JP2000034902A - Cooling rotor blade for gas turbine - Google Patents

Abstract

PROBLEM TO BE SOLVED: To enhance cooling efficiency of a cooling air and sealing performance of a sealing air. SOLUTION: Cooling airs 100, 101, and 102 flow in from cooling passages 3, 4, and 5 to a rotor blade 1. The cooling air 100 passes through a passage 3a, with its cooling efficiency enhanced by turbulators 9a and 9b, cools a front- rim part, and flows out of the front end. The cooling airs 101 and 102 flow in from the back-rim side, passes through a serpentine channel consisting of passages 6a, 6b, 6c, 6d, and 6e, cools the rotor blade 1 together with the working of a turbulator 8, and flows out from the front end of the passage 6e. Through the introduction of cooling airs 100, 101, 102 the cooling efficiency is improved. Also, sealing airs 103 and 104 pass through gaps of knife edges 13 and 14 and flow out of passages 10 and 11 obliquely upwards cambering in the flowing direction of the combustion gas. Thereby, sealing performance is enhanced, and also, an edge part 2a of a platform 2 and a shroud edge part inside a stator blade at a rear stage are cooled from the outlowing air.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a cooling blade for a gas turbine, and more particularly to a cooling blade having a structure in which the flow of cooling air is improved to enhance the cooling efficiency, and the flow of sealing air is also improved to improve the sealing performance. It is a wing.

[0002]

2. Description of the Related Art FIG. 3 is a typical sectional view of a one-stage moving blade of a conventional gas turbine. In the figure, reference numeral 20 denotes a blade, 21 denotes a blade root, and 22 denotes a platform. Cooling passages 23, 24, 25, and 26 are provided independently of each other at the blade root portion. The cooling passage 23 is a passage on the leading edge side, communicates with a cooling passage 23a in the blade, and cools air 40 from the rotor side. It flows in and flows out of the cooling hole 29 while cooling the front edge from the cooling passage 23a, and the front edge is cooled by the shower head film. Cooling air 41 flows from the cooling passage 24, passes through the cooling passage 24a in the blade,
b, flows toward the base from 24b, flows into 24c, flows out from the tip, and in this process flows out from the cooling hole to the blade surface as described later with reference to FIG.

[0003] Cooling air 42 flows from the cooling passage 25, and cooling air 43 flows from the passage 28, merges and flows through the cooling passage 25a, enters the passage 25b from the front end, flows toward the base, and flows into the 25c. In the course of flowing through the passage 25c, as described later with reference to FIG. 4, the film flows out from the cooling hole to the surface to perform film cooling, and the remaining air flows out from the cooling hole 28 in the trailing edge 27 to perform pin fin cooling.

FIG. 4 is a sectional view taken along line BB in FIG.
As shown in the drawing, a part of the cooling air in the cooling passage 23a at the leading edge flows out of the blade through the cooling hole 29, performs showerhead film cooling, and cools the leading edge. Also, cooling passage 2
Part of the cooling air flowing through the inside of the cooling holes 4c obliquely flows out of the cooling holes 30 in the process of flowing, performs film cooling of the surface, and similarly, in the process of flowing through the passages 25c, the cooling holes 31
From the surface of the wing, and the trailing edge is film-cooled. In the example shown, the cooling holes are 29, 30,
Although only 31 locations are shown, a number of cooling holes are provided in addition to this.

FIG. 5 is a diagram showing the flow of sealing air. In the drawing, a stationary blade 50 is arranged at a stage subsequent to the moving blade 20, 51 is an inner shroud, and 52 is an outer shroud. 53 is a cavity inside the inner shroud end 51, 54 is a labyrinth seal 58 which is a seal ring holding ring.
Holding. The labyrinth seal 58 forms a disk and a seal portion of the rotor. Reference numeral 55 denotes a hole formed in the seal ring holding ring, and sealing air flows out as described later. Reference numeral 56 denotes an adjacent moving blade 20 and a stationary blade 50.
57 is a honeycomb seal at the inner shroud end 57a of the stationary blade 50, 58 is a labyrinth seal described above, and 59 is a space formed between the stationary blade 50 and an adjacent preceding moving blade. is there.

In the above configuration, the sealing air 70 passes through the sealing tube 71 provided in the stationary blade 50,
Flows into the cavity 53 in the inner shroud end 51,
The inside of the cavity 53 is made higher than the pressure of the external combustion gas passage, flows out of the hole 55 of the seal ring retaining ring 54 into the space 56, and from the space 56, the end 22 b of the platform 22 of the moving blade 20 and the inner shroud of the stationary blade 50. Passing through the gap of the seal portion formed by the honeycomb seal 57 at the end portion 57a,
It flows out into the combustion gas passage as indicated by 70a. On the other hand, a part of the cooling air flowing out of the hole 55 of the seal ring holding ring 54 passes through the gap between the labyrinth seal 58 and the rotor disk 80, flows out into the space 59, and from the rear end of the shroud 51 inside the space 59. It flows out as 70b as in the previous stage.

