JP2010144656A - Gas turbine blade and gas turbine - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To reduce stress concentration resulting from thermal stress, and to improve a fatigue life. <P>SOLUTION: A gas turbine blade 8 is provided with: a blade body 53 having an outer surface along which combustion gas flowing from an upstream side to a downstream side in a gas turbine passes; and a shroud 60, 61 arranged on at least one of opposite sides in a height direction H of the blade body, supporting the blade body and forming a part of a partition wall of a combustion gas flow path F in which the combustion gas flows. The blade body and the shroud are respectively provided with blade body side connection parts 58, 59 and shroud side connection parts 66, 67, which extend in the height direction outside the combustion gas flow path, and the blade body side connection part and the shroud side connection part are connected in parts of their projection ends 70, 71 projecting outside the combustion gas flow path and are not connected at base ends 72, 73 situated on a combustion gas flow path side. <P>COPYRIGHT: (C)2010,JPO&INPIT

Description

  The present invention relates to a gas turbine blade and a gas turbine including the gas turbine blade.

Conventionally, for example, the following gas turbine blade as shown in Patent Document 1 below is known. That is, the gas turbine blade includes a hollow blade body (air foil) having an outer surface, and the outer surface extends laterally between a leading edge and a trailing edge, and It extends in the height direction (length direction) between the first end and the second end located on opposite sides. The outer surface of the wing body has at least one groove for receiving the brazing material adjacent to the first end, and the groove extends laterally between the leading edge and the trailing edge. It extends at least partially. The gas turbine blade includes a shroud (first band), and the shroud has a shroud side connecting portion (inner surface) that defines a support opening, and the support opening includes the first shroud. The wing body is received in a shape complementary to the wing body at the end of the wing. In addition, the shroud side connecting portion has at least one groove, and the groove defines an enlarged clevis for receiving the brazing material between the groove of the blade body and the groove. Extending along substantially the same straight line and facing the groove of the blade body.
Japanese Patent No. 3130739

  However, in the conventional gas turbine blade, the shroud side connecting portion extending in the height direction of the blade body is brazed to the outer surface of the blade body over the entire surface. Therefore, when high-temperature combustion gas flows in the combustion gas flow path in which the shroud forms a part of the partition wall, thermal stress due to thermal deformation acts on the blade body and the shroud, and the combustion gas in the shroud side connection portion. Stress concentration occurs at the base end located on the flow path side. As a result, there is a problem that damage due to fatigue is promoted at the base end portion, and the fatigue life of the gas turbine blade is shortened.

  The present invention has been made in view of the above-described circumstances, and an object of the present invention is to relieve stress concentration caused by thermal stress and improve fatigue life, and the gas turbine. It is to provide a gas turbine with blades.

In order to solve the above problems, the present invention proposes the following means.
The gas turbine blade according to the present invention is disposed at least one of a blade main body through which the combustion gas flowing from the upstream side toward the downstream side in the gas turbine passes the outer surface, and both ends in the height direction of the blade main body. And a shroud that constitutes a part of a partition wall of the combustion gas flow path through which the combustion gas flows, and the blade body and the shroud are arranged in the height direction outside the combustion gas flow path. The blade main body side connecting portion and the shroud side connecting portion that are extended to each other are provided, and the blade main body side connecting portion and the shroud side connecting portion are one of projecting end portions that protrude outside the combustion gas flow path. It is connected by the part, It is not connected in the base end part located in the said combustion gas flow path side, It is characterized by the above-mentioned.

  According to the gas turbine blade, the blade main body side connecting portion and the shroud side connecting portion are connected to each other at a part of the projecting end portion and not connected to the base end portion. However, the thermal deformation of each of the blade body and the shroud can be absorbed by the base end side portion located on the combustion gas flow path side from the projecting end portion in the blade body side connection portion and the shroud side connection portion. Therefore, stress concentration caused by thermal stress can be relaxed, and the fatigue life of the gas turbine blade can be improved.

  Further, the shroud may be provided with a first cooling hole for discharging a cooling gas into the combustion gas flow path and cooling a portion of the blade body connected to the blade body side connecting portion.

According to this configuration, since the first cooling hole is provided in the shroud, it is possible to cool the portion connected to the blade body side coupling portion of the blade body. Therefore, it is possible to further suppress the thermal stress acting on the blade body, and to further improve the fatigue life of the gas turbine blade.
Here, when the first cooling hole is provided in the vicinity of the base end portion of the shroud side connecting portion, the stress increases in the vicinity of the first cooling hole, but the stress is reduced in combination with the effect of relaxing the thermal stress. The lifetime can be improved.

  Moreover, the thick part formed thicker than the other part may be provided in the part connected to the said wing body side connection part of the said wing | blade main body.

According to this configuration, since the thick portion is provided in the portion connected to the blade body side coupling portion of the wing body, the rigidity of the portion connected to the blade body side coupling portion of the wing body can be increased. It becomes possible. This suppresses the occurrence of gas bending even when a force that causes gas bending that plastically deforms from the base end of the blade body side coupling part due to the pressure of the combustion gas acts on the blade body. can do. Accordingly, it is possible to suppress the occurrence of gas bending while improving the fatigue life of the gas turbine blade.
Here, by providing the thick part, thermal stress is likely to occur in the vicinity of the thick part, but the thermal stress in the thick part is relieved by cooling the first cooling hole, so that the fatigue life can be improved. it can.

  In addition, a cooling gas introduction portion into which the cooling gas discharged from the first cooling hole is introduced may be provided in the vicinity of a portion connected to the blade body side coupling portion of the blade body.

  According to this configuration, since the cooling gas introduction part is provided, when the cooling gas released from the first cooling hole cools the part connected to the blade body side coupling part of the blade body, the combustion gas It can suppress that the above-mentioned cooling is inhibited by the combustion gas which distribute | circulates a flow path. Therefore, the thermal stress acting on the blade body can be further suppressed, and the fatigue life of the first stage stationary blade can be further improved.

  Further, at least one of the mutually opposing surfaces of the blade main body side connecting portion and the shroud side connecting portion has a relief portion that suppresses contact between the base ends of the blade main body side connecting portion and the shroud side connecting portion. May be provided.

