Turbomachine stator vane, blade segment body and gas turbine

Abstract

A turbine vane, a vane segment body, and a gas turbine improve the cooling efficiency in a shroud of a turbine vane. A turbine vane of at least one embodiment of the present disclosure includes: an airfoil portion; and a shroud provided on at least one of one side and the other side in the blade height direction of the airfoil portion. The shroud has a recess formed on a surface on the opposite side of the airfoil portion across a gas passage surface. The recess includes: a first region covered by an impact plate; and a second region not covered by the impact plate. The shroud has: a first opening portion formed in the first region; a first side portion passage formed from the leading edge side to the trailing edge side at a first side end portion on one side in the circumferential direction of the shroud, with one end portion connected to the first opening portion; a second opening portion formed in the second region; and a second side portion passage formed from the leading edge side to the trailing edge side at a second side end portion on the other side in the circumferential direction of the shroud, with one end portion connected to the second opening portion.

Description

Technical Field

[0001] This disclosure relates to turbine stator blades, blade segments, and gas turbines. Background Technology

[0002] For example, in the turbine stator blades of a gas turbine, it is known to have a technique of providing shrouds on the outer and inner sides of the airfoil section in the blade height direction. The surface of the shroud where the airfoil section is located is a gas passage surface and is exposed to high-temperature combustion gases. Therefore, for the shroud, cooling air is used to impinge cool the surface opposite to the airfoil section across the gas passage surface. Furthermore, by circulating the cooling air after impinging cooling of the shroud in a cooling passage provided within the shroud, the cooling air is used efficiently for cooling the shroud (see, for example, Patent Document 1).

[0003] Prior art literature

[0004] Patent documents

[0005] Patent Document 1: Japanese Patent No. 6763636

[0006] From the perspective of improving the efficiency of gas turbines, it is desirable to utilize cooling air more effectively in order to cool the shroud more efficiently. Summary of the Invention

[0007] In view of the above, the purpose of at least one embodiment of the present disclosure is to improve the cooling efficiency in the shroud of the turbine stator blades.

[0008] (1) The turbine stator blade of at least one embodiment of the present disclosure comprises:

[0009] The wing-shaped part; and

[0010] A protective cover, which is disposed on at least one of the two sides of the airfoil in the direction of blade height.

[0011] The protective cover has a recess on the surface opposite to the airfoil, which is separated from the gas passage surface.

[0012] The recess includes: a first region covered by the impact plate; and a second region not covered by the impact plate.

[0013] The protective cover has the following features:

[0014] A first opening is formed in the first region;

[0015] A first side passage is formed from the front edge to the rear edge at one of the first side ends in the circumferential direction of the shield, and one end is connected to the first opening.

[0016] A second opening is formed in the second region; and

[0017] The second side passage is formed from the front edge to the rear edge at the second side end on the other side of the circumference of the shield, and one end is connected to the second opening.

[0018] (2) The blade segment body of at least one embodiment of the present disclosure includes a first segment body and a second segment body of a turbine stator blade having the structure described in (1) above.

[0019] The second side end of the first segment body is bolted to the first side end of the second segment body.

[0020] (3) The gas turbine of at least one embodiment of this disclosure has:

[0021] A compressor, which compresses the incoming air;

[0022] A burner that supplies fuel to compressed air compressed by the compressor and burns the fuel; and

[0023] A turbine that obtains rotational power from combustion gases obtained by combustion in the burner, the turbine having turbine stator blades with the structure described above (1).

[0024] (4) The gas turbine of at least one embodiment of this disclosure has:

[0025] A compressor, which compresses the incoming air;

[0026] A burner that supplies fuel to compressed air compressed by the compressor and burns the fuel; and

[0027] A turbine that obtains rotational power from combustion gases obtained by burning in the burner.

[0028] The turbine has a blade section body with the structure described in (2) above.

[0029] Invention Effects

[0030] According to at least one embodiment of the present disclosure, the cooling efficiency in the shroud of the turbine stator blades can be improved. Attached Figure Description

[0031] Figure 1 This is a schematic diagram showing the overall structure of a gas turbine, an example of rotating machinery.

[0032] Figure 2 This is a cross-sectional view showing the gas flow path of a turbine.

[0033] Figure 3This is a diagram showing the blade section consisting of two turbine stator blades in one embodiment, viewed from the radial inside.

[0034] Figure 4 yes Figure 3 Sectional view in direction IV-IV.

[0035] Figure 5 yes Figure 3 A schematic V-shaped sectional view.

[0036] Explanation of reference numerals in the attached figures:

[0037] 2...protective shield;

[0038] 2a...gas passage surface;

[0039] 10... Gas turbine;

[0040] 11...compressor;

[0041] 12...burner;

[0042] 13... Turbine;

[0043] 21... Turbine stator blades (stator blades);

[0044] 23...wing-shaped portion;

[0045] 23a...front edge;

[0046] 23b... trailing edge;

[0047] 23c...ventral blade surface;

[0048] 23d...dorsal blade surface;

[0049] 25...Inner protective cover;

[0050] 25a...gas passage surface;

[0051] 27...Outer protective cover;

[0052] 27a...gas passage surface;

[0053] 32... Combustion gas flow path;

[0054] 70... Collision plate (impact plate);

[0055] 100...blade segment body;

[0056] 101...section body;

[0057] 101A...First Segment Body;

[0058] 101B...Second segment body;

[0059] 105...Air passage;

[0060] 111...First opening;

[0061] 112...Second opening;

[0062] 131...First lateral passage;

[0063] 132...Second lateral passage;

[0064] 135... Circumferential pathway;

[0065] 151...First side end;

[0066] 152...Second side end;

[0067] 161... Bolt hole;

[0068] 171... Bolt;

[0069] 180... trailing edge end passage;

[0070] 251...First side end;

[0071] 252...Second side end;

[0072] 255...inner region;

[0073] 255a... Bottom surface of the inner region;

[0074] 257... Spatial part (recessed part). Detailed Implementation

[0075] Hereinafter, several embodiments of the present invention will be described with reference to the accompanying drawings. The dimensions, materials, shapes, and relative arrangements of the constituent components described in the embodiments or shown in the drawings are not intended to limit the scope of the present invention, but are merely illustrative examples.

