CN212535776U - 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
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- CN212535776U CN212535776U CN202022245666.7U CN202022245666U CN212535776U CN 212535776 U CN212535776 U CN 212535776U CN 202022245666 U CN202022245666 U CN 202022245666U CN 212535776 U CN212535776 U CN 212535776U
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Abstract
A turbine stator blade of a gas turbine and a gas turbine using the same, the turbine stator blade comprising: the blade body comprises a pressure surface and a suction surface; wherein, a pressure surface and a suction surface surround to form a blade inner cavity; the impact flow guide bushing is arranged in the inner cavity of the blade; the area between the blade inner cavity and the impact guide sleeve is defined as an annular cavity; at least one current limiting structure arranged in the annular cavity; wherein, the impact guide bush is provided with an impact cooling jet hole, and the blade body of the blade is provided with a cold air hole; cooling air in the impact flow guide bush is subjected to impact cooling jet flow to the annular cavity through the impact cooling jet flow hole, and impact cooling exhaust air is formed and is discharged out of the blade body of the blade through the cold air hole; the flow limiting structure is used for limiting free flow of the impingement cooling exhaust gas in the annular cavity, and the impingement cooling effect is obviously improved.
Description
Technical Field
The utility model relates to a technical field of gas turbine design especially relates to a turbine stator blade of gas turbine and adopt its gas turbine.
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.
When the cooling design of the stator blade of the turbine blade of the gas turbine with the temperature of 1000-1200 ℃ (corresponding to the first stage of an E-stage gas turbine or the second stage of an F-stage gas turbine), the most common cooling design scheme is that the inside of the blade adopts impingement cooling, the outside adopts tail edge injection (or pressure surface tail edge injection), and the suction surface near the leading edge region adopts air film cooling (or leading edge spray air film cooling). The internal impingement cooling working medium needs to be discharged into main stream fuel gas from the air film holes or the jet holes on the surface of the blade, and under the condition that the number of the air film hole rows on the surface of the blade is small, characteristic parameters, positions and the like of the air film holes can obviously influence the distribution of the internal cooling working medium among the air film hole rows, so that the transverse flow of fluid on the inner surface of the blade is influenced, and the transverse flow can obviously influence the heat exchange on the inner surface of the blade. How to reasonably control the transverse flow of the cold air in the blades according to the design intention so as to have a positive effect on cooling is a problem which is concerned by cooling designers of gas turbines.
SUMMERY OF THE UTILITY MODEL
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, so as to at least partially solve at least one of the above mentioned technical problems.
In order to achieve the above object, the technical solution of the present invention includes:
as an aspect of the present invention, there is provided a turbine stator blade of a gas turbine, comprising:
a blade body comprising a pressure side and a suction side; wherein a blade inner cavity is formed by the pressure surface and the suction surface;
the impact flow guide bushing is arranged in the inner cavity of the blade; the area between the blade inner cavity and the impact guide sleeve is defined as an annular cavity;
at least one flow-limiting structure disposed within the annular cavity;
wherein, the impact guide bush is provided with an impact cooling jet hole, and the blade body of the blade is provided with a cold air hole; cooling air in the impact flow guide bush is subjected to impact cooling jet flow to the annular cavity through the impact cooling jet flow hole, and impact cooling exhaust air is formed and is discharged out of the blade body of the blade through the cold air hole; wherein the flow restricting structure is configured to restrict free flow of the impingement cooling off-gas within the annular cavity.
As another aspect of the present invention, there is also provided a gas turbine including the turbine stator blade as described above.
