JP2021017127A - Structure of movable body, and method of manufacturing the same - Google Patents

Abstract

To establish compatibility between saving in weight of a structure of a movable body and an increase in strength of a region requiring reinforcement.SOLUTION: A structure 1 of a movable body comprises a metal plate 2 that is a base material for the structure 1 of the movable body, and a plurality of fiber-reinforced composite materials 3 that serve as reinforcement materials for the metal plate 2 and that are joined like a line such as a curved line to at least one surface of the metal plate 2. The fiber-reinforced composite materials 3 are arranged in accordance with stress acting on the metal plate 2 during the movement of the movable body.SELECTED DRAWING: Figure 1

Description

The present invention relates to a moving body structure and a method for manufacturing the moving body structure.

In the automobile industry, as an alternative to steel materials, which have been the mainstream for automobile structural materials, lighter metal materials such as aluminum and titanium, and lightweight carbon fiber reinforced plastics (CFRP: Carbon Fiber Reinforced Plastic) are used. Multi-material technology that reduces the weight of the vehicle body by combining it with a composite material is drawing attention. However, CFRP has the problem of (1) lower workability than metal, and (2) lower conductivity, so there is a risk of significant damage when a large current flows through the structural material due to a lightning strike or the like. There is also the problem that there is. Further, when a brittle material such as (3) CFRP is used as the structural material, there is a problem that stress concentration is likely to occur at the joint portion between the structural materials, and a strength defect may occur.

Therefore, in structural materials such as automobiles, instead of constructing all members with CFRP, it is necessary to use metal and CFRP in the right place and aim to reduce the weight of the structural material while ensuring productivity and safety. It has been considered. For example, in Patent Document 1, in a structural material in which a stretching member made of an aluminum alloy and a CFRP member are composited, a CFRP member is provided for a part of the back surface of the stretching member in which tensile stress is generated when a compressive bending load is applied. It is disclosed to bond and reinforce.

On the other hand, in the aircraft industry, CFRP is already widely used as a structural material, and a technique of reinforcing a CFRP structural material (base material) with a metal material (reinforcing material) has also been proposed. For example, Patent Document 2 discloses a technique for reinforcing a force introduction region and a connection region of a component formed of a CRP prepreg material with a reinforcing material made of a metal foil in a large-sized aircraft component. Further, in Patent Document 3, an integrated lightning protection material sheet in which a surface finishing layer, a fiber reinforced material layer, a conductive layer made of a metal foil, an insulating adhesive layer, and a carrier paper layer are laminated is manufactured. A technique for attaching the integrated lightning protection material sheet in a strip shape on a composite skin which is a composite structural material made of CFRP is disclosed.

Japanese Unexamined Patent Publication No. 2017-119422 Special Table 2009-526670A Special Table 2011-524265

However, as described in Patent Documents 2 and 3, structural materials such as aircraft using a CFRP composite material as a base material and a metal material such as aluminum or titanium as a reinforcing material are lighter and stronger. In order to achieve this, there was no choice but to improve the properties of the CFRP composite material and the metal material itself. For this reason, simply combining a CFRP composite material (base material) and a metal material (reinforcing material) limits the weight reduction of the entire aircraft structure and the improvement of the strength of the parts that need reinforcement. There was a problem. Furthermore, when the CFRP composite material is used as the base material of the structural material, the above-mentioned material problems (1) to (3) of CFRP still remain.

Therefore, an object of the present invention is to achieve both weight reduction of a structure of a moving body such as an aircraft or a vehicle and improvement of strength of a portion requiring reinforcement.

In order to solve the above problems, according to a certain viewpoint of the present invention, with respect to at least one surface of the metal plate as a base material of the structure of the moving body and as a reinforcing material of the metal plate. , A plurality of fiber-reinforced composites joined in a linear shape including a curved shape, and the fiber-reinforced composites are arranged according to the stress acting on the metal plate when the moving body moves. A moving body structure is provided.

The fiber-reinforced composite material may be arranged in a linear shape including a curved shape along the main stress direction acting on the metal plate when the moving body moves.

The fiber-reinforced composite material may be arranged at a stress concentration portion where the stress acting on the metal plate is concentrated when the moving body moves.

The moving body is an aircraft, and the stress concentration site may include a joint portion between the fuselage and the main wing of the aircraft.

The moving body may be an aircraft, and the structure of the moving body may be an exterior material of at least one of the fuselage or main wings of the aircraft.

In order to solve the above problems, according to another aspect of the present invention, as a reinforcing material of a metal plate as a base material of a structure of a moving body, a curved shape is formed with respect to at least one surface of the metal plate. Provided is a method for manufacturing a structure of a moving body, which comprises a step of arranging a plurality of linearly joined fiber-reinforced composite materials including a plurality of fiber-reinforced composite materials according to a stress acting on the metal plate when the moving body moves. ..

