A method for evaluating the weight of a composite wing panel structure

Abstract

The application belongs to the field of aircraft structure design, and particularly relates to a composite wing panel structure weight evaluation method. The method comprises the following steps: step one, calculating the average equivalent elastic modulus of the skin, stringers and spars in the tension or compression direction respectively; step two, obtaining a wing box section simplified model, and calculating the equivalent height of each section and the load bearing efficiency of each section according to the average equivalent elastic modulus and the wing box section simplified model; step three, calculating the equivalent thickness of the upper skin and the lower skin of each section according to the equivalent height of each section and the load bearing efficiency of each section; step four, calculating the design thickness of the upper skin and the lower skin of each section after removing the stringers according to the equivalent thickness of the upper skin and the lower skin of each section; step five, determining the final thickness of the upper skin and the lower skin of each section according to the design thickness of the upper skin and the lower skin of each section after removing the stringers and the minimum thickness limit; and step six, calculating the structure weight of the upper panel and the lower panel according to the final thickness of the upper skin and the lower skin of each section.

Description

Technical Field

[0001] This application belongs to the field of aircraft structural design, and specifically relates to a method for weight assessment of composite material wing panel structures. Background Technology

[0002] In the aircraft design phase, structural weight assessment of each section is a crucial task, as structural weight significantly impacts flight performance, fuel consumption, and cost. For monoplane wings, the wing panels, as the primary load-bearing structure, bear over 90% of the wing's bending moment and, along with spars and other structures, share the wing's torque. Typically, the wing panel structural weight accounts for more than 50% of the wing box section's weight. The wing panel structural weight is influenced by multiple factors, including wing aerodynamic shape, loads, structural layout, strength, stiffness, and material selection, requiring comprehensive optimization design. Therefore, rapid assessment of the wing panel weight is essential during the design phase to facilitate analysis and evaluation of multiple options.

[0003] Existing methods for assessing the weight of wing panels have the following drawbacks:

[0004] 1. During the design phase, multiple schemes are being demonstrated for the wing shape, structural layout, and structural materials. Using the finite element method to evaluate the panel weight is time-consuming and costly, and cannot meet the design cycle requirements of this phase.

[0005] 2. Composite material wing panels typically adopt a stringer + skin structure to improve the stability and load-bearing capacity of the panels. Traditional panel structure weight assessment based on engineering beam methods often ignores the influence of factors such as the stiffness ratio and area ratio of the stringer on the panel weight, resulting in low weight accuracy in the assessment.

[0006] 3. The stringers and skins of composite wing panels often use different layup ratios, which leads to differences in the equivalent elastic modulus of the two materials, making it more difficult to assess the weight of composite wing panels. There is little existing data on the weight assessment of composite wing panels.

[0007] In summary, there is currently no relevant technology that proposes a method for assessing the weight of composite wing panels that can simultaneously take into account panel load-bearing efficiency, panel strength control, and stringer area ratio control.

[0008] Therefore, there is an urgent need for a technical solution to overcome or mitigate at least one of the aforementioned defects in the existing technology. Summary of the Invention

[0009] The purpose of this application is to provide a method for evaluating the structural weight of composite material wing panels to solve at least one problem existing in the prior art.

[0010] The technical solution of this application is:

[0011] A method for assessing the structural weight of a composite material wing panel includes:

[0012] Step 1: Calculate the average equivalent modulus of elasticity of the skin, stringer, and wing beam along the tensile or compressive direction, respectively.

[0013] Step 2: Obtain the simplified model of the wing box section. Based on the average equivalent elastic modulus and the simplified model of the wing box section, calculate the equivalent height of each section and the load-bearing efficiency of each section.

[0014] Step 3: Calculate the equivalent thickness of the upper and lower skins of each section based on the equivalent height and bearing efficiency of each section.

[0015] Step 4: Based on the equivalent thickness of the upper and lower skins on each section, calculate the design thickness of the upper and lower skins after removing the stringers on each section.

