The present invention relates to a method for operating a lean premix burner of an aircraft gas turbine, where fuel and primary supporting air are supplied by means of a supporting burner (pilot burner) arranged centrically to the burner axis, where secondary air surrounding the supporting burner is supplied, and where fuel and air are supplied by means of a main burner, characterized in that the primary supporting air is supplied in an amount of 5 vol % to 10 vol % of the total air quantity, that the secondary supporting air is supplied in an amount of 5 vol % to 20 vol % and that 35 vol % to 75 vol % of the total air quantity are supplied via the main burner in the partial load range and in the full load range.

A fuel nozzle for a gas turbine engine that includes: an elongated centerbody; an elongated peripheral wall formed about the centerbody so to define a primary flow annulus; a primary fuel and air supply in fluid communication with the primary flow annulus; and a pilot nozzle. The pilot nozzle, formed in the centerbody, may have axially elongated air and mixing tubes extending between inlets and outlets defined, respectively, through upstream and downstream faces of the pilot nozzle. A secondary air supply may be communicate with the inlets of the air and mixing tubes. A fuel port may be positioned in the mixing tubes for connecting each to a secondary fuel supply. An uninterrupted sidewall sealing structure in each of the air tubes may segregate an air flow therethrough from the secondary fuel supply. The air and mixing tubes may be configured as canted tubes so to induce a downstream swirling flow.

A pilot burner assembly for a combustion volume in a gas turbine engine includes a pilot burner, a pilot-fuel supply line, and a pilot-air supply line. The pilot burner has a burner face that includes a plurality of pilot-fuel injection holes. The pilot-fuel injection holes provide a pilot-fuel to the combustion volume. The pilot-fuel supply line is fluidly connected to the pilot-fuel injection holes for supplying the pilot-fuel to the pilot-fuel injection holes. The pilot-air supply line provides a pilot-air to the pilot burner. The pilot-air is supplied to the combustion volume through the burner face. Pilot-air injection holes are located on the burner face and fluidly connected to the pilot-air supply line. The pilot-air injection holes inject the pilot-air into the combustion volume. A gas turbine has the pilot burner assembly.

An embodiment of a torch ignitor system for combustor of a gas turbine engine includes a torch ignitor, the torch ignitor having a combustion chamber oriented about an axis, the combustion chamber having axially upstream and downstream ends defining a flow direction through the combustion chamber, along the axis. The torch ignitor system also includes a cap defining the axially upstream end of the combustion chamber and oriented about the axis, wherein the cap is configured to receive a fuel injector and at least one glow plug, a tip at a downstream end of the combustion chamber, and a passage for pressurized oxygen containing gas passing through the cap from an exterior of the combustion chamber and in fluid communication with the combustion chamber. An embodiment of a method for starting a gas turbine engine is also disclosed.

Adhesion of particulate matters to the burner accompanying combustion in a lean-combustion gas turbine combustor is suppressed, and the structural reliability is improved. In a gas turbine combustor including: a tubular liner that forms a combustion chamber; and a burner including an air hole plate that is arranged at an inlet of the liner and includes a plurality of air holes for guiding compressed air to the combustion chamber, and a plurality of fuel nozzles that are arranged on a side opposite to the combustion chamber with the air hole plate being sandwiched therebetween, the plurality of fuel nozzles each injecting a fuel toward a corresponding air hole, the air holes and the fuel nozzles forming a plurality of concentric annular lines, a plurality of small holes having opening diameters smaller than those of the air holes are provided through the air hole plate such that the plurality of small holes are positioned in an inner area of an innermost annular line of the air holes.

The present invention reduces the concentration of thermal stress on a burner. A gas turbine combustor receiving compressed air from a compressor, mixing the compressed air with a fuel, burning the mixture to generate a combustion gas, and supplying the combustion gas to a turbine. The combustor includes: an inner cylinder internally forming a combustion chamber; an outer cylinder covering the inner cylinder and forming a cylindrical outer circumferential flow path between the inner and outer cylinders to allow the compressed air to flow; and a burner mounted on an end of the outer cylinder, which is positioned on an opposite side to a turbine side, and facing the combustion chamber. The burner includes a cylindrical base frame including a cavity distributing the fuel, and fuel nozzles circularly arranged as viewed from the combustion chamber and connected to the cavity. When viewed from the combustion chamber, slits extending radially are formed in the base frame such that each separate the circumferentially adjacent fuel nozzles from each other.

A combustor assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a combustion chamber, and a fuel injector assembly in communication with the combustion chamber that has a swirler body situated about a nozzle to define an injector passage that converges to a throat. The throat is defined at a distance from the combustion chamber. The nozzle includes a primary fuel injector and an array of secondary plain jet fuel injectors.

A combustor assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a combustion chamber, and a fuel injector assembly in communication with the combustion chamber that has a swirler body situated about a nozzle to define an injector passage that converges to a throat. The throat is defined at a distance from the combustion chamber. The nozzle includes a primary fuel injector and an array of secondary plain jet fuel injectors.

An embodiment of a torch ignitor system for combustor of a gas turbine engine includes a torch ignitor, the torch ignitor having a combustion chamber oriented about an axis, the combustion chamber having axially upstream and downstream ends defining a flow direction through the combustion chamber, along the axis. The torch ignitor system also includes a cap defining the axially upstream end of the combustion chamber and oriented about the axis, wherein the cap is configured to receive a fuel injector and at least one glow plug, a tip at a downstream end of the combustion chamber, and a passage for pressurized oxygen containing gas passing through the cap from an exterior of the combustion chamber and in fluid communication with the combustion chamber. An embodiment of a method for starting a gas turbine engine is also disclosed.

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.