gas turbine combustor

Abstract

To seal a cooling chamber while facilitating manufacturing of a CMC liner in a combustor for a gas turbine.SOLUTION: A combustor for a gas turbine includes: a metallic shell; a liner disposed on the inner side of the shell and made of a ceramic matrix composite material, the liner including an inner surface defining a combustion chamber, an outer surface facing a side opposite to the combustion chamber and at least one cooling hole opened toward the combustion chamber; and a cooling chamber defined between the shell and the liner. The shell includes: a cylindrical shell body; and a sealing wall structure projecting from the inner surface of the shell body toward the liner.SELECTED DRAWING: Figure 3

Description

Technical Field

[0001] The present disclosure relates to a combustor of a gas turbine.

Background Art

[0002] In a combustor of a gas turbine, in order to achieve high-temperature and high-pressure combustion for improving fuel efficiency, an improvement in the heat-resistant temperature of components is required. It is known that a liner defining a combustion chamber of a combustor can be formed of a ceramic matrix composite (CMC) instead of metal. CMC is lighter than metal. Since CMC has a high heat-resistant temperature, the amount of cooling air can be reduced while achieving high-temperature and high-pressure combustion. Therefore, by making the liner of CMC, the fuel efficiency of the gas turbine can be improved.

[0003] Patent Document 1 discloses a configuration for cooling a liner (thermal shield) disposed on the radially inner side of a shell of a combustor. A cooling chamber is formed between the shell and the liner. An opening for introducing cooling air into the cooling chamber is formed in the shell. Cooling holes for allowing the cooling air to flow out from the cooling chamber into the combustion chamber are formed in the liner.

Prior Art Documents

Patent Documents

[0004]

Patent Document 1

Summary of the Invention

Problems to be Solved by the Invention

[0005] In the configuration of Patent Document 1, in order to sufficiently discharge the cooling air from the cooling holes of the liner to cool the inner peripheral surface of the liner, a sufficient differential pressure is required between the cooling chamber and the combustion chamber. Therefore, it is preferable that no unintended leakage occurs in the cooling air in the cooling chamber. However, when the liner is made of CMC, the degree of freedom in the shape of the liner is reduced, so it is difficult to prevent leakage of the cooling air in the cooling chamber.

[0006] Therefore, the present disclosure aims to seal the cooling chamber in a gas turbine combustor while facilitating the fabrication of the CMC liner. [Means for solving the problem]

[0007] A combustor for a gas turbine according to one aspect of the present disclosure comprises a metal shell, a liner made of a ceramic matrix composite material disposed inside the shell, the liner having an inner surface defining a combustion chamber, an outer surface facing away from the combustion chamber, and at least one cooling hole opening toward the combustion chamber, and a cooling chamber defined between the shell and the liner. The shell includes a cylindrical shell body and a seal wall structure protruding from the inner surface of the shell body toward the liner. [Effects of the Invention]

[0008] According to one aspect of this disclosure, the sealing wall structure protrudes from the metal shell, and the CMC liner does not need to be complicated, thus simplifying the manufacturing of the CMC liner while still allowing the cooling chamber to be sealed by the sealing wall structure. [Brief explanation of the drawing]

[0009] [Figure 1] Figure 1 is a cross-sectional view of the combustor of a gas turbine according to the first embodiment, viewed from the downstream side in the flow direction. [Figure 2] Figure 2 is a cross-sectional view of the combustor shown in Figure 1, taken along line II-II. [Figure 3] Figure 3 is a partial cross-sectional view of the combustor shown in Figure 2, viewed from the circumferential direction. [Figure 4] Figure 4 is a partial cross-sectional view of the combustor in Figure 2, seen from the downstream side in the flow direction. [Figure 5] Figure 5 is a partially unfolded view of the seal wall structure of the combustor shown in Figure 3, viewed from its protruding direction. [Figure 6] Figure 6 is a partial unfolded view of the first modified example of the seal wall structure shown in Figure 5, viewed from its protruding direction. [Figure 7] Figure 7 is a partial unfolded view of a second modified example of the seal wall structure shown in Figure 5, viewed from its protruding direction. [Figure 8] Figure 8 is a partial unfolded view of a third modified example of the seal wall structure shown in Figure 5, viewed from its protruding direction. [Modes for carrying out the invention]

