CN114396316A - Turbine stator blade and turbine stator blade - Google Patents
Turbine stator blade and turbine stator blade Download PDFInfo
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- CN114396316A CN114396316A CN202111564331.4A CN202111564331A CN114396316A CN 114396316 A CN114396316 A CN 114396316A CN 202111564331 A CN202111564331 A CN 202111564331A CN 114396316 A CN114396316 A CN 114396316A
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- 238000001816 cooling Methods 0.000 claims abstract description 170
- 238000005192 partition Methods 0.000 claims description 25
- 238000003776 cleavage reaction Methods 0.000 claims 2
- 230000007017 scission Effects 0.000 claims 2
- 230000000694 effects Effects 0.000 abstract description 11
- 238000005516 engineering process Methods 0.000 description 10
- 239000011120 plywood Substances 0.000 description 8
- 238000010146 3D printing Methods 0.000 description 6
- 238000004519 manufacturing process Methods 0.000 description 5
- 230000000712 assembly Effects 0.000 description 4
- 238000000429 assembly Methods 0.000 description 4
- 239000000654 additive Substances 0.000 description 3
- 230000000996 additive effect Effects 0.000 description 3
- 230000001788 irregular Effects 0.000 description 3
- 238000002485 combustion reaction Methods 0.000 description 2
- 238000005495 investment casting Methods 0.000 description 2
- 239000000463 material Substances 0.000 description 2
- 238000003466 welding Methods 0.000 description 2
- 230000004075 alteration Effects 0.000 description 1
- 238000004891 communication Methods 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 239000000446 fuel Substances 0.000 description 1
- 230000004927 fusion Effects 0.000 description 1
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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
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention discloses a turbine stationary blade and a turbine stationary blade, which comprise a blade body, wherein the blade body is provided with a front edge side and a rear edge side, a first cavity and a second cavity are arranged in the blade body, the first cavity and the second cavity are sequentially arranged along the direction from the front edge side to the rear edge side, the rear edge side of the blade body is provided with a split seam, the second cavity is communicated with the outer side of the blade body through the split seam, the blade body is provided with a plurality of air film cooling holes, and the first cavity and the second cavity are communicated with the outer side of the blade body through the air film cooling holes. The turbine stationary blade of the invention ensures high inlet temperature, good cooling effect and high cooling efficiency of the gas turbine.
Description
Technical Field
The invention relates to the technical field of heavy-duty gas turbines, in particular to a turbine stator blade for a gas turbine and a turbine stator blade applying the turbine stator blade.
Background
The gas turbine is an internal combustion type power machine for converting energy of fuel into useful work, and in order to improve the working efficiency of the gas turbine, it is an effective scheme to improve the inlet temperature of the gas turbine.
Disclosure of Invention
The present invention is directed to solving, at least to some extent, one of the technical problems in the related art.
Therefore, the embodiment of the invention provides the turbine stator blade which is high in inlet temperature, good in cooling effect and high in cooling efficiency.
The embodiment of the invention also provides a turbine stator blade applying the turbine stator blade.
The turbine stationary blade comprises a blade body, wherein the blade body is provided with a front edge side and a rear edge side, a first cavity and a second cavity are arranged in the blade body, the first cavity and the second cavity are sequentially arranged along the direction from the front edge side to the rear edge side, a cleft is arranged on the rear edge side of the blade body, the second cavity is communicated with the outer side of the blade body through the cleft, the blade body is provided with a plurality of air film cooling holes, and the first cavity and the second cavity are communicated with the outer side of the blade body through the air film cooling holes.
First cooling module, first cooling module locates the inner wall of blade shaft, first chamber the intracavity of second is equipped with first cooling module, first cooling module includes plywood and a plurality of post rib, and is a plurality of the post rib is connected the plywood with between the inner wall of blade shaft, the plywood with form between the inner wall of blade shaft and strike the chamber, be equipped with the impingement cooling hole on the plywood, first chamber the cooling air in the second intracavity be suitable for through strike the cooling hole strike the chamber the air film cooling hole flows the blade shaft.
A second cooling assembly disposed on an inner wall of the blade body and located within the second cavity, the second cooling assembly located behind the first cooling assembly, the second cooling assembly including a plurality of first ribs and a plurality of second ribs, the plurality of first ribs being arranged in parallel at intervals, the plurality of second ribs being arranged in parallel at intervals, the first ribs and the second ribs being arranged in a cross-over manner.
The turbine stationary blade provided by the embodiment of the invention has the advantages of high inlet temperature, good cooling effect and high cooling efficiency.
In some embodiments, the blade body includes a pressure section and a suction section which are oppositely arranged, one side of the pressure section and one side of the suction section are connected to form a leading edge side, the other side of the pressure section and the other side of the suction section are parallelly arranged at intervals to form a trailing edge side, the first cavity, the second cavity and the cleft are located between the pressure section and the suction section, the outer surface of the pressure section forms a pressure surface, the outer surface of the suction section forms a suction surface, and the first cooling assembly and the second cooling assembly are arranged on the inner wall of the pressure section and/or the inner wall of the suction section.