[0076] For example, expressions such as "in a certain direction", "along a certain direction", "parallel", "orthogonal", "center", "concentric" or "coaxial" indicate relative or absolute configurations, not only in a strict sense, but also in a state of relative displacement by angle or distance with tolerance or to the extent that the same function can be obtained.

[0077] For example, expressions such as "same," "equal," and "homogeneous" that indicate the state of equality of things not only indicate a state of strict equality, but also indicate a state of difference in the degree to which the same function can be obtained due to the existence of tolerances.

[0078] For example, the descriptions of shapes such as quadrilaterals and cylindrical shapes not only refer to quadrilaterals and cylindrical shapes in a strict geometric sense, but also to shapes that include concave and convex parts, chamfered parts, etc., within the range where the same effect can be obtained.

[0079] On the other hand, expressions such as "possessing," "containing," "equipped with," "including," or "having" a constituent element are not exclusive expressions that exclude the existence of other constituent elements.

[0080] Figure 1 This is a schematic diagram showing the overall structure of a gas turbine, an example of rotating machinery. Figure 2 This is a cross-sectional view showing the gas flow path of a turbine.

[0081] In this embodiment, such as Figure 1 As shown, the gas turbine 10 is composed of a compressor 11, a burner 12, and a turbine 13 arranged coaxially on a rotor 14, with a generator 15 connected to one end of the rotor 14. It should be noted that in the following description, the direction in which the axis of the rotor 14 extends is defined as the axial direction Da, the circumferential direction around the center of the axis of the rotor 14 is defined as the circumferential direction Dc, and the direction perpendicular to the axis Ax of the rotor 14 is defined as the radial direction Dr. It should be noted that the radial direction Dr is referred to as the blade height direction.

[0082] In compressor 11, air AI drawn in from the air inlet is compressed by multiple stationary and moving vanes to generate high-temperature / high-pressure compressed air AC. Combustor 12 supplies a specified amount of fuel FL to the compressed air AC and combusts it to generate high-temperature / high-pressure combustion gas FG. Turbine 13 drives rotor 14 to rotate by passing the high-temperature / high-pressure combustion gas FG generated by combustor 12 through multiple stationary and moving vanes, thereby driving generator 15 connected to rotor 14.

[0083] In addition, such as Figure 2 As shown, in the turbine 13, the turbine stator blade 21 is configured such that the hub side of the airfoil 23 is fixed to the inner shroud 25, and the front end side is fixed to the outer shroud 27. The turbine moving blade 41 is configured such that the base end of the airfoil 43 is fixed to the platform 45. Furthermore, the outer shroud 27 and the dividing ring 51 disposed on the front end side of the moving blade 41 are supported in the turbine housing 30 via a heat insulation ring 53, and the inner shroud 25 is supported by a support ring 31. Therefore, the combustion gas flow path 32 through which the combustion gas FG passes is formed along the axial direction Da as a space surrounded by the inner shroud 25, the outer shroud 27, the platform 45, and the dividing ring 51.

[0084] Figure 3This is a view of the blade section 100, which is composed of two turbine stator blades 21 according to one embodiment, as viewed from the radial inside of Dr.

[0085] Figure 4 yes Figure 3 Sectional view in direction IV-IV.

[0086] Figure 5 yes Figure 3 A schematic V-direction sectional view shows a turbine stator blade 21.

[0087] In one embodiment, the turbine stator 21 comprises a section body 101 with an airfoil 23 disposed relative to an outer shroud 27 and an inner shroud 25. In another embodiment, two section bodies 101 are connected by bolts to form a blade section body 100.

[0088] It should be noted that, as Figure 3 As shown, regarding the airfoil portion 23, an airfoil portion 23 is formed by connecting and integrating the ventral blade surface 23c formed by a concave surface as a pressure surface and a dorsal blade surface 23d formed by a convex surface as a negative pressure surface, and the ventral blade surface 23c and the dorsal blade surface 23d at the leading edge 23a on the upstream side and the trailing edge 23b on the downstream side of the axial direction Da.

[0089] It should be noted that the inner shield 25 and the outer shield 27 function as gas passage surface forming components. Gas passage surface forming components refer to components that divide the combustion gas flow path 32 and have gas passage surfaces 2a (25a, 27a) that contact the combustion gas FG. When there is no need to specifically distinguish between the inner shield 25 and the outer shield 27, they are sometimes simply referred to as shield 2.

[0090] (Regarding blade segment 100)

[0091] As described above, one embodiment of the blade segment body 100 includes two bolted segments 101. In the following description, for ease of explanation, the segment body 101 configured such that its dorsal blade surface 23d faces the ventral blade surface 23c of the opposite segment body 101 is designated as the first segment body 101A, and the segment body 101 configured such that its ventral blade surface 23c faces the dorsal blade surface 23d of the opposite segment body 101 is designated as the second segment body 101B. Figure 3 In the diagram, the segment 101 on the left is the first segment 101A, and the segment on the right is the second segment 101B.

