Manufacturing method for joint parts

Abstract

To achieve weight savings and reduced manufacturing time in coupled parts with complex shapes in aircraft structures.SOLUTION: The present invention includes: a hot-press process (S200) in which a first thermoplastic composite material including first reinforcing fibers is hot-press molded to produce a first intermediate material having a main structural region in which the first reinforcing fibers are arranged to extend along a direction corresponding to a main load direction acting on the joint component; an injection molding process (S300) to produce a second intermediate material having a complex structural region and a main structural region made of a second thermoplastic composite material by injection molding the second thermoplastic composite material containing second reinforcing fibers shorter than the first reinforcing fibers and bonding it to an end face of the first intermediate material; a fusion bonding process (S400) in which end faces of a plurality of second intermediate materials to each other to produce a joint part having a hollow structure.SELECTED DRAWING: Figure 5

Description

[Technical field]

[0001] The present invention relates to a method for manufacturing a joint component. [Background technology]

[0002] Conventionally, metallic materials such as titanium have generally been used for connecting parts that connect multiple components of an aircraft structure. In order to reduce the weight of the aircraft structure, excess parts of the connecting parts made of metallic materials have been machined. However, there is a limit to how much weight can be reduced by machining, and the machining time required results in low yields. In recent years, composite materials such as carbon fiber reinforced plastics (CFRP), which are lighter than metallic materials, have been used in aircraft structures (for example, Patent Document 1). By using composite materials, it is possible to achieve a lighter aircraft structure than when metallic materials are used. [Prior art documents] [Patent documents]

[0003] [Patent Document 1] Japanese Patent Application Publication No. 9-254172 Summary of the Invention [Problem to be solved by the invention]

[0004] However, in general, the manufacture of aircraft structures using composite materials requires a lamination process of laying up composite sheets called prepregs and a curing process of curing the prepregs using a pressurizing and heating device called an autoclave, which takes a lot of time.In addition, the joining parts of aircraft structures often have complex three-dimensional shapes, and it is difficult to form such complex three-dimensional shapes by laminating prepregs.

[0005] Therefore, an object of the present invention is to provide a method for manufacturing a joined part, which can achieve weight reduction and shortening of manufacturing time for a joined part having a complex shape of an aircraft structure. [Means for solving the problem]

[0007] In order to solve the above problems, the manufacturing method of the joined part of the present invention is a manufacturing method of a joined part for joining a plurality of aircraft members, and includes a hot-pressing process of manufacturing a first intermediate material having a main structural region in which the first reinforcing fibers are arranged to extend along a direction corresponding to a main load direction acting on the joined part by hot-pressing a first thermoplastic composite material containing first reinforcing fibers, an injection molding process of manufacturing a second intermediate material in which a complex structural region and a main structural region made of a second thermoplastic composite material are joined by injection molding a second thermoplastic composite material containing second reinforcing fibers shorter than the first reinforcing fibers and joining the second thermoplastic composite material to an end portion of the first intermediate material, and a fusion process of manufacturing a joined part having a hollow structure by fusion-bonding the end faces of a plurality of second intermediate materials to each other, The main structural region and the complex structural region are scarf-jointed, and the end surface of the main structural region of the first intermediate material includes a scarf slope. nothing .

[0008] The method may further include a lamination step of stacking a plurality of sheets made of a first thermoplastic composite material containing first reinforcing fibers, and in the hot press step, a first intermediate material having a main structural region is produced by hot press molding a stack of the plurality of sheets, and a scarf slope may be formed on an end face of the main structural region after hot press molding by providing a difference in width between the plurality of sheets to be stacked.

[0009] In order to solve the above problems, the manufacturing method of the joined part of the present invention is a manufacturing method of a joined part for joining a plurality of aircraft members, and includes a hot-pressing process of manufacturing a first intermediate material having a main structural region in which the first reinforcing fibers are arranged to extend along a direction corresponding to a main load direction acting on the joined part by hot-pressing a first thermoplastic composite material containing first reinforcing fibers, an injection molding process of manufacturing a second intermediate material in which a complex structural region and a main structural region made of a second thermoplastic composite material are joined by injection molding a second thermoplastic composite material containing second reinforcing fibers shorter than the first reinforcing fibers and joining the second thermoplastic composite material to an end portion of the first intermediate material, and a fusion process of manufacturing a joined part having a hollow structure by fusion-bonding the end faces of a plurality of second intermediate materials to each other, The complex structure region includes a curved section having a larger curvature than the main structure region. nothing .