In some embodiments, the first chamber is provided in plurality, the first chambers are sequentially arranged in the direction from the leading edge side to the trailing edge side, and the first chambers are provided in plurality on the front side of the second chamber.
In some embodiments, the plurality of first cavities include a first sub-cavity and a second sub-cavity, the second sub-cavity is located between the first sub-cavity and the second cavity, a first partition plate and a second partition plate are arranged in the blade body, the first partition plate is arranged between the first sub-cavity and the second sub-cavity, and the second partition plate is arranged between the second sub-cavity and the second cavity.
In some embodiments, the airfoil has a mean camber line connecting between the leading edge side and the trailing edge side, and any point on the mean camber line is spaced from the pressure section by an equal distance as the point is spaced from the suction section.
In some embodiments, a ratio of a mean camber line length between the first diaphragm and the leading edge side to a total length of the mean camber line is 20% to 40%, and a ratio of a mean camber line length between the second diaphragm and the leading edge side to a total length of the mean camber line is 40% to 80%.
In some embodiments, a second cooling assembly is disposed on the pressure section and the suction section, the cleft is formed between the second cooling assembly of the pressure section and the second cooling assembly of the suction section, and a ratio of a width dimension of the cleft to a height dimension of the second cooling assembly is no more than 1.
In some embodiments, the camber line has a mounting section, a ratio of a camber line length between any point on the mounting section and the leading edge side to a total length of the camber line is 50% to 95%, and the second cooling unit is provided at a position of the blade body portion corresponding to the mounting section.
In some embodiments, the extending direction of the first rib and the cross section of the blade body form an included angle α, the extending direction of the second rib and the cross section of the blade body form an included angle β, and the included angle α and the included angle β are 30 degrees to 160 degrees.
In some embodiments, the ratio of the distance between two adjacent first ribs to the width dimension of the first ribs is 4 to 15, and the ratio of the distance between two adjacent second ribs to the width dimension of the second ribs is 4 to 15.
In some embodiments, the first cooling assembly has a spanwise direction in which a ratio of a pitch of adjacent two impingement cooling holes to a diameter of the impingement cooling holes is 4 to 10, a ratio of a pitch of adjacent two column ribs to a diameter of the column ribs is 2.5 to 5, and a flow direction in which a ratio of a pitch of adjacent two impingement cooling holes to a diameter of the impingement cooling holes is 4 to 12, and a ratio of a pitch of adjacent two column ribs to a diameter of the column ribs is 2.5 to 10.
In some embodiments, a ratio of a spacing of the plies and an inner wall of the airfoil to the impingement cooling holes is from 1 to 3.
The turbine stator blade of the embodiment of the invention comprises a blade, an upper end wall and a lower end wall, wherein the blade is connected between the upper end wall and the lower end wall, and the blade is the turbine stator blade in any embodiment.
Drawings
FIG. 1 is a schematic structural view of a turbine vane of an embodiment of the present invention.
FIG. 2 is a schematic view of a blade internal cooling configuration of an embodiment of the present invention.
Fig. 3 is a schematic structural view of a first cooling module according to an embodiment of the present invention.
FIG. 4 is a schematic view of an expanded configuration of a laminate according to an embodiment of the present invention.
Fig. 5 is a schematic cross-sectional view of fig. 2 taken in the transverse direction at B.
Fig. 6 is a schematic cross-sectional view of fig. 2 taken along the vertical direction B.
FIG. 7 is a cross-sectional line schematic view of a blade of an embodiment of the present invention.
FIG. 8 is a schematic flow diagram of cooling air within a bucket according to an embodiment of the present invention.
Fig. 9 is a partially enlarged schematic view of a second cooling assembly of an embodiment of the present invention.
Reference numerals:
a turbine vane 100; a blade 1;
a blade body 11; a first cavity 111; a first sub-cavity 1111; a second subchamber 1112; a second chamber 112; a slit 113; film cooling holes 114; a pressure section 115; a suction section 116;
a first cooling assembly 12; a laminate 121; the column ribs 122; an impingement cavity 123; impingement cooling holes 124;
a second cooling assembly 13; the first ribs 131;
a first separator 2; a second separator 3; a mean camber line 4; an upper end wall 5; a lower end wall 6; .
Detailed Description
Reference will now be made in detail to embodiments of the present invention, examples of which are illustrated in the accompanying drawings. The embodiments described below with reference to the drawings are illustrative and intended to be illustrative of the invention and are not to be construed as limiting the invention.