[0010] The embodiments will be described below with reference to the drawings. Note that the axis X of the gas turbine 1 is the same as the axis of the combustor 2. The direction perpendicular to the axis X of the combustor 2 is referred to as the radial direction R of the combustor 2, and the direction extending around the axis X is referred to as the circumferential direction C of the combustor 2. In the flow direction F of the combustion chamber 9, the side with the combustion injection device 10 is referred to as the upstream side, and the side with the outlet 9a is referred to as the downstream side.

[0011] Figure 1 is a schematic diagram of the combustor 2 of the gas turbine 1 according to an embodiment, viewed from the downstream side in the flow direction. As shown in Figure 1, the combustor 2 is the combustor of the gas turbine 1 used as an aircraft engine. The combustor 2 mixes compressed air supplied from the compressor with fuel to produce a mixture, and burns this mixture to generate high-temperature, high-pressure combustion gas. The generated combustion gas is supplied to the turbine to drive the turbine.

[0012] The combustor 2 has a cylindrical shape and is, for example, an annular type formed in an annular shape surrounding the axis X of the gas turbine 1. However, the combustor 2 may be of a type other than an annular type. The combustor 2 includes a casing 3. The casing 3 has an annular outer casing 4 and an annular inner casing 5 arranged concentrically inside the outer casing 4. The outer casing 4 and the inner casing 5 form an annular internal space. The inside of the combustor 2 means the inside of the outer casing 4 in the radial direction R and the outside of the inner casing 5 in the radial direction R.

[0013] A combustion chamber shell 6 is arranged concentrically with the casing 3 within the annular internal space of the casing 3. The shell 6 is made of metal. The shell 6 has an annular outer shell 7 and an annular inner shell 8 arranged concentrically inside the outer shell 7. The shell 6 defines an annular space with the outer shell 7 and the inner shell 8. The annular space enclosed by the outer shell 7 and the inner shell 8 is used as the combustion chamber 9. The inside of the shell 6 means the inside of the radial radius R of the outer shell 7 and the outside of the radial radius R of the inner shell 8.

[0014] Upstream of the combustion chamber 9, a plurality of fuel injectors 10 are arranged in an annular manner along the combustion chamber 9 to inject fuel into the combustion chamber 9. The plurality of fuel injectors 10 are arranged in the circumferential direction C on a virtual circle concentric with the shell 6. The shell 6 is provided with a spark plug 11 that generates a spark to ignite the fuel mixture in the combustion chamber 9 when the gas turbine 1 is started. An outer liner 14 is arranged inside the cylindrical outer shell 7 in the radial direction R. An inner liner 15 is arranged outside the cylindrical inner shell 8 in the radial direction R. The outer liner 14 can be an assembly of a plurality of panels 18 arranged adjacent to each other in the circumferential direction to form a cylinder. The inner liner 15 can also be an assembly of a plurality of panels 19 arranged adjacent to each other in the circumferential direction to form a cylinder. Note that one or both of the outer liner 14 and the inner liner 15 may be a cylindrical body integrally molded into a cylindrical shape without being divided into a plurality of panels arranged in the circumferential direction.

[0015] FIG. 2 is a cross-sectional view taken along line II-II of the combustor 2 in FIG. 1. As shown in FIG. 2, a diffuser 12 for taking in the compressed air generated by the compressor into the casing 3 is provided in the upstream portion of the casing 3. The fuel injection device 10 is supported by a stem 13 fixed to the casing 3. A part of the compressed air taken into the casing 3 is supplied to the fuel injection device 10 for combustion. The remaining part of the compressed air taken into the casing 3 cools the outer surface of the shell 6, and a part of it is supplied into the shell 6 as cooling air from an opening 20c (see FIG. 3) formed in the shell 6.

