Composite flap control surface structure

Abstract

The utility model discloses a kind of composite aileron control surface structures in the technical field of aircraft wing structure design, including fixedly installed in the aileron control surface of aircraft flying wing rear suspension bracket, the aileron control surface is composed of inner aileron and aileron adjusting piece;The inner aileron includes sandwich wallboard component and first skeleton component;The sandwich wallboard component includes upper wallboard and lower wallboard, the first skeleton component is set between upper wallboard and lower wallboard, and its upper and lower sides are connected with upper wallboard and lower wallboard respectively.The utility model is designed by using honeycomb sandwich structure for the sandwich wallboard component of inner aileron, aileron adjusting piece uses full height honeycomb structure, and the main load-bearing structure of inner aileron and aileron adjusting piece all uses carbon fiber composite material, fully utilizes the characteristics of high specific strength, high specific stiffness and low density of carbon fiber composite material, significantly reduces structure weight, improves overall efficiency of structure, solves the problem of large density, heavy weight and low structure efficiency of traditional metal structure aileron control surface.

Description

Technical Field

[0001] This utility model relates to the field of aircraft wing structure design technology, and in particular to a composite material aileron control surface structure. Background Technology

[0002] The ailerons are a core component of the aircraft's primary control surfaces, typically arranged symmetrically on the outer trailing edges of the left and right wings. Their core function is to respond to the pilot's lateral control commands by differentially deflecting the ailerons. When one aileron deflects upwards to reduce lift, the other deflects downwards to increase lift, thereby generating a control torque that causes the aircraft to roll around its longitudinal axis. This roll motion is fundamental to the aircraft's coordinated turns, maintaining lateral stability, and control.

[0003] In traditional aircraft design, aileron control surfaces typically employ a "metal frame and thin metal skin" structure. This structure is implemented as follows: the internal frame mainly consists of metal beams as longitudinal reinforcements and metal ribs as lateral supports; these key components are usually precision-machined using CNC machine tools. The external aerodynamic shape is formed by a thin metal skin, which is produced through sheet metal stamping. It is worth noting that because the skin dimensions of the upper and lower aileron surfaces are often large, to meet manufacturing requirements, the entire skin is often divided into multiple pieces, formed separately, and then assembled and installed on-site.

[0004] However, the "metal frame + thin metal skin" structure has certain problems. First, in terms of structural efficiency, the high density of metal materials results in a large overall weight of the control surface, which contradicts the aviation industry's continuous pursuit of weight reduction and efficiency improvement, directly affecting the aircraft's payload capacity and fuel economy. Second, in terms of manufacturing and assembly, this structure involves a large number of independent components such as beams, ribs, and multiple skin panels, resulting in a large number of fasteners (such as rivets and bolts). This not only makes the assembly process numerous and complex but also significantly increases manufacturing costs and time. Finally, in terms of durability and maintainability, metal structures, especially at their joints, have relatively poor corrosion resistance in complex environments such as humidity and salt spray. This not only reduces the overall durability of the structure but also makes it more susceptible to fatigue cracks induced by stress corrosion, posing a potential threat to the fatigue life and safety reliability of the control surface and increasing the maintenance cost throughout its life cycle. Therefore, we urgently need a composite material aileron control surface structure to solve the above problems. Utility Model Content

[0005] To address the shortcomings of existing technologies, this invention provides a new technical solution for a composite material aileron control surface structure. It employs carbon fiber composite materials, a honeycomb sandwich structure, and a full-height honeycomb structure, combined with integrated adhesive bonding and simplified joint connections. This achieves weight reduction and simplified assembly processes for the aileron control surface, while simultaneously improving corrosion resistance and fatigue performance. It solves the problems of heavy weight, complex assembly, and poor durability associated with traditional metal structures. The modular design of the inner aileron and trim tabs, combined with a composite laminate skeleton and honeycomb core sandwich structure, ensures uniform load transfer and aerodynamic precision, significantly improving structural efficiency and meeting the aircraft's requirements for lightweight and high reliability in lateral control.

[0006] The purpose of this utility model is achieved as follows: A composite material aileron control surface structure includes an aileron control surface fixedly installed on the rear suspension bracket of an aircraft wing. The aileron control surface is composed of an inner aileron and an aileron trim tab. The inner aileron includes a sandwich panel assembly and a first frame assembly. The sandwich panel assembly includes an upper panel and a lower panel. The first frame assembly is disposed between the upper panel and the lower panel, and its upper and lower sides are respectively connected to the upper panel and the lower panel via countersunk bolts. Both the upper panel and the lower panel consist of an outer skin, an inner skin, and a frame sandwiched therebetween. The first honeycomb core is formed by adhesive bonding, and the first skeleton assembly is connected to the rear spars of the outer wing of the aircraft via a suspension joint and a drive joint; the aileron trim piece includes a second skeleton assembly and a second honeycomb core bonded within the second skeleton assembly; the aileron trim piece also includes an upper skin and a lower skin, the second skeleton assembly is disposed between the upper skin and the lower skin, and its upper and lower sides are respectively connected to the upper skin and the lower skin; the second skeleton assembly is connected to the inner aileron via a connecting joint, and an actuation joint connected to the actuation cylinder on the aircraft wing is installed on the lower skin.

[0007] Optionally, the first frame assembly includes a first front beam, a first rear beam, a wedge-shaped trailing edge strip, a first support rib, and a reinforcing angle box; the first front beam and the first rear beam are arranged parallel to each other, the first support rib is arranged between the first front beam and the first rear beam, and the two ends of the first support rib are fixedly connected to the first front beam and the first rear beam respectively, the wedge-shaped trailing edge strip is fixedly connected to the first rear beam, the suspension joint and the drive joint are both fixedly installed on the first front beam, the first rear beam is equidistantly equipped with suspension joints for connecting aileron trim tabs, and the reinforcing angle box is fixedly connected to the connection points of the components of the first frame assembly.

