A combined dissimilar material turbine stator blade assembly

Abstract

The present invention belongs to the technical field of aero engines, and particularly relates to a combined dissimilar material turbine stator blade assembly. In the present invention, one end that is subjected to high-temperature gas erosion is made of a ceramic matrix blade, while a metal blade is used for the high-flow velocity diversion region with a complex structure and high processing difficulty. The present invention combines the excellent temperature resistance of the ceramic matrix composite material blade, the reliable mechanical properties of the superalloy blade, and the excellent aerodynamic performance, improves the temperature-bearing capacity of the leading edge of the high-pressure turbine guide vane, solves the problem of overheating of the leading edge of the turbine guide vane of a new generation of advanced engines, realizes the light weight of the turbine blade, extends the working life of the blade, and generally improves the overall performance of the aero engine.

Description

Technical Field

[0001] The present invention belongs to the technical field of aeroengines, and particularly relates to a combined dissimilar material turbine stator blade assembly. Background Art

[0002] The turbine guide vane converts part of the thermal energy of the high-temperature gas flowing through it into kinetic energy and changes the gas flow direction to meet the inlet requirements of the working blade. It is a key component determining the turbine efficiency and an important part of the turbine component of an aeroengine. With the development of aeroengines, the thrust-to-weight ratio index and the turbine inlet temperature of the engine are constantly increasing. The more severe high temperature and the high thrust-to-weight ratio requirement of the engine have brought problems such as a decrease in the life of the turbine component and an overweight engine. The leading edge of the high-pressure turbine guide vane is directly scoured by the high-temperature gas, and the over-temperature problem at the leading edge position of the blade is serious. Ceramic matrix composites have characteristics such as low density and outstanding temperature resistance. Using them as the material of the turbine guide vane can better solve problems such as over-temperature and excessive weight at the leading edge of the engine turbine guide vane. However, limited by the forming and manufacturing process capabilities of ceramic matrix composites, it is difficult to manufacture ceramic matrix high-pressure turbine guide vanes with large turning of upper and lower flange plates and complex internal cavities. At the same time, the middle and rear parts of the blade are high-flow regions, and the surface quality of the blade profile has a great impact on the aerodynamic efficiency. Moreover, the greater the trailing edge thickness of the blade body, the greater the wake loss of the blade, which has a decisive impact on the turbine efficiency. Usually, the maximum trailing edge thickness requirement of the blade body is ≯1.2 mm. However, limited by the material characteristics and process level of ceramic matrix composites, it is difficult to meet the requirements of the profile quality and trailing edge thickness in the middle and rear parts of the blade. Summary of the Invention

[0003] In view of this, the present invention proposes a combined dissimilar material turbine stator blade assembly, using the end of the ceramic matrix blade to withstand the scouring of high-temperature gas, while using a metal blade for the high-flow guide region with complex structure and large processing difficulty. The present invention combines the superior temperature resistance of the ceramic matrix composite blade and the reliable mechanical properties and excellent aerodynamic performance of the superalloy blade, improves the temperature-bearing capacity of the leading edge of the high-pressure turbine guide vane, solves the problem of over-temperature at the leading edge of the turbine guide vane of the new generation of advanced engines, realizes the lightweight of the turbine blade, extends the working life of the blade, and generally improves the overall performance of the aeroengine.

[0004] In order to achieve the above technical objectives, the specific technical solutions adopted by the present invention are as follows:

[0005] A combined dissimilar material turbine stator blade assembly, comprising:

[0006] A ceramic matrix blade, including a ceramic matrix blade body; the ceramic matrix blade body is a hollow air-cooled structure of ceramic matrix, is arranged at one end of the turbine stator blade that withstands the scouring of high-temperature gas, and the windward surface faces the incoming gas direction of the high-temperature gas;

