CN112081632A - Turbine stator blade of gas turbine and gas turbine adopting same - Google Patents
Turbine stator blade of gas turbine and gas turbine adopting same Download PDFInfo
- Publication number
- CN112081632A CN112081632A CN202011114521.1A CN202011114521A CN112081632A CN 112081632 A CN112081632 A CN 112081632A CN 202011114521 A CN202011114521 A CN 202011114521A CN 112081632 A CN112081632 A CN 112081632A
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- endwall
- trailing edge
- cavity
- blade
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- 238000001816 cooling Methods 0.000 claims abstract description 83
- 238000007789 sealing Methods 0.000 claims description 5
- 238000005266 casting Methods 0.000 claims description 4
- 238000003754 machining Methods 0.000 claims description 4
- 229910003460 diamond Inorganic materials 0.000 claims description 2
- 239000010432 diamond Substances 0.000 claims description 2
- 238000000034 method Methods 0.000 claims description 2
- 238000000465 moulding Methods 0.000 claims description 2
- 238000003466 welding Methods 0.000 claims description 2
- 239000007789 gas Substances 0.000 description 33
- 238000002347 injection Methods 0.000 description 7
- 239000007924 injection Substances 0.000 description 7
- 238000002679 ablation Methods 0.000 description 5
- 230000000052 comparative effect Effects 0.000 description 5
- 230000003647 oxidation Effects 0.000 description 5
- 238000007254 oxidation reaction Methods 0.000 description 5
- 238000002485 combustion reaction Methods 0.000 description 4
- 230000000694 effects Effects 0.000 description 4
- 239000000243 solution Substances 0.000 description 4
- 239000000567 combustion gas Substances 0.000 description 3
- 238000005516 engineering process Methods 0.000 description 3
- IJGRMHOSHXDMSA-UHFFFAOYSA-N nitrogen Substances N#N IJGRMHOSHXDMSA-UHFFFAOYSA-N 0.000 description 2
- 229910052757 nitrogen Inorganic materials 0.000 description 2
- 238000005728 strengthening Methods 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000005336 cracking Methods 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000009826 distribution Methods 0.000 description 1
- 230000007613 environmental effect Effects 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000005192 partition Methods 0.000 description 1
Images
Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/12—Cooling
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/10—Stators
- F05D2240/12—Fluid guiding means, e.g. vanes
- F05D2240/122—Fluid guiding means, e.g. vanes related to the trailing edge of a stator vane
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A turbine stator blade of a gas turbine and the gas turbine adopting the same are provided, wherein the turbine stator blade comprises a blade body and end walls positioned at two ends of the blade body; the outer surface of the blade body of the blade consists of a suction surface and a pressure surface, and the junction areas of the suction surface and the pressure surface are the front edge and the tail edge of the blade respectively; the end wall part corresponding to the tail edge of the blade is defined as the tail edge of the end wall; wherein the end wall trailing edge is provided with an end wall trailing edge cavity, which end wall trailing edge cavity is provided with at least one cold air through hole in the area corresponding to the end wall trailing edge. The invention provides an endwall trailing edge cavity in the endwall trailing edge region to reduce the temperature and stress level in the endwall trailing edge region of the turbine stator blade without increasing the amount of cooling air.
Description
Technical Field
The invention relates to the technical field of gas turbine design, in particular to a turbine stator blade of a gas turbine and the gas turbine adopting the turbine stator blade.
Background
With the increasing level of gas turbine design technology, the gas turbine inlet gas temperature is increasing continuously, and the thermal load of turbine parts is extremely high, and the limit that high-temperature materials can bear is already exceeded. In order to ensure safe and reliable operation of the turbine blade, it is necessary to design the turbine blade with a complex cooling system to maintain the temperature and stress distribution of the blade body at a reasonable level.
In order to meet the increasing environmental requirements, the DLN low-nitrogen combustion technology is generally adopted in the combustion system of the gas turbine, and the technology enables the inlet temperature profile of the gas turbine to be flatter, namely the inlet temperature peak value of the gas turbine is reduced, but the temperature level of the end wall area is increased. Therefore, in the low-nitrogen combustion gas turbine, the design difficulty of the end wall is further increased, the cooling arrangement of the tail edge area of the end wall of the turbine stator blade is difficult, and the failure phenomena such as cracks, high-temperature oxidation and even ablation are easy to occur.
Disclosure of Invention
In view of the above, the main object of the present invention is to provide a turbine stator blade of a gas turbine and a gas turbine using the same, which are intended to at least partially solve at least one of the above mentioned technical problems.