[0008] Optionally, the number of suspension joints is two sets, and the drive joint is disposed between the two sets of suspension joints.

[0009] Optionally, a first front edge rib is fixedly connected to the front edge of the first front beam, a first skin is fixedly connected to the first front edge rib, and maintenance cover is provided on the upper and lower wing surfaces of the first skin.

[0010] Optionally, the first skeleton assembly includes a first front beam, a first rear beam, a first support rib, a reinforcing corner box, a suspension joint, a first leading edge rib, and a first skin, all of which are composite material laminate structures.

[0011] Optionally, the second frame assembly includes a second front beam, a second rear beam, a second support rib, and a second front edge rib; the second front beam and the second rear beam are arranged parallel to each other, the second support rib is located between the second front beam and the second rear beam, the two ends of the second support rib are fixedly connected to the second front beam and the second rear beam respectively, and the second front edge rib is fixedly connected to the front edge of the second front beam.

[0012] Optionally, a second skin is fixedly connected to the second leading edge rib.

[0013] Optionally, the second front beam, the second rear beam, the second support rib, the second leading edge rib, and the second skin are all composite laminate structures.

[0014] Optionally, the number of connecting joints is four sets, and the four sets of connecting joints are installed equidistantly on the second front beam, and the connecting joints are connected to the suspension joints.

[0015] Optionally, the first honeycomb core is a honeycomb core material used in sandwich structures, and the second honeycomb core is a full-height honeycomb core. Both the first and second honeycomb cores are made of carbon fiber composite material.

[0016] Compared with the prior art, the beneficial effects of this utility model are as follows: 1. According to an embodiment of this disclosure, the composite material aileron control surface structure adopts a honeycomb sandwich structure of "outer skin + inner skin + honeycomb core" for the sandwich panel assembly (upper panel + lower panel) of the inner aileron, and adopts a full-height honeycomb structure of "skeleton assembly + full-height honeycomb core + upper skin + lower skin" for the aileron trim piece. Moreover, the main load-bearing structures (such as skeleton assembly) of the inner aileron and the aileron trim piece are all designed with carbon fiber composite materials, which makes full use of the high specific strength, high specific stiffness and low density characteristics of carbon fiber composite materials. Compared with the traditional "metal frame + thin metal skin" aileron control surface, it significantly reduces the structural weight, improves the overall structural efficiency, and solves the problems of high density, heavy weight and low structural efficiency of traditional metal structure aileron control surfaces.

[0017] 2. According to one embodiment of this disclosure, the composite material aileron control surface structure is integrated by using countersunk pins to connect the inner aileron skeleton assembly and the sandwich panel assembly. The aileron trim piece skeleton assembly, full-height honeycomb core, upper skin, and lower skin are integrally glued together after being positioned by adhesive bonding fixtures. At the same time, the inner aileron and aileron trim piece are integrated through only four sets of connecting joints and suspension joints to achieve the overall design. This leverages the advantages of composite material integration, significantly reduces the number of components and fasteners required by traditional metal structures, simplifies the assembly and manufacturing process, and avoids the cumbersome process of traditional metal skins requiring segmented molding and multiple fastener splicing due to their large size. This solves the problem of numerous components and complicated assembly processes in traditional metal structure aileron control surfaces. Attached Figure Description

[0018] To more clearly illustrate the technical solutions in the embodiments of this utility model or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only embodiments of this utility model. For those skilled in the art, other drawings can be obtained based on the provided drawings without creative effort.

[0019] Figure 1 This is a schematic diagram of the overall exploded structure of this utility model.

[0020] Figure 2 This is a schematic diagram of the first honeycomb core structure of this utility model.

[0021] Figure 3 This is a schematic diagram of the second honeycomb core structure of this utility model.

[0022] Figure 4 This is a front view structural diagram of the present invention.

[0023] Figure 5 This is a side view of the structure of this utility model.

[0024] Figure 6 This is a schematic diagram of the first skeleton component of this utility model.

[0025] Figure 7 This is a schematic diagram of the second skeleton component of this utility model.

[0026] The diagram shows: 1. Inner aileron; 2. Aileron trim tab; 3. Sandwich panel assembly; 31. Upper panel; 32. Lower panel; 33. Outer skin; 34. Inner skin; 35. First honeycomb core; 4. First frame assembly; 41. Suspension joint; 42. Drive joint; 43. First front beam; 44. First rear beam; 45. Wedge-shaped tail flange; 46. First support rib; 47. Reinforcing angle box; 48. Suspension joint; 49. First leading edge rib; 410. First skin; 411. Service port cover; 5. Second frame assembly; 51. Second front beam; 52. Second rear beam; 53. Second support rib; 54. Second leading edge rib; 55. Second skin; 6. Second honeycomb core; 7. Upper skin; 8. Lower skin; 9. Connecting joint; 10. Actuating joint. Detailed Implementation

[0027] The technical solutions of the present utility model will be clearly and completely described below with reference to the accompanying drawings of the embodiments. Obviously, the described embodiments are only some embodiments of the present utility model, and not all embodiments. Based on the embodiments of the present utility model, all other embodiments obtained by those of ordinary skill in the art without creative effort are within the protection scope of the present utility model.

[0028] like Figure 1-7 As shown, a composite material aileron control surface structure includes an aileron control surface fixedly installed on the rear suspension bracket of an aircraft wing. The aileron control surface is composed of an inner aileron 1 and an aileron trim tab 2.