[0010] In order to solve the above problems, the manufacturing method of the joined part of the present invention is a manufacturing method of a joined part for joining a plurality of aircraft members, and includes a hot-pressing process of manufacturing a first intermediate material having a main structural region in which the first reinforcing fibers are arranged to extend along a direction corresponding to a main load direction acting on the joined part by hot-pressing a first thermoplastic composite material containing first reinforcing fibers, an injection molding process of manufacturing a second intermediate material in which a complex structural region and a main structural region made of a second thermoplastic composite material are joined by injection molding a second thermoplastic composite material containing second reinforcing fibers shorter than the first reinforcing fibers and joining the second thermoplastic composite material to an end portion of the first intermediate material, and a fusion process of manufacturing a joined part having a hollow structure by fusion-bonding the end faces of a plurality of second intermediate materials to each other, In the main structural region, the first reinforcing fibers are arranged to extend along a direction corresponding to the main load direction, and in the complex structural region, the second reinforcing fibers are arranged in an unspecified direction. R . Effect of the Invention

[0011] According to the present invention, it is possible to achieve weight reduction and shortening of manufacturing time for joined parts having complex shapes of aircraft structures. [Brief description of the drawings]

[0012] [Figure 1] 1 is a schematic perspective view of an aircraft according to an embodiment of the present invention; [Diagram 2] FIG. 2 is a schematic perspective view showing a part of a structural member of a connecting mechanism that connects the engine and the main wing according to the embodiment. [Diagram 3] FIG. 2 is a schematic perspective view showing a structure of a mounting member according to the embodiment. [Figure 4]FIG. 2 is a schematic perspective view showing the structure of a first divided body and a second divided body according to the embodiment. [Diagram 5] 4 is a flowchart showing a method for manufacturing a joined part of an aircraft according to the embodiment. [Figure 6] FIG. 2 is a schematic cross-sectional view showing a state in which the laminated prepregs according to the embodiment are hot-press molded. [Figure 7] FIG. 4 is a schematic perspective view showing the structure of a first intermediate material after hot press molding according to the same embodiment. [Figure 8] FIG. 2 is a schematic cross-sectional view showing a state in which a thermoplastic composite material containing short fibers according to the embodiment is injection molded. [Figure 9] 4 is a schematic cross-sectional view showing the structure of a second intermediate material after injection molding according to the same embodiment. FIG. [Figure 10] 4 is a schematic cross-sectional view showing the structure of a scarf joint portion of a second intermediate material according to the same embodiment. FIG. [Figure 11] 5 is a schematic cross-sectional view showing a state in which end surfaces of a pair of second intermediate materials according to the embodiment are fused together. FIG. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0013] A preferred embodiment of the present invention will be described in detail below with reference to the accompanying drawings. The dimensions, materials, and other specific values ​​shown in the embodiment are merely examples for facilitating understanding of the invention, and do not limit the present invention unless otherwise specified. In this specification and drawings, elements having substantially the same functions and configurations are given the same reference numerals to avoid repeated explanations, and elements not directly related to the present invention are not shown.

[0014] [1. Overall configuration of the aircraft] First, the overall configuration of an aircraft 1 according to an embodiment of the present invention will be described with reference to Fig. 1. Fig. 1 is a schematic perspective view of an aircraft 1 according to an embodiment of the present invention.

[0015] 1, the aircraft 1 includes a fuselage 3, a main wing 5, a horizontal tail 7, and a vertical tail 9. Hereinafter, the main wing 5, the horizontal tail 7, and the vertical tail 9 will also be referred to simply as wings.

[0016] The fuselage 3 is the central structural member of the body of the aircraft 1, and its length in the front-to-rear direction (roll axis direction) is longer than its lengths in the left-to-right direction (pitch axis direction) and up-down direction (yaw axis direction). Inside the fuselage 3, a boarding space is formed in which passengers can board, and various devices such as engines and other driving sources, fuel tanks, driving devices, and measuring instruments are mounted.

