Disclosed is a fuel mixed injection method for a gas turbine. The method includes the following steps: arranging a secondary fuel injection nozzle and a secondary air injection nozzle on a secondary combustion section, wherein the secondary fuel injection nozzles is closer to a main combustion section than the secondary air injection nozzle; and injecting secondary fuel and secondary primary air in sequence through the secondary fuel injection nozzle and the secondary air injection nozzle, respectively, thus enabling the secondary fuel to spontaneously combust in a mainstream high-temperature flue gas atmosphere to form a transverse jet flame and increase the flame lift-off height.

A turbine engine can include a compressor section, a combustion section, and a turbine section in serial flow arrangement. A combustor in the combustion section can include a combustion chamber, a fuel supply fluidly coupled to the combustion chamber, and a fuel nozzle assembly. The fuel nozzle assembly can include an air flow passage and a fuel flow passage.

A turbine engine can include a compressor section, a combustion section, and a turbine section in serial flow arrangement. A combustor in the combustion section can include a combustion chamber, a fuel supply fluidly coupled to the combustion chamber, and a fuel nozzle assembly. The fuel nozzle assembly can include an air flow passage and a fuel flow passage.

Fuel nozzle assemblies are provided. For example, a fuel nozzle assembly for a combustor system comprises a fuel nozzle having a pilot swirler and an outlet defined in an outlet end, as well as a main mixer attached to the outlet end and extending about the outlet. A total combustor airflow through the combustor system comprises a pilot swirler airflow that is greater than about 14% and a main mixer airflow that is less than about 50% of the total combustor airflow. In further embodiments, the fuel nozzle also comprises main and pilot fuel injectors that each are configured to receive a portion of a fuel flow to the fuel nozzle. The fuel nozzle provides less than about 80% of the fuel flow to the main fuel injector at a high power operating condition of a gas turbine engine in which the fuel nozzle assembly is installed.

A burner arrangement has a plurality of mixing channels which extend parallel to the main axis of the burner arrangement and are arranged in at least two concentric circles, in which mixing channels fuel and discharge air from the compressor are mixed during the operation of the burner arrangement. The mixing channels are grouped together into fuel stages so as to produce an irregular staging in the peripheral direction of the burner arrangement.

A burner arrangement has a plurality of mixing channels which extend parallel to the main axis of the burner arrangement and are arranged in at least two concentric circles, in which mixing channels fuel and discharge air from the compressor are mixed during the operation of the burner arrangement. The mixing channels are grouped together into fuel stages so as to produce an irregular staging in the peripheral direction of the burner arrangement.

The invention relates to a thrust chamber device comprising a thrust chamber with a thrust space having a first portion, a second portion adjacent thereto, and a third portion adjacent to the second portion, the thrust space being delimited in all three portions by an outer nozzle wall having an outer thrust space surface, which outer thrust space surface tapers in the first and second portion toward the third portion, widens in the third portion away from the second portion, and has a narrowest point at the transition from the second portion to the third portion, the first portion being delimited by an inner nozzle wall with an inner thrust space surface, which tapers toward the second portion, an annular combustion chamber being formed between the inner thrust space surface and the outer thrust space surface and extending over the first portion.

A turbine engine and method of operation with a combustor as shown and described. The turbine engine including a combustor comprising an annular array of rich cups and an annular array of lean cups. A method for controlling nitrogen oxides including injecting a fuel/air mixture from the lean cups and the rich cups.

A turbine engine and method of operation with a combustor as shown and described. The turbine engine including a combustor comprising an annular array of rich cups and an annular array of lean cups. A method for controlling nitrogen oxides including injecting a fuel/air mixture from the lean cups and the rich cups.

A Brayton cycle engine including a longitudinal wall extended along a lengthwise direction. The longitudinal wall defines a gas flowpath of the engine. An inner wall assembly is extended from the longitudinal wall into the gas flowpath. The inner wall assembly defines a detonation combustion region in the gas flowpath upstream of the inner wall assembly.