[0092] therefore, Figure 4 and Figure 5 The figures shown are of the first segment 101A.

[0093] The first section body 101A includes: a first airfoil 23A; a first outer shield 27A disposed on the front end side of the first airfoil 23A, i.e. the outer end in the blade height direction (the outer end in the radial direction Dr); and a first inner shield 25A disposed on the hub side of the first airfoil 23A, i.e. the inner end in the blade height direction (the inner end in the radial direction Dr).

[0094] Similarly, the second segment body 101B includes: a second airfoil 23B; a second outer protective cover (illustrated) disposed on the front end side of the second airfoil 23B; and a second inner protective cover 25B disposed on the hub side of the second airfoil 23B.

[0095] It should be noted that in the following description, when describing the common structure of the first segment body 101A and the second segment body 101B, and when there is no need to specifically distinguish between the first segment body 101A and the second segment body 101B, and when referring to the first segment body 101A and the second segment body 101B collectively, the last letter of the reference numerals will be omitted.

[0096] like Figure 3 As shown, in several embodiments, the stator 21, for example, has a leading-edge retainer 61 and a trailing-edge retainer 63 in the inner shroud 25, which face the gas passage surface 25a (see reference). Figure 4 It extends radially inward on the side opposite to the airfoil 23. A leading-edge retainer 61 is formed on the leading edge 23a side of the airfoil 23, and a trailing-edge retainer 63 is formed on the trailing edge 23b side of the airfoil 23. The leading-edge retainer 61 and the trailing-edge retainer 63 are connected via a support ring 31 (see reference). Figure 2 Installed in the machine room 30.

[0097] The passage of cooling air in the inner shroud 25 of the stator blade 21 in several embodiments will be described below.

[0098] In one embodiment, the inner shroud 25 of the stator blade 21 has an inner region 255 on the surface opposite to the gas passage surface 25a, i.e., radially inside Dr, which serves as a space for storing cooling air CA supplied from the outside. The inner region 255 is a region surrounded by the periphery of the inner shroud 25, i.e., the first end portion 251 on the ventral blade surface 23c side, the second end portion 252 on the dorsal blade surface 23d side, the leading edge end portion 253 on the axial leading edge 23a side, and the trailing edge end portion 254 on the trailing edge 23b side of the two ends forming the circumferential Dc of the inner shroud 25. It forms a space portion 257 and an impact space 256 (described later) that are recessed in the radially inside Dr direction. The bottom surface 255a of the inner region 255 forms the surface radially inside Dr of the gas passage surface 25a. That is, the space 257 and the impact space 256 are formed by the bottom surface 255a of the inner region and the outer wall portion extending from the bottom surface 255a of the inner region along the blade height direction (radial Dr), namely the first side end 251, the second side end 252, the leading edge end 253 and the trailing edge end 254.

[0099] In one embodiment of the gas turbine 10, cooling air CA is supplied from the radially outer side of the stator blade 21 via a through hole 24 that penetrates the airfoil 23 in the blade height direction to the space 257.

[0100] In the inner region 255, a collision plate (impact plate) 70 with multiple through holes 71 is arranged such that it covers the area located on the ventral side of the airfoil 23 when viewed from the blade height direction within the bottom surface 255a of the inner region (first region R1). Figure 5 It should be noted that, in Figure 3 In the text, the first region R1 is represented by a shaded line, and the description of the collision plate 70 is omitted.

[0101] The first region R1 within the inner region 255 of the space 257 is divided by the impact plate 70 into the space 257 on the radially inner side (Dr) and the impact space 256 on the radially outer side (Dr). The space 257 and the impact space 256 are connected via the through hole 71 of the impact plate 70.

[0102] When viewed from the blade height direction within the inner region 255 of the forming space 257, the region located on the back side of the airfoil 23 (second region R2) is not covered by the collision plate 70.

[0103] A portion of the cooling air CA supplied to the space 257 is supplied to the impact space 256 through the through hole 71 to perform impact cooling (collision cooling) on ​​the bottom surface 255a of the inner region of the first region R1. By performing impact cooling on the bottom surface 255a of the inner region of the first region R1, overheating caused by combustion gas FG in the gas passage surface 25a located on the ventral side of the airfoil 23 is suppressed.

[0104] Cooling air CA, after being subjected to impact cooling on the bottom surface 255a of the inner region of the first region R1, is supplied to the first side passage 131 described later.

[0105] In addition, a portion of the cooling air CA supplied to the space section 257 cools the bottom surface 255a of the inner region of the second region R2. By cooling the bottom surface 255a of the inner region of the second region R2, overheating caused by combustion gas FG in the gas passage surface 25a located on the back side of the airfoil 23 is suppressed.

[0106] Cooling air CA, after cooling the bottom surface 255a of the inner region of the second region R2, is supplied to the second side passage 132 described later.

[0107] A portion of the cooling air CA supplied to the space section 257 flows into the leading edge side flow path 191, which is formed at intervals along the circumferential direction Dc and extends along the axial direction Da at the leading edge end 253, mainly to cool the leading edge end 253. After flowing through the leading edge end 253, the cooling air CA is discharged into the combustion gas flow path 32 through the opening in the gas passage surface 25a of the leading edge end 253.

