Unequal-thickness ultrathin ceramic matrix composite airfoil

Abstract

The invention discloses an unequal-thickness ultrathin ceramic matrix composite airfoil, belongs to the technical field of aircraft airfoils, and solves the problem that an existing aircraft airfoil cannot be ultrathin due to the fact that a heat insulation structure and the like are designed in the aircraft airfoil. The airfoil specifically comprises an airfoil framework, wherein one side of the airfoil framework is connected with an airfoil front edge; the airfoil framework is wrapped with a skin assembly; the airfoil framework comprises a two-dimensional laying box-shaped piece splicing framework and a three-dimensional needling solid framework; the two-dimensional laying box-shaped piece splicing framework and the three-dimensional needling solid framework are wrapped with the skin assembly. According to the aircraft, the skin assembly bears pneumatic pressure borne by the aircraft in the flight process and transmits the pneumatic pressure to the airfoil framework, and the airfoil framework plays a role in supporting the skin assembly; the airfoil front edge is arranged at the front end of the airfoil framework, has an anti-ablation function and can bear high-temperature ablation generated by friction with air in the flying process; a heat insulation structure does not need to be designed in the whole airfoil, unequal-thickness ultra-thin design can be achieved, and the requirements for high temperature resistance and high bearing performance are met at the same time.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft wing surfaces, and in particular to an unequal-thickness ultra-thin ceramic matrix composite wing surface. Background Art

[0002] With the demand of aircraft for higher flight speeds, the air resistance of their wing surfaces continuously increases. In order to reduce the air resistance suffered by the wing surfaces during flight, the wing surface shapes are usually designed into unequal-thickness ultra-thin structures, and the thickness requirement for the ultra-thin region is less than 10 mm.

[0003] There are usually two types of existing aircraft wing surfaces. One is the traditional metal structure wing surface. Although the metal structure wing surface has high strength, its high-temperature resistance is poor. In a high-temperature environment, the metal structure wing surface has disadvantages such as easy deformation, low high-temperature mechanical properties, and poor ablation resistance, so it can no longer meet the requirements for the development of future aircraft. The other is the wing surface combined with metal materials and composite materials. This type of wing surface can use composite materials as a protective layer or heat insulation layer to protect the metal structure, and can also block the heat transfer generated by air friction from reaching the metal structure; however, for the ultra-thin structure wing surface, there is not enough space inside it to design protection or heat insulation, so the combination of metal materials and composite materials is also not applicable to the preparation of ultra-thin structure wing surfaces. Continuous fiber reinforced ceramic matrix composites can replace metal materials for the preparation of aircraft wing surfaces due to their characteristics such as high specific strength, high specific modulus, large damage tolerance, and low density; and it can be used as a high-temperature resistant and ablation-resistant functional material without the need to design an additional heat insulation structure. Therefore, the present invention provides an unequal-thickness ultra-thin wing surface structure of an aircraft prepared by using ceramic matrix composites. Summary of the Invention

[0004] Aiming at the deficiencies existing in the prior art, the present invention provides an unequal-thickness ultra-thin ceramic matrix composite wing surface, which solves the problem that the existing aircraft wing surface cannot be ultra-thin due to the internal design of heat insulation and other structures.

[0005] In order to achieve the above object, the technical solution adopted by the present invention is as follows:

[0006] An unequal-thickness ultra-thin ceramic matrix composite wing surface, comprising a wing surface skeleton, a wing surface leading edge is connected to one side of the wing surface skeleton; a skin assembly is wrapped on the wing surface skeleton; the wing surface skeleton includes a two-dimensional ply box-shaped part splicing skeleton, and a three-dimensional needle-punched solid skeleton is connected to the front end of the two-dimensional ply box-shaped part splicing skeleton; the wing surface leading edge is connected to the two-dimensional ply box-shaped part splicing skeleton; the skin assembly is wrapped on the two-dimensional ply box-shaped part splicing skeleton and the three-dimensional needle-punched solid skeleton.

