JP2011111946A - Blade body and gas turbine equipped with blade body - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To inhibit overcooling of pin fins and variation of temperature distribution near the trailing edge of a blade body by improving the cooling efficiency of the pin fins in the body of the blade body and thereby improve the performance and efficiency of a gas turbine. <P>SOLUTION: The blade body 10 includes: the body 20 of the blade body in which a belly side wall 21 and a back side wall 22 are connected at a leading edge 23a and the trailing edge 23b, the inside is formed in a hollow portion 24 and a ventilation hole 27 is formed in the trailing edge 23b; and inserts 30 with impingement holes 32 formed in the hollow portion 24. In the blade body 10, the pin fins 28 lying near the trailing edge 23b are cooled with cooling air A in a trailing edge cooling area 29. Between the insert 30 and the trailing edge cooling area 29, a flow regulation plate 40 formed separately from the body 20 of the blade body and formed with a flow regulation hole 42 by machining is mounted to the body 20 of the blade body. <P>COPYRIGHT: (C)2011,JPO&INPIT

Description

  The present invention relates to a blade body used in a rotating machine including a turbine including a gas turbine and a blade body, and a gas turbine including the blade body.

  As is well known, for example, a blade body used in a gas turbine (for example, a turbine stationary blade and a turbine blade) is exposed to high-temperature combustion gas, so that cooling air is guided to the base end portion of the blade body and the cooling air is used. The wing body is cooled by communicating with a cooling air passage in the wing body.

  In such a wing body, it is necessary to ensure a sufficient thickness, but in the vicinity of the trailing edge of the wing body, it is difficult to ensure a sufficient thickness because the thickness of the wing body is thin, and the trailing edge Since the vicinity of the part tends to be relatively hotter than the center side, cooling may be provided by providing a cooling pin fin in the vicinity of the rear edge of the wing body. (For example, refer to Patent Document 1).

  FIG. 7 is a diagram illustrating an example of a turbine stationary blade 10 having pin fins, in which FIG. 7A is a cross-sectional view orthogonal to the blade height direction, and FIG. 7B is an arrow in FIG. The outline of the film cooling hole of the turbine stationary blade surface shown by the view S and the impingement hole (broken line) of the insert surface is shown. The blade height direction refers to the direction in which the turbine stationary blade extends, and substantially coincides with the direction in which the blade body standing in the turbine casing faces the rotation axis of the rotor.

The turbine stationary blade 10 includes a blade body 20 and an insert 30, and the blade body 20 has a cross-section orthogonal to the blade height direction, which has a concave outer surface as shown in FIG. The abdominal side wall portion 21 and the back side wall portion 22 having a convex outer shape are connected by a front edge portion 23a and a rear edge portion 23b, and a wing shape is formed with a hollow portion 24 formed therein.
Further, a plurality of pin fins 28 are formed in the vicinity of the rear edge portion 23b of the wing body 20 to connect the abdominal side wall portion 21 and the back side wall portion 22. The pin fins 28 are provided with cooling air A from the abdominal side wall portion. It cools by passing the rear edge part cooling area | region 29 formed between 21 and the back side wall part 22. As shown in FIG.

  The insert 30 is disposed, for example, in each of three cavities C1, C2, and C3 in which the hollow portion 24 is partitioned by the two partition plates 25, and the cooling air A guided to the flow path 31 inside the insert 30 is It is injected from the impingement hole 32 of the insert 30 onto the inner surfaces of the abdominal wall 21 and the back wall 22, and then flows out from the film cooling hole 26 of the turbine vane 10 and the air hole 27 of the rear edge 23 b of the turbine vane 10. When the cooling air A flowing out from the vent hole 27 passes through the trailing edge cooling region 29, the pin fins 28 are cooled, and consequently the vicinity of the trailing edge 23b of the wing body 20 is cooled.

