Wing inclined supporting rod made of composite material and forming method of wing inclined supporting rod

Abstract

The invention discloses a composite material wing inclined supporting rod and a forming method thereof, the inclined supporting rod comprises the following structures: a structure main body, the cross section of the structure main body is of an I-shaped structure, and the two opposite sides of the structure main body are respectively provided with clamping grooves which extend along the longitudinal direction and have opposite openings; the transverse section of the front edge plate is of a C-shaped structure, the two edges of the front edge plate are provided with sunken first clamping plates in the longitudinal direction, and the first clamping plates are clamped in the clamping grooves and fixed; the transverse section of the rear edge plate is of a C-shaped structure, the two edges of the rear edge plate are provided with sunken second clamping plates in the longitudinal direction, and the second clamping plates are clamped in the other clamping grooves and fixed; connecting holes are formed in the metal joints, and the metal joints are connected to the two ends of the structure body and used for being connected with the wings and the fuselage. The inclined supporting rod has the advantages of being light in weight, high in strength, resistant to fatigue and resistant to corrosion, the arrangement of the front edge plate and the rear edge plate can facilitate reduction of air resistance when an aircraft flies, and the fuel economy of the aircraft can be improved.

Description

Technical Field

[0001] This invention relates to the field of composite material molding technology, specifically to a composite material wing strut and its molding method. Background Technology

[0002] Wing struts are critical load-bearing components in aircraft wing structures, primarily used to transfer and distribute aerodynamic loads to ensure wing stiffness and stability. Currently, traditional wing struts are mostly made of metallic materials (such as aluminum or titanium alloys) through machining or forging. However, this method suffers from problems such as heavy strut weight, poor corrosion resistance, and high manufacturing costs. Furthermore, heavy struts are detrimental to aircraft performance. With the continuous development of composite material technology, there is an urgent need for a lightweight, high-strength wing strut structure to improve aircraft performance and reliability. Summary of the Invention

[0003] The purpose of this invention is to design a composite material wing strut, which aims to solve the problems of large weight, poor corrosion resistance and high manufacturing cost of traditional metal wing struts, thereby improving aircraft performance and reliability.

[0004] To achieve the above objectives, the present invention is implemented through the following technical solution: This invention provides a composite material wing diagonal brace, which includes the following structural configuration: The main structure has an I-shaped cross-section. The main structure has slots on opposite sides that extend longitudinally and have opposite openings. The main structure is a laminate structure formed from composite materials. The front edge plate has a "C" shaped cross section. The two edges of the front edge plate are provided with recessed first clamping plates along the longitudinal direction. The first clamping plates are engaged and fixed in the clamping groove. The front edge plate is a sandwich structure formed by molding composite material. The rear edge plate has a "C" shaped cross-section. Two edges of the rear edge plate are longitudinally provided with recessed second clamping plates. The second clamping plates are engaged in another slot and fixed. The rear edge plate is a sandwich structure formed by molding composite material. And a metal connector with connection holes, the metal connector being connected to the opposite ends of the main body of the structure, for connection to the wing and fuselage respectively.

[0005] Furthermore, a composite material wing diagonal brace: the main body of the structure is formed by connecting an upper edge strip, a web plate, and a lower edge strip to form an "I"-shaped cross section, and the groove is formed between the upper edge strip, the web plate, and the lower edge strip.

[0006] Further, a composite material wing brace: the upper flange and the lower flange have the same size and thickness.

[0007] Further, a composite material wing brace: the depth of sinking of the first clamping plate and the second clamping plate is the same as the thickness of the upper flange or the lower flange, so that after the first clamping plate and the second clamping plate are respectively clamped in the clamping grooves, the surfaces of the upper flange and the lower flange can be flush with the edges of the front edge plate and the rear edge plate respectively.

[0008] Further, a composite material wing brace: after the first clamping plate and the second clamping plate are clamped in the clamping grooves, they are respectively fixed by bonding with structural adhesive.

[0009] Further, a composite material wing brace: the metal joint and the end of the structural body are connected by adhesive screw.