As shown in fig. 1 and 2, a turbine stator blade according to an embodiment of the present invention includes a blade body 11, the blade body 11 has a leading edge side and a trailing edge side, a first cavity 111 and a second cavity 112 are provided in the blade body 11, the first cavity 111 and the second cavity 112 are sequentially provided in a direction from the leading edge side to the trailing edge side, a split gap 113 is provided on the trailing edge side of the blade body 11, and the split gap 113 connects the second cavity 112 with an outer side of the blade body 11.
It should be noted that in the embodiment of the present invention, the leading edge side of the blade body 11 is the front side of the blade body 11, the trailing edge side of the blade body 11 is the rear side of the blade body 11, the inner direction of the blade body 11 is defined as the inner side, and the outer direction of the blade body 11 is defined as the outer side.
Specifically, as shown in fig. 2, the inner wall surface and the outer wall surface of the blade body 11 may be arc surfaces, the cross section of the blade body 11 may be in an airfoil shape, the inside of the body of the blade body 11 is a hollow structure, the inside of the blade body 11 may be divided into a first cavity 111 and a second cavity 112 by a partition plate disposed between the leading edge side and the trailing edge side, a split gap 113 is disposed at the trailing edge side of the blade body 11 and communicates the second cavity 112 with the outside of the blade body 11, and when cooling is performed, the split gap 113 is used for allowing part of cooling air in the second cavity 112 to flow out. As shown in FIG. 6, the blade body 11 is provided with a plurality of film cooling holes 114, and the film cooling holes 114 communicate the first and second cavities 111, 112 with the outside of the blade body 11.
Specifically, the cross section of the film cooling hole 114 may be square or circular, a plurality of film cooling holes 114 are formed in the blade body 11, the film cooling holes 114 may be disposed in the wall of the blade body 11 to be inclined with the inner wall or the outer wall of the blade body 11, the film cooling holes 114 may be uniformly arranged in the wall of the blade body 11, the film cooling holes 114 communicate the first cavity 111 and the second cavity 112 with the outer side of the blade body 11, the film cooling hole 114 in the first cavity 111 is used for flowing out all of the cooling air, and the film cooling hole 114 in the second cavity 112 is used for flowing out of part of the cooling air.
It is understood that in other embodiments, film cooling holes 114 in the wall of blade body 11 may be irregularly spaced.
Specifically, as shown in fig. 2 to 6, the first cooling module 12 is disposed on the inner wall of the blade body 11 in the first cavity 111 and the second cavity 112, the first cooling module 12 is composed of a laminate 121 and a plurality of column ribs 122, the curvature of the laminate 121 may be consistent with the curvature of the inner wall surface of the first cavity 111 or the second cavity 112, the plurality of column ribs 122 may be uniformly arranged between the laminate 121 and the inner wall of the blade body 11, the column ribs 122 may be cylindrical or prismatic, the laminate 121, the column ribs 122 and the blade body 11 may be integrally formed by 3D printing, an impingement cavity 123 is formed between the laminate 121 and the inner wall of the blade body 11, a plurality of impingement cooling holes 124 are disposed on the laminate 121, the cross section of the impingement cooling hole 124 may be circular or square, the plurality of impingement cooling holes 124 may be uniformly arranged on the laminate 121, and the cooling air in the first cavity 111 is suitable for passing through the impingement cooling hole 124, the impingement cavity 123, and the impingement cavity 11, The film cooling holes 114 exit the airfoil 11, and a portion of the cooling air within the second chamber 112 is adapted to exit the airfoil 11 through the impingement cooling holes 124, the impingement chamber 123, the film cooling holes 114, and another portion exits the airfoil 11 at the location of the cleave gap 113.
It is understood that in other embodiments, the layer plates 121 may have other regular or irregular plate surface shapes, and accordingly, the length of the column ribs 122 may be different due to different distances between the layer plates 121 and the inner wall of the blade body 11 at different connection positions, the column ribs 122 between the layer plates 121 and the inner wall of the blade body 11 may be arranged at irregular intervals, and the impingement cooling holes 124 on the layer plates 121 may be arranged at irregular intervals.
And a second cooling assembly 13, wherein the second cooling assembly 13 is arranged on the inner wall of the blade body 11 and positioned in the second cavity 112, the second cooling assembly 13 is positioned at the rear side of the first cooling assembly 12, the second cooling assembly 13 comprises a plurality of first ribs 131 and a plurality of second ribs, the plurality of first ribs 131 are arranged in parallel at intervals, the plurality of second ribs are arranged in parallel at intervals, and the first ribs 131 and the second ribs are arranged in a cross manner.
Specifically, as shown in fig. 2 and 9, the second cooling block 13 is disposed at the inner wall of the second cavity 112, and the second cooling block 13 is located at the rear side of the first cooling block 12 in the second cavity 112, the second cooling block 13 is composed of a plurality of first ribs 131 and a plurality of second ribs, the plurality of first ribs 131 may be arranged in parallel at intervals, the plurality of second ribs may also be arranged in parallel at intervals, the first ribs 131 and the second ribs may be arranged in a cross, and the first ribs 131 and the second ribs may be integrally formed with the blade body 11 through 3D printing.