JP 2000-356104 A

However, when casting the blade body of a blade body such as a turbine stationary blade, the hollow portion is formed using the core, so that the thickness of the core is increased in order to prevent damage to the core. Then, the cooling air passage in the trailing edge cooling region becomes large, the pin fins are cooled by a large amount of cooling air, and the pin fins are overcooled.
Due to the overcooling of the pin fins, the temperature distribution of the blade body becomes unbalanced, resulting in thermal stress and thermal deformation, and the flow rate of the cooling air becomes excessive, resulting in a decrease in performance and thermal efficiency of the entire gas turbine.
Further, when the number of impingement holes and the opening area are reduced in order to reduce the cooling air flow rate, there is a problem that a local hot spot is generated.
Therefore, there is a strong technical demand to efficiently cool the pin fins in order to suppress overcooling of the pin fins and improve the performance and thermal efficiency of the gas turbine.

  The present invention has been made in consideration of such circumstances, and by optimizing the cooling efficiency of the pin fins in the wing body, the pin fin is supercooled and the temperature distribution in the vicinity of the trailing edge of the wing body varies. The purpose is to improve the performance and efficiency of the gas turbine.

In order to solve the above problems, the present invention proposes the following means.
According to the first aspect of the present invention, the abdominal side wall portion that is extended in the blade height direction and has a concave outer surface in the cross section orthogonal to the blade height direction and the back side wall portion that is convex in the outer surface are provided. A wing body that is connected at the front edge portion and the rear edge portion and has a hollow portion formed therein, and has a wing body body in which a vent hole communicating with the outside is formed in the vicinity of the rear edge portion; and The cooling wall is disposed between the inner surface of the abdominal sidewall portion and the inner surface of the back sidewall portion, and is directed from the internal flow path to at least one of the inner surface of the abdominal sidewall portion and the inner surface of the back sidewall portion. And a pin fin connecting the abdominal side wall portion and the back side wall portion in the vicinity of the rear edge portion of the hollow portion, in the vicinity of the rear edge portion. In the rear edge cooling region formed in the impingement A blade body that is cooled by cooling air injected from a hole, and adjusts a flow rate of cooling air injected from the impingement hole between the insert and the trailing edge cooling region. A plate is provided, and the flow rate adjusting plate is formed separately from the wing body main body, and a flow rate adjusting hole for adjusting the flow rate of the cooling air is formed by machining, and is attached to the wing body main body. It is characterized by.

  The invention described in claim 4 is a gas turbine, and includes the wing body described in any one of claims 1 to 3.

According to the wing body and the gas turbine according to the present invention, the flow rate adjusting plate for adjusting the flow rate of the cooling air injected from the impingement hole is formed separately from the wing body main body. A flow rate adjusting hole having a desired position, size and shape can be machined before being attached to the main body. By adjusting the flow rate of the cooling air flowing through the rear edge cooling region through the appropriately formed flow rate adjustment hole, the flow rate and the flow rate of the cooling air are appropriately adjusted, and the efficiency is reduced by a small amount of cooling air. Thus, the pin fin can be cooled.
As a result, over-cooling of the pin fins is suppressed to reduce variations in the temperature distribution near the rear edge of the wing body, and excessive cooling air flow is suppressed to improve the performance and thermal efficiency of the gas turbine. can do.

  Invention of Claim 2 is the wing | blade body of Claim 1, Comprising: The said blade height direction position of the said flow volume adjustment hole is formed so that it may overlap with the said blade height direction position of the said pin fin. It is characterized by being.

  According to the wing body of the present invention, since the blade height direction position of the flow rate adjusting hole is formed so as to overlap the pin fin, the jet of cooling air flowing from the flow rate adjusting hole into the trailing edge cooling region is The heat exchange between the cooling air and the pin fins can be performed efficiently. As a result, the cooling efficiency of the pin fins can be improved, and the performance and efficiency of the entire gas turbine can be further improved by reducing the flow rate of the cooling air.

  The invention according to claim 3 is the wing body according to claim 2, wherein the blade height direction position of the flow rate adjusting hole is formed at the same position as the blade fin direction position of the pin fin. It is characterized by being.

  According to the blade body according to the present invention, the blade height direction position of the flow rate adjusting hole is the same position as the blade height direction position of the pin fin, so the jet of cooling air flowing into the trailing edge cooling region is By colliding with the center of the pin fin, it is possible to further improve the heat exchange efficiency between the cooling air and the pin fin. As a result, the flow rate of the cooling air can be reduced to further improve the performance and efficiency of the entire gas turbine.

