CN115370423A - Gas turbine blade manufacturing method, gas turbine blade, and gas turbine - Google Patents
Gas turbine blade manufacturing method, gas turbine blade, and gas turbine Download PDFInfo
- Publication number
- CN115370423A CN115370423A CN202211170578.2A CN202211170578A CN115370423A CN 115370423 A CN115370423 A CN 115370423A CN 202211170578 A CN202211170578 A CN 202211170578A CN 115370423 A CN115370423 A CN 115370423A
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- Prior art keywords
- blade body
- spoiler
- gas turbine
- wall surface
- blade
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- 238000004519 manufacturing process Methods 0.000 title claims abstract description 38
- 238000001816 cooling Methods 0.000 claims abstract description 68
- 238000005192 partition Methods 0.000 claims description 7
- 230000000149 penetrating effect Effects 0.000 claims description 4
- 238000004891 communication Methods 0.000 claims description 3
- 238000007789 sealing Methods 0.000 claims description 3
- 238000005520 cutting process Methods 0.000 claims description 2
- 230000000694 effects Effects 0.000 abstract description 9
- 238000000034 method Methods 0.000 abstract description 6
- 239000007789 gas Substances 0.000 description 46
- 239000000112 cooling gas Substances 0.000 description 12
- 238000009434 installation Methods 0.000 description 4
- 238000005266 casting Methods 0.000 description 2
- 239000000463 material Substances 0.000 description 2
- 229910000838 Al alloy Inorganic materials 0.000 description 1
- RYGMFSIKBFXOCR-UHFFFAOYSA-N Copper Chemical compound [Cu] RYGMFSIKBFXOCR-UHFFFAOYSA-N 0.000 description 1
- 229910000640 Fe alloy Inorganic materials 0.000 description 1
- 230000004075 alteration Effects 0.000 description 1
- 230000008602 contraction Effects 0.000 description 1
- 229910052802 copper Inorganic materials 0.000 description 1
- 239000010949 copper Substances 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 239000012530 fluid Substances 0.000 description 1
- 238000003754 machining Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 229910000753 refractory alloy Inorganic materials 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
Images
Classifications
-
- 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
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
Abstract
The invention provides a gas turbine blade manufacturing method, a gas turbine blade and a gas turbine. The manufacturing method of the gas turbine blade comprises the following steps: manufacturing a spoiler according to a preset shape; connecting a plurality of the spoiler components with each other and arranging the spoiler components at intervals along the height direction of a blade body of the gas turbine blade to form a spoiler assembly; and fixing the turbulence component in a cooling cavity of the blade body, wherein the distance between the turbulence component and the inner wall surface of the blade body is smaller than or equal to a preset value. Therefore, the method for manufacturing the gas turbine blade according to the invention facilitates manufacturing the gas turbine blade with the spoiler having a complicated structure and a good spoiler effect.
Description
Technical Field
The invention relates to the technical field of gas turbines, in particular to a gas turbine blade manufacturing method, a gas turbine blade and a gas turbine.
Background
In order to reduce the temperature of the blade, a plurality of cooling structures are generally designed in the blade to cool the wall surface of the blade. In the related art, a relatively simple cooling structure, such as turbulence columns, fins, etc., is generally adopted due to the influence of the casting technology. For complex and efficient cooling structures, machining and casting cannot be generally performed.
Disclosure of Invention
The present invention is directed to solving, at least to some extent, one of the technical problems in the related art. To this end, embodiments of the present invention provide a gas turbine blade manufacturing method, a gas turbine blade, and a gas turbine.
The manufacturing method of the gas turbine blade comprises the following steps:
manufacturing a spoiler according to a preset shape;
connecting a plurality of the spoiler components with each other and arranging the spoiler components at intervals along the height direction of a blade body of the gas turbine blade to form a spoiler assembly;
and fixing the turbulence component in a cooling cavity of the blade body, wherein the distance between the turbulence component and the inner wall surface of the blade body is smaller than or equal to a preset value.
Therefore, the method for manufacturing the gas turbine blade according to the embodiment of the invention facilitates manufacturing the gas turbine blade with the spoiler having a complicated structure and a good spoiler effect.
In some embodiments, a plurality of the spoilers are connected by a bracket.
In some embodiments, the spoiler assembly is welded to an inner wall surface of the blade body and the spoiler abuts against the inner wall surface of the blade.
