JP2013015062A - Gas turbine blade - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a gas turbine blade having a high performance blade cooling structure applicable to the blade regardless of the size of the blade.SOLUTION: The gas turbine blade includes: a cooling channel provided inside a gas turbine moving blade; a ventral rib 2a that is hung from a leading edge 51 side to a trailing edge side with respect to a ventral wall surface in the cooling channel and of which an end 72a on the trailing edge side is inclined to be positioned downstream in a flow direction of a cooling medium 10 than an end 71a on the leading edge side; a back side rib 2b that is hung from the leading edge side to the trailing edge side with respect to a back side surface wall in the cooling channel and of which an end 72b on the trailing edge side is inclined to be positioned downstream in the flow direction of the cooling medium than an end 71b on the leading edge side; and film cooling holes 1a, 1b, and 1c for allowing the cooling medium flowing in the cooling channel to flow out the outside of the gas turbine blade.

Description

  The present invention relates to a gas turbine blade used in a gas turbine.

  In a gas turbine, compressed air compressed by a compressor is mixed with fuel in a combustor and burned to generate a high-temperature and high-pressure working medium (combustion gas). The working medium is divided into a plurality of turbine blades and turbine stationary blades. It is a heat engine which guides to a turbine provided with, and drives the turbine to obtain rotational power (kinetic energy) of the turbine. Accordingly, since the surface of the gas turbine blade is exposed to a high-temperature working medium, it is necessary to forcibly cool the blade to prevent the blade material from being hot-corroded and the structural strength is lowered and ensure soundness.

  The gas turbine blade cooling technology includes (1) an internal cooling method in which a cooling medium is forcedly convected in a cooling flow path provided inside the blade, and (2) pores (film cooling holes) provided on the blade surface. There is a film cooling (external cooling) method that suppresses heat input from a working medium by ejecting a cooling medium from the inside of the blade and covering the blade surface with the cooling medium.

  In recent years, the combustion temperature has risen in order to improve the thermal efficiency of gas turbines. In cooling of moving blades, there is a cooling structure called shower head cooling in which multiple rows of film cooling holes are arranged on the leading edge of the blade where the heat load is high. It has been adopted. The following patent documents are disclosed as prior art relating to the cooling structure of the moving blade leading edge portion.

  US Pat. No. 5,387,086 shows a cooling structure in which two meandering channels for convection cooling are provided inside a moving blade, and a cavity having a film hole is provided at the front edge. An impingement cooling hole is provided in the flow path on the most downstream side (blade leading edge) side of the meandering flow path on the leading edge side, and the air (cooling medium) flowing through the meandering flow path and cooling the inside of the blade is the cooling hole. From the inner surface, impingement cooling the blade leading edge from the inner surface. Thereafter, the cooling air in the cavity is ejected from the film cooling hole, and the outer surface is covered from the blade leading edge to suppress heat input from the high temperature working medium.

  In JP-A-11-257005, as a method of internal cooling that is not impingement cooling, a meandering channel for convection cooling is provided inside a moving blade, a turbulator (rib) is provided on the inner wall of the meandering channel, A cooling structure in which film cooling holes are arranged between turbulators is shown. In this gas turbine blade, the relative position between the turbulator and the film cooling hole is defined, the generation of the peeling area on the downstream side of the turbulator is suppressed, the low heat transfer area is eliminated, and the internal cooling performance in the convection cooling channel is improved. It has become.

US Pat. No. 5,387,086 JP 11-257005 A

  However, in the case of a relatively small blade such as a small gas turbine, a large number of cooling channels cannot be provided inside the blade, and a cavity for impingement cooling as shown in Patent Document 1 is provided. Is difficult. That is, with this type of cooling structure, it is difficult to ensure internal cooling performance depending on the size of the gas turbine blade.