  According to the blade body and the gas turbine according to the present invention, it is possible to suppress the overcooling of the pin fins and the temperature distribution variation in the vicinity of the trailing edge portion of the blade body resulting from this, and to suppress the excessive flow rate of the cooling air. Thus, the performance and thermal efficiency of the entire gas turbine can be improved.

It is a longitudinal section showing a schematic structure of a gas turbine concerning one embodiment of the present invention. It is a principal part expanded sectional view of the gas turbine which concerns on one Embodiment. It is a principal part expanded sectional view of the turbine stationary blade which concerns on one Embodiment, and is a figure which shows the II cross section in FIG. It is a figure which shows an example of the flow volume adjusting plate which concerns on one Embodiment, and is the XX sectional view taken on the line in FIG. (A) It is a figure which shows an example of arrangement | positioning of the flow regulating plate and pin fin which concern on one Embodiment, and is the YY sectional view of FIG. (B), (C), (D) is a figure which shows the modification of (A). It is a figure which shows the modification of the flow volume adjustment board which concerns on this invention. It is a figure which shows an example of the conventional blade body, (A) is sectional drawing orthogonal to a blade height direction, (B) is a film cooling hole and impingement hole at the time of seeing SS in (A) FIG.

Hereinafter, an embodiment of the present invention will be described with reference to FIGS.
FIG. 1 is a longitudinal sectional view showing a schematic configuration of a gas turbine 1 according to an embodiment of the present invention.
As shown in FIG. 1, the gas turbine 1 includes a compressor 2 that generates compressed air c, and a plurality of fuels that supply the compressed air c supplied from the compressor 2 to generate combustion gas G1 that is a working fluid. And a turbine 4 having four stages of a set of turbine stationary blades (blade bodies) 10 and turbine rotor blades 6 and generating rotational power by the combustion gas G1 supplied from the combustor 3. I have.

  Further, 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 is a bearing provided at one end on the upstream side of the compressor 2. The other end 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 7b provided on the downstream side of the turbine 4. Has been. Hereinafter, in the axial direction D of the rotor 7, the compressor 2 side is a front side, and the turbine 4 side is a rear side.

  The compressor 2 includes a compressor casing 2b in which an air intake port 2a for taking in air is disposed on the front side, a plurality of compressor stationary blades 2c and a plurality of compressor rotor blades 2d disposed in the compressor casing 2b. It has. The compressor stationary blades 2 c are fixed to the inner peripheral surface of the compressor casing 2 b and extended toward the rotor 7, and are arranged at equal intervals in the circumferential direction R of the turbine 4. The compressor blades 2d are fixed to the outer peripheral surface of the rotor 7 and extended toward the inner peripheral surface of the compressor casing 2b, and are arranged at equal intervals in the circumferential direction R of the turbine 4. . The compressor stationary blades 2c and the compressor moving blades 2d are arranged in multiple stages so as to alternate along the axial direction D.

  The combustor 3 includes an inner cylinder 3a having a burner (not shown) therein, an outer cylinder 3b that guides compressed air c supplied from the compressor 2 to the inner cylinder 3a, and a fuel injection (not shown) that supplies fuel to the inner cylinder 3a. And a tail cylinder 3c that guides the combustion gas G1 from the inner cylinder 3a to the turbine 4. The plurality of combustors 3 are disposed in the circumferential direction R of the turbine 4 and are disposed in a combustor casing 3d having a front end portion connected to a rear end portion of the compressor casing 2b.

The turbine 4 includes a turbine casing 5 whose front end is connected to the rear end of the combustor casing 3d, turbine stationary blades 10 and turbine rotor blades 6 that are alternately arranged in four stages in the axial direction in the turbine casing 5. And.
The turbine stationary blades 10 at each stage are arranged in the circumferential direction R at an equal interval, are fixed to the turbine casing 5 side, and extend radially toward the rotor 7 side. Similarly, the turbine rotor blades 6 at each stage are also arranged in the circumferential direction R at equal intervals, are fixed to the rotor 7 side, and extend radially toward the turbine casing 5 side.
An exhaust chamber 8 that opens toward the rear side is connected to a rear end portion of the turbine casing 5 of the turbine 4. The exhaust chamber 8 is provided with an exhaust diffuser 8a that converts the dynamic pressure of the combustion gas G1 that has passed through the turbine stationary blade 10 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 2a of the compressor 2 passes through the compressor stationary blades 2c and the compressor rotor blades 2d arranged in multiple stages and is compressed. Air c is generated. Next, in the combustor 3, the combustion gas G <b> 1 is generated by supplying the fuel to the compressed air c and burning it as described above, and the combustion gas G <b> 1 is guided to the turbine 4. The combustion gas G1 passes through the range where the turbine stationary blades 10 and the turbine rotor blades 6 are arranged as a combustion gas flow path, so that the rotor 7 is rotationally driven. The exhaust gas after rotationally driving the rotor 7 is converted into a static pressure by the exhaust diffuser 8a in the exhaust chamber 8, and then released to the atmosphere.