[0029] Here, the aileron control surface is fixedly mounted on the suspension bracket of the aircraft's outer wing rear spars. The entire surface consists of the inner aileron 1 and the aileron trim tab 2, connected by an adapter joint to form a complete control surface. It can rotate flexibly under the control of the control system, providing lateral balance and turning torque for the aircraft. The inner aileron 1 includes a sandwich panel assembly 3 and a first frame assembly 4. The sandwich panel assembly 3 consists of an upper panel 31 and a lower panel 32. The first frame assembly 4 is positioned between the upper panel 31 and the lower panel 32, connected to both panels via countersunk pins. Both the upper panel 31 and the lower panel 32 employ a structure where the outer skin 33, inner skin 34, and first honeycomb core 35 are bonded together. A co-bonding process is used: first, the outer skin 33 is cured; then, the first honeycomb core 35 is placed and the inner skin 34 is laid; finally, the entire assembly is cured. The first frame assembly 4 is connected to the aircraft's outer wing rear spars via a suspension joint 41 and a drive joint 42, enabling load transfer between the inner aileron 1 and the aircraft.

[0030] Furthermore, the design employs a split structure of the inner aileron 1 and aileron trim piece 2, combined with a composite honeycomb sandwich design. Compared to the traditional metal "frame + segmented skin" structure, this significantly reduces the number of components and fasteners, simplifying the assembly process. The co-bonding process and the sandwich structure formed by the first honeycomb core 35 reduce internal defects and improve the overall structural integrity. At the same time, the low density of the composite material reduces the structural weight, solving the problems of heavy weight and cumbersome assembly in traditional metal structures. The composite material also has good corrosion resistance, avoiding the corrosion defects of traditional metal structures, extending the structural fatigue life, and increasing the aircraft's service life.

[0031] The inner aileron 1 includes a sandwich panel assembly 3 and a first frame assembly 4; the sandwich panel assembly 3 includes an upper panel 31 and a lower panel 32, and the first frame assembly 4 is disposed between the upper panel 31 and the lower panel 32, with its upper and lower sides respectively connected to the upper panel 31 and the lower panel 32 by countersunk pins.

[0032] Here, when the first frame assembly 4 is embedded between the upper wall panel 31 and the lower wall panel 32, it needs to be positioned using a tooling fixture to ensure that the edges of the first front beam 43, the first rear beam 44, and the upper wall panel 31 and the lower wall panel 32 are aligned. The first frame assembly 4 consists of the first front beam 43, the first rear beam 44, the wedge-shaped tail edge strip 45, the first support rib 46, and the reinforcing corner box 47. The first front beam 43, the first rear beam 44, the first support rib 46, and the reinforcing corner box 47 are all composite laminate structures. The wedge-shaped tail edge strip 45 is molded and then machined to achieve its shape. Countersunk pins are used to connect the first frame assembly 4 to the upper wall panel 31 and the lower wall panel 32. The connection positions are distributed along the first front beam 43, the first rear beam 44, and the first support rib 46 to ensure uniform connection strength and avoid local stress concentration.

[0033] Furthermore, the tooling frame positioning and countersunk riveting connection ensure the assembly accuracy of the first frame component 4 with the upper wall panel 31 and lower wall panel 32, reducing assembly errors. The frame components such as the first front beam 43 and the first rear beam 44, made of composite material laminate, are lighter and stronger than traditional metal frames. The molding and machining process of the wedge-shaped tail strip 45 simplifies the forming process and improves production efficiency. The countersunk riveting connection avoids exposed fasteners, ensures the flatness of the wall panel surface, reduces aerodynamic resistance, and the detachable design facilitates later maintenance, solving the problems of multiple fastener riveting and difficult maintenance in traditional metal structures.

[0034] The upper wall panel 31 and the lower wall panel 32 are both formed by bonding an outer skin 33, an inner skin 34 and a first honeycomb core 35 sandwiched therein. The first frame assembly 4 is connected to the rear beam of the aircraft outer wing through a suspension joint 41 and a drive joint 42.

[0035] Here, the outer skin 33 and inner skin 34 of the upper wall panel 31 and the lower wall panel 32 are both composite laminates, each with a thickness of 1mm. The first honeycomb core 35 is machined and requires positioning with the aid of a projector during assembly to ensure the fit between the first honeycomb core 35 and the outer skin 33 and inner skin 34. The bonding process adopts a co-bonding process: the outer skin 33 is cured separately first, and after cooling, the first honeycomb core 35 is placed on the outer skin 33, then the inner skin 34 is laid, and finally the whole is cured into one piece to form a complete sandwich panel structure, reducing defects such as air bubbles and delamination inside the panel.

[0036] Furthermore, the sandwich structure of the outer skin 33, inner skin 34, and first honeycomb core 35 utilizes the high specific strength and high specific stiffness characteristics of composite materials to significantly reduce weight while ensuring the strength of the wall panel, resulting in a significant weight reduction compared to traditional thin metal skins. Projector-assisted positioning improves the assembly accuracy of the first honeycomb core 35, reduces internal defects, and enhances the fatigue resistance of the wall panel. Co-bonding process enhances the overall integrity of the wall panel, avoids the weak points caused by the segmented splicing of traditional metal skins, and reduces assembly steps, thereby improving production efficiency.

[0037] The first frame assembly 4 includes a first front beam 43, a first rear beam 44, a wedge-shaped tail edge strip 45, a first support rib 46, and a reinforcing corner box 47. The first front beam 43 and the first rear beam 44 are arranged parallel to each other. The first support rib 46 is arranged between the first front beam 43 and the first rear beam 44, and the two ends of the first support rib 46 are fixedly connected to the first front beam 43 and the first rear beam 44, respectively. The wedge-shaped tail edge strip 45 is fixedly connected to the first rear beam 44. The suspension joint 41 and the drive joint 42 are both fixedly installed on the first front beam 43.