  On the other hand, in the structure of Patent Document 2 that does not include a cavity for impingement cooling, the internal cooling performance can be improved by suppressing separation by defining the relative positions of the turbulator and the film cooling hole in the convection cooling flow path. Compared with impingement cooling, the internal cooling performance is still inferior. Therefore, for cooling the entire blade, it is necessary to enhance external cooling (film cooling) by increasing the amount of cooling air. If the amount of cooling air is increased when the compressed air extracted from the compressor is used as cooling air, the gas turbine efficiency may decrease.

  In addition, although the rib (turbulator) is provided in the cooling flow path of the gas turbine blade of patent document 1 and patent document 2, the said rib promotes the generation | occurrence | production of the turbulent flow in a cooling flow path, and is only internal cooling. This is intended to improve performance.

  An object of the present invention is to provide a gas turbine blade having a high-performance blade cooling structure that can be applied regardless of the blade size.

  In order to achieve the above object, the present invention provides a cooling channel provided inside a gas turbine blade for flowing a cooling medium, and a leading edge side with respect to a wall surface on the ventral side of the gas turbine blade in the cooling channel. A rib extending from the rear edge side to the rear edge side, wherein the end part on the rear edge side is inclined so that the end part on the rear edge side is located downstream of the front edge side in the flow direction of the cooling medium, and the cooling A rib extending from the front edge side to the rear edge side with respect to the wall surface on the back side of the gas turbine blade in the flow path, wherein the end portion on the rear edge side is more in the flow direction of the cooling medium than the end portion on the front edge side And a film cooling hole for allowing the cooling medium flowing through the cooling flow path to flow out of the gas turbine blades.

  According to the present invention, since the cooling medium is hardly peeled off from the blade surface and the film cooling efficiency is improved, the blade cooling performance can be improved regardless of the blade size.

The expanded view along CC cross section of the gas turbine rotor blade which concerns on embodiment of this invention. Sectional drawing in the DD cross section of the gas turbine rotor blade which concerns on embodiment of this invention. The structural diagram of the rib 2 and the film cooling hole 1c which were provided in the inner wall of the flow path 41 which concerns on embodiment of this invention. The AA sectional view which followed the ventral film cooling hole 1a about channel 41c concerning an embodiment of the invention. BB sectional drawing along the back side film cooling hole 1b about the flow path 41c which concerns on embodiment of this invention. Sectional drawing which expanded the flow path 41c part of FIG.

  Hereinafter, embodiments of the present invention will be described with reference to the drawings. FIG. 1 is a developed view along a CC section (see FIG. 2) of a gas turbine blade according to an embodiment of the present invention, and FIG. 2 is a diagram D of the gas turbine blade according to the embodiment of the present invention. It is sectional drawing in -D cross section (refer FIG. 1: the surface of the substantially circumferential direction of a gas turbine rotor). That is, FIG. 1 shows the inner wall surface on the back side of the gas turbine rotor blade.

  A plurality of gas turbine rotor blades 50 shown in these drawings are annularly arranged on the outer periphery of a gas turbine rotor (hereinafter sometimes referred to as a rotor), and a blade root portion 57 located on the rotation center side of the rotor, The blade portion 58 extends from the blade root portion 57 to the blade tip portion 59 toward the outer side in the rotor radial direction. The blade portion 58 is disposed in a flow path (gas path) through which the working medium (combustion gas) 30 flows in the gas turbine, and is exposed to a high temperature and high pressure environment.

  Two systems of cooling flow paths 41, 42 are provided inside the gas turbine rotor blade 50 in the present embodiment, and a plurality of partition walls 60 that are part of the cooling flow paths 41, 42 are formed from the front edge 51. It is provided over the trailing edge 52. In the present embodiment, five partition walls 60a, 60b, 60c, 60d, and 60e are provided from the front edge 51 to the rear edge 52 (when it is not necessary to distinguish the five partition walls 60a to 60e). The subscripts (a to e) are omitted and indicated as “partition wall 60. In the following, the other parts may be indicated in the same manner).