In some embodiments, the inner wall surface of the blade body includes a first inner wall surface and a second inner wall surface that are connected, the first inner wall surface and the second inner wall surface are oppositely arranged in the thickness direction of the blade body, the first inner wall surface and the second inner wall surface define the cooling cavity, and both the first inner wall surface and the second inner wall surface are arc-shaped surfaces;
the plurality of spoilers comprises a plurality of first spoilers and a plurality of second spoilers;
the step of connecting a plurality of spoilers to each other and arranging along the height direction interval of blade body includes:
arranging a plurality of first spoilers on the bracket at intervals along the height direction of the blade body;
arranging a plurality of second spoilers on the bracket at intervals along the height direction of the blade body;
the step of fixing the spoiler assembly in the cooling cavity and enabling the distance between the spoiler and the inner wall surface of the blade body to be smaller than or equal to a preset value comprises the following steps of:
the turbulent flow component and the cooling cavity are arranged in the cooling cavity in an interference fit mode, the first turbulent flow piece abuts against the first inner wall face, and the second turbulent flow piece abuts against the second inner wall face.
In some embodiments, the bracket includes a connecting column, a plurality of first connecting rods, and a plurality of second connecting rods, a length direction of the connecting column is a height direction of the blade body, each of the first spoiler is connected to the connecting column through at least one of the first connecting rods, and each of the second spoiler is connected to the connecting column through at least one of the second connecting rods;
the step of fixing the spoiler assembly in the cooling cavity and enabling the distance between the spoiler and the inner wall surface of the blade body to be smaller than or equal to a preset value comprises the following steps:
fixing the turbulence assembly on the inner wall surface of the blade;
and cutting off the part of the bracket extending out of the cooling cavity.
In some embodiments, the cooling cavity is provided with cold air inlets at two sides of the blade body in the height direction,
the method of manufacturing a gas turbine blade further comprises the steps of:
the blade body is provided with an end cover for sealing the cold air port, and the end cover is provided with a plurality of cold air holes which penetrate through the blade body in the height direction.
The method for manufacturing the gas turbine blade of the embodiment of the invention further comprises the following steps:
a partition plate is arranged in the cooling cavity so as to divide the cooling cavity into a plurality of sub-cavities which are sequentially arranged in the length direction of the blade body;
disposing the flow perturbation assembly within at least one of the plurality of sub-cavities.
The method for manufacturing the gas turbine blade of the embodiment of the invention further comprises the following steps:
the blade body is provided with impact holes penetrating through the blade body, and the impact holes are distributed on the inner wall surface of the blade in a staggered manner with the spoiler.
The invention also provides a gas turbine blade, which comprises
A blade body having a cooling cavity and an impingement hole in communication with the cooling cavity;
the vortex subassembly, the vortex subassembly is established the cooling intracavity, the vortex subassembly includes support and a plurality of vortex spare, and is a plurality of the vortex spare is established on the support and follow the direction of height interval arrangement of blade body, the vortex spare with distance less than or equal to default between the internal face of blade body, at the perpendicular to on the cross section of the thickness direction of blade body, the projection of vortex spare is at least one of rectangle, ripple shape, triangle-shaped, trapezoidal.
The invention also provides a gas turbine which comprises the gas turbine blade.
Drawings
FIG. 1 is a schematic view of a gas turbine blade according to an embodiment of the invention.
FIG. 2 is a top view of a gas turbine blade according to an embodiment of the invention.
FIG. 3 is a schematic view of a spoiler and a bracket according to an embodiment of the present invention.
Reference numerals are as follows:
a gas turbine blade 100;
the structure comprises a blade body 1, a cooling cavity 11, a first inner wall surface 12, a second inner wall surface 13, impact holes 14, sub-cavities 15 and a partition plate 16;
the spoiler 2, the first spoiler 21, the second spoiler 22;
the support 3, the connecting column 31, the first connecting rod 32 and the second connecting rod 33.
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.
A method of manufacturing a gas turbine blade according to an embodiment of the present invention will be described below with reference to the drawings. As shown in fig. 1 to 3, a method for manufacturing a gas turbine blade according to an embodiment of the present invention includes the steps of:
the spoiler 2 is manufactured in a predetermined shape. Specifically, independently make spoiler 2 according to predetermineeing the shape in the outside of blade body 1 to can reduce spoiler 2's the manufacturing degree of difficulty so that carry out meticulously to spoiler 2, and then be convenient for make the spoiler 2 that the structure is complicated and the vortex is effectual.