The same is true between the moving blade 20 and the preceding stationary vane, and the sealing air 70c similarly flows out between the front end 22a of the platform 22 and the preceding stationary vane.
In this manner, the inside of the inner shroud 51 of the stationary blade 50 and the inside of the platform 22 of the moving blade 20 are set at a higher pressure than the combustion gas passage, thereby preventing high-temperature combustion gas from flowing into these inside spaces.

[0008]

As described above, in a single-stage bucket of a gas turbine, cooling air is guided into the blades to perform cooling. The cooling air flows from a cooling hole to a blade surface in a process of flowing through an internal cooling passage. To cool the showerhead film on the leading edge of the wing surface and cool the film on the ventral and dorsal sides of the wing. In recent gas turbines, the temperature of combustion gas has been increasing, and the temperature at the combustor outlet has been increased to about 1150 ° C. In recent years, plants having a temperature of 1300 ° C. or more are being developed. Along with this, the first stage rotor blade is the part most exposed to particularly high temperature combustion gas,
It is necessary to cool the blade more efficiently, and further improvement of the cooling structure is desired.

In addition, the sealing air is also improved in sealing performance as the temperature of the combustion gas is increased, and a sufficient sealing pressure is ensured so that the combustion gas does not enter the inside of the platform or the shroud. You have to use it well.

Accordingly, the present invention provides a gas turbine cooling blade,
Especially in the first stage rotor blades, the cooling air passage was improved,
Cooling air efficiently flows through the cooling passage to increase cooling efficiency,
Further, it is an object of the present invention to improve the cooling performance and the sealing performance by adopting a shape for increasing the sealing effect also in the flow path of the sealing air.

[0011]

The present invention provides the following means (1) and (2) to solve the above-mentioned problems.

(1) A cooling air passage is provided inside the moving blade, cooling air is guided from the lower part of the platform through the passage to cool the blade, and the inner shroud ends of the adjacent front and rear stationary vanes and the platform are arranged. In a gas turbine cooling blade having a seal portion provided between the end portion and a sealing air from the seal portion to a combustion gas passage, the cooling air passage allows cooling air to flow to a front edge portion and to flow to a tip end. A first passage for guiding the cooling air from the trailing edge side to allow a portion to flow out from a number of slots provided on the trailing edge, and a second passage to flow the serpentine flow path toward the leading edge side and to the leading end; And a gas turbine cooling blade.

(2) In the above invention (1), the seal portion forms a seal with a knife edge provided at the platform end and an inner shroud end of an adjacent stationary blade, and meanders from the seal. A gas turbine cooling blade having a flow path formed therein, wherein the sealing air flows from the meandering flow path to the combustion gas flow path inclining upward in the combustion gas flow direction.

In (1) of the present invention, in the first cooling passage, the cooling air is exposed to the hottest combustion gas, cools the leading edge which is severely affected by heat, and flows out from the tip as it is, so that the cooling air is cooled. The leading edge is effectively cooled. In the second cooling passage, the cooling air is cooled at the trailing edge first with cold air, cools the hub portion where the fatigue strength is reduced due to the thermal influence, prevents the fatigue strength from being lowered, and partially cools the rear portion. The air flows out of the slot at the edge to cool the trailing edge fin, then flows into the serpentine flow path, cools the main part of the blade, flows toward the leading edge, and flows out from the tip. Since the first and second passages are provided independently from the leading edge side and the trailing edge side, the cooling efficiency is improved.

In (2) of the present invention, since the sealing air forms the seal by the knife edge, the sealing pressure is increased at this portion, and the sealing air flowing out of the seal flows meandering to reduce the amount of leakage. To reduce the number and improve the sealing performance. In addition, since the outflowing sealing air flows into the combustion gas passage obliquely upward in the flow direction of the combustion gas, the flow hits the end of the platform or the end of the inner shroud of the stationary blade, and acts to cool these portions. Is also combined.

[0016]

Embodiments of the present invention will be specifically described below with reference to the drawings. FIG. 1 is an internal cross-sectional view of a gas turbine cooling blade according to an embodiment of the present invention, and particularly shows an example of a single-stage blade, and FIG. 2 is an AA cross-sectional view thereof.