  According to this configuration, since the relief portion is provided on at least one of the mutually opposing surfaces of the blade main body side connecting portion and the shroud side connecting portion, the base ends of the blade main body side connecting portion and the shroud side connecting portion are provided. It is possible to suppress fretting of each other (repetitive relative slip with frictional force and small amplitude). Therefore, fretting wear and fretting fatigue can be suppressed, and the fatigue life of the gas turbine blade can be further improved.

  Further, surface treatment layers having different thicknesses may be provided on the outer surface of the blade body and the surface of the shroud on the combustion gas flow path side.

  In this configuration, surface treatment layers having different thicknesses are provided on the outer surface of the blade body and the surface of the shroud on the combustion gas flow path side. Here, after the surface treatment layer is formed in advance on each of the blade main body and the shroud, for example, the blade main body and the shroud are integrally formed by connecting via the blade main body side connecting portion and the shroud side connecting portion. As compared with the above, surface treatment layers having different thicknesses can be formed easily and accurately. Therefore, according to this gas turbine blade, it is possible to easily form a surface treatment layer having a thickness suitable for the temperature conditions of the blade body and the shroud while improving the fatigue life.

  Further, between the blade main body side connecting portion and the shroud side connecting portion, a cooling gas is discharged into the combustion gas flow path to cool a portion connected to the blade main body side connecting portion of the blade main body. A second cooling hole may be provided.

  According to this configuration, since the second cooling hole is provided between the blade body side coupling portion and the shroud side coupling portion, the portion connected to the blade body side coupling portion of the blade body can be cooled. . Therefore, it is possible to further suppress the thermal stress acting on the blade body, and to further improve the fatigue life of the gas turbine blade.

  Moreover, the positioning part which positions the said wing | blade main body in the said height direction with respect to the said shroud may be provided in the said shroud side connection part.

  According to this configuration, since the positioning portion is provided in the shroud, the assembly of the blade body and the shroud can be facilitated when the gas turbine blade is manufactured.

  In addition, a gas turbine according to the present invention includes a compressor that generates compressed air, a combustor that generates fuel by supplying fuel to the compressed air supplied from the compressor, and the gas turbine according to the present invention. A turbine having blades and generating rotational power by the combustion gas supplied from the combustor.

According to the gas turbine, since the turbine has the gas turbine blades according to the invention, the fatigue life of the gas turbine as a whole can be improved.
In the turbine vanes, in particular, in the first-stage vanes and the two-stage vanes positioned on the upstream side, the high-temperature combustion gas passes through the combustion gas flow path. Therefore, the gas turbine blades according to the present invention are applied to these vanes. Employment is effective against fatigue caused by thermal stress.

According to the gas turbine blade of the present invention, stress concentration caused by thermal stress can be relaxed, and the fatigue life can be improved.
Moreover, according to the gas turbine which concerns on this invention, since this gas turbine blade is provided, the improvement of a fatigue life as a whole can be aimed at.

(First embodiment)
Hereinafter, a first embodiment according to the present invention will be described with reference to the drawings. FIG. 1 is a schematic half sectional view showing a gas turbine according to a first embodiment of the present invention.
As shown in FIG. 1, the gas turbine 1 includes a compressor 2 that generates compressed air, and a plurality of fuel gas G1 (see FIG. 2) that supplies fuel to the compressed air supplied from the compressor 2. A combustor 3 and a turbine 4 having at least one stage of turbine stationary blades 5 and turbine rotor blades 6 and generating rotational power by the combustion gas G1 supplied from the combustor 3 are provided.

  In addition, a rotor 7 extending in the axial direction D is integrally attached to the gas turbine 1 from the compressor 2 to the turbine 4, and the rotor 7 has a bearing portion 23 provided at one end in the compressor 2. Is supported rotatably in the circumferential direction R of the turbine 4 around the axis O, and the other end is supported rotatably in the circumferential direction R of the turbine 4 by a bearing portion 41 provided in the turbine 4. Hereinafter, the compressor 2 side is defined as the front side and the turbine 4 side is defined as the rear side along the axial direction D of the rotor 7.

The compressor 2 includes a compressor casing 20 disposed with the air intake port 20a for taking in air facing forward, a plurality of compressor vanes 21 and a plurality of compressor movements disposed in the compressor casing 20. And wings 22.
The compressor vane 21 includes a compressor vane body 51 that is fixed to the inner peripheral surface of the compressor casing 20 and extends toward the rotor 7. The compressor vane main bodies 51 are arranged at equal intervals in the circumferential direction R of the turbine 4. The compressor blade 22 includes a compressor blade main body 52 that is fixed to the outer peripheral surface of the rotor 7 and extends toward the inner peripheral surface of the compressor casing 20. The compressor blade main bodies 52 are arranged at equal intervals in the circumferential direction R of the turbine 4. The compressor stationary blades 21 and the compressor rotor blades 22 are arranged in multiple stages so as to alternate along the axial direction D.

The combustor 3 includes an inner cylinder 30 having a burner (not shown) therein, an outer cylinder 31 that guides compressed air supplied from the compressor 2 to the inner cylinder 30, and a fuel injector (not shown) that supplies fuel to the inner cylinder 30. And a tail cylinder 32 that guides the combustion gas G <b> 1 from the inner cylinder 30 to the turbine 4. According to the combustor 3 configured as described above, in the inner cylinder 30, the compressed air guided from the outer cylinder 31 and the fuel supplied from the combustion injector are mixed, and the mixed fluid is burned by the burner. Thus, the combustion gas G1 can be generated, and the combustion gas G1 can be guided to the turbine 4 through the tail cylinder 32.
The plurality of combustors 3 are arranged in the circumferential direction R of the turbine 4 and are disposed inside a combustor casing 33 whose front end is connected to the rear end of the compressor casing 20.

The turbine 4 includes a turbine casing 40 having a front end connected to a rear end of the combustor casing 33, and the turbine stationary blade 5 and the turbine rotor blade 6 disposed in the turbine casing 40. Yes.
The turbine vane 5 includes a turbine vane main body 53 that is fixed to the inner peripheral surface of the turbine casing 40 and extends toward the rotor 7. The turbine stationary blade bodies 53 are arranged at equal intervals in the circumferential direction R of the turbine 4. The turbine blade 6 includes a turbine blade main body 54 that is fixed to the outer peripheral surface of the rotor 7 and extends toward the inner peripheral surface of the turbine casing 40. The turbine rotor blade main bodies 54 are arranged at equal intervals in the circumferential direction R of the turbine 4. The turbine stationary blades 5 and the turbine rotor blades 6 are arranged in multiple stages so as to alternate along the axial direction D.