[0017] A pair of main wings 5, 5 are provided on both the left and right sides of the center of the fuselage 3. The pair of main wings 5, 5 are arranged to protrude in the left and right directions from the center of the fuselage 3. The main wings 5 ​​generate upward lift for the aircraft 1.

[0018] A pair of horizontal stabilizers 7, 7 are provided on the left and right sides of the rear of the fuselage 3. The pair of horizontal stabilizers 7, 7 are arranged to protrude in the left and right directions from the rear of the fuselage 3. The horizontal stabilizers 7 have the function of maintaining stability of the aircraft 1 around the pitch axis.

[0019] A vertical tail 9 is provided above the rear of the fuselage 3. The vertical tail 9 is disposed so as to protrude upward from the rear of the fuselage 3. The vertical tail 9 has the function of maintaining the stability of the aircraft 1 about the yaw axis.

[0020] A plurality of (here, four) engines 100 are attached to the main wing 5. The plurality of engines 100 are attached to the main wing 5 via connecting mechanisms 200. The engines 100 generate force to propel the aircraft 1.

[0021] The connected parts of the aircraft 1 according to this embodiment are applied to, for example, a connection part between the engine 100 and the main wing 5, and therefore the connected parts between the engine 100 and the main wing 5 will be described in detail below. However, the connected parts of the aircraft 1 of the present invention may be applied to any connection part that connects a plurality of members that constitute the aircraft 1, and may be applied to, for example, a connection part between the fuselage 3 and the wing (the main wing 5, the horizontal stabilizer 7, the vertical stabilizer 9) other than the connection part between the engine 100 and the main wing 5.

[0022] [2. Internal structure of the connecting mechanism] 2 is a schematic perspective view showing a part of a structural member of a connection mechanism 200 that connects the engine 100 and the main wing 5, as an example of a connecting part of the aircraft 1 according to this embodiment. As shown in Fig. 2, the main wing 5 includes an inner structure 300. The inner structure 300 has a plurality of spars 310 and a plurality of ribs 320.

[0023] The spar 310 is disposed so as to extend in the wingspan direction from the root of the fuselage 3 and the main wing 5 toward the wing tip. The rib 320 has an airfoil shape. The rib 320 is disposed so as to extend in a direction intersecting (approximately perpendicular to) the spar 310.

[0024] The coupling mechanism 200 includes an engine pylon 210, a mount member (connecting part) 220, and a beam member 230. The engine pylon 210 is disposed on an upper portion of the engine 100, and supports the engine 100 in a suspending manner.

[0025] The mount member 220 is attached to the engine pylon 210, and is directly or indirectly connected to the internal structure 300 (spars 310, ribs 320) of the main wing 5. The mount member 220 is configured as a connecting part that connects the engine pylon 210 and the internal structure 300.

[0026] The mount member 220 is made of a composite material containing reinforcing fibers and resin, for example, carbon fiber reinforced plastics (CFRP). Examples of the resin that serves as the base material include thermosetting resins such as epoxy resin, polyamide resin, phenol resin, benzoxazine resin, bismaleimide resin, and unsaturated polyester, and thermoplastic resins such as ABS resin, AS resin, and acrylic resin. The material of the fiber reinforced composite of the present invention is not limited to CFRP, and various fiber reinforced plastics (FRP: Fiber Reinforced Plastics) can be used, such as glass fiber reinforced plastic (GFRP: Glass FRP) using glass fiber as a reinforcement, boron fiber reinforced plastic (BFRP: Boron FRP), resin fiber reinforced composites using resin fibers such as aramid, Kevlar, Dyneema, Zylon, etc. as reinforcement (AFRP: Aramid FRP, KFRP: Kevlar FRP, DFRP: Dyneema FRP, ZFRP: Zylon FRP, etc.), silicon carbide fiber reinforced composites (SiCFRP: Silicon Carbide FRP) made by fiberizing a compound of carbon and silicon, etc. This allows the specific strength to be increased and the weight to be reduced compared to when the mounting member 220 is made of a metal material.

