Turbine rotor blade, and gas turbine

Abstract

A turbine rotor blade of a gas turbine according to one embodiment comprises: a plurality of cooling passages that extend along a blade height direction inside an airfoil portion and have an elongated hole shape extending along the camber line of the airfoil portion when viewed from the blade height direction; and at least one hollow portion that is provided inside a shroud portion and communicates with any one of the plurality of cooling passages. The plurality of cooling passages include a first cooling passage that is located further to the leading edge side than a seal fin that protrudes from the shroud portion toward the blade tip side, and a second cooling passage that is located further to the trailing edge side than the seal fin and is adjacent to the first cooling passage. The at least one hollow portion includes a single first hollow portion that communicates with the first cooling passage and the second cooling passage.

Description

Turbine rotor blade and gas turbine

[0001] The present disclosure relates to a turbine rotor blade and a gas turbine. The present application claims priority based on Japanese Patent Application No. 2025-037870 filed with the Japan Patent Office on March 11, 2025, the content of which is incorporated herein by reference.

[0002] In a turbine rotor blade of a gas turbine, a plurality of cooling passages extending in the blade height direction are formed inside an airfoil section for cooling with compressed air (see, for example, Patent Document 1).

[0003] Japanese Unexamined Patent Publication No. 2007-327493

[0004] For example, in the case of a turbine rotor blade having a relatively long blade length, such as a rear-stage rotor blade in a gas turbine turbine, the plurality of cooling passages formed inside the airfoil section are cooling passages called multi-holes having a circular cross-sectional shape. In such cooling passages, the pressure loss of compressed air (cooling air) is relatively large, so it is desired to reduce the pressure loss. However, increasing the passage cross-sectional area of the cooling passage increases the stress acting on the cross-section of the airfoil section, so it is desirable to also reduce the weight of the shroud portion located on the blade tip side.

[0005] In view of the above circumstances, at least one embodiment of the present disclosure aims to provide a turbine rotor blade and a gas turbine capable of reducing the pressure loss of cooling air.

[0006] (1) A turbine blade according to at least one embodiment of the present disclosure is a turbine blade for a gas turbine, comprising: an airfoil portion extending in the blade height direction; a shroud portion located closer to the blade tip than the airfoil portion; a plurality of cooling passages having an elongated shape extending along the blade height direction within the airfoil portion and extending along the camber line of the airfoil portion when viewed from the blade height direction; and at least one hollow portion provided inside the shroud portion and communicating with any of the plurality of cooling passages, wherein the plurality of cooling passages include: a first cooling passage located closer to the leading edge than a seal fin protruding from the shroud portion toward the blade tip; and a second cooling passage located closer to the trailing edge than the seal fin and adjacent to the first cooling passage, and the at least one hollow portion includes a single first hollow portion communicating with the first cooling passage and the second cooling passage.

[0007] (2) A gas turbine according to at least one embodiment of the present disclosure comprises a turbine blade having the configuration of (1) above, and a combustor for generating combustion gas flowing through a combustion gas passage on which the turbine blade is provided.

[0008] According to at least one embodiment of the present disclosure, it is possible to provide a turbine blade and a gas turbine that can reduce the pressure loss of cooling air.

[0009] This is a schematic diagram of a gas turbine to which a turbine blade according to one embodiment is applied. This is a schematic diagram of the turbine blade according to one embodiment, viewed in the direction from the negative pressure surface toward the pressure surface (direction along the rotor circumferential direction). This is a schematic diagram showing a cross-section along the camber line near the blade tip of the turbine blade according to one embodiment, showing the state before the through-hole is closed. This is a schematic diagram showing a cross-section along the camber line near the blade tip of the turbine blade according to one embodiment, showing the state after the through-hole is closed. This is a view of the shroud portion located on the blade tip side of the turbine blade according to one embodiment, viewed from the blade height direction. This is a view of the shroud portion located on the blade tip side of the turbine blade according to one embodiment, viewed from the blade height direction, with the cooling holes omitted. This is a cross-sectional view taken along the line VII-VII in Figure 6. This is a cross-sectional view taken along the line VIII-VIII in Figure 6. This is a cross-sectional view taken along the line IX-IX in Figure 6. This is a cross-sectional view taken along the line X-X in Figure 6.

[0010] Hereinafter, several embodiments of this disclosure will be described with reference to the attached drawings. However, the dimensions, materials, shapes, relative arrangements, etc., of the components described as embodiments or shown in the drawings are not intended to limit the scope of this disclosure, but are merely illustrative examples. For example, expressions describing relative or absolute arrangements such as "in a certain direction," "along a certain direction," "parallel," "orthogonal," "center," "concentric," or "coaxial" should not only describe such arrangements strictly, but also represent states of relative displacement with tolerances, or angles or distances to the extent that the same function is achieved. For example, expressions describing things being in an equal state such as "identical," "equal," and "homogeneous" should not only describe states of being strictly equal, but also represent states where tolerances, or differences to the extent that the same function is achieved, exist. For example, expressions describing shapes such as square shapes or cylindrical shapes should not only describe geometrically precise square shapes or cylindrical shapes, but also represent shapes including concave and concave parts, chamfered parts, etc., to the extent that the same effect is achieved. On the other hand, expressions such as "to possess," "to be equipped with," "to have," "to include," or "to have" a single component are not exclusive expressions that exclude the existence of other components.

[0011] <Gas Turbine Configuration> First, a gas turbine to which the turbine blades according to one embodiment are applied will be described. Figure 1 is a schematic diagram of a gas turbine to which the turbine blades according to one embodiment are applied. As shown in Figure 1, the gas turbine 1 comprises a compressor 2 for generating compressed air, a combustor 4 for generating combustion gas using compressed air and fuel, and a turbine 6 configured to be rotationally driven by the combustion gas. In the case of a gas turbine 1 for power generation, a generator (not shown) is connected to the turbine 6.