[0038] Here, the first front beam 43 and the first rear beam 44 are arranged parallel to each other, both being composite laminate structures with a base thickness of 1.6mm, and the thickness is increased to 2mm in the connection area with the suspension joint 41 and the drive joint 42; the first support rib 46 is set between the first front beam 43 and the first rear beam 44, with a thickness of 1mm, and both ends are fixedly connected to the first front beam 43 and the first rear beam 44 through standard parts; the wedge-shaped tail strip 45 is fixed on the first rear beam 44, and is formed by molding and then machined to obtain the required shape; the reinforcing corner box 47 has a thickness of 1.2mm and is fixed at the connection points of the components of the first frame assembly 4, especially at the connection points of the suspension joint 41, the drive joint 42 and the suspension joint 48 with the first front beam 43 and the first rear beam 44, to ensure the connection strength at the joints. Suspension joints 48 are equidistantly installed on the first rear beam 44 for connecting the aileron trim tabs 2. The first front beam 43, the first rear beam 44, the wedge-shaped tail fin strip 45, the first support rib 46, and the reinforcing angle box 47 are connected by standard parts to form a complete first frame assembly 4.

[0039] Furthermore, the thickened design of the connection area of ​​the first front beam 43 can improve the load-bearing capacity of the suspension joint 41 and drive joint 42 and avoid stress concentration; the setting of the reinforced corner box 47 further strengthens the strength of the connection and joint of each component, solving the problem of fatigue damage of traditional metal frame joints; the frame components with standard parts connection are flexible to assemble and easy to maintain, reducing welding stress and improving structural durability compared with traditional metal frame welded structures; the first frame component 4 made of composite material laminate is lightweight and corrosion resistant, significantly reducing structural weight and extending fatigue life compared with traditional metal frames.

[0040] Suspension joints 48 for connecting aileron adjustment pieces 2 are equidistantly installed on the first rear beam 44. The reinforcing angle box 47 is fixedly connected to the connection points of each component of the first frame assembly 4. There are two sets of suspension joints 41, and the drive joint 42 is located between the two sets of suspension joints 41.

[0041] Here, there are two sets of suspension joints 41, symmetrically fixedly installed on the outer side of the web of the first front beam 43; the drive joint 42 is fixedly installed on the first front beam 43 and located between the two sets of suspension joints 41. The suspension joints 41 are used to connect the inner aileron 1 to the rear spars of the outer wing of the aircraft, and the drive joints 42 are used to connect the drive rod to realize the up and down deflection of the inner aileron 1. Four sets of suspension joints 48 are equidistantly installed on the first rear beam 44 for connecting the aileron trim tabs 2. The connection points of the suspension joints 41, drive joints 42, and suspension joints 48 with the first front beam 43 and the first rear beam 44 are all fixed with reinforcing angle boxes 47. The reinforcing angle boxes 47 improve the connection strength between the joint and the beam and prevent structural damage at the joint.

[0042] Furthermore, the symmetrical arrangement of the two sets of suspension joints 41 and the centrally positioned drive joint 42 ensures that the inner aileron 1 is subjected to uniform force and the load is smoothly transferred to the rear spars of the outer wing, avoiding structural deformation caused by concentrated load on one side and improving the flexibility and stability of the inner aileron 1's deflection. The four sets of equidistant suspension joints 48 can evenly distribute the load of the aileron trim piece 2 to the first rear spars 44, ensuring a reliable connection between the aileron trim piece 2 and the inner aileron 1. The addition of reinforcing corner boxes 47 at the joints solves the problem of insufficient connection strength of traditional metal structure joints, improves structural durability, and extends service life.

[0043] A first front edge rib 49 is fixedly connected to the front edge of the first front beam 43, and a first skin 410 is fixedly connected to the first front edge rib 49. Maintenance cover 411 is provided on the upper and lower wing surfaces of the first skin 410.

[0044] Here, the first leading edge rib 49 is fixedly connected to the leading edge of the first front beam 43 and is a composite laminate structure; the first skin 410 is also a composite laminate structure, fixedly connected to the first leading edge rib 49, covering the leading edge area of ​​the inner aileron 1, and playing a role in resisting airflow impact. Maintenance hatches 411 are provided on the upper and lower wing surfaces of the first skin 410. The maintenance hatches 411 are detachably connected to the first skin 410, which facilitates the later inspection and maintenance of the internal structures such as the first frame assembly 4, suspension joint 41, and drive joint 42.

[0045] Furthermore, the composite material laminate structure of the first leading edge rib 49 and the first skin 410 is lightweight and high-strength, and can effectively resist airflow impact and foreign object impact during flight. Compared with the traditional metal leading edge structure, it is lighter and has better impact resistance. The maintenance cover 411 allows for internal maintenance without disassembling the inner aileron 1 as a whole, simplifying the maintenance process, shortening maintenance time, reducing maintenance costs, and avoiding the cumbersome problem of disassembling the whole structure for maintenance of traditional metal structures.

[0046] The aileron trim piece 2 includes a second skeleton assembly 5 and a second honeycomb core 6 glued within the second skeleton assembly 5.

[0047] Here, the aileron adjustment piece 2 is based on the second frame assembly 5 and the second honeycomb core 6 as its core structure. The second honeycomb core 6 is a full-height honeycomb core, formed by CNC machining equipment. During processing, the thickness needs to be adjusted according to the digital model. Because the thickness of the second honeycomb core 6 will shrink during the bonding and curing process, the actual processed thickness needs to be greater than the design value, and the thickness tolerance is set to (0, +0.4mm). The second frame assembly 5 consists of the second front beam 51, the second rear beam 52, the second support rib 53, and the second leading edge rib 54. After each component is positioned and assembled by the tooling fixture, it is bonded to the second honeycomb core 6 as a whole. During the bonding process, it needs to be positioned on the bonding fixture to ensure the relative positional accuracy of the second honeycomb core 6 and the second frame assembly 5.

