Composite material wall panel structure of aircraft control surface and forming system and method thereof

Abstract

The present invention provides an aircraft rudder composite material wall panel structure and its molding system and method, which relate to the field of aviation manufacturing technology, can realize fast and low-cost manufacturing, and can increase the wall panel rigidity based on the force characteristics of the rudder surface, and design the force transmission path at the same time; the structure is a plate-like structure with a plurality of geometric groove structures; the molding method comprises: placing the ribs in the grooves of the concave grid template, and laying a layer of skin thereon; after preheating, under vacuum conditions, pressing the convex grid template down to close the mold, pressurizing and maintaining the pressure, cooling and shaping, and then demolding to obtain the wall panel structure. The technical solution provided by the present invention is applicable to the process of designing and manufacturing aircraft rudder wall panels.

Description

[Technical field]

[0001] The invention relates to the technical field of aviation manufacturing, and in particular to an aircraft control surface composite material wall plate structure and a molding system and method thereof. [Background technology]

[0002] The design of new aircraft fuselage components requires ultra-lightweight structures within the allowed limited weight. Fiber-reinforced composite materials are widely used in the primary and secondary load-bearing structures of modern aircraft due to their excellent specific stiffness and specific strength. The rudder structure, as the secondary load-bearing structure of the aircraft wing, has been replaced by composite materials relatively early in order to reduce weight and improve vibration characteristics.

[0003] There are two main types of existing aircraft rudder panel structures. One is a reinforced panel structure formed by co-bonding or co-curing long spars such as "T", "I" or "π" with a composite skin; the other is a sandwich panel structure formed by co-curing a foam core and a composite skin. Among them, the reinforced panel structure is more widely used. Compared with traditional metal structures, the above two panel structures can effectively improve the rudder stiffness, reduce the number of fastener connections, reduce fuel consumption, and increase economy. However, the main disadvantage is that the skin is thick and heavy, only regular structures can be prepared, the structural efficiency is limited, and the process cost is high. Moreover, based on the force characteristics of the rudder structure, the force transmission path of the panel structure is difficult to design.

[0004] Among the existing technical solutions, there are four technical solutions that are closest to the proposal of this application:

[0005] The first one is Figure 1 As shown, it is from the patent "Composite stringer and method of manufacturing a composite stringer", application number WO2011121340. This solution involves the forming method and manufacturing process of composite reinforced wall panels. The long stringer is T-shaped, and the web includes an insert and a twist strip design. Two rows of nails are used at the end of the reinforced plate long stringer as a crack stop design. This solution belongs to a typical T-shaped long stringer wall panel co-bonded reinforced wall panel structure.

[0006] Second Figure 2 As shown, it comes from the patent "Process of manufacturing composite panels with U-shaped stiffening members", application number US7959753B2. This solution mainly involves a co-curing integral molding method for a U-shaped long-stringer reinforced wall panel configuration, which is suitable for two typical structures of intermediate interlayer and no interlayer, and does not involve a positioning device between the long stringer and the skin. This solution belongs to a typical U-shaped non-continuous long stringer wall panel co-curing reinforced wall panel structure.

[0007] The third type Figure 3 As shown, it comes from the patent "Composite material reinforced plate liquid molding device and method", application number CN102990940A. This patent mainly designs a composite material reinforced plate liquid molding device and method, including placing the rib structure reinforced fiber material in the groove of the molding mold for molding, and then laying it with the skin structure reinforced fiber material for molding, and then injecting liquid resin for overall molding. This patent does not use the molding process, and the fiber volume content is difficult to control.

[0008] The fourth type Figure 4 As shown, it is from the patent "A composite elevator", application number CN103754353B. This patent mainly relates to a composite elevator structure, which is equipped with a counterweight beam, innovatively designs the docking of the elevator and the trim plate, and adopts a transition joint connection. The rudder surface structure adopts an advanced honeycomb wall panel sparse rib structure and a simple and reliable joint connection form.

[0009] The above four structures all have their drawbacks and cannot meet the requirements for aircraft structure design in the new era.

[0010] Therefore, it is necessary to study a rudder panel structure that is quick, easy to form, low-cost, and highly adaptable, and its forming system and method to address the shortcomings of the prior art and to solve or alleviate one or more of the above problems. [Summary of the invention]

[0011] In view of this, the present invention provides an aircraft rudder composite material wall panel structure and its forming system and method, which can not only achieve fast and low-cost manufacturing, but also increase the wall panel stiffness based on the rudder surface force characteristics and design the force transmission path.