[0016] Step 5: Based on the design thickness and minimum thickness limit of the upper and lower skins after removing the stringers in each section, determine the final thickness of the skins and lower skins on each section.

[0017] Step 6: Calculate the structural weight of the upper and lower wall panels based on the final thickness of the skin and lower skin on each section.

[0018] In at least one embodiment of this application, in step one, the average equivalent elastic modulus of the skin, stringer and wing beam in the 0° direction is taken as the average equivalent elastic modulus along the tensile or compressive direction, and the average equivalent elastic modulus of the skin, stringer and wing beam in the 0° direction is calculated by weighted average.

[0019] The average equivalent elastic modulus of the skin in the 0° direction is:

[0020] ;

[0021] In the formula, The average equivalent elastic modulus of the skin in the 0° direction. This is the sum of the equivalent elastic moduli in the 0° direction of all plies in the skin ply library. The number of different layers in the skin ply library;

[0022] The average equivalent elastic modulus of the stringer in the 0° direction is:

[0023] ;

[0024] In the formula, Let be the average equivalent elastic modulus of the stringer in the 0° direction. This is the sum of the equivalent elastic moduli in the 0° direction of all plies in the stringer ply library. The number of different plies in the stringer ply library;

[0025] The average equivalent elastic modulus of the wing spars in the 0° direction is:

[0026] ;

[0027] In the formula, The average equivalent elastic modulus of the wing spars in the 0° direction is given. This is the sum of the equivalent elastic moduli in the 0° direction of all plies in the wing sparple ply library. This refers to the number of different plies in the wing beam ply library.

[0028] In at least one embodiment of this application, in step two, the equivalent height of each cross-section is:

[0029] ;

[0030] In the formula, Let i be the equivalent height of the i-th profile. Let i be the area of ​​the i-th cross-section. The length of the line connecting the midpoints of the front and rear beams in the height direction at the i-th cross-section;

[0031] The bearing capacity of each section is:

[0032] ;

[0033] In the formula, Let i be the bearing capacity of the i-th profile. This represents the ratio of the girder area to the skin area of ​​the upper panel. This represents the ratio of the girder area to the skin area of ​​the lower wall panel. The width of the front beam's edge strip. The width of the beam edge strip of the rear beam. This is the thickness coefficient of the beam edge strip.

[0034] In at least one embodiment of this application, in step three, the equivalent thickness of the skin and the lower skin on each cross-section is:

[0035] ;

[0036] ;

[0037] ;

[0038] ;

[0039] In the formula, Let be the equivalent thickness of the skin on the i-th cross-section. The bending moment in the cross section. The allowable stress of the upper wall panel, The allowable strain for the wall panel under pressure. For the allowable stress of the lower wall panel, For the allowable strain of the wall panel, To account for the reduction value caused by the Poisson effect in composite wall panels.

[0040] In at least one embodiment of this application, in step four, the design thickness of the upper and lower skins after removing the stringers in each cross-section is:

[0041] ;

[0042] ;

[0043] In the formula, The design thickness of the upper skin after removing the stringer in the i-th section. The design thickness of the lower skin after removing the stringer in the i-th section.

[0044] In at least one embodiment of this application, in step five, the final thickness of the skin and the lower skin on each cross-section is:

[0045] ;

[0046] ;

[0047] In the formula, Let be the final thickness of the skin on the i-th cross-section. Let be the final thickness of the skin under the i-th profile. This is a minimum thickness limit.

[0048] In at least one embodiment of this application, in step six, the structural weights of the upper and lower wall panels are:

[0049] ;

[0050] ;

[0051] In the formula, This refers to the structural weight of the upper wall panel. Let the structural weight of the upper wall panel element at the i-th section be . This is the structural conversion factor for the upper wall panel. The density of the upper wall panel material. This refers to the structural weight of the lower wall panel. Let the weight of the lower wall panel element be the weight of the i-th section. This is the structural conversion factor for the lower wall panel. The density of the lower wall panel material. The distance between the (i+1)th profile and the ith profile is... The distance between the i-th profile and the (i-1)-th profile is denoted by n, where n is the number of profiles.

