CN212481376U - Front end wall cooling structure of first-stage blade of gas turbine - Google Patents
Front end wall cooling structure of first-stage blade of gas turbine Download PDFInfo
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- CN212481376U CN212481376U CN202021081293.8U CN202021081293U CN212481376U CN 212481376 U CN212481376 U CN 212481376U CN 202021081293 U CN202021081293 U CN 202021081293U CN 212481376 U CN212481376 U CN 212481376U
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- end wall
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- cooling hole
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Abstract
A cooling structure for the front end wall of the first-stage blade of a gas turbine especially relates to a cooling structure for the combustion chamber and turbine interface. The utility model provides a carry out refrigerated problem to the anterior end wall of gas turbine first order blade at present. The utility model discloses a passageway is formed between combustor changeover portion and the blade to combustor changeover portion and blade, processing has last cooling hole on the combustor upper wall of combustor changeover portion, and processing has down the cooling hole on the combustor lower wall of combustor changeover portion, the blade includes end wall under blade upper end wall and the blade, and coolant gets into main gas channel cooling blade upper end wall from last cooling hole, and coolant gets into main gas channel cooling blade lower end wall from cooling hole down. The utility model discloses form the gas mould with coolant through upper and lower cooling hole input and cover on the upper and lower end wall of blade, prevent that main gas from forming high temperature damage to upper and lower end wall of blade.
Description
Technical Field
The utility model relates to a front end wall cooling structure of a first-stage blade of a gas turbine, in particular to a cooling structure of a combustion chamber and a turbine interface.
Background
At present, for a gas turbine unit, an effective method for improving the cycle efficiency of the gas turbine unit is to improve the gas inlet temperature of a gas turbine, and an effective method for improving the initial temperature is to cool and protect high-temperature parts of the gas turbine through reasonable cooling air system design;
in order to solve the problem of the cooling of the front end wall of the first-stage blade of the gas turbine, the utility model provides a front end wall cooling structure of the first-stage blade of the gas turbine.
SUMMERY OF THE UTILITY MODEL
The utility model provides a carry out refrigerated problem to the anterior end wall of gas turbine first order blade at present, to the not enough of prior art, the utility model discloses a "anterior end wall cooling structure of gas turbine first order blade". A brief summary of the present invention is provided below in order to provide a basic understanding of some aspects of the present invention. It should be understood that this summary is not an exhaustive overview of the invention. It is not intended to identify key or critical elements of the invention or to delineate the scope of the invention.
The technical scheme of the utility model:
a cooling structure for the front end wall of a first-stage blade of a gas turbine comprises a combustion chamber transition section and the blade, wherein a channel is formed between the combustion chamber transition section and the blade;
an upper cooling hole is processed on the upper wall of the combustion chamber transition section, and a lower cooling hole is processed on the lower wall of the combustion chamber transition section;
the blade comprises a blade upper end wall and a blade lower end wall, a cooling medium enters the channel from the upper cooling hole to cool the blade upper end wall, and the cooling medium enters the channel from the lower cooling hole to cool the blade lower end wall.
Furthermore, a combustion chamber cooling hole is arranged on the combustion chamber transition section.
Furthermore, the angle formed by the upper cooling hole and the axis of the central line of the combustion engine is 50-60 degrees.
Furthermore, the angle formed by the lower cooling hole and the axis of the central line of the combustion engine is 12-18 degrees.
The utility model has the advantages that:
1. cooling medium is sprayed into the channel through the upper cooling hole and the lower cooling hole, so that high-temperature damage of main fuel gas to the upper end wall and the lower end wall of the blade is effectively isolated;
2. the angle of the cooling hole is matched with the flow characteristics of the channel, so that the consumption of high-pressure cooling air is reduced to the maximum extent, the mixing loss is reduced, and the efficiency of the gas turbine is improved;
3. the cooling holes are processed on the upper wall and the lower wall of the transition section of the combustion chamber, so that a cooling structure system for processing the turbine blade is omitted, and the complexity of the mechanism design of the blade cooling system is reduced.
Drawings
FIG. 1 is a schematic illustration of a gas turbine first stage bucket forward endwall cooling configuration.
In the figure, 1-a combustor transition section, 2-a blade, 3-an upper cooling hole, 4-a lower cooling hole, 5-a combustor upper wall, 6-a combustor lower wall, 7-a combustor cooling hole, 10-a blade upper end wall, 11-a blade lower end wall and 12-a channel.
