Design method and device for force-bearing covering cap of aircraft wing

Abstract

The invention provides a method and a device for designing a force-bearing covering cap of an aircraft wing, which are used for calculating static strength based on a finite element analysis method and aims to conveniently, quickly and accurately calculate the static strength of the force-bearing covering cap of the aircraft wing and determine a connecting gap between the covering cap and a wing box structure. According to the method provided by the invention, the rigidity and static strength of the aircraft wing with the force-bearing covering cap can be accurately and quickly calculated; the reduction of the rigidity of the wing by the gap of the force-bearing covering cap is fully considered and converted to the force-bearing covering cap, so that the accuracy of static strength analysis is improved; the method provided by the invention is simple to operate, the standardization of the calculation method and the flow is realized, the relatively large result error during calculation by different research and development personnel is avoided, and a standardized calculation method is provided for similar design problems.

Description

Technical Field

[0001] The invention belongs to the field of aircraft structure strength design, and in particular relates to a design method and device for an aircraft wing load-bearing cover. Background Art

[0002] Aircraft wing load-bearing covers specifically refer to various types of covers installed on the wing box, the main load-bearing structure of the wing. Their main function is to meet the functional requirements of the covers such as inspection and maintenance, while also transmitting the wing structure load, thereby ensuring that the wing structure meets the strength design requirements.

[0003] At present, in the design of domestic aircraft wing load-bearing flaps, the impact of the load-bearing flap on the wing load-bearing structure is usually multiplied by a uniformly specified reduction coefficient based on engineering experience, without considering the specific impact of the connection gap. However, when the above design method is used in actual model work, warping and cracks may occasionally occur in the flap structure and joints, affecting the structural safety.

[0004] Considering that the traditional design method lacks consideration of the actual impact of the actual connection between the aircraft wing load-bearing cover and the wing structure, the connection gap between the cover and the wing box structure will cause the cover to not transmit 100% of the load. The traditional design method only multiplies a unified reduction coefficient when analyzing the strength of the load-bearing cover. This design method is not accurate enough for the cover design. Therefore, it is necessary to establish a new design method to conveniently and accurately design the aircraft wing load-bearing cover. Summary of the Invention

[0005] The present invention provides a design method and device for an aircraft wing load-bearing flap cover, which can accurately and conveniently calculate and verify the static strength of the wing load-bearing flap cover and accurately determine the connection gap between the flap cover and the wing box structure.

[0006] A first aspect of the present invention provides a method for designing an aircraft wing load-bearing cover, comprising:

[0007] S1. Establish an overall finite element model of the wing structure, constrain the wing-body joint according to actual boundary conditions, and apply all working loads; the overall finite element model of the wing structure is provided with a flap opening;

[0008] S2. Calculate the displacement difference of the flap opening edge in the wing box structure under all working conditions using finite element calculation software;

[0009] S3. Subtract the clearance of the bolt connection of the flap in the design state from the displacement difference of the flap opening edge to obtain the difference, and obtain the load transfer ratio based on the difference;

[0010] S4. Establish a finite element model of the flap. Reduce the modulus of the flap material in the finite element model proportionally according to the load transfer ratio to obtain a revised finite element model of the flap. This model is then incorporated into the overall finite element model of the wing structure for solution to obtain the actual load boundary conditions of the flap.

[0011] S5. Based on the actual load boundary conditions of the mouth cover, solve the static and stability of the mouth cover and calculate the safety margin of the outlet cover; if the margin does not meet the design requirements, increase the strength and stiffness of the mouth cover until the safety margin is met.

[0012] Optionally, obtain the transmission ratio based on the difference, including:

[0013] The load transfer ratio is obtained by dividing the difference by the displacement difference of the opening edge of the flap.

[0014] Optionally, the displacement difference is a maximum spanwise displacement difference.

[0015] Optionally, the modulus of the flap material in the finite element model of the flap is proportionally reduced according to the load ratio, including:

[0016] The Young's modulus of the mouth cover material is multiplied by the load transfer ratio to obtain the Young's modulus of the load-bearing mouth cover material after the ratio is reduced.

[0017] Optionally, the design requirement is that the safety margin of the wing box structure and the flap structure body and connections must be no less than 1.0.

[0018] Optionally, S4 also includes:

[0019] The overall finite element model of the wing structure is brought into the solution to obtain the stress and strain state of the overall finite element model of the wing structure; the stress and strain state is affected by the flap installed in the flap opening area;

[0020] The method further comprises:

[0021] Based on the stress and strain state of the overall finite element model of the wing structure, check whether the wing structure meets the design requirements.

[0022] A second aspect of the present invention provides a design device for an aircraft wing load-bearing cover, for executing the method described in any one of the first aspects.