[0052] The invention has at least the following beneficial technical effects:

[0053] The composite material wing panel structure weight assessment method of this application rapidly calculates and assesses the structural parameters and weight of the simplified composite material wing box section panel. The assessment process can simultaneously consider factors such as panel strength control, stringer stiffness ratio control, stringer and skin material selection, and minimum size constraints. This method is applicable to the weight assessment of composite material panels and improves the accuracy of composite material wing panel structure weight assessment. Attached Figure Description

[0054] Figure 1 This is a flowchart of a composite material wing panel structure weight assessment method according to one embodiment of this application;

[0055] Figure 2 This is a schematic outline of the wing box segment according to one embodiment of this application;

[0056] Figure 3 This is a simplified schematic diagram of a wing box segment according to one embodiment of this application;

[0057] Figure 4 This is a schematic diagram of a wall panel unit according to one embodiment of this application.

[0058] in:

[0059] 101-Upper skin; 102-Upper girder; 103-Lower skin; 104-Lower girder; 105-Front beam; 106-Rear beam; 107-Front beam flange; 108-Rear beam flange. Detailed Implementation

[0060] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below with reference to the accompanying drawings. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are some, but not all, embodiments of this application. The embodiments described below with reference to the accompanying drawings are exemplary and intended to explain this application, and should not be construed as limiting this application. All other embodiments obtained by those skilled in the art based on the embodiments of this application without creative effort are within the scope of protection of this application. The embodiments of this application will be described in detail below with reference to the accompanying drawings.

[0061] In the description of this application, it should be understood that the terms "center", "longitudinal", "lateral", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limiting the scope of protection of this application.

[0062] The following is in conjunction with the appendix Figures 1 to 3 This application will be described in further detail.

[0063] This application provides a method for evaluating the weight of composite material wing panel structures, such as... Figure 1 As shown, it includes the following steps:

[0064] Step 1: Calculate the average equivalent modulus of elasticity of the skin, stringer, and wing beam along the tensile or compressive direction, respectively.

[0065] Step 2: Obtain the simplified model of the wing box section. Based on the average equivalent elastic modulus and the simplified model of the wing box section, calculate the equivalent height of each section and the load-bearing efficiency of each section.

[0066] Step 3: Calculate the equivalent thickness of the upper and lower skins of each section based on the equivalent height and bearing efficiency of each section.

[0067] Step 4: Based on the equivalent thickness of the upper and lower skins on each section, calculate the design thickness of the upper and lower skins after removing the stringers on each section.

[0068] Step 5: Based on the design thickness and minimum thickness limit of the upper and lower skins after removing the stringers in each section, determine the final thickness of the skins and lower skins on each section.

[0069] Step 6: Calculate the structural weight of the upper and lower wall panels based on the final thickness of the skin and lower skin on each section.

[0070] The weight assessment method for composite material wing panel structures in this application uses the following input conditions: wing box section frame arrangement (including spars and ribs), and cross-sectional bending moments at each rib location along the spanwise direction of the wing. The ratio of the girder area to the skin area of ​​the upper and lower wall panels .

[0071] The specific implementation steps are as follows:

[0072] First, in step one, the average equivalent modulus of elasticity of the wing skin, stringer, and spars along the tensile or compressive direction is calculated based on the material selection, ply ratio, and ply number range of the composite wing skin, stringer, and spars. For the wing skin, stringer, and spars, the tensile or compressive direction is generally the same as the 0° direction or the angle between the two is very small. Therefore, the average equivalent modulus of elasticity of the skin, stringer, and spars in the 0° direction can be used as the average equivalent modulus of elasticity along the tensile / compressive direction.