[0012] On the one hand, the present invention provides an aircraft control surface composite material wall panel structure, characterized in that the wall panel structure is a plate-like structure provided with a plurality of geometric groove structures.

[0013] According to the aspects described above and any possible implementation manner, an implementation manner is further provided, wherein the geometric groove is a combination of one or more of a square groove, a triangular groove, a polygonal groove, and a trapezoidal groove.

[0014] According to the above aspects and any possible implementation manner, there is further provided an implementation manner, wherein the geometric grooves are distributed on the same surface of the plate-like structure.

[0015] According to the above aspects and any possible implementation, there is further provided an implementation, wherein the geometric grooves are distributed on two surfaces of the plate-like structure, and the shapes of the grooves on the two surfaces correspond to each other.

[0016] According to the above aspects and any possible implementation, there is further provided an implementation, wherein the material of the wall panel structure includes a skin and ribs.

[0017] On the other hand, the present invention provides a molding system for a composite material wall panel structure of an aircraft control surface, characterized in that the molding system comprises: a female mold with a plurality of concave grid unit blocks on the upper surface, a male mold with a plurality of convex grid unit blocks on the lower surface, a heating device, a cold water circulation device, a vacuum bag and a press;

[0018] The male mold is suspended above the female mold, and the two are opposite to each other; the male mold is connected to the press and completes the pressing and rising actions under the action of the press; a plurality of heating devices for heating the molds and a cold water circulation device for cooling down during shaping are arranged near the male mold and the female mold;

[0019] The concave grid unit block is placed in the vacuum bag, and the vacuum bag is connected to a vacuum pumping device.

[0020] According to the above aspects and any possible implementation manner, an implementation manner is further provided, wherein an expansion pad is provided on the inner wall of the groove of the concave grid unit block.

[0021] In another aspect, the present invention provides a molding method for a composite material wall panel structure of an aircraft control surface, characterized in that the molding method is implemented by using any molding system as described above;

[0022] The steps of the molding method include:

[0023] S1, cutting ribs and preforming;

[0024] S2, placing the preformed ribs in the grooves of the concave grid unit blocks;

[0025] S3, cutting and preforming skin panel materials;

[0026] S4, start the heating device to heat the mold;

[0027] S5, laying the preformed skin panel material on the surface of the concave grid unit block;

[0028] S6, after the vacuum bag is evacuated, the male mold with a plurality of convex grid unit blocks on the lower surface is pressed down to close the mold;

[0029] S7, pressurizing and holding the pressure, then cooling and shaping;

[0030] S8, demoulding, and obtaining the aircraft control surface composite material wall panel structure.

[0031] According to the aspects described above and any possible implementation method, an implementation method is further provided, wherein the shape of the concave grille unit block groove is a combination of one or more of a square groove, a triangular groove, a polygonal groove, and a trapezoidal groove; and the shape of the protrusion of the convex grille unit block corresponds to the shape of the groove.

[0032] According to the above aspects and any possible implementation, an implementation is further provided, wherein the concave grille unit blocks are evenly and regularly laid on the female mold base plate, and the laying method of the convex grille unit blocks corresponds to that of the concave grille unit blocks.

[0033] Compared with the prior art, the present invention can achieve the following technical effects: the aircraft rudder composite material wall panel grid structure and the forming method thereof of the present invention are light in weight, have a large design space, high bending stiffness, good stability, and are easy to maintain while ensuring the stiffness of the wall panel; the molding process is also more likely to achieve rapid, low-cost, and high-quality manufacturing; the molding system is suitable for a variety of composite material grid wall panel structures, and can improve the molding efficiency of composite material parts and reduce manufacturing costs; the addition of ribs, on the one hand, increases the stiffness of the wall panel, and on the other hand, constitutes a force transmission path, so that the force borne by the wall panel has a fixed and effective transmission path.

[0034] Of course, any product implementing the present invention does not necessarily need to achieve all of the above-mentioned technical effects at the same time.

