JP2020094508A - Turbine stationary blade and gas turbine - Google Patents

Abstract

To provide a turbine stationary blade that can suppress combustion vibration between outlet parts of a combustor due to acoustic propagation via a clearance between the combustor and a stationary blade, and to provide a gas turbine.SOLUTION: A turbine stationary blade includes a hollow blade body. The blade body includes: a first part partially forming an airfoil profile including a pressure surface, a suction surface and a trailing edge; and a second part positioned on a side of a leading-edge of the airfoil profile relative to the first part and having a projection part. The second part includes a recess part at the outside surface and the inside surface thereof.SELECTED DRAWING: Figure 6

Description

The present disclosure relates to a turbine vane and a gas turbine

In a gas turbine including a plurality of combustors, combustion vibration may occur near the outlet of the combustor due to acoustic propagation between the combustors. Since such combustion vibration can be a factor that hinders stable operation of the gas turbine, measures for suppressing combustion vibration that occurs near the outlet of the combustor are being studied.

For example, in the gas turbine described in Patent Document 1, the combustor transition portion (tail tube) has a sidewall extension portion that extends axially beyond the combustor outlet portion and into the space on the turbine side. The side wall extension extends to the vicinity of the turbine vane upstream of the turbine vane arranged on the most upstream side of the turbine, and extends in the annular space between the combustor outlet and the turbine vane in the circumferential direction. It is partially partitioned in. As a result, thermoacoustic pulsation due to thermoacoustic connection between the outlets of the adjacent combustors is suppressed.

Japanese Patent No. 5726267

In the gas turbine described in Patent Document 1, there is an axial position where no partition exists between the transition piece of the combustor and the turbine vane, and at the axial position, a gap between the combustor and the turbine vane. Are linearly formed in the circumferential direction, the effect of suppressing acoustic propagation between adjacent combustors can be limited.

In view of the above-mentioned circumstances, at least one embodiment of the present invention provides a turbine vane capable of suppressing combustion vibration between the outlets of the combustor due to acoustic propagation through a gap between the combustor and the turbine vane. And to provide a gas turbine.

(1) A turbine vane according to at least one embodiment of the present invention comprises:
A turbine stationary blade having a hollow blade body,
The wing body is
A first portion partially forming an airfoil including a pressure surface, a suction surface and a trailing edge;
A second portion located on the leading edge side of the airfoil with respect to the first portion and having a convex portion;
Including,
The outer or inner surface of the second portion includes a recess.

According to the turbine vane described in (1) above, when the convex portion of the second portion of the turbine vane is fitted to the radial wall of the combustor, the convex portion of the turbine vane and combustion The radial wall of the container can be axially overlapped. As a result, it is possible to prevent the gap between the combustor and the turbine vane from being linearly formed in the circumferential direction. Therefore, it is possible to reduce the combustion vibration between the outlets of the combustor due to the acoustic propagation through the gap between the combustor and the turbine vane.

On the other hand, the temperature of the portion of the turbine vane that is not exposed to the flow of the combustion gas of the turbine tends to be lower than the temperature of the portion of the turbine vane that is exposed to the flow of the combustion gas. Therefore, the thermal expansion of the portion of the turbine vane that is exposed to the flow of combustion gas is restricted to the portion that is not exposed to the flow of combustion gas (the portion where the temperature is relatively low), and the blade surface of the blade body (particularly the pressure surface). Large compressive stress is likely to occur on the side).

In this regard, according to the turbine vane described in (1) above, since the outer surface or the inner surface of the second portion including the convex portion includes the concave portion, the rigidity of the second portion including the convex portion can be reduced. .. As a result, the constraint on the thermal expansion is weakened, the compressive stress generated on the blade surface is reduced, and the damage of the turbine vane can be suppressed without increasing the cooling air for cooling the blade body.

(2) In some embodiments, in the turbine vane described in (1) above,
The recess is formed in a range including the central position in the blade height direction.

According to the turbine vane described in (2) above, it is possible to reduce the compressive stress generated on the blade surface while suppressing the influence of forming the recess on the rigidity of the fillet portion.

(3) In some embodiments, in the turbine vane described in (1) or (2) above,
A shroud wall portion which is connected to an end portion of the wing body to form a flow path wall,
A fillet portion provided at a corner formed by the wing surface of the wing body and the wall surface of the shroud wall portion,
Equipped with
In the blade height direction, the range S1 does not overlap with at least a part of the range S2, where S1 is the range of the recess and S2 is the range of the upstream end surface of the fillet portion.

If the concave portion is formed over the entire range of the blade body in the blade height direction, the rigidity of the fillet portion, where the compressive stress is likely to increase, may decrease, and the stress of the fillet portion may increase. Therefore, by forming a recess so that the range S1 does not overlap at least a part of the range S2 as in the above (3), the compressive stress generated on the blade surface is suppressed while suppressing the influence on the rigidity of the fillet part. Therefore, the turbine vane can be prevented from being damaged.
The "shroud wall portion" in (3) above may be, for example, an "outer shroud wall portion 60" or an "inner shroud wall portion 62" described later.

