JP2000345802A - Gas turbine blade manufacturing method - Google Patents

Abstract

(57) [Summary] [Problem] To provide a method for manufacturing a gas turbine blade that can be easily manufactured even when the blade height is increased. A method for manufacturing a gas turbine blade according to the present invention includes a blade effective portion root portion side cooling passage formed on a blade effective portion root portion and a solid formed on a blade effective portion tip portion. The shape is produced by precision casting. Next, the blade effective portion tip side cooling passage 28 is formed by machining into a solid shape formed on the blade effective portion tip portion 23 side.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a method for manufacturing a gas turbine blade which can be applied to a subsequent stage of a turbine stage.

[0002]

2. Description of the Related Art Generally, as shown in FIG. 8, a gas turbine plant applied to a power plant or the like has a high pressure air compressed by driving an air compressor 2 provided coaxially with a gas turbine 1 as a gas turbine. The fuel is supplied to the combustor 3 together with the fuel, and generates combustion gas in a combustion chamber 5 defined by a combustor liner 4. The combustion gas is transferred via a transition piece 6 to a turbine stationary blade 7 or a turbine blade of the gas turbine 1. The turbine blade 8 is rotated by the combustion gas that has been guided by the turbine vanes 8 and expanded by the turbine vanes 7, and at that time, a rotational torque is generated on the turbine shaft 9.

In order to improve the thermal efficiency of a gas turbine plant, it is desirable to raise the temperature of the gas turbine inlet, as is well known, and in practice, the gas turbine inlet temperature is raised. With the rise of the gas turbine inlet temperature, the necessity of using materials that can withstand high temperatures has also increased for the gas turbine combustor 3, the turbine stationary blades 7 and the turbine blades 8, and heat-resistant superalloy materials have been used as gas turbine components. Has come to be used.

However, while the temperature limit of the heat-resistant superalloy material currently applied to the gas turbine 1 as a high temperature member is 800 to 900 ° C., the gas turbine inlet temperature is 1300 ° C. or more, The temperature has been significantly exceeded. Therefore, in order to guarantee the operation of the gas turbine plant, a gas turbine blade having a structure capable of cooling to the above-mentioned limit temperature has been required.

Recently, for example, as shown in FIG. 9, a turbine blade 8 is being applied to a gas turbine 1 of a class having a gas turbine inlet temperature of 1300 ° C. or higher. That is,
In FIG. 9, cooling air mainly extracted from the air compressor is supplied to the shank portion 10 a of the turbine rotor blade 8 and the blade implant portion 1.
0b flows continuously from the blade root portion (blade root portion) 10 integrally formed into the passages 11a and 11b in the blade effective portion 10c to perform convection cooling and to reverse the meandering passages 12a and 12a.
After the return flow is repeated at 12b, exits 13a,
13b is discharged outside the wing.

The leading edge 14 of the blade is cooled by cooling air blown from a shower head cooling air blow hole 15 formed in the passage 11a. Further, the blade surface is cooled by cooling air blown out from a film cooling air blowing hole (not shown). Further, the cooling air having passed through the passage 11 b is cooled from the blade trailing edge 17 by the pin fins 16, and is then blown out from the cooling air blowout holes 18 provided in the blade trailing edge 17.

[0007]

However, the conventional gas turbine blade shown in FIG. 9 has the following problems.

In recent gas turbine plants, the rise of the gas turbine inlet temperature and the adoption of a reburner also require cooling of the gas turbine blades installed after the third stage of the turbine, which did not require cooling. The gas turbine blades installed after the third stage of the turbine are higher in blade height than the gas turbine blades installed in the first stage of the turbine and the second stage of the turbine. 60cm or more.

As described above, when the blade height becomes large and the cooling passages 11a and 11b become reversing meandering passages, the structure is too complicated with the current manufacturing technology, and there is a limit to manufacturing by precision casting. And it has become difficult to implement all in-foil cooling structures by machining.

Further, during operation, a large centrifugal force acts on the blade root portion (blade root portion) as compared with the tip portion (blade tip portion) of the effective blade portion 10c. However, as the blade height increases, the centrifugal force also increases. Increase. For this reason, if the blade height of the gas turbine blade after the third stage of the turbine is large, the blade root portion has few hollow portions such as passages (cooling holes) 11 and many blade base material portions in terms of centrifugal force strength. A structure is required.

