JPH10280908A - Stator blade of gas turbine - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a stator blade for a gas turbine which can be effectively cooled with less flow rate of a cooling medium, can recover heat in another heat engine by recovering the cooling medium used for cooling a blade and can meet crude fuel. SOLUTION: In a stator blade 10 for a gas turbine having an outside end wall 12 and an inside end wall 13 provided integrally with both radial sides of a blade effective part 11, a cooling medium recovery type cooling passage 20 is provided in the blade effective part 11. The cooling passage 20 comprises two independent systems 21 and 22. The two systems are formed so as to pass the cooling medium in opposed flows through a partition wall at least in a part.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a gas turbine stationary blade capable of efficiently cooling.

[0002]

2. Description of the Related Art A gas turbine system used in a power plant is generally configured as shown in FIG. That is, the compressed air compressed by the compressor 2 provided coaxially with the gas turbine 1 is supplied to the combustor 3, and the fuel is burned in the liner portion 3a of the combustor 3. Then, the high-temperature combustion gas generated by the combustion is guided from the transition piece 4 to the moving blade 6 via the stationary blade 5 of the gas turbine 1, and the rotating blade 6 is given a rotational force to perform work as a gas turbine. It is configured to be performed.

It is known that the higher the gas inlet temperature of the gas turbine 1 is, the better the thermal efficiency of the gas turbine system is to be improved. In fact, there is a tendency to increase turbine inlet temperatures to improve thermal efficiency. As described above, as the turbine inlet temperature rises, it is necessary to use materials capable of withstanding high temperatures for the combustor 3 and the stationary blades 5 and the moving blades 6 of the gas turbine 1. It is used as a component material.

However, at present, the limit temperature of a heat-resistant superalloy material that can be used as a gas turbine material is 80%.
0-900 ° C. On the other hand, the inlet temperature of the gas turbine has reached about 1300 ° C., far exceeding the limit temperature of the heat-resistant superalloy material. Therefore, in order to maintain the reliability of the gas turbine, it is essential to use a blade having a cooling structure for cooling the blade of the gas turbine to the limit temperature of the heat-resistant superalloy material.

At present, an air cooling system is employed in which the blades are cooled by using a part of the air discharged from the compressor. However, this air cooling system has essentially low cooling characteristics. Therefore, when the inlet temperature of the gas turbine exceeds 1300 ° C., the amount of air required for cooling the blades increases significantly. Moreover, sufficient cooling effect cannot be obtained only by convection cooling inside the blade, and a film cooling system that blows cooling air from the small holes formed on the surface of the effective blade to the outside of the blade must be used together. Absent. When this method is adopted, the temperature of the mainstream gas decreases due to the blown cooling air. Therefore,
The significance of raising the gas turbine inlet temperature is halved.

As described above, it is difficult to improve the thermal efficiency of the gas turbine in the air cooling system, and there is a possibility that the thermal efficiency of the combined power generation system as a whole is reduced. In addition, there is a problem that a poor fuel containing impurities may not be applied because there is a risk of clogging of a film cooling hole formed on the blade surface.

[0007]

As described above, in a conventional gas turbine blade, cooling using both a method of convectively cooling the inside of the blade with a cooling medium and a film cooling method of injecting the cooling medium onto the blade surface is used. Since the method is adopted,
Not only the type of the cooling medium is limited, but if the flow rate of the cooling medium is increased to increase the degree of cooling, the temperature of the mainstream gas may be reduced, and the thermal efficiency of the entire system may be reduced.
There was also a problem that the fuel used was restricted.

Accordingly, the present invention provides an efficient cooling with a smaller cooling medium flow rate without lowering the temperature of the mainstream gas, and recovering the cooling medium used for cooling the blades and recovering heat with another heat engine. It is another object of the present invention to provide a gas turbine vane that can be used and can cope with poor fuel.

