Gas turbine rotor blade and gas turbine

Abstract

To suppress heat stress occurring in a gas turbine rotor vane effectively.SOLUTION: A platform of a gas turbine rotor vane according to one embodiment has a groove part recessed from an end surface at a rear edge side to a front edge side. A bottom part of the groove part overlaps with at least an airfoil part when viewed from a radial direction. When an end of the bottom part at a belly side of the platform is referred to as a first point, a tangent line of the bottom part which extends along a surface intersecting with the radial direction in the first point is referred to as a first tangent line. When an intersection point between a line segment, connecting a rear edge side end of a serpentine cooling passage received by the interior of the airfoil part when viewed from the radial direction, and a rear edge side end of the airfoil part and the bottom part is referred to as a second point, a tangent line of the bottom part which extends along the surface in the second point is referred to as a second tangent line, and an intersection point of the first tangent line and the second tangent line as seen in the radial direction is referred to as a third point. The third point exists at an opposite side of the rear edge side end of the airfoil part across a straight line connecting the first point with the second point.SELECTED DRAWING: Figure 3A

Description

[Technical Field]

[0001] The present disclosure relates to gas turbine blades and gas turbines. [Background technology]

[0002] In gas turbine rotor blades, a temperature difference is likely to occur between the airfoil portion and the platform during transient states such as when the gas turbine starts or stops operating, which can easily cause thermal stress. It is known that this thermal stress is particularly likely to be large near the root portion where the trailing edge of the airfoil portion and the platform meet. For this reason, a gas turbine rotor blade is known that is configured to reduce the above-mentioned thermal stress by forming a groove in the platform that is recessed from the trailing edge side of the platform toward the leading edge side and extends in the circumferential direction of the rotor (see, for example, Patent Document 1). [Prior art documents] [Patent documents]

[0003] [Patent Document 1] Japanese Patent Application Publication No. 8-254103 Summary of the Invention [Problem to be solved by the invention]

[0004] In a turbine rotor blade, a serpentine cooling channel is formed inside the airfoil portion to cool the airfoil portion. This serpentine cooling channel extends over a range in the blade height direction, including at least a part of the platform. Therefore, the groove formed in the platform needs to be formed at a position that ensures a wall thickness between the serpentine cooling channel and the platform. Therefore, there is a limit to the depth of the groove.

[0005] In view of the above circumstances, at least one embodiment of the present disclosure has an object to provide a gas turbine rotor blade that can effectively suppress thermal stress occurring in the gas turbine rotor blade. [Means for solving the problem]

[0006] (1) A gas turbine blade according to at least one embodiment of the present disclosure comprises: a base end portion fixed to the rotor; an airfoil portion extending radially of the rotor and having pressure and suction surfaces defining an airfoil shape between a leading edge and a trailing edge; a platform disposed between the base end and the airfoil; Equipped with the platform has a groove portion recessed from the end surface on the trailing edge side toward the leading edge side and extending in the circumferential direction of the rotor, a bottom of the groove overlaps with at least the airfoil portion when viewed in the radial direction; When the end of the platform on the ventral side of the bottom is defined as a first point, a tangent to the bottom extending along a plane intersecting the radial direction at the first point is defined as a first tangent; a second tangent is a tangent to the bottom portion that extends along the surface at a second point where a line segment connecting a trailing edge side end of a serpentine cooling flow passage received inside the airfoil portion when viewed from the radial direction and the trailing edge side end of the airfoil portion intersects with the bottom portion; When the intersection of the first tangent line and the second tangent line when viewed from the radial direction is defined as a third point, When viewed from the radial direction, the third point is located on the opposite side of the straight line connecting the first point and the second point from the end portion of the airfoil portion on the trailing edge side.

[0007] (2) A gas turbine according to at least one embodiment of the present disclosure includes: The rotor; a gas turbine rotor blade having the configuration (1) above, the base end of which is fixed to the rotor; Equipped with. [Effects of the Invention]

[0008] According to at least one embodiment of the present disclosure, thermal stresses occurring in a gas turbine rotor blade can be effectively suppressed. [Brief explanation of the drawings]

[0009] [Figure 1] 1 is a schematic configuration diagram of a gas turbine to which a gas turbine rotor blade according to some embodiments is applied. [Figure 2] 1 is a view of a rotor blade (gas turbine rotor blade) according to some embodiments, viewed from the suction side. FIG. [Figure 3A] FIG. 3 is a diagram showing an example of a cross section AA in FIG. 2. [Figure 3B] FIG. 3 is a diagram showing another example of the cross section taken along the line AA in FIG. 2. [Figure 3C] FIG. 3 is a diagram showing yet another example of the cross section taken along the line AA in FIG. 2. [Figure 3D] FIG. 3 is a diagram showing yet another example of the cross section taken along the line AA in FIG. 2. [Figure 4] FIG. 10 is a diagram for explaining the shape of a groove portion. DETAILED DESCRIPTION OF THE INVENTION

