Composite structure and its manufacturing method

Abstract

To shorten assembling time and reduce weight of a composite structure.SOLUTION: A method for manufacturing a composite structure of an aircraft provided with a skin and a reinforcement material comprises: a first layering step of forming a skin inside layer including a lock part for locking both width ends of a flange of the reinforcement material in a skin by layering an uncured composite sheet in a region adjacent to a holding part of a tool for holding the reinforcement material; a fitting step of bringing both the width ends of the flange of the reinforcement material into contact with the lock part of the skin inside layer and fitting the reinforcement material to the holding part of the tool; a second layering step of forming a skin outside layer in the skin by layering an uncured composite sheet on the outer surface of the flange of the reinforcement material and the outer surface of the skin inside layer; and a curing step of curing the skin inside layer and the skin outside layer.SELECTED DRAWING: Figure 10

Description

[Technical field]

[0001] The present invention relates to a composite structure and a method for its manufacture. [Background technology]

[0002] In recent years, aircraft structures have been required to be lighter in weight due to demands for improved fuel efficiency. To reduce weight, carbon fiber composite materials are being used and adhesive bonding instead of conventional bolt bonding is being used for the joints between components. However, in the current composite structural assembly, the reliability of the adhesive is insufficient, so in consideration of fail-safety, it is now mandatory to reinforce the components with bolt fastenings (chicken fasteners, arresting fasteners) in the same way as conventional metal components. For this reason, hundreds of thousands of bolts are fastened throughout the aircraft.

[0003] For example, Patent Document 1 discloses a composite structure that can form the skin of an aircraft wing. The composite structure of Patent Document 1 includes a panel (skin) and a stringer (reinforcement) bonded to the surface of the panel. A recess is formed in the panel, and a pad engageable with the recess is formed at the runout (a pair of longitudinal ends) of the stringer. The stringer pad engages with the recess of the panel. An adhesive is provided between the stringer and the panel. Also, the stringer pad and the recess of the panel are fastened together by a fastener. [Prior art documents] [Patent documents]

[0004] [Patent Document 1] Special Publication No. 2010-525979 Summary of the Invention [Problem to be solved by the invention]

[0005] However, in order to fasten the fasteners as described in Patent Document 1, drilling and bolting work is required for the stringer pad and the recesses of the panel, which requires a huge amount of assembly time. In addition, the weight of the composite structure increases due to the bolt fastening.

[0006] SUMMARY OF THE PRESENT EMBODIMENTS An object of the present invention is to provide a composite material structure and a manufacturing method thereof that can achieve a reduction in assembly time and weight. [Means for solving the problem]

[0007] In order to solve the above problems, the manufacturing method of the composite structure of the present invention is a manufacturing method of a composite structure of an aircraft having a skin and a reinforcement, and includes a first lamination step of forming an inner skin layer of the skin including locking portions for locking both ends of the flange of the reinforcement in the width direction of the skin by laminating an uncured composite sheet in an area of ​​a jig adjacent to a holding portion that holds the reinforcement, a mounting step of mounting the reinforcement to the holding portion of the jig so that both ends of the flange of the reinforcement abut against the locking portions of the inner skin layer, a second lamination step of forming an outer skin layer of the skin by laminating an uncured composite sheet on the outer surface of the flange of the reinforcement and the outer surface of the inner skin layer, and a curing step of curing the inner skin layer and the outer skin layer. The skin has a recessed portion at a portion where the skin inner layer is not formed, and the inner surface of the flange is located outside the recessed portion of the skin and is not covered by the skin inner layer. .

[0008] In order to solve the above problems, the composite structure of the present invention is a composite structure for an aircraft composed of a skin and a stiffener, and the skin includes a skin inner layer including fastening portions that abut and fasten to both ends of the flange of the stiffener in the width direction, and a skin outer layer that is laminated on the outer surface of the flange of the stiffener and the outer surface of the skin inner layer and is bonded to the outer surface of the flange of the stiffener. The skin has a recessed portion at a location where the skin inner layer is not formed, and the inner surface of the flange is located outside the recessed portion of the skin and is not covered by the skin inner layer. .

[0009] vinegar With the flange of the reinforcing material embedded in the recess of the kin, both widthwise ends of the flange may be locked by the locking portions of the skin inner layer.

[0010] An end face in the width direction of the flange may have a first tapered surface, and the engaging portion of the skin inner layer may have a second tapered surface that abuts against the first tapered surface.

