JP7709355B2 - Composite laminate design method and aircraft - Google Patents

Description

The present invention relates to a method for designing composite laminates and aircraft.

Patent Document 1 discloses a method for evaluating the strength of adhesive joint structures using composite laminates.

JP 2008-108242 A

In recent years, aircraft have been using composite laminates in which composite materials containing reinforcing fibers and resins are laminated. Various loads are applied to the composite laminates depending on the part of the aircraft. The composite laminates need to have a laminate structure that can withstand the loads of each part of the aircraft. For example, the laminate structure of the composite laminate needs to be optimized in terms of the combination of thickness, orientation ratio, orientation direction, etc. for each part of the aircraft. Conventionally, optimizing the laminate structure of a composite laminate required a great deal of time, making it difficult to improve the design efficiency of composite laminates.

The present invention aims to provide a method for designing a composite laminate that can improve the design efficiency of the composite laminate, and an aircraft equipped with a composite laminate designed by the design method.

In order to solve the above problems, the design method of the composite laminate of the present invention includes the steps of: deriving an allowable load region that shows in three dimensions a combination of an allowable X-direction load in a predetermined x-direction in a plane perpendicular to the stacking direction in which each layer of the composite laminate is stacked, an allowable Y-direction load in a y-direction perpendicular to the x-direction in a plane perpendicular to the stacking direction in which each layer of the composite laminate is stacked, and an allowable shear load in a shear direction in a plane perpendicular to the stacking direction in which each layer of the composite laminate is stacked, based on allowable strains corresponding to multiple fiber orientation angles in a composite laminate having multiple fiber orientation angles; and determining whether or not the applied load acting on the composite laminate is included in the allowable load region.

The allowable load region may include a step of selecting a magnification of a basic stacking configuration in which the number of layers of each fiber orientation angle is the smallest unit based on an orientation ratio, which is the ratio of the thickness of the composite layer of each fiber orientation angle to the thickness of the composite laminate, and an orientation ratio that varies with the orientation ratio and is the smallest such that the applied load is included within the allowable load region.

An aircraft equipped with a composite laminate designed using the above design method.

The present invention can improve the design efficiency of composite laminates.

FIG. 1 is a block diagram showing a design device for a composite laminate according to an embodiment of the present invention. FIG. 2 is an image diagram of an allowable load range that can be applied to a composite laminate. FIG. 1 is an image diagram of an allowable load region that changes depending on the orientation ratio of a composite laminate. FIG. 1 is an image diagram of the allowable load range that changes depending on the structure of a composite laminate. FIG. 1 is a first image diagram of an allowable load region that changes depending on the magnification of a basic stacking configuration of a composite laminate. FIG. 11 is a second image diagram of the allowable load region that changes depending on the magnification of the basic stack configuration of the composite laminate. FIG. 2 is a flow chart showing a method for designing a composite laminate according to the present embodiment.

The preferred embodiment of the present invention will be described in detail below with reference to the attached drawings. The dimensions, materials, and other specific values shown in the embodiment are merely examples to facilitate 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 duplicated explanations, and elements not directly related to the present invention are not illustrated.

Figure 1 is a block diagram showing a design device 100 for a composite laminate according to the present embodiment. The composite laminate according to the present embodiment has a laminated structure in which composite materials containing reinforcing fibers and resin are laminated. In the present embodiment, the composite laminate is composed of a laminate of a plurality of prepregs. Examples of prepregs include woven fabrics in which reinforcing fibers made of long fibers or continuous fibers are used as warp and weft threads and impregnated with resin, and so-called UD (Uni-Direction) materials in which reinforcing fibers made of long fibers or continuous fibers are oriented in one direction and impregnated with 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 thermosetting resins and thermoplastic resins are used as the resins. Specifically, the composite laminate is composed of, for example, carbon fiber reinforced plastics (CFRP: Carbon Fiber Reinforced Plastics). However, without being limited thereto, the composite laminate may be made of other fiber-reinforced plastics such as glass-fiber-reinforced plastics (GFRP) or aramid-fiber-reinforced plastics (AFRP). In addition, in this embodiment, the design of a composite laminate used in an aircraft is described. That is, a composite laminate designed by the design method described below is used in an aircraft. In other words, an aircraft includes a composite laminate designed by the design method described below. However, without being limited thereto, a composite laminate designed by the design method described below may be used for something other than an aircraft.