The plurality of spoilers 2 are connected to each other and arranged at intervals in the height direction of the blade body 1 of the gas turbine blade 100 to form a spoiler assembly. The spoiler assembly is fixed in the cooling cavity 11 of the blade body 1 and the distance between the spoiler 2 and the inner wall surface of the blade body 1 is less than or equal to a preset value, so that the gas turbine blade 100 can be manufactured.
Specifically, a plurality of spoiler 2 link to each other and form the spoiler assembly and can reduce the installation degree of difficulty of installing spoiler 2 to improve the installation effectiveness and the installation accuracy of spoiler 2. That is, the plurality of spoiler 2 are connected to form the spoiler assembly, which can reduce the manufacturing difficulty of the gas turbine blade 100 and improve the installation accuracy of the plurality of spoiler 2 in the gas turbine blade 100.
The direction of height of the blade body 1 of the gas turbine blade 100 is the direction of conveyance of the cooling gas, a plurality of vortex pieces 2 are arranged along the direction of height of the blade body 1 of the gas turbine blade 100 at intervals and the distance between the vortex piece 2 and the inner wall surface of the blade body 1 is less than or equal to a preset value, when the cooling gas is conveyed to the cooling cavity 11, a plurality of vortex pieces 2 of the vortex component are convenient for sequentially carrying out vortex on the cooling gas flow, so that more cooling gas in the cooling cavity 11 can easily exchange heat with the inner wall surface of the blade body 1, the cooling effect of the cooling gas on the blade body 1 is further improved, the cooling effect of the cooling gas on the blade body 1 can be further improved by the vortex piece 2 with a complex structure and a good vortex effect, and the temperature of the vortex piece 100 during operation is reduced.
Therefore, the gas turbine blade manufacturing method according to the embodiment of the invention facilitates manufacturing the gas turbine blade 100 having the spoiler 2 having a complicated structure and a good spoiler effect.
The present invention also provides a gas turbine blade 100, and the following describes a method for manufacturing a gas turbine blade according to an embodiment of the present invention in detail with reference to the gas turbine blade 100 according to an embodiment of the present invention.
As shown in fig. 1 to 3, a gas turbine blade 100 according to an embodiment of the present invention includes a blade body 1 and a spoiler assembly.
The blade body 1 has a cooling cavity 11 and an impingement hole 14 communicating with the cooling cavity 11. Specifically, the inner wall surface of the blade body 1 defines a cooling cavity 11, and the extending direction of the cooling cavity 11 coincides with the height direction of the blade body 1. The height direction of the blade body 1 may be an up-down direction as indicated by an arrow in fig. 1.
As shown in fig. 2, in some embodiments, the inner wall surface of the blade body 1 includes a first inner wall surface 12 and a second inner wall surface 13 connected to each other, the first inner wall surface 12 and the second inner wall surface 13 are disposed opposite to each other in a thickness direction of the blade body 1, the first inner wall surface 12 and the second inner wall surface 13 define the cooling cavity 11, and both the first inner wall surface 12 and the second inner wall surface 13 are arc-shaped surfaces. Specifically, the blade body 1 is a long and narrow arc-shaped cylinder, and the thickness direction and the length direction of the blade body 1 are both horizontal directions. For example, the thickness direction (substantially) of the blade body 1 is the front-rear direction, and the first inner wall surface 12 is located forward of the second inner wall surface 13. The longitudinal direction (substantially) of the blade body 1 is the left-right direction, the front edge of the blade body 1 is positioned to the right of the rear edge of the blade body 1, and the front edge and the rear edge of the blade body 11 are determined according to the flow direction of the external high-temperature fluid. The front-back direction and the left-right direction are shown by arrows in fig. 2.
The vortex subassembly is established in cooling chamber 11, and the vortex subassembly includes support 3 and a plurality of vortex piece 2. On a cross section perpendicular to the thickness direction of the blade body 1, the projection of the spoiler 2 is at least one of rectangular, corrugated, triangular, and trapezoidal. Specifically, the spoiler 2 is manufactured in a predetermined shape so as to manufacture the spoiler 2 having a complicated structure and a good spoiler effect.
