Turbine blade and method of manufacturing a turbine blade, turbine blade assembly

Abstract

Provided is a method for manufacturing a turbine blade with cold air cavities, comprising the following steps: winding a plurality of layers of fiber cloth around an inner cavity core mold to form an inner cavity preform; splicing the inner cavity preform and a filling preform; winding a plurality of layers of fiber cloth around the outer part of the inner cavity preform and the filling preform to obtain a blade preform with a blade outer shape, and performing curing and shaping; and removing the inner cavity core mold to obtain a turbine blade with a plurality of cold air cavities. The above-mentioned method for manufacturing a turbine blade is based on the manufacturing process of ceramic matrix composites, and through the design of the fiber layup of the cooling cavity and the fiber preform preparation forming mode, a plurality of cooling cavities and rotating channels are introduced into the turbine blade, so that the cold air flow path design of the blade is optimized, and a reasonable temperature distribution is provided for the blade.

Description

Technical Field

[0001] This invention relates to the field of aero-engines, and more specifically to turbine blades. Background Technology

[0002] Turbine blades are key components that convert the energy of high-temperature, high-pressure gases after combustion into mechanical power. Due to their long-term exposure to high-temperature environments, some turbine blades are made of ceramic matrix composites, which have a higher operating temperature compared to conventional high-temperature metals.

[0003] The significant differences in heat load across different regions of the blade result in low overall cooling efficiency. Therefore, an effective fabrication method is needed to prepare ceramic matrix composite turbine blades with cooling structures. Summary of the Invention

[0004] One object of the present invention is to provide a method for manufacturing turbine blades with a cooling gas cavity.

[0005] The preparation method to achieve the above objective includes the following steps: wrapping a multi-layer fiber cloth around an inner cavity mandrel to form an inner cavity preform; splicing the inner cavity preform and the filling preform; wrapping the multi-layer fiber cloth around the outside of the inner cavity preform and the filling preform to obtain a blade preform with a blade shape, and then curing and shaping it; removing the inner cavity mandrel to obtain a turbine blade with multiple cooling gas cavities.

[0006] In one or more embodiments, a radial channel preform and an embedded body are prepared, the embedded body being embedded in the radial channel preform; the radial channel preform is spliced ​​together with the inner cavity preform and the filling preform; the inner cavity preform, the filling preform and the radial channel preform are wrapped with multi-layer fiber cloth to obtain a blade preform with a blade shape; the embedded body is removed to form a cooling channel.

[0007] In one or more embodiments, the radial channel preform includes an inlet channel, a turning channel, and a flow channel that are interconnected with each other, and a plurality of the pre-embedded bodies are embedded in the inlet channel, the turning channel, and the flow channel, wherein the end of the inlet channel is connected to the outside of the radial channel preform, and removing the pre-embedded bodies forms a rotary cooling channel.

[0008] In one or more embodiments, an opening is machined in the embedded body, and the filling preform is filled into the opening; the embedded body is removed, and the filling preform forms a turbulence structure.

[0009] In one or more embodiments, film cooling holes are machined on the turbine blades, the film cooling holes communicating with the cold air cavity and / or cooling structure.

[0010] In one or more embodiments, the inner cavity preforms are spliced ​​along the chordal direction of the leaf shape.

[0011] Another object of the present invention is to provide a turbine blade prepared by the above-described turbine blade preparation method.

[0012] In one or more embodiments, the turbine blade further includes a rotary cooling channel, the rotary cooling channel including a radial cold air inlet, a radial flow section and an end bend section, and a plurality of the radial flow sections are connected through the end bend section.

[0013] In one or more embodiments, the rotary cooling channel is located in the middle and rear part of the blade along the chord length direction.

[0014] In one or more embodiments, the turbine blade includes a plurality of cooling gas cavities arranged along the blade chord direction.

[0015] In one or more embodiments, the turbine blades further include film cooling holes that communicate with the rotary cooling channel and / or the cooling air cavity.

[0016] Another object of the present invention is to provide a turbine blade assembly, including the aforementioned turbine blade, and further including a sleeve disposed in at least one cooling gas cavity, the sleeve including an air supply cavity and a cooling gas impact hole.