[0010] Hereinafter, several embodiments of the present disclosure will be described with reference to the accompanying drawings. However, the dimensions, materials, shapes, relative arrangements, etc. of components described as embodiments or shown in the drawings are merely illustrative examples and are not intended to limit the scope of the present disclosure. For example, expressions expressing relative or absolute arrangement such as "in a certain direction," "along a certain direction," "parallel," "orthogonal," "center," "concentric," or "coaxial" not only express such an arrangement exactly, but also express a state in which there is a relative displacement with a tolerance or an angle or distance to the extent that the same function is obtained. For example, expressions such as "identical," "equal," and "homogeneous" that indicate that something is in an equal state not only indicate a state of strict equality, but also indicate a state in which there is a tolerance or a difference to the extent that the same function is obtained. For example, expressions representing shapes such as a square shape or a cylindrical shape not only represent shapes such as a square shape or a cylindrical shape in the strict geometric sense, but also represent shapes including uneven portions, chamfered portions, etc., to the extent that the same effect can be obtained. On the other hand, the expressions "comprise," "include," "have," "includes," or "have" of one element are not exclusive expressions that exclude the presence of other elements.

[0011] (Gas Turbine 1) First, a gas turbine to which gas turbine rotor blades according to some embodiments are applied will be described.

[0012] Fig. 1 is a schematic configuration diagram of a gas turbine to which gas turbine rotor blades according to some embodiments are applied. As shown in Fig. 1, the gas turbine 1 includes a compressor 2 for generating compressed air, a combustor 4 for generating combustion gas using the 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.

[0013] The compressor 2 includes a plurality of stator vanes 16 fixed to the compressor casing 10 side, and a plurality of rotor blades 18 implanted in the rotor 8 so as to be arranged alternately with respect to the stator vanes 16 . Air taken in through an air intake 12 is sent to the compressor 2, and this air is compressed as it passes through a plurality of stator vanes 16 and a plurality of rotor blades 18, becoming high-temperature, high-pressure compressed air.

[0014] The combustor 4 is supplied with fuel and compressed air generated by the compressor 2, and the fuel is combusted in the combustor 4 to generate combustion gas, which is the working fluid of the turbine 6. As shown in FIG. 1 , a plurality of combustors 4 may be arranged in the casing 20 along the circumferential direction around the rotor 8.

[0015] The turbine 6 has a combustion gas flow path 28 formed in the turbine casing 22 and includes a plurality of stator vanes 24 and rotor blades 26 provided in the combustion gas flow path 28 . The stator vanes 24 are fixed to the turbine casing 22 side, and a plurality of the stator vanes 24 arranged along the circumferential direction of the rotor 8 constitute a stator vane row. The moving blades 26 are implanted in the rotor 8, and a plurality of the moving blades 26 arranged along the circumferential direction of the rotor 8 constitute a moving blade row. The stator vane rows and moving blade rows are arranged alternately in the axial direction of the rotor 8. In the turbine 6, the combustion gas from the combustor 4 flows into a combustion gas flow path 28 and passes through the plurality of stator vanes 24 and the plurality of rotor blades 26, thereby driving the rotor 8 to rotate, which in turn drives a generator connected to the rotor 8 to generate electricity. After driving the turbine 6, the combustion gas is discharged to the outside via an exhaust chamber 30.

[0016] In some embodiments, the blades 26 of the turbine 6 may be gas turbine blades 40, described below.

[0017] (Gas turbine rotor blade 40) FIG. 2 is a view of a rotor blade 26 (gas turbine rotor blade 40) according to some embodiments, as viewed from the suction side. FIG. 3A is a diagram showing an example of a cross section taken along line AA in FIG. FIG. 3B is a diagram showing another example of the cross section taken along line AA in FIG. FIG. 3C is a diagram showing yet another example of the cross section taken along line AA in FIG. FIG. 3D is a diagram showing yet another example of the cross section taken along line AA in FIG. FIG. 4 is a diagram for explaining the shape of the groove 70, taking the groove 70 shown in FIG. 3A as an example. The AA cross section shown in Figures 3A, 3B, 3C, and 3D represents a cross section of the blade portion 42 at the outer end of the fillet portion 36 described later in the radial direction (hereinafter also simply referred to as the "radial direction") of the rotor 8.