[0073] In aircraft structural design, the ply ratio of each component remains essentially constant across different thicknesses. Therefore, a weighted average method can be used to calculate the average equivalent elastic modulus of each component in the 0° direction. The formula for calculating the average equivalent elastic modulus of the skin in the 0° direction is as follows:

[0074] ;

[0075] In the formula, The average equivalent elastic modulus of the skin in the 0° direction. This is the sum of the equivalent elastic moduli in the 0° direction of all plies in the skin ply library. The number of different layers in the skin ply library;

[0076] Similarly, the average equivalent elastic modulus of the stringer in the 0° direction can be calculated as follows:

[0077] ;

[0078] In the formula, Let be the average equivalent elastic modulus of the stringer in the 0° direction. This is the sum of the equivalent elastic moduli in the 0° direction of all plies in the stringer ply library. The number of different plies in the stringer ply library;

[0079] The average equivalent elastic modulus of the wing spars in the 0° direction is:

[0080] ;

[0081] In the formula, The average equivalent elastic modulus of the wing spars in the 0° direction is given. This is the sum of the equivalent elastic moduli in the 0° direction of all plies in the wing sparple ply library. This refers to the number of different plies in the wing beam ply library.

[0082] A layup library for composite parts refers to the predefined combinations of layup sequences, thickness gradients, and layup ratios at the initial design stage. Designers can directly find the corresponding layup sequence from the library based on the layup ratio and thickness of the part, greatly improving design efficiency and ensuring quality. The calculation of the equivalent elastic modulus of laminates can be found in Appendix A1.3 of the "Guideline for Stability Analysis of Composite Structures".

[0083] Secondly, in step two, a simplified model of the wing box segment is obtained, such as... Figure 2-3 As shown. The equivalent height of each section is:

[0084] ;

[0085] In the formula, Let i be the equivalent height of the i-th profile. Let i be the area of ​​the i-th cross-section. The length of the line connecting the midpoints of the front and rear beams in the height direction at the i-th cross-section;

[0086] The bearing capacity of each section is:

[0087] ;

[0088] In the formula, Let i be the bearing capacity of the i-th profile. This represents the ratio of the girder area to the skin area of ​​the upper panel. This represents the ratio of the girder area to the skin area of ​​the lower wall panel. The width of the front beam's edge strip. The width of the beam edge strip of the rear beam. This is the thickness coefficient of the beam edge strip.

[0089] The ratio of the girder area to the skin area of ​​the upper and lower wall panels , Defined as the ratio of the total area of ​​the stringers to the area of ​​the skin in the panel cross-section. In a monoplane wing, the stringers need to provide simply supported boundary conditions for the skin and should buckle later than the skin. Therefore, there are certain control requirements for the ratio of stringer to skin area. This ratio is related to the cross-sectional shape of the stringers, the spacing between stringers, and the design requirements for stability and damage tolerance. In the design process, it can be determined based on reference aircraft and engineering experience. For I-shaped stringers, the ratio of stringer to skin area is generally between 0.45 and 0.65, and for T-shaped stringers, the ratio is generally between 0.65 and 0.85.

[0090] Width of the flange strips of the front and rear beams During the design phase, the size, spacing, and edge distance of the fasteners on the beam edge strips can be estimated based on these principles.

[0091] Beam flange thickness coefficient The ratio of the thickness of the beam edge strip to the local skin thickness is generally between 0.8 and 1.2.

[0092] Then, in step three, the equivalent thickness refers to the total thickness of the skin after converting the load-bearing capacity of the girder into skin thickness. Specifically, the equivalent thickness of the skin and lower skin on each section is:

[0093] ;

[0094] ;

[0095] In the formula, Let be the equivalent thickness of the skin on the i-th cross-section. The bending moment in the cross section. The allowable stress of the upper wall panel, The allowable strain for the wall panel under pressure. For the allowable stress of the lower wall panel, For the allowable strain of the wall panel, To account for the reduction value caused by the Poisson effect in composite wall panels.

