Manufacturing method of ceramic matrix composite turbine blade and turbine blade

Abstract

According to the manufacturing method of the ceramic matrix composite turbine blade, a blade inner core prefabricated part comprising a front edge sub-part, a middle sub-part and a tail edge sub-part is manufactured, the front edge sub-part and the tail edge sub-part are manufactured by pressing one-way belt molds arranged in the blade height direction, and the sectional areas of the front edge sub-part and the tail edge sub-part are kept unchanged at the positions of different blade heights; the middle sub-component is formed by mould pressing of a plurality of layers of ceramic fiber spread materials which are spread at different angles; laying a plurality of layers of ceramic fiber spreading materials on the peripheral side of the blade inner core prefabricated part to a preset external size, and carrying out mold pressing and low-temperature curing to obtain a blade prefabricated part; and the blade prefabricated part and the margin plate prefabricated part are spliced to obtain a turbine blade prefabricated part, and then high-temperature carbonization and molten silicon permeation densification treatment are conducted to obtain the finished ceramic matrix composite turbine blade. The method is simple in process and good in formability, and the forming quality and the production efficiency of the turbine blade can be improved. The invention further provides the ceramic matrix composite material turbine blade.

Description

Technical Field

[0001] This invention belongs to the field of aero-engines, specifically relating to a method for manufacturing a ceramic matrix composite turbine blade and the turbine blade itself. Background Technology

[0002] Ceramic matrix composites (CMCs) possess excellent high-temperature mechanical properties, significantly increasing the upper limit of operating temperature for aero-engine turbine blades compared to traditional high-temperature alloys. However, with increasing demands for engine power and efficiency, the structure of CMC turbine blades has become more complex. Newer engines feature smaller and more complex CMC turbine blades, leading to increased molding difficulties. Simultaneously, design requirements necessitate higher strength in the blade height direction, further complicating the design and molding of CMC fiber preforms. Existing CMC component manufacturing processes are increasingly unable to meet the performance and manufacturability requirements of aero-engine turbine blades. Therefore, providing a manufacturing method for ceramic matrix composite turbine blades that improves blade molding quality is of positive significance for enhancing aero-engine performance and reducing manufacturing costs. Summary of the Invention

[0003] The purpose of this invention is to provide a method for manufacturing ceramic matrix composite turbine blades, thereby improving the forming quality of ceramic matrix composite turbine blades. This invention also provides a ceramic matrix composite turbine blade.

[0004] According to one aspect of the present invention, a method for manufacturing a ceramic matrix composite turbine blade is provided. The method includes the following steps: manufacturing a blade core preform, the blade core preform including a leading edge sub-component, an intermediate sub-component, and a trailing edge sub-component, the interface of the leading edge sub-component, the intermediate sub-component, and the trailing edge sub-component extending along the blade height direction; wherein the leading edge sub-component and the trailing edge sub-component are molded from unidirectional strips arranged along the blade height direction, so that the cross-sectional area of ​​the leading edge sub-component and the trailing edge sub-component remains constant at different blade height positions; the intermediate sub-component is molded from multilayer ceramic fiber layup laid at different angles; the leading edge sub-component, the intermediate sub-component, and the trailing edge sub-component are spliced ​​together and integrally molded and low-temperature cured to form a whole. Multilayer ceramic fiber layup is laid at different angles on the outer periphery of the blade core preform to a preset external dimension, and then molded and low-temperature cured to obtain the blade preform. A rim plate preform is manufactured. The blade preform and the rim plate preform are spliced ​​together to obtain a turbine blade preform. The turbine blade preform is subjected to high-temperature carbonization and molten silicon infiltration densification treatment to obtain the finished ceramic matrix composite turbine blade.