[0012] The compressor 2 includes a plurality of stationary vanes 16 fixed to the compressor casing 10 side, and a plurality of rotor blades 18 mounted on the rotor 8 so as to be alternately arranged with respect to the stationary vanes 16. Air taken in from the air intake 12 is supplied to the compressor 2, and this air is compressed by passing through the plurality of stationary vanes 16 and the plurality of rotor blades 18 to become high-temperature, high-pressure compressed air.

[0013] The combustor 4 is supplied with fuel and compressed air generated by the compressor 2. In the combustor 4, the fuel and compressed air are mixed and burned to produce combustion gas, which is the working fluid for the turbine 6. As shown in Figure 1, multiple combustors 4 may be arranged within the casing 20 along the circumferential direction with the rotor 8 at its center.

[0014] The turbine 6 has a combustion gas passage 28 formed within the turbine casing 22, and includes a plurality of stator blades 24 and turbine rotor blades (rotor blades) 26 provided in the combustion gas passage 28. The stator blades 24 are fixed to the turbine casing 22 side, and a plurality of stator blades 24 arranged along the circumferential direction of the rotor 8 constitute a stator blade row. The rotor blades 26 are mounted on the rotor 8, and a plurality of rotor blades 26 arranged along the circumferential direction of the rotor 8 constitute a rotor blade row. The stator blade row and the rotor blade row are arranged alternately in the axial direction of the rotor 8.

[0015] In the turbine 6, combustion gas from the combustor 4 flows into the combustion gas passage 28 and passes through multiple stationary blades 24 and multiple rotor blades 26, thereby driving the rotor 8 to rotate. This drives a generator connected to the rotor 8, generating electricity. After driving the turbine 6, the combustion gas is discharged to the outside through the exhaust chamber 29.

[0016] <Configuration of Turbine Blades> The turbine blades 26 according to one embodiment will be described in more detail below. Figure 2 is a schematic diagram of the turbine blades 26 according to one embodiment, viewed in the direction from the negative pressure surface to the pressure surface (direction along the rotor circumference). Figure 3 is a schematic diagram showing a cross-section along the camber line Lca near the blade tip of the turbine blades 26 according to one embodiment, showing the state before the through-hole 72a, which will be described later, is closed. Figure 4 is a schematic diagram showing a cross-section along the camber line Lca near the blade tip of the turbine blades 26 according to one embodiment, showing the state after the through-hole 72a, which will be described later, is closed. Figure 5 is a view of the shroud portion 52 located on the blade tip side of the turbine blades 26 according to one embodiment, viewed from the blade height direction. Figure 6 is a view of the shroud portion 52 located on the blade tip side of the turbine blades 26 according to one embodiment, viewed from the blade height direction, and the cooling holes 81, which will be described later, are omitted. Figure 7 is a cross-sectional view taken along the line VII-VII in Figure 6. Figure 8 is a cross-sectional view taken along the line VIII-VIII in Figure 6. Figure 9 is a cross-sectional view taken along the line IX-IX in Figure 6. Figure 10 is a cross-sectional view taken along the line X-X in Figure 6.

[0017] In one embodiment, the turbine blade 26 comprises a platform 32, an airfoil portion 34 and a blade root portion 36 connected to the platform 32, a shroud portion 52 located closer to the blade tip than the airfoil portion 34, and a fillet portion 40 connected to the shroud portion 52. The turbine blade 26 also includes a seal fin 54 to reduce fluid leakage at the blade tip of the turbine blade 26. In one embodiment, the turbine blade 26 is, for example, a three-stage blade for a turbine 6 with four stages, but it may also be a one-stage blade, a two-stage blade, or a four-stage blade. Note that the number of stages of the turbine 6 is not limited to four.

[0018] The airfoil section 34 extends in the wing height direction (span direction) and has a base end 38 and a tip end 39, which are the two ends in the wing height direction, and is connected to the platform 32 at the base end 38. The airfoil section 34 also has a leading edge 42 and a trailing edge 44 that extend along the wing height direction, and a pressure surface 46 and a negative pressure surface 48 that extend between the leading edge 42 and the trailing edge 44. The airfoil section 34 may have a twisted shape as it moves from the base end 38 to the tip end 39 in the wing height direction.

[0019] The blade root 36 is located on the opposite side of the airfoil 34 from the platform 32 in the blade height direction. The blade root 36 includes an engagement portion having an uneven shape, and the turbine blade 26 is attached to the rotor 8 of the turbine 6 by the engagement portion engaging with a blade groove provided in a rotor disk (not shown) that rotates with the rotor 8.

[0020] Furthermore, when the turbine blades 26 are attached to the rotor 8, the blade height direction is aligned with the radial direction of the turbine 6. In other words, the blade height direction of the turbine blades 26 and the radial direction of the turbine 6 are approximately the same.

[0021] The fillet portion 40 is formed by a curved surface 40a and is connected to the end of the shroud portion 52 on the airfoil portion 34 side. The fillet portion 40 may also be connected to a flat surface 52a of the shroud portion 52 that extends in the direction perpendicular to the airfoil height. The fillet portion 40 formed by the curved surface 40a can alleviate stress concentration at the connection point of the shroud portion 52 to the airfoil portion 34.

[0022] The seal fins 54 protrude from the shroud portion 52 toward the blade tip and are provided to extend along the circumferential direction. The seal fins 54 of the multiple turbine blades 26 arranged in the circumferential direction form an annular seal portion.

