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.

An assembly is provided for a turbine engine. This turbine engine assembly includes a combustor wall. The combustor wall includes a shell, a heat shield and an annular body. The annular body extends through the combustor wall and at least partially defines a quench aperture along a centerline through the combustor wall. The shell defines a first cooling aperture radially outwards of the annular body relative to the centerline and is configured to direct air to impinge against a portion of the annular body between the heat shield and the shell.

A main mixer including a swirler along an axis, the swirler including an outer swirler with a multiple of outer vanes, and a center swirler with a multiple of center vanes and a swirler hub along the axis, the swirler hub including a fuel manifold and an inner swirler with a multiple of inner vanes that support a centerbody, the multiple of inner vanes interconnect the fuel manifold and the centerbody.

An assembly is provided for a gas turbine engine. This assembly includes a multi-walled structure including a cold wall, a hot wall and a cooling cavity vertically between the cold wall and the hot wall. The cold wall includes a plurality of cold wall apertures fluidly coupled with the cooling cavity. The cold wall apertures are configured to subject the cold wall to a cold wall pressure drop vertically across the cold wall. The hot wall includes a plurality of hot wall apertures fluid coupled with the cooling cavity. The hot wall apertures are configured to subject the hot wall to a hot wall pressure drop vertically across the hot wall that is greater than or equal to the cold wall pressure drop.

An assembly is provided for a gas turbine engine. This assembly includes a combustor wall, a case and an inlet. The combustor wall is configured with a quench aperture. The case is displaced from the combustor wall such that a plenum is formed by and extends between the combustor wall and the case. The case includes a case wall, a deflector and a conduit. The deflector projects out from the case wall into the plenum towards the combustor wall. The deflector is arranged upstream of the quench aperture. The inlet to the conduit is arranged next to and downstream of the deflector.

The present invention relates to a method for producing an engine component having a cooling duct arrangement which has a plurality of cooling ducts, each having an inflow opening, the inflow openings being arranged according to a predefined pattern in an inflow surface of the engine component, and each cooling duct opening into a recess in a wall of the engine component, along which wall a cooling film is to be formed. According to the invention, the pattern is formed in at least one subregion of defined size of the inflow surface, from a plurality of identical isosceles triangles, which are defined by a minimum spacing (k) and by a mean diameter (a) of the inflow openings correlating to the minimum spacing (k). This procedure reduces the complexity of the design process.

Adhesion of particulate matters to the burner accompanying combustion in a lean-combustion gas turbine combustor is suppressed, and the structural reliability is improved. In a gas turbine combustor including: a tubular liner that forms a combustion chamber; and a burner including an air hole plate that is arranged at an inlet of the liner and includes a plurality of air holes for guiding compressed air to the combustion chamber, and a plurality of fuel nozzles that are arranged on a side opposite to the combustion chamber with the air hole plate being sandwiched therebetween, the plurality of fuel nozzles each injecting a fuel toward a corresponding air hole, the air holes and the fuel nozzles forming a plurality of concentric annular lines, a plurality of small holes having opening diameters smaller than those of the air holes are provided through the air hole plate such that the plurality of small holes are positioned in an inner area of an innermost annular line of the air holes.

A dome-deflector structure for a combustor of a gas turbine includes a dome portion and a deflector portion, the dome portion and the deflector portion being connected together to form a dome-deflector cavity therebetween. The dome portion includes a plurality of dome-side cooling airflow passages therethrough arranged in a plurality of groups of dome-side cooling airflow passages, each group of dome-side cooling airflow passages being arranged to provide a flow of air therethrough in a respective group swirl direction, and adjacent groups of the dome-side cooling airflow passages providing the flow of air in a different group swirl direction with respect to one another.

A swirler assembly includes a swirler having a primary swirler with a primary swirler venturi, a swirler ferrule plate connected upstream to the primary swirler, and a fuel nozzle disposed in the swirler ferrule plate. The swirler ferrule plate has an annular pressure drop cavity with oxidizer inlet orifices in fluid communication with the swirler, and at least one outlet orifice in fluid communication with the primary swirler venturi. A second flow of oxidizer to the swirler incurs a first pressure drop, a third flow of the oxidizer from the swirler to the annular pressure drop cavity incurs a second pressure drop, and a fourth flow of the oxidizer from the annular pressure drop cavity to the primary swirler venturi incurs a third pressure drop.

A liner for a combustor includes a first wall 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 liner 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 and the first wall. The first insert is formed by a tubular body portion extending through the first wall orifice and the second wall orifice, and a shoulder extending around the body portion that abuts the first side of the second wall. The shoulder has a first diameter and the first wall orifice has a second diameter, greater than the first diameter.