Fuselage skin layering method considering falling adaptability

Abstract

The invention relates to the technical field of fuselage structure design, and discloses a fuselage skin layering method considering falling adaptability, which comprises the following steps: S1, establishing a fuselage skin model and a ground model; s2, static constraint and falling adaptability constraint of the fuselage skin model are established; s3, the laying layer angles of the laying layer structure are set to be N1, the laying layer thicknesses are set to be N2, N1 is equal to N2, the N1 laying layer angles are formed by combining N1 numerical values selected from first discrete variables, every two adjacent laying layer angles are not equal, and the N2 laying layer thicknesses are formed by combining N2 numerical values selected from second discrete variables; s4, acquiring layering angles of a plurality of layering structures conforming to a multi-objective optimization algorithm; s5, calculating equivalent buckling factors for a plurality of paving layer structures conforming to the multi-objective optimization algorithm and the paving layer structures formed by the conventional paving layer angle paving layers, and selecting the paving layer angle of the maximum equivalent buckling factor; and S6, composite material layering is carried out according to the layering angle selected in the S5 to form the fuselage skin.

Description

Technical Field

[0001] The present invention relates to the technical field of fuselage structure design, and in particular to a fuselage skin laying method taking crashworthiness into consideration. Background Art

[0002] Composite materials, with their advantages such as high specific strength and stiffness, strong designability, excellent fatigue resistance, ease of large-scale integral forming, and easy maintainability, have become one of the most promising and popular technologies today. The low density and high performance of composite materials make them ideal for aircraft structures. The use of advanced composite materials in aircraft can reduce weight by 20%-30%, a significant advantage unattainable with traditional aviation aluminum alloys. Currently, the primary method for crashworthiness research is still to conduct crash tests on the entire aircraft or fuselage frame sections to determine the transient response, failure modes, and energy absorption capacity of key components during a crash, thereby assessing the crashworthiness of the aircraft structure. However, as an aircraft is a highly integrated structural system, the material properties of its components vary, and the connections between them are complex, making it difficult to ensure the consistency of experimental data. Furthermore, realistic crash conditions are difficult to accurately replicate in the laboratory, resulting in high experimental risks, long preparation times, and significant costs.

[0003] Therefore, there is an urgent need for a fuselage skin laying method that takes crashworthiness into consideration to solve the above problems. Summary of the Invention

[0004] The object of the present invention is to provide a fuselage skin layup method taking crashworthiness into consideration, which can quickly, accurately and at a low cost obtain a suitable composite material layup angle to improve the crashworthiness of the aircraft.

[0005] As conceived above, the technical solution adopted by the present invention is:

[0006] A method for laying out fuselage skin considering crashworthiness is provided, comprising the following steps:

[0007] S1. Establish the fuselage skin model and ground model in the finite element analysis software;

[0008] S2. Establishing static constraints and crashworthiness constraints for the fuselage skin model, wherein the static constraints include fixing one end of the fuselage skin model and applying a static load to the fuselage skin model, and the crashworthiness constraints include applying an initial impact velocity to the fuselage skin model;

[0009] S3. Establishing multiple layer structures of the fuselage skin model, including:

[0010] S31, setting the ply structure to adopt symmetrical ply;

[0011] S32, set the ply angle of the ply structure as N1, N1 ply angles are formed by selecting N1 value combinations in the input first discrete variable, adjacent two ply angles are not equal, the first discrete variable at least includes a group of conventional ply angles, the conventional ply angle is 0°, 45°, -45° and 90°;

[0012] S33, set the ply thickness of the ply structure as N2, N2 ply thicknesses are formed by selecting N2 value combinations in the input second discrete variable, N1=N2;

[0013] S4, obtain the ply structure meeting the multi-objective optimization algorithm, the optimization objective of the multi-objective optimization algorithm is to minimize the fuselage skin mass, minimize the maximum tensile strain, maximize the minimum compressive strain and minimize the maximum vibration comprehensive acceleration, the maximum vibration comprehensive acceleration is less than 100g;

[0014] S5, calculate the equivalent buckling factor of the ply structure meeting the multi-objective optimization algorithm and the ply structure formed by the conventional ply angle, and select the ply angle of the ply structure with the maximum equivalent buckling factor;

[0015] S6, according to the ply angle selected in S5, perform composite ply layering.

[0016] Optionally, before S5, it also includes S50, the ply thickness of the ply structure meeting the multi-objective optimization algorithm and the ply structure formed by the conventional ply angle is optimized, so that the sum of N2 values of the ply thickness of the ply structure meeting the multi-objective optimization algorithm and the ply structure formed by the conventional ply angle is equal.