[0017] In one or more embodiments, the turbine blade includes a plurality of cooling gas cavities distributed along the blade chord direction, the sleeve is disposed in the first cooling gas cavity along the blade chord direction, and a rotary cooling channel is also provided on the rear side of the last cooling gas cavity along the blade chord direction.

[0018] The aforementioned turbine blade manufacturing method, under certain cooling gas volume constraints, is based on the manufacturing process of ceramic matrix composites. By designing the fiber layup and fiber preform preparation method of the cooling chamber, multiple cooling cavities and rotating channels are introduced into the turbine blade, thereby optimizing the cooling gas flow path design of the blade. For the multiple cooling cavity structure, the cooling gas in the front cooling gas cavity provides effective cooling protection for both the outer and inner walls of the blade through impact holes and film cooling holes. The rotating channel located at the rear is closer to the blade trailing edge, enabling the cooling gas to effectively reduce the internal and external temperatures of the trailing edge. The cooling gas in the rear cooling gas cavity, with a lower temperature, flows directly out through the film cooling holes and provides effective film cooling protection for the downstream blade trailing edge, thus providing a reasonable temperature distribution for the entire blade. Attached Figure Description

[0019] The above and other features, properties and advantages of the present invention will become more apparent from the following description taken in conjunction with the accompanying drawings and embodiments, wherein:

[0020] Figure 1 This is a cross-sectional layup diagram of the first embodiment of the turbine blade;

[0021] Figure 2 This is a cross-sectional schematic diagram of a first embodiment of the turbine blade assembly;

[0022] Figure 3 This is a longitudinal cross-sectional view of the first embodiment of the turbine blade assembly;

[0023] Figure 4 This is a cross-sectional layup diagram of the second embodiment of the turbine blade;

[0024] Figure 5 This is a cross-sectional schematic diagram of a second embodiment of the turbine blade assembly;

[0025] Figure 6 This is a longitudinal cross-sectional layup diagram of the second embodiment of the turbine blade;

[0026] Figure 7 This is a longitudinal cross-sectional schematic diagram of the second embodiment of the turbine blade;

[0027] Figure 8 This is a cross-sectional layup diagram of the third embodiment of the turbine blade;

[0028] Figure 9 This is a cross-sectional schematic diagram of a third embodiment of the turbine blade assembly;

[0029] Figure 10 This is a longitudinal cross-sectional layup diagram of the third embodiment of the turbine blade;

[0030] Figure 11 This is a longitudinal cross-sectional schematic diagram of the third embodiment of the turbine blade;

[0031] Figure 12A-12B This is a schematic diagram of the embedded body for turbine blades. Detailed Implementation

[0032] The present invention will be further described below with reference to specific embodiments and accompanying drawings. More details are set forth in the following description in order to provide a full understanding of the present invention. However, the present invention can obviously be implemented in many other ways different from those described herein. Those skilled in the art can make similar extensions and derivations based on actual application situations without departing from the spirit of the present invention. Therefore, the scope of protection of the present invention should not be limited by the content of this specific embodiment.

[0033] It should be noted that these and other accompanying drawings are merely examples and are not drawn to scale, and should not be construed as limiting the scope of protection of the present invention.

[0034] Reference Figure 2 and Figure 3 As shown, the turbine blade assembly 1 includes a turbine blade 2 and a sleeve 3, and also includes cover plates located at the top and bottom of the turbine blade 2 and the sleeve 3. The sleeve 3 is generally made of metal, and the turbine blade 2 is generally made of ceramic matrix composite (CMC).

[0035] Ceramic matrix composites are a new type of structural material with excellent thermal properties such as lightweight and high temperature resistance. Compared with conventional high-temperature metals, they have a higher operating temperature, making them suitable for use in hot-end components of aero engines. However, the inlet temperature of high-pressure turbines in today's advanced aero engines is still higher than the operating temperature range of ceramic matrix composites. Therefore, high-pressure turbine components made of ceramic matrix composites still require cooling designs to ensure their safe and reliable operation.

[0036] Conventional metal turbine blades are manufactured using casting processes, which allow for the placement of impact cooling at the leading edge, pressure surface, and suction surface, as well as the arrangement of turbulence columns and slit channels at the trailing edge. This maximizes the internal heat transfer process of the turbine blade and improves the overall cooling efficiency. However, due to manufacturing constraints, typical ceramic matrix composite turbine blades can only have simple cooling chambers, thus limiting the utilization of cool air.