[0007] A metal blade, including a metal blade body based on metal; the metal blade body is a hollow double-wall air-cooled structure, and the two blade body walls form a blade body structure with a hollow and gradually changing thickness; the front end with a large thickness of the metal blade body is spliced with the leeward surface of the ceramic-based blade body, and the converging tail end with a small thickness of the metal blade body is the high-flow-rate guiding area of the turbine stator blade;

[0008] Metal upper edge plates; a plurality of the metal upper edge plates are sequentially spliced to form an outer ring structure;

[0009] Metal lower edge plates; a plurality of the metal lower edge plates are sequentially spliced to form an inner ring structure;

[0010] Wherein: the ceramic-based blade and the metal blade are combined into one turbine stator blade; the outer ring structure and the inner ring structure are coaxially arranged; the cavity between the outer ring structure and the inner ring structure forms a part of the diversion cavity; a plurality of the turbine stator blades are symmetrically and uniformly arranged in the diversion cavity with the central axis of the diversion cavity as the center.

[0011] Further, the combined dissimilar material turbine stator blade assembly further includes an upper edge plate and upper hanging pins; the upper edge plate is spliced at one end of the metal upper edge plate in the incoming gas direction of the high-temperature gas; a plurality of the upper edge plates are sequentially spliced to form a front end ring; the front end ring is coaxially arranged with the diversion cavity; the inner ring surface of the front end ring and the metal lower edge plate form another part of the diversion cavity;

[0012] The ceramic-based blade is provided with an upper mounting edge fixed to the ceramic-based blade body on the top surface that does not contact the high-temperature gas and the metal blade body; the ceramic-based blade is spliced between two adjacent upper edge plates based on the upper mounting edge;

[0013] The upper hanging top is used to fix two adjacent upper edge plates at the outer ring surface of the front end ring.

[0014] Further, the ceramic-based blade is provided with a lower positioning boss fixed to the ceramic-based blade body on the bottom surface that does not contact the high-temperature gas and the metal blade body;

[0015] The metal lower edge plate is provided with a groove at one end in the incoming gas direction of the high-temperature gas

[0016] The ceramic-based blade is spliced in the groove based on the positioning boss.

[0017] Furthermore, a cold air cavity is provided inside the ceramic matrix blade body; the wall thickness of the ceramic matrix housing of the ceramic matrix blade body is uniform and is formed by weaving a ceramic matrix composite material. A plurality of ceramic matrix blade body film holes for air conduction between the cold air cavity and the diversion cavity are provided on the main body, and an air inlet channel is provided at the upper mounting edge.

[0018] Furthermore, the wall thickness of the ceramic matrix housing is 2-3 mm; the aperture of the ceramic matrix blade body film hole is less than 0.6 mm.

[0019] Furthermore, the metal blade is integrally cast from a single crystal superalloy at high temperature; a gas collecting cavity is provided inside the metal blade body, and a plurality of impact cavities are provided at the interlayer between the double-layer walls;

[0020] The impact cavity is in air communication with the gas collecting cavity through the blade body impact holes;

[0021] A plurality of metal blade body film holes are provided at the blade basin and blade back of the metal blade body in contact with the high-temperature gas; the metal blade body film holes are in air communication with the impact cavity and the diversion cavity;

[0022] A tail split seam is provided at the convergent tail end of the metal blade body; the tail split seam is in air communication with the impact cavity and the diversion cavity.

[0023] Furthermore, the aperture of the blade body impact hole is 1-1.5 mm; the aperture of the metal blade body film hole is less than 0.5 mm.

[0024] Furthermore, a plurality of front-end impact holes are provided at the splicing portion of the metal blade body and the ceramic matrix blade; the front-end film holes are in air communication with the gas collecting cavity.

[0025] Furthermore, a plurality of front-end film holes are provided at the splicing portion of the metal blade body and the ceramic matrix blade; one end of the front-end film hole is in air communication with the splicing portion between the metal blade body and the ceramic matrix blade, and the other end is in communication with the diversion cavity.

[0026] Furthermore, the metal blade is spliced between the metal upper edge plate and the metal lower edge plate based on the top surface and the bottom surface that do not contact the high-temperature gas and the metal blade body.