In both figures, 1 is a rotor blade and 2 is its platform. Reference numerals 3, 4, and 5 denote cooling passages, which are formed in the blade root portion 16 and are supplied with cooling air supplied from the rotor side. The cooling passage 3 is a passage 3 in the blade.
a, the cooling passages 4 and 5 join together in the middle of the blade root 16 to form a cooling passage 6. These cooling passages 3, 4,
5 is smaller in number than the conventional one shown in FIG.
In No. 5, the flow path is narrowed by a rib to adjust the inflow amount.

The cooling passage 6 is provided on the trailing edge side and communicates with the passages 6a, 6b, 6c, 6d and 6e in the blade to form a serpentine cooling passage in the blade. 7 is the trailing edge,
A number of slots 15 are drilled so as to be oriented substantially in the axial direction. Reference numeral 8 denotes an oblique turbulator provided on the wall surface of each of the passages 6a to 6e, and 9a and 9b represent passages 3a on the front edge side.
It is a turbulator provided on the inner wall surface.

Reference numeral 10 denotes an inclined portion, and the platform 2
This is a portion in which the upper surface of the front end 2a on the gas passage side is formed obliquely upward and forms a passage 17 between the adjacent front stationary vane. Numeral 11 also denotes an inclined portion, which is a portion in which the inner lower surface of the rear end 2c of the platform 2 is formed obliquely upward, and forms a passage 18 between the adjacent front stationary vane. Reference numerals 12a, 12b, and 12c denote seal pins that seal between circumferentially adjacent blade platforms.

Reference numerals 13 and 14 denote knife edges, 13 is provided at the front end 2a of the platform 2 and 14 is provided at the rear end 2d, respectively, and is located close to the seal portion of the adjacent stationary vane. Is sealed. 100, 10
1, 102 is cooling air supplied from the rotor side,
The coolant flows into the cooling passages 3, 4, and 5, respectively. The surface of the rotor blade 1 is made of TBC (Thermal
Barrier Coating is applied to deal with the effects of heat from the hot combustion gases.

In the cooling blade having the above configuration, the cooling air 1
00 flows linearly from the cooling passage 3 to the passage 3a, is disturbed by the turbulators 9a and 9b, improves the heat transfer coefficient, and is exposed to a high temperature. It flows out to the combustion gas passage.

The cooling air 101, 102 is supplied to the cooling passages 4, 5
The turbulator 8 flows into the cooling passage 6 and flows toward the front end thereof. The turbulator 8 cools the trailing edge while increasing the heat transfer effect. 7 flows out of the large number of slots 15 provided substantially in the axial direction and cools the rear edge.

The cooling air from the passage 6a enters the passage 6b from the leading end, flows toward the platform 2, cools the blades in the same manner, inverts from the lower part of the platform 2 and flows into the passage 6c, and the leading end of the passage 6c. Flows from the side into the passage 6d, flows into the passage 6e from the lower part of the platform 2, and flows out from the tip. Thus, the cooling air 10
In the process of flowing through the serpentine flow path composed of the passages 6a, 6b, 6c, 6d, and 6e, the turbulators 8 obliquely disturb the flow and agitate the flow to enhance the cooling effect. The part is cooled, and the cooled air flows out from the tip.

As described above, the cooling air 101 and 102 are made to flow in from the trailing edge side of the rotor blade 1 to flow cool air to the trailing edge hub portion, and in particular, to improve the fatigue strength of the trailing edge hub portion due to heat. Since the cold air flows into the passage 6a first, the fin portion at the trailing edge is effectively cooled, the number of the oblique turbulators 8 in the passage 6a is smaller than that of the other passages, and the passage 6a itself is also more than the other passages. Can be bigger.
Further, in the moving blade 1 adopting such a cooling method, the cooling holes 29, 30, and 31 as in the related art become unnecessary, and the cooling of the shower head film becomes unnecessary.

As described with reference to FIG. 5, the sealing air 103 flows from inside the adjacent stationary blade, flows between the adjacent stationary blade on the front stage side, and flows between the knife edge 13 and the sealing portion of the adjacent stationary blade. It flows out, flows in a meandering manner, and flows out obliquely from the passage 17 formed by the inclined portion 10. The meandering flow enhances sealing performance and reduces loss. The outflowing sealing air flows out obliquely upward in the flow direction of the combustion gas, and hits the end 2 a of the platform 2.
It also has the effect of cooling this part.

Further, on the downstream side, the sealing air 104 similarly flows out of a gap formed by the sealing portion of the downstream stationary blade and the knife edge 14, and the combustion gas flows from the passage 18 formed by the inclined portion 11 while meandering. Flows upward in the direction of flow,
In the same manner as described above, the sealing performance is enhanced, and the leaked sealing air also has the effect of cooling the inner shroud end of the adjacent vane.