[0017] Optionally, the first discrete variable includes a plurality of groups of ply angles, and the number of values included in each group of ply angles can be unequal.

[0018] Optionally, at least one group of ply angles includes M1 pairs of a1 and b1, and the sum of the absolute values of a1 and b1 is 90°, M1≥1, at least one group of ply angles includes M2 pairs of a2 and b2, and the sum of the absolute values of a2 and b2 is 80°, M2≥1.

[0019] Optionally, the second discrete variable is set as a group, and the second discrete variable includes 1, 2 and 3.

[0020] Optionally, the ply angle is 10, and the ply thickness is 10.

[0021] Optionally, the structure of the fuselage skin model is cylindrical, the outer diameter of the fuselage skin model is 3400mm, and the length is 3000mm.

[0022] The static load includes a first working condition and a second working condition, the internal pressure of the first working condition is 0.06895MPa, the bending moment is 7.611e9N·mm, the internal pressure of the second working condition is 0.1379MPa, and the impact initial speed is 7000mm / s.

[0023] Optionally, the composite layer adopts a DD layer.

[0024] The beneficial effects of the present application are:

[0025] The present invention proposes a fuselage skin layer laying method considering crashworthiness, comprising the following steps: S1, establishing a fuselage skin model and a ground model in finite element analysis software; S2, establishing static constraints and crashworthiness constraints of the fuselage skin model, wherein the static constraints include fixing constraints on one end of the fuselage skin model and applying a static load to the fuselage skin model, and the crashworthiness constraints include applying an initial impact velocity to the fuselage skin model; S3, establishing multiple layer laying structures of the fuselage skin model, comprising the following steps: S31, setting the layer laying structure to adopt symmetrical layer laying; S32, setting the layer laying structure to have N1 layer laying angles, wherein the N1 layer laying angles are formed by selecting N1 numerical combinations from the first discrete variable input, and the adjacent layer laying angles are not equal, and the first discrete variable includes at least one set of conventional layer laying angles. Conventional layup angles are 0°, 45°, -45° and 90°; S33, set the layup thickness of the layup structure to N2, and the N2 layup thicknesses are formed by selecting N2 numerical combinations from the second discrete variable of the input, N1=N2; S4, obtain the layup structure that meets the multi-objective optimization algorithm, and the optimization objectives of the multi-objective optimization algorithm are minimizing the fuselage skin mass, minimizing the maximum tensile strain, maximizing the minimum compressive strain and minimizing the maximum vibration integrated acceleration, and the maximum vibration integrated acceleration is less than 100g; S5, calculate the equivalent buckling factor for the layup structure that meets the multi-objective optimization algorithm and the layup structure formed by conventional layup angles, and select the layup angle of the layup structure with the largest equivalent buckling factor; S6, perform composite layup according to the layup angle selected in S5. Using a crashworthiness-considered fuselage skin layup method, multiple layup structures can be set up in finite element analysis software and their crashworthiness simulations performed on these layup structures to obtain the layup structure with the smallest fuselage skin mass, the smallest maximum tensile strain, the largest minimum compressive strain, and the smallest maximum vibration integrated acceleration, with the maximum vibration integrated acceleration being less than 100g. These selected layup structures and layup structures formed by conventional layup angles are then simulated with equivalent buckling factors, ultimately selecting the layup angle with the largest equivalent buckling factor. Comparing these selected layup structures with conventional layup structures further confirms that the equivalent buckling factor of the selected layup structure is greater than that of the conventional layup structure, ultimately determining the optimal layup angle. This crashworthiness-considered fuselage skin layup method eliminates the need for crash testing and instead uses simulation analysis to quickly, accurately, and cost-effectively determine the appropriate composite layup angle, thereby improving the aircraft's crashworthiness. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] Figure 1 is a flow chart of a fuselage skin layup method considering crashworthiness performance provided by an embodiment of the present invention;

[0027] Figure 2Fig. 1 is a schematic diagram of a fuselage skin model and a ground model provided by an embodiment of the present application. DETAILED DESCRIPTION

[0028] In order to make the technical problems solved by the present application, the technical solutions adopted and the technical effects reached more clear, the technical solutions of the present application will be further described below in conjunction with the accompanying drawings and specific embodiments. It can be understood that the specific embodiments described here are only used to explain the present application, but not limit the present application. In addition, it should be noted that, for the convenience of description, only the parts related to the present application are shown in the drawings, not all.