[0048] Furthermore, the application of the full-height second honeycomb core 6, compared to the foam core or segmented honeycomb core of traditional metal trim pieces, results in more uniform structural stiffness, which can improve the buckling resistance of the aileron trim piece 2. CNC machining and thickness tolerance design ensure that the thickness of the second honeycomb core 6 after adhesive bonding and curing meets the design requirements, avoiding aerodynamic shape deviations caused by shrinkage. The tooling fixture and adhesive bonding fixture positioning improve the assembly accuracy of the second frame assembly 5 and the second honeycomb core 6, ensuring the structural integrity of the aileron trim piece 2, while utilizing the properties of composite materials to reduce the weight of the trim piece and improve the aircraft's handling flexibility.

[0049] The aileron trim piece 2 also includes an upper skin 7 and a lower skin 8. The second frame assembly 5 is disposed between the upper skin 7 and the lower skin 8, and its upper and lower sides are respectively connected to the upper skin 7 and the lower skin 8.

[0050] Here, both the upper skin 7 and the lower skin 8 are composite laminate structures with a thickness of 0.8mm. To ensure aerodynamic shape accuracy, a negative mold laying process is used. After the second skeleton component 5 and the second honeycomb core 6 are bonded together, they are placed between the upper skin 7 and the lower skin 8. The three are positioned by a bonding fixture to ensure the fit between the upper skin 7, the lower skin 8 and the second skeleton component 5 and the second honeycomb core 6. Then, the whole assembly is placed in a thermostatic precipitator for bonding to form a complete aileron adjustment piece 2 structure. The bonding process requires control of temperature and pressure to ensure bonding quality and avoid defects such as delamination and bubbles.

[0051] Furthermore, the negative mold laying molding process can accurately guarantee the aerodynamic shape of the upper skin 7 and lower skin 8. Compared with the traditional metal skin stamping molding, it reduces the segmented splicing process, improves the shape accuracy, and reduces aerodynamic drag. The overall adhesive bonding molding process makes the aileron trim piece 2 an integrated structure, reducing the number of components and fasteners, simplifying the assembly process, and improving production efficiency. The combination of the composite material upper skin 7 and lower skin 8 with the full-height second honeycomb core 6 significantly reduces the weight of the aileron trim piece 2, while improving its fatigue resistance and corrosion resistance, solving the problems of heavy weight and easy corrosion of traditional metal trim pieces.

[0052] The second frame assembly 5 includes a second front beam 51, a second rear beam 52, a second support rib 53, and a second front edge rib 54. The second front beam 51 and the second rear beam 52 are arranged parallel to each other. The second support rib 53 is located between the second front beam 51 and the second rear beam 52. The two ends of the second support rib 53 are fixedly connected to the second front beam 51 and the second rear beam 52, respectively. The second front edge rib 54 is fixedly connected to the front edge of the second front beam 51, and a second skin 55 is fixedly connected to the second front edge rib 54.

[0053] Here, the second front beam 51 and the second rear beam 52 are arranged parallel to each other, both being composite laminate structures. The second support rib 53 is located between the second front beam 51 and the second rear beam 52, with both ends fixedly connected to the second front beam 51 and the second rear beam 52, respectively. The second leading edge rib 54 is fixedly connected to the leading edge of the second front beam 51, also being a composite laminate structure. The second front beam 51, the second rear beam 52, the second support rib 53, and the second leading edge rib 54 are positioned by a tooling frame and pre-riveted together with rivets to form a complete second skeleton assembly 5. The second skin 55 is a composite laminate structure, fixedly connected to the second leading edge rib 54, and seamlessly connected to the leading edges of the upper skin 7 and the lower skin 8, ensuring the aerodynamic continuity of the leading edge of the aileron trim piece 2.

[0054] Furthermore, the tooling fixture positioning and pre-riveting assembly ensure the relative positional accuracy of each component of the second frame assembly 5, reduce assembly errors, and improve the structural stability of the aileron trim 2. The second frame assembly 5 and the second skin 55, made of composite material laminate, are lightweight and high-strength. Compared with traditional metal frames, the weight is significantly reduced, while also having good corrosion resistance and extending the structural life. The seamless connection between the second skin 55 and the upper skin 7 and lower skin 8 avoids leading-edge airflow separation, reduces aerodynamic drag, and improves the aerodynamic control efficiency of the aileron trim 2.

[0055] The second front beam 51, the second rear beam 52, the second support rib 53, the second front edge rib 54, and the second skin 55 are all composite material laminate structures.

[0056] Here, the second front beam 51, second rear beam 52, second support rib 53, second leading edge rib 54, and second skin 55 are all made of carbon fiber composite laminate structure. The laminate is made by prepreg laying, molding and curing. The layup direction is designed according to the stress requirements of each component to ensure that the strength and stiffness meet the usage requirements. Among them, the second front beam 51 has increased thickness in the connection area with the connecting joint 9 by increasing the number of prepreg layers to enhance the load-bearing capacity at the joint; the layup design of the second skin 55 must take into account both aerodynamic shape and surface strength to avoid surface damage during flight.

[0057] Furthermore, the carbon fiber composite laminate has the characteristics of high specific strength, high specific stiffness, and low density. Compared with traditional metal materials, it significantly reduces the weight of components such as the second front beam 51 and the second rear beam 52, while improving the structural strength and stiffness, ensuring that the aileron trim 2 has small deformation when subjected to aerodynamic and inertial loads. The carbon fiber composite material has excellent corrosion resistance and fatigue resistance, which can effectively resist the erosion of harsh environments such as high-altitude salt spray and high humidity, reduce structural corrosion damage, extend the fatigue life of the aileron trim 2, and reduce the total life maintenance cost.

[0058] The second frame assembly 5 is connected to the inner aileron 1 via a connecting joint 9, and an actuation joint 10 connected to the actuation cylinder on the wing is installed on the lower skin 8.