  The cooling flow path 41 is a meandering flow path formed on the front edge 51 side in the gas turbine rotor blade 50, and includes a wall surface on the ventral side 53 of the gas turbine rotor blade 50 and a back side 54 of the gas turbine rotor blade 50. And three partition walls 60a, 60b, 60c that divide the two wall surfaces. The cooling flow path 41 is a flow path 41a located at a substantially central part in the chord direction in the wing portion 58, a flow path 41c located on the leading edge 51 side, and a flow located between the two flow paths 41a and 41c. A path 41b is provided. Each flow path 41a, 41b, 41c extends in the blade length direction (rotor radial direction) of the gas turbine rotor blade 50, and the adjacent flow paths among these flow paths 41a, 41b, 41c are the blade tip 59 or The wing root portion 57 is connected, and the flow direction of the cooling medium is changed at the connection portion.

  The inlet 40a of the cooling channel 41 is provided in the vicinity of the blade center on the blade root 57 side, and the cooling medium 15a is introduced into the blade from the channel inlet 40a. As the cooling medium 15a, compressed air extracted from the gas turbine compressor can be used. In addition to air, a refrigerant such as steam may be used.

  A plurality of film cooling holes 1 for allowing the cooling medium 15a flowing through the cooling flow path 41 to flow out of the gas turbine blade 50 are provided in the blade inner wall surface forming the flow path 41c. The cooling medium 15a flows in the flow path 41c from the blade root portion 57 side toward the blade tip portion 59 side, and is ejected to the outside of the blade as the cooling medium 11 through the film cooling hole 1.

  From the viewpoint of improving the film cooling performance, the flow path 41c located on the leading edge 51 side allows the cooling medium to flow from the blade root portion 57 side toward the blade tip portion 59 side as in the present embodiment. It is preferable to configure.

  The cooling flow path 42 is a meandering flow path formed on the rear edge 52 side in the gas turbine rotor blade 50, and a wall surface on the ventral side 53 of the gas turbine rotor blade 50 and a back side 54 of the gas turbine rotor blade 50. And three partition walls 60c, 60d, and 60e that divide the two wall surfaces. In the chord direction, the cooling flow path 42 includes a flow path 42a located at a substantially central portion of the wing portion 58 in the chord direction, a flow path 42c located on the trailing edge 52 side, and two flow paths 42a and 42c. A flow path 42b is provided between them. Each flow path 42a, 42b, 42c extends in the blade length direction (rotor radial direction) of the gas turbine rotor blade 50. Among these flow paths 42a, 42b, 42c, adjacent flow paths are the blade tip 59 or It connects with the blade root part 57, and the distribution direction of the cooling medium is changed in the said connection part.

  A plurality of ribs 2 and 3 are provided on the wall surfaces of the ventral side 53 and the back side 54 in each cooling channel 41 and 42. The types of the ribs 2 and 3 provided in the flow paths 41 and 42 are such that the local temperature of the blade 50 is equal to or lower than the limit value for high temperature corrosion with respect to the heat load distribution from the high temperature working medium 30 to the blade 50. The entire temperature distribution is appropriately selected so as to be equal to or less than the thermal stress limit value. In addition, the ribs 2 and 3 in the gas turbine rotor blade 50 of the present embodiment are provided not only on the back side 54 but also on the ventral side 53 as shown in the cross-sectional view of FIG. Further, as shown in FIG. 2, the ribs 2 and 3 are not provided on each partition wall 60.

  V-shaped staggered ribs in which two ribs (front edge side rib 3a and rear edge side rib 3b) having different inclination directions are combined in a substantially V shape on the wall surfaces of the ventral side 53 and the back side 54 in the flow paths 41a, 41b, 42a. 3 is provided. Of the two ribs forming the V-shape in the staggered rib 3, the leading edge side rib 3a is a rib disposed on the leading edge 51 side, and the trailing edge side rib 3b is a rib disposed on the trailing edge 52 side. The front edge side rib 3a is inclined so that the end portion on the front edge 51 side is located downstream of the end portion on the rear edge 52 side in the flow direction of the cooling medium. On the other hand, the rear edge side rib 3b is inclined so that the end portion on the rear edge 53 side is located downstream of the end portion on the front edge 51 side in the flow direction of the cooling medium. For example, taking the channel 41a in FIG. 1 as an example, the leading edge side rib 3a is inclined upward and the trailing edge side rib 3b is inclined upward.