[0027] The beam member 230 is formed in a rod shape, and is connected to the mount member 220 (or the engine pylon 210) and the internal structure 300.

[0028] Metallic materials such as titanium have generally been used for the connecting parts that connect multiple components of such aircraft structures. In order to reduce the weight of the aircraft structure, excess parts of the connecting parts made of metallic materials have been cut. However, there is a limit to the weight reduction achieved by cutting, and the yield is poor because of the time required for processing. In this regard, by using composite materials for the connecting parts, it is possible to achieve a lighter aircraft structure than when metallic materials are used. However, in general, the manufacture of aircraft structures using composite materials requires a lamination process in which composite sheets called prepregs are laid up (laminated), and a curing process in which the prepregs are cured using a pressurized heating device called an autoclave, which takes a lot of time. In addition, the connecting parts of the aircraft structure often have a complex three-dimensional shape, and it is difficult to form such a complex three-dimensional shape by laminating prepregs.

[0029] The connecting parts of the aircraft 1 according to the present embodiment are molded by hot press molding using a thermoplastic composite material containing reinforcing fibers and a thermoplastic resin. This allows the molding time to be reduced compared to conventional autoclave molding using a thermosetting composite material containing reinforcing fibers and a thermosetting resin. In addition, in the present embodiment, the main structural region of the connecting parts that receives a load in the main load direction is molded by hot press molding a first thermoplastic composite material containing a first reinforcing fiber. Meanwhile, the complex structural region of the connecting parts having a complex three-dimensional shape is molded by injection molding a second thermoplastic composite material containing a second reinforcing fiber shorter than the first reinforcing fiber. This allows the weight of the connecting parts having a complex three-dimensional shape of the aircraft structure to be reduced and the manufacturing time to be reduced. The connecting parts of the aircraft 1 according to the present embodiment and a manufacturing method thereof will be described in detail below.

[0030] [3. Structure of mounting components] 3 is a schematic perspective view showing the structure of the mount member 220 according to this embodiment. As shown in FIG. 3, the mount member 220 includes a first connecting part 221 that connects to the internal structure 300 of the main wing 5 or the beam member 230, and a second connecting part 223 that connects to the engine pylon 210.

[0031] The mount member 220 includes a main structural region R1 (heavy load region) and a complex structural region R2. The main structural region R1 is a region that receives a load acting in the main load direction D of the mount member 220 when the engine pylon 210 that supports the engine 100 is connected to the internal structure 300 of the main wing 5. The complex structural region R2 is a region of the mount member 220 other than the main structural region R1, has a three-dimensional shape, and includes a curved portion 215 that has a larger curvature than the main structural region R1.

[0032] The mount member 220 of this embodiment includes a first divided body 220a and a second divided body 220b. The mount member 220 is configured by joining an end of the first divided body 220a and an end of the second divided body 220b.

[0033] Fig. 4 is a schematic perspective view showing the structure of the first divided body 220a and the second divided body 220b. As shown in Fig. 4, the first divided body 220a and the second divided body 220b have divided surfaces 227, 227 extending along the main load direction D, and are joined by fusing the divided surfaces 227, 227. The portions of the first divided body 220a and the second divided body 220b other than the divided surfaces 227, 227 are separated from each other to form a hollow portion 229. In this way, the mounting member 220 of this embodiment has a monocoque (hollow) structure.

[0034] [4. Manufacturing method of joining parts] Next, a method for manufacturing the connected parts of the aircraft 1 according to this embodiment will be described with reference to Fig. 5. Fig. 5 is a flowchart showing the method for manufacturing the connected parts of the aircraft 1 according to this embodiment.

[0035] (S100) Lamination process First, a plurality of sheets (prepregs) made of a first thermoplastic composite material containing unidirectional carbon fibers (long fibers, first reinforcing fibers) are laminated (step S100). Here, the first reinforcing fibers are, for example, long fibers having a fiber length longer than 10 mm. In addition, in order to suppress the displacement of the laminated prepregs (laminate) during hot press molding, the central parts of the laminated prepregs are point-fused by ultrasonic welding. The part to be point-fused is a part having a small curvature in the molded product after hot press molding. The curvature of the inner diameter side of the part having a curvature in the molded product after hot press molding is different from the curvature of the outer diameter side. Therefore, if the part of the prepreg where the curvature is formed after hot press molding is point-fused, the prepreg may be displaced during hot press molding depending on the difference in the curvature between the inner diameter side and the outer diameter side. Therefore, in this embodiment, the prepreg is point-fused in the part having a small curvature in the molded product after hot press molding.