(4) In some embodiments, in the turbine vane according to any one of (1) to (3) above,
The recess is formed on an outer surface or an inner surface of a portion of the second portion that is not exposed to the flow of the combustion gas of the turbine.

According to the turbine vane described in (4) above, the rigidity of the portion of the second portion that is not exposed to the flow of combustion gas can be reduced, so that the thermal expansion of the portion that is exposed to the flow of combustion gas is It is possible to weaken the restraint by the portion not exposed to the flow of the combustion gas (the portion having a relatively low temperature). As a result, the compressive stress generated on the blade surface can be reduced, and damage to the turbine vane can be suppressed.

(5) In some embodiments, in the turbine stationary blade according to any one of (1) to (4) above,
The concave portion is formed in a portion of the inner surface of the second portion that corresponds to the convex portion (a portion behind the convex portion).

By fitting the convex portion to the radial wall portion of the combustor, the convex portion is particularly unlikely to be affected by the flow of the combustion gas, and the temperature tends to be lower than that around the convex portion. In this regard, according to the turbine vane described in (5) above, since the concave portion is formed in the portion corresponding to the convex portion on the inner surface of the second portion, the convex portion against the thermal expansion of the portion around the convex portion is formed. Can restrain the restraint by. As a result, the compressive stress generated around the convex portion can be effectively reduced, and the turbine vane can be prevented from being damaged.

(6) In some embodiments, in the turbine stationary blade according to any one of (1) to (5) above,
The recess is formed on the inner surface of the second portion,
In the circumferential direction of the rotor, the range S3 is located inside the range S4, where S3 is the range of the recesses and S4 is the range of the protrusions.

According to the turbine vane described in (6) above, since the range S3 is located inside the range S4, it is possible to weaken the restraint by the convex portion on the thermal expansion of the portion around the convex portion. As a result, the compressive stress generated around the convex portion can be effectively reduced, and the turbine vane can be prevented from being damaged.

(7) In some embodiments, in the turbine vane according to (5) or (6) above,
When the depth of the concave portion is d and the thickness of the portion of the second portion adjacent to the convex portion is t, d>t is satisfied.

According to the turbine vane described in (7) above, the rigidity around the convex portion and the convex portion can be effectively reduced. As a result, the compressive stress generated around the convex portion can be effectively reduced, and the turbine vane can be prevented from being damaged.

(8) In some embodiments, in the turbine vane according to any one of (1) to (4) above,
The concave portion is formed on the outer surface of the convex portion.

According to the turbine vane described in (8) above, since the concave portion is formed on the outer surface of the convex portion, it is possible to reduce the rigidity of the convex portion in which the temperature tends to be relatively low, and the circumference of the convex portion can be reduced. It is possible to weaken the restraint of the convex portion on the thermal expansion of the portion. Thereby, the compressive stress generated around the convex portion can be effectively reduced, and the damage of the turbine vane A can be suppressed.

(9) In some embodiments, in the turbine vane according to any one of (1) to (4) above,
The second portion includes a flat portion in which a flat surface adjacent to the convex portion is formed on an outer surface,
At least a part of the concave portion is formed on the inner surface of the flat portion.

According to the turbine vane described in (9) above, the rigidity of the second portion including the flat portion can be reduced by forming at least a part of the concave portion on the inner surface of the flat portion adjacent to the convex portion. .. As a result, the compressive stress generated on the blade surface can be reduced, and damage to the turbine vane can be suppressed.

(10) A gas turbine according to at least one embodiment of the present invention is
A plurality of combustors arranged in the circumferential direction of the rotor, including an outlet portion having a radial wall portion along the radial direction of the rotor;
The turbine stationary blade according to any one of (1) to (9), which is located on a downstream side of a pair of the radial wall portions facing each other of the outlet portions of the combustors adjacent to each other in the circumferential direction. ,
Equipped with.

According to the gas turbine described in (10) above, since the turbine vane described in any one of (1) to (9) is provided, the combustion vibration caused by the acoustic propagation between the outlets of the plurality of combustors. Can be reduced. Further, it is possible to reduce the compressive stress generated on the blade surface and suppress the damage of the turbine vane.

(11) In some embodiments, in the gas turbine according to (10) above,
The radial wall portion is fitted with the convex portion.

According to the gas turbine described in (11) above, it is possible to prevent the gap between the combustor and the turbine stationary blade from being linearly formed in the circumferential direction, and therefore, between the outlet portions of the plurality of combustors. It is possible to reduce combustion vibration due to sound propagation.

(12) In some embodiments, in the gas turbine according to (10) or (11) above,
The downstream end surface of the radial wall portion includes a convex portion receiving groove along the radial direction of the rotor, and the convex portion receiving groove and at least a part of the convex portion partially overlap in the axial direction of the rotor. doing.

According to the gas turbine described in (12) above, it is possible to prevent the gap between the combustor and the turbine vane from being linearly formed in the circumferential direction, and therefore, between the outlet portions of the plurality of combustors. It is possible to reduce combustion vibration due to sound propagation.