[0023] The turbine blade 26 further comprises a plurality of cooling passages 60, at least one cooling cavity 70, and a plurality of cooling holes 81.

[0024] Each of the multiple cooling passages 60 extends within the airfoil 34 along the airfoil height direction. Typically, the multiple cooling passages 60 are arranged along the camber line Lca of the airfoil 34. The multiple cooling passages 60 have an elongated hole shape that extends along the camber line Lca of the airfoil 34 when viewed from the airfoil height direction. The dimension of each of the multiple cooling passages 60 along the camber line Lca is greater than the airfoil thickness in the region of the airfoil 34 on the tip side, for example, in the region where the cooling passages 60 are formed. In one embodiment of the turbine blade 26, for example, four cooling passages 60 are provided. The four cooling passages 60 are also referred to as the first pass 61, the second pass 62, the third pass 63, and the fourth pass 64, in order from the leading edge 42 side.

[0025] The first pass 61 and the second pass 62 are located on the leading edge 42 side of the seal fin 54. The third pass 63 and the fourth pass 64 are located on the trailing edge 44 side of the seal fin 54. The third pass 63 is adjacent to the second pass 62 along the camber line Lca.

[0026] Each of the multiple cooling passages 60 extends in the wing height direction from the bottom surface of the wing root 36 to the cooling cavity 70, which will be described later. The first pass 61 is connected to the first cooling cavity 71, which will be described later. The second pass 62 and the third pass 63 are connected to the second cooling cavity 72, which will be described later. The fourth pass 64 is connected to the third cooling cavity 73, which will be described later.

[0027] The cooling cavity 70 is at least partially provided inside the shroud portion 52. In one embodiment of the turbine blade 26, the cooling cavity 70 is formed from the shroud portion 52 to the fillet portion 40 in the blade height direction. In one embodiment of the turbine blade 26, a first cooling cavity 71, a second cooling cavity 72, and a third cooling cavity 73 are formed in order from the leading edge 42 side.

[0028] The first cooling cavity 71 is formed in a position that overlaps with at least a portion of the first pass 61 when viewed from the wing height direction. The shroud portion 52 has a through hole 71a that opens at the wingtip side end face 52b of the shroud portion 52 and communicates with the first cooling cavity 71. The through hole 71a is formed in a position that overlaps with at least a portion of the first pass 61 when viewed from the wing height direction.

[0029] The second cooling cavity 72 is formed in a position that overlaps with at least a portion of the second pass 62 and the third pass 63 when viewed from the wing height direction. The second cooling cavity 72 extends from a region on the leading edge 42 side of the seal fin 54 to a region on the trailing edge 44 side of the seal fin 54 when viewed from the wing height direction. A portion of the second cooling cavity 72 overlaps with a portion of the seal fin 54 when viewed from the wing height direction.

[0030] The shroud portion 52 has two through holes 72a that open at the wingtip end face 52b of the shroud portion 52 and communicate with the second cooling cavity 72. These through holes 72a are closed by inserting and welding a member such as a plug 72b, as shown in Figure 4, or simply by welding. The through holes 72a are formed at positions that overlap with at least a part of the second pass 62 and at least a part of the third pass 63 when viewed from the wing height direction.

[0031] The third cooling cavity 73 is formed in a position that overlaps with at least a portion of the fourth pass 64 when viewed from the wing height direction. The shroud portion 52 has a through hole 73a that opens at the wingtip end face 52b of the shroud portion 52 and communicates with the third cooling cavity 73. The through hole 73a is formed in a position that overlaps with at least a portion of the third pass 63 when viewed from the wing height direction.

[0032] Each of the multiple cooling holes 81 is connected to the second cooling cavity 72 and opens onto the surface of the shroud portion 52. The cooling holes 81 may open onto the wingtip end face 52b (see Figure 5) of the shroud portion 52, or they may open onto the side surface of the shroud portion 52.

[0033] Multiple cooling passages 60 are supplied with compressed air from the compressor 2, which is the cooling medium, through inlet openings 58 that open at the ends of the blade roots 36 of the turbine blades 26. The compressed air supplied to the cooling passages 60 flows toward the blade tips and, after passing through the cooling passages 60, is retained in the cooling cavities 70. The compressed air in the first cooling cavity 71 is released to the outside of the turbine blades 26 through through holes 71a. The compressed air in the second cooling cavity 72 flows through cooling holes 81 and is released to the outside of the turbine blades 26 through openings 83 located on the surface of the shroud portion 52. The compressed air in the third cooling cavity 73 is released to the outside of the turbine blades 26 through through holes 73a. In this way, the turbine blades 26, including the airfoil portion 34 and the shroud portion 52, are cooled by flowing compressed air inside the turbine blades 26.

[0034] In the case of turbine blades with relatively long blade lengths, such as the downstream blades of the turbine 6 of a gas turbine 1, the multiple cooling passages formed inside the airfoil are conventionally called multi-hole cooling passages, which have a circular cross-sectional shape. In such cooling passages, the pressure loss of compressed air is relatively large, so it is desirable to reduce the pressure loss. However, increasing the cross-sectional area of ​​the cooling passages increases the stress acting on the cross-section of the airfoil, so it is also desirable to reduce the weight of the shroud located on the blade tip side. Therefore, in the turbine blade 26 according to one embodiment, the pressure loss of the cooling passage 60 and the weight of the shroud 52 are reduced as follows.

[0035] In one embodiment of the turbine blade 26, the number of cooling passages 60 is reduced compared to the number of cooling passages called multi-holes having a conventional circular cross-sectional shape, and the cross-sectional shape of the cooling passages 60 in the direction perpendicular to the blade height direction is made into an elongated hole shape extending along the camber line Lca, thereby increasing the cross-sectional area of ​​each cooling passage 60. This reduces the pressure loss in the cooling passages 60.