[0005] The aforementioned method, by segmenting the blade core preform, effectively reduces the difficulty of design layup and molding, resulting in preforms with excellent density. The use of unidirectional strip layup in the leading and trailing edge regions ensures the highest strength performance of the turbine blade's leading and trailing edges in the blade height direction, improving the overall centrifugal strength of the finished blade. The intermediate components are manufactured using ceramic fiber layup in different directions, reducing the difficulty of layup design and improving both chordal strength and bending / torsional strength while maintaining strength in the blade height direction, enabling the turbine blade to meet strength requirements in different directions. After the blade core preform has solidified, winding ceramic fiber layup around the outer periphery further enhances the integrity of the fiber layup.

[0006] Furthermore, in some embodiments, the blade core preform also includes a tenon component. One end of the leading edge component, the intermediate component, and the trailing edge component is provided with a forked structure, and the tenon component is embedded in the forked structure. The forked structure ensures the integrity of the blade and the fibers in the tenon area in the blade height direction, thereby further improving the strength in the blade height direction. By embedding the tenon component, the difficulty of the layup design of the blade core preform can also be reduced, and the formability of the part can be improved.

[0007] Furthermore, in some embodiments, the rim plate preform includes two end plates. After the blade preform and the rim plate preform are spliced, the two end plates are connected to both ends of the tenon component. The fibers of the end plates extend to the tenon area, improving the bonding strength between the rim plate and the blade and the overall integrity of the blade, which is beneficial to improving the rim plate's ability to resist centrifugal loads.

[0008] Furthermore, in some embodiments, the interface between the leading edge sub-component, the intermediate sub-component, and the trailing edge sub-component is perpendicular to the outer surface of the blade inner core preform.

[0009] Furthermore, in some embodiments, the cross-sectional areas of the leading edge sub-component and the trailing edge sub-component account for 5%-40% of the total cross-sectional area of ​​the blade inner core preform, respectively.

[0010] Furthermore, in some embodiments, the intermediate sub-component is configured as a plate with variable thickness. Depending on the structural design of different turbine blades, the thickness of the turbine blade varies from the leading edge to the trailing edge. Using an independent intermediate sub-component to achieve the thickness transition helps avoid shape boundary requirements between the leading and trailing edges, reduces the difficulty of ply layup design, effectively reduces the number of small layers, improves manufacturing efficiency, and reduces manufacturing difficulty.

[0011] Furthermore, in some embodiments, the ceramic fiber layup is configured as a unidirectional ceramic fiber tape.

[0012] Furthermore, in some embodiments, the flange preform includes sealing wings disposed on the axial front and rear sides.

[0013] Furthermore, in some embodiments, after the rim plate preform is laid and shaped, a step of forming through holes is included, the through holes being used to allow the blade preform to pass through. The blade preform passes through the through holes and is spliced ​​and positioned with the rim plate preform through the through holes to facilitate subsequent overall processing and forming.

[0014] According to another aspect of the present invention, a ceramic matrix composite turbine blade is provided, the blade comprising an oriented structure formed after carbonization and melting of ceramic fiber layup, wherein the ceramic matrix composite turbine blade comprises an inner oriented structure and an outer oriented structure; wherein the inner oriented structure comprises a leading edge structure, an intermediate structure, and a trailing edge structure, the oriented structures in the leading edge structure and the trailing edge structure extending along the blade height direction of the ceramic matrix composite turbine blade and having the same cross-sectional area at different blade height positions, the intermediate structure comprising the oriented structures extending in different directions; the outer oriented structure surrounds the inner oriented structure circumferentially, the outer oriented structure comprising the oriented structures extending in different directions.

[0015] Furthermore, in some embodiments, the inner layer orientation structure further includes a tenon structure at one end of the tenon, the tenon structure having an extension direction different from that of the orientation structure in the leading edge structure, and the tenon structure having an extension direction different from that of the orientation structure in the trailing edge structure.