[0010] The present invention also provides a forming method for a composite material wing brace, and this method includes the following steps: S1. Forming of the structural body: Provide a pair of forming molds with three continuous laying surfaces, and an R angle is formed between adjacent laying surfaces. After applying a release agent on the laying surfaces of the two forming molds, first lay a layer of the first prepreg, and then lay a plurality of second prepregs to respectively form a prepreg blank with a "C" - shaped structure on the two forming molds; Align the two forming molds so that the two prepreg blanks are butted, and fill a composite material wick at the butted R angle. Then lay the first prepreg at the non - butted part of the prepreg blank to form a covering layer connecting the two prepreg blanks, and then cure and mold, and remove the forming molds to obtain an "I" - shaped structural body; S2. Forming of the front edge plate: Provide a laying tooling whose cross - sectional shape is adapted to the transverse cross - section of the front edge plate. The surface of the laying tooling is used as the front edge plate laying surface. After applying a release agent, first lay the first prepreg, then fill a foam core material, and then lay the first prepreg again to form a sandwich - structured front edge plate blank. Then make a vacuum bag, and after vacuum compaction and curing, obtain the front edge plate; S3. Forming of the rear edge plate: Provide a laying tooling whose cross - sectional shape is adapted to the transverse cross - section of the rear edge plate. The surface of the laying tooling is used as the rear edge plate laying surface. After applying a release agent, first lay the first prepreg, then fill a foam core material, and then lay the first prepreg again to form a sandwich - structured rear edge plate blank. Then make a vacuum bag, and after vacuum compaction and curing, obtain the rear edge plate; S4. Processing of the metal joint: Process a metal joint with a connecting hole by using a light metal material; S5. Assembly: First, place the metal connectors at both ends of the main structure and position them to ensure coaxiality, and then connect and fix them to the main structure. Insert the first card plate on the front edge plate into the card slot and connect and fix it, and then insert the second card plate on the rear edge plate into the card slot and connect and fix it to obtain the composite material wing diagonal brace.

[0011] Furthermore, a method for forming a composite material wing strut: the first prepreg is a T300 grade carbon fiber fabric, and the second prepreg is a T700 grade carbon fiber unidirectional material.

[0012] Furthermore, a method for forming a composite material wing strut: when the second prepreg is laid, the layup ratios in the 0° / ±45° / 90° directions are 60% / 30% / 10%, respectively.

[0013] The beneficial effects of this invention are: The composite material wing strut of this invention overcomes the problems of traditional metal struts, such as large weight, poor corrosion resistance, and high manufacturing cost. At the same time, thanks to the optimized design of the strut structure and materials of this invention, it has the advantages of light weight, high strength, fatigue resistance, and corrosion resistance. Furthermore, the structural arrangement of the leading edge plate and trailing edge plate can significantly reduce air resistance during aircraft flight and improve the aircraft's fuel economy. Attached Figure Description

[0014] To more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings used in the following description of the embodiments will be briefly introduced. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0015] Figure 1 This is a schematic diagram of the composite material wing diagonal brace designed in Embodiment 1 of the present invention; Figure 2 The end cross-sectional view of the composite material wing diagonal brace designed for Embodiment 1 of the present invention; Figure 3 This is a schematic diagram of the main structure of the composite material wing diagonal brace designed in Embodiment 1 of the present invention; Figure 4 This is a schematic diagram of the leading edge plate and trailing edge plate in the composite material wing diagonal brace designed in Embodiment 1 of the present invention; Figure 5 This is a schematic diagram showing the two molding molds being joined together after being laid out in Embodiment 2 of the present invention.

[0016] The markings in the image are as follows: 1-Structural body, 2-Front edge plate, 3-Rear edge plate, 4-Metal joint, 5-Forming mold, 11-Upper edge strip, 12-Body plate, 13-Lower edge strip, 14-Slot, 15-Prepreg blank, 16-Composite material twist strip, 17-Covering layer, 21-First clamping plate, 31-Second clamping plate, 41-Connecting hole, 51-Paving surface. Detailed Implementation

[0017] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. The following description of at least one exemplary embodiment is merely illustrative and is in no way intended to limit the present invention or its application or use. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention.

[0018] In the description of this invention, it should be understood that the terms "upper," "lower," "left," "right," "top," and "bottom," etc., indicating orientation or positional relationships, are merely for the convenience of describing the invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation of the invention. Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of indicated technical features. Thus, a feature defined with "first" and "second" may explicitly or implicitly include one or more of that feature. Moreover, the terms "first," "second," etc., are used to distinguish similar objects and are not necessarily used to describe a specific order or sequence. It should be understood that such data can be interchanged where appropriate so that the embodiments of the invention described herein can be implemented in orders other than those illustrated or described herein.