[0007] In this solution, the skin component bears the aerodynamic pressure received during the flight of the aircraft and transmits it to the wing surface framework, and the wing surface framework plays a role in supporting the skin component; the leading edge of the wing surface is arranged at the front end of the wing surface framework and has an anti-ablation function, capable of withstanding the high-temperature ablation generated by the friction with the air during flight; there is no need to design a heat insulation structure inside the entire wing surface, and an unequal-thickness ultra-thin design can be achieved, while meeting the requirements of high temperature resistance and high load-bearing performance at the same time.

[0008] Furthermore, the two-dimensional laminated box-shaped part splicing framework includes a first main frame, a second main frame, and a third main frame; there are a first sub-frame and a second sub-frame connected to each other between the first main frame and the second main frame, the first main frame is connected to the first sub-frame, and the second main frame is connected to the second sub-frame; there are a third sub-frame and a fourth sub-frame connected to each other between the second main frame and the third main frame, the second main frame is connected to the third sub-frame, and the third main frame is connected to the fourth sub-frame;

[0009] On the side of the first main frame away from the second main frame, there is also a side frame connected;

[0010] The side frame, the first main frame, the first sub-frame, the second sub-frame, and the second main frame are all connected to the three-dimensional needle-punched solid framework; the second main frame, the fourth sub-frame, and the third main frame are all connected to the leading edge of the wing surface.

[0011] In this solution, the two-dimensional laminated box-shaped part splicing framework is mainly applied to the places where the thickness of the unequal-thickness ultra-thin ceramic matrix composite material wing surface is relatively thick. Among them, the first main frame, the second main frame, and the third main frame serve as the main load-bearing structures, and sub-frames are arranged between adjacent two main frames, with a dense distribution, which can support the skin component more evenly and comprehensively.

[0012] Furthermore, the rear ends of the side frame and the first main frame are connected by a first box-shaped part; the middle parts of the side frame and the first main frame are connected by a second box-shaped part;

[0013] The rear ends of the first main frame and the first sub-frame, the rear ends of the first sub-frame and the second sub-frame, and the rear ends of the second sub-frame and the second main frame are respectively connected by three third box-shaped parts; the middle parts of the first main frame and the first sub-frame, the middle parts of the first sub-frame and the second sub-frame, and the middle parts of the second sub-frame and the second main frame are respectively connected by three fourth box-shaped parts;

[0014] The rear ends of the second main frame and the third sub-frame, the rear ends of the third sub-frame and the fourth sub-frame, and the rear ends of the fourth sub-frame and the third main frame are respectively connected by three fifth box-shaped parts; the middle parts of the second main frame and the third sub-frame are connected by a sixth box-shaped part;

[0015] The third main frame and the leading edge of the wing surface are connected by a seventh box-shaped part.

[0016] In this solution, adjacent main frames or between the auxiliary frame and the main frame are connected by a box-shaped part with a box-shaped structure. When the box-shaped structure is connected, the contact area is large and the connection strength is high, which can improve the overall load-bearing capacity of the splicing skeleton of the two-dimensional laying box-shaped part; moreover, the box-shaped part can be directly prepared by flanging the two-dimensional plain carbon fiber cloth, and the preparation is simple; both sides of each main frame are box-shaped structures, and when connected to the three-dimensional needle-punched solid skeleton, the connection head part can be directly inserted into the inside of the box-shaped structure, and the assembly is simple and convenient.

[0017] Furthermore, the three-dimensional needle-punched solid skeleton includes a first solid skeleton and a second solid skeleton, and the first solid skeleton and the second solid skeleton are spliced through a stepped structure arranged on the connection surface;

[0018] The first solid skeleton is connected to the side frame and the first main frame; the second solid skeleton is connected to the first main frame, the first auxiliary frame and the second auxiliary frame.

[0019] In this solution, the three-dimensional needle-punched solid skeleton is mainly applied to the places where the thickness of the wing surface of the non-uniform thickness ultra-thin ceramic matrix composite material is ultra-thin. The solid skeleton in the thickness direction is used to support the skin assembly, which can improve the wing surface strength in the ultra-thin area; the first solid skeleton and the second solid skeleton are overlapped through a stepped structure, and the generation of thermal stress is reduced in sections, which can avoid large internal stress caused by thermal load and cause structural damage.

