Combustion chamber of a combustor for a gas turbine

Abstract

A combustion chamber of a combustor for a gas turbine is provided. A combustion chamber includes 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 extends between the radial inner wall portion and radial outer wall portion of the segment.

Claims

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.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) FIG. 1 is a schematic diagram of a gas turbine; and

(2) FIG. 2 is a schematic diagram of a combustor and its combustion chamber, in accordance with aspects of the present technique.

(3) FIG. 3 is a schematic end view of the annular combustor looking at a second section.

(4) FIG. 4 is a schematic end view of the annular combustor looking at a first section.

DETAILED DESCRIPTION OF INVENTION

(5) FIG. 1 is a schematic diagram of a gas turbine 10 depicting internal components. The gas turbine 10 includes a rotor 13 which is mounted such that it can rotate along an axis of rotation 12, has a shaft 11 and is also referred to as a turbine rotor.

(6) The gas turbine 10 includes an intake housing 14, a compressor 15, a combustor 16 having a combustion chamber 20, a turbine 18, and an exhaust-gas housing 19 following one another along the rotor 13. The combustion chamber 20 is an annular combustion chamber with a plurality of coaxially arranged burners 17.

(7) The annular combustion chamber 20 is in communication with an annular hot-gas passage 21, where, by way of example, four successive turbine stages 22 form the turbine 18.

(8) It may be noted that each turbine stage 22 is formed, for example, from two blade or vane rings. As seen in the direction of flow of a working medium 23 from the combustion chamber 20 to the turbine 18, in the hot gas passage 21 a row 25 of guide vanes 40 is followed by a row 35 formed from rotor blades 30. The guide vanes 40 are secured to an inner housing 48 of a stator 53, whereas the rotor blades 30 of the row 35 are fitted to the rotor 13 for example by means of a turbine disk 43.

(9) A generator not shown in FIG. 1 is coupled to the rotor 13. During the operation of the gas turbine 10, the compressor 15 sucks in air 45 through the intake housing 14 and compresses it. The compressed air provided at the turbine-side end of the compressor 15 is passed to the burners 17, where it is mixed with a fuel. The mix is then burnt in the combustion chamber 20, forming the working medium 23. From there, the working medium 23 flows along the hot-gas passage 21 past the guide vanes 40 and the rotor blades 30. The working medium 23 is expanded at the rotor blades 30, transferring its momentum, so that the rotor blades 30 drive the rotor 13 and the latter in turn drives the generator coupled to it.

(10) In addition, while the gas turbine 10 is in operation, the components which are exposed to the hot working medium 23 are subjected to thermal stresses. The guide vanes 40 and the rotor blades 30 of the first turbine stage 22, as seen in the direction of flow of the working medium 23, together with the heat shield bricks which line the annular combustion chamber 20, are subject to the highest thermal stresses. These components are typically cooled by a coolant, such as oil.

(11) As will be appreciated, the components of the gas turbine 10 are made from a material such as superalloys which are iron-based, nickel-based or cobalt-based. More particularly, the turbine vanes 40 and/or blades 30 and components of the combustion chamber 20 are made from the superalloys mentioned hereinabove.

(12) The combustion chamber 20 which is an annular combustion chamber 20 in the presently contemplated configuration includes a multiplicity of burners 17 arranged circumferentially around the axis of rotation 12 and open out into a common combustion chamber space and generates flames. To achieve a high efficiency, the combustion chamber 20 is designed for a temperature of the working medium 23 of approximately 1000 degree Celsius to 1600 degree Celsius. To allow a long service life even with these operating parameters, which are unfavorable for the materials, the combustion chamber wall is provided, on its side which faces the working medium 23, with an inner lining formed from heat shield elements.

(13) Referring now to FIG. 2, a schematic diagram of the combustor 16 and its combustion chamber 20, respectively, is depicted in accordance with aspects of the present technique. The combustor 16 includes the combustion chamber 20 which in the presently contemplated configuration is an annular combustion chamber which includes a plurality of segments arranged circumferentially around the axis 12. FIG. 2 shows a cross section through one of those segments. As an example, a total of twenty segments would form the combustion chamber 20. Each segment includes an inner wall portion 54 and an outer wall portion 56.

(14) It may be noted that the inner wall portion 54 and the outer wall portion 56 are positioned radially outwards from the axis 12.

(15) In accordance with aspects of the present technique, the segment has a first section 62 and a second section 64, with the burner installed at an opening 63 at the first section 62 and an airfoil 52 such as a guide vane at the second section 64.

(16) It may be noted however, that the first section may be at the first end and the second section may be at the second end, wherein the first end and the second end are opposing each other. For the purpose of explanation the terms first section and first end and the second section and second end are used interchangeably.