Based on the above technical scheme can know, compared with prior art, the utility model discloses one of them or one of them part that has following beneficial effect at least:
(1) at least one flow limiting structure is arranged in an annular cavity between the inner surface of the blade body of the blade and the impact flow guide bushing, and the flow limiting structure is tightly matched with the inner surface of the blade body of the blade and the impact flow guide bushing, so that the aim of limiting the free flow of the impingement cooling exhaust gas can be fulfilled;
(2) the impact cooling exhaust gas purposefully changes or limits the flow according to a design target, so that the impact cooling effect can be obviously improved, the convection cooling level is improved, the flow parameters in cold air holes on the surface of the blade are optimized, and the gas film covering effect on the surface of the blade is optimized, so that the optimal cooling effect and the gas turbine performance are achieved on the premise of not improving the cold air amount;
(3) the turbine blade structure is simplified and the cost is reduced.
Drawings
FIG. 1 is a schematic design of a turbine stator blade of a gas turbine of comparative example 1;
FIG. 2 is an expanded view A-A of FIG. 1;
FIG. 3 is a schematic view of a turbine stator blade design of a gas turbine according to embodiment 1 of the present invention;
FIG. 4 is an expanded view B-B of FIG. 3;
FIG. 5 is a schematic design of a turbine stator blade of a gas turbine of comparative example 2;
fig. 6 is a schematic view of a turbine stator blade design of a gas turbine according to embodiment 2 of the present invention.
In the above drawings, the reference numerals have the following meanings:
1-a suction surface; 2-pressure surface; 3-leading edge; 4-trailing edge; 5-leading edge stagnation line; 6-impacting the flow guide bush; 7-suction surface front edge air film hole; 8-pressure surface trailing edge jet hole; 9-an annular cavity; 10-impingement cooling jets; 11-impingement cooling dead gas stagnation line; 12. 13-impingement cooling the exhaust gas; 14. 17-high pressure cavity; 15-a current limiting structure; 16-a cooling chamber divider plate; 18-through hole.
Detailed Description
The utility model provides a turbine stator blade carries out rational design to turbine stator blade inside cooling channel under the condition that does not increase total cooling air volume, and control lateral flow makes its orientation development towards the direction that is favorable to cooling design to furthest's improvement blade cooling efficiency.
In order to make the objects, technical solutions and advantages of the present invention more apparent, the present invention will be described in detail with reference to the accompanying drawings.
As an aspect of the present invention, there is provided a turbine stator blade of a gas turbine, comprising:
the blade body comprises a pressure surface and a suction surface; wherein, a pressure surface and a suction surface surround to form a blade inner cavity;
the impact flow guide bushing is arranged in the inner cavity of the blade; the area between the blade inner cavity and the impact guide sleeve is defined as an annular cavity;
at least one current limiting structure arranged in the annular cavity;
wherein, the impact guide bush is provided with an impact cooling jet hole, and the blade body of the blade is provided with a cold air hole; cooling air in the impact flow guide bush is subjected to impact cooling jet flow to the annular cavity through the impact cooling jet flow hole, and impact cooling exhaust air is formed and is discharged out of the blade body of the blade through the cold air hole; wherein the flow restricting structure is adapted to restrict free flow of impingement cooling exhaust gas within the annular cavity.
In an embodiment of the present invention, the length of the flow-limiting structure in the blade height direction is less than or equal to the blade height of the blade body.
More specifically, in the annular cavity between the inner surface of the blade body and the impact flow guide bushing, the flow limiting structure can occupy the whole blade height direction and also can occupy a part of the blade height direction, and the length of the blade height direction of the flow limiting structure is adjusted according to the arrangement of the cold air holes.
In an embodiment of the present invention, the cross-sectional shape of the current limiting structure includes a rectangle, a trapezoid, or an arc.
More specifically, in the annular cavity between the inner surface of the blade body and the impact flow guide bushing, the cross-sectional shape of the flow limiting structure may be any polygon including but not limited to rectangle, trapezoid, arc segment, etc. on the premise of ensuring good sealing.
In the embodiment of the present invention, both ends of the flow-limiting structure are in sealing contact with the inner surface of the blade body and the outer surface of the impact flow-guiding bushing, respectively.
