JP3271957B2 - Manufacturing method of composite material - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a method for producing a composite material comprising a fiber reinforced material impregnated with a synthetic resin.

[0002]

2. Description of the Related Art Fiber reinforced composite materials (commonly known as FRP) have advantages such as light weight, high elastic modulus and high strength, and advantages of being able to arbitrarily design rigidity and strength. Since it can be expressed, its importance as an aircraft material is increasing. FIG. 9 is a view showing a part of the fuselage panel 13 constituting the fuselage of the passenger aircraft. The torso panel 13 includes the outer panel 3.
A stringer 4 extending in the machine axis direction of the fuselage is attached to an inner wall of the body.

[0003] The body panel 13 is formed by laminating a semi-cured prepreg 10 impregnated with a synthetic resin such as an epoxy resin on a reinforcing fiber layer such as a carbon fiber layer to form an outer panel 3, and further laminating the prepreg 10. The stringer 4 is integrally attached to the outer plate 3. This is inserted into an autoclave, pressurized and heated to cure the synthetic resin, and a body panel 13 made of a composite material is manufactured. The body panel 13 is subjected to a compressive stress as shown by an arrow 6 or a tensile stress as shown by an arrow 11. When such stress acts, stress concentrates on the step 8 between the outer plate 3 and the stringer 4, and the boundary layer 14 between the outer plate 3 and the stringer 4 is formed.
Thus, the outer plate 3 and the stringer 4 may be separated from each other.

As a method for preventing this peeling, a stepped portion 8
As shown by a broken line 12 between the outer skin 3 and the stringer 4 in the vicinity, all the prepregs are moved along the step 8 in the machine direction (perpendicular to the plane of FIG. 9) in the thickness direction of the fuselage panel 13 (vertical direction in FIG. 9). Stitching in different directions).

[0005]

In the prior art, as shown in FIG. 9, the layers other than the boundary layer 14 between the outer plate 3 and the stringer 4 which are not involved in the separation are stitched to increase the peeling resistance. I was Such useless stitching damages the reinforcing fiber layer of the prepreg 10 which is not involved in peeling, and has a problem in that the strength in the machine axis direction along the step is reduced.

Further, in the prior art, the stitching is performed after the lamination of a plurality of layers of the prepreg 10 is completed. Therefore, if the composite material to be manufactured is large, such as the Stitching had to be performed using a stitching device.

An object of the present invention is to provide a method for producing a composite material which does not easily peel off even when stress is applied.

[0008]

According to a first aspect of the present invention, there is provided a composite material for use in an aircraft, wherein one part where a plurality of layers of reinforcing fibers are laminated and the other where a plurality of layers of reinforcing fibers are laminated. In the method of manufacturing a composite material that is sewn together and laminated via a step, a stress is analyzed by applying a tensile and compressive load to a model of the composite material by computer simulation, and the stress is applied when the load is applied. Determining a region near the stepped portion where the steep increase increases as a stitching region, and laminating a reinforced fiber layer included in the stitched region, or a prepreg in which the reinforcing fibers are impregnated with a synthetic resin, among the one portion. Step, of the other part, the reinforcing fiber layer or the prepreg included in the stitched area, A step of layering, a part of the part of the one part and the part of the other part which are stacked, a part included in the stitching area, and a step of stitching and joining the part along the stepped part; And a step of laminating the remaining reinforcing fiber layer or prepreg at the portion of the composite material.

[0009]

[0010]

[0011]

[0012]

According to the present invention, first, a load is applied to the model of the composite material, and a stress analysis is performed by computer simulation to determine a region where the stress sharply increases. Next, a plurality of reinforcing fiber layers or prepregs at one portion of the composite material included in the region determined as described above are laminated. Next, on one of the laminated portions, a plurality of reinforcing fiber layers or prepregs included in the region determined as described above are laminated via a step, and the other portion is formed on one portion. Next, one part and the other part are stitched along the step near the step part, and a reinforcing fiber layer or a prepreg is further laminated on the one part and the other part. When laminating a prepreg, the synthetic resin is cured at a high temperature and a high pressure, and when laminating a reinforcing fiber layer, the synthetic resin is impregnated with the resin, and then the resin is cured at a high temperature and a high pressure. Form a composite material.