As shown in FIG. 1, the design device 100 includes a database 102, a derivation unit 104, a judgment unit 106, and a selection unit 108. The design device 100 is, for example, a personal computer equipped with a semiconductor integrated circuit including a central processing unit (CPU), a ROM in which programs and the like are stored, and a RAM as a work area, and functions as the derivation unit 104, the judgment unit 106, and the selection unit 108 by executing a program.

The database 102 is, for example, a hard disk or a flash memory, and includes a non-volatile storage element. The database 102 stores, for example, the allowable strain (allowable strain) corresponding to the orientation direction (fiber orientation angle) of the reinforcing fibers of the layers that make up the composite laminate. Here, the orientation direction of the reinforcing fibers is, for example, 0°, ±45°, and 90° when the longitudinal direction of the aircraft wing is set to 0°. The allowable strain corresponding to each orientation direction is, for example, a value obtained by a compression test or a tensile test of the composite laminate. The database 102 also stores data on the magnitude and direction of the applied load acting on each part of the aircraft.

The derivation unit 104 derives the allowable load that can be applied to the composite laminate. In general, the distortion and bending curvature that occur when a load or bending moment is applied to the composite laminate can be calculated using a stiffness matrix based on lamination theory. Specifically, the distortion and bending curvature that occur in the composite laminate can be calculated based on the following formula 1.

In the above formula 1, Nx is the load applied in the x direction among the in-plane directions of the composite laminate, Ny is the load applied in the y direction among the in-plane directions of the composite laminate, and Nxy is the shear load in the in-plane directions (x and y directions) of the composite laminate. Here, the in-plane direction of the composite laminate is any direction in a plane perpendicular to the direction in which the prepregs of the composite laminate are stacked (hereinafter also simply referred to as the stacking direction). The x direction is a specified direction in that plane, and the y direction is a direction perpendicular to the x direction in that plane. Also, Mx is the bending moment applied in the x direction of the composite laminate, My is the bending moment applied in the y direction of the composite laminate, and Mxy is the torsional moment applied in the x and y directions of the composite laminate.

In addition, in the above formula 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 (shear direction) of the composite laminate. Also, 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.

Generally, the composite laminates used in each part of an aircraft are designed to prevent deformation in the out-of-plane direction. For example, the composite laminates used in the main wings of an aircraft are constructed symmetrically in the lamination direction (thickness direction) so that deformation in the out-of-plane direction is almost nonexistent.

In other words, the composite laminate of this embodiment has a stacking configuration that is symmetrical with respect to the center in the thickness direction. Therefore, in this embodiment, the design does not take into account deformation in the out-of-plane direction of the composite laminate. However, this is not limited to this, and the composite laminate may be configured asymmetrically in the stacking direction, or may be designed to take into account deformation in both the in-plane direction and the out-of-plane direction of the composite laminate.

When the out-of-plane deformation of the composite laminate is not taken into consideration, it is not necessary to consider D, Mx, My, Mxy, κx, κy, and κxy in the above formula 1. Furthermore, when the composite laminate is configured symmetrically, B=0, and therefore the above formula 1 can be abbreviated to the following formula 2. In formula 2, the subscript of A represents an index in the stiffness matrix.

The following Equation 3 is a conversion of the above Equation 2, where the left side of the above Equation 2 represents the load applied to the composite laminate, whereas the left side of the below Equation 3 represents the strain acting on the composite laminate. In Equation 3, a represents the inverse matrix of the stiffness matrix.

