JP3546134B2 - Gas turbine blade - Google Patents

Description

[0001]
TECHNICAL FIELD OF THE INVENTION
The present invention relates to a gas turbine blade used for thermal power generation and the like, and more particularly to a structure for simplifying a shroud cooling structure and improving cooling performance.
[0002]
[Prior art]
7A and 7B show a conventional typical gas turbine rotor blade, wherein FIG. 7A is a longitudinal sectional view of the rotor blade, and FIG. 7B is an EE sectional view thereof. In the figure, 21 is a rotor blade, 22 is a shroud at the tip, and 23 is a fin provided on the shroud 22. Reference numeral 24 denotes a multi-hole, that is, a small hole formed in the rotor blade 22. 25 is a pin fin provided on the inner wall of the rotor blade 21, and 26 is a rib for supporting the cavity 29. 27 is a hub portion, 28 is a blade root portion, and 29 is the aforementioned cavity.
[0003]
8 is a sectional view taken along line FF in FIG. 7, and FIG. 9 is a sectional view taken along line GG in FIG. 8. In both figures, two cavities 30 and 31 are independently formed inside the shroud 22. Plugs 32 and 33 are inserted into the upper and lower portions 30 and 31, respectively, to seal the inside. Multi holes 24 of the moving blade 21 communicate with the cavities 30 and 31, respectively, to supply cooling air. A plurality of cooling holes 34 communicate with the cavities 30 and 31 toward both sides, respectively, and the cooling holes 34 are open at both end surfaces to allow cooling air to flow out.
[0004]
In the rotor blade having the above-described configuration, cooling air flows into the cavity 29 from the blade root portion 28 as shown by the arrow, the heat transfer coefficient is improved by the pin fins 25, the base portion is cooled, and the cooling air flows through the multi-hole 24 to form the tip portion. Led to. From the tip end, it flows into the cavities 30, 31 of the shroud 22, passes through the cooling holes 34 from the cavities 30, 31, and flows out to both end surfaces of the shroud 22 facing each other, and cools the entire shroud.
[0005]
In the moving blade described above, an integral shroud 22 is formed integrally with the moving blade 21 at the tip of the moving blade 21 as described above. The shroud 22 reduces the gas leaking from the tip of the moving blade 21 and improves the vibration resistance of the moving blade 21 by forming a series of group blades by pressing the end face of the shroud 22 against the end face of the adjacent shroud 22. Let me. The rotor blade 21 generates vibrations in two directions, that is, a rotation axis direction and a circumferential direction. By forming the end face of the shroud 22 obliquely, vibrations in both directions are suppressed. Further, fins 23 are cut out from the shroud 22 in order to reduce gas leaking from the tip of the rotor blade 21 and to prevent contact with the casing.
[0006]
[Problems to be solved by the invention]
As described above, in the conventional gas turbine blade, the cooling air flows out of the cooling holes 34 on both sides of the shroud 22 opposite to each other and cools the entire shroud 22 to lower the temperature. From the cooling air, the cooling air merges from the multi-hole 24 of the rotor blade 21 into the cavities 30 and 31, flows from the cavities 30 and 31 into the cooling holes 34, which are divided into both sides and oppose each other, and flow to both sides. Therefore, a plurality of cooling holes 34 are arranged on both sides from each cavity, and due to the difference in the resistance of each cooling hole 34, the flow rate differs for each cooling hole, the cooling air does not flow evenly, and the cooling air is distributed uniformly. Adjustment is difficult and uniform cooling is not possible.
[0007]
Accordingly, the present invention provides a shroud for a gas turbine blade, in which a cooling structure that facilitates adjustment of a flow rate of cooling air flowing into a cooling hole provided in the shroud and has a uniform cooling effect, and a temperature of the shroud is uniformly reduced. It is an object to provide such a gas turbine blade.
[0008]
[Means for Solving the Problems]
The present invention provides the following means (1) to (3) to solve the above-mentioned problems.
