Method for designing composite laminate and aircraft

Abstract

To improve design efficiency of a composite laminate.SOLUTION: A design method of a composite laminate comprises: a step of, on the basis of an allowable strain corresponding to a plurality of fiber orientation angles in a composite laminate comprising the plurality of fiber orientation angles, deriving an allowable load region in which, a combination of an allowable X-direction load in a prescribed X-direction in a plane orthogonal to a lamination direction where each layer of the composite laminate is laminated, an allowable Y-direction load in a y-direction orthogonal to the X-direction in the plane orthogonal to the lamination direction where each layer of the composite laminate is laminated, and an allowable shear load in a shear direction in the plane orthogonal to the lamination direction where each layer of the composite laminate is laminated, is indicated in a three-dimensional manner; and a step of determining whether there is included an action load acting to the composite laminate, in the allowable load region.SELECTED DRAWING: Figure 7

Description

Technical Field

[0001] The present invention relates to a method for designing a composite laminate and an aircraft.

Background Art

[0002] Patent Document 1 discloses performing strength evaluation of an adhesive joint structure using a composite laminate.

Prior Art Documents

Patent Documents

[0003]

Patent Document 1

Summary of the Invention

Problems to be Solved by the Invention

[0004] In recent years, in aircraft, composite laminates in which composite materials containing reinforcing fibers and resins are laminated have been used. Depending on the part of the aircraft, various loads are applied to the composite laminate. The composite laminate needs to have a laminated structure that can withstand the loads for each part of the aircraft. For example, for the laminated structure of the composite laminate, it is necessary to optimize combinations such as thickness, orientation ratio, and orientation direction for each part of the aircraft. Conventionally, in order to optimize the laminated structure of the composite laminate, it has taken a great deal of time and it has been difficult to improve the design efficiency of the composite laminate.

[0005] An object of the present invention is to provide a method for designing a composite laminate capable of improving the design efficiency of the composite laminate and an aircraft including the composite laminate designed by the design method.

Means for Solving the Problems

[0006] In order to solve the above problems, the method for designing a composite material laminate according to the present invention includes: based on the allowable strain corresponding to a plurality of fiber orientation angles in a composite material laminate having a plurality of fiber orientation angles, the allowable X-direction load in a predetermined x-direction in a plane orthogonal to the lamination direction in which each layer of the composite material laminate is laminated, the allowable Y-direction load in a y-direction orthogonal to the x-direction in a plane orthogonal to the lamination direction in which each layer of the composite material laminate is laminated, and deriving an allowable load region showing in three dimensions a combination of the allowable shear load in the shear direction in a plane within the plane orthogonal to the lamination direction in which each layer of the composite material laminate is laminated; and determining whether the applied load acting on the composite material laminate is included within the allowable load region.

[0007] The allowable load region may include steps of: a magnification of a basic lamination structure in which the number of layers of each fiber orientation angle is configured in a minimum unit based on an orientation ratio which is the ratio of the thickness of the composite material layer of each fiber orientation angle to the thickness of the composite material laminate; and selecting an orientation ratio that changes according to the orientation ratio and at which the magnification of the basic lamination structure is minimized when the applied load is included within the allowable load region.

[0008] An aircraft including a composite material laminate designed by the above design method.

Advantages of the Invention

[0009] According to the present invention, the design efficiency of the composite material laminate can be improved.

Brief Description of the Drawings

[0010]

Figure 1

Figure 2

Figure 3

Figure 4

Figure 5

Figure 6

Figure 7

Mode for Carrying Out the Invention

[0011] Hereinafter, preferred embodiments of the present invention will be described in detail with reference to the accompanying drawings. The dimensions, materials, and other specific numerical values shown in such embodiments are merely examples for facilitating the understanding of the invention, and do not limit the present invention unless otherwise specified. In the present specification and drawings, elements having substantially the same functions and configurations are denoted by the same reference numerals, and redundant descriptions are omitted, and elements not directly related to the present invention are not shown.