[0029] In the description of the present application, unless otherwise explicitly specified and limited, the terms "connected", "connected", "fixed" should be understood broadly, for example, it can be fixedly connected, or it can be detachably connected, or it can be integrated; it can be mechanically connected, or it can be electrically connected; it can be directly connected, or it can be indirectly connected through an intermediate medium; it can be the internal communication of two elements or the interaction relationship between two elements. For those skilled in the art, the specific meaning of the above terms in the present application can be understood according to the specific circumstances.

[0030] In the present application, unless otherwise explicitly specified and limited, the first feature "on" or "under" the second feature can include that the first and second features are in direct contact, or that the first and second features are not in direct contact but are in contact through another feature between them. Moreover, the first feature "on", "above" and "above" the second feature includes that the first feature is directly above and obliquely above the second feature, or only indicates that the horizontal height of the first feature is higher than that of the second feature. The first feature "under", "below" and "below" the second feature includes that the first feature is directly below and obliquely below the second feature, or only indicates that the horizontal height of the first feature is less than that of the second feature.

[0031] In the description of the present embodiment, the terms "up", "down", "left", "right" and other orientation or position relationship are based on the orientation or position relationship shown in the drawings, which is only for the convenience of description and simplification of operation, and does not indicate or imply that the device or element referred to must have a particular orientation, be constructed and operated in a particular orientation, therefore cannot be understood as a limitation of the present application. In addition, the terms "first", "second" are only used to distinguish in description, and have no special meaning.

[0032] The technical solutions of the present application will be further described below in conjunction with the accompanying drawings and specific embodiments.

[0033] As Figure 1 and Figure 2As shown, the embodiment provides a fuselage skin layering method considering crashworthiness performance, comprising the following steps: S1, establishing a fuselage skin model and a ground model in a finite element analysis software; S2, establishing static force constraints and crashworthiness constraints of the fuselage skin model, the static force constraints including fixing constraints on one end of the fuselage skin model and applying static force load to the fuselage skin model, and the crashworthiness constraints including applying impact initial velocity to the fuselage skin model; S3, establishing a plurality of layering structures of the fuselage skin model, comprising: S31, setting that the layering structure adopts symmetric layering; S32, setting that the layering angle of the layering structure is N1, N1 layering angles are formed by selecting N1 numerical combinations in an input first discrete variable, adjacent two layering angles are not equal, the first discrete variable at least includes a group of conventional layering angles, and the conventional layering angles are 0°, 45°, -45° and 90°; S33, setting that the layering thickness of the layering structure is N2, N2 layering thicknesses are formed by selecting N2 numerical combinations in an input second discrete variable, and N1=N2; S4, obtaining a layering structure meeting a multi-objective optimization algorithm, and optimization objectives of the multi-objective optimization algorithm are minimization of fuselage skin mass, minimization of maximum tensile strain, maximization of minimum compressive strain and minimization of maximum vibration comprehensive acceleration, and the maximum vibration comprehensive acceleration is less than 100g; S5, calculating equivalent buckling factors of the plurality of layering structures meeting the multi-objective optimization algorithm and layering structures formed by the conventional layering angles, and selecting layering angles of a layering structure with the maximum equivalent buckling factor; S6, performing composite layering according to the layering angles selected in S5. The fuselage skin layering method considering crashworthiness performance provided by the embodiment sets a plurality of layering structures in the finite element analysis software, and simulates the crashworthiness performance of the plurality of layering structures, so as to obtain a layering structure with the minimum fuselage skin mass, the minimum maximum tensile strain, the maximum minimum compressive strain and the minimum maximum vibration comprehensive acceleration, and the maximum vibration comprehensive acceleration is less than 100g, and further simulate equivalent buckling factors of the layering structures screened out and layering structures formed by the conventional layering angles, finally select layering angles of a layering structure with the maximum equivalent buckling factor, and compare the layering structures screened out with the conventional layering structures, which can further prove that the equivalent buckling factor of the layering structure screened out is greater than the equivalent buckling factor of the conventional layering structure, so as to finally determine the optimal layering angle. The fuselage skin layering method considering crashworthiness performance can quickly, accurately and low-cost obtain suitable composite layering angles through simulation analysis without crash test, so as to improve the crashworthiness performance of the aircraft.