[0036] In the case of conventional cooling passages called multi-holes with a circular cross-sectional shape, the cooling passages were formed by machining after the turbine blades were cast. In one embodiment of the turbine blade 26, the cooling passages 60 are formed during the casting of the turbine blade 26. That is, the cooling passages 60 are formed by installing a core at the position corresponding to the cooling passages 60 during the casting of the turbine blade 26. When installing the core during the casting of the turbine blade 26, the end of the core on the blade tip side is supported so that it protrudes from the through holes 71a, 72a, and 73a towards the blade tip side.

[0037] Of these through holes 71a, 72a, and 73a, the through hole 71a communicating with the first cooling cavity 71 and the through hole 73a communicating with the third cooling cavity 73 are left open after casting to allow compressed air to be discharged. The two through holes 72a communicating with the second cooling cavity 72 are closed after casting as described above.

[0038] In one embodiment of the turbine blade 26, the second cooling cavity 72, through which the second pass 62 and the third pass 63 of the plurality of cooling passages 60 communicate, is a single cooling cavity 70. That is, the second cooling cavity 72 is formed to extend along the camber line Lca from the second pass 62 to the third pass 63 when viewed from the blade height direction. By forming the second cooling cavity 72 to extend along the camber line Lca from the second pass 62 to the third pass 63 in this way, the volume of the second cooling cavity 72 can be increased compared to the case where the second cooling cavity 72 is divided into a cooling cavity 70 communicating with the second pass 62 and a cooling cavity 70 communicating with the third pass 63. As a result, the shroud portion 52 can be made lighter, and thus the stress from the airfoil portion 34 to the shroud portion 52 can be reduced. Furthermore, according to one embodiment of the gas turbine 1, the pressure loss in the cooling passage 60 in the turbine blade 26 can be reduced, thereby improving the efficiency of the gas turbine 1.

[0039] In one embodiment of the turbine blade 26, as shown in Figures 2 to 4, the second pass 62 and the third pass 63 are arranged such that the distance between the second pass 62 and the third pass 63 increases as you approach the blade tip, near the end 60a on the blade tip side (for example, the blade tip region including the end 60a). In the turbine blade 26, it has been found that in the blade tip region of the airfoil 34, the temperature in region RM (see Figure 2) between the leading edge 42 region RL and the trailing edge 44 region RT (see Figure 2) tends to be lower than the temperature in the leading edge 42 region RL and the trailing edge 44 region RT (see Figure 2). If such a temperature difference becomes large, it may lead to undesirable deformation of the airfoil 34. According to the turbine blade 26 of one embodiment, the above temperature difference can be reduced, and the possibility of undesirable deformation of the airfoil 34 can be reduced.

[0040] In one embodiment of the turbine blade 26, the distance between the second pass 62 and the third pass 63 is greater than the distance between the second pass 62 and the first pass 61 in the vicinity of the end 60a on the blade tip side of the first pass 61 to the third pass 63 (for example, the blade tip region including the end 60a). That is, in the region relatively close to the blade tip, the first pass 61 and the second pass 62 are arranged relatively close together, but the second pass 62 and the third pass 63 are arranged relatively far apart. As described above, in the turbine blade 26, in the blade tip region of the airfoil 34, the temperature of region RM between the region RL on the leading edge 42 side and the region RT on the trailing edge 44 side tends to be lower than that of the region RL on the leading edge 42 side and the region RT on the trailing edge 44 side. According to the turbine blade 26 of one embodiment, as described above, by arranging the first pass 61 and the second pass 62 relatively close together, and arranging the second pass 62 and the third pass 63 relatively far apart, the temperature difference between the region RL on the leading edge 42 side and the region RM between the region RL on the leading edge 42 side and the region RT on the trailing edge 44 side can be reduced in the region on the tip side of the airfoil 34, thereby reducing the possibility of undesirable deformation of the airfoil 34.

[0041] In one embodiment of the turbine blade 26, the distance between the second pass 62 and the third pass 63 is greater than the distance between the third pass 63 and the fourth pass 64 in the vicinity of the end 60a on the blade tip side of the second pass 62 to the fourth pass 64 (for example, the blade tip region including the end 60a). That is, in the region relatively close to the blade tip, the third pass 63 and the fourth pass 64 are arranged relatively close together, while the second pass 62 and the third pass 63 are arranged relatively far apart. As described above, in the turbine blade 26, in the blade tip region of the airfoil 34, the temperature of region RM between the region RL on the leading edge 42 side and the region RT on the trailing edge 44 side tends to be lower than that of the region RL on the leading edge 42 side and the region RT on the trailing edge 44 side. According to the turbine blade 26 of one embodiment, as described above, by arranging the third pass 63 and the fourth pass 64 relatively close together and the second pass 62 and the third pass 63 relatively far apart, the temperature difference between the region RT on the trailing edge 44 side and the region RM between the region RL on the leading edge 42 side and the region RT on the trailing edge 44 side can be reduced in the region on the tip side of the airfoil 34, thereby reducing the possibility of undesirable deformation of the airfoil 34.

[0042] In one embodiment of the turbine blade 26, the shroud portion 52 preferably has through holes 72a formed at positions where the second pass 62 and the third pass 63 are extended in the blade height direction. Furthermore, the through holes 72a preferably are closed. This makes it possible to support the core for forming the second pass 62 and the third pass 63 by having it protrude from the through holes 72a toward the blade tip side of the airfoil portion 34 during the casting of the turbine blade 26, thus facilitating the installation of the core. According to the turbine blade 26 of one embodiment, the through holes 72a, which become unnecessary after casting, can be closed.

