Provided is a gas turbine combustor that can achieve improvement of the combustion stability. A gas turbine combustor includes a pilot burner of the diffusion combustion type, a pilot flow control valve that regulates a flow rate of fuel to be supplied to the pilot burner, a main burner of the premix combustion type arranged on an outer circumference side of the pilot burner, main flow control valves that regulate flow rates of fuel to be individually supplied to burner sectors into which the main burner is sectioned in a circumferential direction, and a controller configured to control the pilot flow control valve and the main flow control valves. The controller controls the main flow control valves such that, when fuel is to be supplied to all the burner sectors, a difference in fuel flow rate occurs between at least one burner sector and the other burner sectors among the burner sectors.

An integrated combustor nozzle includes a combustion liner that extends between an inner liner segment and an outer liner segment along a radial direction. The combustion liner including a forward end portion, an aft end portion, a first side wall, and a second side wall. An impingement panel having an impingement plate disposed along an exterior surface of one of the inner liner segment or the outer liner segment. The impingement plate defines a plurality of impingement apertures that direct coolant in discrete jets towards the exterior surface of the inner liner segment or the outer liner segment. The impingement panel includes an inlet portion that extends from the impingement plate to a collection duct. The impingement panel further includes a plurality of supports spaced apart from one another. The plurality of supports extend between, and are coupled to, the inlet portion, the collection duct, and the impingement plate.

A main mixer for an engine. The main mixer includes a plurality of mixer vanes located circumferentially around a mixer body. Each mixer vane includes a waveform profile. The waveform profile detaches or trips a boundary layer of an air flow across the mixer vane such that the waveform profile introduces turbulence into the air flow.

A gas turbine combustor includes a main mixer for providing an air flow to mix with a fuel flow in a combustor. The main mixer includes an annular mixer body, a plurality of mixer vanes located circumferentially around the annular mixer body, and a plurality of wedges extending radially outward from each of the plurality of mixer vanes. At least one of the wedges includes serrations. The air flow through the main mixer enters the main mixer at a leading edge of the mixer vane, flows over the plurality of wedges, and exits the mixer vanes at a trailing edge of the main mixer. The wedges create vortices in the air flow to provide a uniform fuel-air flow.

To suppress separation of air flows in air holes that are positioned in the outermost circumferential row in an air hole plate, but are far from a turn guide. A gas turbine combustor includes: an inner cylinder forming a combustion chamber; an outer cylinder forming a cylindrical outer circumference flow path through which compressed air flows, between the outer cylinder and the inner cylinder; an end cover closing an end portion of the outer cylinder; an air hole plate having air holes introducing the compressed air passed through the outer circumference flow path, into the combustion chamber; a plurality of fuel nozzles injecting a fuel into the combustion chamber via the air holes; a turn guide provided to the inner cylinder or the air hole plate and guiding a turn of the compressed air passed through the outer circumference flow path; and an auxiliary guide provided at an outer circumferential portion of a surface of the air hole plate facing the fuel nozzles, so as to be positioned on an inner circumference side with respect to the turn guide.

A fuel mixer configured to provide a fuel-air mixture to a combustor of an engine. The fuel mixer may include a mixer body having a mixer outer wall, a center body, an annular passageway defined between the mixer outer wall and the center body, and a fuel tube assembly placed circumferentially about the mixer body. The fuel tube assembly may include at least one fuel channel for injecting a fuel flow into the annular passageway. The fuel tube assembly may be configured to cool a boundary layer flow present in the annular passageway. The fuel tube assembly may be configured to cool the mixer outer wall, the center body, or both the mixer outer wall and the center body. Heat from the mixer outer wall, the center body, or both the mixer outer wall and the center body, may pass to the fuel flow in the fuel tube assembly.

A combustor includes an outer can into which fuel is introduced, an outer head disposed on a front side of the outer can, an inner can disposed inside of the outer can and having a combustion chamber in which a fuel-air mixture is combusted, and an inner head disposed to mix the fuel and the compressed air and supply the mixture into the inner can. The inner head includes a head plate covering a front side of the inner can, and nozzle assemblies disposed to mix the fuel and the compressed air and supply the mixture rearwards. The nozzle assembly includes a nozzle head into which fuel is introduced and nozzles. The nozzles each is coupled between the nozzle head and the head plate to mix the fuel and the compressed air and supply the mixture rearwards. The nozzles each has a shape with a diameter decreasing and increasing toward the rear side thereof

A combustor nozzle and a combustor for a gas turbine including the same are provided. The combustor nozzle may include a plurality of vanes disposed radially on an outer circumferential surface thereof, each vane including an internal cavity to which fuel is supplied, wherein the vane may include an airfoil in cross-section, wherein at least one of a pressure surface and/or a suction surface of the airfoil is disposed in a height direction thereof with a plurality of nozzle holes communicating with the cavity, wherein the vane may include a V-shaped nozzle hole including a pair of outlets on both sides of the airfoil based on a leading edge of the airfoil, wherein the V-shaped nozzle hole may form an angle (α) diverging outward with respect to a horizontal plane across the leading edge, and an acute angle (β) radially outward with respect to a vertical plane perpendicular to the horizontal plane.

A fuel nozzle device (100) for injecting liquid fuel into a combustion chamber (52) of a gas turbine engine (10) includes an outer tube (101), a rear end wall (102) closing a base end of the outer tube, a tapered conical tube (104) defining a first air passage (112) therein, and a second air passage (114) having an annular cross section jointly with the outer tube, a fuel passage (108) axially passed through the rear end wall, and leading to a fuel ejection port (109) directed toward an inner circumferential surface of a base end of the conical tube, a first air introduction passage (111) passed through the outer tube to communicate with the first air passage, and a second air introduction passage (113) passed through the outer tube to communicate with the second air passage.

Combustor systems are provided. For example, a combustor system comprises a combustor having forward and aft ends and including annular inner and outer liners that each extend generally along an axial direction and define a combustion chamber therebetween. The combustor system also comprises a fuel nozzle having an outlet defined in an outlet end of the fuel nozzle and including a pilot swirler. The outlet is positioned at the forward end of the combustor to direct a fuel-air mixture into the combustion chamber. The combustor system further comprises a main mixer attached to the outlet end of the fuel nozzle and extending about the outlet. A total combustor airflow through the combustor comprises a pilot swirler airflow that is greater than about 14% of the total combustor airflow and a main mixer airflow that is less than about 50% of the total combustor airflow.