[0034] Optionally, S5 is followed by S50, which optimizes the ply thickness of the ply structure conforming to the multi-objective optimization algorithm and the ply structure formed by the conventional ply angle, so that the sum of N2 values of the ply thickness of the ply structure conforming to the multi-objective optimization algorithm and the ply structure formed by the conventional ply angle is equal. The ply structure preliminarily screened by the multi-objective optimization algorithm considers multiple optimization objectives such as minimization of the aircraft skin mass, minimization of the maximum tensile strain, maximization of the minimum compressive strain, and minimization of the maximum vibration comprehensive acceleration. Therefore, the number of ply layers of the screened ply structure is not necessarily consistent. To ensure the accuracy of the comparison of the equivalent buckling factors, the number of ply layers is optimized so that the number of ply layers of each ply structure is the same when the equivalent buckling factors are compared.

[0035] Optionally, the first discrete variable includes multiple sets of ply angles, and the number of values included in each set of ply angles can be unequal. For example, the first set of ply angles is 0°, -45°, 45°, and 90°. The second set of ply angles is 0°, -15°, -30°, -45°, -60°, -75°, 15°, 30°, 45°, 60°, 75°, and 90°. That is, there are multiple sets of ply angles, and the number of values included in each set of ply angles is not necessarily equal. However, N1 ply angles are screened for each set of ply angles to form a combination, and the adjacent two ply angles in the combination are not equal. For example, N1 = 10, and the following combinations can be selected from the first set of ply angles: [0° / -45° / 45° / 90° / 0° / -45° / 45° / 90° / 0° / -45°], [90° / -45° / 45° / 0° / 90° / -45° / 45° / 0° / 90° / -45°], and the like. Each set of ply angles can generate a series of ply structures, and multiple sets of ply angles can generate multiple series of ply structures to expand the sample size of the ply structure simulation.

[0036] Optionally, to further expand the value range of the discrete variable of the ply angle, at least one set of ply angles includes M1 pairs of a1 and b1, and the sum of the absolute values of a1 and b1 is 90°, and M1 ≥ 1. For example, in a set of ply angles, a1 = -22°, 0°, 15°, 17°, 30°, 45°, and b1 = -68°, 90°, 75°, 73°, 60°, 45°. In addition, at least one set of ply angles includes M2 pairs of a2 and b2, and the sum of the absolute values of a2 and b2 is 80°, and M2 ≥ 1. For example, in a set of ply angles, a2 = -13°, 13°, and b2 = 67°, -67°.

[0037] Optionally, a set of second discrete variables is provided, and the set of second discrete variables includes 1, 2, and 3. Wherein 1 refers to the thickness of one layer, i.e. 0.127 mm, 2 refers to the thickness of two layers, i.e. 0.254 mm, and 3 refers to the thickness of three layers, i.e. 0.381 mm. N2 layer thickness combinations are screened out from the set of second discrete variables. For example, N2 = 10, and the following combinations can be selected from the set of layer thicknesses: [2 / 2 / 1 / 2 / 2 / 2 / 3 / 2 / 1 / 1], [1 / 1 / 1 / 2 / 1 / 2 / 2 / 2 / 1 / 1], and the like.

[0038] In the establishment of the fuselage skin model, only the skin structure of the fuselage is retained, and the static force constraint and the crashworthiness constraint borne by the actual fuselage structure are applied, so the thickness of the fuselage skin model is particularly increased compared to the actual skin structure to ensure that the simulation analysis effect is equivalent to that of the actual skin structure. In this embodiment, the layer thickness is preferably 10, and the layer angle is 10, i.e. the maximum layer thickness of the fuselage skin model is 60 layers, i.e. 7.62 mm, and the minimum layer thickness of the fuselage skin model is 20 layers, i.e. 2.54 mm.

[0039] In this embodiment, the structure of the fuselage skin model is selected to be a straight section of the fuselage, i.e. the fuselage skin model is cylindrical, the outer diameter of the fuselage skin model is 3400 mm, and the length is 3000 mm. The static force load applied to the fuselage skin model includes a first working condition and a second working condition, the internal pressure of the first working condition is 0.06895 MPa, and the bending moment is 7.611e9 N·mm, and the internal pressure of the second working condition is 0.1379 MPa, to realize the static force two-working condition analysis of the fuselage skin model. In addition, the initial impact velocity applied to the fuselage skin model is 7000 mm / s, to realize the crashworthiness analysis of the fuselage skin model.

[0040] In summary, the specific steps of the fuselage skin layering method considering the crashworthiness provided in this embodiment are as follows:

[0041] Step one, establish the fuselage skin model and the ground model in the finite element analysis software, and set the ground model to be a rigid surface.