[0009]
(1) A gas turbine rotor blade having a shroud at the blade tip, providing a cooling passage inside the blade from the base to the tip, flowing cooling air into the shroud, and flowing out from around the shroud. Two cavities are independently formed inside the shroud, and the two cavities communicate with cooling passages of the bucket, and a plurality of cooling holes through which cooling air flows toward opposite sides in the shroud. Wherein the cooling hole is linearly inclined obliquely downward toward the periphery of the shroud , extends to the periphery of the shroud, and opens on the lower surface of the periphery of the shroud. .
[0010]
(2) In the invention of the above (1), the cooling passage of the blade is an integral cavity extending over the entire length of the blade, and a number of pin fins are provided on the inner wall of the gas turbine. .
[0011]
(3) In the invention of the above (1), the cooling passage of the rotor blade is characterized in that a base side of the blade includes an integral cavity, a number of pin fins provided in the cavity, and a tip side has a number of small holes toward the tip. Gas turbine blades.
[0012]
In (1) of the present invention, two cavities are provided in the shroud, and the cooling holes communicating with the cavities are linearly inclined obliquely downward toward the periphery of the shroud and extend to the periphery of the shroud. In the periphery, the cooling air is opened obliquely to the lower surface of the shroud and flows out obliquely, so it is cooled up and down by the cooling holes inclined in the thickness direction of the shroud. Is effectively cooled, and the entire shroud is uniformly cooled.
[0013]
The shroud of (1) according to (2) of the present invention is provided at the tip of the blade formed by a pin fin provided in the cavity with an integral cavity over the entire length of the rotor blade, and the invention of ( 3 ). The shroud of the above (1) is provided with a cavity and pin fins on the base side and a blade tip having a number of small holes on the tip side, as in the conventional example, so that the blade has any type of cooling structure. The cooling effect by improving the heat transfer coefficient of the moving blade and the uniform cooling effect of the entire shroud enhance the cooling effect of the entire blade.
[0014]
BEST MODE FOR CARRYING OUT THE INVENTION
Hereinafter, embodiments of the present invention will be specifically described with reference to the drawings. FIG. 1 is a longitudinal sectional view of a gas turbine blade according to a first embodiment of the present invention. In the figure, 1 is a blade, 2 is a shroud at the tip, and 3 is a blade root. Reference numeral 4 denotes a rib, which is formed in the internal cavity 6 at the time of manufacturing and supports the cavity, and is not particularly required for the present invention. Reference numeral 5 denotes a number of pin fins fixed to both wall surfaces in the internal cavity 6. Reference numeral 6 denotes an internal cavity formed in the rotor blade 1 described above.
[0015]
As described above, the rotor blade according to the first embodiment of the present invention has an internal cavity 6 over its entire length, and has a structure in which a large number of pin fins 5 are provided to improve the flow and convection of cooling air inside, thereby enhancing the cooling effect. Further, as described below, the cooling of the tip shroud is also characterized.
[0016]
2 is a sectional view taken along the line AA in FIG. 1, and FIG. 3 is a sectional view taken along the line BB in FIG. In both figures, cavities 19 and 20 are independently formed in the shroud 2, and plugs 32 and 33 are mounted on the upper portions thereof as in the conventional case, and cooling air is separated from the cavity 6 of the blade. Is introduced. Cooling passages 7 to 12 communicate with the cavity 19 in directions opposite to both sides, and the cooling passages 7, 8, 9 are provided on one side, and the cooling passages 10, 11, 12 are provided with cooling air on the opposite side opposed thereto. It is a structure to let out.
[0017]
Similarly, the cooling passages 13 to 18 communicate with the cavity 20 on both sides, and the cooling passages 13, 14, 15 flow cooling air to one side, and the cooling passages 16, 17, 18 flow cooling air to the opposite side opposite to them. It has a structure to make it.
[0018]
FIG. 4 is a cross-sectional view taken along the line CC in FIG. 3, in which the cavity 20 communicates with the internal cavity 6 of the rotor blade 1, and the cooling passages 12, 16 are respectively inclined from the top to the bottom in the thickness direction of the shroud. The cooling air in the cavity 20 is obliquely discharged to the lower surface of the shroud 2. In addition, a plug 33 is provided above the cavity 20.
