Transition parts, combustors and gas turbine engines

Abstract

The present invention aims to provide a transition part, a combustor, and a gas turbine engine that can alleviate the temperature gradient and stress of a transition part and extend its service life. A transition part includes a first flow path group formed by arranging a plurality of in-wall flow paths extending inside a plate forming the transition part, and a second flow path group located on the combustor liner side relative to the first flow path group, wherein each in-wall flow path of the first flow path group and the second flow path group has an inlet facing the main flow path of compressed air at an end on one side in the flow direction of the combustion gas, and has an outlet facing the combustion gas flow path at an end on the other side in the flow direction of the combustion gas, and the setting area of ​​the first flow path group and the setting area of ​​the second flow path group partially overlap with a predetermined overlap amount in the flow direction of the combustion gas, and the overlap amount is set larger at a relatively large shape change than at a relatively small shape change of the transition part.

Description

Technical Field

[0001] The invention relates to a transition piece, a combustor and a gas turbine engine. Background Art

[0002] The gas turbine engine burns fuel together with compressed air compressed by a compressor in a combustor, and uses the combustion gas generated thereby to drive the gas turbine. A plurality of combustors are arranged along the circumference of the casing of the gas turbine engine, and in each combustor, the combustion gas is supplied to the gas turbine via a transition part formed into a cylindrical shape by a metal (alloy) plate.

[0003] Among the transition parts through which high-temperature combustion gas undergoing a combustion reaction passes, there are transition parts that have a plurality of in-wall flow paths extending in the flow direction of the combustion gas in the circumferential direction inside a metal plate constituting the transition part (Patent Document 1, etc.). A portion of the compressed air for combustion flows into these in-wall flow paths, cools the transition part, and then is ejected into the combustion gas flow path inside the transition part.

[0004] Prior art literature

[0005] Patent Literature

[0006] Patent Document 1: Japanese Patent Application Publication No. 2014-98352 Summary of the invention

[0007] Problems to be solved by the invention

[0008] The compressed air flowing in the wall flow path of the transition part increases in temperature as it moves along the wall flow path. Therefore, if each wall flow path is long, the cooling effect of the downstream portion of the wall flow path decreases. Therefore, for example, it is considered to divide the area of ​​the transition part into multiple areas along the flow direction of the combustion gas and form a wall flow path in each area. In this way, the length of each wall flow path can be shortened, and the cooling capacity of the transition part as a whole can be improved.

[0009] However, there is a certain temperature difference between the heated compressed air flowing in the wall flow path and ejected to the inside of the transition part and the unheated compressed air flowing into the wall flow path from the outside of the transition part. Therefore, the temperature gradient of the plate constituting the transition part may increase in both the outlet of the wall flow path in the area close to the downstream side (gas turbine side) when viewed along the flow direction of the combustion gas in the transition part and the inlet of the wall flow path on the upstream side (combustor liner side) of the combustion gas of the transition part. In addition, the transition part is a structure in which the cross-sectional shape gradually changes from the circular inlet to the quadrilateral outlet corresponding to the shape of the combustor liner, and the difference in curvature becomes larger depending on the location, and even if the thermal conditions are the same, the stress is different.

[0010] An object of the present invention is to provide a transition part, a combustor, and a gas turbine engine that can alleviate temperature gradients and stresses and extend service life.

[0011] Solutions to Solve Problems

[0012] 14. The burner engine of claim 13 wherein the burner engine is located adjacent to the combustion chamber and wherein the burner engine is located adjacent to the combustion chamber. The first flow path group is located on a side close to the combustor liner, and is formed by arranging a plurality of in-wall flow paths extending from a side close to the gas turbine to a side close to the combustor liner inside the plate along the circumferential direction of the transition part. Each in-wall flow path of the first flow path group and the second flow path group has an inlet facing the main compressed air flow path at an end on one side in the flow direction of the combustion gas, and an outlet facing the combustion gas flow path at an end on the other side in the flow direction of the combustion gas. The setting area of ​​the first flow path group and the setting area of ​​the second flow path group partially overlap with a specified overlap amount in the flow direction of the combustion gas, and the overlap amount is set to be larger at a location where the shape change is relatively large than at a location where the shape change of the transition part is relatively small.

[0013] Effects of the Invention

[0014] According to the present invention, the temperature gradient and stress can be alleviated and the life of the transition part can be prolonged. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] Figure 1 This is a schematic configuration diagram schematically showing an example of a gas turbine plant including a transition part according to the first embodiment of the present invention.

[0016] Figure 2 It is a perspective view of the transition component according to the first embodiment of the present invention.

[0017] Figure 3 This is a schematic diagram of a cross section of the transition part according to the first embodiment of the present invention obtained by cutting it with a plane passing through the center line of the gas turbine.

[0018] Figure 4 is a schematic representation of the Figure 3 A directional view of a portion of the outer peripheral surface of the transition part of the first embodiment of the present invention as viewed in the direction of arrow IV in FIG.

[0019] Figure 5 is based on Figure 4 Sectional view along line VV in the figure.

[0020] Figure 6 is based on Figure 4 Sectional view along line VI-VI.

[0021] Figure 7 is based on Figure 4 Cross-sectional view along line VII-VII.