It is worth mentioning that the sealing contact means a tight fit, achieving a sealing effect.
In an embodiment of the present invention, the flow-limiting structure and the blade body are integrally cast.
In other embodiments of the present invention, the flow-limiting structure and the impact flow-guiding bushing are integrally formed.
In other embodiments of the present invention, the flow-limiting structure is formed by welding the independent structure to the blade body or the impact flow-guiding bushing.
In embodiments of the present invention, the location, number, shape, size and arrangement of the flow-limiting structures may be determined according to the inelegant needs of the cooling design profession.
The utility model discloses an in the embodiment, stator blade processing mode is precision casting, and the blade body can be single blade independent casting structure, also can be a plurality of blades integral casting structure.
In an embodiment of the present invention, the cooling air holes on the blade body include a plurality of rows of injection holes and/or film holes arranged in an array.
In an embodiment of the invention, the injection holes comprise trailing edge injection holes and/or pressure face trailing edge injection holes; the film holes comprise suction surface film holes and/or suction surface leading edge film holes.
In the embodiment of the utility model, the number of the impact flow guide bushings is one or more;
when a plurality of impact flow guide bushings are arranged in the cavity in the blade body, a cooling cavity partition plate with a through hole is arranged between every two adjacent impact flow guide bushings.
In the embodiment of the utility model, when the turbine stator blade comprises a single blade inner cavity, the blade inner cavity is provided with an impact flow guide bush, and the cold air hole comprises a tail edge jet hole and/or a pressure surface tail edge jet hole; after the cooling gas is used for carrying out impact cooling on the inner surface of the blade body of the blade through the impact flow guide bushing, the impact cooling exhaust gas is discharged into main flow fuel gas through a tail edge jet hole and/or a pressure surface tail edge jet hole. The impingement cooling exhaust gas in the annular cavity between the inner surface of the blade body and the impingement flow guide bushing is separated by the flow limiting structure arranged in the annular cavity, so that the impingement cooling exhaust gas in the annular cavity area cannot pass through the flow limiting structure, and only main flow gas can be discharged from the tail edge jet hole and/or the pressure surface tail edge jet hole according to a flow path formed by the flow limiting structure.
In the utility model discloses a preferred embodiment, when turbine stator blade includes single blade inner cavity, blade inner cavity has arranged one and has strikeed the water conservancy diversion bush, and when the air conditioning hole includes one row of suction surface leading edge air film hole and one row of pressure surface trailing edge jet orifice, the current limiting structure includes two, and one of them current limiting structure sets up between the suction surface that is close to blade trailing edge and strikeed the water conservancy diversion bush, and another one current limiting structure sets up between the pressure surface that is close to blade leading edge and strikeed the water conservancy diversion bush.
More specifically, after cooling air is used for carrying out impact cooling on the inner surface of the blade body of the blade through the impact flow guide bushing, impact cooling exhaust gas is discharged into mainstream fuel gas through at least one exhaust film hole in the surface of the blade body of the blade and a tail edge spray hole of the pressure surface. The impingement cooling exhaust gas in the annular cavity between the inner surface of the blade body and the impingement flow guide bushing is separated by at least two flow limiting structures arranged in the annular cavity, so that the impingement cooling exhaust gas in a part of the annular cavity area can be discharged from the cold air hole close to the blade body, and the impingement cooling exhaust gas in the other part of the annular cavity area can be discharged into main flow fuel gas only from the tail edge jet hole of the pressure surface.
In the preferred embodiment of the present invention, when the number of the impact guide bushings is plural, and the air conditioning hole includes one row of suction surface leading edge air film holes and one row of pressure surface trailing edge injection holes, the two flow limiting structures are provided, and the two flow limiting structures are respectively provided between the suction surface near the blade leading edge and the impact guide bushing, and between the pressure surface near the blade leading edge and the impact guide bushing.