  A plurality of pin fins 4 for cooling the pin fins are provided in the flow path 43 communicating with the trailing edge 52 side. The pin fin 4 contributes to the improvement of the internal cooling performance by generating vortices and disturbances in the cooling medium while ensuring the strength of the flow path 43.

  FIG. 3 is a structural diagram of the rib 2 and the film cooling hole 1 provided on the inner wall of the flow path 41 according to the embodiment of the present invention. The left side in FIG. 3 corresponds to the ventral side 53 of the wing 50, and the right side corresponds to the back side 54. Note that the same reference numerals are given to the same parts as those in the previous figure, and description thereof will be omitted as appropriate (the same applies to the subsequent figures).

  As shown in this figure, a ventral rib 2a is provided on the wall surface of the ventral side 53 in the channel 41c located on the front edge 51 side, and a back rib 2b is provided on the wall surface of the dorsal side 54. Yes.

  The ventral rib 2a extends from the front edge 51 side to the rear edge 52 side with respect to the inner wall surface of the flow path 41, and the rear edge side end portion 72a of the rib 2a is cooled more than the front edge side end portion 71a. The medium 10 is inclined in one direction so as to be located downstream of the flow direction of the medium 10. In the example of FIG. 3, the abdominal rib 2a has an end 72a on the trailing edge side located on the blade tip 59 side with respect to the end 71a on the front edge side, and is inclined upward to the left in FIG.

  The back rib 2b extends from the front edge 51 side to the rear edge 52 side with respect to the inner wall surface of the flow path 41, and the rear edge side end 72b of the rib 2b is cooled more than the front edge side end 71b. It is inclined in one direction so as to be located downstream of the medium flow direction. In the example of FIG. 3, the rear rib 2b has a trailing edge 72b positioned closer to the blade tip 59 than the leading edge 71b, and is inclined upward in FIG.

  A predetermined interval is provided along the front edge 51 between the front edge side end portion 71a of the ventral rib 2a and the front edge side end portion 71b of the back side rib 2b. A film cooling hole 1 c is arranged along the front edge 51 between the edge side end portion 71 b.

  The film cooling hole 1 in the present embodiment is mainly provided with three types (1a, 1b, 1c (see FIG. 2, FIG. 3, etc.)) depending on the difference in the position of the opening end, respectively. It is formed in a substantially cylindrical shape. As described above, the film cooling hole (front edge film cooling hole) 1c has an opening end disposed on the front edge 51, and a cross section appears on the blade development view shown in FIG. The film cooling hole 1a on the ventral side 53 has an opening end disposed on the ventral side 53 with the front edge 51 portion as a boundary, and the film cooling hole 1b on the back side 54 has a back side with the front edge 51 portion as a boundary. 54 has an open end. The same type of film cooling holes 1a, 1b, and 1c are arranged at predetermined intervals along the flow direction of the cooling medium.

  In the gas turbine rotor blade 50 configured as described above, the cooling medium 15a introduced into the blade from the inlet 40a of the cooling flow path 41 is appropriately turned in the flow direction at the blade tip 59 or the blade root 57. It advances through the meandering flow paths 41a and 41b, and is introduced into the flow path 41c located closest to the front edge 51 while being gradually heated by the heat of the working medium. The cooling medium 10 introduced into the flow path 41 c flows in the flow path 41 c from the blade root portion 57 toward the blade end portion 59. At this time, a rotational component is given to the cooling medium 10 by the ventral rib 2a and the back rib 2b, and vortices 20a and 20b are generated. That is, on the ventral side 53, when viewed from the upstream side in the flow direction of the cooling medium 10 (when viewed from the blade root portion 57 to the blade tip portion 59 in the flow path 41c), the clockwise vortex 20a is formed. Occur. On the other hand, on the back side 54, a vortex 20b in the opposite direction, that is, a counterclockwise vortex 20b as seen from the same direction is generated.