FIG. 2 is an enlarged cross-sectional view of a main part of the gas turbine 1.
As shown in FIG. 2, the rotor 7 includes rotor disks 7 </ b> A to 7 </ b> D that fix first to fourth stage turbine blades 6 </ b> A to 6 </ b> D on the outer periphery thereof. A seal disk 11 is coaxially connected to the upstream side of the first stage rotor disk 7A. In this seal disk 11, disk holes 11a penetrated to supply a part of the compressed air c from the upstream side toward the turbine rotor blades 6A to 6D are equiangularly spaced from each other around the axis. A plurality of them are formed.

  Further, the radial holes 7A1 that take in a part of the compressed air c that has flowed through the respective disk holes 11a and guide them into the interior of each turbine blade 6A are equiangular with the first stage rotor disk 7A. A plurality are formed at intervals. Furthermore, a plurality of disk holes 7A2 for supplying the remaining compressed air c toward the second stage are formed in the rotor disk 7A at equal angular intervals.

  Similarly to the rotor disk 7A, a plurality of radial holes 7B1 and 7C1 and a plurality of disk holes 7B2 and 7C2 are formed in each of the rotor disks 7B and 7C. A plurality of radial holes 7D1 are formed in the rotor disk 7D.

The turbine casing 5 is formed with compressed air passages 12 corresponding to the respective turbine stationary blades 10, and a part of the compressed air c compressed by the compressor 2 via the compressed air passages (not shown). Is supplied to the flow path 31 inside the insert 30 of the turbine vane 10 of each stage as the compressed air c as cooling air A.
2 indicates the radial height of the rotor when the turbine stationary blade 10 is erected on the inner periphery of the turbine casing 5 toward the inner periphery in the radial direction. It is called the direction.

FIG. 3 is an enlarged cross-sectional view of a main part of the turbine stationary blade 10 and is a view showing a cross section taken along the line II in FIG.
As shown in FIG. 3, the turbine stationary blade 10 includes a blade body 20, an insert 30, and a flow rate adjustment plate 40. The blade body 20, the insert 30, and the flow rate adjustment plate 40 are respectively in the blade height direction. It extends to.

The wing body 20 is formed by, for example, casting, and in a cross section orthogonal to the wing height direction, an abdominal side wall portion 21 having a concave shape on the outer surface and a back side wall portion 22 having a convex shape on the outer surface, The front edge portion 23a and the rear edge portion 23b are connected to each other, and the airfoil has a hollow portion 24 formed therein.
In this embodiment, the surface of the abdominal side wall 21 and the back side wall 22 is spaced from the hollow portion 24 to the outside at intervals in the blade height direction and in the direction from the front edge 23a toward the rear edge 23b. A plurality of film cooling holes 26 communicating with each other are formed.
Further, a plurality of vent holes 27 communicating from the hollow portion 24 to the outside are formed in the trailing edge portion 23b in the blade height direction.

  The hollow portion 24 is formed in the wing body main body 20 so as to extend in the wing height direction, and is formed by three partition plates 25 that connect the abdominal side wall portion 21 and the back side wall portion 22 to form three cavities C1, C2, It is partitioned into C3.

  The insert 30 is formed so as to extend in the blade height direction, the cross section orthogonal to the blade height direction has a rectangular shape or a substantially triangular shape, and a flow path 31 through which the cooling air A flows is formed. In the outer surface, a plurality of impingement holes 32 communicating from the flow path 31 to the outside of the insert 30 are formed at intervals in the blade height direction and the direction from the front edge portion 23a toward the rear edge portion 23b. Has been.