Further, the turbine 4 is provided with a bypass passage (not shown) through which air inside the compressor 2 is supplied from the compressor 2 by bypassing the combustor 3. The air supplied to the turbine 4 through this bypass flow path is circulated inside the turbine stationary blade body 53 and the turbine rotor blade body 54 as a cooling gas G2 (see FIG. 2).
An exhaust chamber 42 that opens toward the rear side is connected to the rear end portion of the turbine casing 40. The exhaust chamber 42 is provided with an exhaust diffuser 42a that converts the dynamic pressure of the combustion gas G1 that has passed through the turbine stationary blade 5 and the turbine rotor blade 6 into a static pressure.

In the gas turbine 1 configured as described above, first, air taken in from the air intake port 20a of the compressor 2 passes through the compressor stationary blades 21 and the compressor rotor blades 22 arranged in multiple stages and is compressed. Compressed air is generated. Next, in the combustor 3, the combustion gas G <b> 1 is generated from the compressed air as described above, and this combustion gas G <b> 1 is guided to the turbine 4. The combustion gas G1 passes through the turbine stationary blade 5 and the turbine rotor blade 6, whereby the rotor 7 is rotationally driven, and the gas turbine 1 can output rotational power.
The exhaust gas after rotationally driving the rotor 7 is converted into a static pressure by the exhaust diffuser 42a in the exhaust chamber 42 and then released to the atmosphere.

Next, in the gas turbine 1 configured as described above, the configuration of the first stage stationary blade 8 positioned on the foremost side of the turbine stationary blade 5 will be described in detail. FIG. 2 is a perspective view of a part of the first stage stationary blade of the gas turbine shown in FIG. 3 is a cross-sectional view taken along line AA shown in FIG.
As shown in FIG. 2, the first stage stationary blade 8 is configured such that the combustion gas G1 flowing from the front side (upstream side) to the rear side (downstream side) of the gas turbine 1 passes through the outer peripheral surface (outer surface) 53a. The main body (blade main body) 53 is disposed at both ends in the height direction H of the turbine stationary blade main body 53 that coincides with the radial direction of the turbine 4, supports the turbine stationary blade main body 53, and the combustion gas G1 flows. And shrouds 60 and 61 constituting a part of the partition wall of the combustion gas flow path F.

As shown in FIG. 3, the turbine stationary blade body 53 is formed in a cylindrical shape extending in the height direction H, and an outer peripheral surface 53 a and an inner peripheral surface 53 b in a cross-sectional view along the height direction H. Is formed in parallel with the height direction H.
Further, the turbine stationary blade body 53 is provided with a blade body side coupling portion extending in the height direction H. In the present embodiment, as the blade main body side connecting portion, from the end portion on one side H1 in the height direction H (outside in the radial direction) of the turbine stationary blade main body 53 toward the one side H1 in the height direction H. From the extended first blade body side connecting portion 58 and the end of the other side H2 in the height direction H (inner side in the radial direction) of the turbine stationary blade body 53 toward the other side H2 in the height direction H. And a second blade main body side connecting portion 59 extended.
The first blade main body side connecting portion 58 and the second blade main body side connecting portion 59 are both formed in a cylindrical shape connected to the turbine stationary blade main body 53, and any of the outer peripheral surfaces 58a and 59a and the inner peripheral surfaces 58b and 59b are It is formed so as to be continuous with the outer peripheral surface 53a and the inner peripheral surface 53b at the end of the corresponding turbine stationary blade body 53.

In the present embodiment, as shown in FIG. 2, the shroud is disposed on the first shroud 60 disposed on the first blade main body side connecting portion 58 side and on the second blade main body side connecting portion 59 side. A second shroud 61.
As shown in FIG. 3, the first shroud 60 is formed in a plate shape and is disposed substantially parallel to the axis O. The first shroud 60 is formed with a first insertion hole 63 through which the first blade main body side connecting portion 58 is inserted. The first insertion hole 63 is formed in the same shape and the same size as the outer shape of the first blade main body side connecting portion 58, and the turbine stationary blade at the first blade main body side connecting portion 58 is in the first insertion hole 63. A portion (base end portion 72 described later) connected to the main body 53 is inserted without a gap.

  The second shroud 61 is formed in a plate shape and is disposed substantially parallel to the axis O. The second shroud 61 is formed with a second insertion hole 65 through which the second blade main body side connecting portion 59 is inserted. The second insertion hole 65 is formed in the same shape and the same size as the outer shape of the second blade main body side connection portion 59, and the turbine stationary blade in the second blade main body side connection portion 59 is in the second insertion hole 65. A portion (base end portion 72 described later) connected to the main body 53 is inserted without a gap.

  In the gas turbine 1 shown in FIG. 1, by arranging a plurality of turbine stationary blade bodies 53 in the circumferential direction R, each of the first shroud 60 and the second shroud 61 is connected in an annular shape, and the first shroud connected in an annular shape. 60 is fixed to the inner peripheral surface of the turbine casing 40, and the second shroud 61 connected in an annular shape is connected to the rear end of the tail cylinder 32 of the combustor 3, whereby the first shroud 60 and the second shroud 61 are connected. An annular combustion gas flow path F is formed between the two. Further, as shown in FIG. 3, the exterior A partitioned from the combustion gas flow path F by the shrouds 60 and 61 and the interior B of the turbine stationary blade body 53 communicate with each other.

Further, the first shroud 60 and the second shroud 61 are respectively provided with a first shroud side connecting portion 66 and a second shroud side connecting portion 67 extending in the height direction H.
The first shroud side connecting portion 66 is formed in a cylindrical shape extending from the peripheral wall portion 63a defining the outline of the first insertion hole 63 toward the outside with respect to the combustion gas flow path F that is one side H1 in the height direction H. Is formed. The protruding end of the first shroud side connecting portion 66 extending to the one side H1 in the height direction H has a protruding end located outside the combustion gas flow path F in the first blade main body side connecting portion 58 and the height. The position in the direction H matches. Further, the inner peripheral surface 66a of the first shroud side connecting portion 66 and the outer peripheral surface 58a of the first blade main body side connecting portion 58 are in contact with each other without any gap.