Brief Description of the Drawings

[0035] In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0036] Figure 1-Figure 4 There are four structures in the prior art;

[0037] Figure 5 This is a schematic diagram of the aircraft control surface wall panel grid structure provided by an embodiment of the present invention;

[0038] Figure 6 This is a flow chart of a method for forming an aircraft control surface panel provided by an embodiment of the present invention;

[0039] Figure 7 is a typical grid structure diagram provided by an embodiment of the present invention;

[0040] Figure 8 is a schematic diagram of the grid structure provided by an embodiment of the present invention;

[0041] Fig. 9 is a schematic diagram of a grille mold structure provided by an embodiment of the present invention;

[0042] Fig.10 It is a schematic diagram of the structure of a grille mold expansion pad provided by an embodiment of the present invention;

[0043] Fig.11 It is a schematic diagram of a composite material grid wall panel forming method provided by an embodiment of the present invention;

[0044] Fig.12 A composite material grid wall panel process test piece provided by an embodiment of the present invention.

[0045] Among them, in the figure:

[0046] 1. Wall panel structure; 2. Skin; 3. Edge ribs; 4. Middle ribs; 5. Mold; 6. Mold base plate; 7. Grid unit block; 8. Expansion pad; 9. Vacuum equipment; 10. Female mold; 11. Male mold; 12. Heating device; 13. Cold water circulation; 14. Workbench; 15. Vacuum bag. [Specific implementation method]

[0047] In order to better understand the technical solution of the present invention, the embodiments of the present invention are described in detail below with reference to the accompanying drawings.

[0048] It should be clear that the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0049] The terms used in the embodiments of the present invention are only for the purpose of describing specific embodiments, and are not intended to limit the present invention. The singular forms "a", "said" and "the" used in the embodiments of the present invention and the appended claims are also intended to include plural forms, unless the context clearly indicates other meanings.

[0050] The present invention provides a new aircraft rudder panel grille structure prepared based on a continuous compression molding process and a molding method thereof, which can not only realize rapid and low-cost manufacturing, but also increase the panel stiffness and design the force transmission path based on the rudder surface stress characteristics.

[0051] Aircraft rudder panel grid structure Figure 5As shown. The grille reinforcement structure includes reinforcing ribs and skin 2, and the reinforcing ribs include edge ribs 3 and middle ribs 4. The grille is distributed in a polygonal grid, and the structure exhibits anisotropy. The grille structure has good topological optimization characteristics in geometric form. The aircraft control surface composite wall panel grille unit structure is designed based on the optimized topological relationship. The wall panel grille structure is preliminarily designed according to the load characteristics of the commercial aircraft control surface using commercial optimization software. The detailed design of the grille is carried out using CATIA three-dimensional software, taking into account weight cost, structural strength, manufacturability and maintainability.

[0052] The grid structure includes square, hexagon, square with a cross, positive and negative 45° oblique lines and horizontal lines crossing each other. The crossing shape can be a variety of crossing forms, such as equal and mixed triangles, triangles or polygons of different sizes arranged regularly or irregularly. The grid structure mainly includes but is not limited to Figure 7 The grid structure can be decomposed into several groups of ribs with different orientations, such as Figure 8 As shown, horizontal, vertical and oblique ribs of different angles are combined together. The inclination angle of the oblique ribs is not limited to 45°, but can also be other angles.

[0053] The width and length of the ribs are cut from the prepreg according to the design requirements of the rudder panel 1. The middle ribs 4 are the difficulty in manufacturing the grid structure (the edge ribs refer to Figure 8 The four sides of the grille and the whole shown in the last figure are characterized by being non-intersecting; the middle ribs refer to the ribs that need to intersect. Due to thickness issues, it is more difficult to process the intersecting ribs than the non-intersecting ribs). According to the groove position of the ribs in the mold 5, a grid of the same size is pasted on the aluminum plate using high-temperature PI tape, which is used as a reference for the preformed rib grid. The prepreg tow is prepared into multiple single-layer grille ribs using a fusion welding process to facilitate the subsequent rapid laying of the ribs under the high-temperature mold, saving time and reducing the resin oxidation time.

[0054] The mold 5 adopts a molding process, the upper female mold 10 moves, and the lower male mold 11 is fixed. Fig. 9 As shown. It is mainly composed of a mold base plate 6 and a grid unit block 7. The base plate 6 and the grid unit block 7 are positioned by bolt connection. The composite preformed ribs are filled in the expansion pad groove to form a grid structure.

[0055] The process of forming the composite material wall panel grid structure of the aircraft control surface is as follows Figure 6 As shown, the preparation is carried out by a compression molding process with thermoplastic in-situ heating combined with rapid water cooling, and the steps include: rib cutting → middle rib preforming → edge rib preforming → panel cutting → panel ply preforming → mold preheating → rib laying → panel laying → mold closing and preheating → vacuuming → pressurizing and maintaining pressure → cooling and shaping → demolding.

