United Technologies Corporation

Raytheon Technologies Corporation

James P Bangerter

Kevin Joseph Low

Donald William Peters

Alexander W Williams

Frank J Cunha

Timothy S Snyder

Steven D Porter

James B Hoke

Russell B Hanson

Robert M Sonntag

Nurhak Erbas-Sen

Reza Rezvani

Dennis John Duhamel

Nicholas Mule

Brian S Tryon

Russell Albert Beers

John S Tu

Robert Sze

Monica Pacheco-Tougas

Thomas R Russo

David Kwoka

Jonathan M Jause

David A Burns

Dennis J Duhamel

A liner panel within a gas turbine engine includes a perimeter rail that defines a first height from a cold side. The liner panel also includes an intermediate rail that defines a second height from the cold side, where the second height is less than the first height.

Gas turbine engine wall assembly with offset rail

A seal system is provided. The seal system may comprise a first duct having an annular geometry, a second duct overlapping the first duct in a radial direction, and a seal disposed between the first duct and the second duct. The seal may comprise a groove defined by the first duct and a piston configured to slideably engage the groove.

Composite piston ring seal for axially and circumferentially translating ducts

A wall assembly for a combustor of a gas turbine engine is provided. The wall assembly includes a support shell with a contoured region to define at least one convergent passage between the support shell and a multiple of liner panels. A method of cooling a wall assembly for a combustor of a gas turbine engine is also provided. This method includes a step of directing cooling air into at least one convergent passage between the support shell and a multiple of liner panels.

Gas turbine engine wall assembly with support shell contour regions

A heat shield for a combustor of a gas turbine engine includes an outer edge surface with an outlet of an edge cooling passage, the edge cooling passage oriented to direct cooling air generally upstream relative to a combustion gas flow.

Combustor panel T-junction cooling

In various embodiments, a cowl actuation system may comprise a cowl, an actuator, a bell crank, and a guide arm. The cowl may be configured to modulate at least a portion of the exhaust flow of a gas turbine engine. The bell crank may have a first end and a second end. The first end may be operatively coupled to the actuator. The second end may be operatively coupled to the cowl. The bell crank may be configured to pivot about a first point located between the first end and the second end. The first point may also be located along a diameter. The guide arm may have a third end and a fourth end. The third end may be mounted to a second point. The fourth end may be operatively coupled to the cowl. The second point may be located along the diameter.

System for actuating an exhaust cowl

A liner panel for use in a combustor for a gas turbine engine includes a forward region including a multiple of effusion passages each directed at an orientation consistent with a local swirl direction of combustion gases. Another liner panel for use in a combustor for a gas turbine engine includes a forward region forward of a dilution passage and an aft region aft of the forward region. The forward region includes a multiple of effusion passages each directed generally circumferentially. The aft region includes a multiple of effusion passages each directed at an orientation generally axially. A method of cooling a wall assembly within a combustor of a gas turbine engine includes orienting a multiple of effusion passages within a forward region consistent with a local swirl direction of combustion gases.

Co-swirl orientation of combustor effusion passages for gas turbine engine combustor

A heat shield for a combustor of a gas turbine engine includes a cold side opposite a hot side, the cold side defines a grommet that surrounds an igniter aperture, the grommet includes a multiple of fins.

Finned ignitor grommet for a gas turbine engine

A wall assembly is provided for a gas turbine engine. This wall assembly includes a support shell with a contoured region; and a multiple of liner panels mounted to the support shell. At least one of the multiple of liner panels includes an end rail. The contoured region is deformable to selectively contact at least a portion of the end rail. A method of assembling a wall assembly within a gas turbine engine is also provided. This method includes locating a stud that extends from a cold side of a liner panel through a support shell; and attaching a fastener onto the stud to at least partially close a gap defined between the panel and shell.

Gas turbine engine wall assembly interface

A liner for a combustor of a turbine engine includes a cooling feature which projects from a backside and one or more effusion holes that communicate(s) through the liner. The effusion hole(s) surround an opening through said liner. The cooling feature may include a trip strip or a pyramid pin fin (e.g., a three-side pyramid pin fin, a conical pyramid pin fin). The effusion hole(s) may penetrate said cooling feature. The effusion hole(s) may define an angle less than or equal to ninety (90) degrees with respect to a face of said liner. The effusion hole(s) may be proximate an opening through said liner.

Gas turbine engine combustor liner

A combustor includes a shell that bounds at least a portion of a combustion chamber. The shell includes a wall that has an orifice that opens to the combustion chamber. A grommet includes a body portion and a lip that projects from the body portion. The body portion and the lip carry a surface that extends around a passage through the grommet. At least a portion of the lip extends within the orifice of the wall of the shell.

Combustor with grommet having projecting lip

Aspects of the disclosure are directed to a system of an aircraft, comprising: at least one fairing, a liner, and an actuator configured to cause the at least one fairing to be translated relative to the liner in order to obtain a modulation of a metering area between the liner and the at least one fairing.

Stepped fairing modulated exhaust cooling

A stator vane for a gas turbine engine includes an airfoil that extends radially from a first side of a platform. A pocket is provided in the platform on a second side opposite the airfoil. Forward and aft rails extend from the second side of the platform. One of the forward and aft rails includes a radial surface. An impingement plate is secured to the platform over the pocket. The impingement plate includes a radial flange that engages the radial surface.

Gas turbine engine turbine vane platform cooling

A method for coating a turbine engine component comprises the steps of: providing a turbine engine component having at least one sacrificial attachment on a first side; grasping the turbine engine component via the at least one sacrificial attachment to position a first surface of the turbine engine component relative to a source of coating material; and applying a coating to said first side.

Method and system for holding a combustor panel during coating process

A combustor liner which reduces cooling flow to a combustion chamber and augments pressure drop split between impingement holes and effusion holes is disclosed. The combustor liner may further include accelerating channels, trip strips, pedestals, and cone-shaped effusion holes to provide further cooling of the liner. The combustor liner may reduce NOx production and the temperature of the combustion chamber of a gas turbine engine or the like.

Growing community of inventors

Manchester, CT, United States of America

James P Bangerter