As described above, only the reinforcing fiber layer of the portion to be sewn in advance is firstly sewn and then sewn, and then the remaining reinforcing fiber layer or prepreg is sewn, so that only the predetermined region can be accurately sewn. Further, the reinforcing fiber layer in a region unrelated to the peeling is not damaged.

[0015]

FIG. 1 is a view showing a part of a body panel 21 which is an embodiment of a composite material manufactured by a method for manufacturing a composite material according to the present invention. The composite material of the present invention is suitably applied to a member or the like on which a large stress acts in a normal use environment such as a fuselage panel 21 of a passenger aircraft as shown in FIG. The fuselage panel 21 constituting the fuselage of the passenger aircraft includes an outer plate 23 and a stringer 24 attached to the inner peripheral wall of the outer plate 23 and extending in the machine axis direction of the passenger aircraft (a direction perpendicular to the plane of FIG. 1). The composite material used for such a body panel 21 includes a step portion, in this embodiment, the outer plate 23 and the stringer 2.
A step portion 30 is formed at a joint portion between the step portion 4 and the step 4. When an external force acts on this composite material, stress concentrates near this step,
In a region separated from the step, little stress acts in the interlayer direction. Therefore, in the body panel 21 of the present embodiment, only the region near the step portion 30 where the stress is concentrated is included in the region near the step portion 30 where a stress equal to or more than a predetermined value acts to selectively increase the strength. Stitch only the reinforcing fiber layer. This eliminates the need to sew up to the reinforcing fiber layer in a region where a stress equal to or greater than a predetermined value does not act.
Damage to the reinforcing fiber layer can be prevented.

Next, a method of manufacturing the body panel 21 will be described with reference to FIGS. First, in a first step, a prepreg impregnated with a synthetic resin is laminated to form a part of an outer plate 23 which is one part of a body panel 21 formed in a stepped manner, and then in a second step, First
A prepreg is further laminated on a part of the outer plate 23 laminated in the step via the step portion 30 to form a part of the stringer 24 which is the other part. The reinforcing fiber layer of the prepreg laminated in the first step and the second step is a reinforcing fiber layer in a region where a stress equal to or more than the above-mentioned predetermined value acts.
Therefore, in the next third step, a part of the stringer 24 near the step portion 30 and a part of the outer plate 23 are sewn (stitched), and in a fourth step, a part of the outer plate 23 and the stringer are stitched. A prepreg is further laminated on a part of 24. FIG. 2 is a view showing a state immediately after the end of the third step.

The prepreg 27 is obtained by impregnating a reinforcing fiber layer with a synthetic resin and preliminarily polymerizing, that is, a curing reaction of the synthetic resin is advanced to some extent, the temperature is lowered, and the molding material is stopped in the middle of the reaction to form a semi-cured material. And are generally wound on a roll. The prepreg 27 includes a one-way prepreg having a reinforcing fiber layer in which fibers are aligned in one direction, a cross prepreg having a reinforcing fiber layer in which fibers are crossed, and the like. Examples of the reinforcing fiber layer include a glass fiber layer, a carbon fiber layer, and an aramid fiber layer. Examples of the synthetic resin include a thermosetting resin such as an unsaturated polyester resin, an epoxy resin, and a vinyl ester resin, and AB.
There are thermoplastic resins such as S resin, polyamide resin and fluororesin. In the present embodiment, as an example, a carbon fiber layer is used as a reinforcing fiber layer, and a unidirectional prepreg using an epoxy resin as a synthetic resin is used. However, other prepregs using a reinforcing fiber layer and a synthetic resin may also be used. The composite material of the invention and the method for producing the composite material can be suitably implemented.

First, as a step before the first step, the prepreg sheet wound into a roll is cut into a desired shape according to the dimensions of the body panel 21.

Next, in the first step, one or more (two in this embodiment) prepregs 27a and 27b which are to be a part of the outer plate 23 cut into a desired shape are laminated on a jig. . At this time, the two prepregs 27 laminated
The reinforcing fibers a and 27b are laminated so as to cross each other at a predetermined angle, for example, 45 degrees.

Next, in the second step, the prepreg 27a
One or more layers (two layers in this embodiment) of the prepregs 27A and 27B which are part of the stringer 24 are laminated thereon. At this time, the two prepregs 27A, 2
The reinforcing fibers of 7B are laminated so as to cross each other at 45 degrees.