[0016] Inside the radial direction R of the cylindrical outer shell 7, a cylindrical outer liner 14 is arranged as a liner. Outside the radial direction R of the cylindrical inner shell 8, a cylindrical inner liner 15 is arranged as a liner. The outer liner 14 and the inner liner 15 are separated from the shell 6. The outer liner 14 and the inner liner 15 define a combustion chamber 9. The gas combusted in the combustion chamber 9 is discharged toward the turbine from an exhaust port 9a defined by the downstream ends 14e, 15e in the flow direction F of the outer shell 7 and the inner shell 8.

[0017] The outer liner 14 and the inner liner 15 are made of a ceramic matrix composite (CMC). There are gaps between both ends 14d, 14e in the flow direction F of the outer liner 14 and the outer shell 7, and there are also gaps between both ends 15d, 15e in the flow direction F of the inner liner 15 and the inner shell 8. In FIG. 2, the illustration of a seal wall structure 21 described later is omitted.

[0018] Hereinafter, since the seal wall structures of the outer shell 7 and the inner shell 8 have similar configurations to each other, the outer shell 7 will be described as a representative. FIG. 3 is a partial cross-sectional view seen from the circumferential direction C of the combustor 2 in FIG. 2. FIG. 4 is a partial cross-sectional view seen from the downstream side in the flow direction F of the combustor 2 in FIG. 2. As shown in FIGS. 3 and 4, the outer shell 7 includes a shell body 20 and a seal wall structure 21. The shell body 20 has a cylindrical shape. The shell body 20 has an outer surface 20a facing outward in the radial direction R and an inner surface 20b facing inward in the radial direction R.

[0019] The seal wall structure 21 protrudes from the inner surface 20b of the shell body 20 toward the outer liner 14. That is, the seal wall structure 21 protrudes inward in the radial direction R from the inner surface 20b of the shell body 20. Note that the shell body 2 and the seal wall structure 21 may be formed as an integral body, or a separate seal wall structure 21 may be fixed (for example, joined, fastened, etc.) to the shell body 20.

[0020] The outer liner 14 is attached to the outer shell 7 by a fixture 16 with a gap G between the outer shell 7. That is, a tip 21a of the seal wall structure 21 on the inner side in the radial direction R has a gap G from the outer surface 14a of the outer liner 14. The gap G is larger than the elastic deformation amount of the spring S sandwiched between the nut N and the outer shell 7.

[0021] The outer shell 7 is made of metal and the outer liner 14 is made of CMC, and the thermal expansion coefficients of the outer shell 7 and the outer liner 14 are different from each other. Therefore, the size of the gap G can change according to the temperature of the combustion chamber 9 during the operation of the gas turbine 1. In the present embodiment, the gap G is set to a size such that the tip 21a of the seal wall structure 21 is kept separated from the outer liner 14 throughout the assumed temperature range of the combustion chamber 9 during the operation of the gas turbine 1.

[0022] The outer liner 14 has an outer surface 14a facing outward in the radial direction R and opposite to the outer shell 7, and an inner surface 14b facing inward in the radial direction R and defining the combustion chamber 9. The outer liner 14 has a plurality of cooling holes 14c that open toward the combustion chamber 9. The space between the outer shell 7 and the outer liner 14 is used as a cooling chamber 30. That is, the inner surface 20b of the shell body 20 of the outer shell 7 and the outer surface 14a of the outer liner 14 define the cooling chamber 30.

[0023] The configuration of the mounting fixture 16 is not particularly limited, but as an example, the mounting fixture 16 includes a bolt B, a nut N, and a spring S. For example, the bolt B is attached from the inner surface 14b side of the outer liner 14, and the nut N is attached from the outer surface 20a side of the shell body 20 of the outer shell 7. The spring S is sandwiched between the outer surface 20a of the shell body 20 and the nut N in an elastically deformable state. The spring S allows displacement of the outer liner 14 relative to the outer shell 7 in the radial direction R. The spring S may be, for example, a disc spring.