It is understood that in other embodiments, the plurality of first ribs 131 and the plurality of second ribs may be arranged at intervals that are not parallel.
In the turbine vane blade according to the embodiment of the present invention, the impingement cooling technology of the laminate 121 adopted by the first cooling assembly 12 has a higher cooling performance than the liner impingement film cooling technology applied to the heavy-duty gas turbine in the related art, and the staggered rib cooling technology adopted by the second cooling assembly 13 also has a higher cooling performance than the cooling technology of the column rib 122 or other turbulent flow structures applied to the heavy-duty gas turbine in the related art, and the fusion of the first cooling assembly and the second cooling assembly can achieve a better cooling effect on one hand, so that the re-combustion turbine can meet the use requirement of a higher inlet temperature, and on the other hand, under the condition of the same cooling efficiency, a smaller amount of cooling air can be consumed, so that the power efficiency of the gas turbine can be improved, and an economic value can be created.
It should be noted that the first cooling assembly 12, the second cooling assembly 13 and the blade airfoil 11 can be formed by 3D printing additive manufacturing technology, and by using the feature that the additive manufacturing technology has less limitation on the cooling structure compared with the conventional investment casting, a more efficient and complex cooling structure can be applied compared with the conventional investment casting technology, and the manufacturing cost is not increased, which makes the heavy-duty gas turbine more economical.
In some embodiments, the airfoil 11 includes a pressure section 115 and a suction section 116 disposed opposite one another, one side of the pressure section 115 and one side of the suction section 116 are connected and form a leading edge side, the other side of the pressure section 115 and the other side of the suction section 116 are spaced apart in parallel and form a trailing edge side, the first cavity 111, the second cavity 112, the cleft 113 are located between the pressure section 115 and the suction section 116, an outer surface of the pressure section 115 forms a pressure surface, an outer surface of the suction section 116 forms a suction surface, and the first cooling assembly 12 and the second cooling assembly 13 are disposed on an inner wall of the pressure section 115 and/or an inner wall of the suction section 116.
Specifically, as shown in FIG. 2, the blade airfoil 11 includes a pressure segment 115 and a suction segment 116 arranged oppositely, the pressure segment 115 and the suction segment 116 are arranged oppositely in a left-right direction, a front side of the pressure segment 115 and a front side of the suction segment 116 are connected to form a front edge side, a rear side of the pressure segment 115 and a rear side of the suction segment 116 are arranged in parallel at intervals to form a rear edge side, a first cavity 111, a second cavity 112, and a cleft 113 are sequentially provided between the pressure segment 115 and the suction segment 116 from the front side to the rear side of the blade airfoil 11, the first cooling assembly 12 and the second cooling assembly 13 may be provided on an inner wall of the pressure segment 115 and an inner wall of the suction segment 116 at the same time, or one of the arrangements may be selected on the inner wall of the pressure section 115 and on the inner wall of the suction section 116, and, the pressure section 115, the suction section 116, the first cooling assembly 12 and the second cooling assembly 13 can be formed by selecting a 3D printing additive manufacturing technology.
In some embodiments, the first cavities 111 are plural, the plural first cavities 111 are sequentially arranged in a direction from the leading edge side to the trailing edge side, and the plural first cavities 111 are arranged at the front side of the second cavity 112.
Specifically, as shown in fig. 2, a plurality of first cavities 111 may be included in the blade body 11, for example, 2 first cavities 111 are included, and 2 first cavities 111 are sequentially disposed along a direction from the leading edge side to the trailing edge side, wherein the first cavity 111 on the rear side is located on the front side of the second cavity 112, and the plurality of first cavities 111 are disposed, so that the cooling effect may be improved.
It will be appreciated that in some other embodiments, 3, 4, and other numbers of first cavities 111 may also be included within the main body 11.
In some embodiments, the plurality of first cavities 111 includes a first sub-cavity 1111 and a second sub-cavity 1112, the second sub-cavity 1112 is located between the first sub-cavity 1111 and the second cavity 112, a first separator 2 and a second separator 3 are disposed within the blade body 11, the first separator 2 is disposed between the first sub-cavity 1111 and the second sub-cavity 1112, and the second separator 3 is disposed between the second sub-cavity 1112 and the second cavity 112.
Specifically, as shown in fig. 2, two first cavities 111 are provided in the blade body 11, the two first cavities 111 are a first sub-cavity 1111 and a second sub-cavity 1112 respectively, the second sub-cavity 1112 is located at the rear side of the first sub-cavity 1111, and the second sub-cavity 1112 is located at the front side of the second cavity 112, the first sub-cavity 1111 and the second sub-cavity 1112 are separated by a first partition plate 2, the second sub-cavity 1112 and the second cavity 112 are separated by a second partition plate 3, wherein the first partition plate 2 and the second partition plate 3 can be arranged in parallel, and the first partition plate 2, the second partition plate 3 and the blade body 11 can be integrally manufactured by a 3D printing technology, so that the cost is lower and the blade is more economical.