Here, the composite laminate of this embodiment has multiple layers with different orientation directions of reinforcing fibers. In other words, the composite laminate in which the above-mentioned UD material is layered has at least multiple layers with different orientation directions of reinforcing fibers. Specifically, the composite laminate has multiple layers in which the reinforcing fibers are oriented at 0°, ±45°, and 90° directions, for example, when the longitudinal direction of the aircraft wing is 0°.

If the direction in which the reinforcing fibers are oriented (orientation direction) is axial, the direction perpendicular to the orientation direction of the reinforcing fibers is trans, and the shear direction of the reinforcing fibers is share, then the above formula 3 is converted to the following formula 4 and formula 5.

Here, the above formula 4 represents the strain acting on a composite layer whose reinforcing fiber orientation directions are 0° and 90°, and the above formula 5 represents the strain acting on a composite layer whose reinforcing fiber orientation directions are ±45°.

In the above formulas 4 and 5, T indicates the coordinate transformation matrix, and the subscript of T indicates the orientation direction of the reinforcing fibers. In the above formula 4, the equation for the strain acting on the layer whose reinforcing fiber orientation direction is 90° is omitted because it is the same as the equation for the strain acting on the layer whose reinforcing fiber orientation direction is 0°. Similarly, in the above formula 5, the equation for the strain acting on the layer whose reinforcing fiber orientation direction is -45° is omitted because it is the same as the equation for the strain acting on the layer whose reinforcing fiber orientation direction is +45°.

By expanding the above formulas 4 and 5, the relational expressions between the strain and the load acting on the composite laminate are obtained as shown in the following formulas 6 and 7. In the following formulas 6 and 7, OHT is the allowable tensile strain depending on the orientation direction of the reinforcing fibers of the composite laminate, and OHC is the allowable compressive strain depending on the orientation direction of the reinforcing fibers of the composite laminate.

In the above formulas 6 and 7, the formulas on the right side of the arrows are obtained by substituting the allowable value of tensile strain or the allowable value of compressive strain according to the orientation direction of the reinforcing fibers into the formulas on the left side of the arrows. By substituting the allowable value of tensile strain or the allowable value of compressive strain, the allowable load (Nx, Ny, Nxy) that can be applied to the composite laminate can be obtained. Hereinafter, the allowable load (Nx) in a specified x direction in a plane perpendicular to the stacking direction in which the prepregs of the composite laminate are stacked is referred to as the allowable X-direction load. In addition, the allowable load (Ny) in the y direction perpendicular to the x direction in a plane perpendicular to the stacking direction in which the prepregs of the composite laminate are stacked is referred to as the allowable Y-direction load. In addition, the allowable load (Nxy) in the shear direction in a plane perpendicular to the stacking direction in which the prepregs of the composite laminate are stacked is referred to as the allowable shear load. In other words, the allowable X-direction load refers to the allowable load (Nx) in a specified x direction in a plane perpendicular to the stacking direction in which each layer of the composite laminate is stacked. The allowable Y-direction load refers to the allowable load (Ny) in the y-direction perpendicular to the x-direction in a plane perpendicular to the stacking direction in which each layer of the composite laminate is stacked. The allowable shear load refers to the allowable load (Nxy) in the shear direction in a plane perpendicular to the stacking direction in which each layer of the composite laminate is stacked.

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

Surface S1 is the surface that represents the maximum allowable tensile load in the 0° orientation direction. Surface S2 is the surface that represents the maximum allowable tensile load in the +45° orientation direction. Surface S3 is the surface that represents the maximum allowable tensile load in the 90° orientation direction. Surface S4 is the surface that represents the maximum allowable tensile load in the -45° orientation direction. Surface S5 is the surface that represents the maximum allowable compressive load in the 0° orientation direction. Surface S6 is the surface that represents the maximum allowable compressive load in the +45° orientation direction. Surface S7 is the surface that represents the maximum allowable compressive load in the 90° orientation direction. Surface S8 is the surface that represents the maximum allowable compressive load in the -45° orientation direction.