[0011] The skin may further include an adhesive layer that bonds the outer layer of the skin to the flange of the reinforcement, and the rigidity of the adhesive structure that bonds the skin and the reinforcement by the adhesive layer may be higher than the rigidity of the mechanical bonding structure that connects the skin and the reinforcement by the locking portion. Effect of the Invention

[0012] According to the present invention, it is possible to achieve a reduction in assembly time and weight of a composite structure. [Brief description of the drawings]

[0013] [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 portion of a structural member of the fuselage according to the embodiment. [Diagram 3] 4 is a schematic cross-sectional view showing a joining structure between a skin and a stringer according to the embodiment. FIG. [Figure 4] FIG. 4 is a load-displacement diagram of the composite material structure according to the embodiment. [Diagram 5] FIG. 4 is a cross-sectional view showing a first lamination step for forming an inner skin layer according to the embodiment. [Figure 6] 11 is a cross-sectional view showing a mounting step of mounting a stringer to a jig according to the embodiment. FIG. [Figure 7] 10 is a cross-sectional view showing a bonding process for attaching an adhesive layer to an outer surface of a flange of the stringer according to the embodiment. FIG. [Figure 8] FIG. 4 is a cross-sectional view showing a second lamination step for forming an outer skin layer according to the embodiment. [Figure 9] 4 is a cross-sectional view showing a curing process of the skin after lamination of the outer skin layer according to the embodiment. FIG. [Figure 10] 4 is a flowchart showing a method for manufacturing a composite material structure according to the embodiment. [Figure 11]FIG. 11 is a schematic cross-sectional view showing a composite material structure according to a modified example. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0014] 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.

[0015] [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.

[0016] 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.

[0017] 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.

[0018] 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.

[0019] 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.

[0020] 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.

[0021] The composite material structure of the aircraft 1 according to this embodiment is preferably applied to, for example, structural members of the fuselage 3, and therefore an example of a structural member of the fuselage 3 will be described in detail below. However, the composite material structure of the aircraft of the present invention may also be applied to wings (for example, the main wing 5) and the like in addition to the fuselage 3.

[0022] [2. Internal structure of the fuselage] Fig. 2 is a schematic perspective view showing some of the structural members of the fuselage 3 as an example of a composite material structure of the aircraft 1 according to this embodiment. As shown in Fig. 2, the fuselage 3 includes a skin 11 (outer plate), a plurality of stringers 13 (reinforcements), a plurality of frames 15, and a plurality of floor beams 17. Here, the stringers 13 are an example of reinforcements that make up the composite material structure of the fuselage 3.

[0023] The skin 11 has a roughly cylindrical shape. The skin 11 forms an outer shell of the fuselage 3. The outer surface of the skin 11 is exposed to the exterior space of the fuselage 3, and the inner surface forms the interior space of the fuselage 3.

[0024] The multiple stringers 13 are attached to the inner surface of the skin 11, and are arranged to extend in the fore-and-aft direction (roll axis direction) of the fuselage of the aircraft 1. The multiple stringers 13 are arranged at intervals in the circumferential direction of the inner circumferential surface of the skin 11. The multiple stringers 13 may be arranged at equal intervals or at unequal intervals in the circumferential direction of the inner circumferential surface of the skin 11.

[0025] The multiple stringers 13 are reinforcing materials that reinforce the skin 11 of the fuselage 3, and are an example of a reinforcing material that reinforces a skin of the present invention. In this embodiment, a structural example in which multiple stringers 13 (reinforcing materials) are attached to the inner surface of the skin 11 of the fuselage 3 will be described as a composite material structure of the aircraft 1, but the composite material structure of the aircraft of the present invention is not limited to such an example. For example, the skin of the present invention may be the outer skin of the wings (main wing 5, horizontal tail 7, vertical tail 9) of the aircraft 1, or the skins (outer plate, inner plate, etc.) of other parts, other than the skin 11 of the fuselage 3. Furthermore, the reinforcing material of the present invention may be a reinforcing material that reinforces the wings (main wing 5, horizontal tail 7, vertical tail 9) of the aircraft 1, or the skins of other parts, other than the stringers 13 of the fuselage 3.