[0019]
In the rotor blade according to the first embodiment having the above-described configuration, the cooling air 30 flows into the blade from the base of the blade, convection is promoted by the pin fins 5 to cool the base, flow to the tip, and cool the shroud after cooling the blade. 2 flows into.
[0020]
In the shroud 2, the cooling air from the internal cavity 6 flows into the cavities 19 and 20, respectively, from the cavities 19 to the cooling passages 7, 8 and 9, and from the cooling passages 10, 11 and 12 to face these. 4 and flows out of the shroud 2, but the cooling air flows obliquely from the lower part since the respective cooling passages 7 to 12 are formed to be inclined in the shroud 2 as shown in FIG. 4. leak.
[0021]
Similarly, the fluid flows from the cavity 20 to one side from the cooling passages 13, 14, 15 and from the cooling passages 16, 17, 18 to the other opposite side, and flows out obliquely from the lower surface of the shroud 2. In this way, the cooling air flows obliquely downward from one side of the shroud 2 through the cooling passages 7-9, 12-15 and from the other side through the cooling passages 10-12, 16-18 to the periphery. Therefore, the entire surface can be uniformly cooled from the center to the periphery of the shroud 2.
[0022]
Further, since each cooling passage extends obliquely below the shroud 2 toward the periphery of the shroud, the cooling air is thermally transferred from the top to the bottom in the thickness direction of the shroud 2 and obliquely toward the periphery. Since the air flows in the direction having a large influence, the peripheral portion of the shroud can be effectively cooled.
[0023]
FIG. 5 is a cross-sectional view of a gas turbine blade according to a second embodiment of the present invention. The blade portion has the same configuration as that of the conventional example described with reference to FIG. This is a rotor blade composed of a multi-hole 24 composed of pin fins 25 and a number of small holes on the tip side from the base. Therefore, FIG. 5 is a cross-sectional view corresponding to the FF section shown in FIG. 7, and the structure of the rotor blade is the same as that of the conventional example shown in FIG. The shroud, which is a feature of the present invention, will be described in detail below.
[0024]
In FIG. 5, reference numeral 2 denotes a shroud, and reference numerals 19 and 20 denote cavities, which are independently formed in the shroud 2. Cooling passages 7, 8, and 9 communicate with the cavity 19 toward one side, and cooling passages 10, 11, and 12 communicate with each other toward the opposite side.
[0025]
Cooling passages 13, 14, 15 communicate with the cavity 20 toward one side, and cooling passages 16, 17, 18 communicate with each other toward the opposite side. 6, the upper portions of the cavities 19 and 20 are sealed by plugs 32 and 33, as in the first embodiment. These structures are the same as those of the second embodiment shown in FIG.
[0026]
As described in the conventional example of FIG. 7, a number of multiholes 24 are formed in the bucket 21 from the middle of the base toward the tip, and are divided into two groups, one into the cavity 19 and the other into the cavity 19. The cooling air 30 is led to each of the cavities 20.
[0027]
FIG. 6 is a cross-sectional view taken along the line DD in FIG. 5. The cavity 20 is formed in the shroud 2, the multi-hole 24 of the moving blade 21 communicates with the cavity 20 in the shroud 2, and cooling air flows from the moving blade 21 to the cavity. Leading to 2. Cooling passages 12 and 16 communicate with the cavity 20 on both sides, and the cooling passages 12 and 16 are provided to be inclined downward, respectively, and have a structure in which cooling air flows out obliquely from the lower surface of the shroud 2. The upper portion of the cavity 20 is sealed with a plug 33 as in the conventional example.
[0028]
In the second embodiment of the above configuration, as described in the conventional example, the cooling air flows into the blade from the base of the blade and promotes convection by the pin fins 25 to cool the base side, and the multihole 24, The tip cools through a number of small holes and then flows into shroud 2.