[0027] Adopting the above technical solution, the present invention can bring the following beneficial effects:

[0028] a) Compared with a blade made of a single superalloy material, the present invention can effectively increase the temperature resistance level by more than 200 °C;

[0029] b) Compared with a single crystal superalloy integrally cast blade, the present invention can reduce the weight by more than 20%;

[0030] c) Compared with metal blades, the present invention can reduce the cooling air consumption of turbine guide vanes and improve the engine efficiency;

[0031] d) The ceramic matrix composite part of the present invention has a simple structure and good manufacturability.

[0032] e) The blade structure of the assembly method of the present invention is reliable, has good assemblability, and can effectively solve the thermal matching problem between the ceramic matrix and metal parts. Description of the Drawings

[0033] In order to more clearly illustrate the technical solutions of the embodiments of the present disclosure, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings in the following description are only some embodiments of the present disclosure. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without creative efforts.

[0034] Figure 1 It is an exploded view of a combined dissimilar material turbine stator vane assembly in a specific embodiment of the present invention;

[0035] Figure 2 It is an assembled view of a combined dissimilar material turbine stator vane assembly in a specific embodiment of the present invention;

[0036] Figure 3 It is an axial sectional view of a combined dissimilar material turbine stator vane assembly in a specific embodiment of the present invention;

[0037] Figure 4 It is a radial sectional view of a combined dissimilar material turbine stator vane assembly in a specific embodiment of the present invention;

[0038] Wherein: 1, ceramic matrix vane; 101, ceramic matrix blade body; 102, positioning boss; 103, upper mounting edge; 104, mounting hole; 105, ceramic matrix blade body film hole; 106, intake passage; 107, cooling air cavity; 108, ceramic matrix fiber; 2, metal vane; 201, metal blade body; 202, metal upper flange; 203, metal lower flange; 204, groove; 205, front end impact hole; 205a, blade body impact hole; 206, metal blade body film hole; 206a, front end film hole; 207, tail split seam; 208, gas collecting cavity; 209, impact cavity; 3, upper flange; 301, mounting boss; 302, mounting edge; 303, upper flange bolt hole; 304, inner ring surface; 4, upper hanging ring; 401, upper hanging ring bolt hole; 5, screw. Detailed Embodiments

[0039] The embodiments of the present disclosure will be described in detail below with reference to the drawings.

[0040] The following describes the embodiments of the present disclosure through specific specific examples. Those skilled in the art can easily understand other advantages and effects of the present disclosure from the content disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of the present disclosure, rather than all the embodiments. The present disclosure can also be implemented or applied through other different specific implementation manners. Various details in this specification can also be modified or changed based on different viewpoints and applications without departing from the spirit of the present disclosure. It should be noted that, without conflict, the following embodiments and the features in the embodiments can be combined with each other. All other embodiments obtained by those of ordinary skill in the art based on the embodiments in the present disclosure without creative efforts belong to the scope of protection of the present disclosure.

[0041] It should be noted that the following describes various aspects of the embodiments within the scope of the appended claims. It should be obvious that the aspects described herein can be embodied in a wide variety of forms, and any specific structure and / or function described herein is illustrative only. Based on the present disclosure, those skilled in the art should understand that one aspect described herein can be implemented independently of any other aspect, and two or more of these aspects can be combined in various ways. For example, any number of aspects described herein can be used to implement the device and / or practice the method. Additionally, this device and / or this method can be implemented using other structures and / or functionality in addition to one or more of the aspects described herein.

[0042] It should also be noted that the diagrams provided in the following embodiments only illustrate the basic concept of the present disclosure schematically. The diagrams only show the components related to the present disclosure, rather than being drawn according to the number, shape, and size of the components in actual implementation. The type, quantity, and ratio of each component in actual implementation can be an arbitrary change, and the component layout type may also be more complex.