It is understood that in other embodiments, the first partition plate 2 and the second partition plate 3 may be arranged in a non-parallel manner, and the first partition plate 2 and the second partition plate 3 may be connected to the blade body 11 by welding or the like.
In some embodiments, body 11 has a mean camber line 4, mean camber line 4 being connected between the leading edge side and the trailing edge side, and any point on mean camber line 4 is spaced from pressure segment 115 by an amount equal to the spacing between that point and suction segment 116.
Specifically, as shown in FIG. 7, the cross-section of the airfoil 11 is an airfoil in which a series of circles are drawn that are tangent to both the pressure and suction surfaces, the centers of which are connected by a mean camber line 4, the mean camber line 4 being defined to facilitate locating structural features of the first cooling module 12 and the second cooling module 13.
In some embodiments, the ratio of the length of the mean camber line 4 between the first partition 2 and the leading edge side to the total length of the mean camber line 4 is 20% to 40%, and the ratio of the length of the mean camber line 4 between the second partition 3 and the leading edge side to the total length of the mean camber line 4 is 40% to 80%.
Specifically, as shown in fig. 2, the length of the mean camber line 4 between the first partition board 2 and the leading edge side accounts for 20% to 40% of the total length of the mean camber line 4, the length of the mean camber line 4 between the second partition board 3 and the leading edge side accounts for 40% to 80% of the total length of the mean camber line 4, and this arrangement can improve the heat exchange effect in the blade body 11, for example, the length of the mean camber line 4 between the first partition board 2 and the leading edge side accounts for 20%, 25%, 30%, 35%, 38%, 40%, etc. of the total length of the mean camber line 4, and the length of the mean camber line 4 between the second partition board 3 and the leading edge side accounts for 40%, 45%, 50%, 55%, 60%, etc. of the length of the mean camber line 4.
In some embodiments, the pressure section 115 and the suction section 116 are provided with the second cooling assembly 13, the cleft 113 is formed between the second cooling assembly 13 of the pressure section 115 and the second cooling assembly 13 of the suction section 116, and the ratio of the width dimension of the cleft 113 to the height dimension of the second cooling assembly 13 is not more than 1.
Specifically, as shown in fig. 2, in the second cavity 112, the second cooling assemblies 13 are simultaneously disposed on the inner walls of the pressure section 115 and the suction section 116, the channel formed between the two second cooling assemblies 13 is the split gap 113, the first rib 131 and the second rib in the two second cooling assemblies 13 can be disposed at the same height, it should be noted that, in the implementation of the present invention, the left-right direction of the blade body 11 is defined as the height direction of the blade body 11, that is, the left-right direction of the first rib 131 and the second rib is the height of the first rib 131 and the second rib, the width of the split gap 113 is the distance between the inner side surfaces of the two second cooling assemblies 13, and the ratio of the width dimension of the split gap 113 and the height dimension of the first rib 131 (or the second rib) is not more than 1, so as to improve the heat exchange effect in the blade body 11, for example, the ratio of the width dimension of the split gap 113 and the height dimension of the first rib 131 (or the second rib) can be 0.5, 0.6, 0.7, 0.8, 0.9, etc. This can further improve the cooling effect of the blade.
In some embodiments, the camber line 4 has a mounting section, a ratio of a length of the camber line 4 between any point on the mounting section and the leading edge side to a total length of the camber line 4 is 50% to 95%, and the second cooling unit 13 is provided at a position corresponding to a portion of the blade body 11 of the mounting section.
Specifically, as shown in fig. 2, the mean camber line 4 of the second cavity 112 is provided with a mounting segment, the length of the mean camber line 4 between any point on the mounting segment and the leading edge side accounts for 50% to 95% of the total length of the mean camber line 4, and the second cooling assembly 13 is provided at a portion of the inner wall of the pressure section 115 and (or the suction section 116) of the blade body 11 corresponding to the mounting segment, so that the heat exchange effect in the blade body 11 can be effectively improved.
For example, the second cooling assembly 13 is mounted at a position corresponding to a ratio of the length of the mean camber line 4 between the front side and the leading edge side of the mean camber line 4 to the total length of the mean camber line 4 of 50%, 55%, 60%, 65%, 70%, 75%, 80%, 90%, 95%, etc., and the second cooling assembly 13 is mounted at a position corresponding to a ratio of the length of the mean camber line 4 between the rear side and the leading edge side of the mean camber line 4 to the total length of the mean camber line 4 of 90%.
In some embodiments, the first rib 131 extends at an angle α to the cross-section of the blade body 11, and the second rib extends at an angle β to the cross-section of the blade body 11, the angles α and β being 30 to 160 degrees.
