JPH1061406A - Cooling structure for gas turbine blade - Google Patents

Abstract

PROBLEM TO BE SOLVED: To hardly generate the cracking of a blade wall, enhance the soundness of a blade, and extent the life of the blade, by arranging in a gas turbine blade wall, a slit which is communicated with a cavity inside the blade, formed slantingly against a blade surface, and extended in the height direction of the blade. SOLUTION: A gas turbine stator blade comprises a blade part 1 and a shroud 2. A slit 3 is arranged over the whole height of the blade in the blade part 1. The wake side of the stator blade is cooled by flowing cooling air and so on through the slit 3. A cavity for flowing the cooling air and so on is arranged inside the blade, and this cavity is partitioned by a rib 4. The rib 4 is arranged in the vicinity of the slit 3, so that blade walls in front and rear of the slit 3 may not be deformed. As to a blade wall part 8 on a rear edge side further from the slit 3, a rib 4' is arranged just behind the slit 3, and the length of the blade wall part 8 is shortened. Hereby, a deformational problem can be solved.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

The present invention belongs to the field of gas turbine blade cooling technology.

[0002]

2. Description of the Related Art Conventionally, there has been a strong demand for a high turbine inlet temperature in order to realize a highly efficient operation of a gas turbine. Those that reach are also being developed. As the turbine inlet temperature increases, various methods and structures for cooling turbine blades have been devised.

FIG. 2 shows the structure of a conventional gas turbine stationary blade. In FIG. 1A, the stationary blade includes a wing portion 01 and a shroud 02. FIG. 2 (b)
The perspective view when the wing part 01 is cut | disconnected perpendicularly to the height direction of the wing is shown. Inside the wing portion 01, a cavity for flowing air or the like to cool the wing from the inside is provided, and a rib 04 for partitioning the cavity is provided.

[0004] Downstream of the wing portion 01, a plurality of cooling holes 06 communicating with a downstream cavity 05 partitioned by a rib 04 are provided to be inclined toward the downstream side of the wing. And cavity 05
The cooling air flowing out of the wing through the cooling holes 06 from
The cooling hole 06 flows along the blade surface due to the inclination thereof, thereby compensating for insufficient cooling near the blade trailing edge.

[0005]

In the conventional blade cooling structure, since the cooling holes 06 are provided at a certain inclination angle toward the downstream side of the blades, as shown in FIG. As a result, an edge portion 07 having an acute angle is formed, and a remarkable stress is concentrated on the edge portion 07 as compared with the inside of the blade base material. Incidentally, the degree of stress concentration in the edge portion when the cooling hole 06 is perpendicular to the blade surface and in the edge portion when the angle between the cooling hole 06 and the blade surface is 30 degrees are as follows. The former is three times higher and the latter is 6.85 times higher than the case where no is provided.

The high stress concentration on the sharp edge portion 07 causes local cracks in the edge portion 07, and once a crack occurs, the crack easily spreads to other cooling holes 06. This may cause a reduction in the life of the gas turbine blade.

[0007]

SUMMARY OF THE INVENTION In order to solve the above-mentioned problems, the present invention communicates with a wing internal cavity, is formed to be inclined with respect to the wing surface, and extends in the height direction of the wing. The slit is provided on the gas turbine blade wall.

[0008] A rib inside the blade is provided immediately before or immediately after the slit.

As described above, since the slit is formed over the entire height of the wing, the wing surface becomes a free boundary surface having no structural discontinuity, so that the disturbance of the stress flow is eliminated. Therefore, the stress in the mainstream gas flow direction at the slit end portion of the blade surface is 0, and the stress in the blade height direction at the same portion is 1, so that no concentration of stress occurs.

[0010]

FIG. 1 shows an embodiment of a blade cooling structure according to the present invention. In FIG. 1A, the gas turbine stationary blade is constituted by a blade portion 1 and a shroud 2. The wing 1 is provided with a slit 3 over the entire height of the wing. The slit 3 is provided along a portion where the plurality of cooling holes 6 are formed in the conventional stationary blade, and is inclined at an angle similar to that of the conventional cooling hole 6 with respect to the blade wall. Then, cooling air or the like flows through the slit 3 to cool the downstream side of the stationary blade.

FIG. 1B is a perspective view when the wing 1 is cut in the gas flow direction. As in the case of the conventional blade, there is a cavity for flowing cooling air and the like inside the blade, and the cavity is partitioned by a rib 4. The ribs 4 are used to prevent the wing walls before and after the slit 3 from being deformed.
Is provided in the vicinity of. With respect to the wing wall portion 8 on the trailing edge side of the slit 3, the above-mentioned deformation problem can be solved by installing the rib 4 'immediately after the slit 3 or reducing the length.

Basically, since the blade wall has a blade thickness and structure capable of sufficiently withstanding gas pressure, there is no problem in strength even if the slit 3 is provided over the entire height of the blade.

[0013]

According to the present invention, the degree of stress concentration on the sharp edge portion in the opening is reduced, and the blade wall is less likely to crack. Therefore, the soundness of the blade is improved, and the life of the blade can be extended.

[Brief description of the drawings]

FIG. 1 is a diagram showing a cooling structure of a gas turbine stationary blade according to an embodiment of the present invention.

FIG. 2 is a cooling structure diagram of a conventional gas turbine stationary blade.

[Explanation of symbols]

 01,1 wing part 02,2 shroud 3 slit 04,4,4 'rib 05,5 cavity 06 cooling hole 07,7 sharp edge part 8 trailing edge wing wall part

Claims (2)

[Claims] 1. A gas turbine, wherein a slit is formed in a gas turbine blade wall, the slit being formed in communication with a blade internal cavity, inclined with respect to the blade surface, and extending in the height direction of the blade. Wing cooling structure. 2. The gas turbine blade cooling structure according to claim 1, wherein a rib inside the blade is provided immediately before or immediately after the slit.