[0022] Figure 8 It is a schematic diagram showing the region where the in-wall flow path is provided in the back side portion of the transition component according to the first embodiment of the present invention.

[0023] Fig. 9 It is a schematic diagram showing the region where the in-wall flow path is provided in the side portion of the transition component according to the first embodiment of the present invention.

[0024] Fig.10 It is a schematic diagram showing the region where the in-wall flow path is provided in the belly side portion of the transition component according to the first embodiment of the present invention.

[0025] Fig.11 This is a diagram schematically showing a portion of the outer peripheral surface of a transition component according to a second embodiment of the present invention.

[0026] Fig.12 It is a cross-sectional view of a plate material of a transition part according to one structural example of the third embodiment of the present invention.

[0027] Fig.13 It is a cross-sectional view of a plate material of a transition part according to another structural example of the third embodiment of the present invention.

[0028] Description of Reference Numerals

[0029] 10…compressor, 20…combustor, 21…combustor liner, 23…transition part, 23a…combustion gas flow path, 25…plate, 26-28…inner wall flow path, 26a, 27a, 28a…inlet, 26b, 27b, 28b…outlet, 26G…first flow path group, 27G…second flow path group, 28G…third flow path group, 30…gas turbine, 101a…main compressed air path, 100…gas turbine engine, a…compressed air, g…combustion gas, L1, L2…overlap amount. DETAILED DESCRIPTION

[0030] Hereinafter, embodiments of the present invention will be described using the drawings.

[0031] (First Embodiment)

[0032] - Gas turbine engine -

[0033] Figure 1 1 is a schematic structural diagram schematically showing an example of a gas turbine plant equipped with a transition part according to the first embodiment of the present invention. The gas turbine plant shown in the figure is configured to include a gas turbine engine 100 and a load device 200 driven by the gas turbine engine 100. A typical example of the load device 200 is a generator, but a pump or a compressor (a compressor other than the compressor 10 provided in the gas turbine engine 100) may be used as the load device 200 instead of the generator, and the gas turbine engine 100 may be used to drive these compressors and pumps.

[0034] The gas turbine engine 100 is a prime mover for driving the load device 200, and is configured to include a compressor 10, a combustor 20, and a gas turbine 30. The compressor 10 is configured to suck in and compress air, and generate high-temperature and high-pressure compressed air a. The combustor 20 is configured to combust fuel together with the compressed air a discharged from the compressor 10 via the diffuser 11 to generate combustion gas g. The gas turbine 30 is driven by the combustion gas g supplied from the combustor 20 to output rotational power. The shafts of the rotors of the gas turbine 30 and the compressor 10 are connected to each other. A part of the output of the gas turbine 30 is used as the power of the compressor 10, and the rest is used as the power of the load device 200. The combustion gas g that has driven the gas turbine 30 is discharged as exhaust gas through an exhaust chamber (not shown).

[0035] In the present embodiment, the case where the gas turbine engine 100 is a single-shaft type is exemplified, but the gas turbine engine 100 may also be a dual-shaft type. In the case of a dual-shaft gas turbine engine, the gas turbine 30 is composed of a high-pressure turbine and a low-pressure turbine whose rotating shafts are separated from each other, the high-pressure turbine is coaxially connected to the compressor 10, and the low-pressure turbine is coaxially connected to the load device 200.

[0036] - Burner -

[0037] The combustor 20 is installed with respect to the engine room (casing) 101 of the gas turbine engine 100 in a plurality of positions (in the direction of rotation of the gas turbine 30). Figure 1Only one burner 20 is shown as a representative in the figure. Each burner 20 is configured to include a burner liner 21, a burner nozzle 22, and a transition part 23. The burner 20 causes the fuel ejected from the burner nozzle 22 to burn together with the compressed air a compressed by the compressor 10 in the interior (combustion chamber 21a) of the burner liner 21 to generate combustion gas g, and supplies the combustion gas g to the gas turbine 30 via the transition part 23.

[0038] The combustor liner 21 is a cylindrical member having a combustion chamber 21a formed inside, and is provided inside the engine room 101. The combustor liner 21 separates the compressed air a introduced from the compressor 10 into the engine room 101 (in other words, the compressed air main flow path 101a outside the combustor liner 21) from the combustion gas g generated by the combustion chamber 21a (in other words, the combustion chamber 21a inside the combustor liner 21). The transition part 23 is inserted into the end of the combustor liner 21 on the gas turbine side (right side in the figure).

[0039] The burner 22 is a device that injects fuel into the combustion chamber 21a via at least one fuel nozzle 22a and forms and maintains flames inside the combustion chamber 21a. Fuel from a fuel source (eg, a fuel tank) is supplied to the fuel nozzle 22a via a fuel system (fuel piping) 22b.

[0040] Next, the structure of the transition part 23 will be described.

[0041] -Transition parts-

[0042] Figure 2 It is a three-dimensional diagram of the transition part. Figure 3 Schematic diagram of a cross section of a transition part obtained by cutting with a plane passing through the center line of the gas turbine 30. Figure 2 In the figure, the wall flow passages 26 to 28 and the dilution holes 29 (described later) are omitted from the figure.