[0042] Step two, establish the static force constraint of the fuselage skin model, the static force constraint includes the fixed constraint to one end of the fuselage skin model and the static force load applied to the fuselage skin model, the static force load includes a first working condition and a second working condition, the internal pressure under the first working condition is 0.06895 MPa, and the bending moment is 7.611e9 N·mm, and the internal pressure under the second working condition is 0.1379 MPa.

[0043] Step three, establish the crashworthiness constraint of the fuselage skin model, the crashworthiness constraint is to apply an initial impact velocity of the impact ground model to the fuselage skin model, and the initial impact velocity is 7000 mm / s.

[0044] Step four, a plurality of layup structures of the fuselage skin model are established, the layup structures change due to different layup angles and layup thicknesses, and thus the number of the layup angles and the layup thicknesses needs to be set, and the first discrete variable corresponding to the layup angle and the second discrete variable corresponding to the layup thickness.

[0045] In this embodiment, the layup angle is set to 10. The first discrete variable customized by the user can include a plurality of groups of layup angles, which are respectively a first group of layup angles 0°, -45°, 45°, 90°; a second group of layup angles 0°, -15°, -30°, -45°, -60°, -75°, 15°, 30°, 45°, 60°, 75°, 90°; a third group of layup angles -13°, 13°, 67°, -67°; a fourth group of layup angles -13°, 13°, 77°, -77°; a fifth group of layup angles -22.5°, 22.5°, 67.5°, -67.5°; and other groups of layup angles can also be included. The finite element analysis software selects 10 layup angles in each group of layup angles for combination, and ensures that the two adjacent layup angles are not equal, and the following combinations can be formed, [0° / -45° / 45° / 90° / 0° / -45° / 45° / 90° / 0° / -45°], [-30° / 45° / 15° / 90° / 0° / -45° / 60° / 75° / 0° / -45°], [22.5° / -22.5° / 67.5° / -67.5° / 22.5° / -22.5° / 67.5° / -67.5° / 22.5° / -22.5°], and the like.

[0046] In this embodiment, the layup thickness is set to 10. The second discrete variable customized by the user includes a group of layup angles 1, 2, 3. The finite element analysis software selects 10 layup thicknesses in the group of layup thicknesses for combination, and the following combinations can be formed, [2 / 2 / 2 / 2 / 1 / 2 / 3 / 2 / 1 / 1], [3 / 1 / 2 / 1 / 1 / 2 / 2 / 2 / 1 / 1], [1 / 2 / 3 / 2 / 1 / 1 / 1 / 2 / 1 / 1], and the like.

[0047] Step five, the finite element analysis software performs a multi-objective optimization algorithm analysis, and obtains a group of layup structures with the minimum fuselage skin mass, the minimum maximum tensile strain, the maximum minimum compressive strain, and the minimum maximum vibration comprehensive acceleration less than 100g through the Pareto frontier optimal solution, as shown in the following table:

[0048]

[0049] In the above table, 1 in the ply thickness value means the thickness of one layer, i.e. 0.127 mm, 2 in the ply thickness value means the thickness of two layers, i.e. 0.254 mm, and 3 in the ply thickness value means the thickness of three layers, i.e. 0.381 mm. In addition, the maximum vibration comprehensive acceleration in the optimization target means the square root of the sum of the squares of the maximum accelerations in the X, Y and Z directions when the aircraft skin model impacts the ground model, i.e. the maximum vibration comprehensive acceleration. In the crashworthiness analysis, the maximum value of the square of the acceleration in the Y direction is much greater than the maximum values of the squares of the accelerations in the X and Z directions, so only the maximum and minimum accelerations in the Y direction are listed in the above table, and the larger one of the square of the maximum acceleration and the square of the minimum acceleration in the Y direction is selected. In addition, to ensure that the crashworthiness of the ply structure formed by the optimized ply angle is better than that of the ply structure formed by the conventional ply angle, the selected ply structure must be further subjected to buckling analysis with the ply structure formed by the conventional ply angle. The specific contents are shown in the following table:

[0050]

[0051] Step six, the ply structure with the maximum equivalent buckling factor is selected, i.e. the ply angle is [22.5° / -22.5° / 67.5° / -67.5° / 22.5° / -22.5° / 67.5° / -67.5° / 22.5° / -22.5°]. It can be seen from the table that the equivalent buckling factor of the ply structure is also greater than the equivalent buckling factor of the conventional ply structure, i.e. greater than 1.