[0016] Furthermore, in some embodiments, the ceramic matrix composite turbine blade is manufactured using the ceramic matrix composite turbine blade manufacturing method provided in any of the foregoing embodiments. Attached Figure Description

[0017] Figure 1 This is a schematic diagram of a ceramic matrix composite turbine blade structure in one embodiment;

[0018] Figure 2a This is a schematic diagram of the blade inner core prefabricated component structure in one embodiment;

[0019] Figure 2b This is a front view of the blade inner core prefabricated structure in one embodiment;

[0020] Figure 2c for Figure 2b Schematic diagram of the AA section structure;

[0021] Figure 3 This is an exploded view of the blade core preform in one embodiment;

[0022] Figure 4aThis is a schematic diagram of the outer layer structure of a ceramic matrix composite turbine blade in one embodiment;

[0023] Figure 4b This is a front view of the outer layer structure of a ceramic matrix composite turbine blade in one embodiment;

[0024] Figure 4c for Figure 4b Schematic diagram of the BB section structure;

[0025] Figure 5 This is a schematic diagram of the prefabricated edge plate structure in one embodiment;

[0026] Figure 6 This is a schematic diagram of the cross-sectional structure of a ceramic matrix composite turbine blade in one embodiment.

[0027] Meaning of reference numerals in the attached diagram: 1-Blade body; 11-Outer blade layer; 111-Outer leading edge; 112-Outer trailing edge; 113-Outer tenon; 12-Inner blade core prefabricated component; 121-Leading edge component; 1211-Inner leading edge; 1212-Bifurcation structure; 122-Intermediate component; 1221-Bifurcation structure; 123-Tailing edge component; 1231-Inner trailing edge; 1232-Bifurcation structure; 124-Tenon component; 2-Flange plate; 21-Flange plate body; 211-Through hole; 22-Sealing wing; 23-End plate section; 3-Tenon.

[0028] The purpose of the above-described drawings is to provide a detailed description of the present invention so that those skilled in the art can understand the technical concept of the invention, and is not intended to limit the invention. For the sake of brevity, the above-described drawings only schematically depict the structures related to the technical features of the present invention, and do not depict the complete structure and all details strictly according to actual scale. Detailed Implementation

[0029] The present invention will now be described in further detail with reference to specific embodiments and accompanying drawings.

[0030] The term "embodiment" as used herein means that a particular feature, structure, or characteristic described in connection with an embodiment may be included in at least one embodiment herein. The phrase appearing in various places in the specification does not necessarily refer to the same embodiment, nor is it limited to mutually exclusive, independent, or alternative embodiments. Those skilled in the art will understand that the embodiments herein can be combined with other embodiments without causing structural conflicts.

[0031] In this description, unless otherwise expressly specified and limited, the technical terms "installation," "connection," "joining," etc., should be interpreted broadly, referring to movable connections, fixed connections, or integration. Those skilled in the art can understand the specific meaning of these terms in the embodiments of this application based on the specific circumstances.

[0032] In this description, terms such as "upper," "lower," "left," "right," "lateral," "longitudinal," "height," "length," and "width," which indicate orientation or positional relationships, are intended to accurately describe the embodiments and simplify the description, rather than limiting the parts or structures involved to have a specific orientation, or to be installed or operated in a specific orientation, and should not be construed as limiting the embodiments in this document.

[0033] In this description, terms such as "first" and "second" are used only to distinguish different objects and should not be construed as indicating relative importance or limiting the number, specific order, or primary / secondary relationship of the described technical features. In this description, "multiple" means at least two.

[0034] As the performance requirements of aero-engines continue to increase, the operating temperature of engine turbine blades has gradually approached or even exceeded the melting point of traditional high-temperature alloys such as nickel-based alloys, limiting further improvements in engine performance. Compared with metallic materials, ceramic matrix composites (CMCs) possess superior high-temperature mechanical properties. Using CMCs to manufacture turbine blades can further raise the upper limit of engine turbine operating temperatures, thus providing further optimization space for aero-engine performance and efficiency. However, CMC blades are often small in size, and to meet the aerodynamic performance requirements of turbine blades, they typically require complex geometries, which leads to certain difficulties in the design and molding of ceramic fiber preforms. Furthermore, due to the high-speed rotation of turbine blades during service, CMC blades require sufficient strength in the blade height direction to withstand the centrifugal force generated by high-speed rotation, which further restricts the design of fiber distribution in CMCs. These factors collectively lead to difficulties in the design and manufacturing of CMC turbine blades, limiting their application in aero-engines.