[0020] Furthermore, a number of weight-reducing holes are provided on the first solid skeleton and the second solid skeleton.

[0021] In this solution, the weight of the three-dimensional needle-punched solid skeleton is reduced to realize the lightweight of the wing surface of the non-uniform thickness ultra-thin ceramic matrix composite material.

[0022] Furthermore, the leading edge of the wing surface includes a first leading edge, a second leading edge and a third leading edge; a first tenon is provided on the side of the first leading edge, and the first tenon is embedded in a first mortise groove on one side of the second leading edge; a second tenon is provided on the other side of the second leading edge, and the second tenon is embedded in a second mortise groove on the side of the third leading edge;

[0023] The second leading edge is connected to the second main frame and the fourth auxiliary frame; the third leading edge is connected to the third main frame and the seventh box-shaped part.

[0024] In this solution, during the flight of the aircraft, the friction between the leading edge of the wing surface and the air is relatively serious, and the high-temperature ablation suffered by the leading edge of the wing surface is also relatively serious. The leading edge of the wing surface is designed in sections, which can avoid structural damage when it is deformed by heat; the first leading edge and the second leading edge are overlapped with each other through a tenon and mortise structure, and the second leading edge and the third leading edge are also overlapped with each other through a tenon and mortise structure to ensure the deformation coordination between each section of the leading edge and guarantee the continuity of the aerodynamic shape of the wing surface.

[0025] Further, the first leading edge and the second leading edge are also connected by pins on the side away from the wing surface framework.

[0026] In this solution, the front ends of the first leading edge and the second leading edge are relatively thin, so pins are used to connect the two together, and the assembly is simple.

[0027] Further, the skin assembly includes an upper skin, a lower skin, side skins and a rear skin; the upper skin covers the upper surfaces of the two-dimensional laminated box-shaped part splicing framework and the three-dimensional needled solid framework; the lower skin covers the lower surfaces of the two-dimensional laminated box-shaped part splicing framework and the three-dimensional needled solid framework; the rear skin covers the rear surfaces of the two-dimensional laminated box-shaped part splicing framework and the three-dimensional needled solid framework, and the rear skin is respectively connected to the upper skin and the lower skin; the side skins cover the side surfaces of the two-dimensional laminated box-shaped part splicing framework.

[0028] Further, the thickness of the two-dimensional laminated box-shaped part splicing framework and the three-dimensional needled solid framework gradually increases from the side close to the leading edge of the wing.

[0029] In this solution, this design can reduce the frictional resistance with the air.

[0030] In a second aspect, based on the unequal-thickness ultra-thin ceramic matrix composite wing surface provided in the first aspect, the present invention provides a method for preparing the leading edge of the wing surface, including the following steps:

[0031] S1: Prepare a leading edge preform by a three-dimensional needling process;

[0032] S2: Prepare an interface layer on the surface of the leading edge preform by a chemical vapor infiltration process;

[0033] S3: Prepare a ceramic matrix inside the leading edge preform by a chemical vapor infiltration process;

[0034] S4: Perform ultra-high temperature modification on the area of 20-30 mm at the front end of the leading edge preform to obtain the leading edge of the wing surface.

[0035] In this solution, during the flight of the aircraft, the leading edge of the wing surface is subjected to high-temperature ablation due to friction with the air, and its front end is the most serious. Therefore, gradient ultra-high temperature modification is performed on the local area of the leading edge of the wing surface to improve its high-temperature resistance performance.

[0036] The beneficial effects of the present invention are:

[0037] The unequal-thickness ultra-thin ceramic matrix composite wing surface provided by the present invention is internally provided with a wing surface framework, and the wing surface framework is wrapped with a skin assembly, which can not only bear the aerodynamic load received by the skin assembly during flight, but also bear the high-temperature ablation generated by the friction between the wing surface and the air during high-speed flight.

