JPH0648081Y2 - gas turbine - Google Patents

Description

TECHNICAL FIELD The present invention relates to a gas turbine applied to a jet engine or the like, and more specifically to a technology for preventing crack generation due to thermal stress in a gas turbine integrated with a stationary blade, a ring and a blade ring. Regarding

2. Description of the Related Art Generally, as shown in FIG. 3, a gas turbine is configured to have a compressor 1, a combustor 2 and a turbine 3, and a common rotary shaft 4 provided in the compressor 1 and the turbine 3 The generator 5 is connected. Then, the air compressed by the compressor 1 is supplied to the combustor 2, and the fuel is supplied to the combustor 2 for combustion, and the combustion gas is introduced into the turbine 3 for work, and the generator 5 is rotated. It is designed to let you.

By the way, in such a gas turbine, in particular, a small-sized gas turbine, a compact structure is desired, and the number of parts is reduced.

As one of the means, conventionally, a stationary blade that constitutes the turbine 3, a ring that is arranged on the inner and outer circumferences of the stationary blade to form a gas passage, and a blade ring that covers the outer peripheral side of the moving blade that is provided on the downstream side of the stationary blade. Is known to be an integral structure.

4 and 5 show an example of a turbine having such a structure. That is, as shown in FIG. 4, in the gas passage 6 of the combustion gas a introduced from the combustor 2, the first-stage and second-stage stationary blades 7 and the moving blades 8 are arranged. Then, as shown in FIG. 5, the stationary blade 7 and a ring 9 arranged on the inner and outer circumferences of the stationary blade 7 to form the gas passage 6 are formed.
And a blade ring 10 that covers the outer periphery of the moving blade 8 are integrally configured.

With such a configuration, the number of parts is reduced and the turbine as a whole is made compact as compared with the assembly method of a single stationary vane.

However, as shown in FIG. 5, the trailing edge portion 7a of the stationary blade 7 is made thin so as to reduce the flow path resistance, and the plate thickness of that portion is smaller than that of each ring 9. Extremely small. Therefore, when these are integrated, the temperature of the trailing edge 7a of the stationary blade 7 rises or falls more rapidly than that of the ring 9 when the gas turbine is started or stopped.

Therefore, in the above-described conventional configuration, the stationary blade 7 and the ring 9 are
Due to a large temperature difference between the roots 7b and the ring 9 of the stationary blade 7, a very large thermal stress is generated in the roots 7b, and the roots 7b are likely to crack due to stress concentration.

Means for Solving the Problems In order to solve the problems of the prior art, the present invention provides a stationary vane, a ring disposed on the inner and outer peripheries of the stationary vane to form a gas passage, and a downstream of the stationary vane. In a gas turbine having an integral structure with a blade ring covering the outer peripheral side of a moving blade provided on the side, at least the stationary blade downstream side portion of each of the rings has a double wall structure, and ,
A slit that opens at an end edge of the ring portion is formed in a portion between the respective vanes of the ring portion on the side facing the gas passage.

By such means, the ring portion close to the trailing edge portion of the stationary blade has a double wall structure and is thin, so that the temperature difference between the trailing edge portion of the stationary blade and the ring due to the difference in plate thickness is reduced. In addition, since a slit is formed in the ring portion of the double-walled structure on the gas passage side, that is, the root portion side of the stationary blade, the temperature difference between the trailing edge portion of the stationary blade and the ring portion when the gas turbine is started or stopped. The heat expansion due to the slits is absorbed by the slits, so that the generation of thermal stress in the root portion of the stationary blade is suppressed, and cracks can be effectively prevented.

Moreover, among the double-walled rings, the slit is formed only on one of the rings on the side facing the gas passage, so the high-pressure gas generated by the compressor is the other ring that forms the double wall. It is blocked by the section and does not leak to the outside through the slit.

Embodiment An embodiment of the present invention will be described in detail below with reference to the drawings.

FIGS. 1 and 2 show an embodiment of the present invention, in which the same parts as those shown in FIGS. 3 and 4 are designated by the same reference numerals and described.