[0024] The shell body 20 has multiple openings 20c through which compressed air taken into the casing 3 (see Figure 1) is introduced. The compressed air introduced through the openings 20c is supplied to the cooling chamber 30 as cooling air. The cooling air in the cooling chamber 30 cools the outer surface 14a of the outer liner 14. Since the cooling air in the cooling chamber 30 is compressed air supplied from a compressor upstream of the combustor 2, the pressure in the cooling chamber 30 is higher than the pressure in the combustion chamber 9. Therefore, the cooling air in the cooling chamber 30 flows out into the combustion chamber 9 through the cooling holes 14c of the outer liner 14. The cooling air that flows out through the cooling holes 14c covers the inner surface 14b of the outer liner 14, reducing the temperature rise of the inner surface 14b of the outer liner 14 due to the combustion gases in the combustion chamber 9.

[0025] In a cross-section of the cooling chamber 30 viewed from the circumferential direction C of the combustor 2, the seal wall structure 21 has multiple wall portions 22 that divide the cooling chamber 30 into multiple cavities P (see Figure 3). The seal wall structure 21 in this embodiment is a single continuous wall as a whole, but in a cross-sectional view viewed from the circumferential direction C, it has multiple wall portions 22 that are separated from each other. The multiple wall portions 22 protrude from the inner surface 20b of the shell body 20 toward the outer liner 14. The protruding ends of the wall portions 22 are the tips 21a of the seal wall structure 21. In a cross-section of the cooling chamber 30 viewed from the circumferential direction C of the combustor 2, the multiple wall portions 22 are arranged with spacing between them in the direction toward the edge of the outer liner 14 from the cavity P that the cooling holes 14c face among the multiple cavities P. In a cross-section of the cooling chamber 30 viewed from the flow direction F of the combustor 2, the seal wall structure 21 also divides the cooling chamber 30 into multiple cavities P (see Figure 4).

[0026] Figure 5 is a partially unfolded view of the seal wall structure 21 of Figure 3, viewed from its protruding direction. As shown in Figure 5, the seal wall structure 21 has a honeycomb shape when viewed from the direction of protrusion of the seal wall structure 21 from the shell body 20 (radial direction R in this embodiment). That is, the seal wall structure 21 includes a portion extending in the circumferential direction C and a portion extending in the flow direction F. Each of the plurality of cooling holes 14c is contained within one of the plurality of cavities P when viewed from the protruding direction. Each of the openings 20c of the shell body 20 is contained within one of the plurality of cavities P when viewed from the protruding direction. The wall portion 22 of the seal wall structure 21 is interposed between the plurality of cooling holes 14c when viewed from the protruding direction.

[0027] At least one of the multiple openings 20c and at least one of the multiple cooling holes 14c face the same cavity P. Specifically, at least one of the multiple openings 20c faces a cavity P that any of the multiple cooling holes 14c face. Note that the cooling holes 14 and the openings 20c may face different cavities P.

[0028] The cooling air introduced into the cooling chamber 30 from the opening 20c is resisted by the presence of the seal wall structure 21 as it attempts to move toward the end of the cooling chamber 30 in the flow direction F. Specifically, in the region from the cavity P facing the cooling hole 14c to both ends 14d, 14e (see Figure 2) of the outer liner 14 in the flow direction F, multiple narrow gaps G are arranged in the flow direction F between the tip 21a of the seal wall structure 21 and the outer surface 14a of the outer liner 14.

[0029] Therefore, the cooling air in the cavity P facing the cooling hole 14c has difficulty moving towards the edge of the cooling chamber 30 in the flow direction F due to the flow resistance of the narrow gap G. As a result, it is less likely for the cooling air in the cooling chamber 30 to leak beyond the edge of the outer liner 14, the pressure in the cooling chamber 30 is kept sufficiently higher than the pressure in the combustion chamber 9, and the cooling air is properly discharged from the cooling hole 14c.

[0030] A similar sealing wall structure is also provided in the inner shell 8, but since it is the same as the sealing wall structure 21 provided in the outer shell 7, a detailed explanation of the sealing wall structure of the inner shell 8 will be omitted.

[0031] According to the configuration described above, the seal wall structure 21 protrudes from the metal shell 6, and the CMC outer liner 14 and inner liner 15 do not need to be complicated. This makes it easier to manufacture the CMC liners 14 and 15, while still allowing the cooling chamber 30 to be sealed by the seal wall structure 21.

