Abstract: A combustor for a gas turbine engine includes a liner enclosing a combustion chamber and defining air passages through the liner, a fuel nozzle fluidly connected to the combustion chamber, and a louver disposed inside the combustion chamber over the air passages. The louver extends circumferentially along the liner and is connected to the liner by a fastener. The fastener spacing at least one of axial edges of the louver from the liner to define an air outlet between the at least one of the axial edges and the liner. A method of manufacturing a combustor of an aircraft engine is also disclosed.

Abstract: An interface for double-skin combustor liner comprises an annular ceramic matrix composite (CMC) body. The annular body extends from a first side to a second side, the first and second sides each having an annular inner joint groove configured to receive a CMC combustor skin and an annular outer joint groove configured to receive a metal combustor skin. The outer joint groove is radially outward of the inner joint groove. A combustor and an gas turbine engine with the interface are also provided.

Abstract: A gas turbine engine combustor has a dome and a shell extending axially from the dome. The dome and the shell cooperates to define a combustion chamber. A dome heat shield is mounted to the dome inside the combustion chamber. A front heat shield is mounted to the shell inside the combustion chamber. The dome heat shield and the front heat shield have axially overlapping portions cooperating to define a flow guiding channel. The flow guiding channel has a length (L) and a height (h). The length (L) is at least equal to the height (h).

Abstract: The described reverse flow combustor of a gas turbine engine includes inner and outer combustor liners defining a combustor chamber therewithin. A large exit duct and a small exit duct are disposed at downstream ends of the outer and inner liner respectively. The small exit duct includes an annular ring removably mounted to a support element of the gas turbine engine by one or more fastening elements that are integrally formed with the annular ring. The fastening elements are mountable with corresponding features of the support element and removably fastened thereto.

Abstract: There is disclosed a double skin combustor for a gas turbine engine. The combustor has a radially inner liner and a radially outer liner defining therebetween an annular combustion chamber. The radially inner liner and the radially outer liners both have a hot skin and a cold skin defining a cooling cavity therebetween. The cold skin has a plurality of segments joined to one another by sliding joints. The cooling cavity between the hot skin and the cold skin may be compartmentalized into individual cavities between the sliding joints.

Abstract: A combustor for a gas turbine engine includes a liner enclosing a combustion chamber and defining air passages through the liner, a fuel nozzle fluidly connected to the combustion chamber, and a louver disposed inside the combustion chamber over the air passages. The louver extends circumferentially along the liner and is connected to the liner by a fastener. The fastener spacing at least one of axial edges of the louver from the liner to define an air outlet between the at least one of the axial edges and the liner. A method of manufacturing a combustor of an aircraft engine is also disclosed.

Abstract: The build-up of carbon deposition on the front face of a combustor heat shield is discouraged by jetting air out from the front face of the heat shield with sufficient momentum to push approaching fuel droplets or rich fuel-air mixture way from the heat shield.

Abstract: A sliding joint in a gas turbine engine between a large exit duct of a combustor and a turbine vane assembly having a leading edge lug. The sliding joint has an elongated flexible arm extending between a first end joined to the distal flange and an opposed free second end disposed radially inward of the distal flange. The flexible arm has a first surface and a second surface spaced radially inward from the first surface. A spacer on the first surface of the flexible arm projects radially away toward the distal flange. The spacer defines a gap and axially displaces with respect to the lug upon thermal expansion of the large exit duct. The leading edge lug of the turbine vane assembly is disposed in the gap, and the second end of the flexible arm is disposed radially inward of the leading edge lug.

Abstract: An interface for double-skin combustor liner comprises an annular ceramic matrix composite (CMC) body. The annular body extends from a first side to a second side, the first and second sides each having an annular inner joint groove configured to receive a CMC combustor skin and an annular outer joint groove configured to receive a metal combustor skin. The outer joint groove is radially outward of the inner joint groove. A combustor and an gas turbine engine with the interface are also provided.

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 sliding joint in a gas turbine engine between a large exit duct of a combustor and a turbine vane assembly having a leading edge lug. The sliding joint has an elongated flexible arm extending between a first end joined to the outer surface of the large entry duct, and an opposed free second end disposed radially inward of the outer surface of the large entry duct. A spacer is joined to the second end of the arm and projects radially away therefrom toward the outer surface of the large entry duct. The spacer is spaced apart from the outer surface and defines a gap therebetween. The spacer, the arm, and the sliding joint axially displace with respect to the lug upon thermal expansion of the large entry duct.