As shown in FIG. 1 and FIG. 2, also in the present embodiment, the stationary blades 7 and the moving blades 8 of the first and second stages are provided in the gas passage 6 of the combustion gas a introduced from the combustor 2. Are arranged. The stationary blade 7, the ring 9 arranged on the inner and outer circumferences of the stationary blade 7 to form the gas passage 6, and the blade ring 10 covering the outer periphery of the moving blade 8 are integrally configured.

In this structure, both ends of each ring 9 in the axial direction, that is, the upstream and downstream parts of the vane are double wall structures 11 and 12, respectively. Of the double wall structures 11 and 12, the one on the downstream side of the vane
One of the ring portions 12a on the side facing the gas passage 6 in 12
A slit 13 having a constant length is formed between the stationary vanes 7 and is open at the edge of the ring portion 12a. In addition, at the toe of each slit 13, the slit 13
A circular hole 14 having a diameter larger than the width is formed so as to communicate with each other.

No slit is formed in the other ring portion 12b on the side not facing the gas passage 6, and it is connected to the blade ring 10. The above configuration is common to the first and second stage vane portions.

According to the above configuration, the ring portion close to the trailing edge portion 7a of the stationary blade 7 has a double wall structure and is thin. Therefore, the temperature difference between the trailing edge portion 7a of the stationary blade and the ring 9 due to the difference in plate thickness. Is reduced. Since the slit 13 and the circular hole 14 are formed in the ring portion 12a of the double wall structure 12 on the gas passage 6 side, that is, on the stationary blade root 7b side, the stationary blade trailing edge is formed when the gas turbine is started or stopped. The thermal expansion due to the temperature difference between the portion 7a and the ring portion 12a is absorbed by the slit 13 and the circular hole 14, so that the generation of thermal stress in the stationary blade root portion 7b is suppressed and cracks can be effectively prevented. In particular, the structure in which the circular hole 14 is provided at the toe of the slit 13 effectively causes the stress dispersion.

Moreover, since the slit 13 of the double wall structure 12 is formed only on one of the ring portions 12a on the side facing the gas passage 6, the high pressure gas generated by the compressor forms a double wall. It is closed by the other ring portion 12b and does not leak to the outside through the slit 13.

Further, since the stationary blade 7, the ring 9 and the blade ring 10 are still integrally formed, the advantage of reducing the number of parts is maintained.

In the above-mentioned embodiment, both the front and rear ends of the ring 9 are the double wall structure portions 11 and 12, but at least the rear edge side portion can have the double wall structure to achieve the intended purpose. .

Further, a circular hole 14 is formed at the toe of the slit 13, but this is the most effective example in terms of preventing cracks from occurring.
Can be omitted in some cases.

Effect of the Invention As described above, according to the present invention, in the structure of the stationary blade, the ring and the blade ring integrated structure, the trailing edge of the ring has a double wall structure and the gas passage side where thermal stress is likely to occur By providing a slit in the, to prevent the occurrence of excessive stress due to the absorption of the difference in thermal expansion at the root portion to the ring of the vane, and thus to prevent the occurrence of cracks, without gas leakage, and without obstructing the compact structure, A gas turbine that can be effectively achieved is provided.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a perspective view showing an embodiment of the present invention, FIG. 2 is a longitudinal sectional view of a central portion showing the embodiment, and FIG. 3 is a general configuration example of a gas turbine. FIG. 4 is a systematic diagram, FIG. 4 is a longitudinal sectional view of a central portion showing a conventional example, and FIG. 6 ... gas passage, 7 ... stator blade, 7a ... stator blade trailing edge, 8 ...
… Motor blades, 9 …… rings, 10 …… blade rings, 13 …… slits.

Claims (1)

[Scope of utility model registration request] 1. A stator blade, a ring disposed on the inner and outer circumferences of the stator blade to form a gas passage, and a blade ring covering the outer peripheral side of a rotor blade provided on the downstream side of the stator blade, are integrally structured. In the gas turbine, at least the vane downstream side portion of each ring has a double-wall structure, and, in the ring of the double-wall structure, at each interblade portion of the ring portion on the side facing the gas passage, A gas turbine characterized in that a slit opening at the edge of the ring portion is formed.