[0009]

SUMMARY OF THE INVENTION In order to achieve the above object, the present invention relates to a gas turbine stationary blade having an outer end wall and an inner end wall integrally provided on both radial sides of an effective blade portion. A cooling medium recovery type cooling channel is provided in the effective portion.

The cooling passage provided in the effective blade portion is constituted by two independent systems, and at least a part of the two systems allows a cooling medium to flow in a counterflow through a partition wall. It may be configured as follows.

Further, the cooling passage provided in the effective blade portion may communicate with at least one of an end wall cooling passage provided in the outer and inner end walls.

Further, the cooling passage provided in the effective blade portion may be provided so as to be able to communicate with the cooling passage of two or more adjacent blades. In addition, a part of the cooling medium that has passed through the cooling passage or a part of the cooling medium that is supplied to the cooling flow path is formed by cooling the blade surface film, cooling the trailing edge of the blade, cooling the blade end wall, and sealing the blade. A means for using at least one of the fluids may be further provided.

The cooling medium is one selected from the group consisting of air, inert gas, steam, water, a gas-liquid mixed fluid of water and steam, a mixed fluid of different gases, or a mixed fluid thereof. Is also good.

Further, at least one or more heat shielding films or corrosion resistant films may be provided on the outer surface of the blade or the inner surface of the blade. In the present invention, a cooling medium recovery type cooling channel is basically provided in the effective blade portion. For this reason, the types of cooling media that can be used can be greatly expanded, and for example, steam or liquid having a higher cooling efficiency than air can be used as the cooling medium. Therefore, the stationary blade can be efficiently cooled with a small cooling medium flow rate. Further, since the cooling medium recovery type is configured, it is easy to recover energy from the cooling medium after passing through the cooling flow path, and it is possible to suppress a decrease in the mainstream gas temperature due to the cooling medium. It can contribute to improvement of thermal efficiency. In addition, since the film cooling system is not focused on, there is no need to consider the clogging of the injection holes. Therefore, it is possible to sufficiently cope with poor fuel.

[0015]

Embodiments of the present invention will be described below with reference to the drawings. FIG. 1 is a perspective view of a stationary blade of a gas turbine according to a first embodiment of the present invention. FIG. 1 is a perspective view of an example in which the present invention is applied to a stationary blade 10 used in a gas turbine of a combined cycle power plant or a thermal power plant. It is shown.

The stationary blade 10 includes a blade effective portion 11, and an outer end wall 12 and an inner end wall 13 provided integrally with the blade effective portion at both radial ends of the blade effective portion 11. .

Inside the outer end wall 12, the wing effective portion 11
Inside and inside the inner end wall 13, a cooling medium recovery type cooling passage 20 for cooling the above three elements is formed.

The cooling passage 20 is provided with a ventral cooling passage 21 mainly used for cooling the ventral side from the leading edge to the trailing edge of the effective blade portion 11, and from the leading edge of the effective blade portion 11. A rear cooling channel 22 mainly used for cooling the rear side toward the rear edge portion.

The ventilated cooling passages 21 each extend in the radial direction inside the effective blade portion 11 and are provided along the outer surface of the ventral side of the effective blade portion 11 with a plurality of radial passages 23, as shown in FIG. As shown, an outer end of each radial channel 23 is provided within the outer end wall 12 to extend along the ventral outer surface.
The cooling medium distribution passages 24 are commonly connected to every other, and the cooling medium distribution passages 24 are also provided in the outer end wall 12 and are commonly connected to the outer end of each of the radial flow passages 23 along the ventral outer surface. The folded-back flow path 26 provided in the inner end wall 13 and commonly connected to the inner ends of the radial flow paths 23 as shown in FIG.
1 and a cooling medium supply pipe 28 provided so as to be connected from the outer surface side of the outer end wall 12 to the supply port 27 of the cooling medium distribution flow path 24 shown in FIG. 2 as shown in FIG. As shown in FIG. 2, a cooling medium recovery pipe 30 is provided so as to be connected from the outer surface side of the outer end wall 12 to the recovery port 29 of the cooling medium recovery channel 25 shown in FIG.