[0018] 2, 3A, 3B, 3C, and 3D, a rotor blade 26, which is a gas turbine rotor blade 40 according to some embodiments, includes an airfoil portion (airfoil portion) 42, a platform 32, and a blade root portion (base end portion) 34. The blade root portion 34 is embedded in the rotor 8 (see FIG. 1), and the rotor blade 26 rotates together with the rotor 8. The platform 32 is configured integrally with the blade root portion 34.

[0019] The blade portion 42 is arranged to extend along the radial direction of the rotor 8 (hereinafter also simply referred to as the "radial direction") and has a base end 50 fixed to the platform 32 and a tip end 48 located on the opposite side of the base end 50 in the blade height direction (radial direction of the rotor 8). Furthermore, the blade portion 42 of the rotor blade 26 has a leading edge 44 and a trailing edge 46 from the base end 50 to the tip 48, and the blade surface of the blade portion 42 includes a pressure surface (pressure surface) 56 and a suction surface (suction surface) 58 extending along the blade height direction (radial direction) between the base end 50 and the tip 48.

[0020] 3A, 3B, 3C, and 3D, a cooling flow passage 60 extending along the blade height direction of the blade 42 is provided inside the blade 42. A cooling fluid (e.g., air) for cooling the gas turbine rotor blade 40 flows through the cooling flow passage 60. By supplying the cooling fluid to the cooling flow passage 60, the blade 42, which is provided in the combustion gas flow passage 28 of the turbine 6 and exposed to high-temperature combustion gas, is cooled. In some embodiments, the cooling channel 60 extends across a span of the airfoil 42 and at least a portion of the platform 32 . The gas turbine rotor blade 40 may have a plurality of cooling channels 60. The cooling channels 60 may further extend across the blade root 34.

[0021] A fillet portion 36 is formed at a base end portion 51, which is a portion on the base end 50 side of the wing portion 42. The wing portion 42 is connected to the platform 32 via the fillet portion 36.

[0022] As shown in Figures 2, 3A, 3B, 3C, and 3D, in the rotor blade 26 according to some embodiments, the platform 32 has a groove 70 that is recessed from the end face 32a on the trailing edge 46 side toward the leading edge 44 side and extends in the circumferential direction of the rotor 8 (hereinafter also simply referred to as the "circumferential direction"). In the rotor blade 26, a temperature difference is likely to occur between the blade portion 42 and the platform 32 during a transient state, such as when the gas turbine 1 starts or stops operating, and thermal stress is likely to occur. It has been found that this thermal stress is particularly likely to become large near the root portion where the trailing edge 46 of the blade portion 42 joins the platform 32. For this reason, a groove 70 is formed in the platform 32 so as to be recessed from the end face 32a on the trailing edge 46 side of the platform 32 toward the leading edge 44 side, and extends in the circumferential direction of the rotor 8, thereby reducing the above-mentioned thermal stress. The grooves 70 will be described in detail later.

[0023] 2, 3A, 3B, 3C, and 3D, in the rotor blade 26 according to some embodiments, the platform 32 has a seal pin groove 81 in which a seal pin (not shown) is disposed to seal a gap between the platform 32 and the platform 32 of another circumferentially adjacent rotor blade 26. In the rotor blade 26 according to some embodiments, the seal pin groove 81 is formed in the end face 32b of the platform 32 on the suction surface 58 side (suction side). In some embodiments of the rotor blade 26, the end face 32c of the platform 32 on the pressure surface 56 side (pressure side) includes a flat surface 32p that can abut against a seal pin (not shown) that is positioned opposite the end face 32c.

[0024] (Regarding the groove 70) As described above, the groove 70 is formed so as to be recessed from the end face 32a of the platform 32 on the trailing edge 46 side toward the leading edge 44 side. In addition, the trailing edge 46 of the wing portion 42 is provided close to the end face 32a of the platform 32 on the trailing edge 46 side. Therefore, the bottom 71 of the groove 70 is formed at a position recessed from the end face 32a on the trailing edge 46 side of the platform 32 toward the leading edge 44 side, relative to the end (trailing edge end 46a) on the trailing edge 46 side of the wing portion 42. In some embodiments of the rotor blade 26, the bottom 71 of the groove 70 is the position closest to the leading edge 44 in the cross section of the rotor blade 26 when viewed circumferentially, i.e., the position closest to the upstream side of the turbine 6 along the axial direction of the rotor 8 (hereinafter also simply referred to as the "axial direction").

[0025] Thermal stress acting on the blade 42 in a transient state such as when the gas turbine 1 starts or stops operation tends to be particularly large near the root portion where the trailing edge 46 of the blade 42 joins the platform 32 . As a result of careful investigations by the inventors, it was found that in order to effectively suppress thermal stress near the root portion between the trailing edge 46 of the wing portion 42 and the platform 32, which tends to be particularly subject to large thermal stress, it is effective to form the groove portion 70 so that the groove portion 70 is located directly below (radially inward from) the trailing edge end 46a of the wing portion 42, thereby suppressing the strength of the area of ​​the platform 32 that restrains the root portion.