On the other hand, the central portion (PCD) of the wing effective portion 10c
Since the temperature of the gas turbine driving gas (mainstream) is higher than that of the root portion and the tip portion, a structure having a high heat transfer coefficient is required.

Further, comparing the cross-sectional shapes of the root portion side, the center portion, and the tip portion side of the effective blade portion 10c, the root portion side is thicker and the blade tip portion side is thinner, and the entire blade has a twisted shape. Has become. In such a case, the insertion depth of the machining jig is limited not only by the length of the jig but also by the point of the minimum thickness.

As described above, in the conventional gas turbine blade,
Manufacturing problems have emerged as the blade height has increased, and new improvements have been required in the manufacture of gas turbine blades.

The present invention has been made in view of such circumstances, and an object of the present invention is to provide a method of manufacturing a gas turbine blade that can be easily manufactured at low cost even if the blade height is increased.

[0015]

According to a first aspect of the present invention, there is provided a method of manufacturing a gas turbine blade, comprising: a first cooling passage formed on a blade effective portion root side of a blade effective portion;
After the solid shape formed on the tip of the effective blade portion of the wing effective portion is manufactured by precision casting, the first cooling passage is machined into a solid shape formed on the tip of the effective wing portion. And a second cooling passage communicating with the second cooling passage.

According to a second aspect of the present invention, there is provided a method of manufacturing a gas turbine blade, comprising: a first cooling passage formed on a blade effective portion root portion side of the blade effective portion; After the pin fin formed across the cooling passage and the solid shape formed on the wing effective portion tip side of the wing effective portion are manufactured by precision casting, the solid formed on the wing effective portion tip portion side The first through the pin fins by machining into shape
The second cooling passage is formed so as to communicate with the cooling passage.

According to a third aspect of the present invention, there is provided a method for manufacturing a gas turbine blade, comprising: a first cooling passage formed on the blade effective portion side of the blade effective portion; A pin fin formed across the cooling passage, a solid shape connected to the pin fin and formed on the wing trailing edge of the wing effective portion, and a solid shape formed on the wing effective portion tip side of the wing effective portion After being manufactured by precision casting, a second cooling passage communicating with the first cooling passage via the pin fins is formed by machining into a solid shape formed on the wing effective portion tip portion side. And forming a trailing-edge-side outlet to be communicated with the pin fin by machining into a solid shape formed at the trailing edge of the blade.

According to a fourth aspect of the present invention, there is provided a method for manufacturing a gas turbine blade, wherein the first cooling passage formed on the blade effective portion side of the blade effective portion, After the inverted meandering passage provided with the heat transfer promoting body traversing the cooling passage and the solid shape formed on the blade effective portion tip side of the effective blade portion are manufactured by precision casting, the effective blade tip is formed. A second cooling passage is formed in the solid shape formed on the part side by machining and communicates with the first cooling passage via the reverse meandering passage.

According to a fifth aspect of the present invention, there is provided a method of manufacturing a gas turbine blade, comprising: a first cooling passage formed on a blade effective portion root side of the blade effective portion; At least one of the second cooling passages formed on the wing effective portion tip portion side is provided with a film cooling blowout hole for blowing a cooling medium out of the wing on a wing wall of the wing effective portion,
The film-cooling blow-out hole communicating with at least one of the first cooling passage and the second cooling passage is formed by machining.

According to a sixth aspect of the present invention, there is provided a method for manufacturing a gas turbine blade, wherein the first cooling passage formed on the blade effective portion side of the blade effective portion includes a first cooling passage formed on a blade root portion. The passage for cooling the blade root portion to be formed is communicated with the first
The cooling passage for the blade root portion communicated with the cooling passage is formed by machining.

[0021]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS An embodiment of a method for manufacturing a gas turbine blade according to the present invention will be described below with reference to the drawings and reference numerals attached to the drawings.

FIG. 1 is a diagram illustrating a method for manufacturing a gas turbine blade according to the present invention.

The gas turbine blade shown in FIG.
A wing root portion (wing root portion) 22 in which a shank portion 20 and a wing implant portion 21 are formed continuously and integrally with a wing effective portion tip portion (wing tip portion) 23 formed on the wing top side, and a wing formed integrally and continuously. An effective unit 24 is provided.