[0043] In one embodiment of the turbine blade 26, the shroud portion 52 is preferably connected to the second cooling cavity 72 and has a plurality of cooling holes 81 opening on the surface of the shroud portion 52. This allows the shroud portion 52 to be cooled by compressed air that has cooled the airfoil portion 34 by flowing through the second pass 62 and the third pass 63, thus enabling effective use of the compressed air that has cooled the airfoil portion 34 by flowing through the second pass 62 and the third pass 63.

[0044] <On thickness t of fillet portion 40> The distance between the cooling cavity 70 and the outer surface (curved surface 40a) of the fillet portion 40 is defined as the thickness t of the fillet portion 40. As shown in Figures 7 and 8, in the turbine rotor blade 26 according to one embodiment, in at least a part of the region on the leading edge 42 side along the camber line Lca, the thickness t of the fillet portion 40 on the suction surface 48 side is preferably larger than the thickness t of the fillet portion 40 on the pressure surface 46 side. As shown in Figures 9 and 10, in the turbine rotor blade 26 according to one embodiment, in at least a part of the region on the trailing edge 44 side along the camber line Lca, the thickness t of the fillet portion 40 on the pressure surface 46 side is preferably larger than the thickness t of the fillet portion 40 on the suction surface 48 side.

[0045] As shown in Figures 5 and 6, in the turbine rotor blade 26 according to one embodiment, when viewed from the blade height direction, the distance from the blade surface (the pressure surface 46 and the suction surface 48) of the airfoil portion 34 to the end surface of the shroud portion 52 tends to be larger on the suction surface 48 side than on the pressure surface 46 side in the region on the leading edge 42 side along the camber line Lca, and tends to be larger on the pressure surface 46 side than on the suction surface 48 side in the region on the trailing edge 44 side along the camber line Lca. Therefore, the centrifugal force of the shroud portion 52 acting on the fillet portion 40 during operation of the gas turbine 1 tends to be larger on the suction surface 48 side than on the pressure surface 46 side in the region on the leading edge 42 side along the camber line Lca, and tends to be larger on the pressure surface 46 side than on the suction surface 48 side in the region on the trailing edge 44 side along the camber line Lca. In Figure 6, a region Rs enclosed by a two-dot chain line is a region where the stress acting on the fillet portion 40 due to the centrifugal force of the shroud portion 52 is relatively higher than that in other regions.

[0046] According to the turbine rotor blade 26 according to one embodiment, between the fillet portion 40 on the pressure surface 46 side and the fillet portion 40 on the suction surface 48 side, increasing the thickness t of the fillet portion 40 on the side where the magnitude of the centrifugal force of the shroud portion 52 acting on the fillet portion 40 during operation of the gas turbine 1 is larger can reduce the stress in the fillet portion 40.

[0047] In the turbine rotor blade 26 according to one embodiment, within a region sandwiched between the fillet portion 40 on the pressure surface 46 side and the fillet portion 40 on the suction surface 48 side, in a region where the stress acting on the fillet portion 40 on the pressure surface 46 side during operation of the gas turbine 1 is larger than the stress acting on the fillet portion 40 on the suction surface 48 side, the cooling passage 60 whose blade tip side end portion 60a is located in the region when viewed from the blade height direction is preferably arranged offset toward the suction surface 48 side of the airfoil portion 34 in the vicinity of the blade tip side end portion 60a of the cooling passage 60 (for example, a blade tip side region including the end portion 60a).

[0048] For example, as shown in FIG. 6, in the turbine rotor blade 26 according to one embodiment, in a region on the trailing edge 44 side along the camber line Lca, the stress acting on the fillet portion 40 on the pressure surface 46 side during operation of the gas turbine 1 is larger than the stress acting on the fillet portion 40 on the suction surface 48 side. Therefore, in the region on the trailing edge 44 side along the camber line Lca, the third pass 63 and the fourth pass 64, which are cooling passages 60 whose blade tip side end portions 60a are located in the region when viewed from the blade height direction, are preferably arranged offset toward the suction surface 48 side of the airfoil portion 34 in the vicinity of the blade tip side end portion 60a of the third pass 63 and the fourth pass 64 (for example, a blade tip side region including the end portion 60a) (see FIGS. 9 and 10).

[0049] Within a region sandwiched between the fillet portion 40 on the pressure surface 46 side and the fillet portion 40 on the suction surface 48 side, in a region where the stress acting on the fillet portion 40 on the suction surface 48 side during operation of the gas turbine 1 is larger than the stress acting on the fillet portion 40 on the pressure surface 46 side, the cooling passage 60 whose blade tip side end portion 60a is located in the region when viewed from the blade height direction is preferably arranged offset toward the pressure surface 46 side of the airfoil portion 34 in the vicinity of the blade tip side end portion 60a of the cooling passage 60 (for example, a blade tip side region including the end portion 60a).

[0050] For example, as shown in Figure 6, in the turbine blade 26 according to one embodiment, in the region on the leading edge 42 side along the camber line Lca, the stress acting on the fillet portion 40 on the negative pressure surface 48 side during operation of the gas turbine 1 is greater than the stress acting on the fillet portion 40 on the pressure surface 46 side. Therefore, in the region on the leading edge 42 side along the camber line Lca, the first pass 61 and the second pass 62, which are cooling passages 60 whose tip end 60a is located in that region when viewed from the blade height direction, are preferably offset towards the pressure surface 46 side of the airfoil portion 34 in the vicinity of the tip end 60a in the first pass 61 and the second pass 62 (for example, the region on the blade tip side including the tip 60a) (see Figures 7 and 8).