[0043] The transition part 23 is a component that guides the combustion gas g generated in the combustion chamber 21a to the gas turbine 30, connects the combustor liner 21 to the gas turbine 30, and is formed into a cylindrical shape by a metal (alloy) plate (transition part plate) 25. The transition part 23 separates the outer compressed air main flow path 101a through which the compressed air a supplied from the compressor 10 to the burner nozzle 22 of the combustor 20 flows, and the inner combustion gas flow path 23a through which the combustion gas g supplied from the combustor liner 21 to the gas turbine 30 flows. As mentioned above, the combustor liner 21 is inserted into the end of the transition part 23 on the combustor liner side, that is, the inlet 23b of the combustion gas g. The end of the transition part 23 on the gas turbine side, that is, the outlet 23c of the combustion gas g faces the inlet 30a ( Figure 1The combustion gas g is supplied from the outlet 23 c of the transition part 23 to the annular working fluid flow path facing the stationary blades (not shown) and the moving blades (not shown) in the gas turbine 30 .

[0044] The inlet 23b of the transition part 23 is connected to the cylindrical combustor liner 21 ( Figure 1 ) has an outlet shape corresponding to Figure 2 The transition part 23 is formed in a circular shape as shown. On the other hand, the outlet 23c of the transition part 23 is formed in a quadrilateral shape in accordance with the shape obtained by equally dividing the inlet 30a of the annular working fluid flow path of the gas turbine 30 into the number of burners 20 along the rotation direction of the gas turbine 30. The outlet 23c of each transition part 23 of the plurality of burners 20 provided in the gas turbine engine 100 is connected in the rotation direction of the gas turbine 30 and is annular in shape in accordance with the shape of the inlet 30a of the gas turbine 30. Therefore, the cross-sectional shape of the transition part 23 gradually changes from the circular inlet 23b to the quadrilateral outlet 23c, and the curvature of the plate 25 constituting the transition part 23 is different depending on the location.

[0045] For example, when the transition part 23 is viewed from the back side, the width (dimension in the rotation direction of the gas turbine 30) of the transition part 23 changes as the inlet 23b approaches the outlet 23c, and the width of the outlet 23c becomes wider than the width of the inlet 23b ( Figure 8 On the other hand, when the transition part 23 is viewed from the side, the width of the transition part 23 (the dimension in the radial direction of the gas turbine 30) becomes narrower as the distance from the inlet 23b to the outlet 23c is approached ( Figure 3 ). Thus, the curvature of the plate 25 constituting the transition part 23 varies depending on the position in the flow direction of the combustion gas g and the circumferential position of the transition part 23. The shape of the transition part 23 is smooth for the purpose of guiding the combustion gas g, but this is complicated.

[0046] It should be noted that the back side of the transition part 23 refers to the outer side of the transition part 23 in the radial direction of the gas turbine 30. Therefore, the inner side of the transition part 23 in the radial direction of the gas turbine 30 is the belly side of the transition part 23. In addition, viewing the transition part 23 from the side means viewing the transition part 23 from the direction along the rotation direction of the gas turbine 30.

[0047] In this embodiment, each transition part 23 is as follows Figure 3As shown in the figure, a plurality of wall flow passages 26-28 and a plurality of dilution holes 29 are provided. It should be noted that, regarding the plurality of dilution holes 29, in the example shown in the figure, a structure in which a single annular row of dilution holes is formed along the circumferential direction of the transition part 23 is arranged, but the number of rows may be multiple. An appropriate number of rows is selected from the viewpoint of combustion stability. The wall flow passages 26-28 and the dilution holes 29 are described in sequence below.

[0048] -Intra-wall flow path-

[0049] Figure 4 is a schematic representation of the Figure 3 A partial view of the outer peripheral surface of the transition part observed in the direction of arrow IV, Figure 5 is based on Figure 4 The cross-sectional view along the VV line in the figure, Figure 6 is based on Figure 4 The cross-sectional view along the VI-VI line in the figure, Figure 7 is based on Figure 4 Cross-sectional view along line VII-VII. Figure 8 It is a schematic diagram showing the region where the flow path in the wall of the back side portion of the transition component is provided. Fig. 9 It is a schematic diagram showing the installation area of ​​the in-wall flow path of the side portion of the transition part. Fig.10 It is a schematic diagram showing the region where the in-wall flow path is provided in the belly side portion of the transition part.

[0050] The transition part 23 is provided with a first flow path group 26G, a second flow path group 27G, and a third flow path group 28G. The first flow path group 26G is a collection of flow paths formed into an annular shape by arranging a plurality of in-wall flow paths 26 along the circumferential direction of the transition part 23, and surrounds the transition part 23. Similarly, the second flow path group 27G and the third flow path group 28G are collections of a plurality of in-wall flow paths 27 and 28, respectively, and surround the transition part 23. The first flow path group 26G is located in the downstream side of the flow direction of the combustion gas g in the transition part 23, that is, in the region close to the gas turbine 30. The second flow path group 27G is located in the central region of the flow direction of the combustion gas g in the transition part 23, and is located on the side close to the combustor liner 21 relative to the first flow path group 26G. The third flow path group 28G is the flow path group located on the most upstream side in the flow direction of the combustion gas g, and is located on the side close to the combustor liner 21 relative to the second flow path group 27G. The in-wall flow paths of the first flow path group 26G, the second flow path group 27G, and the third flow path group 28G (in-wall flow paths 26 and 27, and in-wall flow paths 27 and 28) are not in communication with each other but are independent of each other.