Detailed Description
In order to make the objects, solutions and advantages of the present invention more apparent, the present invention will be described with reference to the accompanying drawings. It should be understood that the description is intended to be exemplary, and not to limit the scope of the invention. Moreover, in the following description, descriptions of well-known structures and techniques are omitted so as to not unnecessarily obscure the concepts of the present invention.
The first embodiment is as follows: the present embodiment is described with reference to fig. 1, and the cooling structure for the front end wall of the first-stage blade of the gas turbine according to the present embodiment includes a combustor transition section 1 and a blade 2, a passage 12 is formed between the combustor transition section 1 and the blade 2, and main gas flows from the combustor transition section 1 to the blade 2 and further flows to the next-stage blade;
an upper cooling hole 3 is processed on the upper wall 5 of the combustion chamber transition section 1, and a lower cooling hole 4 is processed on the lower wall 6 of the combustion chamber transition section 1;
the blade 2 comprises a blade upper end wall 10 and a blade lower end wall 11, a cooling medium enters a channel 12 from the upper cooling hole 3 to cool the blade upper end wall 10, and the cooling medium enters the channel 12 from the lower cooling hole 4 to cool the blade lower end wall 11; the cooling medium is compressed air led out from the exhaust position of the compressor, passes through the combustion chamber cylinder body, is led into the channel 12 from the plurality of upper cooling holes 3 and the lower cooling holes 4 on the upper wall 5 and the lower wall 6 of the combustion chamber, and forms an air mould to cover the upper end wall 10 and the lower end wall 11 of the blade for cooling.
The second embodiment is as follows: in the first-stage blade front end wall cooling mechanism of a gas turbine according to the present embodiment, the combustor transition piece 1 is provided with the combustor cooling holes 7, and the combustor cooling holes 7 are used for cooling the wall surface of the combustor transition piece 1.
The third concrete implementation mode: the present embodiment will be described with reference to fig. 1, in which the upper cooling hole 3 forms an angle of 50 to 60 ° with the central axis of the combustion engine in the front end wall cooling structure of the first stage blade of the gas turbine according to the present embodiment;
specifically, the angle formed by the lower cooling hole 4 and the axis of the central line of the combustion engine is 12-18 degrees; the use amount of high-pressure cooling air is reduced to the maximum extent by processing the angle formed by the upper cooling hole 3 and the axis of the central line of the combustion engine and the angle formed by the lower cooling hole 4 and the axis of the central line of the combustion engine, so that the mixing loss is reduced, and the efficiency of the gas turbine is improved.
This embodiment is only illustrative of the patent and does not limit the scope of protection thereof, and those skilled in the art can make modifications to its part without departing from the spirit of the patent.
Claims (4)
1. A gas turbine first stage blade front end wall cooling structure characterized by: the device comprises a combustion chamber transition section (1) and blades (2), wherein a channel (12) is formed between the combustion chamber transition section (1) and the blades (2);
an upper cooling hole (3) is processed on the upper wall (5) of the combustion chamber transition section (1), and a lower cooling hole (4) is processed on the lower wall (6) of the combustion chamber transition section (1);
the blade (2) comprises an upper blade end wall (10) and a lower blade end wall (11), a cooling medium enters the channel (12) from the upper cooling hole (3) to cool the upper blade end wall (10), and the cooling medium enters the channel (12) from the lower cooling hole (4) to cool the lower blade end wall (11).
2. The gas turbine first stage blade front end wall cooling structure according to claim 1, wherein: and the combustion chamber transition section (1) is provided with a combustion chamber cooling hole (7).
3. The gas turbine first stage blade front end wall cooling structure according to claim 1, wherein: the angle formed by the upper cooling hole (3) and the axis of the central line of the combustion engine is 50-60 degrees.
4. The gas turbine first stage blade front end wall cooling structure according to claim 1, wherein: the angle formed by the lower cooling hole (4) and the axis of the central line of the combustion engine is 12-18 degrees.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202021081293.8U CN212481376U (en) | 2020-06-12 | 2020-06-12 | Front end wall cooling structure of first-stage blade of gas turbine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202021081293.8U CN212481376U (en) | 2020-06-12 | 2020-06-12 | Front end wall cooling structure of first-stage blade of gas turbine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN212481376U true CN212481376U (en) | 2021-02-05 |
Family
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202021081293.8U Active CN212481376U (en) | 2020-06-12 | 2020-06-12 | Front end wall cooling structure of first-stage blade of gas turbine |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN212481376U (en) |
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2020
- 2020-06-12 CN CN202021081293.8U patent/CN212481376U/en active Active
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