[0026] However, in the rotor blade 26, a serpentine cooling channel (cooling channel 60) is formed inside the blade portion 42 to cool the blade portion 42. As described above, the cooling channel 60 extends over a range in the blade height direction that includes at least a portion of the platform 32. Therefore, the groove 70 formed in the platform 32 needs to be formed at a position that ensures a wall thickness between the groove 70 and the cooling channel 60. Therefore, there is a limit to the depth of the groove 70.

[0027] As shown in Figures 3A, 3B, 3C, and 3D, the vicinity of the trailing edge 46 of the wing portion 42 is the portion of the wing portion 42 that is closest to the end face 32a of the platform 32 on the trailing edge 46 side and the end face 32c of the platform 32 on the pressure surface 56 side. 3A, 3B, 3C, and 3D, the cooling passage 60 closest to the end face 32a of the platform 32 on the trailing edge 46 side is the cooling passage 61 closest to the trailing edge 46. The vicinity of the end face (trailing edge end 61a) of this cooling passage 61 on the trailing edge 46 side is the portion closest to the end face 32a of the platform 32 on the trailing edge 46 side and the end face 32c of the platform 32 on the pressure surface 56 side.

[0028] Therefore, in some embodiments of the rotor blade 26, taking into consideration the position of the trailing edge 46 of the blade portion 42 described above and the position of the cooling passage 61 closest to the trailing edge 46, the groove portion 70 is formed so that the depth of the groove portion 70 is relatively deep near the trailing edge end 46a of the blade portion 42 when viewed radially and relatively shallow at a position circumferentially away from the trailing edge end 46a of the blade portion 42, as shown in Figures 3A, 3B, 3C, and 3D. Specifically, groove portion 70 includes a relatively shallow dorsal region 72 located circumferentially away from trailing edge 46a of wing portion 42, a relatively deep ventral region 73 located near trailing edge 46a of wing portion 42, and an intermediate region 74 connecting dorsal region 72 and ventral region 73.

[0029] More specifically, as shown in FIG. 4, the bottom 71 of the groove 70 overlaps at least the wing portion 42 when viewed in the radial direction. When the end 71a of the bottom 71 on the pressure surface 56 side of the platform 32 is defined as the first point P1, the tangent to the bottom 71 at the first point P1 and extending along a plane PL (for example, a plane corresponding to the paper surface in Figure 4) that intersects with the radial direction is defined as the first tangent Lt1. When viewed from the radial direction, the intersection of the line segment Ls connecting the trailing edge 61a of the cooling channel 61 closest to the trailing edge 46, received inside the blade portion 42, and the trailing edge 46a of the blade portion 42 with the bottom 71 is defined as a second point P2. A tangent to the bottom 71 at the second point P2 and extending along the plane PL is defined as a second tangent Lt2. When viewed from the radial direction, the intersection of the first tangent Lt1 and the second tangent Lt2 is defined as a third point P3. When viewed from the radial direction, the third point P3 is located on the opposite side of the straight line SL connecting the first point P1 and the second point P2 from the trailing edge 46a of the blade portion 42. The above conditions are met by any of the grooves 70 shown in FIGS. 3A, 3B, 3C, and 3D.

[0030] Note that the first point P1 is the intersection of the bottom 71 and the end face 32c on the pressure surface 56 side of the platform 32 when viewed from the radial direction, and therefore strictly speaking the first tangent line Lt1 is not uniquely determined. Therefore, the first point P1 is defined as a position that is very close to the intersection and is circumferentially away from the chamfered or lightly chamfered portion at the intersection to a position that is not affected by the chamfering or lightly chamfering.

[0031] By configuring the groove 70 as described above, the bottom 71 is formed at a position that is recessed from the end face 32a on the trailing edge 46 side of the platform 32 toward the leading edge 44 relative to the trailing edge 46a of the blade 42. Therefore, the groove 70 is located at a position that overlaps with the trailing edge 46a of the blade 42 when viewed from the radial direction, which reduces the strength of the platform 32 directly below (radially inside) the trailing edge 46a of the blade 42 and effectively reduces thermal stress that occurs near the trailing edge 46a of the blade 42 when the gas turbine 1 is in a transient state.