In the gas turbine blade manufacturing method, a plurality of spoilers 2 are connected to each other and arranged at intervals in the height direction of the blade body 1 of the gas turbine blade 100 to form a spoiler assembly. Specifically, pass through support 3 with a plurality of spoilers 2 and connect for a plurality of spoilers 2 are established on support 3 and are arranged along the direction of height interval of blade body 1, so that a plurality of spoilers 2 so as to carry out the vortex in proper order to the cooling gas that gets into in the cooling chamber 11. The plurality of spoilers 2 includes a plurality of first spoilers 21 and a plurality of second spoilers 22.
The step of connecting the plurality of spoilers 2 to each other and arranging at intervals in the height direction of the blade body 1 includes:
arranging a plurality of first spoilers 21 on the bracket 3 at intervals in the height direction of the blade body 1;
a plurality of second spoiler 22 are provided on the bracket 3 at intervals in the height direction of the blade body 1.
As shown in fig. 1 to 3, the bracket 3 includes a connection column 31, a plurality of first connection rods 32, and a plurality of second connection rods 33, a length direction of the connection column 31 is a height direction of the blade body 1, each first spoiler 21 is connected to the connection column 31 through at least one first connection rod 32, and each second spoiler 22 is connected to the connection column 31 through at least one second connection rod 33. For example, the longitudinal direction of the connection column 31 is the vertical direction, a plurality of first connection bars 32 are provided on the front side of the connection column 31, and a plurality of second connection bars 33 are provided on the rear side of the connection column 31. The plurality of first spoiler 21 are arranged at intervals in the up-down direction, and each first spoiler 21 is connected to at least one first connecting rod 32. The plurality of second spoiler 22 are arranged at intervals in the up-down direction, and each second spoiler 22 is connected to at least one second connecting rod 33.
Optionally, the spoiler 2, the bracket 3, the first connecting rod 32 and the second connecting rod 33 may all be made of non-refractory alloy, so as to reduce the cost. For example, the spoiler 2, the bracket 3, the first connecting rod 32 and the second connecting rod 33 may be made of iron alloy, aluminum alloy, copper, and the like.
The distance between the spoiler 2 and the inner wall surface of the blade body 1 is smaller than or equal to a preset value. For example, the distance between the spoiler 2 and the inner wall surface of the blade body 1 is 0.5mm or less.
In some embodiments, the spoiler assembly is welded to the inner wall surface of the blade body 1 and the spoiler 2 abuts against the inner wall surface of the blade. Thereby making it possible to enhance the cooling effect of the cooling gas on the gas turbine blade 100. Specifically, the spoiler 2 abuts against the inner wall surface of the blade and welds at least part of the ends of the plurality of spoilers 2 to the inner wall surface of the blade body 1, so that the spoiler assembly is stably connected to the blade body 1.
As shown in fig. 1 and 2, in some embodiments, the step of fixing the spoiler assembly in the cooling cavity 11 and making the distance between the spoiler 2 and the inner wall surface of the blade body 1 be less than or equal to a preset value includes:
the spoiler assembly is arranged in the cooling cavity 11 in an interference fit manner with the cooling cavity 11, the first spoiler 21 abuts against the first inner wall surface 12, and the second spoiler 22 abuts against the second inner wall surface 13. Specifically, after the first spoiler 21 and the second spoiler 22 are connected with the bracket 3, the positions of the first spoiler 21 and the second spoiler 22 are controlled, and then the blade body 1 and the spoiler assembly (the spoiler 2, the bracket 3, the first connecting rod 32 and the second connecting rod 33) are inconsistent in expansibility, so that the spoiler assembly and the cooling cavity 11 are in interference fit through thermal expansion and cold contraction. The inner wall surface of the blade body 1 is smoothed so that the spoiler assembly can easily enter the cooling cavity 11. The first spoiler 21 abuts against the first inner wall surface 12, and the second spoiler 22 abuts against the second inner wall surface 13. Then, the first spoiler 21 is welded to the first inner wall surface 12, and the second spoiler 22 is welded to the second inner wall surface 13, so that the spoiler assembly is stably connected to the blade body 1.