[0059] Here, four sets of connecting joints 9 are equidistantly installed on the second front beam 51, corresponding one-to-one with the suspension joints 48 on the first rear beam 44. Connecting joints 9 and suspension joints 48 are connected by fasteners to fix the second frame assembly 5 to the inner aileron 1, thereby making the aileron trim piece 2 and the inner aileron 1 integrated. The actuation joint 10 is fixedly installed on the lower skin 8 and is used to connect to the actuation cylinder on the aircraft wing. Through the pushing and pulling action of the actuation cylinder, the aileron trim piece 2 is driven to deflect up and down, cooperating with the inner aileron 1 to achieve lateral attitude adjustment of the aircraft. The connection between the actuation joint 10 and the lower skin 8 needs to be strengthened with a reinforced structure to prevent deformation of the lower skin 8 caused by actuation loads.

[0060] Furthermore, the cooperation of the four sets of equidistant connecting joints 9 and the suspension joint 48 ensures that the connection between the aileron trim piece 2 and the inner aileron 1 is evenly stressed, the load is transmitted smoothly, and the joint damage caused by local load concentration is avoided, thus improving the reliability of the connection. The connection between the actuation joint 10 and the actuation cylinder on the aircraft wing makes the aileron trim piece 2 deflection response sensitive, and can precisely fine-tune the lateral attitude of the aircraft. Compared with traditional metal trim pieces, the control precision is higher. The detachable design of the connecting joint 9 makes it easy to disassemble and maintain the aileron trim piece 2 separately without disassembling the entire inner aileron 1, reducing the difficulty and cost of maintenance.

[0061] The first honeycomb core 35 is a honeycomb core material used in sandwich structures, and the second honeycomb core 6 is a full-height honeycomb core. Both the first honeycomb core 35 and the second honeycomb core 6 are made of carbon fiber composite material.

[0062] Here, the first honeycomb core 35 is a special honeycomb core material for sandwich structures, and the second honeycomb core 6 is a full-height honeycomb core. Both are made of carbon fiber composite materials and are machined according to digital models using CNC machining equipment to ensure the accuracy of shape and dimensions. During assembly, the first honeycomb core 35 requires positioning assistance from a projector to ensure its fit with the outer skin 33 and inner skin 34. During the processing of the second honeycomb core 6, a thickness tolerance (0, +0.4mm) must be reserved to compensate for thickness shrinkage after adhesive curing. The hole shape and wall thickness design of the two honeycomb cores must meet the load-bearing requirements of the inner aileron 1 and aileron adjustment piece 2 to avoid collapse or damage under stress.

[0063] Furthermore, the first honeycomb core 35 and the second honeycomb core 6, made of carbon fiber composite material, have low density and high strength, which significantly reduces the structural weight compared to traditional metal honeycomb cores or foam cores, while improving the buckling and fatigue resistance of the sandwich structure. CNC machining ensures the dimensional accuracy of the honeycomb core, and projector-assisted positioning improves the assembly accuracy of the first honeycomb core 35, reduces internal defects, and improves the overall structural integrity. The thickness tolerance design of the second honeycomb core 6 avoids aerodynamic shape deviations caused by shrinkage after adhesive curing, ensuring the aerodynamic performance of the aileron trim piece 2, thereby improving the flight stability of the aircraft.

[0064] Example 1: A composite material aileron control surface structure; The aileron control surface includes an inner aileron 1 and an aileron trim piece 2. The inner aileron 1 has a honeycomb sandwich structure, and the aileron trim piece 2 has a full-height honeycomb structure.

[0065] The sandwich panel assembly 3 of the inner aileron 1 adopts a honeycomb sandwich structure. To improve the quality of the outer surface of the skin panel and reduce the probability of internal defects, a co-bonding method is used for molding. First, the outer skin 33, with a thickness of 1mm, is cured. Then, the first honeycomb core 35 is placed and the inner skin 34, with a thickness of 1mm, is laid on top, and finally cured into a single unit. The first honeycomb core 35 is formed by machining. The placement of the first honeycomb core 35 and the layup of the prepreg are positioned using a projector to improve manufacturing accuracy.

[0066] The first frame assembly 4 of the inner aileron 1 is positioned and installed using a tooling fixture. Two suspension joints 41 and one drive joint 42 are installed on the outer side of the web of the first front beam 43 for connecting and driving the inner aileron 1. Four suspension joints 48 are installed on the outer side of the web of the first rear beam 44 for suspending and installing the aileron trim tabs 2.

[0067] In the first frame assembly 4 of the inner aileron 1, the first front beam 43, the first rear beam 44, the first support rib 46, and the reinforcing corner box 47 are all laminated plate structures formed by a male mold; the thickness of the first front beam 43 and the first rear beam 44 is 1.6 mm, and the thickness in each joint connection area is 2 mm; the thickness of the first support rib 46 is 1 mm, and the thickness of the reinforcing corner box 47 is 1.2 mm.

[0068] The main structure of the aileron trim piece 2 is a full-height honeycomb structure, consisting of a composite material upper skin 7, a lower skin 8, a second skeleton assembly 5, a second honeycomb core 6, and a connector. The upper skin 7 and lower skin 8 are 0.8 mm thick and are formed by a negative mold.

[0069] The second honeycomb core 6 is made of thick plate according to the digital model. When the second honeycomb core 6 is glued and cured, the thickness direction will shrink. The actual thickness of the second honeycomb core 6 should be greater than the design value. The thickness tolerance is set as (0, +0.4mm).

[0070] The second frame assembly 5 consists of parts such as the second front beam 51, the second rear beam 52, and the second support rib 53. These parts are positioned on the tooling and pre-riveted together. Then, all parts are positioned and pre-assembled on the adhesive bonding tooling, and finally assembled into the second frame assembly 5. Four connecting joints 9 are installed on the outer side of the web of the second front beam 51 for connecting the aileron trim tab 2 to the inner aileron 1.