On the other hand, each of the rear cooling passages 22 extends radially in the blade effective portion 11 and
Radial passages 31 provided along the back outer surface of the
And cooling medium distribution channels 32 provided in the outer end wall 12 and commonly connected to the outer end of each radial channel 31 along the back outer surface as shown in FIG.
FIG. 3 shows a cooling medium recovery flow path 33 also provided in the outer end wall 12 and commonly connected to the outer end of each radial flow path 31 every other along the abdominal outer surface. As shown in FIG. 1, formed inside the inner end wall 13 and connecting the inner ends of the radial flow paths 31 in common.
The cooling medium supply pipe 3 provided so as to be connected to the supply port 35 of the cooling medium distribution channel 32 shown in FIG.
6 and the outer end wall 12 as shown in FIG.
2 is connected to the recovery port 37 of the cooling medium recovery channel 33 shown in FIG.
8.

The cooling medium supply pipes 28 and 36 are
A cooling medium supply source (not shown) is connected to a steam supply source of a steam turbine system in this example, and each cooling medium recovery pipe 30, 38 is connected to an energy recovery system (not shown), in this example, a heat exchanger for energy recovery. It is connected. FIG.
In FIG. 2 and FIG. 2, reference numeral 39 denotes a cavity formed like a cavity in the wing effective portion 11.

In the stationary blade 10 of the gas turbine configured as described above, cooling is performed to the ventilated cooling flow path 21 and the back cooling flow path 22 through the cooling medium supply pipes 28 and 36 and the cooling medium recovery pipes 30 and 38. Steam as a medium is supplied, and the steam cools the blade effective portion 11, the outer end wall 12, and the inner end wall 13.

That is, taking the ventral cooling passage 21 as an example, the steam supplied through the cooling medium supply pipe 28 is placed in the radial passage 23 through the cooling medium distribution passage 24 at every other position. Supplied. The steam removes heat from the blade effective portion 11 while flowing through the radial channel 23. The steam flowing through the radial flow path 23 passes through the folded formation flow path 26 formed in the inner end wall 13, and the remaining radial flow path 23
After that, it flows to the cooling medium recovery pipe 30 through the cooling medium recovery channel 25.

In this example, adjacent radial channels 2
The flow direction of the steam flowing through 3 is the opposite direction, that is, the flow direction is the counterflow with the partition wall as a boundary. For this reason, the temperature of the steam flowing in the outward path and the temperature of the steam flowing in the return path can be made substantially the same by heat exchange via the partition. As a result, the wing effective portion 11 is controlled so that the temperature of the wing effective portion 11 becomes substantially uniform.
Can be cooled. Then, the cooling medium recovery pipe 30
The steam flowing into the heat exchanger is guided to a heat exchanger for energy recovery. The same cooling operation is performed on the back cooling channel 22.

As described above, the cooling flow path 20 of the cooling medium recovery type is provided in the effective blade portion 11. Therefore, the types of cooling media that can be used can be greatly expanded, and steam having a higher cooling efficiency than air can be used as the cooling medium as in this example. Therefore, the stationary blade can be efficiently cooled with a small cooling medium flow rate. Further, since the cooling medium is configured as a cooling medium recovery type, it is easy to recover energy from the cooling medium after passing through the cooling flow path 20. In addition, since the temperature of the mainstream gas can be prevented from lowering due to the cooling medium, it is possible to contribute to improvement in thermal efficiency. In addition, since the film cooling system is not focused on, there is no need to consider the clogging of the injection holes. Therefore, it is possible to sufficiently cope with poor fuel.

FIG. 4 is a perspective view of a stationary blade 10a of a gas turbine according to a second embodiment of the present invention. In this figure, the same functional parts as those in FIG. 1 are indicated by the same reference numerals. Therefore, a detailed description of the overlapping part will be omitted.