In some embodiments, the step of fixing the spoiler assembly in the cooling cavity 11 and setting the distance between the spoiler 2 and the inner wall surface of the blade body 1 to be less than or equal to a preset value includes: fixing the turbulence assembly on the inner wall surface of the blade; the portion of the holder 3 protruding outside the cooling chamber 11 is cut off. Specifically, the spoiler assembly stretches into the cooling cavity 11 through the support 3 with a long length, and the portion of the support 3 stretching out of the cooling cavity 11 is cut off so that the spoiler assembly can be matched with the size of the blade body 1.
In some embodiments, the cooling cavity 11 has cooling air ports on both sides of the blade body 1 in the height direction, and cooling air can enter and exit the cooling cavity 11 through the cooling air ports. For example, the cooling chamber 11 has cooling air ports on both sides in the up-down direction.
The method of manufacturing a gas turbine blade further comprises the steps of:
an end cover for sealing the cold air port is arranged on the blade body 1, and a plurality of cold air holes penetrating through the blade body 1 in the height direction are formed in the end cover. So that the cooling gas can enter the cooling cavity 11 through the cold air holes on the end cover.
In some embodiments, the gas turbine blade manufacturing method further comprises the steps of:
a partition plate 16 is provided in the cooling chamber 11 so as to divide the cooling chamber 11 into a plurality of sub-chambers 15 arranged in order in the longitudinal direction of the blade body 1. Specifically, the thickness direction of the partition plate 16 (substantially) coincides with the length direction of the blade body 1, and the partition plate 16 is plural and provided at intervals in the length direction of the blade body 1, so that the plurality of sub-cavities 15 are provided in order in the length direction of the blade body 1.
As shown in fig. 1 and 2, a flow perturbation assembly is provided within at least one of the plurality of sub-cavities 15. Thereby, the cooling gas may be disturbed within at least one of the plurality of sub-cavities 15 (at least part of the cooling cavity 11) in order to increase the cooling effect of the blade body 1. For example, the cooling chamber 11 is divided into four sub-chambers 14 by the partition plates 16, the sub-chamber 14 at the trailing edge has an opening facing away from the leading edge, and an inner casing is arranged in the sub-chamber 15 at the trailing edge, and the inner casing and the blade body 1 define a cooling sandwich chamber. Turbulent flow components are arranged in the sub-cavities 15 at the middle part and the front edge of the blade body 1.
In some embodiments, the gas turbine blade manufacturing method further comprises the steps of:
as shown in fig. 1, the blade body 1 is provided with the impingement holes 14 penetrating therethrough, and the impingement holes 14 are distributed on the inner wall surface of the blade in a staggered manner from the spoiler 2. Thereby, it is facilitated that the cooling gas in the cooling cavity 11 flows out of the outer surface of the blade body 1 through the impingement holes 14 and forms a cooling gas flow layer on the outer surface of the blade body 1.
The invention also provides a gas turbine, and the gas turbine according to the embodiment of the invention comprises the gas turbine blade 100 according to the embodiment of the invention.
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 to implicitly indicate the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one of the feature. In the description of the present invention, "a plurality" means at least two, e.g., two, three, etc., unless specifically limited 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," "above," and "over" a second feature may be directly on or obliquely above the second feature, or simply mean 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 present disclosure, the terms "one embodiment," "some embodiments," "an example," "a specific example," or "some examples" and the like mean that a specific feature, structure, material, or characteristic described in connection with the embodiment or example is included in at least one embodiment or example of the present disclosure. 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. Moreover, various embodiments or examples and features of various embodiments or examples described in this specification can be combined and combined by one skilled in the art without being mutually inconsistent.
Although the above embodiments have been shown and described, it should be understood that they are exemplary and not intended to limit the invention, and that various changes, modifications, substitutions and alterations can be made herein by those skilled in the art without departing from the scope of the invention.
Claims (10)
1. A method of manufacturing a gas turbine blade, comprising the steps of:
manufacturing a spoiler according to a preset shape;
connecting a plurality of the spoiler components with each other and arranging the spoiler components at intervals along the height direction of a blade body of the gas turbine blade to form a spoiler assembly;
and fixing the turbulence assembly in a cooling cavity of the blade body, wherein the distance between the turbulence assembly and the inner wall surface of the blade body is smaller than or equal to a preset value.
2. The method of manufacturing a gas turbine blade according to claim 1, wherein a plurality of spoilers are connected by a bracket.