[0071] The second skeleton component 5 and the second honeycomb core 6 are positioned on the bonding fixture, and then bonded together with the upper skin 7 and the lower skin 8 to form a complete structure of the aileron adjustment piece 2, which is finally connected to the inner aileron 1 to form a whole.

[0072] In this utility model, the basic installation and load-bearing foundation are as follows: The aileron control surface is connected to the aircraft through the first frame assembly 4 of the inner aileron 1: two sets of suspension joints 41 on the first front spar 43 suspend the inner aileron 1 on the rear spar suspension bracket of the outer wing of the aircraft, forming a stable support of "two-point suspension"; at the same time, the drive joint 42 on the first front spar 43 located between the two sets of suspension joints 41 is connected to the drive rod of the aircraft control system, providing a deflection power source for the inner aileron 1.

[0073] The sandwich panel assembly 3 (upper panel 31 + lower panel 32) of the inner aileron 1 and the first frame assembly 4 are formed into a rigid whole by countersunk pins. The sandwich structure of the upper panel 31 and the lower panel 32, consisting of "outer skin 33 + inner skin 34 + first honeycomb core 35", can directly withstand aerodynamic loads (such as airflow pressure and impact force) during flight. The first front beam 43, the first rear beam 44 and the first support rib 46 of the first frame assembly 4 form a grid-like support, which distributes the load transmitted by the sandwich panel assembly 3 to the suspension joint 41 and finally smoothly transmits it to the rear beam of the outer wing of the aircraft, avoiding local stress concentration.

[0074] The aileron trim piece 2 connects to the four sets of suspension joints 48 on the first rear beam 44 of the inner aileron 1 via four sets of connecting joints 9 on the second front beam 51, forming a "four-point equidistant connection" force system, ensuring that the load of the aileron trim piece 2 can be evenly transmitted to the inner aileron 1, and work together with the inner aileron 1 to bear the load.

[0075] Core load transfer path Aerodynamic load transfer: The airflow acts on the first skin 410 of the inner aileron 1, the sandwich panel assembly 3, and the upper skin 7, lower skin 8, and second skin 55 of the aileron trim piece 2. The outer skin 33 / upper skin 7 first transfers the aerodynamic pressure to the inner first honeycomb core 35 / second honeycomb core 6. The honeycomb core evenly distributes the load to the inner skin 34 / lower skin 8 through its own hexagonal hole structure, and then transfers it to the first frame assembly 4 / second frame assembly 5.

[0076] Inertial load transfer: The inertial force generated when the aircraft turns or bumps is transferred to the first front beam 43 through the suspension joint 41 and drive joint 42 of the inner aileron 1, and then distributed by the first front beam 43 to the first rear beam 44 and the first support rib 46. At the same time, it is transferred to the second front beam 51 of the aileron trim piece 2 through the suspension joint 48-connection joint 9, ensuring that the inertial load is not concentrated on a single component.

[0077] Joint reinforcement and load buffering: All joints (suspension joint 41, drive joint 42, suspension joint 48, and connecting joint 9) are equipped with reinforcing corner boxes 47 at the connection points with the skeleton components. The reinforcing corner boxes 47, through the "L-shaped laminate wrapping" structure, convert the concentrated load transmitted by the joint into a surface load, preventing the webs of the first front beam 43, the first rear beam 44, and the second front beam 51 from cracking due to excessive local stress. At the same time, the high specific strength characteristics of carbon fiber composite materials can further buffer the load impact.

[0078] Manipulation response and attitude control logic The core function of the aileron control surfaces is to provide lateral balance and turning moment for the aircraft through deflection. The control process is divided into two levels: "inner aileron main deflection" and "aileron trim tab fine-tuning," which work together to achieve precise control. When the aircraft needs to turn or adjust its lateral attitude, the control system drives the lever to apply a push-pull force to the first front beam 43 through the drive joint 42. Since the first frame assembly 4 and the sandwich panel assembly 3 are a rigid whole, the force drives the inner aileron 1 to rotate around the connection point between the suspension joint 41 and the rear beam of the aircraft's outer wing, thereby achieving up-down deflection (the deflection angle is adjusted according to the control requirements).

[0079] The inner ailerons 1 on both sides adopt a "differential deflection" mode: one inner aileron 1 deflects upward and the other deflects downward, providing the aircraft with a lateral turning moment through the difference in aerodynamic load on the two wing surfaces.

[0080] Fine-tuning control of aileron trim tab 2: To compensate for manufacturing errors, aerodynamic load deviations, or reduce the control drag of the inner aileron 1, the actuating cylinder on the wing applies thrust or pull to the second frame assembly 5 through the actuation joint 10 of the lower skin 8 of the aileron trim piece 2. The second frame assembly 5, the second honeycomb core 6, the upper skin 7, and the lower skin 8 are bonded together as a single structure. The force causes the aileron trim piece 2 to deflect around the connection point between the connecting joint 9 and the suspension joint 48, thereby achieving fine-tuning of attitude.

[0081] The deflection direction of the aileron trim tab 2 is adapted to that of the inner aileron 1 (e.g., when the inner aileron 1 is deflected upward, the trim tab can be finely adjusted upward or downward at a small angle), ensuring that the aerodynamic shape of the entire aileron control surface is smooth, reducing airflow separation, and improving control efficiency.

[0082] Functional synergy of the core structure.

[0083] During operation, the design characteristics and working principles of each core structure are deeply integrated to ensure overall performance: The sandwich support function of the first honeycomb core 35: The first honeycomb core 35 of the inner aileron 1 sandwich panel assembly 3 is formed by machining and bonded together with the outer skin 33 and the inner skin 34. It not only utilizes the lightweight characteristics of the honeycomb structure to reduce the weight of the inner aileron 1 (reduce control inertia), but also uses the buckling resistance of the honeycomb core to prevent the panel from collapsing under aerodynamic loads, thus ensuring the structural stability of the inner aileron 1 when it deflects.