[0037] Meanwhile, high-pressure turbine blades are affected by internal and external heat transfer, resulting in significant differences in heat load across different regions of the blade, leading to low overall cooling efficiency. Therefore, it is essential to consider the cooling structure and fabrication methods of ceramic matrix composite turbine blades in a coordinated manner.

[0038] Based on this, this disclosure proposes a method for preparing and forming a turbine blade with a cooling gas cavity based on ceramic matrix composites. The manufacturing process based on ceramic matrix composites introduces multiple cooling cavities and rotation channels into the turbine blade by designing the fiber layup and fiber preform preparation method of the cooling gas cavity, optimizing the cooling gas flow path design of the blade, and providing a reasonable temperature distribution for the blade.

[0039] The preparation method includes the following steps:

[0040] Multiple layers of fiber cloth are wound around multiple inner cavity core molds to form multiple inner cavity preforms;

[0041] Assemble the precast inner cavity structure and the filling precast structure;

[0042] The inner cavity preform is wrapped with multi-layer fiber cloth and filled with the outside of the preform to obtain a blade preform with a blade shape, and then cured and shaped.

[0043] Remove the inner core mold to obtain turbine blades with multiple cold air cavities.

[0044] Reference Figure 1 or Figure 4 or Figure 8The process involves winding multiple layers of fiber cloth 41 around an inner cavity mandrel 42 to form an inner cavity preform 43, which is then cured and shaped. The fiber cloth 41 includes materials toughened by fibers, such as unidirectional tape, two-dimensional cloth, and three-dimensional fabric.

[0045] After obtaining one or more inner cavity preforms 43, they are spliced ​​with infill preforms 46. The infill preforms 46 are formed by laying multiple layers of fiber cloth 41 to form a preform structure of a specific shape. Multiple layers of fiber cloth 41 are then wound and laid on the outside of one or more inner cavity preforms 43 and infill preforms 46 to obtain an integral blade preform 45 with the designed blade shape. The integral blade preform 45 is then processed through curing, carbonization, silicon infiltration, and other processes to obtain a ceramic matrix composite turbine blade.

[0046] After removing the inner core mold 42, the resulting cavity is the cold air inner cavity 21.

[0047] In some embodiments, film cooling holes 5 can also be machined on the turbine blade, the film cooling holes 5 communicating with the cooling gas cavity 21. The film cooling holes 5 can be arranged on the leading edge 23, pressure surface 24, suction surface 25 and trailing edge 26 of the blade, for providing film cooling to the outer wall surface 28 of the blade. The film cooling holes 5 penetrate the outer wall surface 28 and the inner wall surface 29 of the blade, allowing the cooling gas inside the blade to flow out from the film cooling holes 5.

[0048] The prepared ceramic matrix composite turbine blade 2 is combined with the metal sleeve 3 to form a turbine blade assembly, such as... Figure 2 , Figure 3 , Figure 5 or Figure 9 As shown. The turbine blade 2 has a leading edge 23, a pressure surface 24, a suction surface 25, and a trailing edge 26. A sleeve 3 is fitted inside one or more cooling gas cavities 21 of the ceramic matrix composite blade 2 to provide cooling gas to the blade. The sleeve 3 has at least one air supply cavity 31. The sleeve 3 has an impact hole 6 that penetrates the outer wall surface 32 and the inner wall surface 33 of the sleeve, connecting the sleeve air supply cavity 31 and the blade cooling gas cavity 21. Cooling gas enters the sleeve air supply cavity 31 from the cooling gas inlet 7, and then flows into the blade cooling gas cavity 21 through the impact hole 6, providing impact cooling to the inner wall surface 29 of the blade.

[0049] In some embodiments, radial channel preforms 44 and embedded bodies 47 may also be prepared during the fabrication stage to fabricate cooling channels 22. For example... Figure 6 and Figure 7 As shown, the embedded body 47 is embedded in the radial channel preform 44; the radial preform 44 with the embedded body 47 is spliced ​​together with the inner cavity preform 42 and the filling preform 46; the inner cavity preform 42, the filling preform 46 and the radial channel preform 44 are wrapped with multi-layer fiber cloth 41 to obtain a blade preform with a blade shape; the embedded body 47 is removed to obtain the cooling channel.