The blade effective portion 24 communicates with a blade root portion cooling passage 25 formed in the blade root portion 22, and is formed in the blade effective portion root portion 24a.
A pin fin 27 that collides the cooling medium CF to increase the heat transfer coefficient, a cooling portion 27c for the central portion of the blade effective portion where the pin fin 27 is disposed, and a blade effective portion communicating with the cooling portion 27c for the central portion of the blade effective portion. And a tip-side cooling passage 28.

In manufacturing a gas turbine blade having such a configuration, the following steps are performed.

First, at the time of casting, the blade root portion cooling passage 25 having a relatively large diameter is formed in the blade root portion 22 by precision casting. In addition, the effective blade portion 24 has a relatively large-diameter effective blade route that communicates with the above-mentioned blade root portion cooling passage 25 in a portion of the region A (from the effective blade portion root portion 24a side to the effective blade center side portion). The part-side cooling passage 26 and the pin fins 27 formed in the blade effective part center part cooling passage 27 are manufactured by precision casting, and the area B (from the blade effective part center part to the blade effective part tip side part 23). Side) is made solid.

In this case, the blade effective portion root side cooling passage 26 formed in the region A of the blade effective portion 24 is formed as shown in FIG.
As shown in the figure, the wing is formed between the leading edge 29 and the trailing edge 30.

As shown in FIG. 3, the pin fins 27 formed in the area A of the wing effective portion 24
And the back side 32.

Next, when the casting process is completed, the blade effective portion 24
A portion of the region B formed as a solid is machined, and as shown in FIG. 4, a blade effective portion tip side cooling passage 28 having a relatively small diameter is formed in the arrow L direction shown in FIG.
Also, fill cooling blowout holes 36 formed on each of the ventral side 31 and the backside 32 shown in FIGS. 2, 3 and 4 are also formed in the blade wall by machining.

As described above, in the present embodiment, at the time of casting, the blade root portion cooling passage 25 having a relatively large diameter in the blade root portion 22 and the blade effective portion having a relatively large diameter are formed in the region A of the blade effective portion 24. Root part side cooling passage 26 and pin fin 2
7 is manufactured by precision casting, the region B of the wing effective portion 24 is made of a solid, and after the casting, the region B formed by the solid is machined into a wing effective portion having a relatively small diameter. Since the tip side cooling passage 28 is formed, a cooling structure can be easily manufactured even with a gas turbine blade having a large blade height installed in the latter stage of the turbine, and a cooling structure capable of sufficiently coping with high temperature. A gas turbine blade can be realized.

In the present embodiment, the film cooling blow holes 3 formed on the ventral side 31 and the back side 32, respectively.
Since 6 is drilled in the wing wall by machining, the drilling operation can be easily performed in a short time.

In this embodiment, the pin fins 27 are provided at the center of the effective blade portion 24 of the effective blade portion 24. For example, as shown in FIG. The air passage 26 may be formed, and the blade effective portion tip side cooling passage 28 may be formed in the region B. in this case,
The blade effective section root side cooling passage 26 is formed by precision casting, and the blade effective section tip section side cooling passage 82 is formed by machining.

FIG. 6 is a view for explaining a second embodiment of the method for manufacturing a gas turbine blade according to the present invention. The first
The same reference numerals are given to the same components as those of the embodiment or the corresponding components.

In the method of manufacturing a gas turbine blade according to the present embodiment, first, at the time of casting, a portion of the region A between the blade root portion 22 and the blade effective portion 24 (from the blade effective portion root portion 24a side to the blade effective portion central portion side). ), The cooling passage 25 for the blade root portion having a relatively large diameter and the cooling passage 26 for the blade effective portion root side.
Are manufactured by precision casting, and the area B of the blade effective portion 24 (from the center of the blade effective portion to the tip portion 23 of the blade effective portion) and the region C of the blade effective portion 24 (the center of the blade effective portion 24). Of the wing trailing edge 30 side).