According to at least one embodiment of the present invention, there is provided a turbine vane and a gas turbine capable of suppressing combustion vibration between the outlets of the combustor due to acoustic propagation through a gap between the combustor and the vane. To be done.

It is a schematic structure figure of gas turbine 1 concerning one embodiment. It is a schematic diagram showing an inlet portion of combustor 4 and turbine 6 of gas turbine 1 concerning one embodiment. It is a figure which shows the structure of the exit part 52 of the combustor 4 (combustor assembly 10) which concerns on one Embodiment, and shows two adjacent combustors among the several combustor 4 arrange|positioned in the circumferential direction. .. FIG. 6 is a sectional view taken along the circumferential direction, and FIG. 5 is a sectional view taken along the radial direction. FIG. 3 is a schematic diagram showing a configuration of an outlet portion 52 of a combustor 4 and an inlet portion of a turbine 6 of a gas turbine 1 according to an embodiment. It is an expanded sectional view which shows the specific structural example of 24 A of turbine stationary blades shown in FIG. It is a figure which shows an example of the radial direction range in which the recessed part 18 in the form shown in FIG. 6 is formed. It is an expanded sectional view which shows the specific structural example of 24 A of turbine stationary blades shown in FIG. It is an expanded sectional view which shows the specific structural example of 24 A of turbine stationary blades shown in FIG. It is an expanded sectional view which shows the specific structural example of 24 A of turbine stationary blades shown in FIG. It is a figure which shows an example of the radial direction range in which the recessed part 18 in the form shown in FIG. 10 is formed. It is a figure which shows an example of the radial direction range in which the recessed part 18 in the form shown in FIG. 10 is formed.

Hereinafter, some embodiments of the present invention will be described with reference to the accompanying drawings. However, the dimensions, materials, shapes, relative positions, etc. of the components described as the embodiments or shown in the drawings are not intended to limit the scope of the present invention thereto, but are merely illustrative examples. Absent.
For example, the expressions representing relative or absolute arrangements such as "in a certain direction", "along a certain direction", "parallel", "orthogonal", "center", "concentric", or "coaxial" are strict. In addition to representing such an arrangement, it also represents a tolerance, or a state in which the components are relatively displaced at an angle or distance such that the same function can be obtained.
For example, expressions such as "identical", "equal", and "homogeneous" that indicate that they are in the same state are not limited to a state in which they are exactly equal to each other. It also represents the existing state.
For example, the representation of a shape such as a quadrangle or a cylinder does not only represent a shape such as a quadrangle or a cylinder in a geometrically strict sense, but also an uneven portion or a chamfer within a range in which the same effect can be obtained. The shape including parts and the like is also shown.
On the other hand, the expressions “comprising”, “comprising”, “comprising”, “including”, or “having” one element are not exclusive expressions excluding the existence of other elements.

FIG. 1 is a schematic configuration diagram of a gas turbine 1 according to an embodiment.
As shown in FIG. 1, a gas turbine 1 is driven by a compressor 2 for generating compressed air, a combustor 4 for generating combustion gas using the compressed air and fuel, and rotationally driven by the combustion gas. And a turbine 6 configured as described above. In the case of the gas turbine 1 for power generation, an unillustrated generator is connected to the turbine 6.

Fuel and compressed air generated by the compressor 2 are supplied to the combustor 4, the fuel is combusted in the combustor 4, and combustion gas that is a working fluid of the turbine 6 is generated. To be done. As shown in FIG. 1, the gas turbine 1 has a plurality of combustors 4 arranged in a casing 20 along the circumferential direction centering on a rotor 8.

The turbine 6 has a combustion gas passage 28 formed by the turbine casing 22, and includes a plurality of turbine vanes 24 and turbine rotor blades 26 provided in the combustion gas passage 28. The turbine vanes 24 are supported from the turbine casing 22 side, and the plurality of turbine vanes 24 arranged along the circumferential direction of the rotor 8 form a vane row. The turbine rotor blades 26 are planted in the rotor 8, and the plurality of turbine rotor blades 26 arranged along the circumferential direction of the rotor 8 form a rotor blade row. The stationary blade rows and the moving blade rows are arranged alternately in the axial direction of the rotor 8.

In the turbine 6, the combustion gas from the combustor 4 flowing into the combustion gas passage 28 passes through the plurality of turbine stationary blades 24 and the plurality of turbine moving blades 26, so that the rotor 8 is rotationally driven and connected to the rotor 8. The generator is driven to generate electric power. The combustion gas after driving the turbine 6 is discharged to the outside through the exhaust casing 30.

Hereinafter, the axial direction of the gas turbine 1 (axial direction of the rotor 8) is simply referred to as “axial direction”, and the radial direction of the gas turbine 1 (radial direction of the rotor 8) is simply described as “radial direction”. The circumferential direction of the turbine 1 (the circumferential direction of the rotor 8) will be simply referred to as the “circumferential direction”. Regarding the flow direction of the combustion gas in the combustion gas passage 28, the upstream side in the axial direction is simply referred to as “upstream side”, and the downstream side in the axial direction is simply referred to as “downstream side”.