[0019] Example 1

[0020] like Figures 1-4 As shown, this embodiment 1 provides a composite material wing diagonal brace, which includes the following structural configuration: The main body 1 is a transverse cross-section "I" shaped structure formed by connecting the upper edge strip 11, the web plate 12 and the lower edge strip 13. The upper edge strip 11 and the lower edge strip 13 have the same size and thickness. A groove 14 is formed between the upper edge strip 11, the web plate 12 and the lower edge strip 13. That is, grooves 14 extending longitudinally and with opposite openings are formed on opposite sides of the main body 1. The main body 1 is a laminate structure obtained by molding composite material. The front edge plate 2 has a "C" shaped cross section. The two edges of the front edge plate 2 are respectively provided with a recessed first clamping plate 21 along the longitudinal direction (the recessed depth of the first clamping plate 21 is the same as the thickness of the upper edge strip 11 and the lower edge strip 13). The first clamping plate 21 is snapped into the clamping groove 14 and is bonded and fixed by structural adhesive. The front edge plate 2 is a sandwich structure obtained by molding composite material. The rear edge plate 3 has a "C" shaped cross section. The two edges of the rear edge plate 3 are respectively provided with recessed second clamping plates 31 along the longitudinal direction (the recessed depth of the second clamping plates 31 is the same as the thickness of the upper edge strip 11 and the lower edge strip 13). The second clamping plates 31 are snapped into another clamping groove 14 and fixed by adhesive. The rear edge plate 3 is a sandwich structure obtained by molding composite material. And metal connector 4, which is made of aluminum alloy and has connection holes 41. The metal connector 4 is connected to both ends of the main body 1 and fixed by rubber screws. The metal connectors 4 at both ends are used to connect to the wing and the fuselage respectively.

[0021] Example 2

[0022] like Figure 5 As shown in Embodiment 2, this method provides a method for forming a composite material wing diagonal brace, which includes the following steps: S1. Forming of the main structural element 1: A pair of molding dies 5 with three continuous lay-up surfaces 51 are provided, and an R-angle is formed between adjacent lay-up surfaces 51. After applying a release agent to the lay-up surfaces 51 of the two molding dies 5, a first prepreg (T300 grade carbon fiber fabric) is first laid, and then several second prepregs (T700 grade carbon fiber unidirectional material) are laid to form "U" shaped prepreg blanks 15 on the two molding dies 5 respectively. Then the two molding dies 5 are joined together so that the two prepreg blanks 15 are butted together and composite material twist strips 16 are filled at the R-angle after butting. Then the first prepreg is laid in the non-butting area of ​​the two prepreg blanks 15 to form a covering layer 17 connecting the two prepreg blanks 15. Then the curing is performed (autoclave molding can be used), the molding dies 5 are removed, and the "I" shaped structural body 1 is obtained. The layup ratios of the second prepreg in the 0° / ±45° / 90° directions are 60% / 30% / 10%, respectively; the mating area of ​​the two prepreg blanks 15 forms the web 12, while the non-matting area and the cover layer 17 form the upper edge strip 11 and the lower edge strip 13, respectively. S2, Forming of the leading edge plate 2: A mounting fixture with a cross-sectional shape adapted to the transverse cross-section of the leading edge plate 2 is provided. The surface of the mounting fixture serves as the mounting surface of the leading edge plate. After applying a release agent, the first prepreg is first mounted (which can serve as the lower skin after curing), and then foam core material (PMI foam core material can be used) is filled. Then the first prepreg is mounted again (which can serve as the upper skin after curing) to form a sandwich structure leading edge plate blank. Then a vacuum bag is made, and after vacuum compaction and curing, the leading edge plate 2 is obtained. S3, Forming of the rear edge plate 3: A mounting fixture with a cross-sectional shape adapted to the transverse cross-section of the rear edge plate 3 is provided. The surface of the mounting fixture serves as the mounting surface of the rear edge plate. After applying a release agent, the first prepreg is first mounted (which can serve as the lower skin after curing), and then foam core material (PMI foam core material can be used) is filled. Then the first prepreg is mounted again (which can serve as the upper skin after curing) to form a sandwich structure rear edge plate blank. Then a vacuum bag is made, and after vacuum compaction and curing, the rear edge plate 3 is obtained. S4, Metal Connector 4 Machining: A metal connector 4 with a connection hole 41 is formed by machining a lightweight metal material (aluminum alloy); S5. Assembly: First, place the two metal connectors 4 at both ends of the main body 1 and position them to ensure coaxiality. Then, connect and fix them to the ends of the main body 1 using a screw connection method. The first clamping plate 21 on the leading edge plate 2 is inserted into the clamping slot 14 and bonded to the upper edge strip 11 and the lower edge strip 13 with structural adhesive. Then, the second clamping plate 31 on the trailing edge plate 2 is inserted into the clamping slot 14 and bonded to the upper edge strip 11 and the lower edge strip 13 with structural adhesive to obtain the composite material wing diagonal brace. In order to ensure the quality of the bonding, the structural adhesive is pressurized by internal pressure.

[0023] The composite material wing strut and its molding method provided by this invention have the advantages of being lightweight, high-strength, fatigue-resistant, and corrosion-resistant. Furthermore, through the structural arrangement of the leading edge plate 2 and the trailing edge plate 3, the strut can reduce air resistance during flight and improve the aircraft's fuel economy after being installed on the aircraft.

[0024] The above-described preferred embodiments of the present invention are for illustrative purposes only and are not intended to limit the scope of the invention. Any obvious variations or modifications derived from the technical solutions of the present invention are still within the protection scope of the present invention.

