United Technologies Corporation

Raytheon Technologies Corporation

The Penn State Research Foundation

Raytheon Corporation

Aytheon Technologies Corporation

Rtx Corporation

Scott D Lewis

Mark F Zelesky

Dominic J Mongillo, Jr

Bret M Teller

Christopher Corcoran

Kyle C Lana

Steven Bruce Gautschi

Matthew S Gleiner

Ricardo Trindade

Jeffrey S Beattie

Jeffrey Michael Jacques

Brandon M Rapp

Stephen Lynch

Shane Haydt

San Quach

Matthew A Devore

Atul Kohli

Christopher King

David Donald Chapdelaine

Brandon W Spangler

JinQuan Xu

James B Hoke

Jeffrey T Morton

Thomas J Praisner

Russell Deibel

Matthew Andrew Hough

Lane Mikal Thornton

Jeffrey Jay Lienau

Patrick D Couture

David J Candelori

John W Magowan

Gary A Zess

Yafet Girma

Karen A Thole

Jason David Liles

Max Asterlin

Christoper King

Metthew A Devore

Gaurav M Patel

Jeffrey Scott Hembree

Robert Schroeder

A gas path component for a gas turbine engine includes a film cooling hole disposed in the gas path component. The film cooling hole includes a metering section, a diffuser, and a tapered surface extending between the metering section and the diffuser. The tapered surface is oriented between twenty degrees and seventy degrees with respect to a centerline axis of the metering section. The tapered surface is oriented at an obtuse angle with respect to an immediately adjacent surface of the diffuser, the obtuse angle is open towards the centerline axis. The tapered surface is configured to mitigate flow separation in the diffuser.

Systems and methods for manufacturing film cooling hole diffuser portion

A gas turbine engine component according to an example of the present disclosure includes, among other things, an external wall including adjacent bounding pedestals that extend from an external wall surface to establish a cooling passage, and including a common pedestal situated between the adjacent bounding pedestals to establish a first branched section and a second branched section of the cooling passage that join together at a merged section of the cooling passage. A method of fabricating a gas turbine engine component is also disclosed.

Cooling arrangement including alternating pedestals for gas turbine engine components

A cooling hole arrangement for a combustor panel including a first group of cooling holes, each of the first cooling holes extending from a first inlet to a first outlet. Also included is a second group of cooling holes, each of the second cooling holes extending from a second inlet to a second outlet, each of the first cooling holes and the second cooling holes spaced from an edge of the combustor panel in an axial direction to be arranged substantially parallel to the edge of the combustor panel. At least a portion of the first inlet of each first cooling holes is axially overlapped with a portion of the second inlet of the second cooling holes. The first outlet of each of the first cooling holes is closer to the edge of the combustor panel than the second outlet of the second cooling holes is to the edge.

Combustor panel cooling hole arrangement

An airfoil may include an airfoil body, a root and a platform disposed between the airfoil body and the root. The platform may have a first mating surface and a second mating surface. The platform may include a pocket defined by an inner diameter surface of the platform proximate the first mating surface. A channel may be defined in the platform with an outlet of the channel defined in the second mating surface.

Airfoil platform cooling channels

An airfoil for a gas turbine engine includes pressure and suction side walls joined to one another at leading and trailing edges to provide an exterior airfoil surface. The pressure and suction side walls are spaced apart from one another in a thickness direction. A stagnation line is located near the leading edge. A cooling passage is provided between the pressure and suction side walls. The showerhead cooling holes are arranged at least one of adjacent to or on the stagnation line. At least one of the showerhead cooling holes has a metering hole fluidly connecting the cooling passage to a diffuser arranged at the exterior airfoil surface. At least one showerhead cooling hole is arranged on each of opposing sides of the stagnation line. Each showerhead cooling hole has the diffuser with a first diffuser angle that expands downstream in the thickness direction in opposing directions from one another when separated by the stagnation line.

Gas turbine engine airfoil with showerhead cooling holes near leading edge

An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, at least one of an airfoil section and a platform section including a wall. The wall includes a plurality of pedestals having adjacent pedestals extending from an external wall surface to establish a respective cooling passage, and the cooling passage includes an inlet and an outlet. The adjacent pedestals are dimensioned such that the adjacent pedestals taper inwardly from the inlet in a first direction towards the outlet to establish a throat in the respective cooling passage. A method of fabricating a gas turbine engine component is also disclosed.