[0108] In several embodiments, the stator blade 21 includes an air passage 105 for allowing cooling air CA supplied to the aforementioned space 257 to circulate in the trailing edge end 254 of the inner shroud 25. The air passage 105 has a first side passage 131, a second side passage 132, a circumferential passage 135, and a trailing edge end passage 180.

[0109] The first side passage 131 is an air passage formed from the leading edge 23a side to the trailing edge 23b side of the first side end 251 of the inner shield 25. One end of the leading edge 23a side of the first side passage 131 is connected to the impact space 256 via the first opening 111 formed in the first region R1 at the first side end 251.

[0110] The other end of the rear edge 23b of the first lateral passage 131 is connected to one end of the circumferential passage Dc of the circumferential passage 135, which will be described later.

[0111] The second side passage 132 is an air passage formed from the front edge 23a side to the rear edge 23b side of the second side end 252 of the inner cover 25. One end of the second side passage 132 on the front edge 23a side is connected to the space portion 257 via the second opening 112 formed in the second region R2 at the second side end 252.

[0112] The other end of the rear edge 23b of the second side passage 132 is connected to the other end of the circumferential passage Dc of the circumferential passage 135.

[0113] The circumferential passage 135 is a cooling passage that extends circumferentially along Dc at a position relative to the rear edge 23b of the space portion 257, and connects the other end of the rear edge 23b of the first side passage 131 to the other end of the rear edge 23b of the second side passage 132 as described above.

[0114] The trailing edge end passage 180 has multiple cooling passages spaced apart on the circumferential Dc of the trailing edge end 254. The upstream end 180a is connected to the circumferential passage 135, and the downstream end 180b is open on the trailing edge end face 25c of the inner shield 25.

[0115] In one embodiment of the air passage 105 configured in this way, a portion of the cooling air CA of the space 257 is supplied from the first opening 111 to the first side passage 131 via the impact space 256, and a portion of the cooling air CA of the space 257 is supplied from the second opening 112 to the second side passage 132.

[0116] Cooling air CA supplied to the first side passage 131 flows from the leading edge 23a side to the trailing edge 23b side in the first side passage 131, mainly cooling the first side end 251.

[0117] Cooling air CA supplied to the second side passage 132 flows from the leading edge 23a side to the trailing edge 23b side in the second side passage 132, mainly cooling the second side end 252.

[0118] It should be noted that the first opening 111 and the second opening 112 are preferably located near the leading edge end 253 so that the cooling air CA flowing into the first side passage 131 and the second side passage 132 flows along the axial direction Da in a longer range as much as possible.

[0119] Cooling air CA flowing toward the trailing edge 23b in the first side passage 131 and the second side passage 132 flows into the circumferential passage 135, and then flows into multiple trailing edge end passages 180 via the circumferential passage 135. The cooling air CA flowing into the multiple trailing edge end passages 180 flows from the upstream end 180a of the trailing edge end passage 180 toward the trailing edge end face 25c of the inner shroud 25, primarily cooling the trailing edge end 254. The cooling air CA is discharged into the combustion gases from the trailing edge end face 25c.

[0120] (Regarding the bolted connection between the first section body 101A and the second section body 101B)

[0121] In the first section body 101A of the blade section body 100 in one embodiment, bolt holes 161 are formed in the second side end 252 on the back blade surface 23d side in the first inner shield 25A and the side end 152 on the back blade surface 23d side in the first outer shield 27A, i.e., the second side end 152. The bolt holes 161 are formed in the circumferential direction Dc through the first inner shield 25A and the second side end 252 on the back blade surface 23d side.

[0122] In the first segment 101A of the blade segment body 100 in one embodiment, as Figure 4 As shown, a bolt hole 161 is formed at the second side end 252 of the first inner cover 25A, but multiple bolt holes 161 may also be formed at intervals along the axial direction Da.

[0123] In one embodiment of the blade section body 100, in the first section body 101A, a plurality of bolt holes 161 are preferably formed at intervals along the axial direction Da at the second side end 152 of the first outer outer shield 27A. Figure 4 In the example shown, there are three bolt holes 161, but there can also be two or fewer, or even more than four.

[0124] In the second section body 101B of the blade section body 100 in one embodiment, a bolt hole (not shown) is formed in the first side end 251 on the ventral blade surface 23c side of the second inner shield 25B and the side end (not shown) on the ventral blade surface 23c side of the second outer shield.

[0125] In the second section 101B of the blade section body 100 in one embodiment, a bolt hole is preferably formed at the first side end 251 of the second inner shroud 25B, but multiple bolt holes may also be formed at intervals along the axial direction Da.

[0126] In the second section 101B of the blade section body 100 in one embodiment, a plurality of bolt holes are preferably formed at intervals along the axial direction Da at the first side end 151 of the second outer shield 27B. For example, similar to the first outer shield 27A, the bolt holes are preferably three, but there may be two or less, or even four or more.

[0127] The bolt hole 161 of the first section body 101A and the bolt hole (not shown) of the second section body 101B are positioned such that the bolt 171 can pass through the bolt hole 161 and the bolt hole (not shown).

[0128] In one embodiment of the blade section body 100, the first section body 101A and the second section body 101B are bolted together by passing a bolt 171 through a bolt hole 161 and a bolt hole (not shown) and fitting a nut 172.

[0129] It should be noted that, in the turbine 13 of the gas turbine 10 according to one embodiment, a plurality of blade section bodies 100 are arranged along the circumferential direction Dc. Adjacent blade section bodies 100 on the circumferential direction Dc are not bolted together. A sealing plate (not shown) is arranged between adjacent blade section bodies 100 on the circumferential direction Dc to prevent cooling air CA from leaking between adjacent blade section bodies 100 on the circumferential direction Dc.