  Since the vortices 20a and 20b create a disturbance in the flow of the cooling medium in the flow path 41c, the internal cooling performance of the flow path 41c can be improved as compared with the case where there is no rib. Furthermore, the vortices 20a and 20b contribute to the improvement of the film cooling performance (external cooling performance) as described below.

  FIG. 4: has shown sectional drawing in the AA cross section (refer FIG. 3) along the abdominal film cooling hole 1a about the flow path 41c which concerns on embodiment of this invention. As shown in this figure, the cooling medium 10 is introduced with the vortex 20a into the film cooling hole 1a provided in the blade wall 55a on the ventral side 53, and ejected from the blade surface as the cooling medium 11a with the vortex 21a. .

  FIG. 5: has shown sectional drawing in the BB cross section (refer FIG. 3) along the back side film cooling hole 1b about the flow path 41c which concerns on embodiment of this invention. As shown in this figure, the cooling medium 10 is introduced into the film cooling hole 1b provided in the blade wall 55b on the back side 54 with the vortex 20b, and the vortex 21b rotates in the opposite direction to the vortex 21a on the ventral side 53. Is ejected from the blade surface as a cooling medium 11b with

  FIG. 6 shows a cross-sectional view of the gas turbine rotor blade 50 in which the flow path 41c portion of FIG. 2 is enlarged. As shown in this figure, in the cooling medium flowing through the flow path 41c, the vortex 20a is generated on the ventral side 53 by the ventral rib 2a, and the vortex 20b is generated on the dorsal side 54 by the dorsal rib 2b. . Note that FIG. 6 is a cross-sectional view of the direction of the blade root portion 57 from the blade tip portion 59 side, and therefore the rotation direction of the vortices 20a and 20b is when the direction of the blade tip portion 59 is viewed from the blade root portion 57 side ( The flow direction of the cooling medium 10 in the flow path 41c is opposite to that in FIGS. 3, 4 and 5 when viewed from the upstream side to the downstream side.

  As shown in the present embodiment, the film cooling hole 1 has an external opening end (opening end on the blade surface side) cooled more than an internal opening end (opening end on the blade inside (flow path 41c) side). It is preferable to incline so that it may be located in the downstream of the distribution direction of a medium (refer FIG.1,4,5). When the film cooling hole 1 is inclined as described above, the flow of the vortex 20 generated by the rib 2 can be suppressed from being reduced when passing through the film cooling hole 1. Thereby, since the reduction | decrease of the vortex 21 of the cooling medium ejected from the film cooling hole 1 is suppressed, the fall of film cooling performance can be suppressed.

  Further, as in the present embodiment, the film cooling hole 1 is formed so that the opening end of the film cooling hole 1 on the flow channel 41c side is positioned between two adjacent ribs 2 in the flow direction of the cooling medium 10. It is preferable to arrange. This is based on the viewpoint of maintaining the generation effect of the vortex 20 by the rib 2.

  By the way, in general film cooling, the cooling medium ejected from the film cooling hole is made to flow on the blade surface along the flow of the working medium (combustion gas), and the working medium is covered by covering the blade surface with the cooling medium. The amount of heat input to the blade is reduced. However, just ejecting the cooling medium from the film cooling holes in this way generates a place where the cooling medium tends to peel off from the blade surface compared to other parts (for example, the blade belly side). Cooling performance could not be obtained sufficiently.

