A gas turbine combustor with a relatively simple structure is configured to attenuate pressure fluctuation owing to combustion oscillation while securing mechanical reliability. The gas turbine combustor includes a combustion liner that forms a combustion chamber for generating combustion gas, a combustion casing disposed at an outer circumferential side of the combustion liner, and a burner for supplying air flowing between the combustion liner and the combustion casing, and fuel to be supplied from a fuel supply system to the combustion chamber. The combustor further includes a vane disposed at the outer circumferential side of the combustion liner, a plurality of supports disposed at an inner side of the combustion casing for fixing the vane, and a pressure dynamics damping hole formed in the combustion liner at a position corresponding to the vane for communication with the combustion chamber.

A combustion chamber (18) of a thermodynamic cycle turbine with a recuperator, for electrical energy production, comprising a casing (56) housing a flame tube (64) with a perforated diffuser for passage of the hot compressed air, a primary zone (ZP) that receives part of the hot compressed air flow and where combustion takes place, and a dilution zone (ZD) where the burnt gases from the primary zone mix with the remaining part of the hot compressed air flow, said chamber further comprising an injection means (76) for injecting at least one fuel. The flame tube carries a flame stabilizer (82) comprising perforated diffuser (88), at least one combustion gas recirculation passage (98) and a mixing tube (94).

A combustion chamber (18) of a thermodynamic cycle turbine with a recuperator, for electrical energy production, comprising a casing (56) housing a flame tube (64) with a perforated diffuser for passage of the hot compressed air, a primary zone (ZP) that receives part of the hot compressed air flow and where combustion takes place, and a dilution zone (ZD) where the burnt gases from the primary zone mix with the remaining part of the hot compressed air flow, said chamber further comprising an injection means (76) for injecting at least one fuel. The flame tube carries a flame stabilizer (82) comprising perforated diffuser (88), at least one combustion gas recirculation passage (98) and a mixing tube (94).

A method for operating a rotating detonation engine, having a radially outer wall extending along an axis; a radially inner wall extending along the axis, wherein the radially inner wall is positioned within the radially outer wall to define an annular detonation chamber having an inlet and an outlet, wherein the method includes flowing liquid phase fuel along at least one wall of the radially inner wall and the radially outer wall in a direction from the outlet toward the inlet to cool the at least one wall and heat the liquid fuel to provide a heated liquid fuel; flowing the heated liquid fuel to a mixer at the inlet to reduce pressure of the heated liquid fuel, flash vaporize the heated liquid fuel and mix flash vaporized fuel with oxidant to produce a vaporized fuel-oxidant mixture; and detonating the mixture in the annular detonation chamber.

A method for operating a rotating detonation engine, having a radially outer wall extending along an axis; a radially inner wall extending along the axis, wherein the radially inner wall is positioned within the radially outer wall to define an annular detonation chamber having an inlet and an outlet, wherein the method includes flowing liquid phase fuel along at least one wall of the radially inner wall and the radially outer wall in a direction from the outlet toward the inlet to cool the at least one wall and heat the liquid fuel to provide a heated liquid fuel; flowing the heated liquid fuel to a mixer at the inlet to reduce pressure of the heated liquid fuel, flash vaporize the heated liquid fuel and mix flash vaporized fuel with oxidant to produce a vaporized fuel-oxidant mixture; and detonating the mixture in the annular detonation chamber.

A combustor component of a combustor for combusting fuel to produce a combustion gas includes a combustion cylinder forming a passage for the combustion gas, a first acoustic device which internally includes a first cavity communicating with the passage via a first through hole formed in the combustion cylinder, a second acoustic device which is located on a radially outer side of the first acoustic device so as to cover the first acoustic device, and internally includes a second cavity communicating with the passage via a second through hole formed in the combustion cylinder, and a first communication passage causing the first cavity and an outer space of the combustion cylinder to communicate with each other without via the first through hole and the second cavity.

A combustor for a gas turbine engine includes a liner enclosing a combustion chamber and defining air passages through the liner, a fuel nozzle fluidly connected to the combustion chamber, and a louver disposed inside the combustion chamber over the air passages. The louver extends circumferentially along the liner and is connected to the liner by a fastener. The fastener spacing at least one of axial edges of the louver from the liner to define an air outlet between the at least one of the axial edges and the liner. A method of manufacturing a combustor of an aircraft engine is also disclosed.

A combustor for a gas turbine engine includes a liner enclosing a combustion chamber and defining air passages through the liner, a fuel nozzle fluidly connected to the combustion chamber, and a louver disposed inside the combustion chamber over the air passages. The louver extends circumferentially along the liner and is connected to the liner by a fastener. The fastener spacing at least one of axial edges of the louver from the liner to define an air outlet between the at least one of the axial edges and the liner. A method of manufacturing a combustor of an aircraft engine is also disclosed.

A gas turbine facility of an embodiment includes: a combustor casing; a combustor provided in the combustor casing; a cylinder surrounding a periphery of the combustor and dividing a space between the combustor casing and the combustor; a turbine rotated by combustion gas exhausted from the combustor; a heat exchanger cooling the combustion gas exhausted from the turbine; a pipe through which a part of the combustion gas cooled in the heat exchanger passes in the heat exchanger to be heated, the pipe guiding the combustion gas heated in the heat exchanger into the cylinder; and a pipe guiding another part of the combustion gas cooled in the heat exchanger to a space between the combustor casing and the cylinder.

A gas turbine facility of an embodiment includes: a combustor casing; a combustor provided in the combustor casing; a cylinder surrounding a periphery of the combustor and dividing a space between the combustor casing and the combustor; a turbine rotated by combustion gas exhausted from the combustor; a heat exchanger cooling the combustion gas exhausted from the turbine; a pipe through which a part of the combustion gas cooled in the heat exchanger passes in the heat exchanger to be heated, the pipe guiding the combustion gas heated in the heat exchanger into the cylinder; and a pipe guiding another part of the combustion gas cooled in the heat exchanger to a space between the combustor casing and the cylinder.