[0012] FIG. 1 is a block diagram showing a design apparatus 100 for a composite material laminate according to the present embodiment. The composite material laminate according to the present embodiment has a laminated structure in which a composite material containing reinforcing fibers and a resin is laminated. In the present embodiment, the composite material laminate is composed of a laminate of a plurality of prepregs. As the prepreg, there are a woven fabric material in which a resin is impregnated into a woven fabric having reinforcing fibers made of long fibers or continuous fibers as warp and weft, and a so-called UD (Uni-Direction) material in which reinforcing fibers made of long fibers or continuous fibers are oriented and arranged in one direction and impregnated with a resin. In the present embodiment, an example in which a UD material in which reinforcing fibers made of long fibers or continuous fibers are oriented in one direction is used as the prepreg will be described. Note that as the resin, a thermosetting resin or a thermoplastic resin is used. Specifically, the composite material laminate is composed of, for example, carbon fiber reinforced plastic (CFRP: Carbon Fiber Reinforced Plastics). However, it is not limited thereto, and the composite material laminate may be composed of other fiber reinforced plastics such as GFRP (Glass-Fiber-Reinforced Plastics) or AFRP (Aramid-Fiber-Reinforced Plastics). Further, in the present embodiment, the design of the composite material laminate used in an aircraft will be described. That is, the composite material laminate designed by the design method described below is used in an aircraft. In other words, the aircraft includes a composite material laminate designed by the design method described below. However, it is not limited thereto, and the composite material laminate designed by the design method described below may be used other than in an aircraft.

[0013] As shown in FIG. 1, the design apparatus 100 includes a database 102, a derivation unit 104, a determination unit 106, and a selection unit 108. The design apparatus 100 is, for example, a personal computer including a semiconductor integrated circuit including a central processing unit (CPU), a ROM storing a program, etc., and a RAM as a work area, and functions as the derivation unit 104, the determination unit 106, and the selection unit 108 by executing a program.

[0014] The database 102 is, for example, a hard disk, a flash memory, etc., and includes a non-volatile memory element. In the database 102, for example, the allowable value of strain (allowable strain) corresponding to the orientation direction (fiber orientation angle) of the reinforcing fibers of the layers constituting the composite material laminate is stored. Here, the orientation direction of the reinforcing fibers is, for example, 0°, ±45°, 90° when the longitudinal direction of the wing of an aircraft is set to 0°. The allowable value of strain corresponding to each orientation direction is, for example, a value obtained by a compression test, a tensile test, etc. of the composite material laminate. Also, in the database 102, data regarding the magnitude and direction of the applied load acting on each part of the aircraft is stored.

[0015] The derivation unit 104 derives the allowable load that can be applied to the composite material laminate. Generally, the strain and the curvature of bending that occur when a load or a bending moment is applied to the composite material laminate can be obtained by the stiffness matrix based on the laminate theory. Specifically, the strain and the curvature of bending that occur in the composite material laminate can be obtained based on the following Equation 1.

Equation

[0016] In the above Equation 1, Nx is the load applied in the x direction among the in-plane directions of the composite material laminate, Ny is the load applied in the y direction among the in-plane directions of the composite material laminate, and Nxy is the shear load in the in-plane directions (x, y directions) of the composite material laminate. Here, the in-plane direction of the composite material laminate is an arbitrary direction in a plane orthogonal to the direction in which the prepreg of the composite material laminate is laminated (hereinafter, also simply referred to as the lamination direction). The x direction is a predetermined direction in the plane, and the y direction is a direction orthogonal to the x direction in the plane. Also, Mx is the bending moment applied in the x direction of the composite material laminate, My is the bending moment applied in the y direction of the composite material laminate, and Mxy is the torsional moment applied in the x, y directions of the composite material laminate.

[0017] Also, in the above Equation (1), A, B, and D are stiffness matrices, εx is the strain generated in the x-direction of the composite laminate, εy is the strain generated in the y-direction of the composite laminate, and εxy is the strain generated in the x- and y-directions (shearing direction) of the composite laminate. Further, κx is the curvature of bending generated in the x-direction of the composite laminate, κy is the curvature of bending generated in the y-direction of the composite laminate, and κxy is the curvature of bending generated in the x- and y-directions of the composite laminate.