In order to solve the above problems, according to another aspect of the present invention, a structural optimization simulation including at least one of topology optimization or composite laminated structure optimization is performed to obtain a base material of a moving body structure. Based on the step of designing a frame model of a fiber reinforced composite material to be joined as a reinforcing material to a metal plate, and the linear shape including a curved shape with respect to at least one surface of the metal plate based on the frame model. A method for manufacturing a structure of a moving body, including a step of manufacturing the structure of the moving body, is provided by joining a plurality of the fiber-reinforced composite materials of the above.

Based on the viewpoint of the feasibility of the structure of the moving body, the structure of the moving body further includes a step of thinning out a part of the fiber reinforced composite material from the frame model designed in the step of designing the frame model. In the step of manufacturing the body, linear fiber reinforcement including a curve is applied to at least one surface of the metal plate based on the skeleton model after the partial fiber reinforced composite is thinned out. A plurality of composite materials may be joined.

According to the present invention, it is possible to achieve both weight reduction of a structure of a moving body such as an aircraft or a vehicle and improvement of strength of a portion requiring reinforcement.

It is a schematic diagram which shows the back surface of the structure of the aircraft which concerns on 1st Embodiment of this invention. It is a perspective view which shows the aircraft manufactured using the structure which concerns on this embodiment. It is a flowchart which shows the manufacturing method of the structure of the aircraft which concerns on this embodiment. It is explanatory drawing which shows the model of the structure of the aircraft as the analysis result of the topology optimization which concerns on this embodiment. It is explanatory drawing which shows the simulation result of the structure of the main wing of an aircraft as the analysis result of the composite material laminated composition optimization which concerns on the same embodiment. It is a perspective view which shows the direction of the principal stress acting on the metal plate of the structure of the main wing of the aircraft which concerns on 2nd Embodiment of this invention. It is a perspective view which shows the structure of the main wing which consisted only of the CFRP composite material which concerns on the prior art. It is a perspective view which shows the structure of the main wing which was composed of the metal plate and CFRP composite material which concerns on this embodiment. It is a perspective view which shows the structure of the fuselage of the aircraft which concerns on 3rd Embodiment of this invention, and the joint part of a fuselage and a main wing. It is a perspective view which shows the joint part which concerns on the prior art 1. It is a perspective view which shows the joint part which concerns on the prior art 2. It is a perspective view which shows the joint part which concerns on the same embodiment.

Preferred embodiments of the present invention will be described in detail below with reference to the accompanying drawings. The dimensions, materials, other specific numerical values, etc. shown in the embodiment are merely examples for facilitating the understanding of the invention, and do not limit the present invention unless otherwise specified. In the present specification and drawings, elements having substantially the same function and configuration are designated by the same reference numerals to omit duplicate description, and elements not directly related to the present invention are not shown. To do.

[1. Structure of the aircraft]
First, the configuration of the structure of the aircraft according to the first embodiment of the present invention will be described with reference to FIGS. 1 and 2. FIG. 1 is a schematic view showing the back surface of the structure of the aircraft according to the present embodiment. FIG. 2 is a perspective view showing an aircraft manufactured by using the structure according to the present embodiment.

As shown in FIGS. 1 and 2, in the present embodiment, as an example of the structure of the moving body, an example of the structure 1 of the aircraft, which is the exterior material of the fuselage or the main wing of the aircraft, is given. As described above, the structure of the aircraft of the present invention is not limited to such an example.

As shown in FIG. 1, the aircraft structure 1 according to the present embodiment includes a metal plate 2 as a base material of the structure 1 and a CFRP composite material 3 as a reinforcing material for reinforcing the metal plate 2. Consists of. The structure 1 constitutes, for example, an exterior material such as an aircraft fuselage or a main wing, and functions as an airframe structural material serving as an outer shell of the aircraft body.

The metal plate 2 is molded according to the shape of the fuselage of an aircraft or the like. FIG. 1 shows a state in which the metal plate 2 constituting the exterior material on the upper side of the fuselage of the aircraft is viewed from the back side (inside the fuselage). The metal plate 2 may be formed by processing a single plate, or may be formed by joining a plurality of metal plates made of the same or different materials. From the viewpoint of weight reduction, the material of the metal plate 2 is preferably, for example, aluminum, titanium, magnesium, or an alloy mainly composed of these metals (aluminum alloy, titanium alloy, etc.). However, the material of the metal plate 2 may be steel, stainless steel, or other metal material.

The thickness of the metal plate 2 is, for example, 0.5 to 100 mm, preferably 2 to 10 mm. From the viewpoint of reducing the weight of the structure 1, the thickness of the metal plate 2 is preferably thin, but from the viewpoint of maintaining the strength of the structure 1, the thickness of the metal plate 2 is preferably thick. Therefore, the metal plate 2 may be appropriately changed according to the required strength for each part of the structure 1 used in the airframe of the aircraft.