  In contrast, in the present embodiment, the vortex 20a is generated on the ventral side 53 of the flow path 41c, and the vortex 21a is accompanied with the cooling medium 11a. When the cooling medium 11a is accompanied with the vortex 21a as described above, the cooling medium 11a can flow from the film cooling hole 1a toward the trailing edge 52 and flow toward the blade surface on the ventral side 53 by the action of the vortex 21a. it can. That is, on the ventral side 53 of the gas turbine rotor blade 50, a flow is formed so as to be wound around the blade surface on the ventral side 53 while facing the trailing edge 52 side. This makes it difficult for the cooling medium 11a to be peeled off from the blade surface on the ventral side 53, so that the film cooling performance can be improved as compared with the conventional one.

  Furthermore, in the present embodiment, the vortex 20b is generated on the back side 54 of the flow path 41c as described above, and the vortex 21b is associated with the cooling medium 11b. When the cooling medium 11b is accompanied with the vortex 21b as described above, the cooling medium 11b can flow from the film cooling hole 1b toward the trailing edge 52 and flow toward the blade surface on the back side 54 by the action of the vortex 21b. it can. That is, a flow is formed on the back side 54 of the gas turbine rotor blade 50 so as to be wound on the blade surface of the back side 54 while facing the rear edge 52 side. This makes it difficult for the cooling medium 11b to peel off from the blade surface on the back side 54, so that the film cooling performance can be improved as compared with the conventional one.

  Therefore, according to the present embodiment, it is possible to make it difficult for the cooling media 11a and 11b to be peeled off from the blade surfaces on the ventral side 53 and the back side 54, so that the film cooling performance can be improved. In addition, according to the blade internal structure according to the present embodiment, the film cooling performance is improved, so that the cooling performance can be improved without providing impingement cooling cavities in the flow path located at the blade leading edge. . Thereby, especially the leading edge cooling of the gas turbine rotor blade (for example, a small gas turbine rotor blade) in which it is difficult to increase the internal cooling flow path can be easily enhanced.

  On the other hand, if the blade internal structure according to the present embodiment is applied to a large blade, the impingement cooling holes need not be manufactured, and the internal cooling structure can be simplified. Normally, since a gas turbine rotor blade is manufactured by precision casting, it is necessary to use a core when a hole is provided inside the blade including an impingement cooling hole. However, if the impingement cooling hole is not required to be manufactured as described above, it is possible to avoid a decrease in yield due to breakage of the core, which is likely to occur during the manufacture of the impingement cooling hole, so that the manufacturing cost of the blade can be reduced. .

  As described above, according to the present embodiment, the cooling performance of the blade leading edge portion 51 can be enhanced, and the high-temperature corrosion of the blade material and the decrease in the structural strength can be suppressed to improve the operating rate of the gas turbine. . In addition, the operating cost can be reduced by extending the blade life. Furthermore, when the compressed air extracted from the compressor is used as a cooling medium for the gas turbine rotor blade, the amount of compressed air extracted can be reduced along with the improvement of the film cooling performance, so that the thermal efficiency of the gas turbine is improved. Can be achieved.

  Moreover, the opening part by the side of the flow path 41c in the film cooling hole 1 which concerns on this Embodiment is located between the two ribs 2 adjacent in the distribution direction of the cooling medium 10, as shown in FIG. Therefore, the peeling area can be reduced in consideration of the relative position of the rib 2 (turbulator) and the film cooling hole 1 as in Patent Document 2. However, since the film cooling performance from the outside of the blade can be improved by the effect of the vortices 20a and 20b even if the arrangement of the rib 2 and the film cooling hole 1 is not taken into account, the gas turbine rotor blade having excellent cooling performance can be easily obtained. Can be manufactured.

  Further, when the pair of ventral ribs 2a and back ribs 2b are combined into a substantially V shape at the front edge 51 as in the present embodiment, the same effect as the V-shaped rib is exhibited at the front edge 51. can do. In the V-shaped rib, the heat transfer performance of the narrow portion formed between the pair of ribs forming the V-shape can be improved, but in this embodiment, the narrow portion is positioned at the front edge 51 portion. Since the ribs 2a and 2b are disposed in the inner cooling performance, it is possible to improve the internal cooling performance of the front edge 51 portion having a relatively high thermal load as compared with other portions.

