A method for laying carbon fiber DD in a composite material fuselage stringer structure

Abstract

This invention discloses a carbon fiber double-double layup method for a composite material fuselage stringer structure. The stringer is fixed to the inside of the aircraft fuselage cylinder and adopts a carbon fiber double-double layup structure. First, the two angle parameters φ and ψ of the double-double layup are set within the range of 0º-180º and different combinations are made. Then, under a preset bending moment and a preset number of layers, the optimal combination of angle parameters (±φ, ±ψ) is determined by strength calculation. Finally, the double-double layup is repeated in a fixed cycle along the stacking direction according to the determined combination of angle parameters (±φ, ±ψ). This invention can improve the design stiffness and strength of the stringer with the same material usage, or reduce the amount of composite material used with the same stiffness or strength, resulting in high carbon fiber strength utilization and high material utilization rate.

Description

Technical Field

[0001] This invention relates to fuselage stringer structures, and in particular to a method for laying a carbon fiber composite fuselage stringer structure based on carbon fiber double-double (hereinafter referred to as DD) laying. Background Technology

[0002] Taking advantage of the constant Chua's modulus of laminates, traditional laminate and girder structures can be replaced by four combinations of two angles, ±φ and ±ψ, namely the Double-Double (±φ, ±ψ) layup structure, or DD for short. φ and ψ represent the angle parameters of any two adjacent fiber layers and the angle parameters of the next two fiber layers in a Double-Double layup, respectively.

[0003] An aircraft fuselage is a slender cylindrical structure. Relative to its length and diameter, the thickness of the fuselage skin is very small. Therefore, the fuselage's rigidity is provided by the stringers structure within the fuselage. The stringers primarily bear the bending moments caused by various loads on the fuselage, transmitted through the skin, and a small amount of axial load. The bending moments mainly include axial bending moments along the fuselage's central axis and circumferential bending moments. The skin primarily bears axial and circumferential membrane stresses. According to the conventional design loads of the fuselage, the dominant stress is caused by the difference between the inner and outer surfaces of the fuselage skin.

[0004] Traditional stringers are laid at fixed angles such as (0, ±45, 90), so the fibers parallel to the bending moment direction do not play a significant role when subjected to bending moments. Summary of the Invention

[0005] To address the problems existing in the background art, this invention proposes a carbon fiber DD laying method for a composite material fuselage stringer structure and the same composite material stringer structure. Based on the load conditions on the fuselage, the stringer is designed to have higher stiffness while maintaining the same material usage. Alternatively, compared with traditional stringer structures, the amount of composite material used is significantly reduced while maintaining the same stiffness.

[0006] The technical solution of the present invention is as follows: The stringer is fixed to the inside of the aircraft fuselage cylinder. The stringer adopts a carbon fiber double-double layup structure to form a laminate. S1. First, set the two angle parameters φ and ψ of the carbon fiber Double-Double layup structure within the range of 0º-180º and make different combinations. S2. Then, under the preset bending moment and preset number of layers, the strength calculation method is used to perform optimization calculations to determine the optimal combination of angle parameters (±φ, ±ψ). S3. Finally, according to the determined combination of angle parameters (±φ, ±ψ), the layers are laid repeatedly in a fixed cycle along the stacking direction.

[0007] The strength calculation method adopts the Tsai-Wu strength theory calculation method.

[0008] In step S2, a stiffness matrix is ​​constructed using two angle parameters φ and ψ, and then the following objective function is set based on the stiffness matrix: Min [σ] = [M] / [D] [D]=[D ij ] Where [σ] represents the stress of the laminate, [M] is the bending moment load of the fuselage cylinder, [D] is the bending stiffness matrix of the laminate, which is directly related to the fiber laying angle, i.e., the bending stiffness matrix of the stringer; Min represents the minimization function; φ and ψ represent the angle parameters of two adjacent layers of double-double fiber layup, respectively; D ij This represents the stiffness component element values ​​in the i-th row and j-th column of the stiffness matrix, where i and j represent the i-th row and j-th column of the stiffness matrix, respectively.

[0009] Then, with the goal of minimizing stress, the strength calculation method is used to obtain the optimal angle parameters φ and ψ, forming the optimal angle parameter combination (±φ, ±ψ).