3. The method of manufacturing a gas turbine blade according to claim 2, wherein the spoiler assembly is welded to an inner wall surface of the blade body and the spoiler abuts against the inner wall surface of the blade.
4. The gas turbine blade manufacturing method according to claim 3,
the inner wall surface of the blade body comprises a first inner wall surface and a second inner wall surface which are connected, the first inner wall surface and the second inner wall surface are oppositely arranged in the thickness direction of the blade body, the first inner wall surface and the second inner wall surface limit the cooling cavity, and the first inner wall surface and the second inner wall surface are both arc-shaped surfaces;
the plurality of spoilers comprises a plurality of first spoilers and a plurality of second spoilers;
the step of connecting a plurality of spoilers each other and arranging along the direction of height interval of blade body includes:
arranging a plurality of first spoilers on the bracket at intervals along the height direction of the blade body;
arranging a plurality of second spoilers on the bracket at intervals along the height direction of the blade body;
the step of fixing the spoiler assembly in the cooling cavity and enabling the distance between the spoiler and the inner wall surface of the blade body to be smaller than or equal to a preset value comprises the following steps:
the flow disturbing assembly and the cooling cavity are arranged in the cooling cavity in an interference fit mode, the first flow disturbing piece abuts against the first inner wall face, and the second flow disturbing piece abuts against the second inner wall face.
5. The gas turbine blade manufacturing method according to claim 4,
the support comprises a connecting column, a plurality of first connecting rods and a plurality of second connecting rods, the length direction of the connecting column is the height direction of the blade body, each first spoiler is connected with the connecting column through at least one first connecting rod, and each second spoiler is connected with the connecting column through at least one second connecting rod;
the step of fixing the spoiler assembly in the cooling cavity and enabling the distance between the spoiler and the inner wall surface of the blade body to be smaller than or equal to a preset value comprises the following steps:
fixing the turbulence assembly on the inner wall surface of the blade;
and cutting off the part of the bracket extending out of the cooling cavity.
6. The gas turbine blade manufacturing method according to claim 1,
the two sides of the cooling cavity in the height direction of the blade body are provided with cold air ports,
the method of manufacturing a gas turbine blade further comprises the steps of:
the blade body is provided with an end cover for sealing the cold air port, and the end cover is provided with a plurality of cold air holes which penetrate through the blade body in the height direction.
7. The method of manufacturing a gas turbine blade according to claim 1, further comprising the steps of:
the cooling cavity is internally provided with a partition plate so as to divide the cooling cavity into a plurality of sub-cavities which are sequentially distributed in the length direction of the blade body;
disposing the flow perturbation assembly within at least one of the plurality of sub-cavities.
8. The gas turbine blade manufacturing method according to any one of claims 1 to 7, further comprising the steps of:
the blade body is provided with impact holes penetrating through the blade body, and the impact holes are distributed on the inner wall surface of the blade and the spoiler in a staggered manner.
9. A gas turbine blade, comprising
A blade body having a cooling cavity and an impingement hole in communication with the cooling cavity;
the vortex subassembly, the vortex subassembly is established the cooling intracavity, the vortex subassembly includes support and a plurality of vortex spare, and is a plurality of the vortex spare is established on the support and follow the direction of height interval arrangement of blade body, the vortex spare with distance less than or equal to default between the internal face of blade body, at the perpendicular to on the cross section of the thickness direction of blade body, the projection of vortex spare is at least one of rectangle, ripple shape, triangle-shaped, trapezoidal.
10. A gas turbine engine comprising the gas turbine engine blade of claim 9.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211170578.2A CN115370423A (en) | 2022-09-23 | 2022-09-23 | Gas turbine blade manufacturing method, gas turbine blade, and gas turbine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211170578.2A CN115370423A (en) | 2022-09-23 | 2022-09-23 | Gas turbine blade manufacturing method, gas turbine blade, and gas turbine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN115370423A true CN115370423A (en) | 2022-11-22 |
Family
ID=84071258
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202211170578.2A Pending CN115370423A (en) | 2022-09-23 | 2022-09-23 | Gas turbine blade manufacturing method, gas turbine blade, and gas turbine |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN115370423A (en) |
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2022
- 2022-09-23 CN CN202211170578.2A patent/CN115370423A/en active Pending
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