In the stationary blade 10 a of the gas turbine according to this example, a cooling medium supply pipe 28 and a cooling medium recovery pipe 30 connected to the ventilated cooling flow path 21 are provided on the inner end wall 13 side. Therefore, in this example, the ventral cooling passage 2
1. A folded channel 2 as shown in FIG.
6 is provided in the outer end wall 12.

The same effect as the stationary blade shown in FIG. 1 can be exerted on the stationary blade 10a of the gas turbine configured as described above. FIG. 5 is a perspective view of a stationary blade 10b of a gas turbine according to a third embodiment of the present invention. In this figure, the same functional parts as those in FIG. 1 are indicated by the same reference numerals. Therefore, a detailed description of the overlapping part will be omitted.

In the stationary blade 10b of the gas turbine according to this example, the ventilating cooling passage 21a and the back cooling passage 22a are each separated into two systems. Then, the cooling medium supply pipes 28a and 28b of the ventral side cooling flow path 21a are positioned on the outer end wall 12 side, and the cooling medium collection pipe 30
a, 30b are positioned on the inner end wall 13 side, and the cooling medium supply pipes 36a, 36 of the back side cooling flow path 22a are further positioned.
b is located on the inner end wall 13 side, and the cooling medium recovery pipes 38a, 38b are connected to the outer end wall 12 side.
Side.

Even with this configuration, the same effect as in the above example can be exhibited. FIG. 6 shows a main part of a stationary blade 10c of a gas turbine according to a fourth embodiment of the present invention, in which a flow path configuration in an outer end wall 12 is shown.

In the stationary vane 10c of the gas turbine, the ventilating cooling passage 21 and the back cooling passage 22 for cooling the blade effective portion 11 and the outer end wall 12 independent of the cooling passages 21 and 22 are provided. And cooling passages 40 and 41 for cooling the inner end wall (not shown) (the cooling passage 41 is not shown).

That is, the supply port 42 of the ventral cooling passage 21
The recovery port 43 communicates with a cooling medium supply pipe and a cooling medium recovery pipe (not shown) provided on the outer end wall 12 side. It communicates with a cooling medium supply pipe and a cooling medium recovery pipe (not shown) provided on the wall side. Further, the supply port 44 and the recovery port 45 of the cooling flow path 40 provided in the outer end wall 12 communicate with a cooling medium supply pipe and a cooling medium recovery pipe (not shown) provided on the outer end wall 12 side. Cooling channel 41 provided on an inner end wall (not shown)
Are connected to a cooling medium supply pipe and a cooling medium recovery pipe provided on the inner end wall side.

As described above, the same effect as in the above example can be exerted by employing a configuration in which the system for cooling the blade effective portion 11 and the system for cooling the outer and inner end walls are made independent.

FIG. 7 shows a schematic configuration of a stationary blade 10d of a gas turbine according to a fifth embodiment of the present invention. This example is a modification of the example shown in FIG. In this figure, the same functional parts as those in FIG. 1 are indicated by the same reference numerals.
Therefore, a detailed description of the overlapping part will be omitted.

The stationary blade 10d of the gas turbine according to this embodiment
Has a ventral-side cooling channel 21 and a back-side cooling channel 22 similarly to the one shown in FIG. Then, when actually installed inside the turbine casing, the ventilating cooling passages and the rear cooling passages of the adjacent stationary blades are connected in series.

With this configuration, the same effects as those of the above-described embodiments can be obtained, and the cooling medium supply path and the cooling medium recovery path can be simplified. As shown in FIG. 8 as a modification, in order to make the cooling temperature uniform, the ventral cooling passage 21 and the back cooling passage 22
A heat exchanger 47 for exchanging heat between the cooling media flowing through the two flow paths may be provided in the middle of the process.