[0051] The in-wall flow paths 26-28 extend from the side close to the gas turbine 30 to the side close to the combustor liner 21 inside the plate 25 constituting the transition part 23 (inside the plate thickness), that is, along the flow direction of the combustion gas g. In the first flow path group 26G, the in-wall flow paths 26 adjacent to each other in the circumferential direction of the transition part 23 have the same length. Similarly, in the second flow path group 27G and the third flow path group 28G, the in-wall flow paths 27 and 28 adjacent to each other in the circumferential direction of the transition part 23 have the same length.

[0052] Here, the sheet material 25 constituting the transition part 23 is formed by Figure 5 As shown in the figure, the outer plate 25a facing the compressed air main flow path 101a and the inner plate 25b facing the combustion gas flow path 23a are bonded together. The inner wall flow paths 26-28 form slits on the inner surface of the outer plate 25a, and the inner plate 25b is bonded to the inner surface of the outer plate 25a to block the slits, thereby forming a flow path passing through the inside of the plate 25. It is also possible to set a structure in which a slit is set in the inner plate 25b. In the present embodiment, the inner wall flow paths 26 adjacent to each other in the circumferential direction of the transition part 23 are not connected to each other, but if it is necessary to suppress flow deviation, for example, it is also possible to set a structure in which the adjacent inner wall flow paths 26 are connected to each other at one or more locations. The same is true for the inner wall flow paths 27 and 28.

[0053] In each of the in-wall flow paths 26 of the first flow path group 26G, the inlet 26a and the outlet 26b for the compressed air a are each provided with one ( Figure 3 as well as Figure 4 ). The inlet 26a is provided on the outer plate 25a of the plate 25 and faces the compressed air main flow path 101a, and penetrates the outer plate 25a in the plate thickness direction to connect the compressed air main flow path 101a with the wall inner flow path 26. The outlet 26b is provided on the inner plate 25b of the plate 25 and faces the combustion gas flow path 23a, and penetrates the inner plate 25b in the plate thickness direction to connect the combustion gas flow path 23a with the wall inner flow path 26. During the operation of the gas turbine engine 100, under the action of the differential pressure generated between the inlet 26a and the outlet 26b, a part of the compressed air a flows from the compressed air main flow path 101a into each wall inner flow path 26 as cooling air and is ejected into the combustion gas flow path 23a. In this way, a part of the compressed air a bypasses the burner 22 ( Figure 1 ) flows in the wall flow path 26, thereby cooling the transition part 23.

[0054] It should be noted that the inlet 26a is connected to the end of one side in the flow direction of the combustion gas g in the wall flow passage 26, and the outlet 26b is connected to the end of the other side in the flow direction of the combustion gas g in the wall flow passage 26. Specifically, in each wall flow passage 26, the inlet 26a is provided at the end on the side close to the gas turbine 30, and the outlet 26b is provided at the end on the side close to the combustor liner 21, and the compressed air a is configured to flow in the direction opposite to the flow direction of the combustion gas g in each wall flow passage 26.

[0055] Each of the in-wall flow paths 27 of the second flow path group 27G has the same structure as the in-wall flow paths 26, and each of the inlet 27a and the outlet 27b has one ( Figure 3 as well as Figure 4 Each of the in-wall flow paths 28 of the third flow path group 28G similarly includes one inlet 28a and one outlet 28b ( Figure 3 In the present embodiment, the inlet and outlet arrangements of the in-wall flow passages 27 and 28 are also the same as those of the in-wall flow passage 26, and the compressed air a flows in the in-wall flow passages 27 and 28 in the opposite direction to the combustion gas g.

[0056] like Figure 3-Figure 10 As shown in the figure, the area where the first flow path group 26G is provided overlaps the area where the second flow path group 27G is provided partially by a predetermined overlap amount L1 in the flow direction of the combustion gas g (the direction from the combustor liner 21 toward the gas turbine 30). Specifically, one end of the in-wall flow path 26 of the first flow path group 26G enters between the adjacent in-wall flow paths 27 in the second flow path group 27G, and a band-shaped overlap portion OL1 is formed where the first flow path group 26G and the second flow path group 27G overlap. The overlap portion OL1 exists around the transition part 23 in the circumferential direction.

[0057] Similarly, the area where the second flow path group 27G is provided overlaps the area where the third flow path group 28G is provided partially by a predetermined overlap amount L2 in the flow direction of the combustion gas g. Specifically, one end of the in-wall flow path 27 of the second flow path group 27G enters between the adjacent in-wall flow paths 28 in the third flow path group 28G, and a band-shaped overlap portion OL2 is formed where the second flow path group 27G and the third flow path group 28G overlap. This overlap portion OL2 also exists around the transition part 23 in the circumferential direction.