[0032] Furthermore, by configuring groove portion 70 as described above, bottom portion 71 is formed so as to move from second point P2 toward negative pressure surface 58 toward end face 32a on the trailing edge 46 side of platform 32, that is, toward the axial downstream side of rotor 8. Therefore, bottom portion 71 is prevented from approaching cooling flow passage 61 as it moves from second point P2 toward negative pressure surface 58, making it easier to ensure the wall thickness between groove portion 70 and cooling flow passage 61. Therefore, according to some embodiments of the rotor blade 26, the wall thickness between the groove portion 70 and the cooling flow passage 61 can be ensured while effectively reducing the thermal stress that occurs near the trailing edge 46a of the blade portion 42 during transient conditions of the gas turbine 1. Furthermore, in a gas turbine 1 equipped with a rotor blade 26 having a groove portion 70 configured as described above, the wall thickness between the groove portion 70 and the cooling flow passage 61 can be ensured while effectively reducing thermal stress that occurs near the trailing edge 46a of the blade portion 42 in a transient state of the gas turbine 1, thereby improving the durability of the gas turbine 1.

[0033] In the rotor blade 26 according to some embodiments, when viewed in the radial direction, the bottom portion 71 may intersect with the suction surface 58 and the pressure surface 56 outside a first imaginary circle Cv1 that is centered on the trailing edge 46a of the airfoil portion 42 and passes through the line segment Ls. This condition is satisfied in all of the groove portions 70 shown in Figures 3A, 3B, 3C, and 3D. The radius of the first imaginary circle Cv1 can be determined, for example, by stress analysis or the like, as a size that can effectively reduce thermal stress occurring near the trailing edge 46a of the blade 42 when the gas turbine 1 is in a transient state. This makes it possible to efficiently reduce the strength of the platform 32 directly below (radially inwardly of) the trailing edge 46a of the blade portion 42.

[0034] In the rotor blade 26 according to some embodiments, when viewed in the radial direction, the bottom 71 may intersect with the suction surface 58 outside a second imaginary circle Cv2 that is centered on the trailing edge 61a of the cooling channel 61 and passes through the line segment Ls. This condition is satisfied in all of the grooves 70 shown in Figures 3A, 3B, 3C, and 3D. The radius of the second imaginary circle Cv2 may be, for example, the thickness required as the wall thickness between the groove portion 70 and the cooling flow passage 61. This ensures that the wall thickness between the groove portion 70 and the cooling flow passage 61 is at least equal to the radius of the second imaginary circle Cv2.

[0035] In the rotor blade 26 according to some embodiments, the depth dp of the groove 70 recessed from the end face on the trailing edge 46 side (i.e., the end face 32a of the platform 32 on the trailing edge 46 side) toward the leading edge 44 side may be deeper on the pressure surface 56 side than on the suction surface 58 side. Specifically, the depth dp of the pressure side region 73 may be deeper than the depth dp of the suction side region 72. This condition is satisfied in all of the grooves 70 shown in Figures 3A, 3B, 3C, and 3D. As a result, the depth dp of the groove 70 (i.e., the suction side region 72) at a position circumferentially away from the trailing edge 46a toward the suction surface 58, which contributes less to reducing the strength of the platform 32 immediately below (radially inward from) the trailing edge 46a of the airfoil 42, becomes relatively shallow. Since the groove 70 is generally formed by electrical discharge machining, a shallower depth of the groove 70 can reduce processing costs. Therefore, according to the rotor blade 26 of some embodiments, it is possible to provide a groove portion 70 that can effectively reduce the strength of the platform 32 directly below (radially inwardly of) the trailing edge 46a of the blade portion 42 while reducing processing costs. Furthermore, since the depth dp of the suction side region 72 is relatively shallow, interference between the groove portion 70 and the seal pin groove 81 is unlikely, and it becomes easier to form the seal pin groove 81 in the end face 32b on the negative pressure surface 58 side of the platform 32. As a result, it is not necessary to provide the seal pin groove 81 in the end face 32c on the pressure surface 56 side of the platform 32, and it becomes easier to make the depth dp of the ventral side region 73 relatively deep.

[0036] In the rotor blade 26 according to some embodiments, the depth dp of the groove 70 may be constant at a position farther from the trailing edge 46a of the airfoil 42 than the intersection position P4 between the bottom 71 and the suction surface 58 as viewed in the radial direction. Specifically, the depth dp may be constant in at least a portion of the suction side region 72. This condition is satisfied by the groove 70 shown in FIGS. 3A and 3B. This simplifies the shape of the grooves 70, and reduces the processing costs for the grooves 70.

[0037] In the rotor blade 26 according to some embodiments, the depth dp of the groove 70 may be constant in at least a portion of the region between the first point P1 and the intersection position P5 of the bottom 71 and the pressure surface 56 as viewed from the radial direction. Specifically, the depth dp of the pressure side region 73 may be constant in at least a portion of the region. Note that this condition is satisfied by the groove 70 shown in FIGS. 3A, 3C, and 3D. This makes it easier to avoid interference between the groove 70 and a seal pin (not shown) that seals the gap between the platform 32 of another circumferentially adjacent rotor blade 26, while ensuring the depth of the groove 70 on the ventral side (pressure surface 56 side).