The CFRP composite material 3 is an example of the fiber reinforced composite material of the present invention, and is formed of a composite material made of carbon fiber reinforced plastic (CFRP). CFRP is a material in which carbon fiber is used as a reinforcing material and the carbon fiber is hardened with a resin.

As the base material resin, for example, a thermosetting resin such as an epoxy resin, a polyamide resin, a phenol resin, a benzoxazine resin, a bismaleimide resin or an unsaturated polyester can be used. The material of the fiber reinforced composite material of the present invention is not limited to the example of CFRP, for example, glass fiber reinforced plastic (GFRP: Glass FRP) using glass fiber as the reinforcing material, and boron fiber reinforced plastic (BFRP: Boron). FRP), aramid, Kevlar, Dyneema, Zylon and other resin fibers as reinforcing materials Resin fiber reinforced composite materials (AFRP: Aramid FRP, KFRP: Kevlar FRP, DFRP: Dyneema FRP, ZFRP: Zylon FRP, etc.), carbon Various fiber reinforced plastics (FRP: Fiber Reinforced Plastic) such as silicon carbide fiber reinforced composite material (SiCFRP: Silicon Carbide FRP) obtained by fiberizing a silicon compound can also be used.

The CFRP composite material 3 is, for example, a carbon fiber sheet having a thickness of 0.1 to 100 mm, preferably 0.5 to 10 mm (a sheet-like member in which carbon fibers are impregnated with resin. Also referred to as "prepreg"). It is formed by stacking a plurality of sheets. The carbon fiber sheet has anisotropy in strength and has high strength in a specific direction (orientation direction of reinforcing fibers), but low strength in other directions. Therefore, when laminating a plurality of carbon fiber sheets, it is necessary to laminate the plurality of carbon fiber sheets so that the orientations are different from each other. However, it is difficult to laminate the carbon fiber sheets so that the CFRP composite has the optimum orientation as a whole. Further, when a CFRP composite material is used as a base material for a large structural material as in the prior art, it takes time to laminate a large number of carbon fiber sheets in order to obtain a plate thickness. Will end up.

On the other hand, in the present embodiment, the metal plate 2 having no anisotropic strength is used as the base material of the structure 1, and the CFRP composite material 3 is used as the reinforcing material to the last. As a result, when the CFRP composite material 3 is provided as the reinforcing material, there is an advantage that the number of laminated carbon fiber sheets, the laminating work time, the cost, and the like can be reduced.

In the structure 1 according to the present embodiment, the CFRP composite material 3 is not a conventional surface-shaped (sheet-shaped) member, but a metal as a linear (strip-shaped) elongated member including a curved line, a straight line, and the like. It is characterized in that it is joined to the plate 2. In particular, the present embodiment is characterized in that an elongated strip-shaped CFRP composite material 3 is joined to a wide planar metal plate 2 in a wavy curved line shape. The thickness of the strip-shaped CFRP composite material 3 varies depending on the reinforcing portion, but is, for example, 0.1 to 100 mm, preferably 0.5 to 10 mm. The width of the strip-shaped CFRP composite material 3 varies depending on the reinforcing portion, but is, for example, 5 to 1000 mm, preferably 10 to 200 mm.

As shown in FIGS. 1 and 2, a plurality of linear CFRP composite materials 3 are joined to at least one surface of the metal plate 2 (the back surface in the illustrated example, that is, the inner surface of the machine body). .. These CFRP composite materials 3 are arranged so as to intersect each other, and as a whole, they are arranged in a mesh shape such as a truss structure. In addition, in this specification, "linear" includes a linear shape of an arbitrary shape such as a curved line, a straight line, and a zigzag shape, or a combination thereof. Further, "joining a plurality of CFRP composite materials 3 linearly to a metal plate 2" means, for example, joining a plurality of strip-shaped CFRP composite materials 3 so as not to intersect each other. It includes joining so as to intersect, joining to the CFRP composite 3 in an endless annular shape such as a circle, an ellipse, and a polygon. From the viewpoint of improving the strength, it is preferable to arrange a plurality of CFRP composite materials 3 so as to intersect each other as in the truss structure.

As a method of joining the CFRP composite material 3 to the metal plate 2, for example, as shown in FIG. 1, a joining method that combines heat fusion and crimping can be used. For example, the prepreg 3A is heated by a heating device 21 such as a laser heating device near the joint between the band-shaped prepreg 3A and the metal plate 2, and the resin of the prepreg 3A is melted while the prepreg 3A is metallized by the crimping roller 22. Press against the surface of the plate 2. As a result, the prepreg 3A can be heat-sealed to the metal plate 2 to join the CFRP composite material 3 to the metal plate 2. By repeatedly laminating a plurality of prepregs 3A on the same portion of the metal plate 2, the thickness of the CFRP composite material 3 on the metal plate 2 is increased, and the strength of the CFRP composite material 3 is increased. Reinforcement strength can be increased. Further, by laminating a wide prepreg 3A on the portion to be reinforced of the metal plate 2, the width of the CFRP composite material 3 on the metal plate 2 is increased to increase the strength of the CFRP composite material 3, and the portion concerned. It is also possible to increase the reinforcing strength of.