[0043] In addition, in the following description, specific details are provided to facilitate a thorough understanding of the examples. However, those skilled in the art will understand that the described aspects can be practiced without these specific details.

[0044] In an embodiment of the present invention, a combined dissimilar material turbine stator blade assembly is proposed, as Figures 1-4 shown, including:

[0045] A ceramic matrix blade 1, including a ceramic matrix blade body 101; the ceramic matrix blade body 101 is a ceramic matrix hollow air-cooled structure, arranged at one end of the turbine stator blade that bears the erosion of high-temperature gas, and the windward surface faces the incoming gas direction of the high-temperature gas;

[0046] The metal blade 2 includes a metal blade body 201 based on a metal matrix; the metal blade body 201 is a hollow double-wall air-cooled structure, and the two blade body walls form a blade body structure with a hollow and gradually changing thickness; the thick front end of the metal blade body 201 is spliced with the leeward surface of the ceramic matrix blade body 101, and the thin and convergent tail end of the metal blade body 201 is the high-flow-rate guiding area of the turbine stator blade;

[0047] Metal upper edge plates 202; a plurality of metal upper edge plates 202 are spliced in sequence to form an outer ring structure;

[0048] Metal lower edge plates 203; a plurality of metal lower edge plates 203 are spliced in sequence to form an inner ring structure;

[0049] Wherein: the ceramic matrix blade 1 and the metal blade 2 are combined into a turbine stator blade; the outer ring structure and the inner ring structure are coaxially arranged; the cavity between the outer ring structure and the inner ring structure forms a part of the guiding cavity; a plurality of turbine stator blades are symmetrically and evenly arranged in the guiding cavity with the central axis of the guiding cavity as the center.

[0050] In this embodiment, as Figure 1 , 2 shown, the combined dissimilar material turbine stator blade assembly further includes an upper edge plate 3 and upper hanging pins; the upper edge plate 3 is spliced at one end of the metal upper edge plate 202 in the gas inlet direction of the high-temperature gas; a plurality of upper edge plates 3 are spliced in sequence to form a front end ring; the front end ring is coaxially arranged with the guiding cavity; the inner ring surface 304 of the front end ring and the metal lower edge plate 203 form another part of the guiding cavity;

[0051] The ceramic matrix blade 1 is provided with an upper mounting edge 103 that is fixedly connected to the ceramic matrix blade body 101 on the top surface that does not contact the high-temperature gas and the metal blade body 201; the ceramic matrix blade 1 is spliced between two adjacent upper edge plates 3 based on the upper mounting edge 103;

[0052] The upper hanging top is used to fix two adjacent upper edge plates 3 at the outer ring surface of the front end ring.

[0053] In this embodiment, as Figure 1 shown, the ceramic matrix blade 1 is provided with a lower positioning boss 102 that is fixedly connected to the ceramic matrix blade body 101 on the bottom surface that does not contact the high-temperature gas and the metal blade body 201;

[0054] The metal lower edge plate 203 is provided with a groove 204 at one end in the gas inlet direction of the high-temperature gas

[0055] The ceramic matrix blade 1 is spliced into the groove 204 based on the positioning boss 102.

[0056] In this embodiment, as Figure 3As shown, a cooling air cavity 107 is provided inside the ceramic matrix blade body 101; the wall thickness of the ceramic matrix housing of the ceramic matrix blade body 101 is uniform, which is formed by weaving ceramic matrix composite materials. A plurality of ceramic matrix blade body film holes 105 that are in gas communication with the cooling air cavity 107 are provided on the main body, and an air inlet passage 106 is provided at the upper mounting edge 103.

[0057] In this embodiment, the wall thickness of the ceramic matrix housing is 2 - 3 mm; the aperture of the ceramic matrix blade body film hole 105 is less than 0.6 mm.