[0038] The wing surface framework consists of a spliced framework of two-dimensional laminated box-shaped parts and a three-dimensional needled solid framework. The spliced framework of two-dimensional laminated box-shaped parts is formed by splicing multiple main frames and sub-frames, which can not only well withstand the aerodynamic loads on the wing surface but also well support the skin assembly. The three-dimensional needled solid framework can be prepared by a three-dimensional needled preform. The prepared solid structure has high strength and can strengthen the strength of the ultra-thin parts with unequal thickness of the ultra-thin ceramic matrix composite wing surface. Moreover, the middle of the three-dimensional needled solid framework is hollowed out to reduce the structural weight and achieve overall lightweighting. Description of the Drawings

[0039] Figure 1 Schematic diagram of a wing surface structure of an unequal-thickness ultra-thin ceramic matrix composite material according to the present invention;

[0040] Figure 2 Exploded view of a wing surface of an unequal-thickness ultra-thin ceramic matrix composite material according to the present invention;

[0041] Figure 3 Schematic diagram of the spliced framework of two-dimensional laminated box-shaped parts and the three-dimensional needled solid framework according to the present invention;

[0042] Figure 4 Exploded view of the spliced framework of two-dimensional laminated box-shaped parts and the three-dimensional needled solid framework according to the present invention;

[0043] Figure 5 Exploded view of the leading edge of the wing surface according to the present invention.

[0044] Reference Signs:

[0045] 1, wing surface framework; 11, spliced framework of two-dimensional laminated box-shaped parts; 111, first main frame; 112, second main frame; 113, third main frame; 114, first sub-frame; 115, second sub-frame; 116, third sub-frame; 117, fourth sub-frame; 118, side frame; 12, three-dimensional needled solid framework; 121, first solid framework; 122, second solid framework; 131, first box-shaped part; 132, second box-shaped part; 133, third box-shaped part; 134, fourth box-shaped part; 135, fifth box-shaped part; 136, sixth box-shaped part; 137, seventh box-shaped part;

[0046] 2, leading edge of the wing surface; 21, first leading edge; 211, first tenon; 221, first mortise; 222, second tenon; 231, second mortise; 22, second leading edge; 23, third leading edge; 24, pin;

[0047] 3, skin assembly; 31, upper skin; 32, lower skin; 33, side skin; 34, rear skin; Detailed Embodiment

[0048] The present invention will be further described below in conjunction with the accompanying drawings and specific embodiments. The following describes the specific embodiments of the present invention to facilitate those skilled in the art of this technology to understand the present invention. However, it should be clear that the present invention is not limited to the scope of the specific embodiments. For those of ordinary skill in the art of this technology, as long as various changes are within the spirit and scope of the present invention defined and determined by the appended claims, these changes are obvious, and all inventions and creations using the concept of the present invention are within the scope of protection.

[0049] Embodiment 1

[0050] As Figure 1 and Figure 2 shown, this embodiment provides a wing surface made of an unequal-thickness ultra-thin ceramic matrix composite material. Inside the wing surface made of the unequal-thickness ultra-thin ceramic matrix composite material, there is a wing surface skeleton 1, and a skin assembly 3 is wrapped around the wing surface skeleton 1. It can not only bear the aerodynamic load received by the skin assembly 3 during flight but also bear the high-temperature ablation generated by the friction between the wing surface and the air during high-speed flight. Specifically, it includes:

[0051] The wing surface skeleton 1, the leading edge 2 of the wing surface, and the skin assembly 3;

[0052] Among them, a skin assembly 3 is wrapped around the wing surface skeleton 1; the skin assembly 3 bears the aerodynamic pressure received during the flight of the aircraft and transmits it to the wing surface skeleton 1, and the wing surface skeleton 1 plays a role in supporting the skin assembly 3. One side of the wing surface skeleton 1 is connected to the leading edge 2 of the wing surface; the leading edge 2 of the wing surface has an anti-ablation function and can bear the high-temperature ablation generated by the friction with the air during flight.

