A core engine article includes a combustor liner defining a combustion chamber therein and a turbine nozzle. The combustor liner includes a plurality of injector ports, and the plurality of injector ports have a shape that tapers to a corner on a forward side of the injector ports. The turbine nozzle includes a plurality of airfoils. The combustor liner and turbine nozzle are integral with one another. A method of making a core engine article is also disclosed.

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 gas turbine engine comprising a variable geometry combustor having pilot fuel injectors and main fuel injectors; a fuel metering system configured to control fuel flow to the pilot fuel injectors and the main fuel injectors; a variable geometry airflow arrangement for the variable geometry combustor, which is configured to vary the airflow through the pilot fuel injectors and/or the main fuel injectors; a control system configured to control the variable geometry airflow arrangement in dependence upon airflow delivered to the combustor, the fuel flow to the pilot fuel injectors and the main fuel injectors, and a target index of soot emissions, thereby controlling airflow through the pilot fuel injectors and/or the main fuel injectors and hence the quantity of soot produced by combustion.

A gas turbine engine comprising a variable geometry combustor having fuel injectors, a rich-burn zone, a quick-quench zone, and a lean-burn zone, and further comprising quench ports for admitting quench air to the quick-quench zone; a variable geometry airflow arrangement for the variable geometry combustor, which is configured to vary an airflow through the fuel injectors and/or the quench ports; and a control system configured to control the variable geometry airflow arrangement in dependence upon an airflow delivered to the combustor, a fuel flow to the fuel injectors, and a target index of soot emissions to control the quantity of soot produced by combustion.

A gas turbine includes a rotor that is rotatable about an axis, a casing configured to cover the rotor in a circumferential direction and having an annular space therein, a compressor configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing, a plurality of combustors disposed in the casing at equal intervals in the circumferential direction of the rotor and configured to combust the compressed air and fuel taken in from the casing to generate a combustion gas, a turbine driven by the combustion gas, and an air introduction passage defined by a partition plate configured to divide the space in the casing in the circumferential direction of the rotor and an inner circumferential surface of the casing and configured to introduce the compressed air in the casing into the combustor.

Combustors and gas turbines are provided. A combustor includes an axial centerline and an end cover. The combustor further includes at least one fuel nozzle that extends from the end cover and at is least partially surrounded by a combustion liner. The combustion liner extends between the at least one fuel nozzle and an aft frame and that defines a combustion chamber. An outer sleeve is spaced apart from and surrounds the combustion liner such that an annulus is defined therebetween. The outer sleeve defines at least one aperture. A wake energizer is mounted on the outer sleeve. The wake energizer defines at least one passage that is angled with respect to the axial centerline of the combustor. The at least one passage aligns and is in fluid communication with the at least one aperture of the outer sleeve.

Combustors and gas turbines are provided. A combustor includes an axial centerline and an end cover. The combustor further includes at least one fuel nozzle that extends from the end cover and at is least partially surrounded by a combustion liner. The combustion liner extends between the at least one fuel nozzle and an aft frame and that defines a combustion chamber. An outer sleeve is spaced apart from and surrounds the combustion liner such that an annulus is defined therebetween. The outer sleeve defines at least one aperture. A wake energizer is mounted on the outer sleeve. The wake energizer defines at least one passage that is angled with respect to the axial centerline of the combustor. The at least one passage aligns and is in fluid communication with the at least one aperture of the outer sleeve.

An electrically decoupled jet engine. The electrically decoupled jet engine includes a combustion chamber which creates a toroidal flow of air and a rotational electric motor which drives a fuel supply into the combustion chamber. The toroidal flow of air is mixed with the fuel and combusted in the combustion chamber to create thrust.

An electrically decoupled jet engine. The electrically decoupled jet engine includes a combustion chamber which creates a toroidal flow of air and a rotational electric motor which drives a fuel supply into the combustion chamber. The toroidal flow of air is mixed with the fuel and combusted in the combustion chamber to create thrust.

A system is provided with a fuel sweep system configured to couple to a flow sleeve of a combustor along a first fuel conduit. The flow sleeve is configured to be disposed about a liner of the combustor, and the first fuel conduit is configured to extend along the flow sleeve in a compressor discharge chamber disposed about the flow sleeve. The fuel sweep system includes a first fuel sweep louver adjacent a first fuel sweep opening defined through the flow sleeve.