More specifically, the blade is a plurality of blade internal cavities, and an impact guide bush is arranged in at least one blade internal cavity. After the cooling gas carries out impact cooling on the inner surface of the blade body corresponding to the cavity in the blade through the impact flow guide bush, the impact cooling exhaust gas is discharged into main flow gas through the air cooling holes arranged on the surface of the blade body, and at least one flow limiting structure is arranged in the annular cavity arranged between the blade body and the impact flow guide bush, so that the flow direction of the impact cooling exhaust gas in the annular cavity area is changed, and the main flow gas is discharged into the main flow gas from the air cooling holes arranged on the blade body.
As another aspect of the present invention, there is also provided a gas turbine including the turbine stator blade as described above.
The technical solution of the present invention is further described below with reference to the comparative example and the specific 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
FIG. 1 is a schematic design view of a turbine stator blade of a gas turbine of comparative example 1, and FIG. 2 is an expanded view of FIG. 1 along the A-A cut line. The turbine stator blade body of the gas turbine comprises a suction surface 1, a pressure surface 2, a front edge 3 and a tail edge 4; leading edge stagnation lines 5 are formed in the leading edge 3 area of the outer surface of the blade body by the high-temperature fuel gas; an impact flow guide bushing 6 is arranged inside the blade body; meanwhile, a row of suction surface front edge air film holes 7 are arranged on the suction surface 1; a row of pressure surface trailing edge jet holes 8 are arranged on the pressure surface 2 of the blade body; a high-pressure cavity 14 is formed inside the impact flow guide bush 6; an annular cavity 9 is formed between the impingement flow guiding liner 6 and the blade body.
When high-pressure cooling air in the high-pressure cavity 14 carries out impact cooling on the inner surfaces of the suction surface 1, the pressure surface 2, the front edge 3 and the tail edge 4 of the blade body through impact cooling jet flow 10 formed by impact cooling jet flow holes in the impact flow guide bush 6, the annular cavity 9 is filled with impact cooling exhaust gas, and then the high-pressure cooling air is discharged into main flow fuel gas through the air film holes 7 at the front edge of the suction surface and the jet holes 8 at the tail edge of the pressure surface. According to the mass conservation principle, the mass flow of the cooling air of the impingement cooling jet 10, the annular cavity 9, the suction surface leading edge air film hole 7 and the pressure surface trailing edge jet hole 8 should be equal. The impingement cooling exhaust gas in the annular cavity 9 is automatically divided into two parts, one part is impingement cooling exhaust gas 12 flowing to the suction surface leading edge film hole 7, the other part is impingement cooling exhaust gas 13 flowing to the pressure surface trailing edge jet hole 8, and the two parts are separated by an impingement cooling exhaust gas stagnation line 11. The location of the impingement cooling stagnation line 11 can be obtained from flow analysis and is influenced by the distribution of the impingement cooling jets 10, the geometry of the annular cavity 9, and the size and location of the suction side leading edge film holes 7 and pressure side trailing edge jet holes 8. The impingement cooling dead gas stagnation line 11 also fluctuates in time and space as conditions of the hot gas change. This introduces a great deal of uncertainty in the impingement and convection cooling design within the annular cavity 9, potentially impairing or enhancing internal cooling, thereby altering the designer's original intent; the cooling air jet flow of the suction surface leading edge film holes 7 and the pressure surface trailing edge jet holes 8 is changed, and the film covering effect of the blade body surface is changed.
The above problem is an inherent problem of the turbine stator blade design of such a conventional gas turbine, which may cause the blade cooling design to be not up to standard, and threatens the safe and stable operation of the gas turbine installed with such a turbine stator blade.