[0038] In the groove 70 shown in FIGS. 3A, 3B, 3C, and 3D, the bottom 71 extends linearly when viewed in the radial direction, but may extend in a curved line.

[0039] In the rotor blade 26 according to some embodiments, as described above, the seal pin groove 81 may be formed in the end face 32b of the platform 32 on the suction surface 58 side. The end face 32c of the platform 32 on the pressure surface 56 side may include a flat surface 32p that can come into contact with a seal pin (not shown) that is arranged at a position opposite the end face 32c. This makes it easier to ensure a flat area (flat surface 32p) on the end face 32c of the platform 32 on the pressure surface 56 side. This makes it easier to deepen the groove 70 in the vicinity of the end face 32c of the platform 32 on the pressure surface 56 side. This makes it easier to reduce the strength of the platform 32 directly below (radially inward) the trailing edge 46a of the blade portion 42.

[0040] The present disclosure is not limited to the above-described embodiments, but also includes modifications to the above-described embodiments and appropriate combinations of these modifications. For example, the above-described effects can be achieved even with the rotor blade 26 having the groove portion 70 in which the suction side region 72, the pressure side region 73, and the intermediate region 74 shown in Figures 3A, 3B, 3C, and 3D are appropriately combined.

[0041] The contents described in each of the above embodiments can be understood, for example, as follows. (1) A gas turbine rotor blade 40 (rotor blade 26) according to at least one embodiment of the present disclosure includes a base end (root portion 34) fixed to a rotor 8, an airfoil portion (airfoil portion 42) extending in the radial direction of the rotor 8 and having pressure side and suction side surfaces that form an airfoil shape between a leading edge 44 and a trailing edge 46, and a platform 32 provided between the base end (root portion 34) and the airfoil portion (airfoil portion 42). The platform 32 is recessed from an end face 32a on the trailing edge 46 side toward the leading edge 44 side and has a groove portion 70 extending in the circumferential direction of the rotor 8. A bottom portion 71 of the groove portion 70 overlaps at least the airfoil portion (airfoil portion 42) when viewed in the radial direction. When the end 71a of the bottom 71 on the ventral side (pressure surface 56 side) of the platform 32 is defined as a first point P1, a tangent to the bottom 71 at the first point P1 and extending along a plane PL intersecting the radial direction is defined as a first tangent Lt1. When the intersection of the bottom 71 and a line segment Ls connecting the end (trailing edge end 61a) of the serpentine cooling channel (cooling channel 61) received inside the airfoil section (blade section 42) on the trailing edge 46 side (trailing edge end 46a) of the blade section 42 as viewed from the radial direction is defined as a second point P2. When the intersection of the line segment Ls and the bottom 71 as viewed from the radial direction is defined as a second point P2, a tangent to the bottom 71 at the second point P2 and extending along the plane PL is defined as a second tangent Lt2. When viewed from the radial direction, a third point P3 is defined as the intersection of the first tangent Lt1 and the second tangent Lt2. When viewed from the radial direction, the third point P3 is located on the opposite side of the straight line SL connecting the first point P1 and the second point P2 from the end portion (trailing edge 46a) of the blade portion 42 on the trailing edge 46 side.

[0042] According to the configuration (1) above, the bottom portion 71 is formed at a position recessed from the end face 32a of the platform 32 on the trailing edge 46 side toward the leading edge 44 side relative to the end portion (trailing edge end 46a) on the trailing edge 46 side of the airfoil portion (blade portion 42). Therefore, when viewed from the blade height direction (radial direction), the groove portion 70 is present at a position overlapping with the end portion (trailing edge end 46a) on the trailing edge 46 side of the airfoil portion (blade portion 42). This reduces the strength of the platform 32 directly below (radially inward) the end portion (trailing edge end 46a) on the trailing edge 46 side of the airfoil portion (blade portion 42), and effectively reduces thermal stress occurring in the vicinity of the end portion (trailing edge end 46a) on the trailing edge 46 side of the airfoil portion (blade portion 42) in a transient state of the gas turbine 1. Furthermore, according to the configuration (1) above, the bottom 71 is formed so as to move from the second point P2 toward the blade surface (suction surface 58) on the suction side toward the end face 32a on the trailing edge 46 side of the platform 32, that is, toward the axial downstream side of the rotor 8. Therefore, the bottom 71 is prevented from approaching the serpentine cooling channel (cooling channel 61) as it moves from the second point P2 toward the blade surface (suction surface 58) on the suction side, making it easier to ensure the wall thickness between the groove portion 70 and the serpentine cooling channel (cooling channel 61). Therefore, according to the configuration (1) above, the wall thickness between the groove portion 70 and the serpentine cooling flow passage (cooling flow passage 61) can be ensured, while effectively reducing the thermal stress that occurs near the end portion (trailing edge end 46a) on the trailing edge 46 side of the airfoil portion (airfoil portion 42) during a transient state of the gas turbine 1.