Claims

1. A composite material wing diagonal brace, characterized in that, The diagonal strut includes the following structural settings: A structural body (1) with a "H"-shaped cross-section. The two opposite sides of the structural body (1) respectively have clamping grooves (14) extending longitudinally and with opposite openings. The structural body (1) is formed by molding composite materials; A leading-edge plate (2) with a "C"-shaped cross-section. The two edges of the leading-edge plate (2) are provided with sunken first clamping plates (21) along the longitudinal direction. The first clamping plates (21) are clamped and fixed in the clamping grooves (14). The leading-edge plate (2) is formed by molding composite materials; A trailing-edge plate (3) with a "C"-shaped cross-section. The two edges of the trailing-edge plate (3) are provided with sunken second clamping plates (31) along the longitudinal direction. The second clamping plates (31) are clamped and fixed in the other clamping grooves (14). The trailing-edge plate (3) is formed by molding composite materials; And a metal joint (4) provided with connection holes (41). The metal joint (4) is connected to both ends of the structural body (1) and is respectively used for connecting to the wing and the fuselage.

2. The composite material wing diagonal brace according to claim 1, characterized in that, The structural body (1) is formed by connecting an upper flange (11), a web (12) and a lower flange (13) to form a structure with a "H"-shaped cross-section. The clamping grooves (14) are formed between the upper flange (11), the web (1); 3. A composite material wing diagonal brace according to claim 2, characterized in that, The upper flange (11) and the lower flange (13) have the same size and thickness.

4. A composite material wing diagonal brace according to claim 3, characterized in that, The sunken depths of the first clamping plates (21) and the second clamping plates (31) are the same as the thickness of the upper flange (11) or the lower flange (13).

5. A composite material wing diagonal brace according to claim 1, characterized in that, After the first clamping plates (21) and the second clamping plates (31) are clamped in the clamping grooves (14), they are respectively adhesively fixed by structural adhesive.

6. A composite material wing diagonal brace according to claim 1, characterized in that, The metal joint (4) and the end of the structural body (1) are connected by adhesive-bolt connection.

7. A method for forming a composite material wing diagonal brace according to any one of claims 1 to 6, characterized in that, The method includes the following steps: S1. Molding of the structural body (1): Provide a pair of molding dies (5) with three continuous laying surfaces (51), and an R corner is formed between adjacent laying surfaces (51). After applying a release agent on the laying surfaces (51) of the two molding dies (5), first lay a layer of first prepreg, and then lay several second prepregs to respectively form prepreg blanks (15) with a "匚"-shaped structure on the two molding dies (5); Align the two molding dies (5) so that the two prepreg blanks (15) are butted and composite material tallow strips (16) are filled at the R corner. Then, lay the first prepreg at the non-butted parts of the prepreg blanks (15) to form a covering layer (17) connecting the two prepreg blanks (15). Then, cure and mold, and remove the molding dies (5) to obtain the "H"-shaped structural body (1); S! Molding of the leading-edge plate (2): Provide a laying tooling with a cross-section shape adapted to the cross-section of the leading-edge plate (2). The surface of the laying tooling is used as the leading-edge plate laying surface. After applying a release agent, first lay the first prepreg, then fill the foam core material, and then lay the first prepreg again to form a sandwich-structured leading-edge plate blank. Then, make a vacuum bag, and after vacuum compaction and curing, obtain the leading-edge plate (2); S3. Molding of the trailing-edge plate (3): A mounting fixture with a cross-sectional shape adapted to the transverse cross-section of the rear edge plate (3) is provided. The surface of the mounting fixture is used as the mounting surface of the rear edge plate. After applying a release agent, the first prepreg is first mounted, then the foam core material is filled, and the first prepreg is mounted again to form a sandwich structure rear edge plate blank. Then a vacuum bag is made, and after vacuum compaction and curing, the rear edge plate (3) is obtained. S4. Metal joint (4) processing: A metal connector (4) with a connection hole (41) is formed by processing a lightweight metal material. S5. Assembly: First, place the metal connector (4) at both ends of the main body (1) and position it to ensure coaxiality, and then connect and fix it to the main body (1); The first card plate (21) on the leading edge plate (2) is inserted into the card slot (14) and connected and fixed. Then the second card plate (31) on the trailing edge plate (2) is inserted into the card slot (14) and connected and fixed to obtain a composite material wing diagonal brace.

8. The molding method of a composite material wing diagonal brace according to claim 7, characterized in that, The first prepreg is a T300 grade carbon fiber fabric, and the second prepreg is a T700 grade carbon fiber unidirectional material.

9. A method for forming a composite material wing diagonal brace according to claim 7, characterized in that, When laying the second prepreg, the layup ratios in the 0° / ±45° / 90° directions are 60% / 30% / 10%, respectively.

Images