[0026] The multiple frames 15 are attached to the inner surface of the skin 11, and are arranged so as to extend in a direction intersecting (orthogonal to) the multiple stringers 13. In this embodiment, the multiple frames 15 are arranged so as to extend in the circumferential direction of the inner peripheral surface of the skin 11. The frames 15 are an example of a reinforcing material that reinforces the skin 11 of the fuselage 3.

[0027] The multiple floor beams 17 are attached to the multiple frames 15 one by one. The multiple floor beams 17 are disposed at the bottom in the up-down direction of the fuselage 3. The multiple floor beams 17 form part of the floor of the boarding space where passengers can board. The multiple floor beams 17 have the role of bearing the load applied to the floor.

[0028] Of the fuselage 3, at least the skin 11 and stringers 13 are made of a composite material, for example, carbon fiber reinforced plastics (CFRP). However, the skin 11 and stringers 13 may be made of other fiber reinforced plastics, such as glass-fiber-reinforced plastics (GFRP) or aramid-fiber-reinforced plastics (AFRP). This allows for a higher specific strength and a lighter weight than when the skin 11 and stringers 13 are made of a metal material.

[0029] The composite structure of the aircraft 1 according to this embodiment means a structure of the aircraft that is made of composite materials such as fiber-reinforced plastics. In recent years, there has been a demand for weight reduction in the composite structure of the aircraft 1 due to the demand for improved fuel efficiency. As a solution to this, composite materials such as CFRP are being used as structural materials such as skins and reinforcing materials, and adhesive bonding using an adhesive is being used for the joints between the structural materials, instead of conventional bolt bonding.

[0030] In the composite material structure of the aircraft 1 according to this embodiment, in addition to adhesive bonding using an adhesive, a mechanical bond is also used by devising a molding method for the joints between structural materials. This makes it possible to reduce the weight of the composite material structure, reduce costs by reducing the number of parts, fasteners, jigs, etc., and reduce assembly man-hours while maintaining fail-safety, unlike the adhesive structure that used arresting fasteners. The composite material structure of the aircraft 1 according to this embodiment and its manufacturing method will be described in detail below.

[0031] [3. Skin and stringer joint structure] 3 is a schematic cross-sectional view showing a joining structure between the skin 11 and the stringer 13 according to this embodiment. As shown in FIG. 3, the skin 11 includes a skin inner layer 19 and a skin outer layer 21. The skin inner layer 19 is located on the inner surface side of the fuselage 3 of the skin 11, and the skin outer layer 21 is located on the outer surface side of the fuselage 3 of the skin 11. The inner surface 21b of the skin outer layer 21 is joined to the outer surface 19a of the skin inner layer 19. In this way, the skin inner layer 19 and the skin outer layer 21 are integrated to form the skin 11. In this way, the skin 11 has a two-layer laminate structure consisting of the skin inner layer 19 and the skin outer layer 21.

[0032] The skin outer layer 21 is formed over the entire outer surface of the skin 11. On the other hand, the skin inner layer 19 is not formed over the entire inner surface of the skin 11, and there are regions on the inner surface of the skin 11 where the skin inner layer 19 is not formed. Thus, the skin 11 has recesses 23 on the inner surface side of the fuselage 3 at locations where the skin inner layer 19 is not formed, which form spaces for embedding the flanges 25 of the stringer 13. The flanges 25 of the stringer 13 are disposed so as to be embedded in the recesses 23 of the skin 11, and are engageable with the recesses 23.

[0033] As shown in Fig. 3, the stringer 13 is a structural material having, for example, a T-shaped cross-sectional shape. The stringer 13 has a flange 25 and a web 27. The flange 25 is a long, flat, band-like member, and is joined perpendicularly to the web 27. As shown in Fig. 3, at least a portion of the flange 25 is disposed within the recess 23 of the skin 11. The flange 25 includes an outer surface 25a, an inner surface 25b, and a pair of end portions 25c in the width direction. The width direction of the flange 25 is the left-right direction in Fig. 3 (the direction indicated by the double arrow R).

[0034] The web 27 is a long member in the shape of a flat belt. In this embodiment, the stringer 13 has a T-shaped cross-sectional shape and is a structural member in which the flange 25 and the web 27 are joined in a T-shape. However, the cross-sectional shape of the stringer 13 is not limited to this example, and may have other shapes such as an H-shape, an I-shape, an L-shape, or a U-shape. Also, in this embodiment, the stringer 13 is formed by joining a plurality of composite parts. However, without being limited to this, the stringer 13 may be formed from a single composite part. The composite structure of this embodiment is formed from the skin 11, the stringer 13, and an adhesive layer 29.