[0018] Here, generally, the composite laminate used for each part of an aircraft is designed so that deformation in the out-of-plane direction does not occur. For example, the composite laminate used for the main wing of an aircraft is configured to be symmetric in the stacking direction (thickness direction) so that almost no deformation in the out-of-plane direction occurs.

[0019] That is, the composite laminate of the present embodiment has a stacking structure that is symmetric with respect to the center in the thickness direction. Therefore, in the present embodiment, a design is performed without considering the deformation of the composite laminate in the out-of-plane direction. However, the present invention is not limited to this, and the composite laminate may be configured asymmetrically in the stacking direction, and a design may be performed considering the deformation of the composite laminate in both the in-plane direction and the out-of-plane direction.

[0020] When not considering the deformation of the composite laminate in the out-of-plane direction, in the above Equation (1), it is not necessary to consider D, Mx, My, Mxy, κx, κy, and κxy. Further, when the composite laminate is configured symmetrically, B = 0, so the above Equation (1) can be omitted as the following Equation (2). In Equation (2), the subscript of A represents the index in the stiffness matrix.

Equation

[0021] The following Equation (3) is obtained by transforming the above Equation (2). The left side of the above Equation (2) represents the load applied to the composite laminate, while the left side of the following Equation (3) represents the strain acting on the composite laminate. In Equation (3), a represents the inverse matrix of the stiffness matrix.

Mathematics

[0022] Here, the composite material laminate of the present embodiment includes a plurality of layers in which the orientation directions of the reinforcing fibers are different. That is, the composite material laminate in which the above-described UD material is laminated includes at least a plurality of layers in which the orientation directions of the reinforcing fibers are different. Specifically, the composite material laminate includes, for example, a plurality of layers in which the reinforcing fibers are oriented in the 0°, ±45°, and 90° directions when the longitudinal direction of the wing of the aircraft is set to 0°.

[0023] When the direction in which the reinforcing fibers are oriented (orientation direction) is axial, the direction orthogonal to the orientation direction of the reinforcing fibers is trans, and the shear direction of the reinforcing fibers is share, the above Equation 3 is converted into the following Equation 4 and the following Equation 5.

Mathematics

Mathematics

[0024] Here, the above Equation 4 represents the strain acting on the composite material layer in which the orientation direction of the reinforcing fibers is 0° and 90°, and the above Equation 5 represents the strain acting on the composite material layer in which the orientation direction of the reinforcing fibers is ±45°.

[0025] In the above Equations 4 and 5, T represents a coordinate transformation matrix, and the subscript of T indicates the orientation direction of the reinforcing fibers. In the above Equation 4, the equation for the strain acting on the layer in which the orientation direction of the reinforcing fibers is 90° is omitted because it is the same as the equation for the strain acting on the layer in which the orientation direction of the reinforcing fibers is 0°. Similarly, in the above Equation 5, the equation for the strain acting on the layer in which the orientation direction of the reinforcing fibers is -45° is omitted because it is the same as the equation for the strain acting on the layer in which the orientation direction of the reinforcing fibers is +45°.

[0026] When the above equations (4) and (5) are expanded, relational expressions between the strain and the load acting on the composite material laminate shown in the following equations (6) and (7) are obtained. In the following equations (6) and (7), OHT is the allowable value of tensile strain corresponding to the orientation direction of the reinforcing fibers of the composite material laminate, and OHC is the allowable value of compressive strain corresponding to the orientation direction of the reinforcing fibers of the composite material laminate.

Equation

Equation

[0027] In the above equations (6) and (7), the expressions on the right side of the arrow are obtained by substituting the allowable value of tensile strain or the allowable value of compressive strain corresponding to the orientation direction of the reinforcing fibers into the expressions on the left side of the arrow. By substituting the allowable value of tensile strain or the allowable value of compressive strain, the allowable loads (Nx, Ny, Nxy) that can be applied to the composite material laminate can be obtained. Hereinafter, the allowable load (Nx) in a predetermined x direction in a plane orthogonal to the lamination direction in which the prepreg of the composite material laminate is laminated is referred to as the allowable X-direction load. Also, the allowable load (Ny) in the y direction orthogonal to the x direction in a plane orthogonal to the lamination direction in which the prepreg of the composite material laminate is laminated is referred to as the allowable Y-direction load. Further, the allowable load (Nxy) in the shear direction in a plane orthogonal to the lamination direction in which the prepreg of the composite material laminate is laminated is referred to as the allowable shear load. In other words, the allowable X-direction load refers to the allowable load (Nx) in a predetermined x direction in a plane orthogonal to the lamination direction in which each layer of the composite material laminate is laminated. Also, the allowable Y-direction load refers to the allowable load (Ny) in the y direction orthogonal to the x direction in a plane orthogonal to the lamination direction in which each layer of the composite material laminate is laminated. Further, the allowable shear load refers to the allowable load (Nxy) in the shear direction in a plane orthogonal to the lamination direction in which each layer of the composite material laminate is laminated.