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 risk that the composite laminate will not be able to withstand the applied load.

Figure 3 is an image diagram of the allowable load region R, which changes depending on 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 referred to as a quasi-isotropic laminate).

The allowable load region Rb is the allowable load region when the 90° orientation ratio of the composite laminate is greater than the other orientation ratios (0°, +45°, -45°) (hereinafter 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 area in the Nx direction and a larger area in the Ny direction than the allowable load region Ra.

The allowable load region Rc is the allowable load region when the 0° orientation ratio of the composite laminate is greater than the other orientation ratios (+45°, -45°, 90°) (hereinafter referred to as a 0° rich laminate). For example, the orientation ratios of a 0° rich laminate are 0° = 40%, +45° = 20%, -45° = 20%, 90° = 20%. The allowable load region Rc is larger in the Nx direction and smaller in the Ny direction than the allowable load region Ra. In this way, the allowable load region R changes depending on the orientation ratio of the composite laminate.

Figure 4 is an image diagram of the allowable load region R, which changes depending on 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 the strength is greater when a fastening member is embedded in the through hole of the composite laminate than when no fastening member is embedded in the through hole.

Figure 5 is a first image of the allowable load region R that changes according to the magnification of the basic lamination structure of the composite laminate. Figure 6 is a second image of the allowable load region R that changes according to the magnification of the basic lamination structure of the composite laminate. Here, the basic lamination structure is a composite laminate structure in which the number of layers in each orientation direction is the minimum unit based on the orientation ratio of the composite laminate. Specifically, the basic lamination structure is a composite laminate structure in which the number of layers in each orientation direction is the minimum unit in each of the pseudo-isotropic laminate, 90° rich laminate, and 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 four plies, including one layer (1 ply) in the 0° orientation direction, one layer (1 ply) in the +45° orientation direction, one layer (1 ply) in the -45° orientation direction, and one layer (1 ply) in the 90° orientation direction. Here, the thickness of one ply is the same regardless of the layers in each orientation direction.

In Figures 5 and 6, the allowable load region Rf is the allowable load region when the 0° rich laminate is composed only of the basic laminate structure. Here, the basic laminate structure is, for example, a laminate structure of a total of five plies, including two layers (2 plies) with an orientation direction of 0°, one layer (1 ply) with an orientation direction of +45°, one layer (1 ply) with an orientation direction of -45°, and one layer (1 ply) with an orientation direction of 90°.

The allowable load region Rfa shown in FIG. 5 is an integer multiple (here, a minimum multiple of 2) of the allowable load region Rf so that the applied load (particularly the applied load Fa) is included within the allowable load region. In other words, the allowable load region Rfa is the allowable load region when the multiplication factor of the basic stacking structure (5 plies) of the 0° rich laminate is doubled (10 plies). The allowable load region Rfb shown in FIG. 6 is an integer multiple (here, a minimum multiple of 4) of the allowable load region Rf so that the applied load (particularly the applied load Fb) is included within the allowable load region. In other words, the allowable load region Rfa is the allowable load region when the multiplication factor of the basic stacking structure (5 plies) of the 0° rich laminate is quadrupled (20 plies).

In addition, in Figures 5 and 6, the allowable load region Rg is the allowable load region when the 90° rich laminate is laminated only with the basic laminate structure. Here, the basic laminate structure is, for example, a laminate structure of a total of five plies, including one layer (1 ply) with an orientation direction of 0°, one layer (1 ply) with an orientation direction of +45°, one layer (1 ply) with an orientation direction of -45°, and two layers (2 ply) with an orientation direction of 90°.