  Further, one insert 30 is disposed in each of the three cavities C1, C2, and C3 formed in the hollow portion 24, and each insert 30 extends from the impingement hole 32 to the abdominal side wall portion 21 and the back side wall portion 22. Injected toward the inner surface to cool the abdominal side wall portion 21 and the back side wall portion 22, and from the air cooling hole 26 of the film cooling hole 26 and the rear edge portion 23b to the outside of the blade body 20, that is, the main flow channel of the turbine It has come to leak.

  The cooling air A flowing out from the air holes 27 cools the pin fins 28 when passing through the trailing edge cooling region 29, and as a result, the vicinity of the trailing edge 23b of the wing body 20 is cooled. Yes.

The pin fins 28 are formed in a cylindrical shape. For example, as shown in FIG. 5A, the pin fins 28 are spaced apart from each other in the blade height direction and in the direction from the front edge portion 23a toward the rear edge portion 23b. When viewed from the direction, the abdominal side wall 21 and the back side wall 22 are connected to each other in the wing body 20 in the vicinity of the rear edge 23b so as to form a staggered arrangement, and pass through the rear edge cooling region 29. The cooling air A is cooled.
In this embodiment, the rear edge cooling region 29 is formed between the inner surface of the abdominal side wall portion 21 and the inner surface of the back side wall portion 22, and the position in the direction from the front edge portion 23a toward the rear edge portion 23b is It is assumed that the range in which the pin fins 28 are present is shown.

The flow rate adjusting plate 40 is formed separately from the blade body 20 and is disposed between the insert 30 disposed in the cavity C3 closest to the trailing edge 23b and the trailing edge cooling region 29 to cool the trailing edge. The flow rate of the cooling air A flowing in the region 29 is adjusted by the flow rate adjusting hole 42.
As shown in FIG. 4, the flow rate adjusting plate 40 has a configuration in which a plurality of circular flow rate adjusting holes 42 are arranged in a plate material 41, and the flow rate adjusting holes 42 are formed by machining. Thus, by forming the flow rate adjusting hole 42 by machining, the flow rate adjusting hole 42 can be easily and efficiently set appropriately as compared with the case where it is formed by casting. In addition, since the accuracy of the flow rate adjusting hole 42 as an orifice for adjusting the flow rate can be increased, variations in cooling efficiency and temperature distribution due to flow rate variations can be suppressed.

Further, the flow rate adjusting plate 40 is inserted into the wing body 20 by, for example, being inserted into an attachment groove G formed by electric discharge machining and welded to the wing body 20.
In this embodiment, each flow rate adjustment hole 42 of the flow rate adjustment plate 40 has a pin fin center position in the blade height direction of the first row closest to the flow rate adjustment hole 42 as shown in FIG. The flow control hole 42 is machined to have the same height as the center position 28, and the diameter of the flow rate adjusting hole 42 is the same as the diameter of the pin fin 28. The machining includes wire cutting and electric discharge machining in addition to cutting.
Further, the attachment groove G of the wing body 20 may be formed by machining such as cutting, wire cutting, and electric discharge machining, or may be formed by casting, The flow rate adjustment plate 40 may be attached by brazing or the like in addition to welding.

  Further, the blade height direction position of the flow rate adjusting hole 42 is arranged so as to collide with not only the first row pin fins 28 but also the second row pin fins 28 as shown in FIG. 5B, for example. Alternatively, the arrangement of the pin fins 28 may be changed from the staggered arrangement so as to collide with the pin fins 28 in the third and subsequent rows, for example.

  Further, for example, as shown in FIGS. 5A and 5B, the position of the flow rate adjustment hole 42 in the blade height direction is not limited to coincide with the center of any pin fin 28, and FIG. , (D), the blade height of the flow rate adjusting hole 42 is set such that at least a part of the cooling air A collides with a part of the pin fin 28 so that a part of the pin fin 28 overlaps the flow rate adjusting hole 42. The position in the vertical direction may be set appropriately shifted.