[0035] The outer surface 25a of the flange 25 is located in the recess 23 of the skin 11. In the recess 23, an adhesive layer 29 is provided between the outer surface 25a of the flange 25 and the inner surface 21b of the skin outer layer 21. The adhesive layer 29 bonds the outer surface 25a of the flange 25 to the inner surface 21b of the skin outer layer 21. In other words, the adhesive layer 29 is interposed between the flange 25 of the stringer 13 and the skin outer layer 21 of the skin 11, and bonds the stringer 13 to the skin 11. However, the adhesive layer 29 does not need to be provided between the outer surface 25a of the flange 25 and the inner surface 21b of the skin outer layer 21. For example, the cured outer surface 25a of the flange 25 and the inner surface 21b of the uncured skin outer layer 21 may be directly bonded by co-bonding or the like.

[0036] The inner surface 25b of the flange 25 is the surface opposite to the outer surface 25a, and is located outside the recessed portion 23 of the skin 11. However, the inner surface 25b of the flange 25 may be located inside the recessed portion 23 of the skin 11. A web 27 is joined to the center position in the width direction of the inner surface 25b of the flange 25. The web 27 is erected in a direction perpendicular to the inner surface 25b of the flange 25.

[0037] A part of the end 25c in the width direction of the flange 25 is disposed within the recessed portion 23 of the skin 11, and the remaining part of the end 25c is disposed outside the recessed portion 23. An end face 25d of the end 25c of the flange 25 is a tapered surface. In this embodiment, the width (left-right direction R in FIG. 3) of the flange 25 varies depending on the position in the thickness direction (up-down direction in FIG. 3) of the flange 25, and the width of the flange 25 gradually decreases from the outer surface 25a to the inner surface 25b. Therefore, the end face 25d of the flange 25 is a tapered surface (first tapered surface) that is inclined from the outer surface 25a to the inner surface 25b so as to approach the web 27.

[0038] Here, the skin inner layer 19 has a locking portion 19c at a position facing the recessed portion 23. The locking portion 19c of the skin inner layer 19 abuts against an end portion 25c of the flange 25 located within the recessed portion 23. The locking portion 19c locks the end portion 25c of the flange 25 so that the flange 25 does not come out of the recessed portion 23.

[0039] An end face 19d of the locking portion 19c in the left-right direction R in Fig. 3 is a tapered surface. In this embodiment, the end face 19d of the locking portion 19c is a tapered surface (a second tapered surface) that is inclined from the outer surface 19a to the inner surface 19b of the skin inner layer 19 so as to approach the web 27.

[0040] The end faces 19d (second tapered surfaces) of the pair of locking portions 19c of the skin inner layer 19 are in contact with at least a part of the pair of end faces 25d (first tapered surfaces) of the flange 25. At this time, the end faces 19d of the locking portions 19c are arranged to cover the end faces 25d of the ends 25c of the flange 25, so that the pair of locking portions 19c lock both ends 25c of the flange 25. This restricts the flange 25 from moving in a direction away from the skin 11 (adhesive layer 29) (the up-down direction in FIG. 3). In this way, the flange 25 of the stringer 13 engages with the recess 23 of the skin 11, and both ends 25c of the flange 25 are in contact with and locked to the locking portions 19c of the skin inner layer 19. As a result, the flange 25 of the stringer 13 is mechanically coupled to the skin 11.

[0041] According to this embodiment, even if a load in a direction away from the skin 11 is applied to the stringer 13 and the adhesive structure between the flange 25 of the stringer 13 and the skin 11 by the adhesive layer 29 peels off, the flange 25 of the stringer 13 is mechanically coupled to the skin 11 by the locking portion 19c of the skin 11. In this manner, in this embodiment, a mechanical coupling structure (locking portion 19c, recessed portion 23, etc.) that functions as a fail-safe structure when the adhesive structure between the skin 11 and the stringer 13 peels off is provided. This makes it possible to eliminate the need for a bolt fastening structure (chicken fastener, arresting fastener, etc.) between structural members that has been used as a fail-safe structure in the past. As a result, drilling and bolt fastening work for the bolt fastening structure are no longer necessary, and therefore the assembly time can be significantly reduced. In addition, since a large number of bolts are no longer necessary, the composite structure of the aircraft 1 can be made lighter.