  The second shroud side connecting portion 67 extends from the peripheral wall portion 65a defining the outline of the second insertion hole 65 toward the outside with respect to the combustion gas flow path F that is the other side H2 in the height direction H. It is formed in a shape. The protruding end of the second shroud side connecting portion 67 that extends to the other side H2 in the height direction H has a protruding end located outside the combustion gas flow path F in the second blade body side connecting portion 59, and the height. The position in the direction H matches. Further, the inner peripheral surface 67a of the second shroud side connecting portion 67 and the outer peripheral surface 59a of the second blade main body side connecting portion 59 are in contact with each other without any gap.

  The turbine vane main body 53 is supported by the first shroud 60 via the connecting portions 58 and 66 by connecting the first blade main body side connecting portion 58 and the first shroud side connecting portion 66 to each other. Further, the second blade main body side connecting portion 59 and the second shroud side connecting portion 67 are connected to each other, so that they are supported by the second shroud 61 via these connecting portions 59 and 67.

  Here, the configuration related to the connection in the first blade main body side connecting portion 58 and the first shroud side connecting portion 66 and the configuration related to the connection in the second blade main body side connecting portion 59 and the second shroud side connecting portion 67 are: It has the same configuration. Hereinafter, only the configuration related to the connection in the first blade main body side connection portion 58 and the first shroud side connection portion 66 will be described, and the configuration related to the connection in the second blade main body side connection portion 59 and the second shroud side connection portion 67 will be described. About the component which respond | corresponds with the structure which concerns on the connection of the 1st wing body side connection part 58 and the 1st shroud side connection part 66, the same code | symbol is attached | subjected and the description is abbreviate | omitted.

The first blade main body side connecting portion 58 and the first shroud side connecting portion 66 are connected at the protruding end sides that protrude to the outside of the combustion gas flow path F, and at the base end side positioned on the combustion gas flow path F side. Not connected.
More specifically, the protruding portion 70 of the first blade body side connecting portion 58 is cut out so as to be cut away from the inner peripheral surface 58b toward the outer peripheral surface 58a toward the proximal end side. 70a is formed over the entire circumference. Further, the projecting end portion 71 of the first shroud side connecting portion 66 has a notch portion 71a formed by cutting away from the outer peripheral surface 66b toward the inner peripheral surface 66a toward the proximal end side. It is formed over the entire circumference. And the V-shaped groove | channel is formed of the notch part 70a, 71a in the state in which the 1st blade main body side connection part 58 and the 1st shroud side connection part 66 were assembled | attached.

As shown in FIGS. 2 and 3, the first blade main body side connecting portion 58 and the first shroud side connecting portion 66 are formed with a welded portion 75 at a groove formed by both cutout portions 70 a and 71 a. Are connected to each other.
In the example shown in FIG. 3, the length L of the first shroud side connecting portion 66 protruding from the first shroud 60 along the height direction H is equal to the first blade body side connecting portion 58 and the first shroud side. It is larger than twice the wall thickness δ with the connecting portion 66.

  As shown in FIG. 2, the first stage stationary blade 8 configured as described above is supplied with, for example, a cooling gas G2 from a bypass channel, and this cooling gas G2 is supplied to the combustion gas channel by the shrouds 60 and 61. The outside A partitioned by F and the inside B of the turbine stationary blade body 53 are filled. The pressure P1 of the cooling gas G2 in the external A partitioned from the combustion gas flow path F by the shrouds 60 and 61 is higher than the pressure P2 of the combustion gas G1 in the combustion gas flow path F.

  According to the first-stage stationary blade 8 described above, the blade main body side connecting portions 58 and 59 and the shroud side connecting portions 66 and 67 are connected to each other at the projecting end portions 70 and 71, and are connected at the base end portions 72 and 73. Therefore, even if the high-temperature combustion gas G1 circulates in the combustion gas flow path F, the thermal deformation of the turbine stationary blade body 53 and the shrouds 60 and 61 is caused by the blade body side connecting portions 58 and 59 and the In the shroud side connection parts 66 and 67, it can absorb in the part located in the combustion gas flow path F side from the protrusion parts 70 and 71. FIG. Therefore, stress concentration caused by thermal stress can be relaxed, and the fatigue life of the one-stage stationary blade 8 can be improved.

Further, since the pressure P1 of the cooling gas G2 outside the combustion gas passage F is higher than the pressure P2 of the combustion gas G1 inside the combustion gas passage F, for example, a crack or the like has occurred in the weld 75. Even in this case, the combustion gas G1 does not leak to the outside A, but the cooling gas G2 can be discharged from the outside A into the combustion gas flow path F.
Further, according to the gas turbine 1 described above, since the first stage stationary blade 8 is provided, the fatigue life of the gas turbine 1 as a whole can be improved.

(Second Embodiment)
Next, a second embodiment according to the present invention will be described with reference to the drawings. FIG. 4 is a diagram showing a first stage stationary blade according to a second embodiment of the present invention, and is an enlarged cross-sectional view of a portion corresponding to a portion X shown in FIG.
In the second embodiment, the same components as those in the first embodiment are denoted by the same reference numerals, description thereof is omitted, and only different points will be described.

  Further, as shown in the first embodiment, the first stage stationary blade 90 includes the same components on the first shroud 60 side and the second shroud 61 side. Therefore, in the present embodiment and thereafter, the first shroud 60 is provided. In addition, the configuration of the portion (X portion shown in FIG. 3) located on the first shroud 60 side in the turbine stationary blade body 53 will be described as a representative, and the second shroud 61 in the second shroud 61 and the turbine stationary blade body 53 will be described. The description of the structure of the part located in the side is abbreviate | omitted.