[0035] To address the aforementioned problems, one embodiment of the present invention provides a method for manufacturing ceramic matrix composite turbine blades, which can effectively improve the forming quality and finished product performance of CMC turbine blades, especially the strength performance in the blade height direction.

[0036] Based on the structural characteristics and performance requirements of CMC turbine blades, this method divides the turbine blades into sections, effectively reducing the forming difficulty of the turbine blades and enabling the fiber distribution to be optimized according to performance requirements during the turbine blade design and manufacturing process, so that the performance of CMC turbine blades can meet the design requirements.

[0037] Specifically, this method is used to manufacture structures such as Figure 1The CMC turbine blade shown is a CMC turbine blade. The CMC turbine blade includes a blade body 1, a shroud 2, and a tenon 3. During turbine blade operation, the blade body 1 plays a major aerodynamic role and also bears significant centrifugal force along the blade height direction. The shroud 2 serves a sealing function and also bears a certain amount of centrifugal force. The tenon 3 is used to connect and fix the turbine blade to the bladed disk. The manufacturing process of this blade is as follows:

[0038] Combination Figure 6 The turbine blade is manufactured in sections consisting of an outer blade layer 11, an inner blade core preform 12, and a rim plate 2. The outer blade layer 11 includes the outer portion of the blade body 1 and the tenon 3, while the inner blade core preform 12 includes the inner portion of the blade body 1 and the tenon 3.

[0039] First, the blade inner core preform 12 is manufactured.

[0040] like Figure 2a , Figure 2b and Figure 2c As shown, the blade inner core preform 12 is divided into a leading edge sub-component 121, an intermediate sub-component 122, and a trailing edge sub-component 123. The leading edge sub-component 121, the intermediate sub-component 122, and the trailing edge sub-component 123 are connected sequentially, and their respective interfaces extend along the blade height direction.

[0041] Due to the aerodynamic performance and strength requirements of turbine blades, the leading and trailing edges of turbine blades have complex geometric shapes. If they are manufactured using an integral lay-up method, the lay-up structure of the leading and trailing edge regions would be very complex, and molding would be difficult. Therefore, the leading edge sub-component 121 and trailing edge sub-component 123 in the blade core preform 12 are manufactured separately. They are made by molding using unidirectional ceramic fiber tapes arranged along the blade height direction, so that the cross-sectional area of ​​the leading edge sub-component 121 and trailing edge sub-component 123 remains constant at different positions along the blade height direction. Specifically, it should be understood that extending along the blade height direction means that the distribution and orientation of these unidirectional bands are consistent with the extension direction of the leading and trailing edges of the CMC blade profile along the blade height direction. This means that the unidirectional bands do not necessarily extend in a straight line along the blade height direction, but may be curved or tilted depending on the CMC turbine blade structure design. The unidirectional bands in the leading edge sub-component 121 and the trailing edge sub-component 123 may also not be parallel. Maintaining a constant cross-sectional area means that, within the range considering process and molding errors, the cross-sectional area at different blade height positions remains constant. In a preferred embodiment, the leading edge sub-component 121 and the trailing edge sub-component 123 are specifically obtained by compression molding using unidirectional fibers with the same number of filaments. This effectively reduces the molding difficulty of the leading edge sub-component 121 and the trailing edge sub-component 123, avoiding complex layup designs, especially the use of small layers in these areas, enabling molding and manufacturing through conventional compression molding processes. On the other hand, the unidirectional strips arranged along the blade height direction effectively improve the consistency of composite material fibers in the inner leading edge 1211 and the inner trailing edge 1231, enhancing the tensile strength in the blade height direction, thereby improving the ability of the finished blade to withstand centrifugal loads.