[0058] It should be noted that the arrangement of the in-wall flow paths 26-28 is relatively dense. In the present embodiment, a structure is exemplified in which the interval D between two in-wall flow paths 26 and 27 adjacent to each other in the circumferential direction of the transition part 23 in the overlapping portion OL1 is set to be equal to or smaller than the diameter W of the circular cross section of each in-wall flow path 26 and 27 ( Figure 4 as well as Figure 5). Similarly, in the overlapping portion OL2, the interval D between two adjacent in-wall flow paths 27 and 28 in the circumferential direction of the transition part 23 is also set to be the same as or smaller than the diameter W of the circular cross section of each in-wall flow path 27 and 28.

[0059] The aforementioned overlap amounts L1 and L2 are set larger at locations where the shape change is relatively large, compared to locations where the shape change is relatively small in the transition part 23. The shape change of the transition part 23 mentioned here is, for example, the curvature of the plate 25 forming the transition part 23, the rate of change of the cross-sectional area of ​​the transition part 23, or the rate of change of the width of the transition part 23. The rate of change of the cross-sectional area of ​​the transition part 23 is the ratio of the change in the area of ​​the cross section of the transition part 23 orthogonal to the center line of the combustion gas flow path 23a, corresponding to the change in the position along the center line of the combustion gas flow path 23a. The rate of change of the width of the transition part 23 is the ratio of the change in the size of the transition part 23 used in the rotation direction or radial direction of the gas turbine 30, corresponding to the change in the position along the center line of the combustion gas flow path 23a. For example, the overlap amount L2 partially differs depending on the circumferential position of the transition part 23. In the present embodiment, the overlap amount L2 is wider ( L1 / L2) on the side and ventral side relative to the dorsal side of the transition part 23. Figure 8 one Fig.10 The degree of difference in the overlap amount L2 caused by the circumferential position corresponds to the difference in shape change of the transition part 23 at each position, for example. Figure 8 one Fig.10 The overlap amount L1 can also be changed in value according to the circumferential position, but in the present embodiment, it is substantially constant regardless of the circumferential position of the transition part 23 .

[0060] In addition, in the present embodiment, when compared at the same circumferential position, the overlap amount L2 between the second flow path group 27G and the third flow path group 28G is partially different from the overlap amount L1 between the first flow path group 26G and the second flow path group 27G. Specifically, at the side and ventral side of the transition part 23, the overlap amount L2 is wider than the overlap amount L1 ( Fig. 9 , Fig.10 The degree of difference between the overlap amounts L1 and L2 corresponds to the difference in shape change of the transition part 23 at each position. Fig. 9 , Fig.10 In the example of , the overlap amounts L1 and L2 are doubled. On the back side of the transition part 23, a difference can also be given to the overlap amounts L1 and L2, but in the present embodiment, the overlap amounts L1 and L2 on the back side are equal.

[0061] -Dilution hole-

[0062] The plurality of dilution holes 29 are small holes that penetrate the plate 25 forming the transition part 23 and connect the compressed air main flow path 101a with the combustion gas flow path 23a, and the opening diameter is equal to or smaller than the outlets 26b-28b of the wall flow paths 26-28. These dilution holes 29 are located in the space between the wall flow paths 27 adjacent to each other in the circumferential direction of the transition part 23 of the second flow path group 27G, and are located closer to the inlet 27a of the wall flow paths 27 of the second flow path group 27G than the outlet 27b of the wall flow paths 27 of the second flow path group 27G. In this way, the dilution holes 29 of the same number as the wall flow paths 26 or 27 are alternately arranged with the wall flow paths 27 along the overlap portion OL1, forming an annular row around the transition part 23.

[0063] The distance d between the outlet 26b of the inner wall flow path 26 and the dilution hole 29 closest to the outlet 26b is the same as or smaller than the diameter W of the circular cross section of the inner wall flow path 26-28 ( Figure 4 The distance d between the outlet 26b and the dilution hole 29 is at least smaller than the maximum value of the overlap amount L1 between the first flow path group 26G and the second flow path group 27G. As an example, the distance d is about 10 mm.

[0064] In addition, the portion of the transition part 23 where the dilution hole 29 is located is a position where the shape change of the transition part 23 is relatively large (for example, larger than the average value of the shape change of each part of the transition part 23). As mentioned above, the shape change means, for example, the curvature of the plate 25 forming the transition part 23, the change rate of the cross-sectional area of ​​the transition part 23, or the change rate of the width of the transition part 23. In the transition part 23 that adopts a dimensional change in the radial direction (or rotation direction) of the gas turbine 30 as it approaches the gas turbine 30, such a portion where the dimensional change is extremely large or its vicinity can be cited as an example of a preferred position of the dilution hole 29.

[0065] -action-

[0066] During the operation of the gas turbine engine 100, air is taken into the compressor 10 and compressed, and then discharged from the compressor 10 to the compressed air main flow path 101a via the diffuser 11 as high-pressure compressed air a. The compressed air a discharged to the compressed air main flow path 101a is supplied to the burner 22 and mixed with the fuel system 22b ( Figure 1 ) are ejected into the combustion chamber 21a and burned together. The high-temperature combustion gas g generated in the combustion chamber 21a is supplied to the gas turbine 30 via the transition part 23, and the gas turbine 30 is driven by the combustion gas g. In addition, the rotation output of the gas turbine 30 drives the load device 200.