Specifically, as shown in fig. 9, the angle α formed by the extending direction of the first rib 131 and the cross section of the blade body 11 is between 30 degrees and 160 degrees, and the angle β formed by the extending direction of the second rib and the cross section of the blade body 11 is between 30 degrees and 160 degrees, so that the heat exchange effect in the blade body 11 is the best, for example, the angles α and β are different and can be a combination of any value of 30 degrees, 40 degrees, 50 degrees, 55 degrees, 60 degrees, 65 degrees, 70 degrees, 100 degrees, 110 degrees, 120 degrees, 130 degrees, 140 degrees, 150 degrees, and 160 degrees.
In some embodiments, the ratio of the spacing between two adjacent first ribs 131 to the width dimension of the first ribs 131 is 4 to 15, and the ratio of the spacing between two adjacent second ribs to the width dimension of the second ribs is 4 to 15.
Specifically, as shown in fig. 9, the ratio of the pitch of two adjacent first ribs 131 arranged in parallel to the width dimension of the first rib 131 is between 4 and 15, and the ratio of the pitch of two adjacent second ribs arranged in parallel to the width dimension of the second rib is between 4 and 15, so as to achieve the best heat exchange effect in the blade 11, it should be understood that, in the embodiment of the present invention, the pitch direction of two adjacent first ribs 131 and the width direction of the first rib 131 are the same direction, and the pitch direction of two adjacent second ribs and the width direction of the second rib are the same direction, for example, the ratio of the pitch of two adjacent first ribs 131 to the width dimension of the first rib 131 may be 4, 5, 6, 7, 8, 9, 10, 11, 12, 13, 14, 15, etc., and the ratio of the pitch of two adjacent second ribs to the width dimension of the second rib may be 4, 5, 6, 7, 9, 10, 11, 12, 13, 14, 15, etc, 8. 9, 10, 11, 12, 13, 14, 15, etc.
In some embodiments, the first cooling assembly 12 has a spanwise direction and a streamwise direction, and in the streamwise direction, the ratio of the spacing of two adjacent impingement cooling holes 124 to the diameter of the impingement cooling holes 124 is 4 to 10, the ratio of the spacing of two adjacent column ribs 122 to the diameter of the column ribs 122 is 2.5 to 5, in the spanwise direction, the ratio of the spacing of two adjacent impingement cooling holes 124 to the diameter of the impingement cooling holes 124 is 4 to 12, and the ratio of the spacing of two adjacent column ribs 122 to the diameter of the column ribs 122 is 2.5 to 10.
Specifically, as shown in fig. 4 and 5, the first cooling module 12 has a spanwise direction and a flowing direction, and it is to be understood that, in the embodiment of the present invention, the up-down direction of the first cooling module 12 is defined as the spanwise direction, such as the b direction in fig. 4, the front-side to rear-side direction of the first cooling module 12 is defined as the flowing direction, such as the a direction in fig. 4, and the cross section of the impingement cooling holes 124 and the cross section of the column rib 122 are set to be circular, the impingement cooling holes 124 and the column rib 122 are both uniformly arranged, and in the flowing direction, the ratio of the spacing between two adjacent impingement cooling holes 124 and the diameter of the impingement cooling holes 124 may be between 4 and 10, and the ratio of the spacing between two adjacent column ribs 122 and the diameter of the column rib 122 may be between 2.5 and 5, such as 2.5, 3, 4, 4.5, 5, and the like.
In the spanwise direction, the ratio of the spacing between two adjacent impingement cooling holes 124 to the diameter of the impingement cooling holes 124 may be between 4 and 12, and may be, for example, 4, 5, 6, 7, 8, 9, 10, 11, 12, etc. The ratio of the distance between two adjacent ribs 122 and the diameter of the ribs 122 may be between 2.5 and 10, and for example, may be 2.5, 5, 5.5, 6, 7, 9, 9.5, 10, etc.
The impingement cooling holes 124 and the stud ribs 122 are designed using the above parameters to effectively enhance cooling performance within the airfoil 11.
In some embodiments, the ratio of the distance between the laminae 121 and the inner wall of the blade body 11 to the impingement cooling holes 124 is 1 to 3, specifically, as shown in fig. 6, in the embodiment of the present invention, the inner surface and the outer surface of the laminae 121 are scaled by the inner wall surface of the first cavity 111 (or the second cavity 112), and the ratio of the distance between the laminae 121 and the inner wall of the first cavity 111 (or the second cavity 112) to the diameter of the impingement cooling holes 124 may be 1 to 3, which may effectively improve the cooling performance inside the first cavity 111 and the second cavity 112, for example, the ratio of the distance between the laminae 121 and the inner wall of the first cavity 111 (or the second cavity 112) to the diameter of the impingement cooling holes 124 may be 1, 1.5, 2, 2.5, 3, and the like.
The following describes a turbine vane 100 according to an embodiment of the present invention.