[0130] In one embodiment of the turbine stator blade 21, the cooling air CA after cooling the first region R1 and the second region R2 is preferably further utilized for cooling the inner shroud 25. In particular, the temperature of the cooling air CA after cooling the region that is not subjected to impact cooling, i.e., the second region R2, is relatively low, so it is preferable to utilize it effectively for cooling the inner shroud 25.

[0131] Therefore, the turbine stator 21 is configured as follows in one embodiment. Specifically, in the turbine stator 21 of one embodiment, the inner shroud 25 has: a first opening 111 formed in a first region R1; and a first side passage 131, which extends from the leading edge 23a side to the trailing edge 23b side at a first side end 251 on one side of the circumferential direction Dc of the inner shroud 25, with one end connected to the first opening 111. The inner shroud 25 also has: a second opening 112 formed in a second region R2; and a second side passage 132, which extends from the leading edge 23a side to the trailing edge 23b side at a second side end 252 on the other side of the circumferential direction Dc of the inner shroud 25, with one end connected to the second opening 112.

[0132] Therefore, the cooling air CA after cooling the first region R1 which is subjected to impact cooling, and the cooling air CA after cooling the second region R2 which is not subjected to impact cooling, can be further utilized for cooling the inner shield 25, thereby improving the cooling efficiency of the inner shield 25.

[0133] One embodiment of the gas turbine 10 includes: a compressor 11 that compresses incoming air AI; a combustor 12 that supplies fuel FL to the compressed air AC compressed by the compressor 11 and causes it to burn; and a turbine 13 that obtains rotational power from combustion gas FG obtained by combustion in the combustor 12. The turbine 13 includes the blade section body 100 (i.e., turbine stator blade 21) of the above-described embodiment.

[0134] Therefore, the cooling air CA after cooling the first region R1 which undergoes impact cooling, and the cooling air CA after cooling the second region R2 which does not undergo impact cooling, can be further utilized for cooling the inner shroud 25, thereby improving the cooling efficiency of the inner shroud 25. This also helps to suppress, for example, excessive cooling and excessive use of cooling air CA, thereby improving the efficiency of the gas turbine 10.

[0135] In one embodiment of the turbine stator blade 21, the inner shroud 25 preferably has a circumferential passage 135 that extends circumferentially Dc at a position relative to the trailing edge 23b of the first region R1. Preferably, the other end of the first side passage 131 is connected to one end of the circumferential passage 135, and the other end of the second side passage 132 is connected to the other end of the circumferential passage 135.

[0136] In the second side passage 132, cooling air CA flows in from the second region R2 not covered by the collision plate 70. Therefore, compared to the first side passage 131 where cooling air CA flows in from the first region R1 covered by the collision plate 70, the pressure of the cooling air CA tends to be higher. Consequently, the flow rate of cooling air CA flowing in the second side passage 132 tends to be higher. From the viewpoint of efficiently using the cooling air CA, it is preferable to further use the cooling air CA flowing in the second side passage 132 for cooling the inner shield 25.

[0137] According to one embodiment of the turbine stator 21, the cooling air CA flowing in the second side passage 132 can be further used to cool the area around the circumferential passage 135. Furthermore, according to another embodiment of the turbine stator 21, by further providing a cooling flow path (tail-edge end passage 180) on the downstream side of the circumferential passage 135, the cooling air CA flowing in the first side passage 131 and the second side passage 132 can circulate in this cooling flow path (tail-edge end passage 180) to cool the inner shroud 25, thus enabling efficient use of the cooling air CA.

[0138] In one embodiment of the turbine stator blade 21, it is preferable that the inlet area S2 of the second opening 112 is smaller than the inlet area S1 of the first opening 111.

[0139] As described above, in the second side passage 132, cooling air CA flows in from the second region R2 not covered by the collision plate 70. Therefore, compared to the first side passage 131 where cooling air CA flows in from the first region R1 covered by the collision plate 70, the pressure of the cooling air CA tends to be higher. Consequently, for the cooling air CA flowing into the circumferential passage 135, the cooling air CA from the second side passage 132 is more likely to be more abundant than the cooling air CA from the first side passage 131.

[0140] According to one embodiment, the turbine stator 21 can suppress the flow rate of cooling air CA flowing into the second side passage 132 at the second opening 112, thus suppressing the flow rate of cooling air CA flowing into the second side passage 132 to an appropriate flow rate.

[0141] In one embodiment of the turbine stator blade 21, it is preferred that the inner shroud 25 has a trailing edge end passage 180, a plurality of which are arranged circumferentially Dc on the trailing edge 23b side, and the upstream end 180a is connected to the circumferential passage 135, and the downstream end 180b opens at the trailing edge end face 25c of the inner shroud 25.

[0142] Therefore, the cooling air CA that flows in the first side passage 131 and the second side passage 132 can flow in the multiple trailing edge passages 180 to cool the area (trailing edge end 254) on the trailing edge 23b side of the inner shroud 25, and the cooling air CA can be used efficiently.

[0143] In one embodiment of the turbine stator blade 21, it is preferable that the first region R1 includes a region located on the ventral side of the airfoil 23 when viewed from the blade height direction.