FIG. 2 is a schematic diagram showing an inlet portion of the combustor 4 and the turbine 6 of the gas turbine 1 according to the embodiment.
As shown in FIG. 2, each of the plurality of combustors 4 (see FIG. 1) arranged annularly around the rotor 8 includes a combustor liner 36 provided in a combustor casing 32 defined by a casing 20. , And a plurality of second combustion burners 40 arranged so as to surround the first combustion burner 38 and the first combustion burner 38, respectively. The combustor 4 may include other components such as a bypass pipe (not shown) for bypassing the combustion gas.

The combustor liner 36 includes an inner cylinder 48 arranged around the first combustion burner 38 and the plurality of second combustion burners 40, and a transition cylinder 50 connected to a tip portion of the inner cylinder 48. .. The inner cylinder 48 and the transition cylinder 50 may form an integral combustion cylinder.

Each of the first combustion burner 38 and the second combustion burner 40 includes a fuel nozzle (not shown) for injecting fuel, and a burner cylinder (not shown) arranged so as to surround the fuel nozzle. Fuel is supplied to each fuel nozzle via fuel ports 42 and 44, respectively. Further, the compressed air generated by the compressor 2 (see FIG. 1) is supplied into the combustor casing 32 through the casing inlet 41, and the compressed air is supplied from the combustor casing 32 to each burner cylinder. It is supposed to flow in. Then, in each of the burner cylinders, the fuel injected from the fuel nozzle and the compressed air are mixed, and this air-fuel mixture flows into the combustor liner 36, is ignited, and burns to generate combustion gas. There is.

The first combustion burner 38 may be a burner for generating a diffusion combustion flame, and the second combustion burner 40 may be a burner for burning a premixed gas to generate a premixed combustion flame. .. That is, in the second combustion burner 40, the fuel from the fuel port 44 and the compressed air are premixed, and the premixed air mainly forms a swirling flow by the swirler (not shown) and flows into the combustor liner 36.

Further, the compressed air and the fuel injected from the first combustion burner 38 via the fuel port 42 are mixed in the combustor liner 36, ignited by a pilot fire (not shown) and burned to generate combustion gas. At this time, a part of the combustion gas diffuses to the surroundings with a flame, so that the premixed gas flowing into the combustor liner 36 from each second combustion burner 40 is ignited and burned. That is, the flame holding for stable combustion of the premixed gas (premixed fuel) from the second combustion burner 40 can be performed by the diffusion combustion flame by the diffusion combustion fuel injected from the first combustion burner 38. .. At that time, the combustion region may be formed, for example, in the inner cylinder 48 and not in the transition piece 50.

The combustion gas generated by the combustion of the fuel in the combustor 4 as described above passes through the outlet portion 52 of the combustor 4 located at the downstream end portion of the transition piece 50 (the downstream end portion of the combustor 4) to It flows into the turbine vane 24 of the stage.

FIG. 3 is a diagram showing a configuration of the outlet portion 52 of the combustor 4 (combustor assembly 10) according to the embodiment, and two adjacent combustors among the plurality of combustors 4 arranged in the circumferential direction. Is shown. FIG. 4 and FIG. 5 are schematic diagrams respectively showing configurations of the outlet portion 52 of the combustor 4 and the inlet portion of the turbine 6 of the gas turbine 1 according to the embodiment. Of these, FIG. 4 is a sectional view taken along the circumferential direction, and FIG. 5 is a sectional view taken along the radial direction.

For example, as shown in FIG. 3, the outlet portion 52 of each combustor 4 has radial wall portions 54a and 54b extending in the radial direction and circumferential wall portions 56a and 56b extending in the circumferential direction. And, including. Here, among the outlet portions 52 of the combustors 4 that are adjacent to each other in the circumferential direction, the radial wall portion 54a of one combustor 4 and the radial wall portion 54b of the other combustor 4 are a pair of pairs facing each other. The radial wall portions 54a and 54b. At the outlet 52 of each combustor 4, the corners 132 of the radial walls 54a, 54b and the circumferential walls 56a, 56b are rounded.

For example, as shown in FIG. 4, the gas turbine 1 includes a plurality of combustors 4 arranged in the circumferential direction and a first stage turbine vane 24(1) located downstream of an outlet 52 of the combustor 4. Step vanes), and. That is, the combustor 4 and the turbine vane 24 are provided separately. A plurality of combustors 4 arranged in the circumferential direction constitutes a combustor assembly 10.

The first-stage plurality of turbine vanes 24 arranged along the circumferential direction include the turbine vanes 24A located downstream of the pair of radial wall portions 54a and 54b described above. The plurality of turbine vanes 24 of the first stage further include another turbine vane 24B provided at a circumferential position between the pair of turbine vanes 24A, 24A that are adjacent to each other in the circumferential direction. The first-stage turbine vane 24A extends to the upstream side of the leading edge of the other first-stage turbine vane 24B. The alternate long and short dash line L1 in FIG. 4 indicates the position of the leading edge of the other first stage turbine vane 24B in the axial direction. In the illustrated exemplary embodiment, the turbine vanes 24A and the other turbine vanes 24B are arranged alternately in the circumferential direction.