[0042] Next, the strength of the joint structure between the skin 11 and the stringer 13 according to this embodiment will be described with reference to Fig. 4. Fig. 4 is a load-displacement diagram of the composite material structure according to this embodiment.

[0043] In Fig. 4, the vertical axis represents the load applied in the direction in which the skin 11 and stringer 13 move apart, and the horizontal axis represents the displacement of the stringer 13 relative to the skin 11. In addition, in Fig. 4, the dashed line represents the load-displacement curve when the skin 11 and stringer 13 are joined only by an adhesive structure. The dashed-dotted line represents the load-displacement curve when the skin 11 and stringer 13 are joined only by a mechanically joined structure using the locking portions 19c. The solid line represents the load-displacement curve when the skin 11 and stringer 13 are joined both by an adhesive structure and a mechanically joined structure.

[0044] As shown by the dashed line in Fig. 4, if the skin 11 and the stringer 13 are joined only by an adhesive structure, when a load F1 is applied in a direction in which the skin 11 and the stringer 13 move apart, and a displacement X1 occurs, the adhesive structure formed by the adhesive layer 29 is destroyed. Also, as shown by the dashed line in Fig. 4, if the skin 11 and the stringer 13 are joined only by a mechanically joined structure formed by the locking portions 19c, when a load F2 is applied in a direction in which the skin 11 and the stringer 13 move apart, and a displacement X2 occurs, the mechanically joined structure formed by the locking portions 19c is destroyed. Here, the load F1 is greater than the load F2, and the displacement X1 is smaller than the displacement X2.

[0045] In contrast, according to the composite material structure of the present embodiment, as shown by the solid line in FIG. 4, when a load F1 is applied in a direction in which the skin 11 and the stringer 13 move apart, resulting in a displacement X1, the adhesive structure formed by the adhesive layer 29 is destroyed. However, the composite material structure of the present embodiment includes a mechanically bonded structure formed by the locking portion 19c in addition to the adhesive structure. Therefore, even after the adhesive structure is destroyed, the load is shared by the locking portion 19c of the mechanically bonded structure, so that the progression of cracks at the joint between the skin 11 and the stringer 13 can be suppressed, and the skin 11 and the stringer 13 can be suppressed from immediately breaking. For example, as shown in FIG. 4, even after the adhesive structure is destroyed by the load F1, the flange 25 of the stringer 13 does not come off the recessed portion 23 of the skin 11, and the joint between the skin 11 and the stringer 13 is maintained until the load acting on the joint between the skin 11 and the stringer 13 increases to a load F2 and the displacement increases to a displacement X2.

[0046] In the composite structure of this embodiment, a difference (rigidity difference) is provided between the rigidity of the adhesive structure by the adhesive layer 29 and the rigidity of the mechanically bonded structure by the locking portion 19c. Here, the rigidity of the adhesive structure is higher than the rigidity of the mechanically bonded structure. Therefore, until the adhesive structure is destroyed, the adhesive structure can transmit the load between the skin 11 and the stringer 13 well. On the other hand, after the destruction of the adhesive structure starts, until the adhesive structure is completely broken, the mechanically bonded structure by the locking portion 19c shares part of the load, and both the adhesive structure and the mechanically bonded structure can bear the load. Therefore, the progress of cracks from the destruction start point of the adhesive structure can be suppressed, and immediate breakage can be prevented.

[0047] Here, if the rigidity of the adhesive structure is lower than the rigidity of the mechanically bonded structure, the adhesive structure will also be destroyed immediately when the mechanically bonded structure is destroyed. That is, cracks will immediately grow in the adhesive structure, and the skin 11 and the stringer 13 will immediately break. For this reason, in this embodiment, the rigidity of the adhesive structure formed by the adhesive layer 29 is set higher than the rigidity of the mechanically bonded structure formed by the locking portion 19c. This makes it possible to provide a composite material structure of the aircraft 1 that can suppress the growth of cracks in the joints of the structural members. Therefore, even if there is a defect in the adhesive structure of the structural members and peeling or the like occurs, the mechanically bonded structure becomes a fail-safe structure, and it is possible to prevent the composite material structure of the aircraft 1 from being destroyed early.

[0048] [4. Manufacturing methods for composite structures] Next, a method for manufacturing the composite material structure of the aircraft 1 according to this embodiment will be described with reference to Fig. 5 to Fig. 10. Here, Fig. 10 is a flowchart showing the method for manufacturing the composite material structure according to this embodiment.