[0043] (2) In some embodiments, in the configuration of (1) above, when viewed from the radial direction, the bottom 71 may be centered on the end portion (trailing edge 46a) on the trailing edge 46 side of the airfoil portion (airfoil portion 42) and intersect with the suction side airfoil surface (suction surface 58) and the pressure side airfoil surface (pressure surface 56) outside the first imaginary circle Cv1 that passes through the line segment Ls.

[0044] According to the configuration (2) above, the strength of the platform 32 directly below (radially inward) the end portion (trailing edge end 46a) on the trailing edge 46 side of the airfoil portion (airfoil portion 42) can be efficiently reduced.

[0045] (3) In some embodiments, in the configuration of (1) or (2) above, when viewed from the radial direction, the bottom 71 may be centered on the end (trailing edge end 61a) of the serpentine cooling channel (cooling channel 61) on the trailing edge 46 side and intersect with the suction-side blade surface (suction surface 58) outside a second imaginary circle Cv2 that passes through the line segment Ls.

[0046] According to the above configuration (3), it is possible to ensure a wall thickness between the groove portion 70 and the serpentine cooling flow passage (cooling flow passage 61) that is at least the radius of the second imaginary circle Cv2.

[0047] (4) In some embodiments, in any of the configurations (1) to (3) above, the depth dp of the groove portion 70 recessed from the end face on the trailing edge 46 side (i.e., the end face 32a on the trailing edge 46 side of the platform 32) toward the leading edge 44 side may be deeper on the ventral side (the pressure surface 56 side) than on the suction side (the negative pressure surface 58 side).

[0048] According to the configuration (4) above, the depth dp of the groove 70 at a position circumferentially away from the end (trailing edge 46a) toward the suction side (negative pressure surface 58 side) of the trailing edge 46 side of the airfoil portion (airfoil portion 42) is relatively shallow, which contributes little to reducing the strength of the platform 32 directly below (radially inward) of the end (trailing edge 46a). Since the groove 70 is generally formed by electrical discharge machining, a shallower depth of the groove 70 can reduce processing costs. According to the configuration (4) above, a groove portion 70 can be provided that can effectively reduce the strength of the platform 32 directly below (radially inward) the end portion (trailing edge end 46a) on the trailing edge 46 side of the airfoil portion (airfoil portion 42) while reducing processing costs.

[0049] (5) In some embodiments, in any of the configurations (1) to (4) above, the depth dp of the groove 70 recessed from the end face on the trailing edge 46 side (i.e., the end face 32a on the trailing edge 46 side of the platform 32) toward the leading edge 44 side may be constant at a position farther from the end face on the trailing edge 46 side (trailing edge end 46a) of the airfoil portion (airfoil portion 42) than the intersection position P4 between the bottom 71 and the suction side airfoil surface (suction surface 58) when viewed from the radial direction.

[0050] According to the above configuration (5), the shape of the groove 70 can be simplified, and the processing cost of the groove 70 can be reduced.

[0051] (6) In some embodiments, in any of the configurations (1) to (5) above, the depth dp of the groove 70 recessed from the end face on the trailing edge 46 side (i.e., the end face 32a on the trailing edge 46 side of the platform 32) toward the leading edge 44 side may include the first point P1 and may be constant in at least a portion of the region between the first point P1 and the intersection position P5 of the bottom 71 and the ventral blade surface (pressure surface 56) when viewed radially.

[0052] According to the above configuration (6), it is possible to easily avoid interference between the groove 70 and the seal pin for sealing the gap between the platform 32 of another circumferentially adjacent gas turbine rotor blade 40 (rotor blade 26), while ensuring the depth dp of the groove 70 on the pressure side (pressure surface 56).

[0053] (7) In some embodiments, in any of the configurations (1) to (6) above, the platform 32 may have a seal pin groove 81 in which a seal pin is disposed to seal a gap between the platform 32 and the platform 32 of another circumferentially adjacent gas turbine rotor blade 40 (rotor blade 26). The seal pin groove 81 may be formed in an end face 32b on the suction side (negative pressure surface 58 side) of the platform 32. An end face 32c on the pressure side (pressure surface 56) of the platform 32 may include a flat surface 32p that can abut against a seal pin disposed at a position opposite the end face 32c.