[0058] In this embodiment, the metal blade 2 is integrally cast from a single crystal superalloy at high temperature; as Figure 3 shown, a gas collecting cavity 208 is provided inside the metal blade body 201; a plurality of impact cavities 209 are provided at the sandwich between the double - layer walls;

[0059] The impact cavity 209 is in gas communication with the gas collecting cavity 208 through the blade body impact holes 205a;

[0060] A plurality of metal blade body film holes 206 are provided at the blade basin and blade back of the metal blade body 201 that are in contact with the high - temperature gas; the metal blade body film holes 206 are in gas communication with the impact cavity 209;

[0061] A tail split seam 207 is provided at the convergent tail end of the metal blade body 201; the tail split seam 207 is in gas communication with the impact cavity 209.

[0062] In this embodiment, the aperture of the blade body impact hole 205a is 1 - 1.5 mm; the aperture of the metal blade body film hole 206 is less than 0.5 mm.

[0063] In this embodiment, a plurality of front - end impact holes 205 are provided at the splicing joint of the metal blade body 201 and the ceramic matrix blade 1; the front - end film hole 206a is in gas communication with the gas collecting cavity 208.

[0064] In this embodiment, a plurality of front - end film holes 206a are provided at the splicing joint of the metal blade body 201 and the ceramic matrix blade 1; one end of the front - end film hole 206a is in gas communication with the splicing joint between the metal blade body 201 and the ceramic matrix blade 1, and the other end is in communication with the diversion cavity.

[0065] In this embodiment, the metal blade 2 is spliced between the metal upper edge plate 202 and the metal lower edge plate 203 based on the top surface and bottom surface that do not contact the high - temperature gas and the metal blade body 201.

[0066] The first half of the combined dissimilar - material turbine stator blade assembly in this embodiment is a hollow air - cooled ceramic matrix blade 1, which is formed by integrally weaving ceramic matrix fibers 108 and adding a ceramic matrix, and is mainly used to withstand the erosion of the leading - edge high - temperature gas; as Figure 2 、 4As shown, the ceramic matrix blade 2 matches with the upper rim plate 3 to form the front part of the upstream flow path surface, and is bolted to the upper suspension ring 4 to combine the ceramic matrix blades 2 into a ring;

[0067] The second half of the combined dissimilar material turbine stator blade assembly is a double-wall high-efficiency cooling structure air-cooled metal blade 2 with upper and lower rim plates, which is integrally cast from single crystal superalloy and is mainly used for guiding the mainstream gas of high-temperature gas, ensuring excellent aerodynamic efficiency of the blade, and ensuring reliable transmission of the axial force of the blade to the inner and outer casings; the ceramic matrix blade body 101 and the metal blade body 201 are combined to form a complete aerodynamic profile of the turbine guide blade; and a complete diversion cavity flow path is formed after the inner ring surface 304 of the upper rim plate 3, the metal upper rim plate 202 and the metal lower rim plate 203 are combined into a ring.

[0068] The ceramic matrix blade 1 is mainly composed of a ceramic matrix blade body 101, a lower end positioning boss 102, and an upper mounting edge 103. A simple cold air cavity 107 is arranged inside the ceramic matrix blade body 101, and the cross-section is provided with a uniform wall thickness as required, and the wall thickness is generally 2-3 mm; an air inlet channel 106 is arranged at the top of the ceramic matrix blade body 101, and dense ceramic matrix blade body air film holes 105 are arranged on the main body, and the diameter of the ceramic matrix blade body air film holes 105 is generally not more than 0.6 mm; cold air enters the cold air cavity 107 through the air inlet channel 106 and is discharged through the ceramic matrix blade body air film holes 105 to realize convective and air film cooling of the ceramic matrix blade body 101.

