A combustion chamber (15) comprising a wall at least partially defining a combustion zone and having a first surface (41) facing away from the combustion zone and a second surface (43) facing the combustion zone, the wall having at least one effusion cooling aperture (69, 73) extending there-through from the first surface to the second surface, the effusion cooling aperture having an inlet in the first surface and an outlet in the second surface, the first surface having a particle separator (84) at least partially located upstream of the inlet of the effusion cooling aperture, the particle separator projecting away from the first surface and away from the combustion zone.

A combustor for a gas turbine engine includes a first liner having a first surface, a second surface opposite the first surface, and a wall having a thickness defined between the first surface and the second surface. The first liner defines a plurality of effusion cooling holes, and at least one of the effusion cooling holes includes an inlet section spaced apart from the first surface, and a converging section downstream of the inlet section. The inlet section and the converging section are each defined so as to be outside of the thickness. The at least one of the effusion cooling holes includes a metering section downstream of the converging section and a portion of the metering section is defined within the thickness. The at least one of the effusion cooling holes includes an outlet section downstream of the metering section. The outlet section is proximate to the second surface.

A combustor for a gas turbine engine includes a combustion chamber defined between an inner shell and an outer shell. The combustor further includes a bulkhead extending between the inner shell and the outer shell. The combustor further includes a liner panel mounted to one of the inner shell and the outer shell aft of the bulkhead. The liner panel includes a first section including a first plurality of effusion holes. A first portion of the first plurality of effusion holes extends in a substantially circumferential direction. A second portion of the first plurality of effusion holes, disposed forward of the first portion, transitions from the substantially circumferential direction toward a substantially forward direction. A third portion of the first plurality of effusion holes, disposed aft of the first portion, transitions from the substantially circumferential direction to a substantially aft direction.

An apparatus, system, and method of reducing film cooling scrubbing in a combustor are provided. The apparatus includes a combustor liner that includes a field of a plurality of sparsely-spaced film holes that extend through the combustor liner wherein the field includes a boundary. The combustor liner also includes a patch of relatively more densely-spaced shield holes that extend through the combustor liner at specific predetermined locations within the boundary of the field.

A gas turbine engine component assembly including: a first component having a first surface and a second surface opposite the first surface; and a second component having a first surface, a second surface opposite the first surface of the second component, and an impingement slot extending from the second surface of the second component to the first surface of the second component, the second surface of the first component and the first surface of the second component defining a cooling channel therebetween in fluid communication with the impingement slot, wherein the impingement slot is in fluid communication with the second surface of the first component, wherein the impingement slot includes a slot tab configured to direct airflow into the cooling channel at least partially in a lateral direction parallel to the second surface of the first component such that a cross flow is generated in the cooling channel.

A combustor for use in a gas turbine engine, the combustor enclosing a combustion chamber having a combustion area. The combustor includes: a shell having a kink; and a kinked heat shield panel in facing spaced relationship with the shell, the kinked heat shield panel including a kink located proximate the kink in the shell, wherein the kinked heat shield panel further includes a first surface, a second surface opposite the first surface, and a mounting stud located proximate the kink of the kinked heat shield panel and extending away from the second surface.

Combustor panels of gas turbine engines and gas turbine engines are described. The combustor panels include a hot side configured to be exposed to combustion within a gas turbine engine, a cold side opposite the hot side of the combustor panel, the cold side configured to receive cooling flow thereon, and a peak-valley gridded pattern formed on the cold side, the peak-valley gridded pattern comprising a plurality of recessed cells arranged in a grid pattern, with each recessed cell having a peak, angled sidewalls, and an effusion hole located at a bottom of the angled sidewalls.

A liner assembly for use in a combustor of a gas turbine engine is disclosed. In various embodiments, the liner assembly includes a panel defining a left side and a right side and a hot side and a cold side, the panel having a dilution hole and a plurality of effusion holes extending between the hot side and the cold side. In various embodiments, the plurality of effusion holes includes a first subgrouping of effusion holes disposed downstream of the dilution hole and aligned in a generally left to right orientation toward a dividing line extending downstream of the dilution hole and a second subgrouping of effusion holes disposed downstream of the dilution hole and aligned in a generally right to left orientation toward the dividing line extending downstream of the dilution hole.

A combustor arrangement, the combustor arrangement being annular and being arranged about an axis, the axis defining an axial direction, having an annular housing to house a plurality of burners and an annular combustion chamber, the burners arranged circumferentially about the axis inside the annular housing, wherein an annular space is defined between the housing, the burners and the annular combustion chamber, the annular space arranged to guide a compressed fluid. A plurality of stiffening plates, each arranged within the annular housing, wherein two adjacent ones of the burners are separated by one of the stiffening plates. A combustor separating wall arrangement separates the annular space from the annular combustion chamber provides openings for the burners. The stiffening plates are arranged angled and connected to the combustor separating wall arrangement and two boundary walls of the housing, and further plates extend into the annular space.

A combustor for a turbine engine includes a first liner and a second liner forming a combustion chamber with the first liner. The combustion chamber is configured to receive an air-fuel mixture for combustion therein. The first liner is a first double wall liner with a first wall forming a portion of the combustion chamber and a second wall extending around at least a portion of the first wall to form a liner cavity with the first wall. The first wall defines a first wall orifice and the second wall defines a second wall orifice. The combustor further includes a first insert mounted on the second wall within the second wall orifice and extending through the first wall orifice. The first insert is configured to direct a first air jet through the second wall, through the first wall, and into the combustion chamber.