(17) As previously noted, the combustion chamber 20 includes the opening 63 at the first end 62 as depicted in FIG. 2. A burner 17 is installed at the opening 63 at the first end 62. Air from the compressor 15 is directed via a panel 72 and through the airfoil 52 in to the combustion chamber 20 and mixed with fuel. Fuel is directed into the combustion chamber via a fuel pipe 69. The air and fuel mixture is ignited by the burner 17 to produce the working medium 23.

(18) In accordance with aspects of the present technique, the airfoil 52 is present at the second end 64. The airfoil 52 extends between the inner wall portion 54 and an outer wall portion 56. The compressed air from the compressor 15 is directed into the airfoil 52 as indicated by reference numeral 51. Air 51 in the airfoil 52 could also be swirled to create turbulence.

(19) The combustor segment includes an inner surface 60 and an outer surface 58 forming a channel 70 there between to conduct air from the airfoil 52 to the channel 70. Air is mixed with a fuel supplied through the fuel pipe 69 and is ignited by the burner 17 to generate flames 68 and hence produce the working medium 23 for the turbine. This working medium 23 is guided through an exit by the airfoil 52 present at the second end 64 out of the combustion chamber 20.

(20) Additionally the combustor 16 may include cooling holes, or cooling pipes at the end walls to supply cooling air to cool the walls of the combustion chamber 20.

(21) As previously noted, the panel 72 is located at the first section or the first end 62 inside the combustion chamber 20 which acts as a Helmholtz panel to draw air into the combustion chamber 20. The panel 72 alongwith the airfoil 52 acts as a Helmholtz resonator and will keep the air inside the chamber 20 to ensure effective mixing of the air with the fuel and hence better combustion is achieved.

(22) FIG. 3 is an end view of the annular combustor 16 looking upstream at the second section 64. FIG. 4 is an end view of the annular chamber 16 looking downstream at the first section 62. As previously noted and as can be seen in FIGS. 3 and 4, the combustion chamber 20 includes a plurality of segments 60 separated at respective interfaces 62. In FIG. 3 an example embodiment having sixteen segments 60 are shown. The segments 60 are arranged adjacent to each other in a manner such that two segments 60 are assigned to one burner 17. In addition, each segment 60 includes two airfoils 52 located adjacent to each other. The inner wall portion 54, the outer wall portion 56 and the airfoil 52 in a segment 60 are formed of one piece of a material. More particularly, the airfoil 52, the inner wall portion 54 and the outer wall portion 56 are cast to produce a single piece material.

(23) In accordance with the aspects of the present technique, the airfoil 52 and the wall portions 54, 56 are made of material such as alloys, for example nickel-based superalloy. These alloys are capable of withstanding high temperatures which may exceed 650 degree centigrade. The airfoil 52 and the wall portions 54, 56 are cast from the same type of alloy such as, Nickel-based gamma prime strengthened alloy.

(24) It may be noted that the inner wall 54 and the outer wall 56 may be coated with a thermal barrier coating to protect against the high temperatures of the hot gas. Hence it may be noted that the alloys in the present technique are chosen which are compatible with the thermal barrier coatings. Furthermore, it may be noted that alloys such as Nickel-based gamma prime strengthened alloys include a higher quantity of aluminum than the traditional alloys used in the combustors. The presence of aluminum increases the life time of the thermal barrier coatings that are applied to the wall.

(25) Additionally, the alloys for casting the segments of the combustion chamber are chosen which have a better castability and are capable of casting large components such as the segments of combustion chamber 20, such as IN738LC, which is a nickel-based super alloy and has a chemical composition in wt % as Cobalt 8.59, Chromium 16.08, Aluminum 3.43, Silicon 0.18, Carbon 0.11, Phosphorus 0.01, Iron 0.50, Boron 0.05, Sulfur 0.01, Tungsten 2.67, Tantalum 1.75, Nobelium 0.90, Titanium 3.38, Manganese 0.03, Copper 0.03 and Nickel as remaining.

(26) Alternatively, alloy such as CM247CC, which is also a nickel based superalloy may be used for casting the segment. This alloy has a composition in wt % as Cobalt 10, Chromium 8, Molybdenum 0.5, Tungsten 9.5, Aluminum 5.65, Tantalum 3, Hafnium 1.5, Zirconium 0.1, Carbon 0.1 and Nickel as remaining.

(27) Although the invention has been described with reference to specific embodiments, this description is not meant to be construed in a limiting sense. Various modifications of the disclosed embodiments, as well as alternate embodiments of the invention, will become apparent to persons skilled in the art upon reference to the description of the invention. It is therefore contemplated that such modifications can be made without departing from the embodiments of the present invention as defined.