[0067] Meanwhile, a part of the compressed air a from the compressed air main flow path 101a to the burner 22 bypasses the burner 22 and flows into the wall inner flow paths 26-28 from the inlets 26a-28a. The compressed air a flowing into the wall inner flow paths 26-28 flows in the wall inner flow paths 26-28 respectively to cool the transition part 23, and is ejected to the combustion gas flow path 23a inside the transition part 23 to merge with the combustion gas g. In addition, another part of the compressed air a of the compressed air main flow path 101a bypasses the burner 22 and is ejected to the inside of the transition part 23 from the dilution holes 29. The compressed air a ejected from the plurality of dilution holes 29 as small holes forms a film cooling film along the inner wall surface of the transition part 23 and flows toward the gas turbine 30, protecting the plate 25 of the transition part 23 from the heat of the combustion gas g.

[0068] -Effect-

[0069] (1) In the present embodiment, a plurality of in-wall flow paths 26-28 are provided in the transition part 23, and the compressed air a is made to flow as cooling air to the plate material 25 constituting the transition part 23, so that the transition part 23 through which the high-temperature combustion gas g passes can be effectively cooled. At this time, the compressed air a is heated while flowing through the in-wall flow paths 26-28. Therefore, if each in-wall flow path is extended from one end to the other end of the transition part 23, each in-wall flow path is long, so that the temperature of the compressed air a rises near the outlet of each in-wall flow path, and the cooling effect is reduced.

[0070] Therefore, in this embodiment, the transition part 23 is divided into a plurality of regions in the flow direction of the combustion gas g, and a mutually independent flow path group is formed in each region, thereby reducing the length of each in-wall flow path. As a result, the temperature of the compressed air a near the outlet of each in-wall flow path 26-28 can be reduced, and the cooling effect of the transition part 23 can be improved.

[0071] However, there is a temperature difference between the compressed air a ejected from the outlet 26b of the wall-inner flow path 26 of the first flow path group 26G and the compressed air a flowing into the inlet 27a of the wall-inner flow path 27 of the second flow path group 27G. Therefore, a temperature gradient is generated in the plate 25 near the outlet 26b and the inlet 27a in the flow direction of the combustion gas g, and the stress concentration caused by the existence of the outlet 26b may increase. Therefore, by partially overlapping the setting area of ​​the first flow path group 26G and the setting area of ​​the second flow path group 27G, it is possible to suppress the increase of stress in the vicinity by ensuring the interval between the outlet 26b and the inlet 27a. The same is true for the overlapping structure of the second flow path group 27G and the third flow path group 28G.

[0072] In particular, in the present embodiment, the overlap amounts L1 and L2 are set larger at the portion where the shape change is relatively large than at the portion where the shape change is relatively small. As a result, the distance between the inlet 27a and 28a for the relatively low-temperature compressed air a to flow in and the outlet 26b and 27b for the relatively high-temperature compressed air a to be ejected can be ensured, particularly at the portion where the shape change is relatively large, which is disadvantageous in terms of strength. Therefore, the temperature gradient and stress concentration around the outlets 26b and 27b of the in-wall flow paths 26 and 27 can be alleviated as a whole, and the life of the transition part 23 can be extended.

[0073] (2) In addition, if the amount of compressed air a supplied to the burner 22 becomes excessive under the operating condition where the amount of fuel supplied is small, there is a possibility that the combustion temperature will decrease and the combustion stability will be impaired. In contrast, in the present embodiment, a part of the compressed air a is supplied to the combustion gas flow path 23a inside the transition part 23 via a plurality of small-diameter dilution holes 29, bypassing the burner 22, and to the region where the combustion reaction has ended, thereby achieving an improvement in the combustion stability.

[0074] However, the transition part 23 is in a thermally severe environment because it allows the high-temperature combustion gas g that has undergone a combustion reaction in the combustion chamber 21a to pass through, and is also susceptible to stress increase in shape because the shape is deformed from a circular cross section to a rectangular cross section. When the dilution hole 29 is provided in the transition part 23, stress may be concentrated around the dilution hole 29.

[0075] In contrast, in the present embodiment, in each space between the adjacent in-wall flow paths 27 in the circumferential direction of the second flow path group 27G, the dilution holes 29 are respectively arranged at a position closer to the inlet 27a than the outlet 27b of the in-wall flow paths 27 of the second flow path group 27G. The plate 25 near the inlet 27a of the in-wall flow path 27 is cooled by the relatively low-temperature compressed air a that has just flowed into the in-wall flow path 27, so the metal temperature and stress are low. By providing the dilution holes 29 at this position, it is possible to suppress the stress concentration near the dilution holes 29 and suppress the strength risk associated with the provision of the dilution holes 29. In addition, the compressed air a flowing through the dilution holes 29 can also contribute to the cooling of the transition part 23.