[0028] Figure 2 is an image diagram of the allowable load region R that can be applied to the composite laminate. As shown in Figure 2, the allowable load region R is formed by eight surfaces (S1, S2, S3, S4, S5, S6, S7, S8), and the region surrounded by the eight surfaces is the allowable load region R. In Figure 2, the eight surfaces S1 to S8 respectively correspond to S1 to S8 shown in the above equations (6) and (7).

[0029] Surface S1 is the surface representing the maximum value of the tensile allowable load in the orientation direction of 0°. Surface S2 is the surface representing the maximum value of the tensile allowable load in the orientation direction of +45°. Surface S3 is the surface representing the maximum value of the tensile allowable load in the orientation direction of 90°. Surface S4 is the surface representing the maximum value of the tensile allowable load in the orientation direction of -45°. Surface S5 is the surface representing the maximum value of the compressive allowable load in the orientation direction of 0°. Surface S6 is the surface representing the maximum value of the compressive allowable load in the orientation direction of +45°. Surface S7 is the surface representing the maximum value of the compressive allowable load in the orientation direction of 90°. Surface S8 is the surface representing the maximum value of the compressive allowable load in the orientation direction of -45°.

[0030] When the applied load acting on the composite laminate is within the allowable load region R, it can be determined that the composite laminate can withstand the applied load. On the other hand, when the applied load acting on the composite laminate is outside the allowable load region R, it can be determined that there is a possibility that the composite laminate cannot withstand the applied load.

[0031] Figure 3 is an image diagram of the allowable load region R that changes according to the orientation ratio of the composite laminate. Here, the orientation ratio is the ratio of the thickness of the composite layer with each fiber orientation angle to the total thickness of the composite laminate. In Figure 3, the allowable load region Ra is the allowable load region when the orientation ratios of the composite laminate are approximately equal (0° = 25%, +45° = 25%, -45° = 25%, 90° = 25%) (hereinafter also referred to as a pseudo-isotropic laminate).

[0032] The allowable load region Rb is the allowable load region when the orientation ratio of 90° in the composite laminate is larger than the other orientation ratios (0°, +45°, -45°) (hereinafter also referred to as a 90°-rich laminate). For example, the orientation ratios of a 90°-rich laminate are 0° = 20%, +45° = 20%, -45° = 20%, and 90° = 40%. The allowable load region Rb has a smaller region in the Nx direction and a larger region in the Ny direction compared to the allowable load region Ra.

[0033] The allowable load region Rc is the allowable load region when the orientation ratio of 0° in the composite laminate is larger than the other orientation ratios (+45°, -45°, 90°) (hereinafter also referred to as a 0°-rich laminate). For example, the orientation ratios of a 0°-rich laminate are 0° = 40%, +45° = 20%, -45° = 20%, and 90° = 20%. The allowable load region Rc has a larger region in the Nx direction and a smaller region in the Ny direction compared to the allowable load region Ra. Thus, the allowable load region R changes according to the orientation ratio of the composite laminate.

[0034] Figure 4 is an image diagram of the allowable load region R that changes according to the structure of the composite laminate. In Figure 4, the allowable load region Rd is the allowable load region when the composite laminate has a through-hole for inserting a fastening member (e.g., a fastener). The allowable load region Re is the allowable load region when the composite laminate has a through-hole and a fastening member (fastener) is embedded in the through-hole. As shown in Figure 4, the allowable load region Re is larger than the allowable load region Rd. This is because when a fastening member is embedded in the through-hole of the composite laminate, the strength becomes greater than when the fastening member is not embedded in the through-hole.