  In the above description, the film cooling hole 1 is provided only in the flow path 41c located on the front edge 51 side. However, the inner wall surface of the ventral side 53 or the back side 54 in the other flow paths 41a, 41b, 42 (that is, Even if the rib 2 and the film cooling hole 1 inclined in the same direction as described above are provided on the inner wall surface provided with the rib 3, the same vortex as described above can be generated. Can be improved. In the above description, the gas turbine rotor blade is taken as an example, but the present invention is also applicable to a gas turbine stationary blade.

DESCRIPTION OF SYMBOLS 1 Film cooling hole 2a Abdominal side rib 2b Back side rib 4 Pin fin 10 Cooling medium 11 Cooling medium 12 Cooling medium 15 Cooling medium 20 Cooling medium vortex 21 Cooling medium vortex 30 Working medium 41 Leading edge side cooling flow path 42 Trailing edge side cooling flow path 43 Trailing edge cooling flow path 50 Rotor blade 51 Leading edge 52 Trailing edge 53 Abdominal side 54 Back side 55 Blade wall 71 Front edge side end 72 of rib 2 End of trailing edge side of rib 2

Claims (6)

A cooling flow path provided inside the gas turbine blade for flowing a cooling medium;
A rib extending from the front edge side to the rear edge side with respect to the wall surface on the ventral side of the gas turbine blade in the cooling flow path, wherein the end portion on the rear edge side is more of the cooling medium than the end portion on the front edge side. A ventral rib inclined so as to be located downstream in the flow direction;
A rib that extends from the front edge side to the rear edge side with respect to the wall surface on the back side of the gas turbine blade in the cooling flow path, and the end portion on the rear edge side is more of the cooling medium than the end portion on the front edge side. A dorsal rib inclined so as to be positioned downstream in the flow direction;
A gas turbine blade, comprising: a film cooling hole for allowing a cooling medium flowing through the cooling flow path to flow out of the gas turbine blade. A cooling flow path provided inside the leading edge of the gas turbine blade, for flowing a cooling medium from the blade root toward the blade tip;
A rib extending from the front edge side to the rear edge side with respect to the wall surface on the ventral side of the gas turbine blade in the cooling flow path, wherein the end portion on the rear edge side is more than the end portion on the front edge side. A ventral rib that is inclined so as to be located on the side,
A rib extending from the front edge side to the rear edge side with respect to the wall surface on the back side of the gas turbine blade in the cooling flow path, wherein the end portion on the rear edge side is more than the end portion on the front edge side. A dorsal rib inclined so as to be located on the side of the head,
A gas turbine blade, comprising: a film cooling hole for allowing a cooling medium flowing through the cooling flow path to flow out of the gas turbine blade. In the gas turbine blade according to claim 1 or 2,
The gas cooling blade according to claim 1, wherein the film cooling hole is inclined so that the opening end on the blade surface side is located downstream of the opening end on the blade inner side in the flow direction of the cooling medium. In the gas turbine blade according to claim 1 or 2,
The gas turbine blade according to claim 1, wherein the cooling channel side opening of the film cooling hole is located between two adjacent ribs in the flow direction of the cooling medium. The gas turbine blade according to claim 2,
A gas turbine blade according to claim 1, wherein a gap is provided between the ventral rib and the dorsal rib. A cooling flow path provided inside the gas turbine blade for flowing a cooling medium;
A rib extending from the front edge side to the rear edge side with respect to the inner wall surface of the gas turbine blade in the cooling flow path, wherein the end portion on the rear edge side is in the flow direction of the cooling medium than the end portion on the front edge side. A rib inclined so as to be located on the downstream side of
A gas turbine blade, comprising: a film cooling hole provided in the inner wall surface for allowing a cooling medium flowing through the cooling flow path to flow out of the gas turbine blade.

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