In the structure 1 according to the present embodiment, in the direction of the principal stress acting on the metal plate 2 which is the base material of the structure 1 during the flight of the aircraft (corresponding to the “movement of the moving body” of the present invention). Along the line, the CFRP composite material 3 is joined in a curved shape. The principal stress direction means the direction of the principal stress surface in which the sum of various stresses (compressive stress, tensile stress, shear stress, etc.) acting on the structure 1 during flight is maximized. By joining the CFRP composite material 3 to the metal plate 2 along the main stress direction, the parts of the metal plate 2 that need to be structurally reinforced (hereinafter referred to as "reinforcement required parts") are emphasized. Moreover, it can be reinforced effectively and efficiently.

Further, in the structure 1 according to the present embodiment, the CFRP composite material 3 is prioritized with respect to the portion where the stress acting on the metal plate 2 is concentrated (hereinafter, referred to as “stress concentration portion”) during the flight of the aircraft. Is joined to. Here, the "stress concentration portion" is a portion of the structure 1 in which stress having a higher stress value than other portions acts in a concentrated manner, for example, a joint between the fuselage and wings of an aircraft, or a fuselage. Such as the central spine. Further, "joining intensively to the stress concentrated portion" means that the thickness of the CFRP composite material 3 at the stress concentrated portion is made thicker than that of other portions, and the width of the CFRP composite material 3 at the stress concentrated portion is other than that. It includes making it wider than the portion of the above, and intensively joining a plurality of CFRP composite materials 3 to the stress concentration portion at a high density.

During flight, the stress concentration portion of the metal plate 2 is subjected to a stress having a stress value larger than that of the other portions, so that the structure 1 is easily damaged or deformed. Therefore, it is preferable to intensively reinforce the stress concentration portion. Therefore, by preferentially joining the CFRP composite material 3 to the stress concentration portion of the metal plate 2, the stress concentration portion can be effectively and efficiently reinforced, and damage or deformation of the structure 1 can be suppressed.

As described above, in the present embodiment, the linear CFRP composite material 3 is joined to the metal plate 2 along the main stress direction, and the linear CFRP composite material is focused on the stress concentration portion of the metal plate 2. Join the material 3. As a result, the strip-shaped CFRP composite material 3 is formed in a linear locus including a curve in the required direction (direction along the principal stress direction) only for the necessary portion of the structure 1 where high strength is required. Can be joined and reinforced. That is, the metal plate 2 can be appropriately reinforced with the CFRP composite material 3 according to the magnitude and direction of the load acting on the structure 1 during flight of an aircraft or the like. Therefore, it is possible to achieve both the weight reduction of the aircraft structure 1 and the strength improvement of the portion of the structure 1 that requires reinforcement. Therefore, in a multi-material structure made of metal and CFRP, it is possible to provide a useful structure 1 that can realize weight reduction while ensuring productivity and safety.

[2. Manufacturing method of aircraft structure]
Next, a method of manufacturing the aircraft structure 1 according to the present embodiment will be described with reference to FIG. FIG. 3 is a flowchart showing a manufacturing method of the aircraft structure 1 according to the present embodiment.

As shown in FIG. 3, the manufacturing method of the aircraft structure 1 according to the present embodiment is roughly divided into a step (S10) of designing the metal plate 2 and the CFRP composite material 3 and a step of manufacturing the structure 1. (S20) and is included. The design step (S10) includes a metal plate design step (S12), a topology optimization step (S14), a composite material thinning step (S16), and a composite material laminated configuration optimization step (S18). The manufacturing step (S20) includes a molding step (S22) of the metal plate 2 and a joining (lamination) step (S24) of the composite material. Each step will be described in detail below.

(S12) Metal Plate Design Step First, the configuration of the metal plate 2 which is the base material of the aircraft structure 1 is designed. Here, an example in which the structure 1 is the exterior material of the fuselage and the main wing of the aircraft and the metal plate 2 constituting the exterior material is designed will be described. As shown in FIGS. 1 and 2, the shape, size, material, etc. of the metal plate 2 are designed according to the characteristics such as shape, strength, and function required for the fuselage and main wings of the aircraft.

Next, the configuration of the CFRP composite material 3 is designed as a reinforcing material for the metal plate 2 designed in S12. The structural optimization simulation including the topology optimization (S14) and the composite material laminated structure optimization (S18) is used to design the framework model and the laminated structure of the CFRP composite material 3 to be joined to the metal plate 2. The design of the CFRP composite material 3 preferably includes the following three steps (S14, S16, S18).