In order to achieve the purpose, the technical scheme of the invention is as follows:
as one aspect of the present invention, there is provided a turbine stator blade of a gas turbine, including a blade body, end walls at both ends of the blade body; the outer surface of the blade body of the blade consists of a suction surface and a pressure surface, and the junction areas of the suction surface and the pressure surface are the front edge and the tail edge of the blade respectively; the end wall part corresponding to the tail edge of the blade is defined as the tail edge of the end wall;
wherein the end wall trailing edge is provided with an end wall trailing edge cavity, which end wall trailing edge cavity is provided with at least one cold air through hole in the area corresponding to the end wall trailing edge.
As another aspect of the present invention, there is also provided a gas turbine employing the turbine stator blade as described above.
Based on the technical scheme, compared with the prior art, the invention has at least one or part of the following beneficial effects:
(1) according to the invention, the cavity of the tail edge of the end wall is arranged in the tail edge area of the end wall, so that the temperature and the stress level of the tail edge area of the end wall of the turbine stator blade are reduced under the condition of not increasing the cooling air quantity;
(2) the air inlet arrangement mode of the end wall tail edge cavity is various, the cooling air obtaining mode is flexible, and the realization and the application are easy;
(3) the serpentine channel is arranged in the cavity at the tail edge of the end wall to form convection cooling of a plurality of continuous cooling channels, so that the temperature, the temperature gradient and the stress level of the tail edge area of the end wall can be further reduced;
(4) and a cooling strengthening structure, such as bulges, turbulence columns and the like in various shapes, is arranged in the cavity of the tail edge of the end wall, so that the cooling strengthening effect is further enhanced.
Drawings
FIG. 1 is a schematic view of a turbine stator blade of a gas turbine of comparative example 1;
FIG. 2 is a schematic sectional view taken along line A-A of FIG. 1;
FIG. 3 is a schematic view of a turbine stator blade of a gas turbine according to embodiment 1 of the invention;
FIG. 4 is a schematic sectional view taken along line B-B of FIG. 3;
FIG. 5 is a schematic view of a turbine stator blade of a gas turbine according to embodiment 2 of the invention;
FIG. 6 is a schematic view of the C-C facing section of FIG. 5;
FIG. 7 is a turbine stator blade schematic view of a gas turbine according to embodiment 3 of the invention;
fig. 8 is a schematic view of the D-D facing portion of fig. 7.
In the above drawings, the reference numerals have the following meanings:
1-blade body; 2-an upper end wall; 3-a lower end wall; 4-an impact bushing; 5-a first impingement cooling cover plate; 6-a second impingement cooling cover plate; 7-lower end wall guide rails; 8-upper end wall guide rails; 9-a first cold air injection hole; 10-a second cold air injection hole; 11-the leading edge of the blade; 12-the trailing edge of the blade; 13-a first hook; 14-a second hook; 15-a support ring; 16-end wall trailing edge high temperature zone; 17-an upper end wall impingement cavity; 18-a lower end wall impingement cavity; 19-blade internal impingement cavities; 201. 202, 203-lower end wall trailing edge cavity entrance; 21-lower end wall trailing edge cavity; 22-lower end wall trailing edge cavity cover plate; 23-lower end wall trailing edge cavity through hole; 24-a separator; 25-a first cooling channel; 26-a second cooling channel; 27-third cooling channel.
Detailed Description
The present invention designs turbine stator vane endwall region cooling to reduce temperature and stress levels in the turbine stator vane endwall trailing edge region without increasing total cooling air volume.
In order that the objects, technical solutions and advantages of the present invention will become more apparent, the present invention will be further described in detail with reference to the accompanying drawings in conjunction with the following specific embodiments.
As one aspect of the present invention, there is provided a turbine stator blade of a gas turbine, including a blade body, end walls at both ends of the blade body; the outer surface of the blade body of the blade consists of a suction surface and a pressure surface, and the junction areas of the suction surface and the pressure surface are the front edge and the tail edge of the blade respectively; the end wall part corresponding to the tail edge of the blade is defined as the tail edge of the end wall;
wherein the tail edge of the end wall is provided with an end wall tail edge cavity, and the end wall tail edge cavity is provided with at least one cold air through hole in the area corresponding to the tail edge of the end wall.
In the embodiment of the present invention, the cross-sectional shape of the cold air passing hole includes, but is not limited to, a circle, an ellipse, a square, or a diamond; as long as the effect of the flow cooling can be achieved.