[0035] FIG. 5 is a first image diagram of an allowable load region R that changes according to the magnification of the basic lamination structure of the composite material laminate. FIG. 6 is a second image diagram of the allowable load region R that changes according to the magnification of the basic lamination structure of the composite material laminate. Here, the basic lamination structure is a structure of a composite material laminate in which the number of layers in each orientation direction is configured in the minimum unit based on the orientation ratio of the composite material laminate. Specifically, the basic lamination structure is a structure of a composite material laminate in which the number of layers in each orientation direction is configured in the minimum unit in each of a pseudo-isotropic laminate, a 90°-rich laminate, and a 0°-rich laminate. For example, in the case of a pseudo-isotropic laminate, the basic lamination structure is a lamination structure of a total of 4 plies, with 1 layer (1 ply) in the 0° orientation direction, 1 layer (1 ply) in the +45° orientation direction, 1 layer (1 ply) in the -45° orientation direction, and 1 layer (1 ply) in the 90° orientation direction. Here, the thickness of 1 ply is equal for each layer regardless of the orientation direction.

[0036] In FIGS. 5 and 6, the allowable load region Rf is the allowable load region when the 0°-rich laminate is composed only of the basic lamination structure. Here, the basic lamination structure is, for example, a lamination structure of a total of 5 plies, with 2 layers (2 plies) in the 0° orientation direction, 1 layer (1 ply) in the +45° orientation direction, 1 layer (1 ply) in the -45° orientation direction, and 1 layer (1 ply) in the 90° orientation direction.

[0037] The allowable load region Rfa shown in FIG. 5 is obtained by multiplying the allowable load region Rf by an integer multiple (here, 2 times as the minimum magnification) so that the applied load (particularly the applied load Fa) is included in the allowable load region. That is, the allowable load region Rfa is the allowable load region when the magnification of the basic lamination structure (5 plies) of the 0°-rich laminate is doubled (10 plies). Also, the allowable load region Rfb shown in FIG. 6 is obtained by multiplying the allowable load region Rf by an integer multiple (here, 4 times as the minimum magnification) so that the applied load (particularly the applied load Fb) is included in the allowable load region. That is, the allowable load region Rfa is the allowable load region when the magnification of the basic lamination structure (5 plies) of the 0°-rich laminate is quadrupled (20 plies).

[0038] Also, in FIGS. 5 and 6, the allowable load region Rg is the allowable load region when the 90° rich laminate is laminated only by the basic laminate configuration. Here, the basic laminate configuration is, for example, a laminate configuration of a total of 5 plies, including 1 ply (1 ply) of a layer with an orientation direction of 0°, 1 ply (1 ply) of a layer with an orientation direction of +45°, 1 ply (1 ply) of a layer with an orientation direction of -45°, and 2 plies (2 plies) of a layer with an orientation direction of 90°.

[0039] The allowable load region Rga shown in FIG. 5 is obtained by multiplying the allowable load region Rg by an integer multiple (here, 4 times as the minimum multiple) so that the applied load (especially the applied load Fa) is included in the allowable load region. That is, the allowable load region Rga is the allowable load region when the magnification of the basic laminate configuration (5 plies) of the 90° rich laminate is 4 times (20 plies). Also, the allowable load region Rgb shown in FIG. 6 is obtained by multiplying the allowable load region Rg by an integer multiple (here, 2 times as the minimum multiple) so that the applied load (especially the applied load Fb) is included in the allowable load region. That is, the allowable load region Rgb is the allowable load region when the magnification of the basic laminate configuration (5 plies) of the 90° rich laminate is 2 times (10 plies). Thus, the allowable load region R changes similarly according to the magnification of the basic laminate configuration of the composite laminate.

[0040] As shown in FIG. 5, neither the allowable load region Rf nor the allowable load region Rg can contain the applied load (especially the applied load Fa) acting on the composite laminate within the region. On the other hand, both the allowable load region Rfa and the allowable load region Rga can contain the applied load (especially the applied load Fa) acting on the composite laminate within the region.