(S14) Topology optimization step In this step (S14), the optimum solution of the framework model of the CFRP composite material 3 is analyzed and designed by the topology optimization simulation, and the basic arrangement of the CFRP composite material 3 is determined.

Structural optimization used when designing the shape of a structural material includes analysis methods such as dimensional optimization, shape optimization, and topology optimization. Of these, the basic idea of topology optimization is "Under the structural constraints, loads and constraints assumed when using structure 1, the most efficient material distribution in the set design space. To find it. " Topology optimization is an analysis method that finds the optimum design plan by scraping unnecessary materials under the relevant conditions. In the execution process of topology optimization, simulation is repeated and efficient solution search is performed. ..

In the present embodiment, the optimum skeleton model of the CFRP composite material 3 in the structure 1 is designed by using the analysis tool that executes the above topology optimization. FIG. 4 is an explanatory diagram showing a model of the structure 1 of the aircraft as the analysis result of the topology optimization according to the present embodiment. As shown in FIG. 4, the optimum skeleton model 32 of the CFRP composite material 3 is derived from the optimum model 31 of the structure 1 in which the metal plate 2 and the CFRP composite material 3 are combined. In such a skeleton model 32, a plurality of CFRP composite materials 3 are arranged in a skeleton such as a truss structure. Specifically, a plurality of curved CFRP composite members 3 are arranged so as to intersect with each other along the main stress direction acting on the structure 1 of the fuselage of the aircraft. Since a stress concentration portion is generated in the structure 1 during flight of an aircraft, the CFRP composite materials 3 are densely arranged so that the stress concentration portion can be reinforced intensively.

(S16) Thinning Step of Composite Material In this step (S16), a part of CFRP composite material 3 is obtained from the framework model 32 designed in the above step (S14) based on the viewpoint of the feasibility of the structure 1 of the aircraft. Is thinned out.

Due to restrictions due to the feasibility of the aircraft structure 1, the optimized skeleton model 32 may not be directly applicable to the actual aircraft structure 1. Here, the feasibility of the structure 1 includes, for example, (1) feasibility as a vehicle, (2) feasibility as aerodynamics, (3) feasibility of manufacturing, and the like.

For example, when the CFRP composite material 3 of the skeleton model 32 exists near the entrance / exit of the aircraft or in the center of the cabin, the CFRP composite material 3 cannot be actually arranged in the actual aircraft. That is, due to the above (1) spatial restriction as a vehicle, the arrangement of the CFRP composite material 3 is restricted. Therefore, it is preferable to remove the CFRP composite material 3 near the doorway from the frame model 32. Similarly, from the viewpoint of (2) restrictions on aerodynamics during flight of the aircraft, and (3) restrictions on the manufacture of the aircraft that the CFRP composite material 3 that crosses the cabin inside the aircraft cannot be placed. In some cases, the optimized skeleton model 32 cannot be used as it is. Therefore, from the viewpoint of the feasibility of the structure 1 described above, the feasibility of the structure 1 of the actual machine is ensured by thinning out (removing) a part of the CFRP composite material 3 from the skeleton model 32 as necessary. It is preferable to do so. By thinning out the CFRP composite material 3, the skeleton model 32 is redesigned into a more realistic model.

(S18) Composite Material Laminated Structure Optimization Step Next, in this step (S18), the composite material laminated structure optimization simulation is performed based on the frame model 32 of the CFRP composite material 3 designed in the above steps (S14, S16). , The laminated structure (plate thickness, orientation direction, etc.) of the CFRP composite material 3 at each part of the structure 1 is optimized. Specifically, in the frame model 32 of the CFRP composite material 3, the thickness and width of the CFRP composite material 3, the number of laminated carbon fiber sheets (prepreg 3A) constituting the CFRP composite material 3, the orientation direction of the carbon fibers, and the like. The stacking configuration is determined. For this analysis process, an analysis tool that executes the optimization of the composite laminated structure is used. When a part of the CFRP composite material 3 is thinned out in the above step (S16), the composite material laminated composition is optimized for the frame model 32 after the thinning out.

FIG. 5 is an explanatory diagram showing a simulation result of the structure 1 of the main wing of an aircraft as an analysis result of the composite material laminated structure optimization according to the present embodiment. As shown in FIG. 5, the type and distribution of the laminated structure of the carbon fiber sheets are optimized at each part of the structure 1 of the main wing. In FIG. 5, the thickness of the CFRP composite material 3 is analyzed so as to be relatively thick in the region 41 and relatively thin in the region 42. Based on this analysis result, the laminated structure (thickness, width, orientation direction, etc.) of the CFRP composite material 3 at each part of the actual main wing structure 1 is designed.

By the above steps (S14 to S18), the frame model 32 of the CFRP composite material 3 joined to the metal plate 2 and the laminated structure are completed. After that, a plurality of linear CFRP composite members 3 are joined to the metal plate 2 based on the frame model 32 and the laminated structure to manufacture the structure 1 (S20).