[0076] (3) If the number of dilution holes 29 is reduced and the opening area is increased accordingly, the dilution holes 29 will interfere with the inner wall flow path 27. However, in the present embodiment, the dilution holes 29 are divided into a number equal to the number of the inner wall flow paths 27, and the opening area of ​​each dilution hole 29 can be suppressed to be small. As a result, the interference between the dilution holes 29 and the inner wall flow paths 27 can be avoided, and the cooling effect intended by the inner wall flow paths 27 is not hindered. In addition, since a plurality of dilution holes 29 with small diameters are formed in an annular row, a film cooling film (cooling air layer) covering the inner wall of the transition part 23 can be formed. The compressed air a that passes through the dilution holes 29 for the purpose of bypassing the burner 22 to improve the combustion stability can also be used for film cooling, which also plays a role in protecting the transition part 23 from the heat of the combustion gas g.

[0077] (4) From the viewpoint of preventing the part of the compressed air a that bypasses the burner 22 and merges with the combustion gas g from affecting the combustion reaction of the flame, it is advantageous to position the dilution hole 29 close to the gas turbine 30. However, if the distance between the gas turbine 30 and the dilution hole 29 is too short, the compressed air a having a large temperature difference with the combustion gas g cannot be fully mixed with the combustion gas g, and the combustion gas g flows into the gas turbine 30 in a state of non-uniform temperature distribution, which may increase the stress of the gas turbine 30.

[0078] On the other hand, in the present embodiment, the compressed air a ejected from the dilution holes 29 provided in the intervals of the in-wall flow path 27 is ensured to have a mixing distance with the combustion gas g by the length of the first flow path group 26G before being supplied to the gas turbine 30. Therefore, the compressed air a ejected from the dilution holes 29 to the combustion gas flow path 23a can be sufficiently mixed with the combustion gas g, and the temperature distribution of the combustion gas g can be made uniform, thereby suppressing an increase in stress of the gas turbine 30.

[0079] (Second Embodiment)

[0080] Fig.11 is a diagram schematically showing a portion of the outer peripheral surface of a transition part according to a second embodiment of the present invention, and is equivalent to the first embodiment. Figure 4 Figure. Fig.11 In the present invention, the same reference numerals as those in the first embodiment are given to the same or corresponding elements as those in the first embodiment, and description thereof will be omitted.

[0081] The present embodiment is different from the first embodiment in that the compressed air a flows in opposite directions in the wall flow paths adjacent to each other in the circumferential direction of the transition part 23 in each of the first flow path group 26G to the third flow path group 28G. Specifically, in the first flow path group 26G, the wall flow paths 26 adjacent to each other in the circumferential direction of the transition part 23 have different inlets 26a and outlets 26b. Fig.11In the first and third wall-inner flow passages 26 from the top, the inlet 26a is located on the gas turbine side and the outlet 26b is located on the combustor liner side, but in the second wall-inner flow passage 26 therebetween, the inlet 26a is located on the combustor liner side and the outlet 26b is located on the gas turbine side. Similarly, in the upper wall-inner flow passage 27, the inlet 27a is located on the gas turbine side and the outlet 27b is located on the combustor liner side, but in the lower wall-inner flow passage 27, the inlet 27a is located on the combustor liner side and the outlet 27b is located on the gas turbine side. Although not shown in the figure, the same is true for the wall-inner flow passages 28 of the third flow passage group 28G.

[0082] In other points, this embodiment has the same structure as the first embodiment.

[0083] In this embodiment, in addition to the same effects as the first embodiment, since the flow directions of the compressed air a in the adjacent wall passages in each of the first passage group 26G to the third passage group 28G are different from each other, it has the advantage that it is difficult to generate a temperature gradient itself, which is a cause of stress.

[0084] (Third Embodiment)

[0085] Fig.12 as well as Fig.13 2 is a cross-sectional view of a plate of a transition part according to a third embodiment of the present invention. Fig.12 as well as Fig.13 In the present invention, the same or corresponding elements as those in the first and second embodiments are denoted by the same reference numerals as those in the first and second embodiments, and description thereof will be omitted.

[0086] The present embodiment is different from the first embodiment and the second embodiment in that, in at least one of the first flow path group 26G to the third flow path group 28G, the flow path cross-sectional area on the side close to the outlet is smaller than the flow path cross-sectional area on the side close to the inlet of the intra-wall flow path. Fig.12 In the example shown, each of the in-wall flow paths 26 of the first flow path group 26G is formed in a conical shape, and the flow path cross-sectional area decreases continuously as it moves from the inlet to the outlet. Fig.13 In the example shown, the inner wall of the flow path 26 in each wall of the first flow path group 26G is provided with a step, and the flow path cross-sectional area of ​​the portion on the outlet side decreases in a step-like manner relative to the portion on the inlet side. Fig.12 as well as Fig.13In the example of FIG. 1 , the wall flow passage 26 of the first flow passage group 26G is taken as an application object, but the same structure can also be applied to the wall flow passage 27 of the second flow passage group 27G and the wall flow passage 28 of the third flow passage group 28G. In at least one of the first flow passage group 26G, the second flow passage group 27G, and the third flow passage group 28G, the outlet of the wall flow passage is narrowed relative to the inlet of the wall flow passage.

[0087] In other points, this embodiment has the same structure as the first embodiment or the second embodiment.