[0010] In step S3, the two angle parameters φ and ψ of the Double-Double ply are set in four consecutive layers in four ways: +φ, -φ, +ψ, -ψ or -φ, +φ, -ψ, +ψ or +ψ, -ψ, +φ, -φ or -ψ, +ψ, -φ, +φ, and then the basic ply unit is laid in the same way along the stacking direction.

[0011] Under the action of bending moment on the fuselage cylinder with the central axis of the fuselage cylinder as the reference coordinate, the bending moment is a circumferential load and an axial load with a ratio of 2:1. The angle parameter combination φ and ψ of the carbon fiber DD layup structure is obtained by optimization calculation.

[0012] The cross-section of the stringer is a trapezoidal structure with wingspan sections on both sides, which are fixed to the inside of the fuselage with rivets.

[0013] The length of the stringer is along the central axis of the cylindrical fuselage.

[0014] The spandrel section is fixed to the inner wall panel of the fuselage cylinder with rivets.

[0015] The fuselage cylinder has four long girders arranged circumferentially on its inner circumferential surface, and the four long girders are evenly spaced circumferentially.

[0016] The design of this invention is based on various loads on the fuselage, bending moments transmitted through the skin, a small amount of axial load, and stress caused by the difference between the inner and outer surfaces of the skin. If other loads are taken into account, this design can also obtain a new DD angle by modifying the DD angle.

[0017] The stringer mechanism of the DD laying structure of this invention has φ and ψ angles that can be any combination of two angles, so the design is flexible and the required DD angle can be designed according to the bending moment load, reducing a lot of carbon fiber material that is not used. Therefore, the carbon fiber strength utilization rate is high and the material utilization rate is high.

[0018] The beneficial effects of this invention are: This invention proposes a DD composite material stringer structure and its laying method, which can design the stringer to have higher stiffness under the same material usage, based on the load conditions on the fuselage.

[0019] Compared with traditional stringer structures, stringers using the DD laying method of this invention have a significantly reduced amount of composite materials while maintaining the same stiffness, resulting in a substantial reduction in fuselage weight. This can reduce flight energy consumption and increase payload. Attached Figure Description

[0020] Figure 1 These are the front view and side view of the structure of the present invention; Figure 2 This is a cross-sectional structural diagram of the stringer; Figure 3 This is a schematic diagram of the DD laying angle in an embodiment; Figure 4 A schematic diagram of the 14-story laying angle for a traditional long truss (0, ±45, 90). Figure 5 This is a schematic diagram of the laying angles (±50°, ±80°) of the 14-layer DD angle of the long stringer in the example.

[0021] In the diagram, 1 represents the stringer, 2 represents the fuselage, and 3 represents the rivet. Detailed Implementation

[0022] The specific embodiments are described below with reference to the accompanying drawings.

[0023] like Figure 1 As shown, four long girders 1 are arranged circumferentially on the inner circumferential surface of the fuselage cylinder 2, and the four long girders 1 are evenly distributed at intervals along the circumferential direction.

[0024] Each stringer 1 is fixed to the inside of the cylindrical fuselage 2 of the aircraft. The length of the stringer 1 is along the central axis of the cylindrical fuselage 2, and the cross-section of the stringer 1 is a trapezoidal structure, as shown in the attached figure. Figure 2As shown, the trapezoidal structure has wingspans extending outwards in parallel on both sides.

[0025] The specific implementation details are attached. Figure 1 As shown, the span of the stringer 1 is fixed to the inner wall of the fuselage 2 by rivets 3.

[0026] The stringer 1 adopts a carbon fiber double-double ply structure. In the carbon fiber double-double ply structure, the two angle parameters φ and ψ of the double-double ply are laid in layers in the following manner: they are laid in layers in sequence, with four consecutive layers of +φ, -φ, +ψ, -ψ or -φ, +φ, -ψ, +ψ or +ψ, -ψ, +φ, -φ or -ψ, +ψ, -φ, +φ forming the basic ply unit. Then, the basic ply unit is laid in sequence along the ply direction.

[0027] The laying process is as follows: S1. First, set the absolute values ​​of the two angle parameters φ and ψ of the carbon fiber Double-Double layup structure in the range of 0º-180º and make different combinations. S2. Then, under the preset bending moment and preset number of layers, the optimal combination of angle parameters (±φ, ±ψ) is determined by using the strength calculation method.

