United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

Christopher Whitfield

Dominic J Mongillo, Jr

Carey J Clum

Shawn M McMahon

Nicholas M LoRicco

Steven D Porter

Dennis M Moura

Timothy J Jennings

Kevin Zacchera

Sean D Bradshaw

Noah Wadsworth

Thomas N Slavens

San Quach

Sasha M Moore

John McBrien

Justin M Aniello

Clifford J Musto

Steven Bruce Gautschi

Richard M Salzillo

Thomas Allwood

Vincent E Simms

Jeffery Scott Hembree, Sr

Tracy Phropheter-Hinckley

Tracy A Propheter-Hinckley

David Richard Griffin

Mark Borja

Alex J Schneider

Christopher Cosher

Brett Alan Bartling

Kristopher K Anderson

Christopher Perron

Jason B Moran

David Bitzko

Stephanie Tanner

Dominic Mongillo

Jr Dominic J Mongillo

Jr Dominic Mongillo

Nicholas M Loricco

Sr Jeffery Scott Hembree

Nicholas M Lo Ricco

Noah Wadswoth

Thomas N Stevens

A grommet may define a dilution hole in a combustor panel. The grommet may comprise a ridge having a stepped geometry formed about an inner diameter of the grommet, the ridge comprising a passage. The passage may comprise an outlet. The ridge may further comprise a fillet about the inner diameter of the grommet, wherein the outlet is configured to direct a cooling flow circumferentially along the fillet and fill the ridge with the cooling flow.

Dilution holes with ridge feature for gas turbine engines

A gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a first vane stage aft of the combustor. A plurality of vane cooling passages extending through a forward flange of the first vane stage to the radial inner surface for communicating cooling airflow into the core flow path forward of the leading edge. A seal assembly is disposed between the combustor and the forward flange of the vane stage. The seal includes a plurality of circumferentially spaced apart slots and a plurality of seal cooling passages that extend from the radially outer surface to the radially inner surface at each of the plurality of circumferentially spaced apart slots.

Conformal seal and vane bow wave cooling

A gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a first vane stage aft of the combustor. A seal assembly is disposed between the combustor and the first vane stage. The seal assembly includes a first plurality of openings and the first vane stage includes a second plurality of openings communicating cooling airflow into a gap between an aft end of the combustor and the first vane stage. A first vane stage assembly and a method are also disclosed.

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

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 core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a tip. A skin 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 core passage and the skin passage are configured to have a same direction of fluid flow. A resupply hole fluidly interconnects the core and skin passages. A centerline of the resupply hole is arranged at an acute angle relative to the direction of fluid flow in the core passage and is configured to provide a low turbulence flow region in the skin passage. The resupply hole has an exit at the skin passage and the exit has a diffuser.

Double wall turbine gas turbine engine blade cooling configuration

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 core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a tip. A skin passage is arranged in one of the pressure and suction side walls to form a hot side wall at an outer surface and a cold side wall at an inner surface. The skin passage extends a height in the radial direction, a width in a width direction, and a thickness in a thickness direction. The thickness is less than the width. The hot side wall defines a portion of the exterior airfoil surface and the cold side wall defines a portion of the core passage at a core passage surface. The core passage and the skin passage are configured to have a same direction of fluid flow. A resupply hole fluidly interconnects the core and skin passages. A centerline of the resupply hole is arranged at an first angle relative to the direction of fluid flow in the core passage. The first angle lies in a plane parallel to the thickness direction. The centerline of the resupply hole is arranged at a second angle relative to the direction of fluid flow in the core passage. The second angle lies in a plane parallel to the width direction. The second angle is at an acute angle. The first and second angles are configured to provide a low turbulence flow region in the skin passage.

A gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a first vane stage aft of the combustor. A seal assembly is disposed between the combustor and the first vane stage. The seal assembly includes a plurality of openings communicating cooling airflow into a gap between an aft end of the combustor and the first vane stage. A combustor assembly and method are also disclosed.

Conformal seal bow wave cooling

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 core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a tip. A skin 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 core passage and the skin passage are configured to have a same direction of fluid flow. A resupply hole fluidly interconnects the core and skin passages. The resupply hole has a single inlet that is fluidly connected to multiple discrete outlet apertures. A centerline of the single inlet is arranged at an angle relative to the direction of fluid flow in the core passage and is configured to provide a low turbulence flow region in the skin passage.

A vane includes a pair of airfoils having a plurality of film cooling holes that extend through an exterior surface of the airfoils, the plurality of film cooling holes including at least a first subset of film cooling holes, wherein the first subset of film cooling holes break through the exterior surface at geometric coordinates set forth herein. Each of the geometric coordinates is measured from a reference point on a leading edge rail of a platform of the vane.

Turbine vane cooling arrangement

A vane cluster for a gas turbine engine includes an outer diameter platform and an inner diameter platform. A plurality of vanes span from the outer diameter platform to the inner diameter platform. At least one inbound region is defined between a first vane of the plurality of vanes and a second vane of the plurality of vanes. The first vane includes a suction side facing the inbound region. Each of the vanes includes a leading edge core passage and a trailing edge core passage. A plurality of electrical discharge machined (EDM) holes are disposed within at least 0.500 inches (12.7 mm) of a leading edge of the first vane. Each of the EDM holes connect a leading edge core passage of the vane to an exterior surface of the vane.

Turbine vane cluster including enhanced vane cooling

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 core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a tip. A skin 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 core passage and the skin passage are configured to have a same direction of fluid flow. A slot extends longitudinally in the skin passage. A resupply hole fluidly interconnects the core passage and the slot.

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 core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a tip. A skin 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 core passage and the skin passage are configured to have a same direction of fluid flow. A resupply hole fluidly interconnects the core and skin passages. A centerline of the resupply hole is arranged at an acute angle relative to the direction of fluid flow in the core passage and is configured to provide a low turbulence flow region in the skin passage.

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Manchester, CT, United States of America

Christopher Whitfield