[0054] The configuration of (7) above makes it easier to ensure a flat area (flat surface 32p) on the end face 32c on the ventral side (pressure surface 56) of the platform 32. This makes it easier to deepen the groove 70 in the vicinity of the end face 32c on the ventral side (pressure surface 56) of the platform 32. This makes it easier to reduce the strength of the platform 32 directly below (radially inward) the end portion (trailing edge end 46a) on the trailing edge 46 side of the airfoil portion (airfoil portion 42).

[0055] (8) A gas turbine 1 according to at least one embodiment of the present disclosure includes a rotor 8 and a gas turbine rotor blade 40 (rotor blade 26) having any of the configurations (1) to (7) above, which is fixed to the rotor 8 at its base end (root portion 34).

[0056] According to the configuration (8) above, the wall thickness between the groove portion 70 and the serpentine cooling flow passage (cooling flow passage 61) can be ensured while effectively reducing the thermal stress that occurs near the end portion (trailing edge end 46a) on the trailing edge 46 side of the airfoil portion (airfoil portion 42) during a transient state of the gas turbine 1, thereby improving the durability of the gas turbine 1. [Explanation of symbols]

[0057] 1. Gas turbine 8 rotors 26 Moving blade 32 Platform 32a, 32b, 32c end face 32p plane 34 Wing root (base end) 40 Gas turbine blade 42 Wing section (airfoil section) 44 leading edge 46 Trailing edge 46a Trailing edge 56 Pressure surface (ventral surface) 58 Negative pressure surface (back) 60, 61 Cooling passage (serpentine cooling passage) 61a Trailing edge 70 Groove 71 Bottom 81 Seal pin groove

Claims

1. a base end portion fixed to the rotor; an airfoil portion extending radially of the rotor and having pressure and suction surfaces defining an airfoil shape between a leading edge and a trailing edge; a platform disposed between the base end and the airfoil; Equipped with the platform has a groove portion recessed from the end surface on the trailing edge side toward the leading edge side and extending in the circumferential direction of the rotor, a bottom of the groove overlaps with at least the airfoil portion when viewed in the radial direction; When the end of the platform on the ventral side of the bottom is defined as a first point, a tangent line that is a tangent line to the bottom at the first point and extends along a plane that intersects with the radial direction is defined as a first tangent line; a second tangent is a tangent to the bottom at the second point, the second tangent being a tangent to the bottom at the second point and extending along the surface, when a line segment connecting an end of the serpentine cooling flow passage received inside the airfoil portion on the trailing edge side and an end of the airfoil portion on the trailing edge side is defined as a second point; When the intersection of the first tangent line and the second tangent line as viewed from the radial direction is defined as a third point, When viewed from the radial direction, the third point is located on the opposite side of the trailing edge side end of the airfoil portion across the straight line connecting the first point and the second point. Gas turbine rotor blades.

2. When viewed from the radial direction, the bottom portion intersects with the suction-side blade surface and the pressure-side blade surface at a position outside a first imaginary circle that has a center at the end portion of the trailing edge side of the airfoil portion and passes through the line segment. The gas turbine rotor blade according to claim 1 .

3. when viewed from the radial direction, the bottom portion intersects with the suction-side blade surface outside a second imaginary circle that has a center at an end portion of the serpentine cooling channel on the trailing edge side and passes on the line segment, The gas turbine rotor blade according to claim 1 or 2.

4. a depth of the groove portion recessed from the end surface on the trailing edge side toward the leading edge side is deeper on the ventral side than on the dorsal side; The gas turbine rotor blade according to claim 1 .

5. a depth of the groove recessed from the trailing edge side end surface toward the leading edge side is constant at a position farther from the trailing edge side end of the airfoil portion than an intersection position between the bottom portion and the suction side blade surface when viewed from the radial direction. The gas turbine rotor blade according to claim 1 .

6. a depth of the groove portion recessed from the end surface on the trailing edge side toward the leading edge side includes the first point and is constant in at least a part of a region between the first point and an intersection position of the bottom portion and the blade surface on the pressure side as viewed from the radial direction, The gas turbine rotor blade according to claim 1 .

7. the platform has a seal pin groove in which a seal pin is disposed for sealing a gap between the platform and a platform of another gas turbine rotor blade adjacent to the platform in the circumferential direction, The seal pin groove is formed in the back end surface of the platform, an end surface on the ventral side of the platform includes a flat surface that can come into contact with a seal pin disposed at a position opposite the end surface; The gas turbine rotor blade according to claim 1 .

8. The rotor; a gas turbine rotor blade according to claim 1 fixed at the base end to the rotor; Equipped with Gas turbine.

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