[0049] (1) First lamination process Fig. 5 is a cross-sectional view showing a first lamination step for forming the inner skin layer 19 in the manufacturing method of the composite structure according to this embodiment. As shown in Fig. 5, the inner skin layer 19 is formed by laminating an uncured composite sheet S (prepreg) on ​​the surface of a jig 50 by a laminator 60.

[0050] The laminator 60 laminates the composite sheet S on an area 50b (non-holding area) of the surface of the jig 50, which is different from the holding area 50a for holding the stringer 13, to form the inner skin layer 19 (first lamination step: step S100 in FIG. 10). The holding area 50a is formed at a position recessed inward of the jig 50 with respect to the non-holding area 50b. The area 50b is an area of ​​the surface of the jig 50 adjacent to the holding area 50a.

[0051] Here, the laminator 60 laminates composite sheets S of different lengths in the circumferential direction R of the fuselage 3 for each layer of the skin inner layer 19. The laminator 60 laminates composite sheets S of longer lengths closer to the surface (non-holding region 50b) of the jig 50. As a result, the above-mentioned locking portion 19c and end face 19d (tapered surface) are formed in the skin inner layer 19.

[0052] (2) Installation process Fig. 6 is a cross-sectional view showing the mounting step of mounting the stringer 13 to the jig 50. As shown in Fig. 6, after the skin inner layer 19 is formed, the stringer 13 is mounted to the holding portion 50a recessed inward of the jig 50 relative to the non-holding region 50b (mounting step: step S200 in Fig. 10). Here, for example, a cured stringer 13 is mounted to the holding portion 50a. However, without being limited thereto, an uncured stringer 13 may also be mounted to the holding portion 50a.

[0053] The stringer 13 is held by a holder (not shown) while being positioned within the holder 50a. At this time, both end portions 25c (end faces 25d) of the flange 25 are arranged to overlap the locking portions 19c (end faces 19d) of the skin inner layer 19 in the thickness direction of the flange 25, and the end faces 25d of the flange 25 abut against the end faces 19d of the locking portions 19c (see FIG. 7). Part of the end faces 25d of the flange 25 overlaps with the end faces 19d of the skin inner layer 19 in the thickness direction of the flange 25.

[0054] (3) Application process Fig. 7 is a cross-sectional view showing an attachment step of attaching an adhesive layer 29 to the outer surface 25a of the flange 25 of the stringer 13. As shown in Fig. 7, the adhesive layer 29 in this embodiment is made of an adhesive film and is attached to the outer surface 25a of the flange 25 (attachment step in Fig. 10: step S300). However, the present invention is not limited to this, and the adhesive layer 29 may be made of a liquid adhesive and may be applied to the outer surface 25a of the flange 25. When the adhesive layer 29 is attached to the outer surface 25a of the flange 25, the outer surface 19a of the skin inner layer 19 and the outer surface 29a of the adhesive layer 29 are approximately flush with each other.

[0055] (4)Second lamination process Fig. 8 is a cross-sectional view showing a second lamination step for forming the skin outer layer 21. As shown in Fig. 8, the skin outer layer 21 is formed by laminating an uncured composite sheet S (prepreg) on ​​the outer surface 19a of the skin inner layer 19 and the outer surface 29a of the adhesive layer 29 by a laminator 60 (second lamination step: step S400 in Fig. 10). However, without being limited thereto, when the adhesive layer 29 is not provided, the skin outer layer 21 may be formed (laminate) on the outer surface 19a of the skin inner layer 19 and the outer surface 25a of the flange 25.

[0056] Since the outer surface 19a of the skin inner layer 19 and the outer surface 29a of the adhesive layer 29 are roughly flush with each other, depressions or protrusions are unlikely to be formed on the outer surface 21a of the skin outer layer 21 (the outer surface of the fuselage 3) (see FIG. 9). The material of the composite sheet S constituting the skin outer layer 21 is the same as the material of the composite sheet S constituting the skin inner layer 19. However, without being limited thereto, the material of the composite sheet S constituting the skin outer layer 21 may be different from the material of the composite sheet S constituting the skin inner layer 19.