[0053] The wing surface skeleton 1 includes a two-dimensional laminated box-shaped part splicing skeleton 11 and a three-dimensional needle-punched solid skeleton 12. The front end of the two-dimensional laminated box-shaped part splicing skeleton 11 is connected to the three-dimensional needle-punched solid skeleton 12; the leading edge 2 of the wing surface is connected to the two-dimensional laminated box-shaped part splicing skeleton 11; the skin assembly 3 is wrapped around the two-dimensional laminated box-shaped part splicing skeleton 11 and the three-dimensional needle-punched solid skeleton 12.

[0054] The two-dimensional laminated box-shaped part splicing skeleton 11;

[0055] The splicing skeleton 11 of the two-dimensional laminated box-shaped part includes a first main frame 111, a second main frame 112, and a third main frame 113; between the first main frame 111 and the second main frame 112, there are a first sub-frame 114 and a second sub-frame 115 connected to each other, the first main frame 111 is connected to the first sub-frame 114, and the second main frame 112 is connected to the second sub-frame 115; between the second main frame 112 and the third main frame 113, there are a third sub-frame 116 and a fourth sub-frame 117 connected to each other, the second main frame 112 is connected to the third sub-frame 116, and the third main frame 113 is connected to the fourth sub-frame 117; on the side of the first main frame 111 away from the second main frame 112, there is also a side frame 118 connected; the side frame 118, the first main frame 111, the first sub-frame 114, the second sub-frame 115, and the second main frame 112 are all connected to the three-dimensional needle-punched solid skeleton 12; the second main frame 112, the fourth sub-frame 117, and the third main frame 113 are all connected to the leading edge 2 of the wing surface. The splicing skeleton 11 of the two-dimensional laminated box-shaped part is mainly applied to the places where the thickness of the wing surface of the non-uniform thickness ultra-thin ceramic matrix composite material is relatively thick. Among them, the first main frame 111, the second main frame 112, and the third main frame 113 serve as the main load-bearing structures, and sub-frames are arranged between adjacent two main frames, with a dense distribution, which can support the skin component 3 more evenly and comprehensively.

[0056] The rear ends of the side frame 118 and the first main frame 111 are connected by a first box-shaped part 131; the middle parts of the side frame 118 and the first main frame 111 are connected by a second box-shaped part 132; the rear ends of the first main frame 111 and the first sub-frame 114, the rear ends of the first sub-frame 114 and the second sub-frame 115, and the rear ends of the second sub-frame 115 and the second main frame 112 are respectively connected by three third box-shaped parts 133; the middle parts of the first main frame 111 and the first sub-frame 114, the middle parts of the first sub-frame 114 and the second sub-frame 115, and the middle parts of the second sub-frame 115 and the second main frame 112 are respectively connected by three fourth box-shaped parts 134; the rear ends of the second main frame 112 and the third sub-frame 116, the rear ends of the third sub-frame 116 and the fourth sub-frame 117, and the rear ends of the fourth sub-frame 117 and the third main frame 113 are respectively connected by three fifth box-shaped parts 135; the middle part of the second main frame 112 and the third sub-frame 116 is connected by a sixth box-shaped part 136; between the third main frame 113 and the leading edge 2 of the wing surface, there is a seventh box-shaped part 137 connected. Adjacent two main frames or between a sub-frame and a main frame are connected by box-shaped parts of the box-shaped structure. When the box-shaped structure is connected, the contact area is large and the connection strength is high, which can improve the overall load-bearing capacity of the splicing skeleton 11 of the two-dimensional laminated box-shaped part; and the box-shaped parts can be directly prepared by turning the edge of the two-dimensional plain carbon fiber cloth, and the preparation is simple; on both sides of each main frame is a box-shaped structure, and when connected to the three-dimensional needle-punched solid skeleton 12, the connection head part can be directly inserted into the inside of the box-shaped structure, and the assembly is simple and convenient.