In the embodiment of the present invention, the cold air through-holes may be formed by, but not limited to, a process of casting, electro-machining or laser-machining.
In the embodiment of the present invention, the extension direction of the cold air through hole is parallel to the wall surface of the end wall; the cold air through holes comprise a plurality of cold air through holes which are arranged in an array.
More specifically, the number, shape, size and arrangement of the cooling air holes may be determined based on the actual available amount of cooling air and the cooling design requirements of the endwall trailing edge region.
In the embodiment of the invention, the interior of the blade body of the blade is cooled by air, the cooling air in the cavity at the tail edge of the end wall comes from the air after cooling the interior of the blade body of the blade, and the air after cooling the interior of the blade body of the blade is defined as the exhaust air of the cooling air of the blade body of the blade;
an end wall trailing edge cavity inlet is arranged on the end wall trailing edge and is communicated with the end wall trailing edge cavity and the interior of the blade body, and exhaust air of the blade body enters the end wall trailing edge cavity through the end wall trailing edge cavity inlet.
More specifically, an impact bushing is arranged inside the blade body of the blade, and a space between the impact bushing and the blade body of the blade is defined as an internal impact cavity of the blade; after cooling air enters the inside of the impact bushing, the blade body of the blade is subjected to impact cooling, and dead air of the cooling air of the blade body subjected to impact cooling is in an impact cavity inside the blade. The inlet of the cavity at the tail edge of the end wall is communicated with the cavity at the tail edge of the end wall and the inner impact cavity of the blade, so that the blade body cooling air in the inner impact cavity of the blade is exhausted and enters the cavity at the tail edge of the end wall.
In embodiments of the invention, the blade body may include one or more impingement bushings inside, which take the form of impingement cooling. Blade body cooling air off-air in the lower endwall trailing edge cavity may come from an impingement liner of one or more blade bodies.
In another embodiment of the invention, the endwall portion corresponding to the leading edge of the blade is defined as the endwall leading edge; providing an endwall impingement cavity at an endwall leading edge; the cooling air in the cavity at the tail edge of the end wall comes from the end wall impact cooling air exhaust in the end wall impact cavity;
an end wall guide rail is arranged on the end wall, a first through hole is arranged on the end wall guide rail, and the end wall impingement cooling air exhaust enters the cavity of the tail edge of the end wall through the first through hole of the end wall guide rail. Wherein the first through-hole becomes another arrangement of endwall impingement cavity inlets in the endwall impingement cavity, thereby providing another source of cooling air for cooling of the endwall impingement cavity.
In yet another embodiment of the present invention, an end wall guide rail is provided on the end wall, a second through hole is provided in the end wall guide rail, and compressor discharge air enters the end wall trailing edge cavity through the second through hole of the end wall guide rail. Wherein the second through-hole provides yet another arrangement of endwall impingement cavity inlets in the endwall impingement cavity, thereby providing yet another cooling air source for cooling of the endwall impingement cavity.
In an embodiment of the invention, the end wall trailing edge cavity is provided in the upper end wall and/or the lower end wall. More specifically, the endwall trailing edge cavity is not limited to being disposed on the lower endwall, but may be disposed on the upper endwall and both the upper and lower endwalls.
In an embodiment of the invention, a baffle is disposed within the cavity of the trailing edge of the end wall such that a serpentine channel is formed within the cavity of the trailing edge of the end wall.
In embodiments of the present invention, on the one hand, the serpentine channel may function to improve the cooling effect; on the other hand, the serpentine channel may also serve to restrict the gas flow direction of the cooling air within the cavity at the trailing edge of the endwall.
It is worth mentioning that the shape, size and the like of the serpentine channel can be determined according to the cooling design requirement, the serpentine channel can be in any polygon shape, and the flow area can be changed according to the actual requirement.
In embodiments of the present invention, protrusions and/or turbulators are provided within the cavity at the trailing edge of the end wall for enhanced cooling.
In the embodiment of the invention, the end wall tail edge cavity is formed by integrally casting the end wall and the end wall tail edge cavity cover plate; or the end wall tail edge cavity is formed by welding, sealing and molding the end wall and the end wall tail edge cavity cover plate.
In an embodiment of the invention, the blade body of the turbine stator blade is a single independent casting structure; or the blade body of the turbine stator blade is of an integrally cast structure with a plurality of spaced blades.
As an aspect of the present invention, there is also provided a gas turbine employing the turbine stator blade as described above.