[0069] The metal blade 2 is mainly composed of a metal blade body 201, a metal upper rim plate 202, and a metal lower rim plate 203. The inside of the metal blade body 201 is a double-wall high-efficiency cooling structure, with a large gas collecting cavity 208 in the middle, and several impact cavities 209 are arranged in the interlayer between the inner and outer walls. The gas collecting cavity 208 and the impact cavity 209 are connected by several blade body impact holes 205a, and the diameter of the blade body impact holes is generally 1-1.5 mm; several rows of dense metal blade body air film holes 206 are arranged on both the blade concave and convex sides of the metal blade body 201, and the diameter of the metal blade body air film holes 206 is generally not more than 0.5 mm. Several front end impact holes 205 are arranged at the front end of the metal blade body 201, and a trailing edge split 207 is arranged at the trailing edge. Most of the cold air enters the impact cavity 209 through the blade body impact holes 205a on the inner layer wall from the middle gas collecting cavity 208, while impinging on the outer wall surface of the metal blade body 201, and is discharged through the dense metal blade body air film holes 206 and the trailing edge split 207 at the corresponding positions, and sufficient cooling of the concave, convex sides and trailing edge of the metal blade body 201 is realized through cold air impingement and air film coverage; another stream of cold air in the gas collecting cavity 208 is discharged through the front end impact holes 205 to cool and seal the cavity formed by the ceramic matrix blade body 101 and the metal blade body 201, and is discharged through the front end air film holes 206a on both sides.

[0070] A positioning boss 102 is provided at the lower end of the ceramic matrix blade 1. The positioning boss can be set in the form of a keyway, a convex tooth, a spherical head, etc. according to requirements. A groove 204 adapted to the positioning boss 102 is provided on the metal lower edge plate 203. The lower end of the ceramic matrix blade 1 is nested and installed in the groove 204, and a gap of 0.1 - 0.3 mm is left at the mating part to ensure that the lower end of the ceramic matrix blade 1 has a certain amount of movement space to solve the thermal matching problem between the ceramic matrix blade 1 and the metal part.

[0071] An installation edge 103 is provided at the upper end of the ceramic matrix blade 1 along the axial direction, and an installation through hole 104 is opened on the installation edge 103. The installation hole can be set in structural forms such as circular, racetrack-shaped, dovetail-shaped, etc. according to requirements; an adapted installation boss 301 is provided on one side of the upper edge plate 3, and the ceramic matrix blade 1 is connected and fixed to the upper edge plate 3 through the installation structure; four installation edges 302 are respectively provided at the four corners of the upper edge plate 3, and upper edge plate bolt holes 303 are respectively provided on the installation edges. Four upper hanging ring bolt holes 401 are correspondingly provided on the upper hanging ring 4. The installation edges 302 between adjacent upper edge plates 3 are paired in pairs, and the upper hanging ring 4 is installed and connected to a group of upper edge plates 3 through screws 5; after sequential assembly, the blades are combined into a ring to form a complete turbine guide vane.

[0072] As described above, it is only the specific implementation manner of the present disclosure, but the protection scope of the present disclosure is not limited thereto. Any changes or substitutions that can be easily thought of by those skilled in the art within the technical scope disclosed by the present disclosure should be covered by the protection scope of the present disclosure. Therefore, the protection scope of the present disclosure should be subject to the protection scope of the claims.