The allowable load region Rga shown in FIG. 5 is an integer multiple (here, four times as the minimum multiple) of the allowable load region Rg so that the applied load (particularly the applied load Fa) is included within the allowable load region. In other words, the allowable load region Rga is an allowable load region when the multiplication factor of the basic stacking structure (5 plies) of the 90° rich laminate is four times (20 plies). The allowable load region Rgb shown in FIG. 6 is an integer multiple (here, two times as the minimum multiple) of the allowable load region Rg so that the applied load (particularly the applied load Fb) is included within the allowable load region. In other words, the allowable load region Rgb is an allowable load region when the multiplication factor of the basic stacking structure (5 plies) of the 90° rich laminate is two times (10 plies). In this way, the allowable load region R changes similarly depending on the multiplication factor of the basic stacking structure of the composite laminate.

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

Here, the laminate structure (10 plies) of the allowable load region Rfa is smaller than the laminate structure (20 plies) of the allowable load region Rga. In other words, the thickness of the laminate structure of the allowable load region Rfa is smaller than the thickness of the laminate structure of the allowable load region Rga. Therefore, the thickness of the laminate structure can be made thinner in the allowable load region Rfa than in the allowable load region Rga, while still containing the applied load acting on the composite laminate within the region.

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

Here, the laminate structure (10 plies) of the allowable load region Rgb is smaller than the laminate structure (20 plies) of the allowable load region Rfb. In other words, the thickness of the laminate structure of the allowable load region Rgb is smaller than the thickness of the laminate structure of the allowable load region Rfb. Therefore, the thickness of the laminate structure can be made thinner in the allowable load region Rgb than in the allowable load region Rfb, while still containing the applied load acting on the composite laminate within the region.

The design method for the composite laminate of this embodiment will be described below. FIG. 7 is a flow chart showing the design method for the composite laminate of this embodiment. First, the derivation unit 104 sets the basic laminate configuration for each of the pseudo-isotropic laminate, the 90° rich laminate, and the 0° rich laminate (step S101). The derivation unit 104 obtains the allowable strain value corresponding to the orientation direction of the reinforcing fibers of the composite laminate from the database 102.

Then, the derivation unit 104 derives eight surfaces S1 to S8 based on the above formulas 6 and 7, and derives multiple allowable load regions R from S1 to S8 corresponding to the orientation direction of the reinforcing fibers. In other words, the derivation unit 104 derives the allowable load region R, which shows a combination of the allowable X-direction load, allowable Y-direction load, and allowable shear load in the in-plane direction of the composite laminate in three dimensions, based on the allowable strain value corresponding to the orientation direction of the reinforcing fibers (step S102).

In this embodiment, the derivation unit 104 derives the allowable load region R for each of the 0° rich laminate, the pseudo-isotropic laminate, and the 90° rich laminate when they are configured in a basic laminate configuration. In other words, the derivation unit 104 derives the allowable load region R for each basic laminate configuration for each orientation ratio of the composite laminate.

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

If the applied load is outside the allowable load region R for all orientation ratios, the derivation unit 104 sets the magnification of the basic stacking configuration for each orientation ratio to a larger value (e.g., from 1x to 2x) (step S104). The derivation unit 104 then again derives the allowable load region R for each of the basic stacking configurations for each orientation ratio after changing the magnification (step S102).

The determination unit 106 determines whether the applied load is included in the re-derived allowable load region R (step S103). In this way, if the applied load is not included in any of the allowable load regions R of the basic stacking configurations of each orientation ratio, the magnification of the basic stacking configuration is multiplied by an integer, and the magnification is increased until the applied load is included in the allowable load region R. However, this is not limited to this, and the magnification of the basic stacking configuration may be a decimal or a fraction. In addition, the derivation unit 104 may derive the magnification of the basic stacking configuration of each orientation ratio such that the applied load is included in the allowable load region R, and derive the allowable load region R corresponding to each magnification of the basic stacking configuration of each orientation ratio.