As shown in FIG. 4, in the first stage stationary blade 90 of the present embodiment, the first shroud 60 releases the cooling gas G <b> 2 into the combustion gas flow path F to the first blade body side connection portion of the turbine stationary blade body 53. A first cooling hole 91 for cooling a portion 92 connected to 58 is provided.
In the illustrated example, the first cooling hole 91 is formed to communicate with the first shroud 60 between the combustion gas flow path F and the outside A partitioned by the first shroud 60. The first cooling hole 91 is formed in a portion connected to the first shroud side connecting portion 66 of the first shroud 60.
Here, as described above, the pressure P1 of the cooling gas G2 in the external A partitioned from the combustion gas flow path F by the first shroud 60 is higher than the pressure P2 of the combustion gas G1 in the combustion gas flow path F. Therefore, by forming the first cooling hole 91 as described above, the cooling gas G2 outside the first shroud 60 is released into the combustion gas flow path F.

According to the first stage stationary blade 90 shown above, the same operational effects as those of the first embodiment can be obtained.
In addition, since the first cooling hole 91 is provided in the first shroud 60, the portion 92 connected to the first blade body side coupling portion 58 of the turbine stationary blade body 53 can be cooled. Therefore, the thermal stress acting on the turbine stationary blade body 53 can be further suppressed, and the fatigue life of the first stage stationary blade 90 can be further improved.
Here, when the 1st cooling hole 91 is provided in the part connected to the 1st shroud side connection part 66 of the 1st shroud 60 like this embodiment, although stress becomes high in the 1st cooling hole 91 vicinity. In combination with the effect of thermal stress relaxation, the stress can be reduced to improve the fatigue life.

(Third embodiment)
Next, a third embodiment according to the present invention will be described with reference to the drawings. FIG. 5 is a view showing a first stage stationary blade according to a third embodiment of the present invention, and is an enlarged cross-sectional view of a portion corresponding to a portion X shown in FIG.
Note that in the third embodiment, the same components as those in the second embodiment are denoted by the same reference numerals, description thereof is omitted, and only different points will be described.

  As shown in FIG. 5, in the first stage stationary blade 100 of the present embodiment, the portion 92 connected to the first blade body side coupling portion 58 of the turbine stationary blade body 53 is thicker than other portions. A meat part 101 is provided. In the present embodiment, the thick portion 101 is provided from the portion 92 connected to the first blade body side coupling portion 58 of the turbine stationary blade body 53 to the base end portion 72 of the first blade body side coupling portion 58. ing. In the example shown in the drawing, the thick wall portions 101 are respectively connected to the portion 92 connected to the first blade body side coupling portion 58 of the turbine stationary blade body 53 and the base end portion 72 of the first blade body side coupling portion 58. The outer peripheral surfaces 53a and 58a are formed to be flush with each other while the inner peripheral surfaces 53b and 58b are protruded.

According to the 1st stage stationary blade 100 shown above, there can exist an effect similar to 2nd Embodiment.
Further, since the thick portion 101 is provided in the portion 92 connected to the first blade main body side connecting portion 58 of the turbine stationary blade main body 53, it is connected to the first blade main body side connecting portion 58 of the turbine stationary blade main body 53. It is possible to increase the rigidity of the portion 92 to be formed. As a result, even if a force that causes a gas bending that plastically deforms from the base end portion 72 of the first blade body side coupling portion 58 is applied due to the pressure of the combustion gas G1, the occurrence of the gas bending occurs. Can be suppressed. Accordingly, it is possible to suppress the occurrence of gas bending while improving the fatigue life of the first stage stationary blade 100.
Here, the provision of the thick portion 101 facilitates the generation of thermal stress in the vicinity of the thick portion 101, but the thermal stress in the thick portion 101 is relaxed by cooling the first cooling hole 91, thereby improving the fatigue life. Can be planned.

(Fourth embodiment)
Next, a fourth embodiment according to the present invention will be described with reference to the drawings. 6 is a view showing a first stage stationary blade according to a fourth embodiment of the present invention, and is an enlarged cross-sectional view of a portion corresponding to a portion X shown in FIG.
Note that in the fourth embodiment, the same components as those in the first embodiment are denoted by the same reference numerals, description thereof is omitted, and only different points will be described.

As shown in FIG. 6, in the first stage stationary blade 120 of the present embodiment, a stress concentration relaxation portion 121 that relaxes stress concentration between the first shroud side coupling portion 66 and the first shroud 60 is provided.
In the illustrated example, the stress concentration relaxation portion 121 is a concave curved surface that smoothly connects the outer peripheral surface 66b of the first shroud side connecting portion 66 and the surface 60b facing the opposite side of the combustion gas flow path F in the first shroud 60. It is composed of a shaped fillet.

According to the first stage stationary blade 120 described above, the same operational effects as those of the first embodiment can be obtained.
Further, the stress concentration between the first shroud side connecting portion 66 and the first shroud 60 can be relaxed, and the fatigue life of the first stage stationary blade 120 can be further improved.
Here, in this embodiment, since the stress concentration relaxation portion 121 is formed thick, thermal stress is likely to occur. However, stress concentration relaxation by connecting the projecting end portions 70 and 71, and the stress concentration relaxation portion 121. As a result, it is possible to improve the fatigue life in combination with the effect of relaxing the stress concentration by providing.

(Fifth embodiment)
Next, a fifth embodiment of the invention will be described with reference to the drawings. FIG. 7 is a view showing a first stage stationary blade according to a fifth embodiment of the present invention, and is an enlarged cross-sectional view of a portion corresponding to a portion X shown in FIG.
Note that in the fifth embodiment, the same components as those in the fourth embodiment are denoted by the same reference numerals, description thereof is omitted, and only different points will be described.

As shown in FIG. 7, in the first stage stationary blade 130 of the present embodiment, the first blade main body side connecting portion 58 is provided on the inner peripheral surface 66a of the first shroud side connecting portion 66 (the surface facing the blade main body side connecting portion). And the escape part 131 which suppresses the contact of the base ends of the 1st shroud side connection part 66 is provided.
In the illustrated example, the escape portion 131 is formed by a convex curved fillet that smoothly connects the inner peripheral surface 66a of the first shroud side connecting portion 66 and the surface 60a of the first shroud 60 on the combustion gas flow path F side. It is configured. That is, the escape portion 131 is configured by chamfering a portion located on the combustion gas flow path F side in the peripheral wall portion 63 a that defines the first insertion hole 63 of the first shroud 60.