Claims

1. A combined dissimilar material turbine stator blade assembly, characterized in that Comprising: A ceramic matrix blade, including a ceramic matrix blade body; the ceramic matrix blade body is a hollow air-cooled structure of ceramic matrix, arranged at one end of the turbine stator blade that bears the scouring of high-temperature gas, and the windward surface faces the incoming gas direction of the high-temperature gas; A metal blade, including a metal blade body of metal matrix; the metal blade body is a hollow double-wall air-cooled structure, and the two layers of blade body walls form a blade body structure with a hollow and gradually changing thickness; the front end with a large thickness of the metal blade body is spliced with the leeward surface of the ceramic matrix blade body, and the converging tail end with a small thickness of the metal blade body is the high-velocity flow-resistant diversion area of the turbine stator blade; Metal upper flange plates; a plurality of the metal upper flange plates are sequentially spliced to form an outer ring structure; Metal lower flange plates; a plurality of the metal lower flange plates are sequentially spliced to form an inner ring structure; Wherein: the ceramic matrix blade and the metal blade are combined into one turbine stator blade; the outer ring structure and the inner ring structure are coaxially arranged; the cavity between the outer ring structure and the inner ring structure forms a part of the diversion cavity; a plurality of the turbine stator blades are symmetrically and uniformly arranged in the diversion cavity with the central axis of the diversion cavity as the center; The combined dissimilar material turbine stator blade assembly further includes upper flange plates and upper hanging tops; the upper flange plates are spliced at one end of the metal upper flange plates in the incoming gas direction of the high-temperature gas; a plurality of the upper flange plates are sequentially spliced to form a front end ring; the front end ring is coaxially arranged with the diversion cavity; the inner ring surface of the front end ring and the metal lower flange plate form another part of the diversion cavity; On the top surface of the ceramic matrix blade that does not contact the high-temperature gas and the metal blade body, there is an upper mounting edge that is fixedly connected to the ceramic matrix blade body; the ceramic matrix blade is spliced between two adjacent upper flange plates based on the upper mounting edge; The upper hanging top is used to fix two adjacent upper flange plates at the outer ring surface of the front end ring.

2. The combined dissimilar material turbine stator blade assembly according to claim 1, wherein On the bottom surface of the ceramic matrix blade that does not contact the high-temperature gas and the metal blade body, there is a lower positioning boss that is fixedly connected to the ceramic matrix blade body; The metal lower flange plate is provided with a groove at one end in the incoming gas direction of the high-temperature gas The ceramic matrix blade is spliced in the groove based on the positioning boss.

3. The combined dissimilar material turbine stator vane assembly according to claim 2, wherein A cold air cavity is arranged in the ceramic matrix blade body; the wall thickness of the ceramic matrix shell of the ceramic matrix blade body is uniform, woven and formed by ceramic matrix composite materials, and a plurality of ceramic matrix blade body film holes for air-conducting the cold air cavity to the diversion cavity are arranged on the main body, and an air intake channel is arranged at the upper mounting edge.

4. The combined dissimilar material turbine stator blade assembly according to claim 3, wherein The wall thickness of the ceramic matrix shell is 2-3 mm; the aperture of the ceramic matrix blade body film hole is less than 0.6 mm.

5. The combined dissimilar material turbine stator vane assembly according to claim 4, characterized in that, The metal blade is integrally cast from a high-temperature single crystal alloy; a gas collecting cavity is arranged in the metal blade body, and a plurality of impact cavities are arranged at the interlayer between the double walls; The impact cavity is air-conducted with the gas collecting cavity based on blade body impact holes; A plurality of metal blade body film holes are arranged at the blade basin and blade back of the metal blade body that contact the high-temperature gas; the metal blade body film holes air-conduct the impact cavity to the diversion cavity; A tail split seam is provided at the convergent end of the metal blade body; the tail split seam allows gas to conduct between the impact chamber and the diversion chamber.

6. The combined dissimilar material turbine stator vane assembly according to claim 5, wherein The aperture of the blade body impact holes is 1 to 1.5 mm; the aperture of the gas film holes of the metal blade body is less than 0.5 mm.

7. The combined dissimilar material turbine stator vane assembly according to claim 6, wherein A plurality of front impact holes are provided at the splicing portion of the metal blade body and the ceramic matrix blade; the front impact holes are in gas communication with the gas collecting chamber.

8. The combined dissimilar material turbine stator vane assembly according to claim 7, wherein A plurality of front gas film holes are provided at the splicing portion of the metal blade body and the ceramic matrix blade; one end of the front gas film holes is in gas communication with the splicing portion between the metal blade body and the ceramic matrix blade, and the other end is in communication with the diversion chamber.

9. The combined dissimilar material turbine stator blade assembly according to claim 8, characterized in that The metal blade is spliced between the metal upper edge plate and the metal lower edge plate based on the top surface and the bottom surface that do not contact the high-temperature gas and the metal blade body.

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