[0032] The mounting fixture 16 includes a spring S that allows displacement of the liner 14 relative to the shell 6 in the radial direction R, and the tip 21a of the seal wall structure 21 in the radial direction R leaves a gap G between it and the outer surface 14a of the liner 14, so that the cooling chamber 30 can be sealed by the seal wall structure 21 while allowing displacement of the liner 14.

[0033] The gap G is set to a size that ensures the tip 21a of the seal wall structure 21 remains away from the outer surface 14a of the liner 14 throughout the entire temperature range of the combustion chamber 9 during operation of the gas turbine 1. Therefore, the seal wall structure 21 does not interfere with the liner 14, and the generation of local stress on the liner 14 can be prevented.

[0034] In a cross-section of the cooling chamber 30 viewed from the circumferential direction C of the combustor 2, the seal wall structure 21 divides the cooling chamber 30 into multiple cavities P, thereby making it difficult for cooling air to leak from the cooling chamber 30.

[0035] The multiple wall portions 22 are arranged at intervals from each other in the cross-section of the cooling chamber 30 viewed from the circumferential direction C, in the direction from the cavity P facing the cooling hole 14c toward the end of the liner 14. As a result, multiple portions with high flow resistance are arranged in a row, making it difficult for cooling air to leak from the cooling chamber 30.

[0036] Since the opening 20c of the shell 7 and the cooling holes 14c of the liner 14 face the same cavity P, cooling air can be discharged from the cooling holes 14c into the combustion chamber 9 with sufficient pressure.

[0037] In a cross-section of the cooling chamber 30 viewed from the flow direction F from upstream to downstream of the combustor 2, the seal wall structure 21 divides the cooling chamber 30 into multiple cavities P, thereby making it difficult for cooling air to leak from the cooling chamber 30.

[0038] The seal wall structure 21 includes a wall portion 22 interposed between the multiple cooling holes 14c when viewed from the radial direction R, so that a differential pressure with the combustion chamber 9 can be secured for each cooling hole 14c.

[0039] Since the seal wall structure 21 has a honeycomb shape when viewed from the radial direction R, it is possible to increase the strength of the shell 6 while improving the sealing performance of the cooling chamber 30.

[0040] The seal wall structure 21 is not limited to the above-described form, and various modifications can be adopted. Figure 6 is a partially unfolded view of the first modified form of the seal wall structure 21 of Figure 5, viewed from its protruding direction. As shown in Figure 6, as the first modified form, the seal structure 121 of the shell 107 may have a grid shape when viewed from the direction in which it protrudes from the shell body 20. Figure 7 is a partially unfolded view of the second modified form of the seal wall structure 21 of Figure 5, viewed from its protruding direction. As shown in Figure 7, as the second modified form, the seal wall structure 221 of the shell 207 may have a plurality of wall portions 222 that extend in the circumferential direction C and are spaced apart from each other in the flow direction F, and may not have wall portions that extend in the flow direction F. That is, the seal wall structure 221 may have a stripe shape when viewed from the direction in which it protrudes from the shell body 20. The cavity P1 separated by the wall portions 222 may extend continuously over the entire circumferential direction C.

[0041] Figure 8 is a partially unfolded view of a third modified example of the seal wall structure 21 of Figure 5, viewed from its protruding direction. As shown in Figure 8, in the third modified example, the seal wall structure 321 of the shell 307 may have a plurality of wall portions 322 that extend in the circumferential direction C and are arranged at unequal intervals in the flow direction F. That is, the seal wall structure 321 may have an unequally spaced stripe shape when viewed from the direction in which it protrudes from the shell body 20. For example, the arrangement pitch of the plurality of wall portions 322 in the flow direction F may be set to be larger in the first region D1 corresponding to the cooling hole 14c than in the second region D2 which is located away from the first region D1 in the flow direction F. That is, the width in the flow direction F of the cavity P2 in the portion corresponding to the cooling hole 14c may be larger than the width in the flow direction F of the cavity P3 in the region located away from the cooling hole 14c in the flow direction F. In addition, the seal wall structure may be labyrinth-shaped, separate from the configurations in Figures 5 to 8.