Example 1
Fig. 3 is a modified design of a turbine stator blade of a gas turbine according to embodiment 1 of the present invention, and fig. 4 is an expanded view along the B-B division line of fig. 3. The blade body of a turbine stator blade of a gas turbine comprises a suction surface 1, a pressure surface 2, a leading edge 3 and a trailing edge 4; leading edge stagnation lines 5 are formed in the leading edge 3 areas of the outer surfaces of the blade bodies of the high-temperature fuel gases; an impact flow guide bushing 6 is arranged inside the blade body; meanwhile, a row of suction surface front edge air film holes 7 are arranged on the suction surface 1; a row of pressure surface trailing edge jet holes 8 are arranged on the pressure surface 2 of the blade body; a high-pressure cavity 14 is formed inside the impact flow guide bush 6; an annular cavity 9 is formed between the impact guide bush 6 and the blade body of the blade; the flow-limiting structures 15 shown in fig. 3 and 4 are arranged in the region of the pressure surface near the leading edge and the suction surface near the trailing edge, respectively.
When high-pressure cooling air in the high-pressure cavity 14 carries out impact cooling on the inner surfaces of the suction surface 1, the pressure surface 2, the front edge 3 and the tail edge 4 of the blade body through impact cooling jet flow 10 formed by impact cooling jet flow holes in the impact flow guide bush 6, the annular cavity 9 is filled with impact cooling exhaust gas, and then the high-pressure cooling air is discharged into main flow fuel gas through a suction surface front edge air film hole 7 and a pressure surface tail edge jet hole 8. According to the mass conservation principle, the mass flow of the cooling air of the impingement cooling jet 10, the annular cavity 9, the suction surface leading edge air film hole 7 and the pressure surface trailing edge jet hole 8 should be equal. The impingement cooling exhaust gas in the annular cavity 9 is divided into two parts at the location of the flow restriction 15 by the presence of the flow restriction 15, one part being the impingement cooling exhaust gas 12 flowing to the suction side leading edge film holes 7 and the other part being the impingement cooling exhaust gas 13 flowing to the pressure side trailing edge jet holes 8, separated by the flow restriction 15. The impingement cooling dead-gas stagnation line 11, which is automatically determined according to the principle of conservation of mass, is replaced by a manually arranged flow-limiting structure 15, so that this position is no longer influenced by the distribution of the impingement cooling jets 10, the geometrical characteristics of the annular cavity 9 and the dimensions and positions of the suction side leading edge film openings 7 and the pressure side trailing edge injection openings 8. When the conditions of the hot gas change, the impingement cooling dead gas stagnation line 11 (coinciding with the flow restriction 15) no longer fluctuates in time and space. Through the arrangement of the flow limiting structure 15, the distribution proportion of the impingement cooling exhaust gas 12 flowing to the suction surface front edge air film hole 7 and the impingement cooling exhaust gas 13 flowing to the pressure surface tail edge jet hole 8 can be limited, and the flexibility of cooling design is improved.
Comparative example 2
Fig. 5 shows a comparative example of another conventional approach to the problem of single chamber impingement cooled exhaust cross flow, i.e., by the arrangement of multiple chambers. The high-pressure cavity 14 is now divided into two separate parts, the high-pressure cavity 14 and the high-pressure cavity 17, by the cooling cavity partition plate 16. There is still a single chamber impingement cooling exhaust cross flow problem within the single independent high pressure cavity 14 and high pressure cavity 17.
Example 2
Fig. 6 shows a modified embodiment to that of fig. 5. The impingement cooling exhaust gas is divided into impingement cooling exhaust gas 12 flowing to the suction surface leading edge film holes 7 and impingement cooling exhaust gas 13 flowing to the pressure surface trailing edge injection holes 8 by arranging flow-limiting structures 15 in the high-pressure cavities 14 corresponding to the leading edge 3 region, wherein the impingement cooling exhaust gas 13 flowing to the pressure surface trailing edge injection holes 8 flows into the annular cavities 9 corresponding to the trailing edge 4 region through holes 18 arranged in the cooling chamber partition plate 16.
