Various embodiments include a trapped vortex combustor and a method for operating trapped vortex combustor. In one embodiment, the trapped vortex combustor comprises a trapped vortex combustion zone and at least one secondary combustion zone disposed downstream of the trapped vortex combustion zone. The trapped vortex combustion zone is operable to receive and combust a first fuel and a first air and produce a first combustion product flowing toroidally therein. The at least one secondary combustion zone is operable to receive and combust the first combustion product and at least one second injection consisting of fuel and/or air and produce at least one second combustion product therein. The combustor may reduce the residence time of the highest temperature combustion products and achieve the lower NOx emission.

An annular combustion chamber (10) for a turbomachine (100), the combustion chamber presenting an axial direction (X), a radial direction, and an azimuth direction, and comprising a first annular wall (12) and a second annular wall (14), each wall delimiting at least a portion of the volume of the annular combustion chamber (10), the first and second walls (12, 14) presenting complementary fitting elements (12d, 14d), the first wall (12) presenting at least one first through hole (12f), while the second wall (14) presents at least one second through hole (14f), the combustion chamber (10) also having at least one pin (18) engaged in a pair of holes comprising a first hole (12f) and a second hole (14f), said pin (18) locking the fitting of the first and second walls (12, 14).

An annular combustion chamber (10) for a turbomachine (100), the combustion chamber presenting an axial direction (X), a radial direction, and an azimuth direction, and comprising a first annular wall (12) and a second annular wall (14), each wall delimiting at least a portion of the volume of the annular combustion chamber (10), the first and second walls (12, 14) presenting complementary fitting elements (12d, 14d), the first wall (12) presenting at least one first through hole (12f), while the second wall (14) presents at least one second through hole (14f), the combustion chamber (10) also having at least one pin (18) engaged in a pair of holes comprising a first hole (12f) and a second hole (14f), said pin (18) locking the fitting of the first and second walls (12, 14).

An afterburner is disclosed for use with a gas turbine engine and, in one form, is structured to receive a bypass air. The afterburner can be situated in a bypass duct and can be a toroidal combustor or a can combustor. In one embodiment, the afterburner includes a combustor arranged to receive bypass air and a plurality of vanes distributed downstream of a turbine of the gas turbine engine. The vanes can include one or more exit apertures through which hot combustion flow from the afterburner combustor can be injected. The exit apertures can be protrusions or slots in some forms. In one embodiment, a cooling passage is arranged around the exit apertures. An upstream vane portion can be positioned to inject fuel to be combusted via interaction with hot flow that is discharged through the exit apertures.

An afterburner is disclosed for use with a gas turbine engine and, in one form, is structured to receive a bypass air. The afterburner can be situated in a bypass duct and can be a toroidal combustor or a can combustor. In one embodiment, the afterburner includes a combustor arranged to receive bypass air and a plurality of vanes distributed downstream of a turbine of the gas turbine engine. The vanes can include one or more exit apertures through which hot combustion flow from the afterburner combustor can be injected. The exit apertures can be protrusions or slots in some forms. In one embodiment, a cooling passage is arranged around the exit apertures. An upstream vane portion can be positioned to inject fuel to be combusted via interaction with hot flow that is discharged through the exit apertures.

There is provided a method for improving the combustion efficiency of a combustor of a gas turbine engine powering an aircraft. The method comprises selectively using two distinct fuel injection units or a combination thereof for spraying fuel in a combustion chamber of the combustor of the gas turbine engine. A first one of the two distinct fuel injection units is selected and optimized for high power demands, whereas a second one of the two distinct fuel injection units is selected and optimized for low power level demands. In operation, the fuel flow ratio between the two distinct injection units is controlled as a function of the power level demand.

There is provided a method for improving the combustion efficiency of a combustor of a gas turbine engine powering an aircraft. The method comprises selectively using two distinct fuel injection units or a combination thereof for spraying fuel in a combustion chamber of the combustor of the gas turbine engine. A first one of the two distinct fuel injection units is selected and optimized for high power demands, whereas a second one of the two distinct fuel injection units is selected and optimized for low power level demands. In operation, the fuel flow ratio between the two distinct injection units is controlled as a function of the power level demand.

The present application provides a combustor for use with a gas turbine engine. The combustor may include a liner and a flow sleeve surrounding the liner with the liner and the flow sleeve defining a flow path therebetween. A liner retaining feature extends into the flow path to retain the liner within the flow sleeve. The liner retaining feature may include a bolt therein.

The present application provides a combustor for use with a gas turbine engine. The combustor may include a liner and a flow sleeve surrounding the liner with the liner and the flow sleeve defining a flow path therebetween. A liner retaining feature extends into the flow path to retain the liner within the flow sleeve. The liner retaining feature may include a bolt therein.

A propulsion system includes a first compressor in fluid communication with a fluid source. A first conduit is coupled to the first compressor, and a heat exchanger is in fluid communication with the first compressor via the first conduit. A second conduit is positioned proximal to the heat exchanger. A combustor is in fluid communication with the heat exchanger via the second conduit and is configured to generate a high-temperature gas stream. A third conduit is coupled to the combustor, and a first thrust augmentation device is in fluid communication with the combustor via the third conduit. The heat exchanger is positioned within the gas stream generated by the combustor.