[0051] According to the turbine blade 26 of one embodiment, the distance between the surface of the airfoil portion 34 on the opposite side of the offset direction and the cooling passage 60 can be increased, making it easier to secure the thickness t of the fillet portion 40 on the opposite side of the offset direction. For example, if the third pass 63 and the fourth pass 64 are offset to the negative pressure surface 48 side of the airfoil portion 34, the distance between the pressure surface 46 and the third pass 63 and the fourth pass 64 can be increased, making it easier to secure the thickness t of the fillet portion 40 on the pressure surface 46 side in the region on the trailing edge 44 side. For example, if the first pass 61 and the second pass 62 are offset to the pressure surface 46 side of the airfoil portion 34, the distance between the negative pressure surface 48 and the first pass 61 and the second pass 62 can be increased, making it easier to secure the thickness t of the fillet portion 40 on the negative pressure surface 48 side in the region on the leading edge 42 side.

[0052] This disclosure is not limited to the embodiments described above, and also includes modified forms of the embodiments described above, as well as forms that combine these forms as appropriate. In the embodiments described above, the cooling cavity 70 includes a first cooling cavity 71, a second cooling cavity 72, and a third cooling cavity 73. However, the cooling cavity 70 does not necessarily include the first cooling cavity 71, and the end 60a of the first pass 61 on the blade tip side may open at the blade tip side end face 52b of the shroud portion 52. Also, the cooling cavity 70 does not necessarily include the third cooling cavity 73, and the end 60a of the fourth pass 64 on the blade tip side may open at the blade tip side end face 52b of the shroud portion 52.

[0053] In the embodiment described above, the plurality of cooling passages 60 include a first pass 61, a second pass 62, a third pass 63, and a fourth pass 64, but the plurality of cooling passages 60 do not necessarily include at least one of the first pass 61 or the fourth pass 64.

[0054] The contents of each of the above embodiments can be understood, for example, as follows: (1) A turbine blade 26 according to at least one embodiment of the present disclosure is a turbine blade 26 for a gas turbine 1, comprising: an airfoil portion 34 extending in the blade height direction; a shroud portion 52 located on the blade tip side of the airfoil portion 34; a plurality of cooling passages 60 having an elongated hole shape that extends along the blade height direction inside the airfoil portion 34 and extends along the camber line Lca of the airfoil portion 34 when viewed from the blade height direction; and at least one hollow portion (cooling cavity 70) provided inside the shroud portion 52 and communicating with any of the plurality of cooling passages 60. The multiple cooling passages 60 include a first cooling passage (second pass 62) located on the leading edge 42 side of the seal fin 54 that protrudes from the shroud portion 52 toward the wingtip, and a second cooling passage (third pass 63) located on the trailing edge 44 side of the seal fin 54 and adjacent to the first cooling passage (second pass 62). At least one hollow portion (cooling cavity 70) includes a single first hollow portion (second cooling cavity 72) that communicates with the first cooling passage (second pass 62) and the second cooling passage (third pass 63).

[0055] According to the configuration of (1) above, by making the multiple cooling passages 60 elongated holes that extend along the camber line Lca, the pressure loss in the cooling passages 60 can be reduced. According to the configuration of (1) above, the shroud portion 52 can be made lighter, so the stress from the airfoil portion 34 to the shroud portion 52 can be reduced.

[0056] (2) In some embodiments, in the configuration of (1) above, the first cooling passage (second pass 62) and the second cooling passage (third pass 63) are arranged such that, near the end 60a on the wingtip side of the first cooling passage (second pass 62) and the second cooling passage (third pass 63), the distance between the first cooling passage (second pass 62) and the second cooling passage (third pass 63) increases as it approaches the wingtip side.

[0057] In the turbine blade 26, it has been found that in the region RM between the leading edge RL and the trailing edge RL of the airfoil 34, the temperature tends to be lower than that of the leading edge RL and the trailing edge RT of the airfoil 42. If such a temperature difference becomes large, it may lead to undesirable deformation of the airfoil 34. The configuration of (2) above can reduce the above temperature difference and reduce the possibility of undesirable deformation of the airfoil 34.

[0058] (3) In some embodiments, in the configuration of (1) or (2) above, the plurality of cooling passages 60 may include a third cooling passage (first pass 61) which is located on the leading edge 42 side of the first cooling passage (second pass 62) and is adjacent to the first cooling passage (second pass 62). In the vicinity of the end portion 60a on the blade tip side of the first cooling passage (second pass 62), the second cooling passage (third pass 63), and the third cooling passage (first pass 61), the distance between the first cooling passage (second pass 62) and the second cooling passage (third pass 63) is preferably greater than the distance between the first cooling passage (second pass 62) and the third cooling passage (first pass 61).

[0059] According to the configuration of (3) above, the temperature difference between the region RL on the leading edge 42 side and the region RM between the region RL on the leading edge 42 side and the region RT on the trailing edge 44 side can be reduced in the region on the wingtip side of the airfoil 34, thereby reducing the possibility of undesirable deformation of the airfoil 34.

[0060] (4) In some embodiments, in any of the configurations (1) to (3) above, the plurality of cooling passages 60 may include a fourth cooling passage (fourth passage 64) that is located on the trailing edge 44 side of the second cooling passage (third passage 63) and is adjacent to the second cooling passage (third passage 63). In the vicinity of the end portion 60a on the blade tip side of the first cooling passage (second passage 62), the second cooling passage (third passage 63), and the fourth cooling passage (fourth passage 64), the distance between the first cooling passage (second passage 62) and the second cooling passage (third passage 63) is preferably greater than the distance between the second cooling passage (third passage 63) and the fourth cooling passage (fourth passage 64).