According to the first stage stationary blade 130 shown above, the same operational effects as in the fourth embodiment can be obtained.
Further, since the escape portion 131 is provided on the inner peripheral surface 66a of the first shroud side connecting portion 66, fretting (friction) between the base ends of the first blade body side connecting portion 58 and the first shroud side connecting portion 66 is provided. Repetitive relative slip with small force and small amplitude) can be suppressed. Accordingly, fretting wear and fretting fatigue can be suppressed, and the fatigue life of the first stage stationary blade 130 can be further improved.

Further, by adopting the fillet shown in the present embodiment as the escape portion 131, the stress concentration between the first shroud side connecting portion 66 and the first shroud 60 can be further relaxed, and the one-stage stationary blade The fatigue life of 130 can be further improved.
In the present embodiment, the first stage stationary blade 130 includes the stress concentration relaxation unit 121, but the stress concentration relaxation unit 121 may not be provided. Further, the escape portion may be provided on the outer peripheral surface 58a of the first blade main body side connecting portion 58 (surface facing the shroud side connecting portion).

(Sixth embodiment)
Next, a sixth embodiment of the invention will be described with reference to the drawings. FIG. 8 is a view showing a first stage stationary blade according to a sixth embodiment of the present invention, and is an enlarged cross-sectional view of a portion corresponding to a portion X shown in FIG.
In the sixth embodiment, the same components as those in the second embodiment are denoted by the same reference numerals, description thereof is omitted, and only different points will be described.

As shown in FIG. 8, in the first stage stationary blade 110 of this embodiment, the turbine stationary blade body 53 is discharged from the first cooling hole 91 in the vicinity of the portion 92 connected to the first blade body side coupling portion 58. A cooling gas introduction part 111 into which the cooling gas G1 is introduced is provided.
In the illustrated example, the cooling gas introduction part 111 is formed by bending a part connected to the first shroud side coupling part 66 of the first shroud 60. That is, the first shroud 60 is connected to the first shroud side coupling portion 66 and is bent so as to be directed toward the other side H2 in the height direction H as it is separated from the turbine stationary blade body 53. And a second bent portion 113 that is connected to the first bent portion 112 and bent toward the other side H2 in the height direction H so as to be parallel to the height direction H. Yes. A space formed by the first bent portion 112, the second bent portion 113, and the portion 92 connected to the first blade body side coupling portion 58 of the turbine stationary blade body 53 is formed with the cooling gas introduction portion 111. It has become. In the present embodiment, both the first bent portion 112 and the second bent portion 113 are separated from the portion 92 connected to the first blade body side coupling portion 58 of the turbine stationary blade body 53.

  According to the first stage stationary blade 110 shown above, the same operational effects as those of the second embodiment can be obtained. Further, since both the first bent portion 112 and the second bent portion 113 are separated from the portion 92 connected to the first blade body side connecting portion 58 of the turbine stationary blade body 53, both the bent portions 112, By 113, the same effect as the escape part 131 shown in 5th Embodiment can be show | played.

  Further, since the cooling gas introduction part 111 is provided, the cooling gas G1 discharged from the first cooling hole 91 cools the part 92 connected to the first blade body side connection part 58 of the turbine stationary blade body 53. In doing so, the above-described cooling can be prevented from being inhibited by the combustion gas G1 flowing through the combustion gas flow path F. Accordingly, the thermal stress acting on the turbine stationary blade body 53 can be further suppressed, and the fatigue life of the first stage stationary blade 110 can be further improved.

  In the present embodiment, the cooling gas introduction part 111 includes a first bent part 112, a second bent part 113, and a part 92 connected to the first blade body side coupling part 58 of the turbine stationary blade body 53. Although the space is formed, the present invention is not limited to this. For example, the second bent portion 113 may not be provided. Further, instead of forming the bent portion in the first shroud 60, the cooling gas introduction portion may be formed by forming a concave portion.

  In the present embodiment, the bent portions 112 and 113 are both separated from the portion 92 connected to the first blade body side coupling portion 58 of the turbine stationary blade body 53, but the cooling gas introduction portion However, it is not restricted to this, For example, the part connected to the 1st blade main body side connection part of a turbine stationary blade main body, and the bending part may contact.

(Seventh embodiment)
Next, a seventh embodiment of the invention will be described with reference to the drawings. FIG. 9 is a view showing a first stage stationary blade according to a seventh embodiment of the present invention, and is an enlarged cross-sectional view of a portion corresponding to a portion X shown in FIG.
Note that in the seventh embodiment, the same components as those in the first embodiment are denoted by the same reference numerals, description thereof is omitted, and only different points will be described.

  As shown in FIG. 9, in the first stage stationary blade 140 of the present embodiment, the outer peripheral surface 53a of the turbine stationary blade body 53 and the surface 60a of the first shroud 60 on the combustion gas flow path F side have different thicknesses. A part side thermal barrier coating layer (surface treatment layer) 141 and a shroud side thermal barrier coating layer (surface treatment layer) 142 are provided. In the illustrated example, the shroud side thermal barrier coating layer 142 is formed thicker than the wing portion side thermal barrier coating layer 141.

According to the 1st stage stationary blade 140 shown above, there can exist an effect similar to 1st Embodiment.
Further, even when the thermal barrier coating layers 141 and 142 having different thicknesses are provided as described above, after the thermal barrier coating layers 141 and 142 are formed in advance on the turbine stationary blade body 53 and the first shroud 60, respectively, When the turbine stationary blade body 53 and the first shroud 60 are connected via the blade body side connection portion 57 and the first shroud side connection portion 66, for example, the turbine stationary blade body and the shroud are integrally formed. As compared with the above, the thermal barrier coding layers 141 and 142 having different thicknesses can be formed easily and accurately. Therefore, according to this configuration, while improving the fatigue life of the first stage stationary blade 140, the blade-side thermal barrier coating layer 141 having thicknesses suitable for the temperature conditions of the turbine stationary blade body 53 and the first shroud 60, and The shroud side thermal barrier coating layer 142 can be easily formed.
In the present embodiment, the shroud-side thermal barrier coating layer 142 is formed thicker than the wing-side thermal barrier coating layer 141. However, the present invention is not limited to this.