As described above, in the present embodiment, the cooling air 100 flows straight into the passage 3a on the front edge side to cool only the front edge, and allows the cooling air 101 and 102 to flow from the rear edge side. , Passages 6a, 6b, 6c, 6d, 6e
The blades are effectively cooled by the serpentine flow paths formed by the above and the oblique turbulators 8 provided in these paths, so that the cooling effect is improved.

Further, the outflow path of the sealing air 103, 104 is meandered by the knife edges 13, 14 and the obliquely upward passages 17, 18 formed by the inclined portions 10, 11, so that the outflow path is sealed. Is improved,
The outflowing sealing air can also cool the end 2a of the platform 2 and also cool the inner shroud end of the adjacent vane.

[0029]

As described above, the gas turbine cooling blade of the present invention comprises:
(1) A cooling air passage is provided inside the moving blade, cooling air is guided from the lower part of the platform through the passage to cool the blade, and the inner shroud end of the adjacent front and rear stationary vanes and the platform end are connected to each other. In the gas turbine cooling blade, in which a seal portion is provided between the seal portions and the sealing air flows out to the combustion gas passage, the cooling air passage causes the cooling air to flow to the front edge portion and to flow out to the tip. A passage and a second passage for guiding cooling air from the trailing edge, allowing a part of the cooling air to flow out from a number of slots provided on the trailing edge, and the remainder to flow the serpentine flow path toward the leading edge and to the leading end. It is characterized by being done.
With such a configuration, the front edge portion is formed by the first cooling passage,
The second cooling passage effectively cools the trailing edge portion and the main portion, respectively, and improves the cooling efficiency.

According to a second aspect of the present invention, in the first aspect of the present invention, the seal portion forms a seal with a knife edge provided at an end of the platform and an inner shroud end of an adjacent stationary blade. A flow path meandering from the seal is formed, and sealing air is inclined upward in the flow direction of the combustion gas from the meandering flow path and flows out to the combustion gas flow path. With such a configuration, the sealing pressure is sufficiently ensured by the knife edge, and the resistance of the flow path is increased by the meandering passage, whereby the sealing performance can be improved. Furthermore, the sealing air that flows out also has the effect of cooling the platform and the end of the inner shroud.

[Brief description of the drawings]

FIG. 1 is an internal cross-sectional view of a gas turbine cooling blade according to an embodiment of the present invention.

FIG. 2 is a sectional view taken along line AA in FIG.

FIG. 3 is a cross-sectional view of a conventional gas turbine cooling blade.

FIG. 4 is a sectional view taken along line BB in FIG. 3;

FIG. 5 is a diagram showing a flow of sealing air of a conventional gas turbine.

[Explanation of symbols]

 Reference Signs List 1 rotor blade 2 platform 2a, 2b, 2c, 2d end 3,4,5 cooling passage 3a passage 6a, 6b, 6c, 6d, 6e passage 7 trailing edge 8,9a, 9b turbulator 10,11 inclined portion 13,14 Knife edge 15 Slot 17, 18 Passage 100, 101, 102 Cooling air

 ──────────────────────────────────────────────────続 き Continuing on the front page (72) Inventor Junichiro Masada 2-1-1, Shinhama, Arai-machi, Takasago City, Hyogo Prefecture Inside the Takasago Works, Mitsubishi Heavy Industries, Ltd. No.1 F-term in Takasago Works, Mitsubishi Heavy Industries, Ltd. (reference) 3G002 CA06 CA08 CB03 CB05

Claims (2)

[Claims] 1. A cooling air passage is provided inside a moving blade, cooling air is guided from a lower portion of the platform through the passage to cool the blade, and the inner shroud ends of the adjacent preceding and succeeding stationary vanes and the platform end. In the gas turbine cooling blade where a sealing portion is provided between the cooling portion and the sealing portion to flow out to the combustion gas passage from the sealing portion, the cooling air passage causes the cooling air to flow to the front edge portion and to flow out to the tip. A second passage for guiding the cooling air from the trailing edge side, causing a part of the cooling air to flow out from a number of slots provided on the trailing edge, and leaving the remainder flowing through the serpentine flow path toward the leading edge side to the leading end; And a gas turbine cooling blade. 2. A seal comprising a knife edge provided at an end of the platform and an inner shroud end of an adjacent stator vane, and a flow path meandering from the seal. 2. The gas turbine cooling blade according to claim 1, wherein the sealing air flows from the flow path into the combustion gas flow path inclined upward in the combustion gas flow direction.