Further, as shown as a further modified example in FIG. 9, cooling medium supply ports 48a and 48b may be provided in the middle of the ventral cooling passage 21 and the back cooling passage 22. FIG. 10 shows a schematic configuration of a stationary blade 10e of a gas turbine according to a sixth embodiment of the present invention. This example is a modification of the wing shown in FIG. In this figure, the same functional parts as those in FIG. 5 are indicated by the same reference numerals. Therefore,
A detailed description of the overlapping part will be omitted.

In the stationary blade 10e according to this example, the blade effective portion 1
1, a cavity 50 is provided in a sealed form, the cavity 50 is communicated with, for example, a cooling medium supply pipe 28a or 28b provided on the outer end wall 12 side, and the cavity 50 extends from the cavity 50 to the surface of the blade effective portion 11. A plurality of extending film cooling holes 51 are provided.
The cooling medium is blown out from the nozzle.

Even with such a configuration, the same effects as those of the above examples can be exerted. FIG. 11 shows the seventh embodiment of the present invention.
1 shows a schematic configuration of a stationary blade 10f of a gas turbine according to the embodiment.

In the stationary blade 10f according to this example, the blade effective portion 1
1 and cooling channels 54 and 55 for independently cooling the outer and inner end walls 12 and 13, respectively.
A part of the cooling medium supplied to the nozzles and a part of the cooling medium passing through these flow paths are used as seal fluids 56 and 57 for sealing between a moving blade (not shown).

Such a configuration can be adopted.
FIG. 12 shows a schematic configuration of a stationary blade 10g of a gas turbine according to an eighth embodiment of the present invention.

In the stationary blade 10g according to this example, the blade effective portion 1
1 is cooled by cooling flow paths 58 and 59 of a cooling medium recovery type, and a dust remover 60 is disposed upstream of the cooling flow paths 58 and 59 for the purpose of removing inclusions, scale, and the like in the cooling medium.
Is provided.

Such a configuration can be adopted.
The present invention is not limited to the above-described example, and can be implemented with various modifications. For example, in each of the above examples, steam is used as the cooling medium. However, the cooling medium is not limited to steam, but may be air, an inert gas, water, a gas-liquid mixed fluid of water and steam, a mixed fluid of different gases, or a mixed fluid thereof. One selected from among them can be used.

Further, the rear edge of the blade may be convectively cooled by a part of the cooling medium and then ejected to the outside from the rear edge. Further, at least one or more heat shielding films or corrosion resistant films may be provided on the outer surface or inner surface of the blade.

[0045]

As described above, according to the present invention,
Since the cooling medium recovery type cooling flow path is provided in the effective blade portion, the types of cooling medium that can be used can be greatly expanded, and for example, steam or liquid having higher cooling efficiency than air can be used as the cooling medium. Therefore, the stationary blade can be efficiently cooled with a small cooling medium flow rate. Further, since the cooling medium recovery type is configured, it is easy to recover energy from the cooling medium after passing through the cooling flow path, and it is possible to suppress a decrease in the mainstream gas temperature due to the cooling medium. Thermal efficiency can be improved. Also,
There is no need to pay attention to clogging of the injection holes since the main focus is not on film cooling. Therefore, it is possible to sufficiently cope with poor fuel.

[Brief description of the drawings]

FIG. 1 is a perspective view showing a gas turbine according to a first embodiment of the present invention with a stationary blade partially cut away;

FIG. 2 is a schematic configuration diagram showing an outer end wall of the stator blade cut in a direction orthogonal to a direction in which the blade effective portion extends.

FIG. 3 is a schematic configuration diagram showing an inner end wall of the stator blade cut in a direction orthogonal to a direction in which the blade effective portion extends.

FIG. 4 is a perspective view showing a gas turbine according to a second embodiment of the present invention, with a stationary blade partially cut away;

FIG. 5 is a perspective view showing a gas turbine according to a third embodiment of the present invention, with a stationary blade partially cut away;

FIG. 6 is a schematic configuration diagram of a stationary blade of a gas turbine according to a fourth embodiment of the present invention, in which an outer end wall is cut away.