Next, operation | movement of the gas turbine 1 is demonstrated using figures.
First, as shown in FIG. 2, a part of the compressed air c extracted and supplied from the compressor 2 is supplied toward the rotating seal disk 11, passes through each disk hole 11a, and then the first stage. To the rotor disk 7A, and sequentially reach the radial holes 7B1 to 7D1 of the rotor disks 7B to 7D through the disk holes 7A2 to 7C2. Next, the compressed air c that has reached the radial holes 7A1 to 7D1 flows into the cooling passages in the turbine blades 6A to 6D from the bottom surfaces of the blades, flows from the base ends to the tips, and cools the turbine blades 6.

On the other hand, a part of the compressed air c is sent to the compressed air flow path 12 via a compressed air flow path (not shown) and supplied as cooling air A toward the insert 30 of the turbine stationary blade 10 of each stage. .
The compressed air c supplied to the insert 30 of the turbine vane 10 flows as the cooling air A through the flow path 31, and from the impingement hole 32, the abdominal side wall 21 and the back side wall in each of the cavities C 1, C 2, C 3. 22 is injected toward the inner surface of the blade body 22 to impinge cool the inner surface of the wing body 20.

Next, the cooling air A flows out from the film cooling holes 26 of the blade body 20 to form a film layer of the cooling air A around the blade body 20, and the turbine stationary blade 10 is film-cooled.
A part of the cooling air A injected from the impingement hole 32 flows into the rear edge cooling region 29 through the flow rate adjusting plate 40 to cool the pin fin 28 and is discharged from the vent hole 27 in the rear edge 23b. The Further, the pin fin 28 is cooled, so that the vicinity of the rear edge 23b of the wing body 20 is cooled.
At this time, the cooling air A passing through the flow rate adjusting plate 40 is adjusted to a flow rate and a flow rate so that the cooling air A flows while appropriately cooling without overcooling the pin fins 28 by the appropriately set flow rate adjusting holes 42.

  According to the turbine stationary blade 10 and the gas turbine 1, the flow rate is adjusted by passing the flow rate adjusting plate 42 through the appropriately set flow rate adjusting hole 42 in which the hole diameter and the position are formed with high accuracy by machining. It is easy to efficiently cool the pin fins 28 with a small amount of cooling air A by appropriately setting the flow rate, flow velocity, and basin of the air A. As a result, the cooling air A and the pin fins 28 efficiently exchange heat. Therefore, it is possible to improve the performance and efficiency of the gas turbine 1 by suppressing the flow rate of the cooling air A from becoming excessive.

  Further, overcooling of the pin fins 28 can be suppressed, and variations in the temperature distribution in the vicinity of the rear edge portion 23b of the blade body 20 can be reduced.

  Further, according to the turbine stationary blade 10, the blade height direction position of the flow rate adjustment hole 42 is the same position as the blade height direction position of the pin fin 28, so that the jet of the cooling air A flows from the flow rate adjustment hole 42. By impingement cooling by colliding with the pin fins 28, heat exchange between the cooling air A and the pin fins 28 can be performed efficiently.

  Further, since the flow rate adjusting plate 40 constituting the turbine stationary blade 10 is formed separately from the blade body 20, the flow rate adjusting plate 40 has a desired position before the flow rate adjusting plate 40 is attached to the blade body. The flow rate adjusting hole having a size and a shape can be processed easily and efficiently.

In addition, this invention is not limited to the said embodiment, A various change is possible in the range which does not deviate from the meaning of invention.
For example, in the above-described embodiment, the film cooling holes 26 are formed in both the abdominal side wall 21 and the back side wall 22, and the impingement is provided on the surfaces on the side of the abdominal side wall 21 and the back side wall 22 of each insert 30. In the above description, the film cooling hole 26 is provided on either the abdominal side wall portion 21 or the back side wall portion 22, and on which surface of the insert 30 the impingement hole 32 is formed. It can be set arbitrarily.

  In the above embodiment, the case where the present invention is applied to the turbine stationary blade 10 has been described. However, the present invention may be applied to the turbine rotor blade 6.

Further, the number of inserts 30 arranged in the wing body 20, the number of partition plates 25 for partitioning the hollow portion 24, and the arrangement (pitch, number) of the film cooling holes 26 and impingement holes 32 are arbitrarily set. can do.
Moreover, in the said embodiment, although the case where the turbine stationary blade 10 was used for a gas turbine was demonstrated, it is also possible to apply to turbines other than a gas turbine, and a jet engine.

