The invention relates to a device for supporting and centring a fuel injector in a turbine engine combustion chamber, which includes means for centring a fuel injector along an axis, which are movable in a plane that is radial to the centring axis (52) in supporting means intended for being attached to the bottom of an annular chamber (18). According to the invention, the centring means include at least two radially external tabs (54, 56) each inserted respectively in a circumferential recess (60, 64) of the supporting means, the device including circumferential abutment means (78, 80, 74, 76, 82, 86, 84, 88) of the radial tabs (54, 56) of the centring means in the circumferential recesses (60, 64), the circumferential abutment means being configured such as to enable a greater angular displacement of a first (54, 154) one of the radial tabs in a first circumferential recess (60) relative to a second (56, 156) one of the radial tabs in a second circumferential recess (64).

The invention relates to a device for supporting and centring a fuel injector in a turbine engine combustion chamber, which includes means for centring a fuel injector along an axis, which are movable in a plane that is radial to the centring axis (52) in supporting means intended for being attached to the bottom of an annular chamber (18). According to the invention, the centring means include at least two radially external tabs (54, 56) each inserted respectively in a circumferential recess (60, 64) of the supporting means, the device including circumferential abutment means (78, 80, 74, 76, 82, 86, 84, 88) of the radial tabs (54, 56) of the centring means in the circumferential recesses (60, 64), the circumferential abutment means being configured such as to enable a greater angular displacement of a first (54, 154) one of the radial tabs in a first circumferential recess (60) relative to a second (56, 156) one of the radial tabs in a second circumferential recess (64).

A fuel injection device, for a gas turbine, which enhances uniform distribution in concentration of fuel gas and water vapor in a combustion chamber with simple structure and at low cost to effectively reduce NOx, is provided. The fuel injection device with a fuel nozzle to mix fuel gas and water vapor and inject the fuel gas and water vapor into a combustion chamber, includes: a nozzle housing having a mixing chamber thereinside; a first introduction passage to introduce the fuel gas into the mixing chamber from outside of the nozzle housing; and a second introduction passage to introduce the water vapor into the mixing chamber from an outside of the nozzle housing; and a plurality of reverse passages communicating with a downstream end of the mixing chamber and configured to allow for a plurality of reverses of flow of mixed gas from the mixing chamber.

A jet engine includes an inlet (11) which takes in air and a combustor (12) which combusts fuel with the air. The combustor (12) has an injector (20), a plurality of flame stabilizers (21, 22) and a vanishment section (31). The injector (20) injects the fuel. The plurality of flame stabilizers (21, 22) can maintain the flame (F) used for combustion in the combustor (12). The vanishment section (31) is provided to cover the dent of the first flame stabilizer (21) which is situated on the side near to the inlet in the plurality of flame stabilizers (21, 22), and vanishes with the passage of time in the flight.

A jet engine includes an inlet (11) which takes in air and a combustor (12) which combusts fuel with the air. The combustor (12) has an injector (20), a plurality of flame stabilizers (21, 22) and a vanishment section (31). The injector (20) injects the fuel. The plurality of flame stabilizers (21, 22) can maintain the flame (F) used for combustion in the combustor (12). The vanishment section (31) is provided to cover the dent of the first flame stabilizer (21) which is situated on the side near to the inlet in the plurality of flame stabilizers (21, 22), and vanishes with the passage of time in the flight.

A system includes a gas turbine having a turbine shaft disposed along a rotational axis, a turbine casing disposed circumferentially about the turbine shaft, a combustion gas path disposed between the turbine shaft and the turbine casing, and a turbine stage disposed in the combustion gas path. The turbine stage includes a plurality of turbine vanes disposed upstream from a plurality of turbine blades. The gas turbine includes an isothermal expansion system coupled to the turbine stage, wherein the isothermal expansion system includes a plurality of flame stabilizers configured to vary axial positions of combustion within a turbine stage expansion of the turbine stage to reduce temperature variations over the turbine stage expansion. The flame stabilizers are disposed in different axial positions over an axial length between leading and trailing edges of the turbine blades, wherein at least one flame stabilizer is coupled to a plurality of the turbine blades.

A system includes a gas turbine having a turbine shaft disposed along a rotational axis, a turbine casing disposed circumferentially about the turbine shaft, a combustion gas path disposed between the turbine shaft and the turbine casing, and a turbine stage disposed in the combustion gas path. The turbine stage includes a plurality of turbine vanes disposed upstream from a plurality of turbine blades. The gas turbine includes an isothermal expansion system coupled to the turbine stage, wherein the isothermal expansion system includes a plurality of flame stabilizers configured to vary axial positions of combustion within a turbine stage expansion of the turbine stage to reduce temperature variations over the turbine stage expansion. The flame stabilizers are disposed in different axial positions over an axial length between leading and trailing edges of the turbine blades, wherein at least one flame stabilizer is coupled to a plurality of the turbine blades.

A system includes a gas turbine having a turbine shaft disposed along a rotational axis, a turbine casing disposed circumferentially about the turbine shaft, a combustion gas path disposed between the turbine shaft and the turbine casing, and a turbine stage disposed in the combustion gas path. The turbine stage includes a plurality of turbine vanes disposed upstream from a plurality of turbine blades. The gas turbine includes an isothermal expansion system coupled to the turbine stage, wherein the isothermal expansion system includes a plurality of flame stabilizers configured to vary axial positions of combustion within a turbine stage expansion of the turbine stage to reduce temperature variations over the turbine stage expansion. The flame stabilizers are disposed in different axial positions over an axial length between leading and trailing edges of the turbine blades, wherein at least one flame stabilizer is coupled to each of the turbine blades.

A system includes a gas turbine having a turbine shaft disposed along a rotational axis, a turbine casing disposed circumferentially about the turbine shaft, a combustion gas path disposed between the turbine shaft and the turbine casing, and a turbine stage disposed in the combustion gas path. The turbine stage includes a plurality of turbine vanes disposed upstream from a plurality of turbine blades. The gas turbine includes an isothermal expansion system coupled to the turbine stage, wherein the isothermal expansion system includes a plurality of flame stabilizers configured to vary axial positions of combustion within a turbine stage expansion of the turbine stage to reduce temperature variations over the turbine stage expansion. The flame stabilizers are disposed in different axial positions over an axial length between leading and trailing edges of the turbine blades, wherein at least one flame stabilizer is coupled to each of the turbine blades.

A jet engine has an inlet 11 configured to introduce air, and a combustor 12 having a fuel injection port 30a that injects a fuel, and configured to combust the fuel injected from the fuel injection port 30a by using the air. The combustor 12 has a separation section 14 defining the air passage FA through which the air flows, between a rear end 15 of the inlet and the fuel injection port 30a. A plurality of turbulent flow generating sections (20;25) are arranged in the separation section 14 to makes the air flow turbulent. Each of the plurality of turbulent flow generating sections (20;25) contains a member (21;22;25B) which can restrain the turbulence of the air flow by moving or disappearing. It can be prevented that a high-pressure region reaches the inlet so that the thrust of the jet engine is reduced.