[0061] According to the configuration of (4) above, the temperature difference between the region RT on the trailing edge 44 side and the region RM between the region RL on the leading edge 42 side and the region RT on the trailing edge 44 side can be reduced in the region on the wingtip side of the airfoil 34, thereby reducing the possibility of undesirable deformation of the airfoil 34.

[0062] (5) In some embodiments, in any of the configurations (1) to (4) above, the shroud portion 52 may have through holes 72a formed at positions that extend the first cooling passage (second pass 62) and the second cooling passage (third pass 63) in the wing height direction. The through holes 72a may be closed.

[0063] According to the configuration of (5) above, when casting the turbine blade 26, the core for forming the first cooling passage (second pass 62) and the second cooling passage (third pass 63) can be supported by protruding from the through hole 72a toward the blade tip side of the airfoil portion 34, thus facilitating the installation of the core. According to the configuration of (5) above, the through hole 72a, which becomes unnecessary after casting, can be closed.

[0064] (6) In some embodiments, in any of the configurations (1) to (5) above, the shroud portion 52 may have a plurality of cooling holes 81 that are connected to at least one hollow portion (second cooling cavity 72) and open to the surface of the shroud portion 52.

[0065] According to the configuration of (6) above, the shroud portion 52 can be cooled by compressed air that has cooled the airfoil portion 34 by flowing through the first cooling passage (second pass 62) and the second cooling passage (third pass 63), so the compressed air that has cooled the airfoil portion 34 by flowing through the first cooling passage (second pass 62) and the second cooling passage (third pass 63) can be effectively utilized.

[0066] (7) In some embodiments, the configuration of (1) to (6) above may include a fillet portion 40 formed by a curved surface 40a and connected to the end of the shroud portion 52 on the airfoil portion 34 side. The thickness t of the fillet portion 40 is defined as the distance between at least one hollow portion (cooling cavity 70) and the outer surface (curved surface 40a) of the fillet portion 40. In at least a portion of the region on the leading edge 42 side along the camber line Lca, the thickness t of the fillet portion 40 on the negative pressure surface 48 side of the airfoil portion 34 may be greater than the thickness t of the fillet portion 40 on the pressure surface 46 side of the airfoil portion 34. In at least a portion of the region on the trailing edge 44 side along the camber line Lca, the thickness t of the fillet portion 40 on the pressure surface 46 side may be greater than the thickness t of the fillet portion 40 on the negative pressure surface 48 side.

[0067] In some embodiments, when viewed from the blade height direction, the distance from the blade surface (pressure surface 46 and negative pressure surface 48) of the airfoil 34 to the end face of the shroud 52 tends to be greater on the negative pressure surface 48 side than on the pressure surface 46 side in the region on the leading edge 42 side along the camber line Lca, and tends to be greater on the pressure surface 46 side than on the negative pressure surface 48 side in the region on the trailing edge 44 side along the camber line Lca. Therefore, the centrifugal force of the shroud 52 acting on the fillet 40 during operation of the gas turbine 1 tends to be greater on the negative pressure surface 48 side than on the pressure surface 46 side in the region on the leading edge 42 side along the camber line Lca, and tends to be greater on the pressure surface 46 side than on the negative pressure surface 48 side in the region on the trailing edge 44 side along the camber line Lca. According to the configuration described in (7) above, the thickness t of the fillet portion 40 on the pressure surface 46 side and the fillet portion 40 on the negative pressure surface 48 side is increased on the side where the centrifugal force of the shroud portion 52 acting on the fillet portion 40 during operation of the gas turbine 1 is greater, thereby reducing the stress on the fillet portion 40.

[0068] (8) In some embodiments, the configuration of (1) to (7) above may include a fillet portion 40 formed by a curved surface 40a and connected to the end of the shroud portion 52 on the airfoil portion 34 side. In the region sandwiched between the fillet portion 40 on the pressure surface 46 side of the airfoil portion 34 and the fillet portion 40 on the negative pressure surface 48 side of the airfoil portion 34, in the region where the stress acting on the fillet portion 40 on the pressure surface 46 side is greater than the stress acting on the fillet portion 40 on the negative pressure surface 48 side when the gas turbine 1 is in operation, the cooling passage 60 whose end 60a on the tip side is located in that region when viewed from the airfoil height direction may be offset towards the negative pressure surface 48 side of the airfoil portion 34 in the vicinity of the end 60a on the tip side of the cooling passage 60. In the region sandwiched between the fillet portion 40 on the pressure surface 46 side and the fillet portion 40 on the negative pressure surface 48 side, in the region where the stress acting on the fillet portion 40 on the negative pressure surface 48 side is greater than the stress acting on the fillet portion 40 on the pressure surface 46 side when the gas turbine 1 is in operation, the cooling passage 60 whose tip end 60a is located in this region when viewed from the blade height direction should be offset towards the pressure surface 46 side of the airfoil portion 34 near the tip end 60a of the cooling passage 60.

[0069] According to the configuration of (8) above, the distance between the surface of the airfoil portion 34 on the opposite side of the offset direction and the cooling passage 60 can be increased, making it easier to secure the thickness t of the fillet portion 40 on the opposite side of the offset direction.

[0070] (9) In some embodiments, in the configuration of (8) above, the first cooling passage (second pass 62) is preferably offset towards the pressure surface 46 side of the airfoil portion 34 near the end 60a on the tip side of the first cooling passage (second pass 62), and the second cooling passage (third pass 63) is preferably offset towards the negative pressure surface 48 side of the airfoil portion 34 near the end 60a on the tip side of the second cooling passage (third pass 63).