[0052] Step seven, the composite ply is formed according to the ply angle [22.5° / -22.5° / 67.5° / -67.5° / 22.5° / -22.5° / 67.5° / -67.5° / 22.5° / -22.5°] to form the fuselage skin.

[0053] As can be seen from the above specific embodiments, the fuselage skin ply method considering the crashworthiness can quickly, accurately and at low cost obtain suitable composite ply angles, and can improve the crashworthiness of the aircraft. The new ply obtained can not only improve the dynamic performance of the fuselage structure, but also has a 10% weight reduction potential and better buckling performance under the condition of meeting the static mechanical constraints.

[0054] Optionally, the composite ply method is a DD (Double-Double) ply method.

[0055] The above embodiments only illustrate the basic principles and characteristics of the present application, and the present application is not limited to the above embodiments. Without departing from the spirit and scope of the present application, various changes and modifications can be made to the present application, and these changes and modifications all fall within the scope of the present application. The scope of protection of the present application is defined by the appended claims and their equivalents.

Claims

1. A fuselage skin layup method considering crashworthiness, characterized in that: The following steps are involved: S1. Establish the fuselage skin model and ground model in the finite element analysis software; S2. Establishing static constraints and crashworthiness constraints for the fuselage skin model, wherein the static constraints include fixing one end of the fuselage skin model and applying a static load to the fuselage skin model, and the crashworthiness constraints include applying an initial impact velocity to the fuselage skin model; S3. Establishing multiple layer structures of the fuselage skin model, including: S31, setting the ply structure to adopt symmetrical ply; S32: Setting the number of ply angles of the ply structure to N1, where the N1 ply angles are formed by selecting N1 numerical combinations from a first discrete variable input, and two adjacent ply angles are unequal, wherein the first discrete variable includes at least one set of conventional ply angles, wherein the conventional ply angles are 0°, 45°, -45°, and 90°; S33, setting the ply thickness of the ply structure to N2, wherein the N2 ply thicknesses are formed by selecting N2 value combinations from the input second discrete variable, where N1=N2; S4. Obtaining the laminate structure that complies with a multi-objective optimization algorithm, where the optimization objectives of the multi-objective optimization algorithm are minimizing the fuselage skin mass, minimizing the maximum tensile strain, maximizing the minimum compressive strain, and minimizing the maximum integrated vibration acceleration, wherein the maximum integrated vibration acceleration is less than 100 g; S5. Calculate the equivalent buckling factor for both the layup structure conforming to the multi-objective optimization algorithm and the layup structure formed by conventional layup angles, and select the layup angle of the layup structure with the largest equivalent buckling factor. S6. Laying down the composite material according to the laying angle selected in S5.

2. The fuselage skin laying method considering crashworthiness according to claim 1, characterized in that: Before S5, S50 is also included, which optimizes the ply thickness of the ply structure that meets the multi-objective optimization algorithm and the ply structure formed by conventional ply angle ply laying, so that the sum of N2 values ​​of the ply thickness of the multiple ply structures that meet the multi-objective optimization algorithm and the ply structures formed by conventional ply angle ply laying is equal.

3. The fuselage skin laying method considering crashworthiness according to claim 1, characterized in that: The first discrete variable includes multiple groups of ply angles, and the number of values ​​included in each group of ply angles may be unequal.

4. The fuselage skin laying method considering crashworthiness according to claim 3 is characterized in that: At least one group of the ply angles includes M1 for a1 and b1, and the sum of the absolute value of a1 and the absolute value of b1 is 90°, M1≥1, and at least one group of the ply angles includes M2 for a2 and b2, and the sum of the absolute value of a2 and the absolute value of b2 is 80°, M2≥1.

5. The fuselage skin laying method considering crashworthiness according to claim 1, characterized in that: The second discrete variables are set in a group, and the second discrete variables include 1, 2 and 3.

6. The fuselage skin laying method considering crashworthiness according to claim 1, characterized in that: The ply angle is 10 and the ply thickness is 10.

7. The fuselage skin laying method considering crashworthiness according to claim 1, characterized in that: The structure of the fuselage skin model is cylindrical, and the outer diameter of the fuselage skin model is 3400 mm and the length is 3000 mm; The static load includes a first working condition and a second working condition. Under the first working condition, the internal pressure is 0.06895 MPa and the bending moment is 7.611e9 N·mm. Under the second working condition, the internal pressure is 0.1379 MPa, and the initial impact velocity is 7000 mm / s.

8. The fuselage skin laying method considering crashworthiness according to claim 1, characterized in that: The composite material layer is DD layer.

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