[0028] Taking the stress calculation formula of a laminated plate under bending moment as an example, a stiffness matrix is ​​constructed using two angular parameters φ and ψ, and then the following objective function is set based on the stiffness matrix: Min T=W / G Min [σ] = [M] / [D] [D]= [D ij ] Where [σ] represents the stress matrix of the laminate, [M] is the bending moment load of the fuselage cylinder 2, [D] is the bending stiffness matrix of the laminate, which is directly related to the fiber laying angle, i.e., the bending stiffness matrix of the stringer 1, which is related to the fiber laying angle of each layer; Min represents the minimization function; φ and ψ represent the angle parameters of two adjacent layers of double-double fiber laying, respectively; D ij This represents the stiffness component element values ​​in the i-th row and j-th column of the stiffness matrix, where i and j represent the i-th row and j-th column of the stiffness matrix, respectively, as in D. ij This represents the stiffness element value of row i and column i.

[0029] Then, with the goal of minimizing stress, the strength calculation method is used to obtain the optimal angle parameters φ and ψ, forming the optimal angle parameter combination (±φ, ±ψ).

[0030] Then, with the goal of minimizing stress, the strength calculation method is used to obtain the optimal angle parameters φ and ψ, forming the optimal angle parameter combination (±φ, ±ψ).

[0031] In traditional fuselage panels, [D] is basically fixed. However, under the DD layup of this invention, based on the principle of minimizing [σ], we can find the φ and ψ with the maximum [D] in the combination of φ from 0 to 180 and ψ from 0 to 180, thereby obtaining a (±φ, ±ψ) layup that maximizes [D], that is, minimizes the stress [σ] in the laminate under the same load.

[0032] S3. Finally, according to the determined combination of angle parameters (±φ, ±ψ), the layers are laid repeatedly in a fixed cycle along any stacking direction from top to bottom or from top to bottom.

[0033] Specifically, the two angle parameters φ and ψ of the double-double ply are set in four consecutive layers in four ways: +φ, -φ, +ψ, -ψ or -φ, +φ, -ψ, +ψ or +ψ, -ψ, +φ, -φ or -ψ, +ψ, -φ, +φ. These layers form a basic plying unit, which is then repeated along the stacking direction.

[0034] The methods for obtaining the DD laying angle are compared below: For a stringer laid at a conventional (0, ±45, 90) angle, under the action of two mutually perpendicular fixed bending moments, in order to withstand this bending moment, the total number of carbon fiber layers r required can be calculated according to the Tsai-Wu strength theory of carbon fiber composites. This strength theory can also be the maximum principal stress strength theory of carbon fiber or other similar strength theories to calculate the number of layers.

[0035] As a comparative design, the present invention's DD laying method, under the same two bending moments, using the same number of layers r, and the same Tsai-Wu strength theory, can obtain four combinations (±φ, ±ψ) of two angles φ and ψ. By calculating the carbon fiber laying angle φ from 0º-180º and ψ from 0º-180º, the optimal φ and ψ parameter values ​​can be obtained to construct a DD combination (±φ, ±ψ). Its stringer has low stress under this laying angle and the same bending moment, that is, it has the strongest resistance to the above-set bending moment.

[0036] Comparative Example 1: When the fuselage 2 is subjected to loads such as internal and external pressure difference, ascent, and descent, the stringer 1 will transfer most of the load, especially the bending moment, to the stringer 1.

[0037] Preferably, the stringer is subjected to a bending moment acting on the fuselage cylinder 2 with the central axis of the fuselage cylinder 2 as the reference coordinate. The bending moment is a circumferential load and an axial load with a force ratio of 2:1.

[0038] like Figure 4 As shown, the conventional (0, ±45, 90) laid stringers have a thickness of 14 layers.

[0039] like Figure 3 As shown, this invention uses finite element three-dimensional modeling and calculation of the combined DD angles (±φ, ±ψ) from 0º to 180º to determine that the two angle parameter combinations φ and ψ are 50° and 80° respectively, and establishes the same 14-layer angle combination (±50°, ±80°).

[0040] Figure 3 It is a top view of the angles of each fiber in a set of DD-laid laminates, or it can be seen as the angle combination formed by extracting one fiber from each layup.

[0041] like Figure 5 As shown, four layers are laid continuously from top to bottom at angles of +80°, -80°, +50°, and -50° to form a basic laying unit. This basic laying unit is then repeatedly laid in a cyclical manner. The final solution can improve the bending moment strength by more than 18%.