[0056] The molding tooling includes a mold 5, a mold base plate 6, a grid unit block 7, an expansion pad 8 and a vacuum device 9; the molding equipment includes a female mold 10, a male mold 11, a heating device 12, a cold water circulation 13, a workbench 14, a vacuum bag 15 and a press. The molding mold is provided with an adjustable grid unit block 7 and its expansion pad 8. The molding equipment is provided with a vacuum auxiliary pressure device, an electric heating device, a rapid water cooling device and a press press device. The press presses the male mold 11 from top to bottom, so that the wallboard raw material presents an inwardly concave grid structure under the action of pressure.

[0057] The specific contents include: laying the wall panel prepreg 2 on the upper layer of the mold 5 with the grid structure laid out, making a vacuum bag 15 for the entire mold, and using a vacuum machine 9 and a press machine 16 to simultaneously evacuate and heat 12 the workpiece, and pressurize it to form the required aircraft rudder wall panel shape. Then, a rapid water cooling system 13 is used for shaping and demolding. Generally, compression molding requires preheating for 3 minutes, holding the pressure at 0.7-20MPa for 3 minutes, cooling to 70-100℃ for about 20 minutes, and then demolding.

[0058] The rib structure preform is first placed in the groove of the forming mold, and then the wall panel skin is laid on the mold, vacuum-assisted pressurization, preheating and shape correction, and the molding machine realizes co-curing and one-piece molding. The skin is a layered structure, which includes a conductive layer, a first insulating layer, a heating layer, a second insulating layer and a bearing layer in sequence; the conductive layer is used to provide a lightning path; the first insulating layer and the second insulating layer are used to position the heating layer and protect the bearing layer from different potential corrosion; the heating layer is used for electric heating to prevent deicing; the bearing layer is used to provide bearing function for other layers. The conductive layer is a metal mesh or a conductive polymer mesh material. The heating layer is a copper-based or nickel-based material.

[0059] The aircraft control surface composite material wall panel grid structure and its forming system and method of the present invention are light in weight, have a large design space, high bending stiffness, good stability, and are easy to maintain while ensuring the rigidity of the wall panel. It is also easier to achieve rapid, low-cost, and high-quality manufacturing using a molding process. The proposed forming process method and tooling can be applied to a variety of composite material grid wall panel structures, improve the forming efficiency of composite material parts structures, and reduce manufacturing costs.

[0060] The above is a detailed introduction to an aircraft control surface composite material wall panel structure and its forming method provided by the embodiment of the present application. The description of the above embodiment is only used to help understand the method and its core idea of ​​the present application; at the same time, for those skilled in the art, according to the idea of ​​the present application, there will be changes in the specific implementation method and application scope. In summary, the content of this specification should not be understood as limiting the present application.

[0061] For example, certain words are used in the specification and claims to refer to specific components. Those skilled in the art should understand that hardware manufacturers may use different nouns to refer to the same component. This specification and claims do not use differences in names as a way to distinguish components, but use differences in the functions of components as the criteria for distinction. As mentioned throughout the specification and claims, "including" and "comprising" are open-ended terms, so they should be interpreted as "including / including but not limited to". "Approximately" means that within an acceptable error range, those skilled in the art can solve the technical problem within a certain error range and basically achieve the technical effect. The subsequent description of the specification is a preferred embodiment of the present application, but the description is for the purpose of illustrating the general principles of the present application, and is not used to limit the scope of the present application. The scope of protection of the present application shall be determined by the definition of the attached claims.

[0062] It should also be noted that the terms "include", "comprises" or any other variations thereof are intended to cover non-exclusive inclusion, so that a product or system including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such a product or system. In the absence of more restrictions, the elements defined by the sentence "comprises a ..." do not exclude the existence of other identical elements in the product or system including the elements.

[0063] It should be understood that the term "and / or" used in this article is only a description of the association relationship of associated objects, indicating that there can be three relationships. For example, A and / or B can represent: A exists alone, A and B exist at the same time, and B exists alone. In addition, the character " / " in this article generally indicates that the associated objects before and after are in an "or" relationship.

[0064] The above description shows and describes several preferred embodiments of the present application, but as mentioned above, it should be understood that the present application is not limited to the form disclosed herein, and should not be regarded as excluding other embodiments, but can be used in various other combinations, modifications and environments, and can be modified within the scope of the application concept described herein through the above teachings or the technology or knowledge of the relevant field. The changes and modifications made by those skilled in the art do not deviate from the spirit and scope of the present application, and should be within the scope of protection of the claims attached to the present application.