For example, as shown in FIG. 5, the turbine vane 24A has an outer shroud wall portion that is connected to the blade body 70 and an outer end portion 80 of the blade body 70 in the radial direction to form a flow passage wall surface 81 on the outer side in the radial direction. 60, and an inner shroud wall portion 62 that connects to the inner end portion 82 of the airfoil 70 in the radial direction and forms a flow passage wall surface 83 on the inner side in the radial direction. The turbine vane 24 is supported in the turbine casing 22 (see FIG. 1) via the outer shroud wall portion 60.

An outer fillet portion 103 is provided at a corner portion 102 formed by the blade surface 71 of the blade body 70 (the surface of the blade body 70) and the flow passage wall surface 81 of the outer shroud wall portion 60. An inner fillet portion 107 is provided at a corner portion 106 formed by the blade surface 71 and the flow passage wall surface 83 of the inner shroud wall portion 62. The outer fillet portion 103 has a fillet radius widening portion 104 in which the fillet radius increases toward the upstream side so as to eliminate or reduce a step with the corresponding corner portion 132 (see FIG. 3) in the outlet portion 52 of the combustor 4. including. The inner fillet portion 107 includes a fillet radius widening portion 108 in which the fillet radius increases toward the upstream side so as to eliminate or reduce the step difference with the corresponding corner portion 132 in the outlet portion 52 of the combustor 4.

As shown in FIG. 4, the blade body 70 is formed in a hollow shape, and is configured so that cooling air for cooling the blade body 70 flows inside. The airfoil 70 is located on the leading edge side of the airfoil 78 with respect to the first portion 12, which partially forms an airfoil 78 including a pressure surface 72, a suction surface 74, and a trailing edge 76, The second portion 16 having the convex portion 14 is included.

In the illustrated exemplary embodiment, the second portion 16 is an upstream end portion of the blade body 70 in the axial direction, and includes a pair of flat portions 66 each having a flat surface 64 adjacent to the convex portion 14 formed on the outer surface thereof. .. The convex portion 14 is located between the pair of flat portions 66 in the circumferential direction, and the convex portion 14 projects toward the upstream side in the axial direction with respect to the flat surface 64. The convex portion 14 includes a top surface 88 that is substantially orthogonal to the axial direction and a pair of side surfaces 90 that are substantially orthogonal to the circumferential direction and substantially parallel to each other. The pair of side surfaces 90 are connected to the pair of flat surfaces 64, respectively. There is. As shown in FIG. 5, the convex portion 14 linearly extends in a radial direction from the outer shroud wall portion 60 to the inner shroud wall portion, and the radial wall portion 54 a of the outlet portion 52 of the combustor 4 is formed. , 54b are fitted in convex receiving grooves 58 formed on the downstream end surfaces along the radial direction. As a result, a part of the convex portion 14 and the flat surface 64 becomes a portion 86 that is not exposed to the combustion gas F. Further, the portion 86 that is not exposed to the combustion gas F (flow in the axial direction) refers to a portion of the blade body 70 that has a relatively low temperature.

As shown in FIG. 4, according to the turbine vane 24A, when the convex portion 14 of the second portion 16 of the turbine vane 24A is fitted to the radial wall portions 54a and 54b of the combustor 4. The convex portion 14 of the turbine vane 24A and the convex portion receiving groove 58 of the radial wall portions 54a and 54b of the combustor 4 can partially overlap each other in the axial direction. This can prevent the gap between the combustor 4 and the turbine vane 24A from being linearly formed in the circumferential direction (strictly, in an arc shape), so that the outlet portions 52 of the plurality of combustors 4 are formed. It is possible to reduce combustion vibration due to sound propagation between the two.

FIG. 6 is an enlarged cross-sectional view showing a specific configuration example of the turbine vane 24A shown in FIG.
In some embodiments, for example, as shown in FIG. 6, the inner surface 84 of the second portion 16 of the turbine vane 24A includes the recess 18. In the illustrated form, the recess 18 is formed on the inner surface 84 of the portion 86 of the second portion 16 that is not exposed to the flow F (axial flow) of the combustion gas of the turbine 6.

In the exemplary embodiment shown in FIG. 6, when the existence range of the concave portion 18 is S3 and the existence range of the convex portion 14 is S4 in the circumferential direction, the range S3 is located inside the range S4. That is, the concave portion 18 is formed in the portion 92 (the rear side portion of the convex portion 14) corresponding to the convex portion 14 in the inner surface 84 of the second portion 16. Further, when the depth of the concave portion 18 is d and the thickness of the flat portion 66 of the second portion 16 adjacent to the convex portion 14 is t, d>t is satisfied. Further, the concave portion 18 communicates with a cooling flow path 19 formed inside the blade body 70 for passing cooling air.