  Further, in the above embodiment, the case where the flow rate adjusting hole 42 formed in the flow rate adjusting plate 40 is circular has been described. However, instead of the flow rate adjusting plate 40, as shown in FIG. It is also possible to adopt a configuration in which a plurality of flow rate adjusting holes 47 formed in a straight line are arranged, and in addition to a triangular shape, a flow rate adjusting hole having a polygon other than a square or other shapes is provided. Alternatively, a plurality of shapes of flow rate adjusting holes may be set in one flow rate adjusting plate. Moreover, it is good also as a structure which arrange | positions several flow volume adjusting plates between the insert 30 and the rear edge part cooling area | region 29. FIG.

In addition, the size, shape, number, position, pitch, and the like of the flow rate adjusting holes 42 and 47 can be optimally adjusted according to the temperature distribution in the vicinity of the trailing edge 23b.
Further, although the case where the diameter of the flow rate adjusting hole 42 is the same as the diameter of the pin fin 28 has been described, the present invention is not limited to this and may not be the same diameter.
Further, although the circular pin fin 28 has been described, the present invention is not limited to this, and a shape other than a circle may be used, and a cooling member called an pedestal such as an oval shape may be included in the pin fin 28. May be.

  Further, in the above embodiment, the case where the position of the flow rate adjustment hole 42 in the blade height direction is the same as or overlaps with the pin fin 28 has been described, but some or all of the flow rate adjustment holes in the blade height direction. The position may not be the same position in the blade height direction of the pin fin 28, and there is no overlap, that is, the pin fin 28 may be arranged at a position hidden from the flow rate adjusting hole.

  In the above embodiment, the case where the compressed air c is used as the cooling air A has been described. However, a mechanism for cooling the compressed air c is provided, and the compressed air c is cooled before being supplied to the turbine stationary blade 10. Alternatively, the cooling air may be supplied from the outside instead of the compressor 2.

  According to the wing body and the gas turbine according to the present invention, the supercooling of the pin fins and the variation in the temperature distribution in the vicinity of the rear edge of the wing body can be suppressed, and the performance and efficiency of the gas turbine can be improved. Is possible.

A Cooling air c Compressed air H Blade height 1 Gas turbine 6 Turbine blade 10 Turbine stationary blade (blade)
20 Wing body body 21 Abdominal side wall part 22 Back side wall part 23a Front edge part 23b Rear edge part 24 Hollow part 24a, 24b Cooling space 26 Film cooling hole 27 Ventilation hole 28 Pin fin 29 Rear edge cooling area 30 Insert 31 Channel 32 Impinge Ment hole 40, 45 Flow rate adjusting plate 42, 47 Flow rate adjusting hole

Claims (4)

Extending in the blade height direction and connecting the abdominal side wall portion having a concave outer surface and the back side wall portion having a convex outer surface in a cross section orthogonal to the blade height direction at the front edge portion and the rear edge portion. And a wing body having a hollow portion formed therein, a wing body main body formed with a vent hole communicating with the outside in the vicinity of the rear edge portion, the hollow portion, the inner surface of the abdominal side wall portion and the A plurality of impingements arranged with a cooling space between the inner surface of the back side wall portion and injecting cooling air from an internal flow path toward at least one of the inner surface of the abdominal side wall portion and the inner surface of the back side wall portion And a pin fin connecting the abdominal side wall and the back side wall in the vicinity of the rear edge of the hollow portion, and a rear edge cooling region formed in the vicinity of the rear edge. Cooling air sprayed from the impingement hole at And a flow rate adjusting plate that adjusts the flow rate of the cooling air injected from the impingement hole between the insert and the trailing edge cooling region. The adjusting plate is formed separately from the wing body main body, and a flow rate adjusting hole for adjusting the flow rate of the cooling air is formed by machining, and is attached to the wing body main body. The wing body according to claim 1,
The blade body characterized in that the blade height direction position of the flow rate adjusting hole is formed so as to overlap the blade height direction position of the pin fin. The wing body according to claim 2,
The blade height direction position of the flow rate adjusting hole is formed at the same position as the blade height direction position of the pin fin. A gas turbine comprising the wing body according to any one of claims 1 to 3.

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