[0042] Between the leading-edge sub-component 121 and the trailing-edge sub-component 123 is an intermediate sub-component 122. The intermediate sub-component 122 is manufactured by compression molding of multi-layer ceramic fiber layup with different angles; in a preferred embodiment, it is formed by layup of unidirectional ceramic fiber strips. In a preferred embodiment, according to the turbine blade's structural design, the intermediate sub-component 122 is configured as a variable-thickness plate with curvature, serving as a transition connection from the leading edge to the trailing edge of the blade. Because it is structurally independent of the leading-edge sub-component 121 and the trailing-edge sub-component 123, the intermediate sub-component 122 has greater freedom in its layup design; at different blade heights, the intermediate sub-component 122 can have different thicknesses and widths. The unidirectional strips with different layup angles allow the area containing the intermediate sub-component 122 to better withstand chordal tensile and bending-torsional loads.

[0043] In a preferred embodiment, the connection positions between the leading edge sub-component 121, the intermediate sub-component 122 and the trailing edge sub-component 123 are smoothly transitioned, and the interface is perpendicular to the outer surface of the blade inner core preform 12, so as to reduce the molding difficulty of the CMC turbine blade finished product and improve the aerodynamic performance of the CMC turbine blade finished product.

[0044] In a further preferred embodiment, depending on the overall structure of the CMC turbine blade, in any cross-section, the cross-sectional area occupied by the leading edge sub-piece 121 and the trailing edge sub-piece 123 is 5%-40% of the total cross-sectional area of ​​the blade inner core preform 12.

[0045] In a preferred embodiment, combined with Figure 3 At one end of the blade tenon 3, the blade inner core prefabricated part 12 is also provided with a tenon sub-part 124. The ends of the leading edge sub-part 121, the middle sub-part 122 and the trailing edge sub-part 123 extend separately towards the suction side and the pressure side to form a forked structure 1212, a forked structure 1221 and a forked structure 1232, respectively, and the tenon sub-part 124 is embedded in the above-mentioned forked structures. Specifically, the bifurcated structures 1212 and 1232 are formed by one half of the unidirectional belts that make up the leading edge sub-component 121 and the trailing edge sub-component 123 branching off towards the pressure side and the other half towards the suction side of the CMC turbine blade, respectively, ensuring the continuity and cross-sectional area of ​​the unidirectional belts remain unchanged (taking the leading edge sub-component 121 as an example, the cross-sectional area of ​​the unbranched portion of the leading edge sub-component 121 is equal to the sum of the areas of the suction side and the pressure side portions in any cross-section of the bifurcated structure 1212); the intermediate sub-component 122 is formed by partially plying and bending towards the suction side of the CMC turbine blade, and partially plying and bending towards the pressure difference of the CMC turbine blade. The tenon sub-component 124 passes through the leading edge sub-component 121, the intermediate sub-component 122, and the trailing edge sub-component 123 along the rotor axis to improve the axial integrity of the blade inner core preform 12. In different embodiments, the tenon component 124 can be made by unidirectional belt molding arranged along the rotor axis, or by molding ceramic fiber laid up at different angles.

[0046] The various sub-components of the blade inner core preform 12 are assembled together, molded and cured at low temperature, so that the sub-components are connected into a whole, and a complete blade inner core preform 12 is obtained.

[0047] Using the blade inner core preform 12 as a core mold, multiple layers of ceramic fiber lay-up at different angles are wound and laid on the outer periphery of the blade inner core preform 12 to obtain the structure as shown. Figure 4a , Figure 4b and Figure 4c The blade outer layer 11 shown is substantially the same in shape and size as the finished CMC blade after further molding. In some embodiments, the ceramic fiber layup can also be a unidirectional ceramic fiber strip. According to the design requirements of the CMC turbine blade, the ceramic fibers used in the blade inner core preform 12 and the blade outer layer 11 can be the same type of fiber or different types of fiber. The ceramic fiber layup in the blade outer layer 11 completely wraps around the blade inner core preform 12 along the blade axial direction to strengthen the fiber bonding strength in the circumferential direction. Subsequently, it is molded again and cured at low temperature to obtain the blade preform.