[0036] (S200) Hot press process Fig. 6 is a schematic cross-sectional view showing a state in which the laminated prepregs according to this embodiment are hot-press molded. As shown in Fig. 6(a), a hot-press molding apparatus 400 includes an upper mold 410, a lower mold 420, and a power supply unit 430. Note that the power supply unit 430 is omitted in Figs. 6(b) to (e).

[0037] The upper mold 410 is configured to be movable in a direction approaching or moving away from the lower mold 420. Heaters (not shown) are disposed in the upper mold 410 and the lower mold 420. The power supply unit 430 is electrically connected to the heaters and heats them by passing a current through the heaters.

[0038] As shown in FIG. 6(b), the laminate B, which is a laminated prepreg, is placed between the upper mold 410 and the lower mold 420. Then, as shown in FIG. 6(c), the upper mold 410 is moved in a direction approaching the lower mold 420, so that the laminate B is pressed between the upper mold 410 and the lower mold 420. The power supply unit 430 also applies current to the heater to heat it. After the power supply unit 430 heats the heater for a predetermined time, it stops the heater. The laminate B of this embodiment is an example of a first thermoplastic composite material containing a first reinforcing fiber. However, the first thermoplastic composite material containing a first reinforcing fiber does not have to be a laminate, and may be, for example, a block body.

[0039] Next, as shown in Fig. 6(d), cooling water W is supplied to the upper die 410 and the lower die 420 to cool them. After the upper die 410 and the lower die 420 are cooled by the cooling water W for a predetermined time, the supply of the cooling water W is stopped. Then, as shown in Fig. 6(e), the upper die 410 is moved in a direction away from the lower die 420, and the first intermediate material 500 is taken out as a molded product.

[0040] Fig. 7 is a schematic perspective view showing the structure of the first intermediate material 500 after hot press molding according to this embodiment. As shown in Fig. 7, the first intermediate material 500 forms the main structural region R1 other than the complex structural region R2 of the second divided body 220b shown in Fig. 4. The first reinforcing fibers (long fibers) included in this main structural region R1 are arranged so as to extend along a direction corresponding to the main load direction D acting on the mounting member 220. Although not shown in Fig. 7, the first intermediate material 500 forming the main structural region R1 other than the complex structural region R2 of the first divided body 220a is also molded in the same manner by the hot press molding shown in Fig. 6.

[0041] (S300) Injection molding process 8 is a schematic cross-sectional view showing a state in which the thermoplastic composite material containing short fibers according to this embodiment is injection molded. As shown in FIG. 8(a), an injection molding apparatus 600 includes an upper mold 610, a lower mold 620, and a plunger 630.

[0042] The upper die 610 is configured to be movable in a direction approaching or moving away from the lower die 620. The upper die 610 is formed with a cylinder 611, into which a thermoplastic composite member M made of a second thermoplastic composite material containing short fibers (second reinforcing fibers) is introduced. Here, the second reinforcing fibers are short fibers shorter than the first reinforcing fibers, for example, having a fiber length of 10 mm or less. The cylinder 611 is configured to allow a plunger 630 to be inserted.

[0043] The lower die 620 is formed with a holding portion 621 , and the holding portion 621 holds a pair of first intermediate members 500 .

[0044] 8(b), with the pair of first intermediate members 500 held by the holding portion 621, the upper die 610 is moved in a direction approaching the lower die 620, and the die is clamped. Then, the molten thermoplastic composite material M is introduced into the cylinder 611, and the plunger 630 is inserted into the cylinder 611.

