Combustion chamber of a combustor for a gas turbine

1. A combustion chamber for an annular combustor for a gas turbine, comprising:

a plurality of segments arranged annularly about an axis of the combustion chamber, each segment comprising:

a radial inner wall portion and a radial outer wall portion,

a first section comprising an opening for the installation of a burner, and

a second section at which at least one airfoil connects the radial inner wall portion and the radial outer wall portion of the segment, wherein the inner wall portion, the outer wall portion, and the at least one airfoil are cast to form a one-piece body, wherein the at least one airfoil comprises a first end located at the radial inner wall portion and a second end located at the radial outer wall portion, and wherein the at least one airfoil extends continuously between the first and second ends of the at least one airfoil;

wherein the first section and the second section are located respectively at a first end and an opposing second end of the combustion chamber,

wherein each segment further comprises an inner surface and an outer surface with a channel defined between the inner surface and the outer surface,

wherein compressed air from a compressor of the gas turbine is directed into the at least one airfoil,

wherein compressed air from the at least one airfoil is conducted into the channel, and

wherein the at least one airfoil is located at the opposing second end of the combustion chamber and guides a working medium through an exit located at the opposing second end of the combustion chamber.

2. The combustion chamber according to claim 1 , wherein the opposing second end of the combustion chamber is located downstream from the first end of the combustion chamber.

3. The combustion chamber according to claim 1 , wherein each segment comprises two airfoils, the two airfoils extending between a respective radial inner wall portion and a respective radial outer wall portion.

4. The combustion chamber according to claim 1 , wherein the outer surface is brazed.

5. The combustion chamber according to claim 1 , further comprising a panel located at the first end of the combustion chamber for drawing compressed air into the combustion chamber.

6. The combustion chamber according to claim 5 , wherein the panel along with the at least one airfoil acts as a Helmholtz resonator that is effective to ensure effective mixing of compressed air in the combustion chamber with fuel in the combustion chamber.

7. The combustion chamber according to claim 1 , wherein the at least one airfoil and the radial inner wall portion and the radial outer wall portion are formed from an alloy.

8. The combustion chamber according to claim 7 , wherein the alloy is one of a Nickel based gamma prime strengthened alloy, IN738LC, and CM247CC.

9. The combustion chamber according to claim 1 , wherein two adjacent segments of the plurality of segments are assigned to one burner.

10. A combustor comprising a combustion chamber according to claim 1 .

11. A gas turbine, comprising a combustor comprising the combustion chamber according to claim 1 .

12. A combustion chamber for an annular combustor for a gas turbine, comprising:

a plurality of segments arranged annularly about an axis of the combustion chamber, each segment comprising:

a radial inner wall portion and a radial outer wall portion,

a first section comprising an opening for the installation of a burner, and

a second section at which at least one airfoil connects the radial inner wall portion and the radial outer wall portion of the segment;

wherein the inner wall portion, the outer wall portion, and the at least one airfoil are cast to form a one-piece body, and wherein the at least one airfoil comprises a first end located at the radial inner wall portion and a second end located at the radial outer wall portion, and wherein the at least one airfoil extends continuously between the first and second ends of the at least one airfoil,

wherein the first section and the second section are located respectively at a first end and an opposing second end of the combustion chamber,

wherein each segment further comprises an inner surface and an outer surface with a channel defined between the inner surface and the outer surface, and

wherein each segment is configured to direct compressed air originating from a compressor of the gas turbine into the at least one airfoil at a location at which the at least one airfoil interfaces with the radial outer wall portion, to direct compressed air out of the at least one airfoil into the channel in the radial inner wall portion, and to direct compressed air out of the channel in the radial inner wall portion and into the combustion chamber through an exit located at the second end of the combustion chamber.