On the other hand, if the applied load falls within the allowable load region R of any of the orientation ratios, the selection unit 108 selects the orientation ratio that minimizes the magnification of the basic stacking configuration (step S105). Here, if there are multiple orientation ratios that minimize the magnification of the basic stacking configuration, one orientation ratio may be selected from the multiple orientation ratios based on the priority of the orientation ratios that is set in advance. This makes it possible to select the stacking configuration of the lightest composite laminate that can withstand the applied load.

For example, in the case of FIG. 5, the orientation ratio of the 0° rich laminate is selected when the magnification of the basic laminate structure is doubled. Also, in the case of FIG. 6, the orientation ratio of the 90° rich laminate is selected when the magnification of the basic laminate structure is doubled.

In this way, according to the composite laminate design method of this embodiment, an allowable load region R is derived that shows a combination of allowable X-direction load, allowable Y-direction load, and allowable shear load in the in-plane directions of the composite laminate in three dimensions. In addition, 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 sufficient against the in-plane loads (Nx, Ny, Nxy) applied to the composite laminate for a large number of load cases. Therefore, it is possible to efficiently optimize the thickness, orientation ratio, orientation direction, etc. of the composite laminate, and it is possible to improve the design efficiency of the composite laminate.

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

In the above embodiment, the orientation ratio of the composite laminate is described as 0° rich laminate, pseudo-isotropic laminate, and 90° rich laminate. However, the orientation ratio of the composite laminate is not limited to this, and may be, for example, +45° rich laminate or -45° rich laminate. Here, the basic stacking structure of the +45° rich laminate is, for example, a stacking structure of 5 plies, including one layer (1 ply) with a 0° orientation direction, two layers (2 plies) with a +45° orientation direction, one layer (1 ply) with a -45° orientation direction, and one layer (1 ply) with a 90° orientation direction. In addition, the basic stacking structure of the -45° rich laminate is, for example, a stacking structure of 5 plies, including one layer (1 ply) with a 0° orientation direction, one layer (1 ply) with a +45° orientation direction, two layers (2 plies) with a -45° orientation direction, and one layer (1 ply) with a 90° orientation direction. The basic stacking configuration of a pseudo-isotropic laminate is, for example, one layer (1 ply) with an orientation direction of 0°, one layer (1 ply) with an orientation direction of +45°, one layer (1 ply) with an orientation direction of -45°, and one layer (1 ply) with an orientation direction of 90°, for a total of four plies.

In the above embodiment, an example has been described in which the design device 100 includes the selection unit 108. However, the selection unit 108 is not a required component, and for example, the design device 100 does not need to include the selection unit 108.

R Allowable load region 100 Design device 102 Database 104 Derivation unit 106 Determination unit 108 Selection unit

Claims (3)

a step of deriving an allowable load region that is three-dimensionally expressed by a combination of an allowable X-direction load in a predetermined x-direction in a plane perpendicular to a stacking direction in which each layer of the composite laminate is stacked, an allowable Y-direction load in a y-direction perpendicular to the x-direction in a plane perpendicular to the stacking direction in which each layer of the composite laminate is stacked, and an allowable shear load in a shear direction in a plane perpendicular to the stacking direction in which each layer of the composite laminate is stacked, based on an allowable strain corresponding to the plurality of fiber orientation angles in a composite laminate having the plurality of fiber orientation angles;
determining whether an applied load acting on the composite laminate is included within the allowable load region;
A method for designing a composite laminate comprising: The allowable load region varies depending on the magnification of a basic lamination structure in which the number of layers of each fiber orientation angle is the minimum unit based on an orientation ratio, which is a ratio of the thickness of a composite layer of each fiber orientation angle to the thickness of the composite laminate, and the orientation ratio,
Selecting the orientation ratio at which the applied load is included in the allowable load region and the magnification of the basic stacking configuration is minimized,
A method for designing a composite laminate according to claim 1. An aircraft equipped with a composite laminate designed by the design method of claim 1 or 2.