[0144] The combustion gas FG, which is the working fluid of the turbine 13, is a relatively high-temperature fluid. Therefore, in the region near the ventral side of the airfoil 23 within the combustion gas flow path 32, the flow velocity of the combustion gas flow path 32 tends to be higher than that in the region near the dorsal side. Consequently, the temperature of the inner shield 25 also tends to be higher in the region located closer to the ventral side of the airfoil 23 than in the region located closer to the ventral side of the airfoil 23.

[0145] According to one embodiment, the turbine stator 21 can perform shock cooling on areas where the temperature tends to become higher, thus effectively suppressing the temperature rise of the inner shroud 25.

[0146] In one embodiment of the turbine stator blade 21, it is preferable that the structure associated with the air passage 105 described above is provided in the inner shroud 25.

[0147] According to one embodiment of the turbine stator blade 21, in the inner shroud 25, which is smaller in size than the outer shroud 27 and has a smaller area where the collision plate 70 is disposed, the cooling air CA after cooling the first region R1 which is subjected to impact cooling and the cooling air CA after cooling the second region R2 which is not subjected to impact cooling can be further utilized for cooling the inner shroud 25.

[0148] One embodiment of the blade section body 100 includes a first section body 101A and a second section body 101B, which includes a turbine stator blade 21 of one embodiment. The second side end 252 of the first section body 101A and the first side end 251 of the second section body 101B are bolted together. Therefore, the second side end 252 of the first section body 101A and the first side end 251 of the second section body 101B have bolts 171 for bolting together, making cooling relatively difficult.

[0149] According to one embodiment of the blade section body 100, a second side passage 132 is formed at the second side end 252 of the first section body 101A and a first side passage 131 is formed at the first side end 251 of the second section body 101B. Therefore, the second side end 252 of the first section body 101A and the first side end 251 of the second section body 101B, which are relatively difficult to cool, can be cooled.

[0150] In one embodiment of the blade section body 100, it is preferable that the first side passage 131, when viewed from the blade height direction of the airfoil 23, overlaps with the portion of the first side end 251 that is connected by bolts 171 and nuts 172 as fasteners in the axial direction Da. It is also preferable that the second side passage 132, when viewed from the blade height direction, overlaps with the portion of the second side end 252 that is connected by bolts 171 and nuts 172 as fasteners in the axial direction Da.

[0151] Therefore, the second side end 252 of the first segment body 101A and the first side end 251 of the second segment body 101B, which are relatively difficult to cool, can be cooled efficiently.

[0152] This disclosure is not limited to the above-described embodiments, but also includes modifications to the above-described embodiments and appropriate combinations thereof.

[0153] For example, the structure related to the air passage 105 in the inner shroud 25 of the turbine stator blade 21 in the above-described embodiment can also be provided in the outer shroud 27.

[0154] In the turbine stator blade 21 of the above-described embodiment, the cooling air CA flowing into the first side passage 131 and the second side passage 132 flows from the leading edge 23a side toward the trailing edge 23b side, but it is also possible for the cooling air CA flowing into the first side passage 131 and the second side passage 132 to flow from the trailing edge 23b side toward the leading edge 23a side.

[0155] In this case, the first opening 111 and the second opening 112 may also be provided near the rear edge end 254 so that the cooling air CA flowing into the first side passage 131 and the second side passage 132 flows along the axial direction Da for as long as possible.

[0156] Alternatively, in this case, it is preferable to position the circumferential passage 135 at a position closer to the leading edge 23a than the space portion 257, and connect the circumferential passage 135 to the first side passage 131 and the second side passage 132.

[0157] In the turbine stator blade 21 of the above embodiment, the flow rate of cooling air CA flowing into the first side passage 131 and the flow rate of cooling air CA flowing into the second side passage 132 are appropriately adjusted by appropriately setting the inlet area S1 of the first opening 111 and the inlet area S2 of the second opening 112.

[0158] However, it is also possible to appropriately adjust the flow rate of cooling air CA flowing into the first side passage 131 and the flow rate of cooling air CA flowing into the second side passage 132 by appropriately setting the inlet area S1 of the first opening 111 and the inlet area S2 of the second opening 112, while appropriately setting the inlet area S1 of the first opening 111 and the inlet area S2 of the second opening 112.

[0159] The contents described in the above embodiments shall be understood as follows.

[0160] (1) The turbine stator blade 21 of at least one embodiment of the present disclosure includes: an airfoil 23; and a shroud 2 disposed on at least one side and the other side of the airfoil 23 in the blade height direction (radial Dr). The shroud 2 (inner shroud 25) has a recess (space 257) formed on the surface opposite to the airfoil 23 across the gas passage surface 2a. The recess (space 257) includes: a first region R1 covered by an impact plate (collision plate 70); and a second region R2 not covered by the impact plate (collision plate 70). The shroud 2 (inner shroud 25) has: a first opening 111 formed in the first region R1; and a first side passage 131 formed from the leading edge 23a side to the trailing edge 23b side at a first side end 251 on one side of the shroud 2 in the circumferential direction Dc, and one end is connected to the first opening 111. The shield 2 (inner shield 25) has: a second opening 112 formed in the second region R2; and a second side passage 132 formed from the leading edge 23a side to the trailing edge 23b side at the second side end 252 on the other side of the circumferential direction Dc of the shield 2, and one end is connected to the second opening 112.

[0161] According to the structure described in (1), the cooling air CA after cooling the first region R1 which is subjected to impact cooling and the cooling air CA after cooling the second region R2 which is not subjected to impact cooling can be further utilized for cooling the shield 2 (inner shield 25), thereby improving the cooling efficiency of the shield 2 (inner shield 25).