[0057] Three-dimensional needle-punched solid skeleton 12;

[0058] The three-dimensional needle-punched solid framework 12 includes a first solid framework 121 and a second solid framework 122. The first solid framework 121 and the second solid framework 122 are spliced through a stepped structure arranged on the connection surface. The first solid framework 121 is connected to the side frame 118 and the first main frame 111. The second solid framework 122 is connected to the first main frame 111, the first sub-frame 114, and the second sub-frame 115. The three-dimensional needle-punched solid framework 12 is mainly applied to the places where the thickness of the wing surface of the non-uniform-thickness ultra-thin ceramic matrix composite material is ultra-thin. A solid framework in the thickness direction is used to support the skin assembly 3, which can improve the wing surface strength in the ultra-thin area. The first solid framework 121 and the second solid framework 122 are lapped together through a stepped structure, which can reduce the generation of thermal stress in segments and avoid large internal stress caused by thermal load, resulting in structural damage.

[0059] A number of weight-reducing holes are provided on the first solid framework 121 and the second solid framework 122, which can reduce the weight of the three-dimensional needle-punched solid framework 12 and realize the lightweight of the wing surface of the non-uniform-thickness ultra-thin ceramic matrix composite material.

[0060] The leading edge 2 of the wing surface;

[0061] The leading edge 2 of the wing surface includes a first leading edge 21, a second leading edge 22, and a third leading edge 23. A first tenon 211 is arranged on the side of the first leading edge 21, and the first tenon 211 is embedded in the first mortise 221 on one side of the second leading edge 22. A second tenon 222 is arranged on the other side of the second leading edge 22, and the second tenon 222 is embedded in the second mortise 231 on the side of the third leading edge 23. The second leading edge 22 is connected to the second main frame 112 and the fourth sub-frame 117. The third leading edge 23 is connected to the third main frame 113 and the seventh box-shaped part 137. During the flight of the aircraft, the friction between the leading edge 2 of the wing surface and the air is relatively serious, and the high-temperature ablation suffered by the leading edge 2 of the wing surface is also relatively serious. The segmented design of the leading edge 2 of the wing surface can avoid structural damage when it is deformed by heat. The first leading edge 21 and the second leading edge 22 are mutually lapped through a tenon-mortise structure, and the second leading edge 22 and the third leading edge 23 are also mutually lapped through a tenon-mortise structure, ensuring the deformation coordination between each segment of the leading edge and guaranteeing the continuity of the aerodynamic shape of the wing surface.

[0062] The first leading edge 21 and the second leading edge 22 are also connected by a pin 24 on the side away from the wing surface framework 1. The front ends of the first leading edge 21 and the second leading edge 22 are relatively thin, so the pin 24 is used to connect the two together, and the assembly is simple.

[0063] The skin assembly 3;

[0064] The skin assembly 3 includes an upper skin 31, a lower skin 32, side skins 33, and a rear skin 34; the upper skin 31 covers the upper surfaces of the spliced framework 11 of two-dimensional laminated box-shaped parts and the three-dimensional needled solid framework 12; the lower skin 32 covers the lower surfaces of the spliced framework 11 of two-dimensional laminated box-shaped parts and the three-dimensional needled solid framework 12; the rear skin 34 covers the rear surfaces of the spliced framework 11 of two-dimensional laminated box-shaped parts and the three-dimensional needled solid framework 12, and the rear skin 34 is respectively connected to the upper skin 31 and the lower skin 32; the side skins 33 cover the side surfaces of the spliced framework 11 of two-dimensional laminated box-shaped parts.

[0065] The thickness of the spliced framework 11 of two-dimensional laminated box-shaped parts and the three-dimensional needled solid framework 12 gradually increases starting from the side close to the leading edge 2 of the wing. This design can reduce the frictional resistance with the air.

[0066] In this embodiment, the skin assembly 3 and the spliced framework 11 of two-dimensional laminated box-shaped parts are obtained by using laminated two-dimensional plain carbon fiber cloth as a preform and preparing an interface layer and a ceramic matrix through a chemical vapor infiltration process. The three-dimensional needled solid framework 12 is obtained by using a three-dimensional needled preform and preparing an interface layer and a ceramic matrix through a chemical vapor infiltration process. After all parts are prepared, they are assembled together; after assembly, the surface of the skin assembly 3 is polished to make the surface of the skin assembly 3 smooth and achieve the required aerodynamic shape of the wing.