[0042] As described above, the embodiments have been explained as examples of the technology disclosed in this application. However, the technology in this disclosure is not limited thereto and can be applied to embodiments that have been modified, replaced, added, or omitted as appropriate. Furthermore, it is possible to combine the components described in the embodiments to create new embodiments. For example, some components or methods in one embodiment may be applied to other embodiments, and some components in an embodiment can be separated from other components in that embodiment and extracted as appropriate. In addition, the components described in the attached drawings and detailed description include not only components that are essential for solving the problem, but also components that are not essential for solving the problem, in order to illustrate the technology. [Explanation of Symbols]

[0043] 1 Gas Turbine 2 Combustor 6 Shells 7 Outer shell 8 Inner Shell 9 Combustion chamber 14 Outer Liner 14a Exterior 14b Inner Self 14c cooling hole 15 Inner Liner 16 Mounting hardware Panels 18 and 19 20 Shell body 20a Exterior 20b Inner surface 20c aperture 21,121,221,321 Seal wall structure 21a Tip 22,222,322 Wall 30 Cooling room B bolt C circumferential direction G Gap N Nut P Cavity R radial direction S spring X axis

Claims

1. A fuel injector that injects fuel, A metal shell extending from the fuel injection device, the internal space of which is used as a fuel chamber, A liner made of a ceramic matrix composite material, attached to the shell and disposed inside the shell, having an inner surface defining a combustion chamber, an outer surface facing away from the combustion chamber, and at least one cooling hole opening toward the combustion chamber, A cooling chamber defined between the shell and the liner is provided, The combustor for a gas turbine includes a cylindrical shell body and a sealing wall structure that protrudes from the inner surface of the shell body toward the liner and seals the cooling chamber, In a cross-section of the cooling chamber viewed from the circumferential direction of the combustor, the seal wall structure divides the cooling chamber into a plurality of cavities. The seal wall structure has a plurality of wall portions that protrude from the shell body toward the liner, A gas turbine combustor, wherein the plurality of wall portions are arranged at intervals from each other in the cross-section of the cooling chamber viewed from the circumferential direction of the combustor, in the direction toward the end of the liner from the cavity facing the cooling hole among the plurality of cavities.

2. The gas turbine combustor according to claim 1, wherein the liner includes a plurality of panels arranged adjacent to each other in the circumferential direction of the combustor.

3. The shell further comprises a mounting device for attaching the liner, The mounting fixture includes a spring that allows displacement of the liner relative to the shell in the radial direction of the combustor. The gas turbine combustor according to claim 1 or 2, wherein the tip of the seal wall structure has a gap between it and the outer surface of the liner.

4. The gas turbine combustor according to claim 3, wherein the gap is set such that the tip of the seal wall structure remains separated from the outer surface of the liner throughout the entire temperature range of the combustion chamber during operation of the gas turbine.

5. The shell includes an opening into which compressed air is introduced. The combustor for a gas turbine according to claim 1, wherein the opening and the cooling hole face one of the plurality of cavities.

6. The gas turbine combustor according to any one of claims 1 to 5, wherein, in a cross-section of the cooling chamber viewed from the flow direction from upstream to downstream of the combustor, the seal wall structure divides the cooling chamber into a plurality of cavities.

7. The at least one cooling hole includes a plurality of cooling holes, The combustor for a gas turbine according to any one of claims 1 to 6, wherein the seal wall structure separates the spaces between the plurality of cooling holes when viewed from the direction in which the seal wall structure protrudes.

8. A fuel injection device for injecting fuel, A metal shell extending from the fuel injection device, the internal space of which is used as a fuel chamber, A liner made of a ceramic matrix composite material, attached to the shell and disposed inside the shell, having an inner surface defining a combustion chamber, an outer surface facing away from the combustion chamber, and at least one cooling hole opening toward the combustion chamber, A cooling chamber defined between the shell and the liner is provided, The shell includes a cylindrical shell body and a sealing wall structure that protrudes from the inner surface of the shell body toward the liner and seals the cooling chamber. The seal wall structure, when viewed from the direction in which the seal wall structure protrudes from the shell body, has a honeycomb shape, a grid shape, or a stripe shape, in the combustor of a gas turbine.

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