[0162] (2) In several embodiments, based on the structure described in (1) above, it is preferable that the cover 2 (inner cover 25) has a circumferential passage 135 extending in the circumferential direction Dc at a position relative to the rear edge 23b of the first region R1. The other end of the first side passage 131 is connected to one end of the circumferential passage 135, and the other end of the second side passage 132 is connected to the other end of the circumferential passage 135.

[0163] According to the structure described in (2) above, the cooling air CA flowing in the second side passage 132 can be used to further cool the area around the circumferential passage 135. In addition, according to the structure described in (2) above, by further providing a cooling flow path (rear edge end passage 180) on the downstream side of the circumferential passage 135, the cooling air CA flowing in the first side passage 131 and the second side passage 132 can circulate in this cooling flow path (rear edge end passage 180) to cool the shield 2 (inner shield 25), thus enabling efficient use of the cooling air CA.

[0164] (3) In several embodiments, based on the structure of (2) above, it is preferable that the inlet area S2 of the second opening 112 is smaller than the inlet area S1 of the first opening 111.

[0165] As described above, in the second side passage 132, cooling air CA flows in from the second region R2 not covered by the impact plate (collision plate 70). Therefore, compared to the first side passage 131 where cooling air CA flows in from the first region R1 covered by the impact plate (collision plate 70), the pressure of the cooling air CA tends to be higher. Consequently, for the cooling air CA flowing into the circumferential passage 135, the cooling air CA from the second side passage 132 is more likely to be more abundant than the cooling air CA from the first side passage 131.

[0166] According to the structure described above (3), the flow rate of cooling air CA flowing into the second side passage 132 can be suppressed at the second opening 112, and thus the flow rate of cooling air CA flowing into the second side passage 132 can be suppressed to an appropriate flow rate.

[0167] (4) In several embodiments, based on the structure of (2) or (3) above, it is preferred that the cover 2 (inner cover 25) has a trailing edge end passage 180, which has a plurality of trailing edge end passages 180 arranged circumferentially Dc on the trailing edge 23b side, and one end (upstream end 180a) is connected to the circumferential passage 135, and the other end (downstream end 180b) is open at the trailing edge end face 25c of the cover 2 (inner cover 25).

[0168] According to the structure described above (4), the cooling air CA that flows in the first side passage 131 and the second side passage 132 can flow in the multiple trailing edge passages 180 to cool the area on the trailing edge 23b side of the shield 2 (inner shield 25), and the cooling air CA can be used efficiently.

[0169] (5) In several embodiments, based on any of the structures in (1) to (4) above, it is preferred that the first region R1 includes a region located on the ventral side of the airfoil 23 when viewed from the blade height direction.

[0170] If the working fluid (combustion gas FG) of the turbine 13 is a high-temperature fluid, then in the region near the ventral side of the airfoil 23 within the flow path (combustion gas flow path 32) of the working fluid (combustion gas FG), the flow velocity of the working fluid (combustion gas FG) tends to be higher than that in the region near the dorsal side. Therefore, the temperature of the shield 2 (inner shield 25) also tends to be higher in the region located near the ventral side of the airfoil 23 than in the region located near the ventral side of the airfoil 23.

[0171] According to the structure described above (5), it is possible to perform shock cooling on areas where the temperature tends to become higher, thus effectively suppressing the temperature rise of the shield 2 (inner shield 25).

[0172] (6) In several embodiments, based on any of the structures in (1) to (5) above, it is preferred that the shield 2 (inner shield 25) is disposed on the inner side of the airfoil 23 in the blade height direction.

[0173] According to the structure described above (6), in the inner shield 25, which is smaller in size than the outer shield 27 and has a smaller area where the impact plate (collision plate 70) is disposed, the cooling air CA after cooling the first region R1 which is subjected to impact cooling and the cooling air CA after cooling the second region R2 which is not subjected to impact cooling can be further utilized for cooling the inner shield 25.

[0174] (7) The blade section body of at least one embodiment of the present disclosure includes a first section body 101A and a second section body 101B of a turbine stator blade 21 including any of the structures described in (1) to (6) above. The second side end 252 of the first section body 101A and the first side end 251 of the second section body 101B are bolted together.

[0175] The second side end 252 of the first section body 101A and the first side end 251 of the second section body 101B have bolts 171 for bolting, making them relatively difficult to cool.

[0176] According to the structure described above (7), a second side passage 132 is formed at the second side end 252 of the first segment body 101A and a first side passage 131 is formed at the first side end 251 of the second segment body 101B. Therefore, the second side end 252 of the first segment body 101A and the first side end 251 of the second segment body 101B, which are relatively difficult to cool, can be cooled.

[0177] (8) In several embodiments, based on the structure of (7) above, it is preferable that the portion of the first side passage 131 that is connected to the first side end 251 by fasteners (bolts 171 and nuts 172) in the axial direction Da when viewed from the blade height direction of the airfoil 23, and preferably that the portion of the second side passage 132 that is connected to the second side end 252 by fasteners (bolts 171 and nuts 172) in the axial direction Da when viewed from the blade height direction.

[0178] According to the structure described in (8), the second side end 252 of the first segment body 101A and the first side end 251 of the second segment body 101B, which are relatively difficult to cool, can be cooled efficiently.