[0067] Embodiment 2

[0068] Based on the non-uniform thickness ultra-thin ceramic matrix composite wing surface provided in Embodiment 1, this embodiment provides a method for preparing the leading edge 2 of the wing surface, including the following steps:

[0069] S1: Prepare a leading edge preform by using a three-dimensional needling process;

[0070] S2: Prepare an interface layer on the surface of the leading edge preform by using a chemical vapor infiltration process;

[0071] S3: Prepare a ceramic matrix inside the leading edge preform by using a chemical vapor infiltration process;

[0072] S4: Perform ultra-high temperature modification on the area 20 - 30 mm at the front end of the leading edge preform to obtain the leading edge 2 of the wing surface.

[0073] In this embodiment, during the flight of the aircraft, the leading edge 2 of the wing surface is subjected to high-temperature ablation due to friction with the air, and its front end is the most serious. Therefore, local gradient ultra-high temperature modification is performed on the leading edge 2 of the wing surface to improve its high-temperature resistance performance.

[0074] Those of ordinary skill in the art will realize that the embodiments herein are provided to assist the reader in understanding the principles of the present invention, and it should be understood that the scope of protection of the present invention is not limited to such specific statements and embodiments. Those of ordinary skill in the art can make various other specific deformations and combinations that do not depart from the essence of the present invention based on these technical revelations disclosed in the present invention, and these deformations and combinations are still within the scope of protection of the invention.

Claims

1. An ultra-thin ceramic matrix composite wing surface of unequal thickness, characterized in that: It comprises an airfoil frame (1), one side of which is connected to an airfoil leading edge (2); and a skin component (3) is wrapped on the airfoil frame (1); The wing frame (1) comprises a two-dimensional ply box-shaped spliced ​​frame (11), the front end of which is connected to a three-dimensional needle-punched solid frame (12); the wing leading edge (2) is connected to the two-dimensional ply box-shaped spliced ​​frame (11); and the skin assembly (3) is wrapped around the two-dimensional ply box-shaped spliced ​​frame (11) and the three-dimensional needle-punched solid frame (12).

2. The unequal thickness ultra-thin ceramic matrix composite wing surface according to claim 1, characterized in that: The two-dimensional ply box-shaped component splicing skeleton (11) comprises a first main frame (111), a second main frame (112) and a third main frame (113); a first sub-frame (114) and a second sub-frame (115) connected to each other are arranged between the first main frame (111) and the second main frame (112); the first main frame (111) and the first sub-frame (114) are connected, and the second main frame (112) and the second sub-frame (115) are connected; a third sub-frame (116) and a fourth sub-frame (117) connected to each other are arranged between the second main frame (112) and the third main frame (113); the second main frame (112) and the third sub-frame (116) are connected, and the third main frame (113) and the fourth sub-frame (117) are connected; A side frame (118) is further connected to the first main frame (111) at a side away from the second main frame (112); The side frame (118), the first main frame (111), the first sub-frame (114), the second sub-frame (115) and the second main frame (112) are all connected to the three-dimensional needle-punched solid skeleton (12); the second main frame (112), the fourth sub-frame (117) and the third main frame (113) are all connected to the leading edge (2) of the airfoil.