[0045] By inserting the plunger 630 into the cylinder 611, the molten thermoplastic composite material M is injected into the space formed between the upper die 610 and the lower die 620. After being held and cooled for a certain period of time, the upper die 610 is moved in a direction away from the lower die 620, and the second intermediate material 700 is taken out as a molded product. This second intermediate material 700 corresponds to the first divided body 220a and the second divided body 220b shown in FIG.

[0046] FIG. 9 is a schematic cross-sectional view showing the structure of the second intermediate material 700 after injection molding according to this embodiment. As shown in FIG. 9, the second intermediate material 700 has a main structure region R1 and a complex structure region R2. In the main structure region R1, the long fibers (first reinforcing fibers) are arranged so as to extend along a direction corresponding to the main load direction D. On the other hand, in the complex structure region R2, the short fibers (second reinforcing fibers) are arranged in an unspecified direction. The complex structure region R2 is scarf-jointed to the end of the main structure region R1 (first intermediate material 500). Therefore, in the scarf joint 710 where the main structure region R1 and the complex structure region R2 are scarf-jointed, a transition section is formed in which the proportions of the main structure region R1 and the complex structure region R2 change smoothly. By forming the transition section, the rigidity difference between the main structure region R1 and the complex structure region R2 in the scarf joint 710 can be changed smoothly, and the strength reduction of the second intermediate material 700 can be suppressed.

[0047] FIG. 10 is a schematic cross-sectional view showing the structure of the scarf joint 710 of the second intermediate material 700 according to this embodiment. As shown in FIG. 10(a), in the scarf joint 710, the end surface 720 of the main structural region R1 has a wedge shape. The center 720a of the end surface 720 in the thickness direction of the main structural region R1 protrudes toward the complex structural region R2 side with respect to both ends 720b of the end surface 720 in the thickness direction. The end surface 720 of the main structural region R1 has a scarf slope 721 that slopes from both ends 720b in the thickness direction toward the center 720a. The scarf slope 721 is formed, for example, by providing a difference in width between the prepregs (sheets) to be laminated before hot press molding.

[0048] 10(b), the end face 820 of the main structural region R1 may have a tapered shape. An inner end 820a of the end face 820 in the thickness direction of the main structural region R1 protrudes toward the complex structure region R2 side relative to an outer end 820b of the end face 820 in the thickness direction. The end face 820 of the main structural region R1 has a scarf slope 821 that slopes from the outer end 820b to the inner end 820a in the thickness direction.

[0049] 10(c), both end portions 920a of the end surface 920 in the thickness direction of the main structure region R1 may protrude toward the complex structure region R2 side relative to a central portion 920b of the end surface 920 in the thickness direction. The end surface 920 of the main structure region R1 has a scarf slope 921 that slopes from both end portions 920a to the central portion 920b in the thickness direction.

[0050] (S400) Fusion process Fig. 11 is a schematic cross-sectional view showing a state in which end faces 730 of a pair of second intermediate members 700 according to this embodiment are fused together. As shown in Fig. 11(a), by fusing the end faces 730 of the pair of second intermediate members 700 together, a mounting member 220 (joined component) having a hollow structure is manufactured as shown in Fig. 11(b).

[0051] As described above, the manufacturing method of the connecting part of the aircraft 1 according to the present embodiment includes the hot pressing step (S200), the injection molding step (S300), and the fusion step (S400). By hot pressing the thermoplastic composite material containing long fibers, the molding time can be shortened compared to autoclave molding using a thermosetting composite material. In addition, by injection molding the thermoplastic composite material containing short fibers into the molded product (first intermediate material 500) after hot pressing, a complex structure having a complex three-dimensional shape can be easily molded. Furthermore, by fusion bonding the molded product (second intermediate material 700) after injection molding, the mounting member 220 having a monocoque (hollow) structure can be manufactured, and weight reduction can be achieved. As a result, weight reduction and shortening of the manufacturing time can be achieved in the mounting member 220 (connecting part) having a complex shape of the aircraft structure.

[0052] 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 think of various modifications or alterations within the scope of the claims, and it is understood that these also naturally belong to the technical scope of the present invention.

[0053] In the above embodiment, an example in which the main structural region R1 and the complex structural region R2 are scarf-jointed and a transition section is provided has been described. However, the present invention is not limited to this, and the main structural region R1 and the complex structural region R2 may not be scarf-jointed and a transition section may not be provided at the joint.