Abstract: The described reverse flow combustor of a gas turbine engine includes inner and outer combustor liners defining a combustor chamber therewithin. A large exit duct and a small exit duct are disposed at downstream ends of the outer and inner liner respectively. The small exit duct includes an annular ring removably mounted to a support element of the gas turbine engine and includes a plurality of cooling elements integrally formed with the annular ring and projecting therefrom into impingement airflow. The cooling elements increase the effective surface area of the inner surface of the annular ring, which is adapted to be cooled by the impingement airflow.

Abstract: A gas turbine engine combustor has a dome and a shell extending axially from the dome. The dome and the shell cooperates to define a combustion chamber. A dome heat shield is mounted to the dome inside the combustion chamber. A front heat shield is mounted to the shell inside the combustion chamber. The dome heat shield and the front heat shield have axially overlapping portions cooperating to define a flow guiding channel. The flow guiding channel has a length (L) and a height (h). The length (L) is at least equal to the height (h).

Abstract: A method of minimizing damage to a tip of an igniter in a combustor of a gas turbine engine includes mounting the igniter to a liner of the combustor downstream of the fuel nozzles using a floating collar, and directing pressurized air from a plenum surrounding the combustor through a plurality of purge openings and into an annular cavity surrounding the tip of the igniter. The purge openings are located in one or both of the igniter and the floating collar, and create cooling purge air jets into the annular cavity. The purge openings are formed within an arc that is disposed upstream of a bisecting plane extending along an igniter axis to a chamber axis of the combustor, such that the igniter axis lies in the bisecting plane. The igniter is free from the purge openings downstream of the bisecting plane relative to the stream of hot gases.

Abstract: There is disclosed a double skin combustor for a gas turbine engine. The combustor has a radially inner liner and a radially outer liner defining therebetween an annular combustion chamber. The radially inner liner and the radially outer liners both have a hot skin and a cold skin defining a cooling cavity therebetween. The cold skin has a plurality of segments joined to one another by sliding joints. The cooling cavity between the hot skin and the cold skin may be compartmentalized into individual cavities between the sliding joints.

Abstract: A gas turbine engine combustor has a dome and a shell extending axially from the dome. The dome and the shell cooperates to define a combustion chamber. A dome heat shield is mounted to the dome inside the combustion chamber. A front heat shield is mounted to the shell inside the combustion chamber. The dome heat shield and the front heat shield have axially overlapping portions cooperating to define a flow guiding channel. The flow guiding channel has a length (L) and a height (h). The length (L) is at least equal to the height (h).

Abstract: A gas turbine engine comprising a combustor having a combustor liner assembly and a mounting bracket provided on the combustor liner assembly, a floating collar being slidingly received on the mounting bracket for relative sliding movement in a plane normal to an axis of an igniter opening in the liner assembly. The floating collar includes an annular surface defining a collar opening, and an igniter having an axis concentric with the axis of the collar opening is sealingly engages the annular surface. A plurality of purge openings defined in at least one of the igniter and the floating collar form cooling airflow passages communicating from the plenum to the cavity.

Abstract: A coated component of a gas turbine engine having a hot side and a cold side. The component includes a metallic substrate forming a base structure of the component. The metallic substrate has a first surface on the hot side and a second surface on the cold side. A thermal coating system on the metallic substrate includes a first aluminide layer in direct contact with the second surface on the cold side of the component and a first bond layer overlying the first aluminide layer on the cold side of the component.

Abstract: A combustor heat shield has lips with fins distributed on the lips. The lip-fins have an extended end portion projecting rearwardly from the back face of the heat shield. Impingement jets may be directed against the rearwardly extended end portions of the lip-fins to enhance cooling. The heat shield may define a fuel nozzle opening surrounded by a rail on the back side of the heat shield. Impingement holes or slots may be defined in the rail for allowing cooling air passing therethrough to impinge upon the lip-fins.

Abstract: The described reverse flow combustor of a gas turbine engine includes inner and outer combustor liners defining a combustor chamber therewithin. A large exit duct and a small exit duct are disposed at downstream ends of the outer and inner liner respectively. The small exit duct includes an annular ring removably mounted to a support element of the gas turbine engine and includes a plurality of cooling elements integrally formed with the annular ring and projecting therefrom into impingement airflow. The cooling elements increase the effective surface area of the inner surface of the annular ring, which is adapted to be cooled by the impingement airflow.