Next, in a third step, the four prepregs 27a, 27b, 27A, 27B laminated in the second step
In the vicinity of the step 30, as shown by a broken line 31, the width T2 predetermined by the reinforcing fiber along the step 30 is increased in one or more rows in a direction along the step (a direction perpendicular to the paper surface of FIG. 2). In this embodiment, five rows are stitched.

Prepregs 27a, 27b, 27A, 27
B, using a stitching device, as shown in FIG. 3, performs upper lock 35a and lower thread 35b made of reinforcing fibers to perform lock stitching in each row. Alternatively, hand-sewing may be performed without using the above-described stitching device. Further, as shown in FIG. 4, one needle 33 is alternately reciprocated between the prepregs 27b and 27B to form one line of one thread 35. The stitching method may be used.

Next, in a fourth step, the four prepregs 27 stitched in the third step as shown in FIG.
a, 27b, 27A and 27B, the remaining prepregs 27c to 27h constituting the outer plate 23 are laminated on the prepreg 27b which is a part of the outer plate 23, and then the stringer 24 is formed.
The remaining prepregs 27C to 27M constituting the stringer 24 are bent and laminated in an L-shape from both sides to form the stringer 24 as shown in FIG. A stringer 24 having a T-shaped cross section is formed thereon. The steps are T-shaped, H-shaped, J-shaped, Ω-shaped, L-shaped, and C-shaped.
There is a character shape.

As described above, the prepregs 27a to 27h and the prepregs 27A constituting the body panel 21
The stacking and stitching of ~ 27M are completed. At this time, the prepregs 27a to 27h and the prepregs 27A to 27M are in a semi-cured state.

Next, in a post-process of the fourth process, the body panel 21 in a semi-cured state is inserted into the autoclave.
When pressurized with an inert gas such as nitrogen gas and heated at the same time in the autoclave, the prepregs 27a to 27h and the prepregs 27A which were in a semi-cured state were obtained.
The synthetic resin of ~ 27M is cured, and the body panel 21 is completed.

FIG. 5 omits the projection of the stringer 24,
FIG. 6 is a model diagram showing a state where a compressive load is applied to the simplified body panel 21, and FIG. 6 is a model diagram showing a state where a tensile load is applied. This modeled thickness t2 of the outer panel 23 of the body panel 21 is 2.144 mm,
The total thickness t1 of the stringer 24 and the outer plate 23 is 3.
752 mm.

FIG. 7 is a diagram showing a result of performing a stress analysis when a compressive load is applied to the model shown in FIG. 5 by a computer simulation using a finite element method. FIG. 8 is a diagram showing a result of performing a stress analysis on the model shown in FIG. 6 when a tensile load is applied by computer simulation using a finite element method.

As shown in FIG. 5, arrows 39
When a compressive load of 2.94 N / mm (0.3 kgf / mm) is applied as shown in FIG.
In the vicinity of the step 30, a tensile stress is applied to the areas 42 to 46 around the step 30, and a compressive stress is applied to the area 47 separated from the step 30 as shown in FIG. You. In the area 42, 540-680P
a (0.55 to 0.69 kgf / mm ² ), the region 43 has 440 to 540 Pa (0.45 to 0.55 kgf / mm ² ).
mm ² ), the region 44 has 290 to 440 Pa (0.3
0 to 0.45 kgf / mm ² ).
~ 290 Pa (0.12-0.30 kgf / mm ² ),
In the region 46, 29 to 120 Pa (0.03 to 0.12
kgf / mm ² ) tensile stress is applied. From this result, it can be seen that when a compressive load is applied to the body panel 21, the tensile stress 60 acts on the vicinity of the step 30 in a concentrated manner. It can be seen that the tensile stress 60 may cause the outer plate 23 and the stringer 24 to peel in the direction indicated by the arrow 62 in the boundary layer 32. This tensile stress 6
As can be seen from FIG. 7, 0 is concentrated on the step portion 30, and in a region vertically separated from the step portion 30, the tensile stress 60 sharply decreases and hardly acts. Therefore,
The number of prepregs to be sewn is determined by the tensile stress 60
Are included in the area where the numbers are concentrated. For example, FIG.
As can be seen from FIG. 4, the intervals between the regions 42 to 44 are narrower, the stress is rapidly increased in the direction approaching the step 30, the regions 45 and 46 are wider from each other, and the regions 42 to 44 are wider.
Compared with 44, the stress increases more gradually in the direction approaching the step portion 30. Therefore, in the present embodiment, the region where the stress increases sharply is a region where a stress greater than or equal to a predetermined value acts, and the prepreg included in the region 44 is stitched.