Next, when the casting process is completed, the effective blade portion 24 is machined in each of the region B and the region C which are formed as solid, and the blade effective portion tip portion side cooling having a relatively small diameter is performed. Each of the passage 28 and the trailing edge side outlet 33 of the wing trailing edge 30 is drilled from each of the arrow L direction and the arrow M direction.

As described above, in the present embodiment, at the time of casting, the blade root portion cooling passage 25 having a relatively large diameter in the blade root portion 22 and the blade effective portion having a relatively large diameter are provided in the region A of the blade effective portion 24. In addition to producing the part-route-portion-side cooling passage 26 by precision casting, each of the region B and the region C of the blade effective portion 24 is made of a solid, and after casting, a region formed by a solid. In each of the portion B and the region C, the blade effective portion tip side cooling passage 28 and the trailing edge side outlet 33 of the blade trailing edge portion 30 having relatively small diameters are formed by machining. Therefore, a cooling structure can be easily manufactured even with a gas turbine blade having a large blade height installed in the latter stage of the turbine.

FIG. 7 is a view for explaining a third embodiment of the method for manufacturing a gas turbine blade according to the present invention. The first
The same reference numerals are given to the same components as those of the embodiment or the corresponding components.

The method for manufacturing a gas turbine blade according to the present embodiment is as follows. First, at the time of casting, the blade root portion 22 and the portion of the region A of the blade effective portion 24 (the blade effective portion root portion 24a side) are relatively separated. Each of the large-diameter blade root portion cooling passage 25 and the blade effective portion root portion side cooling passage 26 is manufactured with a precise structure, and is formed in the region D of the blade effective portion 24 (the center of the blade effective portion). For example, a reversing meandering passage 35 provided with a heat transfer promoting body 34 formed of, for example, a projecting piece or the like is manufactured by precision casting, and a portion of a region B of the wing effective portion 24 (the wing effective portion tip portion 23 side) is solid. Make it.

Next, when the casting process is completed, the blade effective portion 2
Reference numeral 4 denotes machining a portion of the region B which is formed as a solid, and drills a relatively small-diameter effective blade tip side cooling passage 28 in the direction of arrow L.

As described above, in this embodiment, at the time of casting, the blade root portion cooling passage 25 having a relatively large diameter in the blade root portion 22 and the blade effective portion having a relatively large diameter are formed in the region A of the blade effective portion 24. The cooling route 26 for the part root portion side is manufactured with a precision structure, and the inverted meandering passage 35 having the heat transfer promoting body 34 in the region D of the blade effective portion 24 is manufactured with the precision structure. 24 is made of a solid region B, and after casting, the blade effective portion tip side cooling passage 28 having a relatively small diameter is drilled by machining in the solid region B portion. A cooling structure can be easily manufactured even with a gas turbine blade having a large blade height installed in a later stage.

[0041]

As described above, the gas turbine blade manufacturing method according to the present invention provides a blade root portion cooling passage formed in a blade root portion and a blade effective portion root side cooling formed in a blade effective portion. At least one of the cooling structures such as the passage for use, the pin fins, and the reversing meandering passage is manufactured with a precision structure, and at least one or more of the tip side and the trailing edge side of the effective blade portion is solid. After casting, after casting
Since the cooling passage is drilled in the solid part by machining, even the complicated cooling structure of the gas turbine blade with a large blade height in the latter stage of the turbine can be easily manufactured, and it can sufficiently cope with high temperature. A cooling structure can be realized.

Further, in the gas turbine blade manufacturing method according to the present invention, the film cooling blowout holes formed at the blade leading edge, the ventral side, the back side, and the blade trailing edge are formed by casting after the blade casting, and then drilling by machining. Since it is provided, the time for the hole drilling operation can be shortened, the hole drilling operation can be easily performed, and the labor of the operator can be reduced.

[Brief description of the drawings]

FIG. 1 shows a first method of manufacturing a gas turbine blade according to the present invention.
FIG. 2 illustrates an embodiment.

FIG. 2 is a cross-sectional view as viewed from the direction of arrows ZZ in FIG. 1;

FIG. 3 is a cross-sectional view taken along the line YY of FIG. 1;

FIG. 4 is a cross-sectional view as seen from the direction of arrows XX in FIG. 1;

FIG. 5 shows a first example of the gas turbine blade manufacturing method according to the present invention.
FIG. 7 is a diagram illustrating a modification of the embodiment.

