Abstract: 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.

Abstract: A combustor panel an increased cooling holes provided at at least one of a pair of circumferential edges, a leading edge, a trailing edge or a hole circumference. The increase may be defined as a reduction in spacing or an increase in density. In another feature, holes at the circumferential edges may extend outwardly to an outlet in alignment with rails.

Abstract: 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,

Abstract: An airfoil includes a platform including a pocket, a continuous solid cover plate bonded over the pocket, and a plurality of apertures in the platform in fluid communication with the pocket and adjacent the continuous solid cover plate. A gas turbine engine including the airfoil and a method of cooling an article of a gas turbine engine are also disclosed.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: 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 wall that engages the radial surface.

Abstract: 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.

Abstract: A turbine engine assembly includes a combustor and a stator vane arrangement having a plurality of stator vanes. The combustor includes a combustor wall that extends axially from a combustor bulkhead to a distal combustor wall end, which is located adjacent to the stator vane arrangement. The combustor wall includes a support shell with a plurality of impingement apertures, and a heat shield with a plurality of effusion apertures. The combustor wall end includes a plurality of circumferentially extending film cooled regions. At least one of the film cooled regions is circumferentially aligned with one of the stator vanes and includes a cooling aperture.

Abstract: 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.

Abstract: 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.

Abstract: A liner for a combustor of a turbine engine includes a cooling feature which projects from a backside and an effusion hole that communicates through the liner

Abstract: A turbine engine assembly includes a combustor and a stator vane arrangement having a plurality of stator vanes. The combustor includes a combustor wall that extends axially from a combustor bulkhead to a distal combustor wall end, which is located adjacent to the stator vane arrangement. The combustor wall includes a support shell with a plurality of impingement apertures, and a heat shield with a plurality of effusion apertures. The combustor wall end includes a plurality of circumferentially extending film cooled regions. At least one of the film cooled regions is circumferentially aligned with one of the stator vanes and includes a cooling aperture.