United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

Jeffrey J DeGray

Mark A Boeke

Richard M Salzillo

Dominic J Mongillo, Jr

Raymond C Surace

Timothy J Jennings

Emma J Place

Christopher J Peters

Jeremy B Fredette

Shawn J Gregg

Sarah Riley

Nicholas M LoRicco

Kevin Rajchel

Tracy A Propheter-Hinckley

Brandon W Spangler

Atul Kohli

Edwin Otero

Ryan Alan Waite

Ky H Vu

Gina Cavallo

Thomas P Dziuba

Allison Mainelli

Guillermo J Fernandez

Jr Richard M Salzillo

Richard M Salzillo Jr

A casting core assembly for a gas turbine engine component according to an implementation includes a skin core corresponding to a first cooling passage of an airfoil. The first cooling passage includes a first section and a tip flag section joined at a first bend. The skin core includes a first portion corresponding to the first section and a tip flag portion corresponding to the tip flag section. The skin core includes at least one arcuate slot corresponding to at least one turning vane of the airfoil. A method of forming an airfoil for a gas turbine engine is also disclosed.

Cooling arrangement for gas turbine engine component

An airfoil for a gas turbine engine according to an example of the present disclosure includes an airfoil section extending in a radial direction from a root section to a tip portion. The airfoil section has an external wall and an internal wall. The airfoil section establishes an internal cooling arrangement including a first cooling passage having a first section and a tip flag section. The tip flag section extends in the chordwise direction along the tip portion from the first section. The first section includes a plurality of branched paths established by at least one turning vane.

Component for gas turbine engines are described. The components include a component wall having an inner wall surface and an outer wall surface and a cooling hole formed within the component wall. The cooling hole has an inlet formed in the inner wall surface and an outlet formed in the outer wall surface and defines a fluid path through the component wall. The cooling hole has a metering section extending from the inlet to a transition point and a diffusing section extending from the transition point to the outlet. The metering section is defined by a uniform geometry and the diffusing section is defined by a lobed portion that extends from the transition point toward the outlet. The lobed portion has a first lobe and a second lobe divided by a ridge surface and each lobe defines continuous curved surfaces along the length of the ridge surface.

Airfoil cooling holes

An airfoil core includes a first core portion that has a hybrid skin core, a tip flag core, and a trailing edge core. A second core portion has a serpentine core and a leading edge core.

Multi-walled airfoil core

A method of forming an airfoil, includes forming a hybrid skin core, a tip flag core, and a trailing edge core. The hybrid skin core, tip flag core, and trailing edge core are connected to form a first core portion. A leading edge core and a serpentine core are formed. The first core portion, the leading edge core, and the serpentine core are assembled together to form an airfoil core. An airfoil is formed around the airfoil core.

A stator vane rail is provided. The stator vane rail may comprise a forward rail, an aft rail located axially opposite the forward rail, a first axial surface extending between the forward rail and the aft rail, a first feather seal slot disposed in the first axial surface, wherein a terminus of the first feather seal slot is radially spaced from a radial surface of the aft rail, a second axial surface extending between the forward rail and the aft rail, and a second feather seal slot disposed in the second axial surface, wherein the second feather seal slot extends from the radial surface of the aft rail.

Vane intersegment gap sealing arrangement

A cooled component for a gas turbine engine includes a plurality of internal ribs extending substantially parallel to a longitudinal axis of the gas turbine engine. The internal ribs are disposed within an internal cavity defining cooling air passages within the cooled component. A plurality of cooling holes are arranged in rows with axial orientations alternating between a radially outboard bias directing cooling air radially outward and a radially inboard bias directing cooling air radially inward. Each of the cooling holes includes an internal opening in communication with one of the cooling air passages and an external opening open to an outer surface of the cooled component. The external opening of each of the plurality of cooling holes is disposed on a side opposite the internal rib relative to a corresponding internal opening. A gas turbine engine and a method of fabricating a turbine airfoil are also disclosed.

Cooled turbine vane with alternately orientated film cooling hole rows

A flow path component includes a platform having at least one radially aligned face. A chordal seal extends axially from the radially aligned face. The chordal seal includes a first curved face configured to prevent edge line contact under deflection conditions while the flow path component is installed in an engine.

Flowpath component for a gas turbine engine including a chordal seal

A vane includes a pair of airfoils that have a plurality of film cooling holes that extend through an exterior surface of the airfoils. Each plurality of film cooling holes break through the exterior surface at geometric coordinates in accordance with Cartesian coordinate values of X, Y and Z as set forth in Table 1. Each geometric coordinates is measured from a reference point on a leading edge rail of a platform of the vane.

Turbine vane cooling arrangement

A section of a gas turbine engine according to an exemplary aspect of this disclosure includes, among other things, a vane assembly including a featherseal slot. The section further includes a featherseal at least partially received in the featherseal slot. The featherseal includes a radial portion having at least one tapered side.

Featherseal having tapered radial portion

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an body portion that extends between a leading edge and a trailing edge. At least one of the leading edge and the trailing edge includes at least one discharge slot having a first portion that includes an oval geometry.

Gas turbine engine component with discharge slot having oval geometry

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a mate face and a feather seal slot axially extending along the mate face, the feather seal slot having a variable width along a portion of its axial length.

Gas turbine engine component having variable width feather seal slot

An airfoil includes an airfoil wall including an exterior airfoil surface and at least partially defines an airfoil cavity. A fillet is on the exterior airfoil surface. A recess is in an interior surface of the airfoil wall adjacent the fillet. A baffle tube is located in the airfoil cavity spaced from the recess.

Gas turbine engine airfoil

A gas turbine engine component assembly comprises a first component and a second component circumferentially spaced from the first component relative to an engine center axis. A first baffle is associated with the first component. A second baffle is associated with the second component. Each of the first and second baffles includes at least one radial baffle tab. A gap is between the first and second baffles to define a cooling air inlet. A first coverplate is associated with the first baffle to cover a first portion of the gap. A second coverplate is associated with the second baffle to cover a second portion of the gap. The first and second coverplates are separate from each other, and include at least one coverplate radial tab that cooperates with an associated at least one baffle radial tab to prevent leakage gaps between the first and second baffle plates and the first and second coverplates.

Tolerance resistance coverplates

An airfoil of a turbine engine may have internal passages to permit the travel of cooling air through the airfoil. Multiple airfoils may be formed together in a stator vane assembly, and cooling air may be directed through the stator vane assembly. The stator vane assembly may be mounted to an engine structure to carry structural loads experienced by the stator vane assembly. Buttresses may be formed in the stator vane assembly, such as beneath each airfoil of the stator vane assembly, and extending into a passage permitting the travel of cooling air through the airfoil. The buttresses may enhance the strength and stability of the stator vane assembly by facilitating the transmission of structural loads to an engine structure. A channel may be defined through one or more buttress to enable the passage of cooling air into the airfoils, and yet allow the buttresses to carry structural loads.

Growing community of inventors

Somers, CT, United States of America

Jeffrey J DeGray