[0054] In the above embodiment, an example has been described in which the scarf slope is formed by providing a difference in width between the laminated prepregs. However, the present invention is not limited to this, and the scarf slope may be formed by, for example, cutting processing.

[0055] In the above embodiment, an example has been described in which the complex structure region R2 has a three-dimensional shape with a larger curvature than the main structure region R1. However, this is not limited to this, and the complex structure region R2 does not necessarily have to include a three-dimensional shape with a larger curvature than the main structure region R1. [Explanation of symbols]

[0056] R1 Main structure area R2 Complex structure area 1 aircraft 5. Wing 100 Engine 200 Connection mechanism 210 Engine Pylon 220 Mounting material 230 Beam member 300 Internal structure 310 columns 320 Rib 500 First intermediate material 700 Second intermediate material 710 Scarf joint 720 End face 721 Scarf Slope 820 End face 821 Scarf Slope 920 End face 921 Scarf Slope

Claims

1. A method for manufacturing a joint part for joining a plurality of components of an aircraft, comprising the steps of: a hot-pressing step of producing a first intermediate material having a main structural region in which the first reinforcing fibers are arranged in a direction corresponding to a main load direction acting on the joint part by hot-pressing a first thermoplastic composite material including first reinforcing fibers; An injection molding process for manufacturing a second intermediate material in which a complex structural region made of the second thermoplastic composite material and the main structural region are joined by injection molding a second thermoplastic composite material including second reinforcing fibers shorter than the first reinforcing fibers and joining the second thermoplastic composite material to an end of the first intermediate material; a fusing step of manufacturing the joined component having a hollow structure by fusing end surfaces of a plurality of the second intermediate materials to each other; Including, The main structural region and the complex structural region are scarf-jointed, A method for manufacturing a joined component, wherein an end surface of the main structural region of the first intermediate material includes a scarf slope.

2. The method further includes a lamination step of laminating a plurality of sheets made of the first thermoplastic composite material containing the first reinforcing fibers, In the hot pressing step, the laminate of the plurality of sheets is hot-press molded to produce the first intermediate material having the main structural region; The method for manufacturing a joined component according to claim 1 , wherein the scarf slope is formed on an end face of the main structural region after hot press molding by providing a difference in width between the plurality of sheets to be stacked.

3. A method for manufacturing a joining part for joining a plurality of components of an aircraft, comprising the steps of: a hot-pressing step of producing a first intermediate material having a main structural region in which the first reinforcing fibers are arranged in a direction corresponding to a main load direction acting on the joint part by hot-pressing a first thermoplastic composite material including first reinforcing fibers; An injection molding process for manufacturing a second intermediate material in which a complex structural region made of the second thermoplastic composite material and the main structural region are joined by injection molding a second thermoplastic composite material including second reinforcing fibers shorter than the first reinforcing fibers and joining the second thermoplastic composite material to an end of the first intermediate material; a fusing step of manufacturing the joined component having a hollow structure by fusing end surfaces of a plurality of the second intermediate materials to each other; Including, A method for manufacturing a joined component, wherein the complex structural region includes a curved portion having a greater curvature than the main structural region.

4. A method for manufacturing a joining part for joining a plurality of components of an aircraft, comprising the steps of: a hot-pressing step of producing a first intermediate material having a main structural region in which the first reinforcing fibers are arranged in a direction corresponding to a main load direction acting on the joint part by hot-pressing a first thermoplastic composite material including first reinforcing fibers; An injection molding process for manufacturing a second intermediate material in which a complex structural region made of the second thermoplastic composite material and the main structural region are joined by injection molding a second thermoplastic composite material including second reinforcing fibers shorter than the first reinforcing fibers and joining the second thermoplastic composite material to an end of the first intermediate material; a fusing step of manufacturing the joined component having a hollow structure by fusing end surfaces of a plurality of the second intermediate materials to each other; Including, In the main structural region, the first reinforcing fibers are arranged to extend along a direction corresponding to the main load direction, A method for manufacturing a bonded component, wherein in the complex structure region, the second reinforcing fibers are arranged in an unspecified direction.

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