[0041] Here, the laminate configuration (10 plies) of the allowable load region Rfa is smaller than the laminate configuration (20 plies) of the allowable load region Rga. That is, the thickness of the laminate configuration of the allowable load region Rfa is smaller than the thickness of the laminate configuration of the allowable load region Rga. Therefore, the allowable load region Rfa can make the thickness of the laminate configuration thinner while containing the applied load acting on the composite laminate within the region, compared to the allowable load region Rga.

[0042] Similarly, as shown in FIG. 6, in neither the allowable load region Rf nor the allowable load region Rg can the applied load (particularly the applied load Fb) acting on the composite material laminate be contained within the region. On the other hand, in both the allowable load region Rfb and the allowable load region Rgb, the applied load (particularly the applied load Fb) acting on the composite material laminate can be contained within the region.

[0043] Here, the lamination structure (10 plies) of the allowable load region Rgb is smaller than the lamination structure (20 plies) of the allowable load region Rfb. That is, the thickness of the lamination structure of the allowable load region Rgb is smaller than the thickness of the lamination structure of the allowable load region Rfb. Therefore, the allowable load region Rgb can keep the applied load acting on the composite material laminate within the region while making the thickness of the lamination structure thinner than that of the allowable load region Rfb.

[0044] Hereinafter, the design method of the composite material laminate of the present embodiment will be described. FIG. 7 is a flowchart showing the design method of the composite material laminate of the present embodiment. First, the derivation unit 104 sets the basic lamination structure in each of the pseudo-isotropic laminate, 90°-rich laminate, and 0°-rich laminate (step S101). The derivation unit 104 acquires the allowable value of strain corresponding to the orientation direction of the reinforcing fibers of the composite material laminate from the database 102.

[0045] Then, based on the above equations 6 and 7, the derivation unit 104 derives S1 to S8 which are eight surfaces, and derives a plurality of allowable load regions R corresponding to the orientation direction of the reinforcing fibers from S1 to S8. That is, the derivation unit 104 derives the allowable load region R showing in three dimensions the combination of the allowable X-direction load, allowable Y-direction load, and allowable shear load in the in-plane direction of the composite material laminate based on the allowable value of strain corresponding to the orientation direction of the reinforcing fibers (step S102).

[0046] In the present embodiment, the derivation unit 104 derives the allowable load region R for each case where the 0°-rich laminate, pseudo-isotropic laminate, and 90°-rich laminate are composed of the basic lamination structure. That is, the derivation unit 104 derives the allowable load region R of the basic lamination structure at each orientation ratio of the composite material laminate.

[0047] Next, the determination unit 106 acquires data regarding the magnitude and direction of the applied load on each part of the aircraft from the database 102. Then, the determination unit 106 determines whether the applied load is included within the allowable load region R derived in step S102 from the acquired data (step S103).

[0048] When the applied load is outside the allowable load region R for all the orientation ratios, the derivation unit 104 greatly sets the magnification of the basic lamination structure for each orientation ratio (for example, 1 time → 2 times) (step S104). Then, the derivation unit 104 re-derives the allowable load region R for each orientation ratio after changing the magnification of the basic lamination structure (step S102).

[0049] The determination unit 106 determines whether the applied load is included within the re-derived allowable load region R (step S103). In this way, when the applied load is not included within any of the allowable load regions R of the basic lamination structure for each orientation ratio, the magnification of the basic lamination structure is multiplied by an integer, and the magnification is increased until the applied load is included within the allowable load region R. However, it is not limited to this, and the magnification of the basic lamination structure may be a decimal or a fraction. Further, the derivation unit 104 may derive the magnification of the basic lamination structure for each orientation ratio such that the applied load is included within the allowable load region R, and derive the allowable load region R corresponding to each magnification of the basic lamination structure for each orientation ratio.

[0050] On the other hand, when the applied load is within the allowable load region R for any of the orientation ratios, the selection unit 108 selects the orientation ratio with the minimum magnification of the basic lamination structure (step S105). Here, when there are a plurality of orientation ratios with the minimum magnification of the basic lamination structure, one orientation ratio may be selected from the plurality of orientation ratios based on the preset priority of the orientation ratios. Thereby, the lamination structure of the lightest composite material laminate that can withstand the applied load can be selected.