[0050] As in Figure 6 and Figure 7 In some embodiments shown, the preform 44 includes an inlet channel, a turning channel, and a flow channel that are interconnected, wherein the end of the inlet channel connects to the outside of the radial channel preform. Multiple embedded bodies 47 are embedded in the inlet channel, the turning channel, and the flow channel, forming a rotary cooling channel 22. The rotary cooling channel has a radial cold air inlet 221, a radial flow section 222, and an end turning section 223, and is formed by splicing the radial channel preform 44 and the embedded bodies 47. The channel preform 44 is the solid area after the blade is formed, and the embedded bodies 47 are the radial flow section 222 and the end turning section 223 after the blade is formed.

[0051] The radial channel preform 44 and the embedded body 47 are spliced ​​together, and then spliced ​​together with the inner cavity preform 43 and the filling preform 46. Multiple layers of fiber cloth 41 are then wound and laid on the outside to obtain the integral blade preform 45 with the designed blade shape. The integral blade preform 45 is then processed through curing, carbonization, and silicon infiltration processes to obtain a ceramic matrix composite turbine blade with cooling channels 22, such as... Figure 9 As shown.

[0052] Preferably, the cooling channel 22 is located in the middle and rear part of the blade, which is closer to the blade trailing edge than the cooling air cavity 21. Utilizing the cooling channel 22 enhances heat transfer of the cooling air on the inner wall surface 33 of the blade, effectively reducing the temperature in the trailing edge 26 region. The cooling channel 22 located at the trailing edge also increases the cooling area within a limited space, further improving heat conduction. The blade trailing edge can be the region behind 70% of the blade chord length.

[0053] The position of the radial channel preform 44 is defined by the embedded body 47. In the splicing and filling design, the embedded body 47 can be parallel or perpendicular to the blade radial direction. The embedded body 47 is made of polymer materials such as nylon and polyimide. During the high-temperature curing process, the polymer material embedded body 47 is heated into gas and loose carbon, which is removed from the integral blade preform 45 to form the cooling channel 22.

[0054] Based on the above embodiments, such as Figure 12A-12B As shown, an opening 470 is machined inside the embedded body 47, and a preform of a specific shape is filled in. After the ceramic matrix composite turbine blade is formed, the embedded body 47 is removed, and the preform 46 forms a turbulence structure (not shown in the figure), located on the inner wall surface 29 of the blade on the pressure surface 24 and / or the suction surface 25. The turbulence structure is used to enhance the heat transfer process of the cold air in the cooling channel 22.

[0055] Understandably, the infill preform 46 can be molded into any desired shape.

[0056] For turbine blade structures with multiple cooling gas cavities, multiple cavity prefabricated bodies 42 are spliced ​​along the chord length direction of the blade profile, such as... Figure 1 and Figure 2 As shown, a partition 27 is formed in the solid area between adjacent cold air cavities 21 or cooling channels 22 inside the blade.

[0057] For turbine blade structures with multiple cooling chambers distributed along the chord length, each chamber needs to provide a different cooling effect. The cooling chamber 21 located in the middle region of the turbine blade 2 is not fitted with a sleeve 3, resulting in relatively weak heat transfer within the blade and a lower cooling temperature. After flowing out through the film cooling holes 5, it effectively provides film cooling protection for the downstream blade trailing edge 26. The cooling chamber 21 located in the leading region of the turbine blade 2 is generally fitted with a sleeve 3, allowing the cooling air to impact and cool the inner wall surface 33 of the blade through the impact holes 6, effectively reducing the temperature of the blade leading edge 23. The cooling chamber 21 has a cooling air inlet 7 at one end and is closed at the other, or both ends have cooling air inlets 7. Cooling air enters through the cooling air inlet 7.

[0058] The above-mentioned preparation method and turbine blade assembly structure are further illustrated below through three embodiments.