(Eighth embodiment)
Next, an eighth embodiment of the invention will be described with reference to the drawings. FIG. 10 is a view showing a first stage stationary blade according to an eighth embodiment of the present invention, and is a perspective view of a part of the first stage stationary blade. 11 is a cross-sectional view taken along arrow BB shown in FIG.
Note that in the eighth embodiment, the same components as those in the first embodiment are denoted by the same reference numerals, description thereof is omitted, and only different points will be described.

As shown in FIG. 11, in the first stage stationary blade 150 of the present embodiment, the cooling gas G <b> 2 is discharged into the combustion gas flow path F between the first blade main body side connecting portion 58 and the first shroud side connecting portion 66. A second cooling hole 151 for cooling a portion 155 connected to the blade body side coupling portion 58 of the turbine stationary blade body 53 is provided. The second cooling hole 151 is formed inside the first blade main body side connecting portion 58 and the first shroud side connecting portion 66 so as to communicate from the protruding end side to the proximal end side.
Further, as shown in FIG. 10, the welded portion 152 that connects the first blade main body side connecting portion 58 and the first shroud side connecting portion 66 is a portion located on the front side of the first blade main body side connecting portion 58. In the edge part 153 and the 1st shroud side connection part 66, it is provided so that the front side part 154 corresponding to the front edge part 153 may be excluded. The second cooling hole 151 is provided between the front edge portion 153 of the first blade main body side connecting portion 58 and the front portion 154 of the first shroud side connecting portion 66.

  Here, as described above, the pressure P1 of the cooling gas G2 in the external A partitioned from the combustion gas flow path F by the first shroud 60 is higher than the pressure P2 of the combustion gas G1 in the combustion gas flow path F. Therefore, the cooling gas G2 outside the first shroud 60 is discharged into the combustion gas flow path F by forming the second cooling holes 151 as described above.

According to the first stage stationary blade 150 shown above, the same operational effects as those of the first embodiment can be obtained.
Further, since the second cooling hole 151 is provided in the first shroud 60, the portion 155 connected to the blade body side coupling portion 58 of the turbine stationary blade body 53 can be cooled. Accordingly, the thermal stress acting on the turbine stationary blade body 53 can be further suppressed, and the fatigue life of the first stage stationary blade 150 can be further improved.

(Ninth embodiment)
Next, a ninth embodiment of the invention will be described with reference to the drawings. FIG. 12 is a diagram showing a first stage stationary blade according to the ninth embodiment of the present invention, and is a perspective view of a part of the first stage stationary blade. 13 is a diagram showing the first stage stationary blade shown in FIG. 12, and is an enlarged cross-sectional view of a portion corresponding to part X shown in FIG.
Note that in the ninth embodiment, the same components as those in the first embodiment are denoted by the same reference numerals, description thereof is omitted, and only different points will be described.

As shown in FIG. 13, in the first stage stationary blade 160 of the present embodiment, the first shroud side connecting portion 66 has a positioning portion 161 that positions the turbine stationary blade body 53 in the height direction H with respect to the first shroud 60. Is provided.
As shown in FIGS. 12 and 13, in the illustrated example, the projecting end portion 71 of the first shroud side coupling portion 66 is bent toward the turbine stationary blade body 53 side over the entire circumference to form the positioning portion 161. ing. As shown in FIG. 13, the positioning part 161 has a surface 162 that faces the other side H <b> 2 in the height direction H and faces the other side H <b> 1 in the height direction H of the turbine vane body 53. , The turbine stationary blade body 53 is positioned in the height direction H. In addition, a notch 164 opened on the protruding end side of the positioning portion 161 is provided on the surface 162 of the positioning portion 161 facing the other side H2 in the height direction H, and the first blade main body side connecting portion 58 is provided. And the 1st shroud side connection part 66 is mutually connected in each protrusion part 70,71 by forming the brazing part 165 in this notch part 164. As shown in FIG.

  When assembling the turbine stationary blade body 53 and the first shroud 60 when manufacturing the first stage stationary blade 160 configured as described above, the surface 162 of the positioning portion 161 facing the other side H2 in the height direction H, the turbine A surface 163 of the stationary blade body 53 facing the one side H1 in the height direction H is brought into contact. Thereby, the turbine stationary blade body 53 and the first shroud 60 can be easily positioned, the assembly of both can be facilitated, and the same operational effects as in the first embodiment can be achieved.

The technical scope of the present invention is not limited to the above embodiment, and various modifications can be made without departing from the spirit of the present invention.
For example, in each of the embodiments described above, the case where the gas turbine blade according to the present invention is adopted as the first stage stationary blade has been described as an example. However, the present invention is not limited to this. The blade 5, the turbine blade 6, the compressor stationary blade 21, and the compressor blade 22 may be employed. In the turbine stationary blade 5, in the first-stage stationary blade and the second-stage stationary blade positioned on the front side, particularly high-temperature combustion gas G1 passes, so that the adoption of the gas turbine blade according to the present invention for the turbine stationary blade 5 It is effective against fatigue caused by thermal stress.
Moreover, in each said embodiment, although both the 1st shroud 60 and the 2nd shroud 61 shall be provided, it is not restricted to this, Only any one may be sufficient.

  In each of the above embodiments, the turbine stationary blade body 53 is formed such that the outer peripheral surface 53a and the inner peripheral surface 53b in a cross-sectional view along the height direction H are parallel to the height direction H. It was supposed to be, but it is not limited to this. FIG. 14 is a cross-sectional view showing a modification of the first stage stationary blade according to the present invention. For example, the turbine stationary blade body 171 may be formed to be curved like a first stage stationary blade 170 shown in FIG. Even in this case, each blade main body side connecting portion 58, 59 and each shroud side connecting portion 66, 67 are extended in the height direction H of the turbine stationary blade main body 53. Regardless of the shape, the connecting portions 58, 59, 66, 67 can be easily connected.

  Moreover, in each said embodiment, although the welding parts 75 and 152 and the brazing part 165 were used as a connection means which each connects the protrusion parts 70 and 71, a connection means is not restricted to these, For example, a bolt and The turbine vane main body and the shroud may be coupled by mechanical coupling using a lock nut or the like. Moreover, in any case, the projecting ends 70 and 71 may be connected by only a part of each.