FIG. 6 shows a second example of the gas turbine blade manufacturing method according to the present invention.
FIG. 2 illustrates an embodiment.

FIG. 7 shows a third method of manufacturing the gas turbine blade according to the present invention.
FIG. 2 illustrates an embodiment.

FIG. 8 is a partially cutaway schematic sectional view showing a conventional gas turbine plant applied for power generation.

FIG. 9 is a schematic longitudinal sectional view of a gas turbine blade applied to a conventional gas turbine plant.

[Explanation of symbols]

 DESCRIPTION OF SYMBOLS 1 Gas turbine 2 Air compressor 3 Gas turbine combustor 4 Combustor liner 5 Combustion chamber 6 Transition piece 7 Turbine stationary blade 8 Turbine rotor blade 9 Turbine shaft 10 Blade root portion 10a Shank portion 10b Blade implant portion 10 Blade effective portion 11a , 11b Passage 12a, 12b Reverse meandering passage 13a, 13b Exit 14 Blade front edge 15 Shower head cooling air outlet 16 Pin fin 17 Blade trailing edge 18 Cooling air outlet 20 Shank portion 21 Blade implant 22 Blade root portion 23 Blade Effective part tip part 24 Blade effective part 24a Blade effective part root part 25 Blade root part cooling passage 26 Blade effective part root part side cooling passage 27 Pin fin 27c Blade effective part center part side cooling passage 28 Blade effective part tip part side Cooling passage 29 Blade leading edge 30 Blade trailing edge 31 Ventral side 32 Back side 33 Trailing edge Side outlet 34 Heat transfer enhancer 35 Reverse meandering passage 36 Film cooling outlet

Claims (6)

[Claims] A first cooling passage formed on the effective blade portion side of the effective blade portion and a solid shape formed on the effective tip portion of the effective blade portion are manufactured by precision casting. A method of manufacturing a gas turbine blade, further comprising forming, by machining, a second cooling passage communicating with the first cooling passage in a solid shape formed on the blade effective portion tip side. 2. A first cooling passage formed on the effective blade portion side of the effective blade portion, a pin fin formed transversely to the first cooling passage, and a tip portion of the effective blade portion of the effective blade portion. After the solid shape formed on the side of the wing is formed by precision casting, the solid shape formed on the tip portion of the wing effective portion is machined to communicate with the first cooling passage via the pin fin. A method for manufacturing a gas turbine blade, comprising: forming a cooling passage. 3. A first cooling passage formed on the blade effective portion root portion side of the effective blade portion, a pin fin formed across the first cooling passage, and a pin fin connected to the pin fin and connected to the pin fin. After the solid shape formed on the wing trailing edge portion and the solid shape formed on the wing effective portion tip side of the wing effective portion are manufactured by precision casting, the solid shape formed on the wing effective portion tip portion side Forming a second cooling passage communicating with the first cooling passage through the pin fin by machining to an actual shape,
A method for manufacturing a gas turbine blade, comprising: forming a trailing edge-side outlet that is connected to the pin fin by machining in a solid shape formed on the trailing edge of the blade. 4. A first cooling passage formed on the effective blade portion side of the effective blade portion on the side of the effective blade portion, a reversing meandering passage provided with a heat transfer promoting body across the first cooling passage, After the solid shape formed on the wing effective portion tip portion side of the portion is manufactured by precision casting, the solid shape formed on the wing effective portion tip portion side is machined through the inverted meandering passage by machining. A method for manufacturing a gas turbine blade, comprising forming a second cooling passage communicating with a first cooling passage. 5. At least one of a first cooling passage formed on the blade effective portion root portion side of the blade effective portion and a second cooling passage formed on the blade effective portion tip portion side of the effective blade portion. A film cooling blow-off hole for ejecting a cooling medium to the outside of the blade is provided on a blade wall of the blade effective portion, and the film cooling blow-off hole communicates with at least one of the first cooling passage and the second cooling passage. Forming the gas turbine blades by machining. 6. A cooling passage formed in the blade root portion and a first cooling passage formed on the blade effective portion root side of the blade effective portion and communicating with the first cooling passage. A method for manufacturing a gas turbine blade, wherein the blade root portion cooling passage is formed by machining.