[0023] A third aspect of the present invention provides a computer storage medium, comprising: a memory and a processor;

[0024] The memory is configured to store executable instructions;

[0025] The processor is configured to implement the method as described in any one of the first aspects when executing the executable instructions stored in the memory.

[0026] The present invention provides a design method and device for an aircraft wing load-bearing flap, which performs static strength calculations based on finite element analysis. The purpose is to conveniently, quickly, and accurately calculate the static strength of the aircraft wing load-bearing flap and determine the connection gap between the flap and the wing box structure. The method provided by the present invention can accurately and quickly calculate the stiffness and static strength of an aircraft wing with a load-bearing flap; it fully considers the reduction in wing stiffness caused by the load-bearing flap gap and converts it to the load-bearing flap, thereby improving the accuracy of the static strength analysis; the method provided by the present invention is simple to operate, achieves standardization of calculation methods and processes, avoids large errors in results when different researchers perform calculations, and provides a standardized calculation method for similar design problems. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] Figure 1 This is a schematic diagram of an aircraft wing load-bearing cover according to the present invention;

[0028] Description of reference numerals:

[0029] 1-Load-bearing cover;

[0030] 2-gap at the edge of the flap;

[0031] 3 is the wing of the aircraft. DETAILED DESCRIPTION

[0032] To make the purpose, technical solutions, and advantages of the embodiments of the present invention more clear, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of them. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making any creative work are within the scope of protection of the present invention.

[0033] The features and illustrative embodiments of various aspects of the present invention will be described in detail below. In the detailed description below, many specific details are proposed in order to provide a comprehensive understanding of the present invention. However, it will be apparent to those skilled in the art that the present invention can be implemented without the need for some of these specific details. The following description of the embodiments is merely intended to provide a better understanding of the present invention by illustrating examples of the present invention. The present invention is in no way limited to any specific arrangement and method proposed below, but rather encompasses any improvements, replacements, and modifications to structures, methods, and devices without departing from the spirit of the present invention. In the accompanying drawings and the following description, well-known structures and techniques are not shown to avoid unnecessary ambiguity in the present invention.

[0034] In the description of the present invention, it should be noted that the directions or positional relationships indicated by terms such as "center," "up," "down," "left," "right," "vertical," "horizontal," "inside," and "outside" are based on the directions or positional relationships described in the accompanying drawings and are intended only to facilitate and simplify the description of the present invention and should not be construed as limiting the present invention. Furthermore, the use of ordinal numbers (e.g., "first and second," etc.) is intended to distinguish between objects and is not limited to this order, and should not be construed as indicating or implying relative importance.

[0035] In the description of the present invention, it should be noted that, unless otherwise specified or limited, the terms "mounted," "connected," and "connected" should be understood broadly, and may refer to direct connection or indirect connection through an intermediary. Those skilled in the art will understand the specific meanings of these terms in the present invention.

[0036] It should be noted that, in the absence of conflict, the embodiments of the present invention and the features in the embodiments can be combined with each other, and the embodiments can refer to and quote each other. The present invention will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0037] The present invention will be described in further detail below with reference to the embodiments and drawings, but the embodiments of the present invention are not limited thereto.

[0038] The present invention provides a design method for an aircraft wing load-bearing cover, comprising:

[0039] S1. Establish an overall finite element model of the wing structure, constrain the wing-body joint according to actual boundary conditions, and apply all working loads; the overall finite element model of the wing structure is provided with a flap opening;

[0040] S2. Calculate the displacement difference of the flap opening edge in the wing box structure under all working conditions using finite element calculation software;

[0041] S3. Subtract the clearance of the bolt connection of the flap in the design state from the displacement difference of the flap opening edge to obtain the difference, and obtain the load transfer ratio based on the difference;

[0042] S4. Establish a finite element model of the flap. Reduce the modulus of the flap material in the finite element model proportionally according to the load transfer ratio to obtain a revised finite element model of the flap. This model is then incorporated into the overall finite element model of the wing structure for solution to obtain the actual load boundary conditions of the flap.

[0043] S5. Based on the actual load boundary conditions of the mouth cover, solve the static and stability of the mouth cover and calculate the safety margin of the outlet cover; if the margin does not meet the design requirements, increase the strength and stiffness of the mouth cover until the safety margin is met.

[0044] Optionally, obtain the transmission ratio based on the difference, including:

[0045] The load transfer ratio is obtained by dividing the difference by the displacement difference of the opening edge of the flap.

[0046] Optionally, the displacement difference is a maximum spanwise displacement difference.

[0047] Optionally, the modulus of the flap material in the finite element model of the flap is proportionally reduced according to the load ratio, including:

[0048] The Young's modulus of the mouth cover material is multiplied by the load transfer ratio to obtain the Young's modulus of the load-bearing mouth cover material after the ratio is reduced.