As shown in fig. 1, a turbine vane 100 according to an embodiment of the present invention includes a blade 1, an upper endwall 5, and a lower endwall 6, the blade 1 being connected between the upper endwall 5 and the lower endwall 6, the blade 1 being a turbine vane blade as described in the above-described embodiment.
Specifically, the upper end wall 5 and the lower end wall 6 are respectively connected to the upper side and the lower side of the blade 1, and the upper end wall 5, the lower end wall 6 and the blade 1 can be integrally formed by adopting a 3D printing manufacturing technology, so that the blade is more economical. It will be appreciated that in other embodiments, the upper and lower end walls 5, 6 may also be joined to the blade 1 by welding.
In the description of the present invention, it is to be understood that the terms "central," "longitudinal," "lateral," "length," "width," "thickness," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," "clockwise," "counterclockwise," "axial," "radial," "circumferential," and the like are used in the orientations and positional relationships indicated in the drawings for convenience in describing the invention and to simplify the description, and are not intended to indicate or imply that the referenced devices or elements must have a particular orientation, be constructed and operated in a particular orientation, and are therefore not to be considered limiting of the invention.
Furthermore, the terms "first", "second" and "first" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one such feature. In the description of the present invention, "a plurality" means at least two, e.g., two, three, etc., unless explicitly specifically defined otherwise.
In the present invention, unless otherwise expressly stated or limited, the terms "mounted," "connected," "secured," and the like are to be construed broadly and can, for example, be fixedly connected, detachably connected, or integrally formed; may be mechanically coupled, may be electrically coupled or may be in communication with each other; they may be directly connected or indirectly connected through intervening media, or they may be connected internally or in any other suitable relationship, unless expressly stated otherwise. The specific meanings of the above terms in the present invention can be understood by those skilled in the art according to specific situations.
In the present invention, unless otherwise expressly stated or limited, the first feature "on" or "under" the second feature may be directly contacting the first and second features or indirectly contacting the first and second features through an intermediate. Also, a first feature "on," "over," and "above" a second feature may be directly or diagonally above the second feature, or may simply indicate that the first feature is at a higher level than the second feature. A first feature being "under," "below," and "beneath" a second feature may be directly under or obliquely under the first feature, or may simply mean that the first feature is at a lesser elevation than the second feature.
In the description herein, references to the description of the term "one embodiment," "some embodiments," "an example," "a specific example," or "some examples," etc., mean that a particular feature, structure, material, or characteristic described in connection with the embodiment or example is included in at least one embodiment or example of the invention. In this specification, the schematic representations of the terms used above are not necessarily intended to refer to the same embodiment or example. Furthermore, the particular features, structures, materials, or characteristics described may be combined in any suitable manner in any one or more embodiments or examples. Furthermore, various embodiments or examples and features of different embodiments or examples described in this specification can be combined and combined by one skilled in the art without contradiction.
Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention, and that variations, modifications, substitutions and alterations can be made to the above embodiments by those of ordinary skill in the art within the scope of the present invention.
Claims (13)
1. A turbine stator blade, comprising:
the blade body is provided with a front edge side and a rear edge side, a first cavity and a second cavity are arranged in the blade body, the first cavity and the second cavity are sequentially arranged along the direction from the front edge side to the rear edge side, a cleft is arranged on the rear edge side of the blade body, the second cavity is communicated with the outer side of the blade body through the cleft, the blade body is provided with a plurality of air film cooling holes, and the first cavity and the second cavity are communicated with the outer side of the blade body through the air film cooling holes;
the first cooling assembly is arranged on the inner wall of the blade body, the first cooling assembly is arranged in the first cavity and the second cavity and comprises a laminate and a plurality of column ribs, the column ribs are connected between the laminate and the inner wall of the blade body, an impact cavity is formed between the laminate and the inner wall of the blade body, impact cooling holes are formed in the laminate, and cooling air in the first cavity and the second cavity is suitable for flowing out of the blade body through the impact cooling holes, the impact cavity and the film cooling holes;
a second cooling assembly disposed on an inner wall of the blade body and located within the second cavity, the second cooling assembly located behind the first cooling assembly, the second cooling assembly including a plurality of first ribs and a plurality of second ribs, the plurality of first ribs being arranged in parallel at intervals, the plurality of second ribs being arranged in parallel at intervals, the first ribs and the second ribs being arranged in a cross-over manner.
2. The turbine stator blade for a gas turbine as claimed in claim 1, wherein the blade body comprises a pressure section and a suction section which are oppositely arranged, one side of the pressure section and one side of the suction section are connected to form a leading edge side, the other side of the pressure section and the other side of the suction section are parallelly arranged at intervals to form a trailing edge side, the first cavity, the second cavity and the cleft are located between the pressure section and the suction section, the outer surface of the pressure section forms a pressure surface, the outer surface of the suction section forms a suction surface, and the first cooling assembly and the second cooling assembly are arranged on the inner wall of the pressure section and/or the inner wall of the suction section.