[0179] (9) The gas turbine 10 of at least one embodiment of the present disclosure includes: a compressor 11 that compresses incoming air AI; a combustor 12 that supplies fuel FL to the compressed air AC compressed by the compressor 11 and burns the fuel FL; and a turbine 13 that obtains rotational power from combustion gas FG obtained by combustion in the combustor 12. The turbine 13 has turbine stator blades 21 with any of the structures described in (1) to (6) above.

[0180] According to the structure described above (9), since the cooling air CA after cooling the first region R1 which undergoes impact cooling and the cooling air CA after cooling the second region R2 which does not undergo impact cooling can be further utilized for cooling the shroud 2 (inner shroud 25), the cooling efficiency of the shroud 2 (inner shroud 25) can be improved. Therefore, excessive cooling and excessive use of cooling air CA can be suppressed, thereby improving the efficiency of the gas turbine 10.

[0181] (10) The gas turbine 10 of at least one embodiment of the present disclosure includes: a compressor 11 that compresses incoming air AI; a combustor 12 that supplies fuel FL to the compressed air AC compressed by the compressor 11 and burns the fuel FL; and a turbine 13 that obtains rotational power from combustion gas FG obtained by combustion in the combustor 12. The turbine 13 has a blade section body 100 with the structure of (7) or (8) described above.

[0182] According to the structure described above (10), since the cooling air CA after cooling the first region R1 which undergoes impact cooling and the cooling air CA after cooling the second region R2 which does not undergo impact cooling can be further utilized for cooling the shroud 2 (inner shroud 25), the cooling efficiency of the shroud 2 (inner shroud 25) can be improved. Thus, excessive cooling and excessive use of cooling air CA can be suppressed, thereby improving the efficiency of the gas turbine 10.

Claims

1. A turbine stator blade, wherein, The turbine stator blades include: The wing-shaped part; and A protective cover, which is disposed on at least one of the two sides of the airfoil in the direction of blade height. The protective cover has a recess on the surface opposite to the airfoil, which is separated from the gas passage surface. The recess includes: a first region covered by an impact plate having multiple through holes and subjected to impact cooling; and a second region not covered by the impact plate and not subjected to impact cooling. The protective cover has the following features: A first opening is formed in the first region; A first side passage is formed from the front edge to the rear edge at one of the first side ends in the circumferential direction of the shield, and one end is connected to the first opening. A second opening is formed in the second region; and A second side passage, which extends from the leading edge to the trailing edge at a second side end on the other side of the circumference of the shield, and one end of which connects to the second opening. Cooling air, after being subjected to impact cooling on the bottom surface of the inner region of the first area, is supplied to the first side passage. Cooling air, after cooling the bottom surface of the inner region of the second region, is supplied to the second side passage.

2. The turbine stator blade according to claim 1, wherein, The shield has a circumferential passage extending circumferentially at a position closer to the rear edge than the first region. The other end of the first side passage is connected to one end of the circumferential passage. The other end of the second side passage is connected to the other end of the circumferential passage.

3. A turbine stator blade, wherein, The turbine stator blades include: The wing-shaped part; and A protective cover, which is disposed on at least one of the two sides of the airfoil in the direction of blade height. The protective cover has a recess on the surface opposite to the airfoil, which is separated from the gas passage surface. The recess includes: a first region covered by an impact plate having multiple through holes; and a second region not covered by the impact plate. The protective cover has the following features: A first opening is formed in the first region; A first side passage is formed from the front edge to the rear edge at one of the first side ends in the circumferential direction of the shield, and one end is connected to the first opening. A second opening is formed in the second region; and A second side passage, which extends from the leading edge to the trailing edge at a second side end on the other side of the circumference of the shield, and one end of which connects to the second opening. The shield has a circumferential passage extending circumferentially at a position closer to the rear edge than the first region. The other end of the first side passage is connected to one end of the circumferential passage. The other end of the second side passage is connected to the other end of the circumferential passage. The inlet area of ​​the second opening is smaller than the inlet area of ​​the first opening.

4. The turbine stator blade according to claim 2 or 3, wherein, The shield has a rear edge end passage, and multiple rear edge end passages are arranged circumferentially on the rear edge side, with one end connected to the circumferential passage and the other end open on the rear edge end face of the shield.

5. The turbine stator blade according to any one of claims 1 to 3, wherein, The first region includes the area located on the ventral side of the airfoil when viewed from the height direction of the blade.

6. The turbine stator blade according to any one of claims 1 to 3, wherein, The protective cover is disposed on the inner side of the airfoil in the direction of blade height.

7. A blade segment body, wherein, The blade section body comprises a first section body and a second section body including the turbine stator blade as described in any one of claims 1 to 6. The second side end of the first segment body is connected to the first side end of the second segment body by a fastener.

8. The blade segment body according to claim 7, wherein, The first side passage overlaps axially with the portion of the first side end that is joined by the fastener when viewed from the blade height direction of the airfoil. The second side passage overlaps axially with the portion of the second side end that is joined by the fastener when viewed from the blade height direction.

9. A gas turbine, wherein, The gas turbine has the following features: A compressor, which compresses the incoming air; A burner that supplies fuel to compressed air compressed by the compressor and burns the fuel; and A turbine that obtains rotational power from combustion gases obtained by burning in the burner. The turbine has turbine stator blades as described in any one of claims 1 to 6.

10. A gas turbine, wherein, The gas turbine has the following features: A compressor, which compresses the incoming air; A burner that supplies fuel to compressed air compressed by the compressor and burns the fuel; and A turbine that obtains rotational power from combustion gases obtained by burning in the burner. The turbine has the blade section body as described in claim 7 or 8.

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