United Technologies Corporation

Raytheon Technologies Corporation

Christopher King

San Quach

Steven Bruce Gautschi

Edward F Pietraszkiewicz

Scott D Lewis

Christopher Corcoran

David Donald Chapdelaine

Tracy A Propheter-Hinckley

Michael G McCaffrey

Thomas N Slavens

Matthew A Devore

Matthew S Gleiner

Ryan Alan Waite

David J Candelori

Karl A Mentz

Stephen D Doll

Aurelia Ledbetter

Steven Bruce Guatschi

Rohan Mehta

A gas turbine engine component includes a wall that has an inner surface and an outer surface. An inlet is defined by the inner surface. At least one non-rectangular slot is defined by the outer surface and includes at least one protrusion extending into the slot. A slot passage fluidly connects the inlet to the at least one non-rectangular slot. The slot passage comprises an inlet portion that extends through the wall from the inlet to an intermediate portion. An outlet portion extends through the wall from the intermediate portion to the at least one non-rectangular slot.

Cooling passages for gas turbine engine component

A process for manufacturing a composite rotor blade includes manufacturing an oversized root region of a root region of a composite rotor blade; fixturing the CMC blade into a machining fixture at a primary Y and Z axis datum located at an attachment fillet radii of the root region; machining V-notches into the oversized root region to form a Y' and Z&#x2032; axis datum of a sacrificial datum system in relation to the primary Y and Z axis datum; applying an oversized coating layer over the attachment fillet radii of the root region; fixturing the CMC blade into a machining fixture at the Y&#x2032; and Z&#x2032; axis datum of the sacrificial datum system; machining the oversized coating layer to a machined coating layer forming a Y&#x2033; and Z&#x2033; axis datum with respect to the Y&#x2032; and Z&#x2032; axis datum of the sacrificial datum system; fixturing the CMC blade into a machining fixture at the Y&#x2033; and Z&#x2033; axis datum; and machining off the sacrificial datum system removing the V-notches.

Ceramic matrix composite rotor blade attachment and method of manufacture therefor

A core for gas turbine engine component comprises a body extending between first and second ends to define a length, and extending between first and second edges to define a width. A plurality of core extensions are formed as part of the body. The plurality of core extensions are positioned to be staggered relative to each other such that at least two adjacent core extensions are variable relative to each other in at least one dimension. A gas turbine engine component is also disclosed.

Staggered core printout

A hanger for mounting an exhaust liner to an exhaust duct is provided. The hanger includes at least one cable, a cable flange, and a J-shaped flange. The at least one cable extends lengthwise between a first end and a second end. The cable flange has at least one cable aperture configured to receive a portion of the cable. The J-shaped liner flange is configured for attachment to the exhaust liner. The liner flange and cable flange are configured to mate with one another to prevent substantial relative motion between the cable flange and the liner flange in a first direction and in a second direction. The first and second directions are substantially orthogonal to one another.

Exhaust liner cable fastener

A component for a gas turbine engine includes a wall that has an exterior surface and an interior surface spaced from the exterior surface. A cooling circuit extends from the interior surface to the exterior surface. The cooling circuit includes a feed passage that has a first end defining an inlet at the interior surface and a second end spaced from the first end. At least one second passage is in fluid communication with the feed passage. At least one slot passage is in fluid communication with the second end of the feed passage and a cooling slot in the exterior surface.

Component for a gas turbine engine

An airfoil for a gas turbine engine includes an airfoil including spaced apart pressure and suction side walls joined at leading and trailing edges to provide an exterior airfoil surface that extends in a radial direction from a platform to a tip. A cavity is provided between the pressure and suction side walls near the trailing edge. The cavity includes an interior region bounded by first and second exit regions arranged at angle relative to one another. The first and second exit regions are respectively in low and high pressure regions relative to one another. First and second pedestal groups respectively are arranged at the first and second exit regions. The second pedestal group has first and second pedestals each terminating in an end. The ends of the second pedestals extend beyond the ends of first pedestals.