[0048] Next, manufacturing such Figure 5 The flange 2 shown includes a flange body 21 with a through hole 211. The through hole 211 allows the portion of the blade preform corresponding to the blade body 1 to pass through and be fixed at the position where the blade body 1 is connected to the tenon 3. The through hole 211 can be formed by machining after the flange body 21 is formed. Two end plate portions 23 extend from the side of the flange body 21 toward the blade disk. After the blade preform is inserted into the through hole 211 and fixed, the two end plate portions 23 are connected to the end of the tenon piece 124 (the end plate portions 23 and the end of the tenon piece 124 are separated by the corresponding portion of the outer layer 11 of the blade). The end plate portions 23 are formed by the extension of some fibers in the flange body 21, which can improve the integrity of the fibers in the blade height direction and improve the flange 2's ability to withstand centrifugal loads.

[0049] In a preferred embodiment, sealing wings 22 are formed on the front and rear sides of the flange 2 along the rotor axis, and the sealing wings 22 can improve the sealing effect of the flange.

[0050] After the rim plate 2 is assembled and spliced ​​with the blade preform, the turbine blade preform is obtained. According to the melting and infiltration process, the turbine blade preform is subjected to high-temperature carbonization treatment, followed by molten silicon infiltration densification reaction, and finally the desired ceramic matrix composite turbine blade is obtained.

[0051] The ceramic matrix composite turbine blade manufacturing method provided in the above embodiments effectively simplifies the manufacturing process of CMC turbine blades with complex structures and precision dimensions by using regional ceramic fiber layup and compression molding, as well as sectional assembly compression molding and low-temperature curing connection. After decomposition, the structure of each sub-component is relatively simple, the layup design and compression molding process are effectively simplified, the preform molding operation is simple and the molding efficiency is high, the molding quality of each area of ​​the blade is good, and the finished blade has higher strength performance in the blade height direction.

[0052] The ceramic matrix composite turbine blades manufactured using the above method exhibit different orientations in different regions after carbonization and melt infiltration due to the different layup methods used for the ceramic fibers in each component of the preform. Figure 6The distribution of the orientation structure in the finished CMC turbine blade is the same as that in the outer layer 11, the inner core preform 12, and the rim plate 2. The inner core preform 12 forms an inner orientation structure, which is further divided into leading-edge, intermediate, and trailing-edge structures. The leading-edge structure is formed by the leading-edge sub-piece 121, the intermediate structure by the intermediate sub-piece 122, and the trailing-edge structure by the trailing-edge sub-piece 123. The orientation structures in the leading-edge and trailing-edge sub-pieces 121 and 123 extend along the blade height direction of the CMC turbine blade and have the same cross-sectional area at different blade height positions. Since the intermediate sub-piece 122 is made of ceramic fiber layup in different directions, the intermediate structure includes orientation structures extending in different directions. The outer layer 11 forms an outer orientation structure, which surrounds the inner orientation structure circumferentially and includes orientation structures extending in different directions. In the area where the tenon 3 is located, the tenon part 124 forms a tenon structure, and the orientation structure in the tenon structure has a different extension direction from the leading edge structure and the trailing edge structure.

[0053] The leading and trailing edge structures extend along the blade height direction, giving the CMC turbine blade good mechanical properties in the blade height direction and the ability to withstand centrifugal loads generated by the high-speed rotation of the rotor. Meanwhile, the orientation structures extending in different directions in the intermediate and outer orientation structures provide good tangential strength and resistance to bending and torsion, and give the CMC turbine blade good integrity, enabling it to serve stably for a long time under more severe operating conditions.

[0054] The purpose of the above embodiments is to provide a further detailed description of the present invention in conjunction with the accompanying drawings, so that those skilled in the art can understand the technical concept of the present invention. Within the scope of the present invention, optimization or equivalent substitution of the method steps and component structures involved, as well as combination of implementation methods in different embodiments without causing structural and principle conflicts, all fall within the protection scope of the present invention.