The technical solution of the present invention is further described below with reference to specific examples and comparative examples, but it should be noted that the following examples are only for illustrating the technical solution of the present invention, but the present invention is not limited thereto.
Comparative example 1
As shown in fig. 1 and 2, this comparative example 1 provides a high-temperature turbine stator blade of a gas turbine. The high-temperature turbine stator blade structure of the gas turbine is as shown in embodiment 1 to embodiment 3, and comprises a blade body 1, an upper end wall 2, a lower end wall 3, an upper end wall guide rail 8, a lower end wall guide rail 7, a first impingement cooling cover plate 5, a second impingement cooling cover plate 6 and an impingement bushing 4. The turbine stator blade is matched with the upper end wall guide rail 8 through a first hook 13 and a second hook 14 and is fixed on the combustion engine body, and the support ring 15 is matched with the lower end wall guide rail 7 and fixes the turbine stator blade component on the support ring 15.
The cooling air cools the upper end wall 2 through the first impingement cooling cover plate 5 of the upper end wall 2, then the cooling air enters the inside of the impingement sleeve 4 through the upper end wall impingement cavity 17 and performs impingement cooling on the blade body 1, and the exhaust air after impingement cooling enters the main flow of combustion gas through the first cold air injection holes 9 in the blade inner impingement cavity 19. The cooling air of the lower end wall 3 cools the lower end wall 3 through the second impingement cooling cover plate 6, and then enters the gas side of the lower end wall 3 through the second cold air injection holes 10 in the wall surface of the lower end wall impingement cavity 18.
Here, the turbine stator vane structure of the gas turbine provided in this comparative example 1 differs from the turbine stator vane structures of examples 1 to 3 in that the endwall trailing edge high temperature region 16 is formed on the upper endwall and/or the lower endwall of the turbine stator vane. This is because: the rear part of the lower end wall guide rail 7 of the lower end wall 3 is lower due to the cold air side pressure, the rear part of the upper end wall guide rail 8 of the upper end wall 2 is also lower due to the cold air side pressure, the impact cooling cannot be arranged, and the gas side is not obvious due to the gas film covering effect of the cold air injection holes 10; resulting in the high temperature region 16 of the endwall trailing edge. This defect can cause problems with ablation, oxidation, cracking, etc. of the high temperature zone 16 at the trailing edge of the end wall.
Example 1
Fig. 3 is a schematic structural view of a turbine stator blade of a gas turbine according to embodiment 1 of the present invention, and fig. 4 is a view taken along direction B-B of fig. 3. As shown in fig. 3 and 4, a turbine stator blade of a gas turbine includes a blade body 1, an upper endwall 2, a lower endwall 3, an upper endwall rail 8, a lower endwall rail 7, a first impingement cooling cover plate 5, a second impingement cooling cover plate 6, and an impingement bushing 4. The turbine stator blade is matched with the upper end wall guide rail 8 through a first hook 13 and a second hook 14 and is fixed on the combustion engine body, and the support ring 15 is matched with the lower end wall guide rail 7 and fixes the high-temperature stator blade assembly on the support ring 15.
The cooling air cools the upper end wall 2 through the first impingement cooling cover plate 5 of the upper end wall 2, then the cooling air enters the inside of the impingement sleeve 4 through the upper end wall impingement cavity 17 and performs impingement cooling on the blade body 1, and the exhaust air after impingement cooling enters the main flow of combustion gas through the first cold air injection holes 9 in the blade inner impingement cavity 19. The cooling air of the lower end wall 3 cools the lower end wall 3 through the second impingement cooling cover plate 6, and then enters the gas side of the lower end wall 3 through the second cold air injection holes 10 in the wall surface of the lower end wall impingement cavity 18.
In the turbine stator blade provided in embodiment 1 of the present invention, the high temperature region 16 at the trailing edge of the end wall is provided with the lower end wall trailing edge cavity 21, the lower end wall trailing edge cavity 21 is formed by fixing the lower end wall trailing edge cavity cover plate 22 to the lower end part 3 in a sealing manner, and the cooling air in the lower end wall trailing edge cavity 21 comes from the blade body cooling air exhaust of the blade internal impact cavity 19 and enters the lower end wall trailing edge cavity 21 through the lower end wall trailing edge cavity inlet 201. The cooling air in the lower end wall tail edge cavity 21 cools the end wall tail edge high-temperature area 16 of the lower end wall 3 through the lower end wall tail edge cavity through hole 23, so that the temperature of the area is effectively reduced, and the problems of ablation, oxidation, cracks and the like caused by high temperature are remarkably relieved.
