A gas turbine includes a rotor that is rotatable about an axis, a casing configured to cover the rotor in a circumferential direction and having an annular space therein, a compressor configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing, a plurality of combustors disposed in the casing at equal intervals in the circumferential direction of the rotor and configured to combust the compressed air and fuel taken in from the casing to generate a combustion gas, a turbine driven by the combustion gas, and an air introduction passage defined by a partition plate configured to divide the space in the casing in the circumferential direction of the rotor and an inner circumferential surface of the casing and configured to introduce the compressed air in the casing into the combustor.

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.

An embodiment of a torch igniter for a combustor of a gas turbine engine comprises a combustion chamber oriented about an axis, a cap defining an axially upstream end of the combustion chamber and oriented about the axis, a tip defining an axially downstream end of the combustion chamber, a structural wall coaxial with and surrounding the igniter wall, an outlet passage defined by the igniter wall within the tip, and a cooling system. The cooling system comprises an air inlet formed within the structural wall, a first flow path disposed between the structural wall and the igniter wall, and an aperture extending through the igniter wall transverse to the flow direction. The aperture directly fluidly connects the first flow path to the combustion chamber.

Combustors and gas turbines are provided. A combustor includes an axial centerline and an end cover. The combustor further includes at least one fuel nozzle that extends from the end cover and at is least partially surrounded by a combustion liner. The combustion liner extends between the at least one fuel nozzle and an aft frame and that defines a combustion chamber. An outer sleeve is spaced apart from and surrounds the combustion liner such that an annulus is defined therebetween. The outer sleeve defines at least one aperture. A wake energizer is mounted on the outer sleeve. The wake energizer defines at least one passage that is angled with respect to the axial centerline of the combustor. The at least one passage aligns and is in fluid communication with the at least one aperture of the outer sleeve.

Methods and apparatus for an additive, single-piece, bore-cooled combustor dome or liner are disclosed. An example combustor dome forms an integral part including: a plurality of first openings; a plurality of second openings; and a plurality of passages formed in the combustor dome connecting respective ones of the plurality of first openings with respective ones of the plurality of second openings. The combustor dome is configured to allow air to enter through the plurality of first openings and travel through the plurality of passages to exit through the plurality of second openings, the air to transfer heat from the combustion section.

An impingement cooling structure is provided. The impingement cooling structure includes a flow channel formed between a first wall and a second wall facing the first wall, a plurality of impingement cooling holes disposed in the first wall such that the plurality of impingement cooling holes are spaced apart from each other along the flow channel, and a flow diverter convexly protruding from a surface of the second wall in each space between injection axes of the plurality of impingement cooling holes.

A liner for a combustor includes a support member, an intermediate member, and a liner member. The intermediate member is positioned intermediate the support member and the liner member and has a plurality of protrusions and a plurality of recesses. The support member is coupled to the intermediate member at a tangent of each protrusion. Additionally, the liner member is comprised of a ceramic matrix composite material. The liner member is coupled to the intermediate member at a tangent of each recess.

Combustors and gas turbines are provided. A combustor includes an axial centerline and an end cover. The combustor further includes at least one fuel nozzle that extends from the end cover and at is least partially surrounded by a combustion liner. The combustion liner extends between the at least one fuel nozzle and an aft frame and that defines a combustion chamber. An outer sleeve is spaced apart from and surrounds the combustion liner such that an annulus is defined therebetween. The outer sleeve defines at least one aperture. A wake energizer is mounted on the outer sleeve. The wake energizer defines at least one passage that is angled with respect to the axial centerline of the combustor. The at least one passage aligns and is in fluid communication with the at least one aperture of the outer sleeve.

A liner for a combustor in a gas turbine engine and a related method. The liner includes a liner body having a cold side and a hot side. The liner includes a dilution array having a plurality of dilution passages, each dilution passage of the plurality of dilution passages having a concatenated geometry repeating in a predetermined pattern and extending circumferentially around the liner body. The dilution passage integrates a first dilution air flow flowing through the dilution passage from the cold side to the hot side and a second dilution air flow flowing through the dilution passage from the cold side to the hot side into an integrated dilution air flow and injects the integrated dilution air flow into a core primary combustion zone of the combustor to attain a predetermined combustion state of the combustor. The dilution array is repeated along an axial length of the liner body.

A plurality of cooling passages extending in an axial direction are formed in a transition piece so as to be aligned in a circumferential direction and the axial direction. One or more downstream side passages are formed in a downstream side region (Rd) within one circumferential region. One or more upstream side passages are formed in an upstream side region Ru within the circumferential region. The total cross-sectional area per unit circumferential length of the one or more downstream side passages is larger than the total cross-sectional area per unit circumferential length of the one or more upstream side passages.