[0042] Comparative Example 2: The advantages of using the DD laying method can also be demonstrated by calculation using another method: that is, under the same load, to obtain the same stiffness and strength with DD angles (±50, ±80), the number of layers can be reduced by 16.7% compared to using the traditional (0, ±45, 90) combined stringer. Taking this embodiment as an example, after the total number of layers is reduced from 14 layers to 12 layers, the same stiffness can be obtained, that is, the weight can be reduced by 2 layers of carbon fiber composite material.

[0043] It should be noted that the above load is only one example. After taking into account other bending moment loads, pressure differential loads, etc., the DD laying method can still find the corresponding optimal DD angle using the same method. Therefore, other DD angle laying methods are also within the scope of protection of this invention.

Claims

1. A method for laying carbon fiber DD in a composite material fuselage stringer structure, characterized in that: The stringer (1) is fixed to the inside of the fuselage cylinder (2) of the aircraft. The stringer (1) adopts a carbon fiber double-double layup structure to form a laminate. S1. First, set the two angle parameters φ and ψ of the carbon fiber Double-Double layup structure within the range of 0º-180º and make different combinations. S2. Then, under the preset bending moment and preset number of layers, the strength calculation method is used to perform optimization calculations to determine the optimal combination of angle parameters (±φ, ±ψ). S3. Finally, according to the determined combination of angle parameters (±φ, ±ψ), the layers are laid repeatedly in a fixed cycle along the stacking direction.

2. The carbon fiber DD laying method for a composite material fuselage stringer structure according to claim 1, characterized in that: The strength calculation method adopts the Tsai-Wu strength theory calculation method.

3. The carbon fiber DD laying method for a composite material fuselage stringer structure according to claim 1, characterized in that: In step S2, a stiffness matrix is ​​constructed using two angle parameters φ and ψ, and then the following objective function is set based on the stiffness matrix: Min [σ] = [M] / [D] [D]=[D ij ] Where [σ] represents the stress of the laminate, [M] is the bending moment load of the fuselage cylinder (2), [D] is the bending stiffness matrix of the laminate, which is directly related to the fiber laying angle, i.e., the bending stiffness matrix of the stringer (1); Min represents the minimization function; φ and ψ represent the angle parameters of two adjacent layers of double-double fiber laying, respectively; D ij This represents the stiffness component element values ​​in the i-th row and j-th column of the stiffness matrix, where i and j represent the i-th row and j-th column of the stiffness matrix, respectively. Then, with the goal of minimizing stress, the strength calculation method is used to obtain the optimal angle parameters φ and ψ, forming the optimal angle parameter combination (±φ, ±ψ).

4. The carbon fiber DD laying method for a composite material fuselage stringer structure according to claim 1, characterized in that: In step S3, the two angle parameters φ and ψ of the Double-Double ply are set in four consecutive layers in four ways: +φ, -φ, +ψ, -ψ or -φ, +φ, -ψ, +ψ or +ψ, -ψ, +φ, -φ or -ψ, +ψ, -φ, +φ, and then the basic ply unit is laid in the same way along the stacking direction.

5. The carbon fiber DD laying method for a composite material fuselage stringer structure according to claim 1, characterized in that: Under the action of bending moment on the fuselage cylinder (2) with the central axis of the fuselage cylinder (2) as the reference coordinate, the bending moment is a circumferential load and an axial load with a ratio of 2:

1. The angle parameter combination φ and ψ of the carbon fiber DD layup structure is obtained by optimization calculation.

6. The carbon fiber DD laying method for a composite material fuselage stringer structure according to claim 1, characterized in that: The cross section of the stringer (1) is a trapezoidal structure with wingspan on both sides. The wingspan is fixed to the inside of the fuselage with rivets (3).

7. The carbon fiber DD laying method for a composite material fuselage stringer structure according to claim 6, characterized in that: The length of the stringer (1) is along the central axis of the cylindrical fuselage (2).

8. The carbon fiber DD laying method for a composite material fuselage stringer structure according to claim 6, characterized in that: The span portion of the stringer (1) is fixed to the inner wall of the fuselage cylinder (2) with rivets (3).

9. The carbon fiber DD laying method for a composite material fuselage stringer structure according to claim 6, characterized in that: The inner circumferential surface of the fuselage cylinder (2) is provided with four long girders (1), which are evenly distributed at intervals along the circumferential direction.

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