As shown in FIG. 5, when a tensile load of 620 N / mm (63 kgf / mm) indicated by an arrow 48 is applied to the outer panel 23, the body panel model 21a
As shown in FIG. 5, the regions 50 to 55 centered on the step 30 are subjected to tensile stress, and the regions 56 to 58 separated from the step 30 are subjected to compressive stress. Note that the region 50 has 6640 to 7160 Pa (6.77 Pa).
７7.30 kgf / mm ² ), 5590
~ 6120Pa (5.70-6.24kgf / m
m ² ), and the region 52 has 4540-5590 Pa (4.
63 to 5.70 kgf / mm ² ).
50 to 4540 Pa (2.50 to 4.63 kgf / mm
² ) In the area 54, 1400-2450 Pa (1.4
3 to 2.50 kgf / mm ² ).
１1400 Pa (0.90 to 1.43 kgf / mm ² )
Tensile stress is applied. From these results, it was confirmed that the fuselage panel 2
It is understood that when a tensile load 48 is applied to 1, a tensile stress 61 acts near the step 30. It can be seen that the tensile stress 61 may cause the outer plate 23 and the stringer 24 to peel off in the boundary layer 32 in the direction indicated by the arrow 63. As can be seen from FIG. 8, the tensile stress 61 concentrates on the step portion 30, and in a region vertically separated from the step portion 30, the tensile stress 60 sharply decreases. Therefore, the number of prepregs to be stitched is
The number of sheets included in the area where the tensile stress 60 is concentrated. For example, as shown in FIG. 8, the intervals between the regions 50 to 52 are narrower, the stress is rapidly increased in the direction approaching the step portion 30, and the intervals between the regions 53 to 55 are wide,
The stress gradually increases in the direction approaching the step portion 30 as compared with the regions 50 to 52. Therefore, in the present embodiment, the region where the stress increases sharply is defined as a region where a stress greater than or equal to a predetermined value acts, and the prepreg included in the region 52 is stitched.

That is, when a compressive load 39 is applied,
All of the prepreg included in the region 44 where the stress is sharply increased and the prepreg included in the region 52 where the stress is rapidly increased when the tensile load 48 is applied. In the body panel 21 of the present embodiment, the prepregs included in the region 44 and the region 52 are prepregs 27a and 27b forming a part of the outer plate 23 and 27A and 27B forming a part of the stringer 24.

As shown in FIGS. 7 and 8, the tensile stress is generated only in the vicinity of the step portion 30 of the body panel 21, and the load is applied to all regions in the thickness direction (the vertical direction in FIGS. 7 and 8). Since the prepreg is not stitched in the conventional stitching technique shown in FIG. 9, the prepreg in the region not involved in the peeling and the prepreg in which the large peeling stress does not act are stitched, and it can be seen that there is waste. On the contrary,
In the stitching technique of the present embodiment, the prepreg is stitched only in the region where the abrupt tensile stress increases, so that unnecessary stitching is omitted. Thereby, other prepregs 27c to 27h and 27C which are not involved in the exfoliation other than the prepregs 27a, 27b, 27A and 27B.
To prevent damage to the reinforcing fiber layer of the prepregs 27c to 27h and 27C to 27M.
(The direction perpendicular to the plane of FIG. 1) of the reinforcing fiber layer is prevented from being reduced. Further, in the stitching technique of the method of manufacturing a complex material according to the present invention, stitching is performed in a state where the number of laminated prepregs is small in the third step, so that a large-scale stitching device is not required unlike the related art, and stitching is performed. This makes it easier to manufacture large components, such as components of passenger aircraft, from composite materials.

The composite material of the present invention can be suitably applied not only to a stepped member but also to a member such as a notch or a recess or a hole where stress concentration occurs when a stress is applied.