Gas turbine engine airfoil cooling configuration with pressure gradient separators

Core assemblies and methods for manufacturing components of gas turbine engines include a first core body having a first trunk configured to attach to a first location of a cavity core structure, a first branch extending from the first trunk and configured to form a first portion of a first cooling circuit, the first branch having a first joining surface, and a second core body having a second trunk configured to attach to a second location of a cavity core structure, a first branch of the second core body extending from the second trunk and configured to form a first portion of a second cooling circuit in the component. The first branches of the core bodies joined to form a junction. The junction defines a merger of the first cooling circuit and the second cooling circuit proximate to an exit of the first and second cooling circuits.

Core assemblies and gas turbine engine components formed therefrom

A cooling circuit for a gas turbine engine comprises a gas turbine engine component having a body with at least one internal cavity defined by a cavity wall. A plurality of cooling holes formed within the cavity wall, wherein each cooling hole is defined by a length extending from a cooling hole inlet to a cooling hole outlet, and wherein the cooling holes are positioned relative to each other such that a minimum allowable ligament distance is maintained between adjacent cooling holes along the entire length of each cooling hole. A gas turbine engine and a method of forming a cooling circuit for a gas turbine engine are also disclosed.

Gas turbine engine component with increased cooling capacity

A turbomachine component includes a body defining an interior cooling channel in fluid communication with the exterior of the body for fluid communication with a cooling flow source. At least one flow modifying pedestal is disposed within the interior cooling channel extending in a first direction. The flow modifying pedestal includes at least one flow feature that extends from the pedestal in a second direction.

Heat transfer pedestals with flow guide features

An airfoil according to an example of the present disclosure includes, among other things, an airfoil section having an external wall and an internal wall. The internal wall defines a reference plane extending in a spanwise direction along a surface of the internal wall, a first cavity and a second cavity separated by the internal wall, and a plurality of crossover passages within the internal wall and connecting the first cavity to the second cavity. Each of the plurality of crossover passages defines a passage axis. The passage axis of each of the plurality of crossover passages is arranged at a radial angle relative to a localized region of the reference plane such that the radial angle of at least some of the plurality of crossover passages differs in the spanwise direction.

Flared crossovers for airfoils

A component for a gas turbine engine is provided. The component includes an internal cooling passage disposed within the component, and an s-shaped trip strip formed on a surface of the internal cooling passage.

S-shaped trip strips in internally cooled components

An airfoil for use in a gas turbine engine is provided. The airfoil having: a pressure surface and a suction surface each extending axially from a leading edge to a trailing edge of the airfoil, at least one of the pressure surface, the suction surface, the leading edge and the trailing edge terminating at an edge of a tip section of the airfoil; a plurality of internal cooling channels located within the airfoil; and at least one cooling hole in fluid communication with at least one of the plurality of internal cooling channels, wherein the at least one cooling hole is aligned with an opening or diffuser that extends directly from the at least one cooling hole and wherein the opening or diffuser is formed in and extends through the edge of the tip section of the airfoil.

Turbine airfoil and method of cooling

A gas turbine engine airfoil includes a body that provides an exterior airfoil surface that extends in a radial direction to a tip. The exterior surface has a leading edge in a forward direction and a trailing edge in an aft direction. The tip includes a squealer pocket that has a recess surface. A cooling passage is arranged in the body. Each of the cooling holes extends from an inlet at the cooling passage to an outlet at the recessed surface. The inlet and outlet are arranged at an angle in an angular direction relative to the recessed surface. The angular direction is toward at least one of the forward and aft directions.

Gas turbine engine airfoil squealer pocket cooling hole configuration

The systems and methods described herein adapt the orientation, position and/or diffusion angles of showerhead cooling holes depending on the external gas/streamline flow direction. In regions, for example, where the stagnation line is on the pressure side, the breakouts of the showerhead holes substantially face aft (e.g., primarily towards the suction side). The location and positioning of the holes may be oriented according to the direction of incoming gas flows.

Growing community of inventors

Bristol, CT, United States of America

Christopher King