Example 2
Fig. 5 is a schematic structural view of a turbine stator blade of a gas turbine provided in embodiment 2 of the present invention, and fig. 6 is a sectional view taken along line C-C of fig. 5. As shown in fig. 5 and 6, embodiment 2 of the present invention is different from embodiment 1 of the present invention in the structure: the end wall trailing edge high temperature zone 16 is provided with a lower end wall trailing edge cavity 21, the lower end wall trailing edge cavity 21 is formed by fixing a lower end wall trailing edge cavity cover plate 22 on the lower end wall 3 in a sealing mode, and cooling air of the lower end wall trailing edge cavity 21 comes from end wall impingement cooling air exhaust of the lower end wall impingement cavity 18 of the lower end wall 3 and enters the lower end wall trailing edge cavity 21 through a lower end wall trailing edge cavity inlet 202 (namely, a first through hole). The cooling air in the lower end wall tail edge cavity 21 cools the end wall tail edge high-temperature area 16 of the lower end wall 3 through the lower end wall tail edge cavity through hole 23, so that the temperature of the area is effectively reduced, and the problems of ablation, oxidation, cracks and the like caused by high temperature are remarkably relieved.
Example 3
Fig. 7 is a schematic structural view of a turbine stator blade of a gas turbine provided in embodiment 3 of the present invention, and fig. 8 is a sectional view taken along line D-D of fig. 7. As shown in fig. 7 and 8, embodiment 3 of the present invention is different from embodiment 1 of the present invention in the structure:
an upper end wall tail edge cavity and a lower end wall tail edge cavity 21 are arranged in the end wall tail edge high-temperature region 16, the lower end wall tail edge cavity 21 is formed by fixing a lower end wall tail edge cavity cover plate 22 on the lower end wall 3 in a sealing mode, cooling air of the lower end wall tail edge cavity 21 comes from compressor exhaust and enters the lower end wall tail edge cavity 21 through a lower end wall tail edge cavity inlet 203 (namely a second through hole). The lower endwall trailing edge cavity 21 is divided by partitions 24 into serpentine channels comprising a first cooling channel 25, a second cooling channel 26 and a third cooling channel 27 along which cooling air is uniformly convectively cooled for the endwall region. The cooling air within the lower endwall trailing edge cavity 21 further cools the endwall trailing edge high temperature zone 16 through the lower endwall trailing edge cavity through-holes 23.
Through the cooling design, the temperature level of the tail edge area of the end wall can be effectively reduced, the temperature gradient level is reduced, and the problems of ablation, oxidation, cracks and the like caused by high temperature are remarkably relieved.
The above-mentioned embodiments are intended to illustrate the objects, technical solutions and advantages of the present invention in further detail, and it should be understood that the above-mentioned embodiments are only exemplary embodiments of the present invention and are not intended to limit the present invention, and any modifications, equivalents, improvements and the like made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (10)
1. A turbine stator blade of a gas turbine is characterized by comprising a blade body and end walls positioned at two ends of the blade body; the outer surface of the blade body of the blade consists of a suction surface and a pressure surface, and the junction areas of the suction surface and the pressure surface are the front edge and the tail edge of the blade respectively; the end wall part corresponding to the tail edge of the blade is defined as the tail edge of the end wall;
wherein the end wall trailing edge is provided with an end wall trailing edge cavity, which end wall trailing edge cavity is provided with at least one cold air through hole in the area corresponding to the end wall trailing edge.
2. The turbine stator blade according to claim 1 wherein the blade body interior is air cooled, the endwall trailing edge cavity cooling air is derived from air cooling the blade body interior, the air cooling the blade body interior is defined as blade body cooling air exhaust;
an endwall trailing edge cavity inlet is provided on the endwall trailing edge, the endwall trailing edge cavity inlet communicating the endwall trailing edge cavity with the blade body interior, the blade body cooling air exhaust entering the endwall trailing edge cavity through the endwall trailing edge cavity inlet.
3. The turbine stator blade as claimed in claim 1, wherein the corresponding endwall portion of the leading edge of the blade is defined as an endwall leading edge; providing an endwall impingement cavity at the endwall leading edge; the cooling air in the cavity at the trailing edge of the end wall comes from the end wall impingement cooling air exhaust in the end wall impingement cavity;
and an end wall guide rail is arranged on the end wall, a first through hole is arranged on the end wall guide rail, and the end wall impingement cooling air exhaust enters the cavity of the tail edge of the end wall through the first through hole of the end wall guide rail.