Claims

1. A method for manufacturing a ceramic matrix composite turbine blade, characterized in that, Includes the following steps: A blade core preform is manufactured, comprising a leading edge component, an intermediate component, and a trailing edge component. The interfaces of the leading edge component, the intermediate component, and the trailing edge component extend along the blade height direction. The leading edge component and the trailing edge component are molded from unidirectional strips arranged along the blade height direction to ensure that the cross-sectional area of ​​the leading edge component and the trailing edge component remains constant at different blade height positions. The intermediate component is molded from multilayer ceramic fiber layup at different angles. The leading edge component, the intermediate component, and the trailing edge component are then spliced ​​together and integrally molded and cured at low temperature to form a whole. Multiple layers of ceramic fiber material are laid at different angles on the outer periphery of the blade core preform to a preset external size, and then molded and cured at low temperature to obtain the blade preform. Manufacturing prefabricated flange components; The blade preform is spliced ​​with the rim plate preform to obtain the turbine blade preform; The turbine blade preform is subjected to high-temperature carbonization and molten silicon infiltration densification treatment to obtain the finished ceramic matrix composite turbine blade.

2. The method for manufacturing ceramic matrix composite turbine blades according to claim 1, characterized in that, The blade inner core prefabricated component also includes a tenon component. One end of the leading edge component, the intermediate component, and the trailing edge component is provided with a forked structure, and the tenon component is embedded in the forked structure.

3. The method for manufacturing ceramic matrix composite turbine blades according to claim 2, characterized in that, The flange prefabricated component includes two end plates. After the blade prefabricated component and the flange prefabricated component are spliced ​​together, the two end plates are connected to both ends of the tenon component.

4. The method for manufacturing ceramic matrix composite turbine blades according to claim 1, 2, or 3, characterized in that, The interface between the leading edge sub-component, the intermediate sub-component, and the trailing edge sub-component is perpendicular to the outer surface of the blade inner core preform.

5. The method for manufacturing ceramic matrix composite turbine blades according to claim 1, 2, or 3, characterized in that, The cross-sectional areas of the leading edge sub-component and the trailing edge sub-component account for 5%-40% of the total cross-sectional area of ​​the blade inner core preform, respectively.

6. The method for manufacturing ceramic matrix composite turbine blades according to claim 1, 2, or 3, characterized in that, The intermediate component is configured as a plate with variable thickness.

7. The method for manufacturing ceramic matrix composite turbine blades according to claim 1, 2, or 3, characterized in that, The ceramic fiber mulch is configured as a unidirectional ceramic fiber tape.

8. The method for manufacturing ceramic matrix composite turbine blades according to claim 1, 2, or 3, characterized in that, The flange prefabricated component includes sealing wings disposed on the front and rear sides in the axial direction.

9. The method for manufacturing ceramic matrix composite turbine blades according to claim 1, 2, or 3, characterized in that, After the blade preform is laid and shaped, the process also includes a step of forming through holes, which are used to allow the blade preform to pass through.

10. A ceramic matrix composite turbine blade, comprising an oriented microstructure formed after carbonization and melting of ceramic fiber layup, characterized in that, The ceramic matrix composite turbine blade includes an inner oriented structure and an outer oriented structure; wherein, the inner oriented structure includes a leading edge structure, a middle structure, and a trailing edge structure, the oriented structures in the leading edge structure and the trailing edge structure extend along the blade height direction of the ceramic matrix composite turbine blade and have the same cross-sectional area at different blade height positions, the middle structure includes the oriented structures extending in different directions; the outer oriented structure surrounds the inner oriented structure circumferentially, the outer oriented structure includes the oriented structures extending in different directions.

11. The ceramic matrix composite turbine blade according to claim 10, characterized in that, The inner oriented structure also includes a tenon structure at one end of the tenon. The tenon structure extends in a different direction than the oriented structure in the leading edge structure, and the tenon structure extends in a different direction than the oriented structure in the trailing edge structure.

12. The ceramic matrix composite turbine blade according to claim 10 or 11, characterized in that, The ceramic matrix composite turbine blade is manufactured using the manufacturing method described in any one of claims 1 to 9.

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