  Moreover, the gas turbine 1 is not restricted to the form shown by the said embodiment. For example, although the axial flow type compressor is adopted as the compressor 2 in the embodiment, a centrifugal type compressor may be adopted. Moreover, in the said embodiment, although the 1 axis type was employ | adopted as a gas turbine, you may employ | adopt a multi-shaft type (free turbine).

  In addition, it is possible to appropriately replace the constituent elements in the embodiment with known constituent elements without departing from the spirit of the present invention, and the above-described modified examples may be appropriately combined.

1 is a schematic half sectional view showing a gas turbine according to a first embodiment of the present invention. FIG. 2 is a perspective view of a part of a first stage stationary blade of the gas turbine shown in FIG. 1. It is AA arrow sectional drawing shown in FIG. It is a figure which shows the 1st stage stationary blade of 2nd Embodiment which concerns on this invention, Comprising: It is an expanded sectional view of the part applicable to the X section shown in FIG. It is a figure which shows the 1st stage stationary blade of 3rd Embodiment which concerns on this invention, Comprising: It is an expanded sectional view of the part applicable to the X section shown in FIG. It is a figure which shows the 1st stage stationary blade of 4th Embodiment which concerns on this invention, Comprising: It is an expanded sectional view of the part applicable to the X section shown in FIG. It is a figure which shows the 1st stage stationary blade of 5th Embodiment which concerns on this invention, Comprising: It is an expanded sectional view of the part applicable to the X section shown in FIG. It is a figure which shows the 1st stage stationary blade of 6th Embodiment which concerns on this invention, Comprising: It is an expanded sectional view of the part applicable to the X section shown in FIG. It is a figure which shows the 1st stage stationary blade of 7th Embodiment which concerns on this invention, Comprising: It is an expanded sectional view of the part applicable to the X section shown in FIG. It is a figure which shows the 1st stage stationary blade of 8th Embodiment which concerns on this invention, Comprising: It is a perspective view of a part of 1st stage stationary blade. It is BB arrow sectional drawing shown in FIG. It is a figure which shows the 1st stage stationary blade of 9th Embodiment which concerns on this invention, Comprising: It is a perspective view of a part of 1st stage stationary blade. It is a figure which shows the 1st stage stationary blade shown in FIG. 12, Comprising: It is an expanded sectional view of the part applicable to the X section shown in FIG. It is sectional drawing which shows the modification of the 1st stage stationary blade which concerns on this invention.

Explanation of symbols

DESCRIPTION OF SYMBOLS 1 Gas turbine 2 Compressor 3 Combustor 4 Turbine 8, 90, 100, 110, 120, 130, 140, 150, 160, 170 1st stage stationary blade (gas turbine blade)
53,171 Turbine stationary blade body (blade body)
53a Outer peripheral surface of turbine vane body (outer surface of blade body)
58 1st wing body side connection (wing body side connection)
58a Outer peripheral surface of first blade main body connecting portion (surface facing shroud side connecting portion)
59 Second wing body side connection (wing body side connection)
60 First shroud (shroud)
60a Combustion gas flow path side 61 Second shroud (shroud)
66 1st shroud side connection part (shroud side connection part)
66a Inner peripheral surface of the first shroud side connecting portion (surface facing the blade body side connecting portion)
67 2nd shroud side connection part (shroud side connection part)
70, 71 Protruding end portion 72, 73 Base end portion 91 First cooling hole 92, 155 Portion connected to first blade body side connecting portion 111 Cooling gas introduction portion 112, 113 Bending portion (relief portion)
101 Thick part 131 Escape part 141 Wing side thermal barrier coating layer (surface treatment layer)
142 Shroud side thermal barrier coating layer (surface treatment layer)
151 2nd cooling hole 161 Positioning part F Combustion gas flow path G1 Combustion gas G2 Cooling gas H Height direction

Claims (9)

A blade body through which the combustion gas flowing from the upstream side toward the downstream side in the gas turbine passes through the outer surface;
A shroud that is disposed at at least one of both end portions in the height direction of the blade body to support the blade body, and constitutes a part of a partition wall of a combustion gas passage through which the combustion gas flows;
With
The blade body and the shroud are each provided with a blade body side connection portion and a shroud side connection portion that extend in the height direction outside the combustion gas flow path,
The blade main body side connecting portion and the shroud side connecting portion are connected at a part of each protruding end protruding outside the combustion gas flow path, and are connected at a base end located on the combustion gas flow path side. A gas turbine blade characterized by not. The gas turbine blade according to claim 1,
The shroud is provided with a first cooling hole for discharging a cooling gas into the combustion gas flow path and cooling a portion of the blade body connected to the blade body side connecting portion. Gas turbine blade. The gas turbine blade according to claim 2,
The gas turbine blade according to claim 1, wherein a portion connected to the blade body side coupling portion of the blade body is provided with a thick portion formed thicker than other portions. In the gas turbine blade according to claim 2 or 3,
A cooling gas introduction part for introducing the cooling gas discharged from the first cooling hole is provided in the vicinity of a portion connected to the blade body side coupling part of the blade body. Gas turbine blade. In the gas turbine blade according to any one of claims 1 to 4,
At least one of the mutually opposing surfaces of the wing body side connecting portion and the shroud side connecting portion is provided with a relief portion that suppresses contact between the base ends of the wing body side connecting portion and the shroud side connecting portion. A gas turbine blade characterized in that: The gas turbine blade according to any one of claims 1 to 5,
The gas turbine blade according to claim 1, wherein a surface treatment layer having a different thickness is provided on an outer surface of the blade body and a surface of the shroud on the combustion gas flow path side. In the gas turbine blade according to any one of claims 1 to 6,
Between the blade main body side connecting portion and the shroud side connecting portion, a cooling gas is discharged into the combustion gas flow path to cool a portion connected to the blade main body side connecting portion of the blade main body. A gas turbine blade provided with a cooling hole. The gas turbine blade according to any one of claims 1 to 7,
The gas turbine blade according to claim 1, wherein the shroud side connecting portion is provided with a positioning portion for positioning the blade body in the height direction with respect to the shroud. A compressor for generating compressed air;
A combustor that supplies fuel to compressed air supplied from the compressor to generate combustion gas;
A turbine having the gas turbine blade according to any one of claims 1 to 8, wherein the turbine generates rotational power by the combustion gas supplied from the combustor,
A gas turbine comprising:

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