In the turbine vane 24A, the convex portion 14 and the flat portion 66 adjacent to the convex portion 14 become a portion 86 that is not exposed to the flow F of the combustion gas of the turbine 6, and the pressure surface 72, the negative pressure surface 74, the trailing edge 76, and the flat surface. The remote side of the convex portion 14 of the portion 66 is a portion 87 exposed to the flow F of the combustion gas of the turbine 6. The temperature of the portion 86 of the turbine vane 24A that is not exposed to the combustion gas flow F of the turbine 6 is likely to be lower than the temperature of the portion 87 of the turbine vane 24A that is exposed to the combustion gas flow F. Therefore, the thermal expansion of the portion 87 of the turbine vane 24A exposed to the flow F of the combustion gas is restrained by the portion 86 not exposed to the flow F of the combustion gas (a portion having a relatively low temperature), and A large compressive stress is likely to occur particularly on the pressure surface 72 side of the blade surface 71.

In this regard, in the embodiment shown in FIG. 6, by forming the concave portion 18 on the inner surface 84 of the second portion 16 including the convex portion 14, it is possible to reduce the rigidity of the second portion 16 including the convex portion 14. As a result, the constraint on the thermal expansion can be weakened and the compressive stress generated on the blade surface 71 can be reduced. Therefore, damage to the turbine vane 24A can be suppressed without increasing the cooling air for cooling the blade body 70. be able to.

Further, since the convex portion 14 is fitted to the radial wall portions 54a and 54b of the combustor 4, the convex portion 14 is particularly unlikely to be affected by the flow F of the combustion gas, and is likely to have a temperature lower than that around the convex portion 14. Therefore, by forming the concave portion 18 so that the range S3 is located inside the range S4 as described above, the restraint by the convex portion 14 on the thermal expansion of the portion around the convex portion 14 can be weakened. Thereby, the compressive stress generated around the convex portion 14 can be effectively reduced, and the damage of the turbine vane 24A can be suppressed.

Further, since the concave portion 18 is formed so as to satisfy d>t, the rigidity around the convex portion 14 and the convex portion 14 can be effectively reduced. Thereby, the compressive stress generated around the convex portion 14 can be effectively reduced, and the damage of the turbine vane 24A can be suppressed.

FIG. 7 is a diagram showing an example of a range in the blade height direction in which the concave portion 18 in the form shown in FIG. 6 is formed.
In the form shown in FIG. 7, the recess 18 extends linearly along the blade height direction. In the blade height direction, if the existence range of the concave portion 18 is S1, the existence range of the upstream end surface 105 of the outer fillet portion 103 is S2a, and the existence range of the upstream end surface 109 of the inner fillet portion 107 is S2b, the range S1 is There is no overlap with each of the range S2a and the range S2b. Further, the range S1 includes the central position M in the blade height direction. Here, the central position M in the blade height direction is the midpoint between the flow passage wall surface 83 of the inner shroud wall portion 62 and the flow passage wall surface 81 of the outer shroud wall portion 60.

If the recess 18 is formed from the inner shroud wall portion 62 to the outer shroud wall portion 60 in the blade height direction, the rigidity of the fillet portions 103 and 107, which tend to increase the compressive stress, decreases, and the fillet portion 103, The stress of 107 may increase. Therefore, by forming the concave portion 18 so that the range S1 does not overlap with each of the range S2a and the range S2b as described above, the influence on the rigidity of the fillet portions 103 and 107 is suppressed and the wing surface 71 is generated. Compressive stress can be reduced, and damage to the turbine vane 24A can be suppressed. Further, since the range S1 includes the center position M in the blade height direction, the compressive stress generated on the blade surface 71 is reduced while suppressing the influence of forming the recess 18 on the rigidity of the fillet portions 103 and 107. be able to.

FIG. 8 is an enlarged cross-sectional view showing a specific configuration example of the turbine vane 24A shown in FIG. The configuration shown in FIG. 8 differs from the configuration shown in FIG. 6 in the arrangement of the recesses 18.
In some embodiments, for example, as shown in FIG. 8, the outer surface 96 of the second portion 16 of the turbine vane 24A includes a recess 18. In the illustrated form, the recess 18 is formed on the outer surface 96 of the portion 86 of the second portion 16 that is not exposed to the axial flow F of the combustion gas of the turbine 6. The concave portion 18 is specifically formed on the top surface 88 of the convex portion 14.

Thus, by forming the concave portion 18 on the outer surface 96 of the second portion 16 including the convex portion 14, it is possible to reduce the rigidity of the second portion 16 including the convex portion 14. As a result, similarly to the embodiment shown in FIGS. 6 and 7, the compressive stress generated on the blade surface 71 is reduced, and the damage of the turbine vane 24A is suppressed without increasing the cooling air for cooling the blade body 70. can do.

Further, since the concave portion 18 is formed on the top surface 88 of the convex portion 14, it is possible to reduce the rigidity of the convex portion 14 in which the temperature tends to be relatively low, and to reduce the thermal expansion of the portion around the convex portion 14. The restraint by the convex portion 14 can be weakened. Thereby, the compressive stress generated around the convex portion 14 can be effectively reduced, and the damage of the turbine vane 24A can be suppressed without increasing the cooling air for cooling the blade body 70.