(S22) Metal Plate Molding Step First, the metal plate 2 designed in the metal plate design step (S12) is actually molded. As a method for forming the metal plate 2, for example, a known processing method such as press working, bending processing, welding processing, or the like is used to form the metal plate 2 which is the base material of the structure 1.

(S24) Joining (Laminating) Step of Composite Material Next, one side of the metal plate 2 formed in the above forming step (S22) based on the skeleton model 32 and the laminating configuration designed in the above steps (S14 to S18). Alternatively, a plurality of linear CFRP composite materials 3 are joined to both sides (S24). As shown in FIG. 1, this joining method can use heat fusion using a heating device 21 and a crimping roller 22. Further, by laminating a plurality of prepregs 3A based on the laminated structure designed above, the thickness and width of the CFRP composite material 3 at each portion of the structure 1 are adjusted. It is preferable to apply an automatic laminating technique using a robot or the like to the laminating process of the prepreg 3A.

As described above, a plurality of linear CFRP composite materials 3 are joined to the metal plate 2, and the production of the aircraft structure 1 is completed. According to the method for manufacturing the aircraft structure 1 according to the present embodiment, as a reinforcing material for the metal plate 2 which is the base material of the aircraft structure 1, a plurality of the metal plate 2 is provided on at least one surface of the metal plate 2. The fiber reinforced composite material (CFRP composite material 3) is arranged according to the stress acting on the metal plate 2 when the aircraft moves. Further, in the manufacturing method according to the present embodiment, a plurality of curved linear CFRP composite materials 3 are formed on a metal plate by using a technique of laminating elongated strip-shaped prepregs 3A in a wavy curved linear shape. Join to the required part such as the stress concentration part of 2 in an appropriate direction along the main stress direction. As a result, the metal plate 2 can be reinforced by using a reinforcing material made of a plurality of CFRP composite materials 3 that intersect each other like a truss structure, and the metal plate 2 (base material) and the CFRP composite material 3 (reinforcing material) can be reinforced. ) Can be manufactured.

[3. Arrangement of composite material along the principal stress direction]
Next, with reference to FIGS. 6 to 8, the structure in which the linear CFRP composite material and the metal plate according to the second embodiment of the present invention are combined will be described in more detail. FIG. 6 is a perspective view showing the direction of the main stress acting on the metal plate 6 of the structure 5 of the main wing of the aircraft according to the second embodiment of the present invention. FIG. 7 is a perspective view showing a main wing structure 105 composed of only the CFRP composite material 103 according to the prior art. FIG. 8 is a perspective view showing a main wing structure 5 composed of the metal plate 6 and the CFRP composite material 3 according to the present embodiment.

As shown in FIG. 6, consider a case where stress acts on a metal plate 6 which is a base material of a structure 5 constituting a main wing during flight of an aircraft. In this case, the main stress direction 7 has a curved shape that extends and curves in the longitudinal direction of the main wing. Therefore, in order to improve the strength of the structure 5, it is preferable to orient the carbon fibers of the CFRP composite material along the curved principal stress direction 7.

However, as shown in FIG. 7, in the structure 105 using only the CFRP composite material 103 according to the prior art as the base material, the CFRP composite material 103 is provided with a curvature along the curved principal stress direction 7. It is difficult to orient the carbon fibers of the CFRP composite material 103 due to manufacturing restrictions.

That is, the CFRP composite material 103 is formed by laminating a plurality of carbon fiber sheets, and as shown in FIG. 7, the orientation directions of the carbon fibers are specific such as 0 °, 45 °, 90 °, and the like. Limited to directions. For example, in the conventional general CFRP composite material 103, a carbon fiber sheet oriented at 45 ° is laminated on the outermost layer in order to prevent peeling of the surface layer, and the orientation direction is the principal stress direction 7 (approximately 0 °). Direction) does not match. As described above, when the structure 105 is composed only of the CFRP composite material 103, it cannot deviate from the specific laminated structure due to the restrictions on the production of the CFRP composite material. As a result, the CFRP composite material 103 becomes a pseudo-isotropic laminated plate, and it is difficult to adjust the orientation direction of the carbon fibers along the curved principal stress direction 7 as shown in the figure. Therefore, in the structure 105 made of only the CFRP composite material 103 according to the prior art, it is difficult to achieve both weight reduction of the structure 105 and securing of strength in the direction requiring reinforcement.

On the other hand, as shown in FIG. 8, in the structure 5 according to the present embodiment, a plurality of curved CFRP composite members 3 are arranged along the curved principal stress direction 7 (generally 0 ° direction). That can be easily achieved. With the curved CFRP composite material 3, the metal plate 6 can be effectively and efficiently reinforced along the principal stress direction 7 that needs to be reinforced. As described above, the structure 5 according to the present embodiment uses the metal plate 6 having no anisotropy in strength as the base material of the structure 5, and the curved CFRP composite material 3 is used as a necessary portion of the metal plate 6. In addition, it is configured to be laminated in the required direction. As a result, the structure 5 can be efficiently designed and manufactured by particularly increasing the strength in the direction substantially matching the principal stress direction 7 and the strength of the stress concentration portion. Therefore, the weight of the structure 5 can be reduced because the CFRP composite material 3 is not arranged in the unnecessary part while ensuring the strength in the necessary part and the direction of the structure 5.