[0071] As described above, the centrifugal force of the shroud portion 52 acting on the fillet portion 40 during operation of the gas turbine 1 tends to be greater on the negative pressure surface 48 side than on the pressure surface 46 side in the region on the leading edge 42 side along the camber line Lca, and tends to be greater on the pressure surface 46 side than on the negative pressure surface 48 side in the region on the trailing edge 44 side along the camber line Lca. According to the configuration of (9) above, it becomes easier to secure the thickness t of the fillet portion 40 on the negative pressure surface 48 side in the region on the leading edge 42 side along the camber line Lca. According to the configuration of (9) above, it becomes easier to secure the thickness t of the fillet portion 40 on the pressure surface 46 side in the region on the trailing edge 44 side along the camber line Lca.

[0072] (10) A gas turbine 1 according to at least one embodiment of the present disclosure comprises a turbine blade 26 having any of the configurations described in (1) to (9) above, and a combustor 4 for generating combustion gas that flows through a combustion gas passage 28 on which the turbine blade 26 is provided.

[0073] According to the configuration described in (10) above, the pressure loss in the cooling passage 60 in the turbine blade 26 can be reduced, thereby improving the efficiency of the gas turbine 1.

[0074] 1 Gas turbine 2 Compressor 4 Combustor 6 Turbine 8 Rotor 24 Stator blade 26 Turbine rotor blade 28 Combustion gas passage 32 Platform 34 Airfoil section 40 Fillet section 40a Curved surface 42 Leading edge 44 Trailing edge 46 Pressure surface 48 Negative pressure surface 52 Shroud section 52b Tip end surface 54 Seal fin 60 Cooling passage 60a End 61 First pass 62 Second pass 63 Third pass 64 Fourth pass 70 Cooling cavity 71 First cooling cavity 71a Through hole 72 Second cooling cavity 72a Through hole 72b Plug 73 Third cooling cavity 73a Through hole 81 Cooling hole 83 Opening

Claims

1. A turbine blade for a gas turbine, comprising: an airfoil portion extending in the blade height direction; a shroud portion located closer to the blade tip than the airfoil portion; a plurality of cooling passages having an elongated shape extending along the blade height direction within the airfoil portion and extending along the camber line of the airfoil portion when viewed from the blade height direction; and at least one hollow portion provided inside the shroud portion and communicating with any of the plurality of cooling passages, wherein the plurality of cooling passages include: a first cooling passage located closer to the leading edge than a seal fin protruding from the shroud portion toward the blade tip; and a second cooling passage located closer to the trailing edge than the seal fin and adjacent to the first cooling passage, and the at least one hollow portion comprising a single first hollow portion communicating with the first and second cooling passages.

2. The turbine blade according to claim 1, wherein the first cooling passage and the second cooling passage are arranged such that the distance between the first cooling passage and the second cooling passage increases as it approaches the blade tip side, near the end of the first cooling passage and the second cooling passage on the blade tip side.

3. The turbine blade according to claim 1 or 2, wherein the plurality of cooling passages include a third cooling passage located on the leading edge side of the first cooling passage and adjacent to the first cooling passage, and the distance between the first cooling passage and the second cooling passage is greater than the distance between the first cooling passage and the third cooling passage in the vicinity of the blade tip end of the first cooling passage, the second cooling passage, and the third cooling passage.

4. The turbine blade according to claim 1 or 2, wherein the plurality of cooling passages include a fourth cooling passage located on the trailing edge side of the second cooling passage and adjacent to the second cooling passage, and in the vicinity of the blade tip end of the first cooling passage, the second cooling passage, and the fourth cooling passage, the distance between the first cooling passage and the second cooling passage is greater than the distance between the second cooling passage and the fourth cooling passage.

5. The turbine blade according to claim 1 or 2, wherein the shroud portion has through holes formed at positions extending the first cooling passage and the second cooling passage in the blade height direction, and the through holes are closed.

6. The turbine blade according to claim 1 or 2, wherein the shroud portion is connected to the at least one hollow portion and has a plurality of cooling holes opening to the surface of the shroud portion.

7. A turbine blade according to claim 1 or 2, comprising: a fillet portion formed by a curved surface and connected to the end of the shroud portion on the airfoil side, wherein when the distance between the at least one hollow portion and the outer surface of the fillet portion is the thickness of the fillet portion, in at least a portion of the leading edge region along the camber line, the thickness of the fillet portion on the negative pressure side of the airfoil is greater than the thickness of the fillet portion on the pressure side of the airfoil, and in at least a portion of the trailing edge region along the camber line, the thickness of the fillet portion on the pressure side is greater than the thickness of the fillet portion on the negative pressure side.

8. A turbine blade according to claim 1 or 2, comprising: a fillet portion formed by a curved surface and connected to the end of the shroud portion on the airfoil side, wherein in a region sandwiched between the fillet portion on the pressure side of the airfoil and the fillet portion on the negative pressure side of the airfoil, in a region where the stress acting on the fillet portion on the pressure side is greater than the stress acting on the fillet portion on the negative pressure side when the gas turbine is in operation, the cooling passage whose end on the tip side is located in that region when viewed from the blade height direction is offset towards the negative pressure side of the airfoil near the end on the tip side of the cooling passage; and in a region where the stress acting on the fillet portion on the negative pressure side is greater than the stress acting on the fillet portion on the pressure side when the gas turbine is in operation, the cooling passage whose end on the tip side is located in that region when viewed from the blade height direction is offset towards the pressure side of the airfoil near the end on the tip side of the cooling passage.

9. The turbine blade according to claim 8, wherein the first cooling passage is offset toward the pressure surface side of the airfoil portion near the tip end of the first cooling passage, and the second cooling passage is offset toward the negative pressure side of the airfoil portion near the tip end of the second cooling passage.

10. A gas turbine comprising: a turbine blade according to claim 1 or 2; and a combustor for generating combustion gas flowing through a combustion gas passage on which the turbine blade is provided.

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