Through the cooling design, the temperature level of the end wall area 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, further detailed description of the objects, technical solutions and advantages of the present invention, it should be understood that the above-mentioned embodiments are only specific embodiments of the present invention, and are not intended to limit the present invention, and any modifications, equivalent substitutions, improvements, etc. made within the spirit and principle of the present invention should be included in the scope of the present invention.
Claims (10)
1. A turbine stator blade for a gas turbine, comprising:
a blade body comprising a pressure side and a suction side; wherein a blade inner cavity is formed by the pressure surface and the suction surface;
the impact flow guide bushing is arranged in the inner cavity of the blade; the area between the blade inner cavity and the impact guide sleeve is defined as an annular cavity;
at least one flow-limiting structure disposed within the annular cavity;
wherein, the impact guide bush is provided with an impact cooling jet hole, and the blade body of the blade is provided with a cold air hole; cooling air in the impact flow guide bush is subjected to impact cooling jet flow to the annular cavity through the impact cooling jet flow hole, and impact cooling exhaust air is formed and is discharged out of the blade body of the blade through the cold air hole; wherein the flow restricting structure is configured to restrict free flow of the impingement cooling off-gas within the annular cavity.
2. The turbine stator blade according to claim 1,
the length of the flow limiting structure in the blade height direction is less than or equal to the blade height of the blade body.
3. The turbine stator blade according to claim 1,
the cross-sectional shape of the flow-limiting structure comprises a rectangle, a trapezoid or an arc.
4. The turbine stator blade according to claim 1,
and two ends of the flow limiting structure are in sealing contact with the inner surface of the blade body of the blade and the outer surface of the impact flow guide bushing respectively.
5. The turbine stator blade according to claim 4,
the flow limiting structure and the blade body are integrally cast and formed;
the flow limiting structure and the impact flow guide bushing are integrally processed and formed; or
And the flow limiting structure is welded with the blade body or the impact flow guide bushing.
6. The turbine stator blade according to claim 1,
the cold air holes on the blade body of the blade comprise a plurality of rows of injection holes and/or air film holes which are arranged in an array.
7. The turbine stator vane of claim 6,
the injection holes comprise a trailing edge injection hole and/or a pressure surface trailing edge injection hole;
the air film holes comprise suction surface air film holes and/or suction surface front edge air film holes.
8. The turbine stator vane of claim 7,
the number of the impact flow guide bushings is one or more;
when a plurality of impact flow guide bushings are arranged in the cavity in the blade body, a cooling cavity partition plate with a through hole is arranged between every two adjacent impact flow guide bushings.
9. The turbine stator blade according to claim 8,
when the number of the impact flow guide bushings is one, and the cold air hole comprises a row of suction surface front edge air film holes and a row of pressure surface tail edge jet holes, the number of the flow limiting structures is two, wherein one flow limiting structure is arranged between the suction surface close to the blade tail edge and the impact flow guide bushing, and the other flow limiting structure is arranged between the pressure surface close to the blade front edge and the impact flow guide bushing; when the number of impacting the guide sleeve is a plurality of, when the air conditioning hole includes one row of suction surface leading edge air film hole and one row of pressure surface tail edge jet orifice, the current-limiting structure includes two, two current-limiting structure sets up respectively between suction surface and the impact guide sleeve that is close to blade leading edge and between the pressure surface and the impact guide sleeve that are close to blade leading edge.
10. A gas turbine comprising the turbine stator blade according to any one of claims 1 to 9.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202022245666.7U CN212535776U (en) | 2020-10-10 | 2020-10-10 | Turbine stator blade of gas turbine and gas turbine adopting same |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202022245666.7U CN212535776U (en) | 2020-10-10 | 2020-10-10 | Turbine stator blade of gas turbine and gas turbine adopting same |
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| CN212535776U true CN212535776U (en) | 2021-02-12 |
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| CN202022245666.7U Active CN212535776U (en) | 2020-10-10 | 2020-10-10 | Turbine stator blade of gas turbine and gas turbine adopting same |
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