Further, as in the embodiment shown in FIGS. 6 and 7, the existence range of the concave portion 18 is S1, the existence range of the upstream end surface 105 of the outer fillet portion 103 is S2a, and the existence range of the upstream end surface 109 of the inner fillet portion 107. Is defined as S2b, the recess 18 may be formed so that the range S1 does not overlap with each of the range S2a and the range S2b, and the range S1 may include the central position M in the blade height direction. As a result, it is possible to reduce the compressive stress generated on the blade surface 71 while suppressing the influence of the formation of the recess 18 on the rigidity of the fillet portions 103 and 107.

FIG. 9 is an enlarged cross-sectional view showing a specific configuration example of the turbine vane 24A shown in FIG. The configuration shown in FIG. 9 differs from the configurations shown in FIGS. 6 and 8 in the arrangement of the recesses 18.
In some embodiments, for example, as shown in FIG. 9, the outer surface 96 of the second portion 16 of the turbine vane 24A includes a recess 18. In the illustrated form, the recess 18 is formed on the outer surface 96 of the portion 86 of the second portion 16 that is not exposed to the axial flow F of the combustion gas of the turbine 6. The concave portion 18 is specifically formed on each of the pair of side surfaces 90 of the convex portion 14.

Thus, by forming the concave portion 18 on the outer surface 96 of the second portion 16 including the convex portion 14, it is possible to reduce the rigidity of the second portion 16 including the convex portion 14. As a result, similarly to the embodiment shown in FIGS. 6 and 7, the compressive stress generated on the blade surface 71 is reduced, and the damage of the turbine vane 24A is suppressed without increasing the cooling air for cooling the blade body 70. can do.

Further, since the concave portion 18 is formed on each of the pair of side surfaces 90 of the convex portion 14, it is possible to reduce the rigidity of the convex portion 14 in which the temperature tends to be relatively low, and to reduce the rigidity of the portion around the convex portion 14. The restraint of the convex portion 14 on the heat expansion can be weakened. Thereby, the compressive stress generated around the convex portion 14 can be effectively reduced, and the damage of the turbine vane 24A can be suppressed without increasing the cooling air for cooling the blade body 70.

Further, as in the embodiment shown in FIGS. 6 and 7, the existence range of the concave portion 18 is S1, the existence range of the upstream end surface 105 of the outer fillet portion 103 is S2a, and the existence range of the upstream end surface 109 of the inner fillet portion 107. Is defined as S2b, the recess 18 may be formed so that the range S1 does not overlap with each of the range S2a and the range S2b, and the range S1 may include the central position M in the blade height direction. As a result, it is possible to reduce the compressive stress generated on the blade surface 71 while suppressing the influence of the formation of the recess 18 on the rigidity of the fillet portions 103 and 107.

FIG. 10 is an enlarged cross-sectional view showing a specific configuration example of the turbine vane 24A shown in FIG. FIG. 11 is a diagram showing an example of a radial range in which the recessed portion 18 in the form shown in FIG. 10 is formed. The configuration shown in FIGS. 10 and 11 is different from the configurations shown in FIGS. 6, 8 and 9 in the arrangement of the recesses 18.

In some embodiments, at least a portion of the recess 18 is formed in the inner surface 84 of the flat portion 66, eg, as shown in FIGS. As shown in FIG. 11, the concave portion 18 is formed in a part of the inner surface 84 of the flat portion 66 in the blade height direction, and therefore the flat portion 66 is partially formed in the blade height direction. It is thin. Further, as shown in FIG. 10, the concave portion 18 is formed so as to extend from the inner surface 84 of one flat portion 66 of the pair of flat portions 66 to the inner surface 84 of the other flat portion 66 through the back side of the convex portion 14. Has been done.

As described above, by forming at least a part of the concave portion 18 on the inner surface 84 of the flat portion 66 adjacent to the convex portion 14, the rigidity of the second portion 16 including the flat portion 66 can be reduced. As a result, similarly to the embodiment shown in FIGS. 6 and 7, the compressive stress generated on the blade surface 71 is reduced, and the damage of the turbine vane 24A is suppressed without increasing the cooling air for cooling the blade body 70. can do.

Further, as shown in FIG. 11, assuming that the existing range of the concave portion 18 is S1, the existing range of the upstream end surface 105 of the outer fillet portion 103 is S2a, and the existing range of the upstream end surface 109 of the inner fillet portion 107 is S2b, the range is The recess 18 may be formed so that S1 does not overlap the range S2a and the range S2b, and the range S1 may include the central position M in the blade height direction. As a result, it is possible to reduce the compressive stress generated on the blade surface 71 while suppressing the influence of the formation of the recess 18 on the rigidity of the fillet portions 103 and 107.

The present invention is not limited to the above-described embodiment, and includes a modified form of the above-described embodiment and a combination of these forms as appropriate.

For example, in some of the embodiments shown in FIGS. 6, 8, 9, and 10, the recess 18 is formed so that the range S1 and the range S2a do not overlap with each other. The range S2a may partially overlap, or the range S1 may entirely overlap the range S2a. However, since the range S1 does not overlap at least a part of the range S2a, it is possible to reduce the compressive stress generated on the blade surface 71 while suppressing the influence of the formation of the recess 18 on the rigidity of the fillet portion 103. it can.