[0029]
Since a large number of multi-holes 24 communicate with the cavities 19 and 20 in the shroud 2, the cooling air fills the cavities 19 and 20, and one side of the shroud 2 is cooled by the cooling passages 7 to 9 and 13 to 15, respectively. And flows obliquely downward under the shroud 2 through the cooling passages 10 to 12 and 16 to 18, and uniformly flows from the center to the periphery of the shroud 2 as in the first embodiment. Can be cooled.
[0030]
【The invention's effect】
The gas turbine rotor blade of (1) of the present invention has a shroud at the tip of the blade, and a cooling passage is provided inside the blade from the base to the tip to flow cooling air into the shroud and guide the air into the shroud. In the gas turbine rotor blade flowing out from the surroundings, two cavities are independently formed inside the shroud, and the two cavities communicate with the cooling passages of the rotor blade, respectively, and are formed on opposite sides in the shroud. The cooling holes communicate with a plurality of cooling holes through which the cooling air flows.The cooling holes incline linearly diagonally downward toward the periphery of the shroud , extend around the shroud, and open at the lower surface around the shroud. It is characterized by doing. With such a configuration, two cavities are provided in the shroud, and the cooling holes communicating with the cavities are linearly inclined obliquely downward toward the periphery of the shroud and extend to the periphery of the shroud. Since the cooling air is opened obliquely to the lower surface of the shroud and flows out obliquely, it is cooled vertically by the cooling holes that are inclined in the thickness direction of the shroud. The lower surface of the shroud is effectively cooled, and the entire shroud is uniformly cooled.
[0031]
According to a second aspect of the present invention, in the first aspect of the present invention, the cooling passage of the rotor blade is formed of an integral cavity over the entire length of the blade, and a plurality of pin fins are provided on the inner wall of the cavity. According to a third aspect of the present invention, in the first aspect of the present invention, the cooling passage of the moving blade is formed by a plurality of pin fins provided in the same cavity as an integral cavity on the base side of the blade, and a plurality of small holes whose leading end side is directed toward the leading end. It is characterized by becoming. With such a configuration, it can be applied to the moving blade having any type of cooling structure, and the cooling effect by improving the heat transfer coefficient of the moving blade and the uniform cooling effect of the entire shroud enhance the cooling effect of the entire blade. It is.
[Brief description of the drawings]
FIG. 1 is a cross-sectional view of a gas turbine bucket according to a first embodiment of the present invention.
FIG. 2 is a sectional view taken along the line AA in FIG.
FIG. 3 is a sectional view taken along line BB in FIG.
FIG. 4 is a sectional view taken along line CC in FIG. 3;
FIG. 5 is a sectional view of a shroud of a gas turbine bucket according to a second embodiment of the present invention.
FIG. 6 is a sectional view taken along line DD in FIG. 5;
7A and 7B show a conventional gas turbine blade, in which FIG. 7A is a longitudinal sectional view, and FIG. 7B is a sectional view taken along the line EE in FIG.
FIG. 8 is a sectional view taken along line FF in FIG. 7;
9 is a sectional view taken along line GG in FIG.
[Explanation of symbols]
1,21 bucket 2 shroud 3 blade root 5,25 pin fin 6 internal cavity 7-18 cooling passage 19,20 cavity

Claims (3)

A gas turbine rotor blade having a shroud at a blade tip, providing a cooling passage from the base to the tip inside the blade, flowing cooling air into the shroud, and flowing out from around the shroud. , Two cavities are formed independently of each other, and the two cavities communicate with the cooling passages of the bucket and the plurality of cooling holes through which cooling air flows toward opposite sides in the shroud. A gas turbine rotor blade characterized in that the cooling hole is linearly inclined obliquely downward toward the periphery of the shroud , extends to the periphery of the shroud, and opens on the lower surface of the periphery of the shroud. 2. The gas turbine rotor blade according to claim 1, wherein the cooling passage of the rotor blade has an integral cavity over the entire length of the blade, and a plurality of pin fins are provided on an inner wall of the cavity. 2. The gas turbine rotor blade according to claim 1, wherein the cooling passage of the rotor blade includes a cavity formed integrally with the blade at the base side thereof, a number of pin fins provided in the cavity, and a number of small holes whose tip end faces the tip.

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