[0088] According to the present embodiment, in addition to the same effects as those of the first embodiment or the second embodiment, the effect of suppressing the temperature gradient of the plate 25 by adjusting the flow rate of the compressed air a flowing in the wall flow path can be expected. That is, the compressed air a flowing in each wall flow path increases in temperature as it approaches the outlet of the wall flow path, but in the present embodiment, the flow rate of the relatively high-temperature compressed air a increases in the wall flow path compared to the relatively low-temperature compressed air a. As a result, the heat transfer rate from the compressed air a flowing in the wall flow path to the plate 25 is adjusted, and the temperature gradient generated in the plate 25 can be suppressed.

[0089] (Variation Example)

[0090] In the first embodiment, the compressed air a flowing in each wall passage is configured to flow in a direction opposite to the flow of the combustion gas g flowing in the combustion gas passage 23a, but the compressed air a may flow in the wall passage in the same direction as the combustion gas g.

[0091] In addition, in the present embodiment, the structure of providing three flow path groups, namely the first flow path group 26G to the third flow path group 28G, in the transition part 23 is exemplified, but the region of the transition part 23 may be divided into two so that there are two flow path groups. The region of the transition part 23 may be divided into four or more so that there are four or more flow path groups. When there are two flow path groups, there is only one overlapping portion, but if the amount of overlap is changed in accordance with the size of the shape change caused by the circumferential position of the transition part 23, the above-mentioned effect (1) can be substantially obtained.

[0092] Although the structure in which the annular row of dilution holes 29 is provided along the overlap portion OL1 is described as an example, the structure in which the annular row of dilution holes 29 is provided along the overlap portion OL2 may be used instead of or in addition to this. Furthermore, the dilution holes 29 may be omitted if not necessary.

[0093] The inlet or outlet of each of the wall flow passages 26-28 may be shared by adjacent wall flow passages, that is, the inlet or outlet may be enlarged or provided as a long hole that is longer in the circumferential direction, and one inlet or outlet may be connected to a plurality of wall flow passages.

[0094] Although the example in which the inner wall flow paths 26 to 28 are formed by bonding the outer plate 25a with slits to the inner plate 25b of the plate 25 has been described, the method of forming the inner wall flow paths 26 to 28 can be changed as appropriate.

Claims

1. A transition part provided in a combustor that burns fuel together with compressed air compressed by a compressor of a gas turbine engine inside a combustor liner and supplies combustion gas to a gas turbine, the transition part connecting the combustor liner to the gas turbine and formed into a cylindrical shape by a plate material, and separating an outer compressed air main flow path for supplying the compressed air from the compressor to the combustor from an inner combustion gas flow path for supplying the combustion gas from the combustor liner to the gas turbine, The transition part is characterized by having: A first flow path group formed by arranging a plurality of in-wall flow paths extending from a side close to the gas turbine to a side close to the combustor liner inside the plate along a circumferential direction of the transition part; and a second flow path group, which is located on a side close to the combustor liner relative to the first flow path group and is formed by arranging a plurality of in-wall flow paths extending from a side close to the gas turbine to a side close to the combustor liner inside the plate along a circumferential direction of the transition part, Each of the in-wall flow paths of the first flow path group and the second flow path group has an inlet facing the main flow path of compressed air at an end on one side in the flow direction of the combustion gas, and has an outlet facing the combustion gas flow path at an end on the other side in the flow direction of the combustion gas. The area where the first flow path group is provided partially overlaps the area where the second flow path group is provided by a predetermined amount in the flow direction of the combustion gas. The overlap amount is set larger at the portion where the shape change is relatively large than at the portion where the shape change is relatively small. The shape change is the curvature of the plate, the change rate of the cross-sectional area of ​​the transition part, or the change rate of the width of the transition part.

2. The transition part according to claim 1, characterized in that The overlap amount partially changes at the circumferential position of the transition part.

3. The transition part according to claim 1, characterized in that The transition part further includes a third flow path group, which is located on a side close to the combustor liner relative to the second flow path group and is formed by arranging a plurality of in-wall flow paths extending from a downstream side close to the gas turbine to an upstream side close to the combustor liner inside the plate along a circumferential direction of the transition part. The area where the second flow path group is provided partially overlaps with the area where the third flow path group is provided by a predetermined amount in the flow direction of the combustion gas. The overlap amount between the second flow path group and the third flow path group is partially different from the overlap amount between the first flow path group and the second flow path group.

4. The transition part according to claim 1, characterized in that The transition part is configured such that the compressed air flows in opposite directions in adjacent in-wall flow paths in each of the first flow path group and the second flow path group.

5. The transition part according to claim 1, characterized in that The flow paths in the walls of the first flow path group and the second flow path group are configured so that the compressed air passes in a direction opposite to the flow direction of the combustion gas. In each of the in-wall flow paths of at least one of the first flow path group and the second flow path group, a flow path cross-sectional area on a side close to the outlet is smaller than a flow path cross-sectional area on a side close to the inlet.

6. A burner, wherein: The combustor comprises the transition piece according to claim 1.

7. A gas turbine engine, wherein: The gas turbine engine comprises: A compressor that compresses air to generate compressed air; The combustor according to claim 6, which generates combustion gas by combusting fuel together with compressed air discharged from the compressor; and A gas turbine is driven by the combustion gas supplied from the combustor.

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