3. The unequal thickness ultra-thin ceramic matrix composite wing surface according to claim 2, characterized in that: The side frame (118) and the rear end of the first main frame (111) are connected via a first box-shaped member (131); the side frame (118) and the middle part of the first main frame (111) are connected via a second box-shaped member (132); The rear ends of the first main frame (111) and the first sub-frame (114), the rear ends of the first sub-frame (114) and the second sub-frame (115), and the rear ends of the second sub-frame (115) and the second main frame (112) are respectively connected via three third box-shaped members (133); the middle parts of the first main frame (111) and the first sub-frame (114), the middle parts of the first sub-frame (114) and the second sub-frame (115), and the middle parts of the second sub-frame (115) and the second main frame (112) are respectively connected via three fourth box-shaped members (134); The rear ends of the second main frame (112) and the third sub-frame (116), the rear ends of the third sub-frame (116) and the fourth sub-frame (117), and the rear ends of the fourth sub-frame (117) and the third main frame (113) are respectively connected via three fifth box-shaped members (135); the middle parts of the second main frame (112) and the third sub-frame (116) are connected via a sixth box-shaped member (136); The third main frame (113) and the airfoil leading edge (2) are connected via a seventh box-shaped member (137).

4. The unequal thickness ultra-thin ceramic matrix composite wing surface according to claim 3, characterized in that: The three-dimensional needle-punched solid skeleton (12) comprises a first solid skeleton (121) and a second solid skeleton (122), wherein the first solid skeleton (121) and the second solid skeleton (122) are spliced ​​via a step-type structure arranged on a connecting surface; The first solid frame (121) is connected to the side frame (118) and the first main frame (111); the second solid frame (122) is connected to the first main frame (111), the first sub-frame (114) and the second sub-frame (115).

5. The unequal thickness ultra-thin ceramic matrix composite wing surface according to claim 4, characterized in that: A plurality of weight-reducing holes are provided on the first solid frame (121) and the second solid frame (122).

6. The unequal thickness ultra-thin ceramic matrix composite wing surface according to claim 4, characterized in that: The front edge (2) of the airfoil comprises a first front edge (21), a second front edge (22) and a third front edge (23); a first tenon (211) is arranged on the side of the first front edge (21), and the first tenon (211) is embedded in a first mortise (221) on one side of the second front edge (22); a second tenon (222) is arranged on the other side of the second front edge (22), and the second tenon (222) is embedded in a second mortise (231) on the side of the third front edge (23); The second front edge (22) is connected to the second main frame (112) and the fourth sub-frame (117); and the third front edge (23) is connected to the third main frame (113) and the seventh box-shaped member (137).

7. The unequal thickness ultra-thin ceramic matrix composite wing surface according to claim 6, characterized in that: The first leading edge (21) and the second leading edge (22) are also connected via a pin (24) at a side away from the airfoil frame (1).

8. The unequal thickness ultra-thin ceramic matrix composite wing surface according to claim 2, characterized in that: The skin assembly (3) comprises an upper skin (31), a lower skin (32), a side skin (33) and a rear skin (34); the upper skin (31) is coated on the upper surface of the two-dimensional ply box-shaped component splicing skeleton (11) and the three-dimensional needle-punched solid skeleton (12); the lower skin (32) is coated on the lower surface of the two-dimensional ply box-shaped component splicing skeleton (11) and the three-dimensional needle-punched solid skeleton (12); the rear skin (34) is coated on the rear surface of the two-dimensional ply box-shaped component splicing skeleton (11) and the three-dimensional needle-punched solid skeleton (12), and the rear skin (34) is respectively connected to the upper skin (31) and the lower skin (32); the side skin (33) is coated on the side surface of the two-dimensional ply box-shaped component splicing skeleton (11).

9. The unequal thickness ultra-thin ceramic matrix composite wing surface according to any one of claims 1 to 8, characterized in that: The thickness of the two-dimensional ply box-shaped splicing skeleton (11) and the three-dimensional needle-punched solid skeleton (12) gradually increases from the side close to the leading edge (2) of the airfoil.

10. A method for preparing the leading edge of an airfoil according to any one of claims 1 to 8, characterized in that: The following steps are involved: S1: Preparation of leading edge preform by three-dimensional needle punching process; S2: preparing an interface layer on the surface of the leading edge preform by chemical vapor infiltration process; S3: preparing a ceramic matrix inside the leading edge preform by chemical vapor infiltration process; S4: The area 20 to 30 mm from the front end of the leading edge preform is subjected to ultra-high temperature modification to obtain the airfoil leading edge (2).

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