3. The turbine vane blade for a gas turbine as claimed in claim 2, wherein the first cavity is plural, the plural first cavities are arranged in series in the direction from the leading edge side to the trailing edge side, and the plural first cavities are provided on the front side of the second cavity.
4. The turbine vane blade for a gas turbine as claimed in claim 3, wherein the plurality of first cavities include a first sub-cavity and a second sub-cavity, the second sub-cavity is located between the first sub-cavity and the second cavity, a first partition plate and a second partition plate are provided in the blade body, the first partition plate is provided between the first sub-cavity and the second sub-cavity, and the second partition plate is provided between the second sub-cavity and the second cavity.
5. The turbine vane blade for a gas turbine as claimed in claim 4, wherein the blade body has a mean camber line, the mean camber line being connected between the leading edge side and the trailing edge side, and a pitch between any point on the mean camber line and the pressure section is equal to a pitch between the point and the suction section.
6. The vane blade for a gas turbine as claimed in claim 5, wherein a ratio of a mean camber line length between the first diaphragm and the leading edge side to a total length of the mean camber line is 20% to 40%, and a ratio of a mean camber line length between the second diaphragm and the leading edge side to a total length of the mean camber line is 40% to 80%.
7. The turbine stationary blade for a gas turbine as set forth in claim 2, wherein a second cooling block is provided on said pressure section and said suction section, said cleavage slit is formed between said second cooling block of said pressure section and said second cooling block of said suction section, and a ratio of a width dimension of said cleavage slit to a height dimension of said second cooling block is not more than 1.
8. The turbine vane blade for a gas turbine as claimed in claim 1, wherein the camber line has a mounting section, a ratio of a camber line length between any point on the mounting section and the leading edge side to a total length of the camber line is 50% to 95%, and the second cooling module is provided at a position of the blade body portion corresponding to the mounting section.
9. The turbine vane blade for a gas turbine as claimed in claim 1, wherein an extending direction of the first rib forms an angle α with a cross section of the blade body, an extending direction of the second rib forms an angle β with a cross section of the blade body, and the angle α and the angle β are 30 degrees to 160 degrees.
10. The turbine vane blade for a gas turbine as claimed in claim 1, wherein a ratio of a pitch of two adjacent first ribs to a width dimension of the first ribs is 4 to 15, and a ratio of a pitch of two adjacent second ribs to a width dimension of the second ribs is 4 to 15.
11. The turbine vane blade for a gas turbine as claimed in claim 1, wherein the first cooling module has a spanwise direction and a flow direction, and in the flow direction, a ratio of a pitch of two adjacent impingement cooling holes to a diameter of the impingement cooling holes is 4 to 10, a ratio of a pitch of two adjacent column ribs to a diameter of the column ribs is 2.5 to 5, and in the spanwise direction, a ratio of a pitch of two adjacent impingement cooling holes to a diameter of the impingement cooling holes is 4 to 12, and a ratio of a pitch of two adjacent column ribs to a diameter of the column ribs is 2.5 to 10.
12. The turbine vane blade for a gas turbine according to any one of claims 1 to 11, wherein a ratio of a pitch of the laminae and an inner wall of the blade body to the impingement cooling hole is 1 to 3.
13. A turbine vane comprising a blade, an upper endwall and a lower endwall, the blade being connected between the upper endwall and the lower endwall, the blade being a turbine vane blade as claimed in any one of claims 1 to 12.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111564331.4A CN114396316A (en) | 2021-12-20 | 2021-12-20 | Turbine stator blade and turbine stator blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111564331.4A CN114396316A (en) | 2021-12-20 | 2021-12-20 | Turbine stator blade and turbine stator blade |
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| Publication Number | Publication Date |
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| CN114396316A true CN114396316A (en) | 2022-04-26 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202111564331.4A Pending CN114396316A (en) | 2021-12-20 | 2021-12-20 | Turbine stator blade and turbine stator blade |
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Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20130034429A1 (en) * | 2010-04-14 | 2013-02-07 | Dave Carter | Blade or vane for a turbomachine |
| CN207879399U (en) * | 2018-01-23 | 2018-09-18 | 中国科学院工程热物理研究所 | A kind of turbine blade cooling structure |
| CN113513371A (en) * | 2021-08-19 | 2021-10-19 | 北京全四维动力科技有限公司 | Double-wall cooling blade, turbine blade using same and gas turbine |
-
2021
- 2021-12-20 CN CN202111564331.4A patent/CN114396316A/en active Pending
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20130034429A1 (en) * | 2010-04-14 | 2013-02-07 | Dave Carter | Blade or vane for a turbomachine |
| CN207879399U (en) * | 2018-01-23 | 2018-09-18 | 中国科学院工程热物理研究所 | A kind of turbine blade cooling structure |
| CN113513371A (en) * | 2021-08-19 | 2021-10-19 | 北京全四维动力科技有限公司 | Double-wall cooling blade, turbine blade using same and gas turbine |
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