[0051] For example, in the case of FIG. 5, the orientation ratio of the 0°-rich laminate is selected where the magnification of the basic laminate structure becomes 2 times. Also, in the case of FIG. 6, the orientation ratio of the 90°-rich laminate is selected where the magnification of the basic laminate structure becomes 2 times.

[0052] Thus, according to the method for designing the composite material laminate of the present embodiment, an allowable load region R that three-dimensionally shows the combination of the allowable X-direction load, the allowable Y-direction load, and the allowable shear load in the in-plane direction of the composite material laminate is derived. Also, it is determined whether or not the applied load is included in the derived allowable load region R. Therefore, it is possible to quickly evaluate whether or not the strength is established for the in-plane loads (Nx, Ny, Nxy) applied to the composite material laminate for a large number of load cases. Accordingly, it is possible to efficiently optimize the thickness, orientation ratio, orientation direction, etc. of the composite material laminate, and improve the design efficiency of the composite material laminate.

[0053] As described above, the preferred embodiments of the present invention have been described with reference to the accompanying drawings. Needless to say, the present invention is not limited to such embodiments. It is obvious that those skilled in the art can conceive of various modification examples or correction examples within the scope described in the claims, and it is naturally understood that they also belong to the technical scope of the present invention.

[0054] In the above embodiment, the case where the orientation ratio of the composite material laminate is a 0°-rich laminate, a pseudo-isotropic laminate, or a 90°-rich laminate has been described. However, it is not limited thereto, and the orientation ratio of the composite material laminate may be, for example, a +45°-rich laminate or a -45°-rich laminate. Here, the basic laminate configuration of the +45°-rich laminate is, for example, a laminate configuration of a total of 5 plies including 1 ply of a layer in the 0° orientation direction, 2 plies of a layer in the +45° orientation direction, 1 ply of a layer in the -45° orientation direction, and 1 ply of a layer in the 90° orientation direction. Further, the basic laminate configuration of the -45°-rich laminate is, for example, a laminate configuration of a total of 5 plies including 1 ply of a layer in the 0° orientation direction, 1 ply of a layer in the +45° orientation direction, 2 plies of a layer in the -45° orientation direction, and 1 ply of a layer in the 90° orientation direction. Note that the basic laminate configuration of the pseudo-isotropic laminate is, for example, a laminate configuration of a total of 4 plies including 1 ply of a layer in the 0° orientation direction, 1 ply of a layer in the +45° orientation direction, 1 ply of a layer in the -45° orientation direction, and 1 ply of a layer in the 90° orientation direction.

[0055] In the above embodiment, the example in which the design device 100 includes the selection unit 108 has been described. However, the selection unit 108 is not an essential configuration, and for example, the design device 100 may not be provided with the selection unit 108.

Explanation of Signs

[0056] R Allowable Load Region 100 Design Device 102 Database 104 Derivation Unit 106 Determination Unit 108 Selection Unit

Claims

1. Based on the allowable strain corresponding to the plurality of fiber orientation angles in a composite material laminate having a plurality of fiber orientation angles, the allowable X-direction load in a predetermined x-direction in a plane orthogonal to the lamination direction in which each layer of the composite material laminate is laminated, the allowable Y-direction load in a y-direction orthogonal to the x-direction in a plane orthogonal to the lamination direction in which each layer of the composite material laminate is laminated, and deriving an allowable load region showing, in three dimensions, a combination of the allowable shear loads in the shear direction within a plane orthogonal to the lamination direction in which each layer of the composite material laminate is laminated; Determining whether the applied load acting on the composite material laminate is included within the allowable load region; A design method for a composite material laminate including the above.

2. The allowable load region is based on the magnification of the basic lamination structure in which the number of layers of each fiber orientation angle is configured in the minimum unit based on the orientation ratio, which is the ratio of the thickness of the composite material layer of each fiber orientation angle to the thickness of the composite material laminate, and changes depending on the orientation ratio, Including the step of selecting the orientation ratio at which the magnification of the basic lamination structure is minimized when the applied load is included within the allowable load region, The design method for a composite material laminate according to Claim 1.

3. An aircraft comprising a composite material laminate designed by the design method according to Claim 1 or 2.

Images