Claims

1. An aircraft control surface composite material wall panel structure, characterized in that: The wall plate structure is a plate-like structure provided with a plurality of geometric groove structures; The geometric groove is a combination of one or more of a square groove, a triangular groove, a polygonal groove, and a trapezoidal groove; The geometric grooves are distributed on two surfaces of the plate-like structure, and the shapes of the grooves on the two surfaces correspond to each other; The material of the wall panel structure includes skin and ribs; The ribs include edge ribs and middle ribs, the middle ribs are arranged to intersect, and during production, space for the ribs is reserved in the mold; According to the groove position of the rib in the mold, a grid of the same size is pasted on the aluminum plate using high-temperature PI tape as a reference for the preformed rib grid, and the prepreg tow is prepared into multiple single-layer grid ribs using a fusion welding process; A plurality of the geometric groove structures constitute a grid structure, and the grid structure includes a cross shape of positive and negative 45° oblique lines and horizontal lines, and the cross shape includes one or more of equal-sized and mixed triangles, regularly or irregularly arranged triangles or polygons of different sizes.

2. A molding system for composite material wall panel structure of aircraft control surface, characterized in that: The molding system comprises: a female mold with a plurality of concave grid unit blocks on the upper surface, a male mold with a plurality of convex grid unit blocks on the lower surface, a heating device, a cold water circulation device, a vacuum bag and a press; The male mold is suspended above the female mold, and the two are opposite to each other; the male mold is connected to the press and completes the pressing and rising actions under the action of the press; a plurality of heating devices for heating the molds and a cold water circulation device for cooling down during shaping are arranged near the male mold and the female mold; The concave grid unit block is placed in the vacuum bag, and the vacuum bag is connected to a vacuum pumping device; The ribs include edge ribs and middle ribs, the middle ribs are arranged to intersect, and during production, space for the ribs is reserved in the mold; According to the groove position of the rib in the mold, a grid of the same size is pasted on the aluminum plate using high-temperature PI tape as a reference for the preformed rib grid, and the prepreg tow is prepared into multiple single-layer grid ribs using a fusion welding process; The wall panel prepreg is laid on the upper layer of the mold with the grid structure laid out, and the vacuum bag is made for the entire mold. The workpiece is simultaneously vacuumed, heated and pressurized by the vacuum equipment and the press to form the required aircraft rudder wall panel shape; the cold water circulation device is used for shaping and demolding; the compression molding is preheated for 3 minutes, maintained at a pressure of 0.7-20MPa for 3 minutes, cooled to 70-100°C, and demolded after 20 minutes.

3. The molding system of composite material wall panel structure of aircraft control surface according to claim 2, characterized in that: An expansion pad is provided on the inner wall of the groove of the concave grid unit block.

4. A method for forming a composite material wall panel structure of an aircraft control surface, characterized in that: The molding method is used to manufacture the composite material wall panel structure of the aircraft control surface as claimed in claim 1; The steps of the molding method include: S1, cutting ribs and preforming; S2, placing the preformed ribs in the grooves of the concave grid unit blocks; S3, cutting and preforming skin panel materials; S4, start the heating device to heat the mold; S5, laying the preformed skin panel material on the surface of the concave grid unit block; S6, after the vacuum bag is evacuated, the male mold with a plurality of convex grid unit blocks on the lower surface is pressed down to close the mold; S7, pressurizing and holding the pressure, then cooling and shaping; S8, demoulding to obtain the aircraft control surface composite material wall panel structure; The ribs include edge ribs and middle ribs, the middle ribs are arranged to intersect, and during production, space for the ribs is reserved in the mold; According to the groove position of the rib in the mold, a grid of the same size is pasted on the aluminum plate using high-temperature PI tape as a reference for the preformed rib grid, and the prepreg tow is prepared into multiple single-layer grid ribs using a fusion welding process.

5. The method for forming a composite material wall panel structure of an aircraft control surface according to claim 4, characterized in that: The shape of the groove of the concave grille unit block is a combination of one or more of a square groove, a triangular groove, a polygonal groove, and a trapezoidal groove; the shape of the protrusion of the convex grille unit block corresponds to the shape of the groove.

6. The method for forming a composite material wall panel structure of an aircraft control surface according to claim 4, characterized in that: The concave grille unit blocks are evenly and regularly laid on the female mold base plate, and the laying method of the convex grille unit blocks corresponds to that of the concave grille unit blocks.

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