This exit seal is for connecting an exit flange of a combustor and a shroud in a stator blade of a turbine. The exit seal is equipped with: a seal body that extends in a circumferential direction; and a lid member disposed at a circumferential end of the seal body. The seal body comprises one or more recessed sections that are recessed in a radial direction or in an axial direction and extend in the circumferential direction. At a position of the circumferential end of at least one recessed section among the one or more recessed sections, the lid member is disposed so as to overlap, in the axial direction and the radial direction, with a recessed space formed by the one recessed section.

A combustion cylinder mounting method, wherein a preparation step, a first contact maintenance step, a second contact maintenance step, and a combustion cylinder fixing step are executed. In the preparation step, a combustion cylinder mounting jig equipped with a spacer portion capable of maintaining a predetermined spacing between a first combustion cylinder and a second combustion cylinder adjacent to each other in the circumferential direction is prepared. In the first contact maintenance step, the spacer portion is brought into contact with the first combustion cylinder attached to the stationary portion of a casing. In the second contact maintenance step, the second combustion cylinder is brought into contact with the spacer portion. In the combustion cylinder fixing step, the second combustion cylinder is attached to the stationary portion while the first combustion cylinder and the second combustion cylinder are in contact with the spacer portion.

The invention relates to a turbine engine combustion module (10), in particular for an aircraft turbine engine, designed to carry out constant-volume combustion, comprising: at least two combustion chambers (12A, 12B) arranged about an axis, each chamber (12A, 12B, 12C) comprising a compressed gas intake port (16) and a burnt gas exhaust port (18); and an ignition means that triggers combustion in the combustion chambers (12A, 12B, 12C). The module (10) comprises at least one duct (80) which establishes a communication between a first combustion chamber (12A) and at least one second combustion chamber (12B) in order to inject burnt gases from the first combustion chamber (12A) into the second combustion chamber (12B) so as to trigger combustion in the second combustion chamber (12B).

The invention relates to a turbine engine combustion module (10), in particular for an aircraft turbine engine, designed to carry out constant-volume combustion, comprising: at least two combustion chambers (12A, 12B) arranged about an axis, each chamber (12A, 12B, 12C) comprising a compressed gas intake port (16) and a burnt gas exhaust port (18); and an ignition means that triggers combustion in the combustion chambers (12A, 12B, 12C). The module (10) comprises at least one duct (80) which establishes a communication between a first combustion chamber (12A) and at least one second combustion chamber (12B) in order to inject burnt gases from the first combustion chamber (12A) into the second combustion chamber (12B) so as to trigger combustion in the second combustion chamber (12B).

This combustor is provided with an outer cylinder, an inner cylinder, a direct injection nozzle, a tail pipe, and a base end side acoustic attenuator. The outer cylinder is disposed inside a gas turbine casing. The inner cylinder is disposed on the inner circumferential side of the outer cylinder. The direct injection nozzle is disposed on the inner circumferential side of the inner cylinder. The tail pipe is connected to the inner cylinder, and fuel injected from the direct injection nozzle is burned on the inner circumferential side of the tail pipe. The base end side acoustic attenuator has an outer cylinder formation portion that is a part of a plate forming the outer cylinder, and an acoustic cover forming a base end side space in the gas turbine casing on the outer circumferential side of the outer cylinder in cooperation with the outer cylinder formation portion.

A gas turbine combustor includes a burner composed of a fuel nozzle group having a plurality of fuel nozzles for fuel supply, a fuel nozzle plate structurally supporting the fuel nozzles and serving to distribute the fuel flowing from an upstream side to the fuel nozzles, and a perforated plate located downstream of the fuel nozzles and having nozzle holes corresponding to the fuel nozzles. The fuel nozzle group includes outer circumferential fuel nozzles and inner circumferential fuel nozzles. Each outer diameter of at least a proximal end of the outer circumferential fuel nozzles is larger than that of the inner circumferential fuel nozzles.

A burner comprises: at least one mixing tube extending inside a fuel plenum and having an interior configured to be supplied with an air; and a plurality of fuel injection holes for injecting a fuel supplied to the fuel plenum into the interior of the at least one mixing tube. When the at least one mixing tube is viewed in an axial direction of the mixing tube, a central axis of each of the plurality of fuel injection holes is oblique in a same direction with respect to a circumferential direction of the mixing tube, to a radial direction of the mixing tube.

A combustor component according to at least one embodiment of the present invention includes a cylindrical body which internally includes a combustion chamber, and includes a weld part where a plurality of through holes opening to the combustion chamber are formed, and a housing which is disposed on an outer circumferential side of the cylindrical body to cover a part of the weld part, and defines an acoustic damping space communicating with the combustion chamber via at least one of the through holes. The plurality of through holes in the weld part has a formation density which is higher in a first region of the weld part covered with the housing than in a second region of the weld part positioned outside the housing.

A combustor includes an end cover and at least one fuel nozzle extending from the end cover and at least partially surrounded by a combustion liner. The combustor further includes an outer sleeve spaced apart from and surrounding the combustion liner such that an annulus is defined therebetween. The combustor further includes a fuel injection assembly. The fuel injection assembly includes a fuel injector that extends through the outer sleeve, the annulus, and the combustion liner to the secondary combustion zone. A fuel supply conduit positioned outside of the outer sleeve. The fuel supply conduit extending to the fuel injector. A shielding assembly coupled to the outer sleeve and at least partially surrounding the fuel supply conduit. The at least one fuel sweep opening is defined in the outer sleeve and disposed within the shielding assembly.

A combustor includes an end cover and at least one fuel nozzle extending from the end cover and at least partially surrounded by a combustion liner. The combustor further includes an outer sleeve spaced apart from and surrounding the combustion liner such that an annulus is defined therebetween. The combustor further includes a fuel injection assembly. The fuel injection assembly includes a fuel injector that extends through the outer sleeve, the annulus, and the combustion liner to the secondary combustion zone. A fuel supply conduit positioned outside of the outer sleeve. The fuel supply conduit extending to the fuel injector. A shielding assembly coupled to the outer sleeve and at least partially surrounding the fuel supply conduit. The at least one fuel sweep opening is defined in the outer sleeve and disposed within the shielding assembly.