A combustor of an aircraft engine comprises a liner defining a primary and a dilution zone having a hot surface exposed to a flow of combustion gases traveling from the primary zone downstream to the dilution zone and a cold surface. Dilution holes extending through the liner from the cold to the hot surface delimit the primary from the dilution zone. Effusion holes extending through the liner from the cold to the hot surface direct cooling air into the dilution zone. Two or more rows of effusion holes positioned within three dilution hole diameters downstream of the dilution holes are oriented relative to the liner to direct the cooling air in a cooling direction that is at least one of normal to the direction of the flow of gases passing adjacent the effusion holes, and against the direction of the flow of gases passing adjacent the effusion holes.

A gas turbine combustion chamber, having a combustion chamber housing, a burner positioned at least partially in the combustion chamber housing and a flame tube positioned in the combustion chamber housing. An upstream end of the flame tube has a flame tube cover. In the region of the flame tube cover radially inside the flame tube cover a mixing tube of the burner extends, which defines a pre-primary combustion zone. Between the mixing tube of the burner and the flame tube cover an axial air flow passage is formed, via which air can be conducted in the axial direction outside along the mixing tube, and into the flame tube cover radial air flow passages are introduced via which the air following the axial air flow passage can be conducted in the radial direction of the flame tube cover for cooling the flame tube cover.

A combustor assembly for use in a gas turbine engine includes a combustor liner that defines a combustion chamber and includes an axial combustion portion and a curved transition portion. The combustion liner also includes an inner surface and an outer surface and a first plurality of cooling channels defined between the inner and outer surfaces. The combustor assembly also includes a sleeve substantially circumscribing the combustor liner such that an annular cavity is defined between the combustor liner and the sleeve. The sleeve includes a second plurality of cooling channels defined therethrough that are configured to channel a fluid against the combustor liner outer surface.

A combustor for a detonation engine includes a radially outer wall extending along an axis; a radially inner wall extending along the axis, wherein the radially inner wall is positioned at least partially within the radially outer wall to define an annular detonation chamber having an inlet for fuel and oxidant and an outlet; a cooling flow passage defined along at least one of the radially outer wall and the radially inner wall and comprising at least two axially spaced cooling flow passage sections, whereby a different cooling rate can be implemented in the at least two axially spaced cooling flow passage sections.

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.

A premixer array for a gas turbine engine includes a plurality of swirled premixer tubes and a plurality of non-swirled premixer tubes. The premixer array further includes a pilot tube for providing a pilot product. The plurality of swirled premixer tubes may guide and spread out the pilot product in a radial direction, in a circumferential direction, or both in a radial direction and a circumferential direction. The plurality of swirled premixer tubes may include clockwise swirled and counterclockwise swirled premixers. Also, included is a method of controlling the flow of pilot products.

A combined combustor and recuperator is formed with the recuperator surrounding the combustor. Cold gas conduits (14, 16, 20) through the recuperator follow along involute paths toward the combustor. Hot has conduits (26) through the recuperator follow counterflow paths along corresponding involute curves outward from the combustor. The openings (18) in the combustion chamber wall through which cold gas enters the combustion chamber may be directed to impart flow direct to the cold gas to support particular desired behaviour of the cold gas in the portions of the combustion chamber concerned, e.g. supporting a stable vortex flame, enhancing mixing, providing a protective barrier layer.

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.

A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a liner at least partially defining a combustion chamber of the gas turbine engine, wherein the liner comprises a looped feature.

A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a liner at least partially defining a combustion chamber, the liner including a dilution slot in fluid communication with the combustion chamber; and one or more guides configured to redirect cooling medium into the dilution slot.