In addition, in some of the embodiments shown in FIGS. 6, 8, 9 and 10, the recess 18 is formed so that the range S1 and the range S2b do not overlap with each other. The range S2b may partially overlap, or the range S1 may entirely overlap the range S2b. However, since the range S1 does not overlap at least a part of the range S2b, it is possible to reduce the compressive stress generated on the blade surface 71 while suppressing the influence of the formation of the recess 18 on the rigidity of the fillet portion 103. it can.

Further, in some embodiments, for example, as shown in FIG. 12, the flow channel wall surface 83 of the inner shroud wall portion 62 includes an inclined surface 85 that is inclined toward the outer side in the radial direction toward the downstream side. Good. Thereby, the velocity of the flow along the inclined surface 85 is reduced to make the pressure distribution in the circumferential direction uniform, and the radial direction from the gap between the circumferential wall portion 56b of the outlet portion 52 of the combustor 4 and the inner shroud wall portion 62 is measured. Entrainment f of the combustion gas into the inside can be suppressed.

1 Gas Turbine 2 Compressor 4 Combustor 6 Turbine 8 Rotor 10 Combustor Assembly 12 First Part 14 Projection 16 Second Part 18 Recess 19 Cooling Channel 20 Casing 22 Turbine Cabin 24, 24A, 24B Turbine Stator 26 Turbine Rotor blade 28 Combustion gas flow path 30 Exhaust casing 32 Combustor casing 38 First combustion burner 40 Second combustion burner 41 Vehicle interior inlet 42, 44 Fuel port 48 Inner cylinder 50 Tail cylinder 52 Outlet 54a, 54b Radial wall Portions 56a, 56b Circumferential wall portion 60 Outer shroud wall portion 62 Inner shroud wall portion 64 Flat surface 66 Flat portion 70 Blade body 71 Blade surface 72 Pressure surface 74 Suction surface 76 Trailing edge 78 Airfoil 80 Outer end portion 81, 83 Flow Road wall surface 82 Inner end portion 84 Inner surface 85 Sloping surface 86, 87, 92, 94 Part 88 Top surface 90 Side surface 96 Outer surface 102, 106, 132 Corner portion 103 Outer fillet portion 104, 108 Fillet radius enlarged portion 105, 109 Upstream end surface 107 Inner fillet

Claims (12)

A turbine stationary blade having a hollow blade body,
The wing body is
A first portion partially forming an airfoil including a pressure surface, a suction surface and a trailing edge;
A second portion located on the leading edge side of the airfoil with respect to the first portion and having a convex portion;
Including,
A turbine vane, wherein the outer or inner surface of the second portion includes a recess. The turbine stationary blade according to claim 1, wherein the concave portion is formed in a range including a central position in a blade height direction. A shroud wall portion which is connected to an end portion of the wing body to form a flow path wall,
A fillet portion provided at a corner formed by the wing surface of the wing body and the wall surface of the shroud wall portion,
Equipped with
The range S1 does not overlap with at least a part of the range S2, where S1 is the range in which the recess is present and S2 is the range in which the upstream end face of the fillet portion is located in the blade height direction. The turbine stationary blade according to 2. The turbine stationary blade according to any one of claims 1 to 3, wherein the recess is formed on an outer surface or an inner surface of a portion of the second portion that is not exposed to a flow of combustion gas of the turbine. The turbine stationary blade according to any one of claims 1 to 4, wherein the concave portion is formed in a portion of the inner surface of the second portion corresponding to the convex portion. The recess is formed on the inner surface of the second portion,
The range S3 is located inside the range S4, where S3 is the range in which the recess is present and S4 is the range in which the protrusion is located in the circumferential direction of the rotor. The turbine vane described. The turbine stationary blade according to claim 5 or 6, wherein d>t is satisfied, where d is a depth of the concave portion and t is a wall thickness of a portion of the second portion adjacent to the convex portion. The turbine stationary blade according to any one of claims 1 to 4, wherein the concave portion is formed on an outer surface of the convex portion. The second portion includes a flat portion in which a flat surface adjacent to the convex portion is formed on an outer surface,
The turbine vane according to any one of claims 1 to 4, wherein at least a part of the recess is formed on an inner surface of the flat portion. A plurality of combustors arranged in the circumferential direction of the rotor, including an outlet portion having a radial wall portion along the radial direction of the rotor;
The turbine stationary blade according to any one of claims 1 to 9, which is located on a downstream side of a pair of the radial wall portions facing each other of the outlet portions of the combustors adjacent to each other in the circumferential direction.
Gas turbine equipped with. The gas turbine according to claim 10, wherein the radial wall portion is fitted with the convex portion. The downstream end surface of the radial wall portion includes a protrusion receiving groove along the radial direction of the rotor, and at least a part of the protrusion receiving groove and the protrusion partially overlap in the axial direction of the rotor. The gas turbine according to claim 10 or 11, wherein

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