[0084] Full-height rigidity protection for the second honeycomb core 6: The second honeycomb core 6 of the aileron trim piece 2 is designed for full height and has a (0, +0.4) mm thickness tolerance to compensate for adhesive shrinkage. This ensures that after the adhesive is cured, it fits tightly with the second frame assembly 5, the upper skin 7 / lower skin 8, and prevents vibration caused by the loosening of the honeycomb core when the trim piece deflects. At the same time, the full height structure improves the overall rigidity of the trim piece and ensures the fine adjustment accuracy.

[0085] The upper skin 7 and lower skin 8 of the aileron trim 2 are formed by negative mold laying, while the outer skin 33 and inner skin 34 of the inner aileron 1 are formed by co-bonding. Both processes ensure the aerodynamic accuracy of the outer surface of the skin (surface roughness Ra≤1.6μm), reduce airflow resistance, avoid control lag caused by skin unevenness, and improve response speed.

[0086] The above description of the embodiments is only for the purpose of helping to understand the method and core idea of ​​this utility model. It should be noted that for those skilled in the art, several improvements and modifications can be made to this utility model without departing from the principle of this utility model, and these improvements and modifications also fall within the protection scope of the claims of this utility model.

Claims

1. A composite material aileron control surface structure, comprising an aileron control surface fixedly mounted on the rear suspension bracket of an aircraft wing, characterized in that: The aileron control surface consists of an inner aileron (1) and an aileron trim tab (2); The inner aileron (1) includes a sandwich panel assembly (3) and a first frame assembly (4). The sandwich panel assembly (3) includes an upper wall panel (31) and a lower wall panel (32). The first frame assembly (4) is disposed between the upper wall panel (31) and the lower wall panel (32), and the first frame assembly (4) is connected to the upper wall panel (31) and the lower wall panel (32) respectively. The upper wall panel (31) and the lower wall panel (32) are both composed of an outer skin (33), an inner skin (34) and a first honeycomb core (35) sandwiched therein. The first skeleton assembly (4) is connected to the rear beam of the outer wing of the aircraft through a suspension joint (41) and a drive joint (42). The aileron trim piece (2) includes a second skeleton assembly (5) and a second honeycomb core (6) connected within the second skeleton assembly (5). The aileron adjustment piece (2) also includes an upper skin (7) and a lower skin (8), and the second skeleton assembly (5) is disposed between the upper skin (7) and the lower skin (8), and the second skeleton assembly (5) is connected to the upper skin (7) and the lower skin (8) respectively; The second skeleton assembly (5) is connected to the inner aileron (1) via a connecting joint (9), and the lower skin (8) is equipped with an actuation joint (10) connected to the actuation cylinder on the wing of the aircraft.

2. The composite material aileron control surface structure according to claim 1, characterized in that: The first frame assembly (4) includes a first front beam (43), a first rear beam (44), a wedge-shaped tail flange (45), a first support rib (46), and a reinforcing corner box (47). The first front beam (43) and the first rear beam (44) are arranged parallel to each other. The first support rib (46) is arranged between the first front beam (43) and the first rear beam (44), and the two ends of the first support rib (46) are fixedly connected to the first front beam (43) and the first rear beam (44) respectively. The wedge-shaped tail edge strip (45) is fixedly connected to the first rear beam (44). The suspension joint (41) and the drive joint (42) are both fixedly installed on the first front beam (43). The first rear beam (44) is equidistantly equipped with suspension joints (48) for connecting the aileron adjustment piece (2). The reinforcing angle box (47) is fixedly connected to the connection of each component of the first frame assembly (4).

3. The composite material aileron control surface structure according to claim 2, characterized in that: The number of suspension joints (41) is two sets, and the drive joint (42) is disposed between the two sets of suspension joints (41).

4. The composite material aileron control surface structure according to claim 3, characterized in that: A first front edge rib (49) is fixedly connected to the front edge of the first front beam (43), and a first skin (410) is fixedly connected to the first front edge rib (49). A maintenance cover (411) is provided on the upper and lower wing surfaces of the first skin (410).

5. The composite material aileron control surface structure according to claim 4, characterized in that: The first skeleton assembly (4) includes a first front beam (43), a first rear beam (44), a first support rib (46), a reinforcing corner box (47), a suspension joint (48), a first leading edge rib (49), and a first skin (410), all of which are composite material laminate structures.

6. The composite material aileron control surface structure according to claim 1, characterized in that: The second frame assembly (5) includes a second front beam (51), a second rear beam (52), a second support rib (53), and a second front edge rib (54). The second front beam (51) and the second rear beam (52) are arranged parallel to each other. The second support rib (53) is located between the second front beam (51) and the second rear beam (52). The two ends of the second support rib (53) are fixedly connected to the second front beam (51) and the second rear beam (52) respectively. The second front edge rib (54) is fixedly connected to the front edge of the second front beam (51).

7. The composite material aileron control surface structure according to claim 6, characterized in that: A second skin (55) is fixedly connected to the second leading edge rib (54).

8. The composite material aileron control surface structure according to claim 7, characterized in that: The second front beam (51), the second rear beam (52), the second support rib (53), the second front edge rib (54), and the second skin (55) are all composite laminate structures.

9. A composite material aileron control surface structure according to claim 6, characterized in that: The number of the connecting joints (9) is four sets, and the four sets of connecting joints (9) are installed at equal intervals on the second front beam (51). The connecting joints (9) are connected to the suspension joints (48).

10. A composite material aileron control surface structure according to claim 1, characterized in that: The first honeycomb core (35) is a honeycomb core material used for sandwich structures, and the second honeycomb core (6) is a full-height honeycomb core. Both the first honeycomb core (35) and the second honeycomb core (6) are made of carbon fiber composite material.

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