[0096] Allowable stress of upper and lower wall panels The allowable stress is typically determined by factors such as aircraft materials, test data, and lifespan. For composite materials, due to the different equivalent elastic moduli of the skin and stringers, tensile and compressive allowable strains are usually used for control. The allowable stresses for the upper and lower panels are calculated as follows:

[0097] ;

[0098] ;

[0099] Considering the effect of Poisson's effect on the shrinkage value of composite wall panels Between 0.85 and 0.95.

[0100] Furthermore, in step four, the design thickness of the upper and lower skins after removing the stringers in each section is calculated as follows:

[0101] ;

[0102] ;

[0103] In the formula, The design thickness of the upper skin after removing the stringer in the i-th section. The design thickness of the lower skin after removing the stringer in the i-th section.

[0104] In step five, since the wing skin often serves as the boundary of the aircraft fuel tank, the lightning protection design of the composite material must meet the minimum thickness requirement. The limit is that this value is between 4 and 6 mm. The final thickness of the skin and lower skin on each section is:

[0105] ;

[0106] ;

[0107] In the formula, Let be the final thickness of the skin on the i-th cross-section. Let be the final thickness of the skin under the i-th profile. This is a minimum thickness limit.

[0108] Finally, in step six, the structural weights of the upper and lower wall panels are calculated.

[0109] like Figure 4 As shown, the structural weight of the upper wall panel element at the i-th section is calculated based on the final thickness of the skin on the i-th rib section.

[0110] ;

[0111] Calculate the structural weight of the lower wall panel element at the i-th section based on the final thickness of the lower skin at the i-th rib section:

[0112] ;

[0113] The structural weights of the upper and lower wall panels are:

[0114] ;

[0115] ;

[0116] In the formula, This refers to the structural weight of the upper wall panel. Let the structural weight of the upper wall panel element at the i-th section be . This is the structural conversion factor for the upper wall panel (this factor represents the estimated ratio of the actual digital model weight of the upper wall panel to the theoretical analysis weight). The density of the upper wall panel material. This refers to the structural weight of the lower wall panel. Let the weight of the lower wall panel element be the weight of the i-th section. This is the structural conversion factor for the lower wall panel (this factor represents the estimated ratio of the actual digital model weight of the lower wall panel to the theoretical analysis weight). The density of the lower wall panel material. The distance between the (i+1)th profile and the ith profile is... The distance between the i-th profile and the (i-1)-th profile is denoted by n, where n is the number of profiles.

[0117] Due to factors such as opening reinforcement, concentrated connection reinforcement such as hanging, composite material delamination and other details in the actual wall panel structure, as well as the difference in load transfer efficiency along the chord direction caused by the sweep effect, the final weight of the wall panel is higher than the actual calculated result. Therefore, the upper wall panel structure conversion factor is used. The value ranges from 1.35 to 1.45. Similarly, the structural conversion factor for the lower wall panel... Its value ranges from 1.35 to 1.45.

[0118] The composite material wing panel structure weight assessment method of this application considers the allowable stress of the panel when the stringer and skin use different materials or different ply ratios. It also considers constraints such as the panel's load-bearing capacity and the control of the stringer stiffness ratio within the panel, enabling rapid calculation of panel parameters and providing a relatively reliable load transfer basis for subsequent panel weight assessment. This application rapidly designs panel structural parameters and assesses weight based on the bending load of the wing section during the wing design phase, providing a relatively accurate wing panel weight reference for multi-scheme iteration and mitigating subsequent design risks.

[0119] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A method of composite wing panel structure weight assessment, characterized by, include: Step 1: Calculate the average equivalent modulus of elasticity of the skin, stringer, and wing beam along the tensile or compressive direction, respectively. Step 2: Obtain the simplified model of the wing box section. Based on the average equivalent elastic modulus and the simplified model of the wing box section, calculate the equivalent height of each section and the load-bearing efficiency of each section. Step 3: Calculate the equivalent thickness of the upper and lower skins of each section based on the equivalent height and bearing efficiency of each section. Step 4: Based on the equivalent thickness of the upper and lower skins on each section, calculate the design thickness of the upper and lower skins after removing the stringers on each section. Step 5: Based on the design thickness and minimum thickness limit of the upper and lower skins after removing the stringers in each section, determine the final thickness of the skins and lower skins on each section. Step 6: Calculate the structural weight of the upper and lower wall panels based on the final thickness of the skin and lower skin on each section.

