is captured by the carrier.

Gas turbine engine ring seal

A gas turbine engine component is provided. The gas turbine engine component comprises a main body having a leading edge and a leading edge wall including an elongated transition portion extending between the leading edge and a proximate flowpath surface of the main body. The elongated transition portion has an axial length that is greater than a radial height. A gas turbine engine is also provided.

Gas turbine engine components with optimized leading edge geometry

A gas turbine engine component is provided. The gas turbine engine component comprises a main body and a leading edge cooling passage defined within the main body. The main body has a leading edge and a leading edge wall including an elongated transition portion extending between the leading edge and a proximate flowpath surface of the main body. The leading edge cooling passage comprises an axial flow cooling passage defined within the main body and adjacent to the leading edge wall and has a leading edge periphery that generally conforms to the elongated transition portion of the leading edge wall.

Axial flowing cooling passages for gas turbine engine components

An interface within a gas turbine engine includes a multiple of segmented components, each with a segment flange with a multiple of apertures, at least one of the multiple of apertures a first slot aperture. A full ring component with a ring flange that defines a multiple ring of apertures, at least one of the multiple of ring apertures a second slot aperture, the second slot aperture transverse to the first slot aperture.

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North Waterboro, ME, United States of America

Anthony B Swift