4. The turbine stator vane of claim 1 wherein an endwall rail is provided on the endwall, a second through-hole is provided in the endwall rail, and compressor discharge air enters the endwall trailing edge cavity through the second through-hole of the endwall rail.
5. The turbine stator blade as claimed in claim 1, wherein the endwall trailing edge cavity is provided in an upper endwall and/or a lower endwall.
6. The turbine stator vane of claim 1 wherein a baffle is disposed within the endwall trailing edge cavity such that a serpentine passage is formed within the endwall trailing edge cavity.
7. The turbine stator blade according to claim 1 wherein a protrusion and/or turbulator posts are provided in the endwall trailing edge cavity for enhanced cooling.
8. The turbine stator blade according to claim 1 wherein the cross-sectional shape of the cold gas through-hole comprises a circle, an ellipse, a square, or a diamond;
the cold air through hole is formed through the processes of casting, electric machining or laser machining.
9. The turbine stator blade as claimed in claim 1 wherein said endwall trailing edge cavity is cast integrally with said endwall and endwall trailing edge cavity cover plate; alternatively, the first and second electrodes may be,
the end wall tail edge cavity is formed by welding, sealing and molding the end wall and the end wall tail edge cavity cover plate.
10. A gas turbine, characterized in that a turbine stator blade according to any one of claims 1 to 9 is employed.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011114521.1A CN112081632A (en) | 2020-10-16 | 2020-10-16 | Turbine stator blade of gas turbine and gas turbine adopting same |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011114521.1A CN112081632A (en) | 2020-10-16 | 2020-10-16 | Turbine stator blade of gas turbine and gas turbine adopting same |
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| Publication Number | Publication Date |
|---|---|
| CN112081632A true CN112081632A (en) | 2020-12-15 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202011114521.1A Pending CN112081632A (en) | 2020-10-16 | 2020-10-16 | Turbine stator blade of gas turbine and gas turbine adopting same |
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| CN (1) | CN112081632A (en) |
Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3628880A (en) * | 1969-12-01 | 1971-12-21 | Gen Electric | Vane assembly and temperature control arrangement |
| JPH02241902A (en) * | 1989-03-13 | 1990-09-26 | Toshiba Corp | Cooling blade of turbine and combined generating plant utilizing gas turbine equipped with this blade |
| US5470198A (en) * | 1993-03-11 | 1995-11-28 | Rolls-Royce Plc | Sealing structures for gas turbine engines |
| US7452184B2 (en) * | 2004-12-13 | 2008-11-18 | Pratt & Whitney Canada Corp. | Airfoil platform impingement cooling |
| US8632298B1 (en) * | 2011-03-21 | 2014-01-21 | Florida Turbine Technologies, Inc. | Turbine vane with endwall cooling |
| US20150030461A1 (en) * | 2012-02-09 | 2015-01-29 | Siemens Aktiengesellschaft | Impingement cooling of turbine blades or vanes |
| CN213205767U (en) * | 2020-10-16 | 2021-05-14 | 北京全四维动力科技有限公司 | Turbine stator blade of gas turbine and gas turbine adopting same |
-
2020
- 2020-10-16 CN CN202011114521.1A patent/CN112081632A/en active Pending
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3628880A (en) * | 1969-12-01 | 1971-12-21 | Gen Electric | Vane assembly and temperature control arrangement |
| JPH02241902A (en) * | 1989-03-13 | 1990-09-26 | Toshiba Corp | Cooling blade of turbine and combined generating plant utilizing gas turbine equipped with this blade |
| US5470198A (en) * | 1993-03-11 | 1995-11-28 | Rolls-Royce Plc | Sealing structures for gas turbine engines |
| US7452184B2 (en) * | 2004-12-13 | 2008-11-18 | Pratt & Whitney Canada Corp. | Airfoil platform impingement cooling |
| US8632298B1 (en) * | 2011-03-21 | 2014-01-21 | Florida Turbine Technologies, Inc. | Turbine vane with endwall cooling |
| US20150030461A1 (en) * | 2012-02-09 | 2015-01-29 | Siemens Aktiengesellschaft | Impingement cooling of turbine blades or vanes |
| CN213205767U (en) * | 2020-10-16 | 2021-05-14 | 北京全四维动力科技有限公司 | Turbine stator blade of gas turbine and gas turbine adopting same |
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