2. The composite wing panel structure weight assessment method of claim 1, wherein, In step one, the average equivalent elastic modulus of the skin, stringer and wing beam in the 0° direction is taken as the average equivalent elastic modulus along the tension or compression direction, and the average equivalent elastic modulus of the skin, stringer and wing beam in the 0° direction is calculated by weighted averaging. The average equivalent elastic modulus of the skin in the 0° direction is: ； wherein E0 is the average equivalent modulus of elasticity in 0° direction of the skin, E0 is the sum of the equivalent modulus of elasticity in 0° direction of all plies in the skin ply library, N is the number of different plies in the skin ply library. The average equivalent elastic modulus of the stringer in the 0° direction is: ； In the formula, The average equivalent elastic modulus of the stringer in the 0° direction is given. This is the sum of the equivalent elastic moduli in the 0° direction of all plies in the stringer ply library. The number of different plies in the stringer ply library; The average equivalent elastic modulus of the wing spars in the 0° direction is: ； In the formula, The average equivalent elastic modulus of the wing spars in the 0° direction is given. This is the sum of the equivalent elastic moduli in the 0° direction of all plies in the wing sparple ply library. This refers to the number of different plies in the wing beam ply library.

3. The composite wing panel structure weight assessment method of claim 2, wherein, In step two, the equivalent height of each section is: ； In the formula, is the equivalent height of the i-th cross section, is the area of the i-th cross section, is the length of the line connecting the midpoints of the height direction of the front and rear spars of the i-th cross section. The bearing capacity of each section is: ； In the formula, Let i be the bearing capacity of the i-th profile. This represents the ratio of the girder area to the skin area of ​​the upper panel. This represents the ratio of the girder area to the skin area of ​​the lower wall panel. The width of the front beam's edge strip. The width of the beam edge strip of the rear beam. This is the thickness coefficient of the beam edge strip.

4. The composite wing panel structure weight assessment method of claim 3, wherein, In step three, the equivalent thickness of the skin and lower skin on each cross-section is: ； ； ； ； In the formula, Let be the equivalent thickness of the skin on the i-th cross-section. The bending moment in the cross section, The allowable stress of the upper wall panel, For the allowable strain of the wall panel, The allowable stress of the lower wall panel is... For the allowable strain of the wall panel, To account for the reduction value caused by the Poisson effect in composite wall panels.

5. The composite wing panel structure weight assessment method of claim 4, wherein, In step four, the design thickness of the upper and lower skins after removing the stringers in each section is as follows: ； ； In the formula, Let the design thickness of the upper skin be the length of the stringer after removing the stringer from the i-th section. The design thickness of the lower skin after removing the stringer in the i-th section.

6. The composite wing panel structure weight assessment method of claim 5, wherein, In step five, the final thickness of the skin and lower skin on each cross-section is as follows: ； ； wherein is the final thickness of the skin on the i-th section, is the final thickness of the skin on the i-th section, is the minimum thickness limit.

7. The composite wing panel structure weight assessment method of claim 6, wherein, In step six, the structural weights of the upper and lower wall panels are: ； ； In the formula, This refers to the structural weight of the upper wall panel. Let the structural weight of the upper wall panel element at the i-th section be . This is the structural conversion factor for the upper wall panel. The density of the upper wall panel material. This represents the structural weight of the lower wall panel. Let be the structural weight of the lower wall panel element at the i-th section. This is the structural conversion factor for the lower wall panel. The density of the lower wall panel material. The distance between the (i+1)th profile and the ith profile is... The distance between the i-th profile and the (i-1)-th profile is denoted by n, where n is the number of profiles.

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