[0059] First Embodiment

[0060] Figures 1 to 3 A first embodiment of a turbine blade and turbine blade assembly is shown. The turbine blade assembly 1 includes two parts: a turbine blade 2 and a sleeve 3. The turbine blade 2 is made of ceramic matrix composite material. The ceramic matrix composite turbine blade 2 has two cooling gas cavities 21 arranged in the front and middle regions of the blade, for example, the front and middle regions are located in the region of 0-70% of the blade chord length. The front cooling gas cavity 21 is larger than the rear cooling gas cavity 21. One end of the front cooling gas cavity 21 is open, and the other end is closed. The sleeve 3 has an air supply cavity 31, which is assembled through the open end of the front cooling gas cavity 21. One end of the rear cooling gas cavity 21 has a cooling gas inlet 7, and the other end is closed.

[0061] Cool air enters the sleeve 3 and the rear cool air cavity 21 from the cool air inlet 7. The cool air from the sleeve supply cavity 31 enters the blade cool air cavity 21 through the impact hole 6 and finally flows out through the film cooling hole 5, effectively cooling and protecting both the outer wall surface 28 and the inner wall surface 29 of the blade. The cool air in the rear cool air cavity 21 has relatively weak heat exchange inside the blade and a lower temperature. After flowing out through the film cooling hole 5, it effectively provides film cooling protection for the trailing edge 26 of the downstream blade.

[0062] The method for preparing and forming the ceramic matrix composite turbine blade according to the first embodiment is as follows: Multilayer fiber cloth 41 is wound around the inner cavity mandrel 42 corresponding to the front and rear inner cavities to form front and rear inner cavity preforms 43. After splicing the aforementioned inner cavity preforms 43 with the filling preforms 46, multilayer fiber cloth 41 is continued to be wound and laid on the outside to obtain an integral blade preform 45 with the designed blade shape. The integral blade preform 45 is processed through curing, carbonization, and silicon infiltration processes to obtain the ceramic matrix composite turbine blade. After removing the inner cavity mandrel 42, two cooling gas cavities 21 are obtained. A partition 27 is formed in the solid area between the front and rear cooling gas cavities 21 of the blade. The blade is perforated to form film gas holes 5 on the leading edge 23, pressure surface 24, suction surface 25, and trailing edge 26. The film gas holes 5 connect to the cooling gas cavities and penetrate the outer wall surface 28 and inner wall surface 29 of the blade.

[0063] Second Embodiment

[0064] Figures 4 to 7 A second embodiment of the turbine blade and turbine blade assembly is shown. The turbine blade 2 has a cooling gas cavity 21 at the front (0-70% of the blade chord length) and a cooling channel 22 at the rear. One end of the front cooling gas cavity 21 is open, and the other end is closed. The sleeve 3 has an air supply cavity 31 and is assembled through the open end of the front cooling gas cavity 21. The cooling channel 22 has a radial cooling gas inlet 221, three radial flow sections 222, and two end bend sections 223.

[0065] Cool air enters the sleeve 3 through the cool air inlet 7 and the cooling channel 22 through the radial cool air inlet 221. Cool air from the sleeve's air supply cavity 31 enters the blade's cool air cavity 21 through the impact hole 6 and finally flows out through the film cooling hole 5. Within the cooling channel 22, the cool air flows from one end of the blade 2 to the other along the radial flow section 222, then flows in the opposite direction through the end bend section 223. After multiple flows, it exits through the film cooling hole 5 on the inner wall surface 29 of the downstream radial flow section 222. The cooling channel 22 ensures sufficient heat exchange within the blade.

[0066] The method for preparing and forming the ceramic matrix composite turbine blade in the second embodiment is as follows: Multilayer fiber cloth 41 is wound around the inner cavity mandrel 42 to form an inner cavity preform 43. The radial channel preform 44 and the embedded body 47 are then spliced ​​together, and then spliced ​​together with the inner cavity preform 43 and the filling preform 46. Multilayer fiber cloth 41 is then wound and laid on the outside to obtain an integral blade preform 45 with the designed blade shape. The integral blade preform 45 is treated with curing, carbonization, and silicon infiltration processes, and then the embedded body 47 is removed. A ceramic matrix composite turbine blade with a cooling channel 22 at the tail is obtained. After removing the inner cavity mandrel 42, a cooling gas cavity 21 is obtained. The blade is perforated to form film cooling holes 5 on the leading edge 23, pressure surface 24, suction surface 25, and trailing edge 26. The film cooling holes 5 connect the cooling gas cavity 21 and the cooling channel 22.