FIG. 7 is a schematic configuration diagram of a vane of a gas turbine according to a fifth embodiment of the present invention.

FIG. 8 is a schematic diagram for explaining a modification of the stationary blade of the gas turbine shown in FIG. 7;

FIG. 9 is a schematic diagram for explaining still another modification of the stationary blade of the gas turbine shown in FIG. 7;

FIG. 10 is a partially cutaway perspective view showing a stationary blade of a gas turbine according to a sixth embodiment of the present invention.

FIG. 11 is a schematic configuration diagram of a stationary blade of a gas turbine according to a seventh embodiment of the present invention.

FIG. 12 is a schematic configuration diagram of a vane of a gas turbine according to an eighth embodiment of the present invention.

FIG. 13 is a schematic diagram showing a main part of the gas turbine system.

[Explanation of symbols]

10, 10a, 10b, 10c, 10d, 10e, 10
f, 10 g ... stationary blade 11 ... blade effective part 12 ... outer end wall 13 ... inner end wall 20, 20a ... cooling channel 21, 21a ... ventral side cooling channel 22, 22a ... back side cooling channel 23, 31 ... Radial flow paths 24, 32 ... Cooling medium distribution flow paths 25, 33 ... Cooling medium recovery flow paths 26, 34 ... Turnback forming flow paths 27, 35, 42, 44 ... Supply ports 28, 28a, 28b, 36, 36a, 36b cooling medium supply pipes 29, 37, 43, 45 recovery ports 30, 30a, 30b, 38, 38a, 38b cooling medium recovery pipes 39, 50 cavities 40, 41, 52, 53, 54, 55,. Cooling channel 47: Heat exchanger 51: Film cooling hole

 ──────────────────────────────────────────────────続 き Continuing on the front page (72) Katsuyasu Ito 2-4, Suehirocho, Tsurumi-ku, Yokohama-shi, Kanagawa Prefecture Inside the Toshiba Keihin Plant (72) Inventor Shigeru Triangle 1-1-1, Shibaura, Minato-ku, Tokyo Toshiba Corporation Head Office

Claims (7)

[Claims] 1. A stationary blade of a gas turbine integrally having an outer end wall and an inner end wall on both sides in the radial direction of an effective blade portion, wherein a cooling flow path of a cooling medium recovery type is provided in the effective blade portion. A stationary vane for a gas turbine, characterized in that: 2. A cooling channel provided in the effective blade portion,
2. The system according to claim 1, wherein the cooling system is configured to include two independent systems, and the two systems are configured to allow a cooling medium to flow through the partition wall in at least a part in a counter-current manner. 3. Gas turbine stationary blades. 3. The cooling passage provided in the effective blade portion,
3. The gas turbine vane according to claim 1, wherein the vane communicates with at least one of an end wall cooling passage provided in the outer and inner end walls. 4. 4. A cooling channel provided in the effective blade portion,
4. The method according to claim 1, wherein the cooling passages are provided so as to communicate with two or more adjacent blades.
A vane of the gas turbine according to the paragraph. 5. A part of the cooling medium passing through the cooling passage or a part of the cooling medium supplied to the cooling passage is cooled by a blade film cooling, a blade trailing edge cooling, a blade end wall cooling, and a moving blade. The vane of a gas turbine according to claim 1, further comprising a unit used for at least one of a sealing fluid between the vanes. 6. The cooling medium is one selected from the group consisting of air, inert gas, steam, water, a gas-liquid mixed fluid of water and steam, a mixed fluid of different gases, and a mixed fluid thereof. The vane of the gas turbine according to claim 1, wherein: 7. The gas turbine vane according to claim 1, wherein at least one or more heat shielding films or corrosion-resistant films are further provided on an outer surface or an inner surface of the blade.