[4. Application example of aircraft structure]
Next, an example in which the aircraft structure according to the third embodiment of the present invention is applied to the joint portion between the fuselage and the main wing of the aircraft will be described with reference to FIGS. 9 to 12. FIG. 9 is a perspective view showing the structure 1 of the fuselage of the aircraft according to the third embodiment of the present invention and the joint portion 10 between the fuselage and the main wing.

As shown in FIG. 9, a plurality of connecting portions 10 for connecting the fuselage and the main wing are projected at both ends of the structure 1 of the fuselage of the aircraft in the left-right direction. The joint portion 10 is also generally referred to as a fastener joint portion. Since the joint portion 10 becomes a stress concentration portion during flight of an aircraft, the joint portion 10 is required to have high strength in order to stably connect the fuselage and the main wing.

Here, the coupling portion 10 according to the present embodiment (see FIG. 12) is compared with the coupling portion 10A (see FIG. 10) according to the prior art 1 and the coupling portion 10B (see FIG. 11) according to the prior art 2. .) Will be described.

As shown in FIG. 10, in the prior art 1, the structure of the fuselage of the aircraft is composed of only the CFRP composite material 120 (base material), and the metal material is used for the joint portion 10A at the end of the fuselage. In the joint portion 10A according to the prior art 1, the end portions of the CFRP composite material 120 of the fuselage are sandwiched between the upper and lower metal metal fittings 110 and 111, and the metal metal fittings 110 and 111 and the CFRP composite material 120 are formed by a plurality of fasteners 113. To fix. Through holes 112 for connecting to the main wing are formed at the tips of the metal metal fittings 110 and 111.

A brittle material such as the CFRP composite material 120 is prone to stress concentration at the joint portion 10A formed by the fastener 113 and is prone to brittle fracture. Therefore, in the joint portion 10A having the above structure, it is necessary to increase the plate thickness of the CFRP composite material 120 in the region 114. Therefore, the weight of the connecting portion 10A increases by the amount of the thickening, and the weight reduction of the connecting portion 10A is hindered.

Further, as shown in FIG. 11, in the prior art 2, the structure of the fuselage of the aircraft is composed of only the CFRP composite material 210 (base material), and the joint portion 10B at the end of the fuselage is also the CFRP composite material 210. It is a structure integrally formed with. A through hole 212 for connecting to the main wing is formed in the tip portion 210a of the CFRP composite material 210 forming the connecting portion 10B.

Even in the joint portion 10B having such a structure, the CFRP composite material 210 is a brittle material, and the peripheral portion of the through hole 212 of the tip portion 210a is easily broken. Therefore, in order to secure the surface pressure strength around the through hole 212, it is necessary to make the plate thickness of the tip portion 210a in the region 214 thicker than that of the base portion 210b. Therefore, the weight of the connecting portion 10B increases by the amount of the thickening, and the weight reduction of the connecting portion 10B is hindered.

On the other hand, as shown in FIG. 12, in the joint portion 10 according to the present embodiment, the structure of the fuselage of the aircraft is a hybrid of a metal plate 11 (base material) and a linear CFRP composite material 3 (reinforcing material). It is composed of members, and the joint portion 10 at the end of the body is also composed of the hybrid member. A through hole 12 for connecting to the main wing is formed at the tip of the metal plate 11 in the connecting portion 10. A curved CFRP composite material 3 is arranged on the metal plate 11 along the periphery of the through hole 12, and the metal plate 11 around the through hole 12 is reinforced.

With such a structure, no stress concentration portion is generated on the CFRP composite material 3 which is a reinforcing material, so that the CFRP composite material 3 is not brittlely fractured. Then, since the strength of the metal plate 11 reinforced by the CFRP composite material 3 is increased, even if stress concentration occurs around the through hole 12 of the metal plate 11, it is possible to suppress the metal plate 11 from being broken or deformed. Further, by forming the member around the through hole 12 with the metal plate 11, the surface pressure strength around the through hole 12 can be secured. Therefore, the thickness of the metal plate 11 can be reduced, and the CFRP composite material 3 may be arranged only in the required portion, and the thickness can be reduced. Therefore, the weight of the joint portion 10 can be reduced as compared with the prior arts 1 and 2.