[0067] Third Embodiment

[0068] Figures 8 to 11 A third embodiment of the turbine blade and turbine blade assembly is shown. The turbine blade 2 has two cooling gas chambers 21 at the front (0-70% of the blade chord length) and a cooling channel 22 at the rear. One end of the front cooling gas chamber 21 is open, and the other end is closed. The sleeve 3 has an air supply chamber 31 and is assembled through the open end of the front cooling gas chamber 21. One end of the rear cooling gas chamber 21 has a cooling gas inlet 7, and the other end is closed. The cooling channel 22 has a radial cooling gas inlet 221, two radial flow sections 222, and an end bend section 223.

[0069] Cool air enters the sleeve 3 and the rear cool air cavity 21 from the cool air inlet 7, and then enters the cooling channel 22 from the radial cool air inlet 221. Cool air from the sleeve's air supply cavity 31 enters the blade's cool air cavity 21 through the impact hole 6, and finally flows out from the film cooling hole 5. Cool air from the rear cool air cavity 21 flows out through the film cooling hole 5. The cool air in the cooling channel 22 flows from one end of the blade 2 to the other along the radial flow section 222, then flows in the opposite direction through the end bend section 223, and flows out from the film cooling hole 5 on the inner wall surface 29 of the downstream radial flow section 222.

[0070] The method for preparing and forming the ceramic matrix composite turbine blade in the third embodiment is as follows: Multilayer fiber cloth 41 is wound around an inner cavity mandrel 42 to form an inner cavity preform 43. The two inner cavity preforms 43, the filling preform 46, the radial channel preform 44, and the embedded body 47 are spliced ​​together. Multilayer fiber cloth 41 is then wound around the outside of the inner cavity preform 43, the filling preform 46, the radial channel preform 44, and the embedded body 47 to form an integral blade preform 45. After curing, carbonization, and silicon infiltration processes, a ceramic matrix composite turbine blade with two cooling gas cavities 21 and a tail cooling channel 22 is obtained.

[0071] Under certain cooling gas volume constraints, the above method, based on the manufacturing process of ceramic matrix composites, introduces multiple cooling cavities and rotating channels into the turbine blade by designing the fiber layup and fiber preform preparation method of the cooling chamber. This optimizes the cooling gas flow path design of the blade and provides a reasonable temperature distribution. Specifically, the cooling cavity located at the front of the blade can achieve impact-film cooling, which helps reduce the thermal load on the leading edge, pressure surface, and suction surface of the blade; the cooling cavity located in the middle of the blade mainly provides lower-temperature film cooling gas to the blade tail; and the rotating channels effectively improve the temperature distribution at the blade tail.

[0072] Based on the above description of the method, it can also be understood that a turbine blade and a turbine blade assembly are prepared by the above method. The turbine blade assembly includes a turbine blade and a sleeve disposed in at least one cold air cavity. The sleeve includes an air supply cavity and a cold air impact hole.

[0073] The turbine blades also include one or more cooling gas cavities 21 and cooling channels 22. The cooling channels 22 can be rotary cooling channels, including radial cooling gas inlets 221, radial flow sections 222, and end bend sections 223. The radial flow sections 222 are connected through the end bend sections 223. Figure 7 or Figure 11 As shown.

[0074] In addition, the cooling channel can be equipped with a turbulence structure on the inner wall of the blades on the pressure surface 24 and / or the suction surface 25 to enhance the heat exchange process of the cold air in the cooling channel 22.

[0075] For a multi-cavity cooling structure, the cooling air in the front cooling air cavity provides effective cooling protection for both the outer and inner surfaces of the blade through the impact holes and film cooling holes; the rotary channel located at the rear is closer to the blade trailing edge, allowing the cooling air to effectively reduce the internal and external temperatures of the trailing edge; the cooling air in the rear cooling air cavity has a lower temperature and flows directly out through the film cooling holes to provide effective film cooling protection for the trailing edge of the downstream blade, thereby providing a reasonable temperature distribution for the entire blade.