Cooling arrangement for gas turbine engine components

A seal for a gas turbine engine includes a plurality of seal arc segments. Each of the seal arc segments includes radially inner and outer sides and sloped first and second circumferential sides. The seal arc segments are circumferentially arranged about an axis such that the sloped first and second circumferential sides define gaps circumferentially between adjacent ones of the seal arc segments. Each of the gaps extends from the radially inner sides along a respective central gap axis that slopes with respect to a radial direction from the axis.

Seal arc segment with sloped circumferential sides

A component for a turbine engine includes a fore edge connected to an aft edge via a first surface and a second surface. Multiple cooling passages are defined within the turbine engine component. A skin core passage is defined immediately adjacent the first surface, and at least one pedestal interrupts a flow path through the skin core passage.

Turbine engine component including an axially aligned skin core passage interrupted by a pedestal

A combustor panel may comprise a dilution hole, a plurality of film cooling holes each of the plurality comprising an inlet and an outlet, a first group of the plurality of film cooling holes arranged circumferentially about the dilution hole, wherein each of film cooling holes of the first group have a first outlet angle oriented radially toward the center of the dilution hole, and a second group of the plurality of film cooling holes arranged radially outward of the first group and relatively circumferentially between the first group.

Dilution/effusion hole pattern for thick combustor panels

A vane according to an exemplary aspect of the present disclosure includes, among other things, a platform extending from an edge face and between spaced apart lateral faces and an airfoil extending outwardly from the platform. The platform includes at least one ejection port in the edge face and at least one passage connected to the at least one ejection port. A method of cooling a component is also disclosed.

BOAS thermal protection

A component for a gas turbine engine, the component having: a cooling slot located on a surface of the component, the cooling slot being defined by a plurality of diffuser portions each extending from a respective one of a plurality of cooling openings providing cooling fluid to the cooling slot.

Turbine blade with slot film cooling

A gas turbine engine blade includes a blade portion having a leading edge and a trailing edge. A first surface connects the leading edge to the trailing edge and a second surface connects the leading edge to the trailing edge. A tip section is located at a first end of the blade portion and includes a pocket protruding into the tip section from an outermost end of the tip section. The pocket has a first side wall adjacent the first surface and a second side wall adjacent the second surface. At least one of the first side wall and the second side wall have a curve distinct from a curve of the corresponding adjacent surface.

Gas turbine engine rotor including squealer tip pocket

A blade for a gas turbine engine includes an airfoil which includes pressure and suction side surfaces joined at leading and trailing edges to provide an exterior airfoil surface. The airfoil extends from a 0% span position to a 100% span position at a tip. A mean chamber line is provided at the tip and extends from the leading edge to the trailing edge. The mean chamber line has a leading edge tangent line at the leading edge at the tip and a trailing edge tangent line at the trailing edge at the tip. The leading edge tangent line and the trailing edge tangent line intersect to provide a total camber angle in a range of &#x2212;35&#xb0; to 35&#xb0;. The pressure side includes a tip shelf.

Turbine blade with tip vortex control and tip shelf

An electrode for electrical discharge machining (EDM) may comprise a diffuser portion and a tapered portion defining the tip of the electrode. A method for forming a film cooling hole may comprise moving a tool with respect to a film cooled gaspath component, forming a diffuser of the film cooling hole in response to the moving, and forming a tapered surface between a metering section and the diffuser of the film cooling hole.

A rotor blade for a gas turbine engine is provided. The rotor blade having: an attachment; an airfoil extending from the attachment to a tip; and a tip shelf located in a surface of the tip proximate to a pressure side of the airfoil, wherein the tip shelf has a ledge portion extending from the pressure side to a wall portion extending upwardly from the ledge portion to the tip and wherein the wall portion is configured to have a convex portion with respect to the pressure side of the airfoil as it extends from a leading edge to a trailing edge of the airfoil.

Turbine blade with airfoil tip vortex control

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Vernon, CT, United States of America

Scott D Lewis