[0057] (5) Curing process Fig. 9 is a cross-sectional view showing the curing step of the skin 11 after lamination of the skin outer layer 21. After lamination of the skin outer layer 21, the skin 11 (skin inner layer 19, skin outer layer 21), stringer 13, and adhesive layer 29 are introduced into an autoclave (not shown) while being held by a jig 50, and heated. At this time, the skin 11 (skin inner layer 19, skin outer layer 21) heated in the autoclave is cured (curing step: step S500 in Fig. 10). By curing the skin 11, the composite structure according to this embodiment shown in Fig. 3 is formed.

[0058] 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.

[0059] In the above embodiment, an example has been described in which the end face 19d of the engaging portion 19c of the skin inner layer 19 and the end face 25d of the flange 25 are tapered surfaces. However, this is not limited thereto, and the end faces 19d, 25d may have a stepped surface instead of a tapered surface.

[0060] FIG. 11 is a schematic cross-sectional view of a composite structure according to a modified example of this embodiment. Components substantially the same as those in the composite structure of the above embodiment are denoted by the same reference numerals and will not be described. As shown in FIG. 11, the end face 125d in the width direction of the flange 25 has a step surface (first step surface). In addition, the end face 119d in the left-right direction R of the locking portion 19c has a step surface (second step surface). The end face 125d (first step surface) of the flange 25 abuts against the end face (second step surface) 119d of the locking portion 19c, and a part of the locking portion 19c is disposed so as to cover a part of the end 25c of the flange 25. According to this modified example, it is possible to achieve the effect of shortening the assembly time and reducing the weight of the composite structure, as in the above embodiment. In addition, the end faces 119d and 125d of this modified example are formed by step surfaces. Therefore, the thickness of the locking portion 19c can be made larger than in the case where the end faces 19d, 25d are configured as tapered surfaces as in the above embodiment, and the strength of the locking portion 19c can be increased. [Explanation of symbols]

[0061] S Composite Sheet 1 aircraft 3. Torso 11. Skin 13 Stringer 15 Frames 17 Floor beam 19 Skin inner layer 19a Exterior 19b Inside 19c Locking part 19d End face (second tapered face) 21 Skin Outer Layer 21a Exterior 21b Inside 23 Recess 25 Flange 25a Exterior 25b Inside 25c end 25d End face (first tapered surface) 27 Web 29 Adhesive layer 50 Jig 60 Laminating Machine

Claims

1. 1. A method for manufacturing a composite aircraft structure having a skin and stiffeners, comprising the steps of: a first lamination step of laminating an uncured composite sheet in an area of ​​the jig adjacent to a holding portion that holds the reinforcing material, thereby forming a skin inner layer of the skin, the skin including a locking portion that locks both ends of the flange of the reinforcing material in the width direction; a mounting step of mounting the reinforcing material on the holding portion of the jig such that both end portions of the flange of the reinforcing material are abutted against the engaging portions of the skin inner layer; a second lamination step of laminating an uncured composite sheet onto an outer surface of the flange of the stiffener and onto an outer surface of the inner skin layer to form an outer skin layer of the skin; a curing step of curing the inner skin layer and the outer skin layer; Including, the skin has a recessed portion at a location where the skin inner layer is not formed; A method for manufacturing a composite structure, wherein an inner surface of the flange is outside the recessed portion of the skin and is not covered by the inner skin layer.

2. A composite structure for an aircraft, comprising a skin and a stiffener, The skin is a skin inner layer including a locking portion that abuts against and locks to both ends of the flange of the reinforcing material in the width direction; a skin outer layer laminated onto an outer surface of the flange of the stiffener and an outer surface of the skin inner layer and bonded to the outer surface of the flange of the stiffener; Equipped with the skin has a recessed portion at a location where the skin inner layer is not formed; A composite structure wherein an inner surface of the flange is outside the recess in the skin and is not covered by the inner skin layer.

3. A composite structure as described in claim 2, wherein with the flange of the reinforcing material embedded in the recess of the skin, both widthwise ends of the flange are engaged by the engaging portion of the inner layer of the skin.

4. The end surface of the flange in the width direction has a first tapered surface, 4. The composite structure of claim 2 or 3, wherein the anchoring portion of the inner skin layer has a second tapered surface that abuts the first tapered surface.

5. an adhesive layer that bonds the outer skin layer of the skin and the flange of the stiffener; The composite structure according to any one of claims 2 to 4, wherein the rigidity of the adhesive structure that bonds the skin and the reinforcing material by the adhesive layer is higher than the rigidity of the mechanical bonding structure that connects the skin and the reinforcing material by the locking portion.

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