[4. Summary]
The aircraft structures 1 and 5 and the manufacturing method thereof according to the present embodiment have been described in detail above. According to the present embodiment, the base material of the structures 1 and 5 is composed of metal plates 2, 6 and 11, and a plurality of curved linear or linear lines are formed on the surfaces of the metal plates 2, 6 and 11. The CFRP composite material 3 is laminated to function as a reinforcing material. As a result, it is possible to achieve both weight reduction of the aircraft structures 1 and 5 and improvement of the strength of the portion requiring reinforcement.

Further, by arranging the curved CFRP composite material 3 along the direction of the stress acting on the structures 1 and 5 (main stress direction), the metal plates 2, 6 and 11 are efficiently reinforced while being efficiently reinforced. Weight reduction can be achieved. Further, the metal plates 2, 6 and 11 can be efficiently made by increasing or decreasing the thickness and width of the linear CFRP composite material 3 according to the magnitude (stress value) of the stress acting on the structures 1 and 5. It is possible to reduce the weight while reinforcing the weight.

Further, since the CFRP composite material 3 is laminated on the surfaces of the metal plates 2, 6 and 11 which are the base materials instead of using the CFRP composite material 3 as the base material, the number of carbon fiber sheets laminated can be reduced. , Productivity is good. Further, by using the metal plates 2, 6 and 11 as the base material, the lightning resistance is also good.

Further, by applying a metal plate 11 made of a metal material instead of a brittle material such as CFRP to the stress concentration part such as the joint portion 10 between the fuselage and the main wing of the aircraft, the strength defect of the stress concentration part can be suppressed. , Weight reduction can be realized.

Although the preferred embodiment of the present invention has been described above with reference to the accompanying drawings, it goes without saying that the present invention is not limited to such an embodiment. It is clear that a person skilled in the art can come up with various modifications or modifications within the scope of the claims, and it is understood that these also naturally belong to the technical scope of the present invention. Will be done.

For example, in the above embodiment, an example of an aircraft has been described as a moving body, but the present invention is not limited to such an example. The structure of the moving body of the present invention includes, for example, various aircraft (passenger aircraft, freight transport aircraft, military aircraft, unmanned aircraft, experimental aircraft, research aircraft, helicopters, etc.), vehicles (passenger vehicles, buses, trucks, motorcycles, three wheels, etc.). Applicable to automobiles, agricultural work vehicles, freight transport vehicles, construction civil engineering vehicles, military vehicles, etc.), railroad vehicles (trains, trains, freight vehicles, linear motor cars, etc.), ships or spacecraft, etc. ..

The present invention can be used for structures of moving bodies such as aircraft and automobiles.

1, 5 Structure 2, 6, 11 Metal plate 3 CFRP composite material 3A prepreg 7 Principal stress direction 10 Joint 12 Through hole 32 Frame model

Claims (8)

The metal plate that is the base material of the structure of the moving body and
As the reinforcing material of the metal plate, a plurality of fiber-reinforced composite materials joined in a linear shape including a curved shape with respect to at least one surface of the metal plate.
With
The fiber-reinforced composite material is a structure of a moving body, which is arranged according to the stress acting on the metal plate when the moving body moves. The structure of the moving body according to claim 1, wherein the fiber-reinforced composite material is arranged in a linear shape including a curved shape along the main stress direction acting on the metal plate when the moving body moves. The structure of the moving body according to claim 1 or 2, wherein the fiber-reinforced composite material is arranged with respect to a stress concentration portion where stress acting on the metal plate is concentrated when the moving body moves. The moving body is an aircraft
The moving body structure according to claim 3, wherein the stress concentration portion includes a joint portion between the fuselage of the aircraft and the main wing. The moving body is an aircraft
The structure of the moving body according to any one of claims 1 to 4, wherein the structure of the moving body is an exterior material of at least one of the fuselage or the main wing of the aircraft. As a reinforcing material for a metal plate that is a base material of a structure of a moving body, a plurality of fiber-reinforced composite materials that are linearly joined to at least one surface of the metal plate, including a curved shape, are moved. A step of arranging according to the stress acting on the metal plate when the body moves,
A method for manufacturing a structure of a mobile body, including. A fiber-reinforced composite skeleton that is joined as a reinforcing material to a metal plate that is the base material of a moving body structure by a structural optimization simulation that includes at least one of topology optimization or composite laminated configuration optimization. Steps to design a model and
Based on the skeleton model, a step of manufacturing a structure of a moving body by joining a plurality of linear fiber reinforced composite materials including a curved shape to at least one surface of the metal plate.
A method for manufacturing a structure of a mobile body, including. Further including a step of thinning out a part of the fiber reinforced composite material from the skeleton model designed in the step of designing the skeleton model based on the viewpoint of the feasibility of the structure of the moving body.
In the step of manufacturing the structure of the moving body,
Based on the skeleton model after the partial fiber-reinforced composite is thinned out, a plurality of linear fiber-reinforced composites including curved lines are joined to at least one surface of the metal plate. , The method for manufacturing a moving body structure according to claim 7.

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