[0033]

Further, as another composite material and a method for producing the composite material of the present invention, not only a case where a plurality of layers of prepreg are laminated to produce a stepped member or a member having a notch but also a reinforcing fiber layer is prepared in advance. The present invention can also be suitably carried out when a stepped member or a member having a notch is manufactured by impregnating a synthetic resin after lamination. For example, in the first step and the second step, a plurality of reinforcing fiber layers constituting the outer panel and a plurality of reinforcing fiber layers constituting the stringer are laminated via a step, and then the outer panel is laminated in the third step. A reinforcing fiber layer,
Stitch the stringer with the reinforcing fiber layer. Thereafter, in a fourth step, a reinforcing fiber layer is further laminated and impregnated with the synthetic resin, and then the synthetic resin is cured at high temperature and high pressure in an autoclave. This also
A composite material having the same strength and peel resistance as a composite material manufactured using a prepreg can be manufactured.

[0035]

[0036]

[0037]

According to the present invention, a region where the stress increases sharply is determined by performing a stress analysis by computer simulation. Next, one part of the composite material is formed by laminating a plurality of predetermined reinforcing fiber layers or prepregs. Next, a plurality of predetermined reinforcing fiber layers or prepregs are laminated via a step on one of the parts laminated in the first step, and the other part is formed on one part. next,
One part and the other part are stitched along the step near the step part, and a reinforcing fiber layer or a prepreg is further laminated on the one part and the other part. When laminating a prepreg, the synthetic resin is cured at a high temperature and a high pressure, and when laminating a reinforcing fiber layer, the synthetic resin is impregnated with the resin and then the resin is cured under a high temperature and a high pressure to form a composite material. It is formed.

As described above, only the reinforcing fiber layer at the portion to be sewn in advance is firstly sewn and sewn, and then the remaining reinforcing fiber layer or prepreg is sewn up, so that only the predetermined region can be accurately sewn. Further, the reinforcing fiber layer in a region unrelated to the peeling is not damaged.

[Brief description of the drawings]

FIG. 1 is a view showing a part of a fuselage panel 21 manufactured by a method for manufacturing a composite material according to the present invention.

FIG. 2 is a view showing a state immediately after the end of a third step.

FIG. 3 is a diagram illustrating a method of performing a lock stitch on prepregs 27a, 27b, 27A, and 27B using a stitching device.

FIG. 4 shows prepregs 27a, 27b, 27A and 27B.
Is hand-sewn with one reinforcing fiber 35.

FIG. 5 is a model diagram showing a state in which a compressive load 39 is applied to the body panel 21 which is simplified by omitting a protrusion of the stringer 24.

FIG. 6 is a model diagram showing a state in which a tensile load 48 is applied to the simplified body panel 21 by omitting the protrusion of the stringer 24.

FIG. 7 shows that the modeled fuselage panel 21 has a compression load 3
9 is a stress analysis result when 9 is loaded.

FIG. 8 shows that the modeled fuselage panel 21 has a tensile load of 4
It is a figure which shows the stress analysis result when 8 is loaded.

FIG. 9 is a diagram showing a method for manufacturing a composite material according to the prior art.

[Explanation of symbols]

 21 Body Panel 23 Skin 24 Stringer 31 Stitching Position 32 Boundary Layer

──────────────────────────────────────────────────続 き Continued on the front page (58) Field surveyed (Int. Cl. ⁷ , DB name) B32B 1/00-35/00 B29C 70/00-70/88

Claims (1)

(57) [Claims] A composite material for use in an aircraft, comprising:
In a method of manufacturing a composite material in which one part where a plurality of layers of reinforcing fibers are laminated and the other part where a plurality of layers of reinforcing fibers are laminated are stitched and laminated via a step, the computer simulation of the composite material Applying a tensile and compressive load to the model to perform a stress analysis, and determining a region near a step portion where the stress increases sharply when a load is applied as a stitching region; and, among the one portion, the stitching A step of laminating a prepreg impregnated with a synthetic resin into the reinforcing fiber layer or the reinforcing fiber included in the region, and, among the other portions, the reinforcing fiber layer or the prepreg included in the stitched region; A step of laminating through a step, and the stitching area of a part of the one part and a part of the other part which are laminated And a step of laminating the remaining reinforcing fiber layers or prepregs of one part and the other part on the one part and the other part. .