[0076] The turbine blades and turbine blade assemblies have excellent cooling performance.

[0077] This application uses specific terms to describe embodiments of the application. Terms such as "an embodiment," "one embodiment," and / or "some embodiments" refer to a particular feature, structure, or characteristic associated with at least one embodiment of the application. Therefore, it should be emphasized and noted that references to "an embodiment," "one embodiment," or "an alternative embodiment" in different locations throughout this specification do not necessarily refer to the same embodiment. Furthermore, certain features, structures, or characteristics in one or more embodiments of the application can be appropriately combined.

[0078] While the present invention has been disclosed above with reference to preferred embodiments, it is not intended to limit the invention. Any variations and modifications can be made by those skilled in the art without departing from the spirit and scope of the invention. Therefore, any modifications, equivalent changes, and alterations made to the above embodiments based on the technical essence of the present invention, without departing from the scope of the invention, fall within the protection scope defined by the claims of the present invention.

Claims

1. A method for manufacturing turbine blades with a cooling gas cavity, characterized in that, Includes the following steps: The inner cavity preform is formed by wrapping multiple layers of fiber cloth around the inner cavity core mold. The inner cavity preform and the filling preform are spliced ​​together; The multi-layer fiber cloth is wrapped around the outside of the inner cavity preform and the filling preform to obtain a blade preform with a blade shape, and then cured and shaped. Remove the inner cavity core mold to obtain a turbine blade with multiple cold air cavities.

2. The turbine blade manufacturing method according to claim 1, characterized in that, Prepare a radial channel prefabricated body and an embedded body, wherein the embedded body is embedded in the radial channel prefabricated body; The radial channel preform is assembled together with the inner cavity preform and the filling preform; The inner cavity preform, the filling preform, and the radial channel preform are wrapped with multi-layer fiber cloth to obtain a blade preform with a blade shape. Remove the pre-embedded body to form a cooling channel.

3. The turbine blade manufacturing method as described in claim 2, characterized in that, The radial channel prefabricated body includes an inlet channel, a turning channel, and a flow channel that are interconnected, and multiple pre-embedded bodies are embedded in the inlet channel, the turning channel, and the flow channel. The end of the inlet channel connects to the outside of the radial channel preform. Remove the pre-embedded body to form a rotary cooling channel.

4. The turbine blade manufacturing method according to claim 2, characterized in that, A hole is drilled in the embedded body, and the filling preform is filled into the hole; After removing the pre-embedded body, the filling prefabricated body forms a turbulence structure.

5. The turbine blade manufacturing method according to claim 2, characterized in that, A film cooling hole is machined on the turbine blade, the film cooling hole connecting the cold air cavity and / or the cooling structure.

6. The turbine blade manufacturing method according to claim 1, characterized in that, Multiple inner cavity preforms are spliced ​​together along the chord length direction of the blade shape.

7. A turbine blade, characterized in that, The turbine blades are prepared by the method described in any one of claims 1-6.

8. The turbine blade as claimed in claim 7, characterized in that, The turbine blades also include a rotary cooling channel, which includes a radial cold air inlet, a radial flow section, and an end bend section. Multiple radial flow sections are connected through the end bend section.

9. The turbine blade as claimed in claim 8, characterized in that, The rotary cooling channel is located in the middle and rear part of the blade along the chord length.

10. The turbine blade as claimed in claim 8, characterized in that, The turbine blades also include film cooling holes, which are connected to the rotary cooling channel and / or the cooling air cavity.

11. The turbine blade as claimed in claim 7, characterized in that, The turbine blade includes multiple cooling gas cavities arranged along the blade chord direction.

12. A turbine blade assembly, characterized in that, The turbine blades as described in any one of claims 7-11 are further comprising a sleeve disposed within at least one cooling gas cavity, the sleeve comprising a supply gas cavity and a cooling gas impact hole.

13. The turbine blade assembly as claimed in claim 12, characterized in that, The turbine blade includes multiple cooling gas cavities distributed along the blade chord length direction, and the sleeve is disposed in the first cooling gas cavity along the blade chord length direction. A rotary cooling channel is also provided on the rear side of the last cold air cavity along the blade chord direction of the turbine blade.

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