[0049] Optionally, the design requirement is that the safety margin of the wing box structure and the flap structure body and connections must be no less than 1.0.

[0050] Optionally, S4 also includes:

[0051] The overall finite element model of the wing structure is brought into the solution to obtain the stress and strain state of the overall finite element model of the wing structure; the stress and strain state is affected by the flap installed in the flap opening area;

[0052] The method further comprises:

[0053] Based on the stress and strain state of the overall finite element model of the wing structure, check whether the wing structure meets the design requirements.

[0054] In a possible embodiment, the present invention is implemented as follows:

[0055] First, attach Figure 1 Taking the wing load-bearing cover 1 in the figure as an example, it is connected to the aircraft wing 3 by bolts on all sides, bearing the entire wing load. The figure shows the wing box upper wall panel, which bears in-plane tensile and compressive loads. In the figure, there is a gap 2 between the cover edge and the aircraft wing 3.

[0056] Second, given the wing numerical model and material properties, a finite element analysis model was established. The skin was simplified into plate-shell elements, and the load-bearing opening cover was treated as an opening area. The relevant loads borne by the wing were applied and calculated.

[0057] Third, extract the displacement data of the nodes of the wing finite element model in the opening area where the load-bearing cover is located;

[0058] Fourth, subtract the displacement data of the corresponding nodes in the front, back, left and right of the opening area where the load-bearing cover is located from each other to obtain the displacement difference of the cover edge relative to the node, that is, the gap of the cover edge;

[0059] Fifth, subtract the gap at the edge of the flap obtained in step 4 from the actual width of the flap, and then divide the result by the displacement difference of the flap opening edge to obtain the gap percentage. Multiply this gap percentage by the Young's modulus of the flap material to obtain the Young's modulus of the load-bearing flap material after geometric conversion.

[0060] Sixth, according to the Young's modulus of the load-bearing flap material after geometric conversion and the load-bearing flap numerical model, a finite element model of the load-bearing flap is established. The edges of the flap finite element model must share nodes with the wing finite element model.

[0061] Seventh, the wing finite element model connected to the load-bearing port cover finite element model is loaded with the relevant loads of step 2 and the data processing and analysis are performed again.

[0062] Eighth, extract the load boundary conditions of the load-bearing port cover finite element model from the wing model in step seven, perform static and stability solutions, and calculate the safety margin of the port cover.

[0063] The advantages and effects of the present invention primarily include: The method provided herein accurately and rapidly calculates the static strength of a finite element model of a wing load-bearing flap. Firstly, it considers the effect of the flap's own stiffness on the wing's stiffness; secondly, it incorporates the effect of flap gaps on force transmission into the flap stiffness calculation, enabling more accurate iterative calculations. The method provided by the present invention is simple to operate and standardizes the calculation method and process, providing a standardized design approach for wing load-bearing flap design.

[0064] The above specific implementation methods are detailed descriptions of the present invention. It cannot be considered that the specific implementation methods of the present invention are limited to these descriptions. For ordinary technicians in the technical field to which the present invention belongs, they can make several simple deductions and substitutions without departing from the concept of the present invention, which should be regarded as falling within the scope of protection of the present invention.

Claims

1. A design method for an aircraft wing load-bearing cover, characterized in that: include: S1. Establish an overall finite element model of the wing structure, constrain the wing-body joint according to the actual boundary conditions, and apply all working loads; A flap opening is provided in the overall finite element model of the wing structure; S2. Calculate the displacement difference of the flap opening edge in the wing box structure under all working conditions using finite element calculation software; S3. Subtract the clearance of the bolt connection of the flap in the design state from the displacement difference of the flap opening edge to obtain the difference, and obtain the load transfer ratio based on the difference; S4. Establish a finite element model of the flap. Reduce the modulus of the flap material in the finite element model proportionally according to the load transfer ratio to obtain a revised finite element model of the flap. This model is then incorporated into the overall finite element model of the wing structure for solution to obtain the actual load boundary conditions of the flap. S5. Based on the actual load boundary conditions of the mouth cover, solve the static and stability of the mouth cover and calculate the safety margin of the outlet cover; if the margin does not meet the design requirements, increase the strength and stiffness of the mouth cover until the safety margin is met.

2. The design method of an aircraft wing load-bearing cover according to claim 1, characterized in that: Obtain the transmission ratio based on the difference, including: The load transfer ratio is obtained by dividing the difference by the displacement difference of the opening edge of the flap.

3. The design method of an aircraft wing load-bearing cover according to claim 1, characterized in that: The displacement difference is the maximum spanwise displacement difference.

4. The design method of an aircraft wing load-bearing cover according to claim 1, characterized in that: According to the load ratio, the modulus of the flap material in the finite element model of the flap is proportionally reduced, including: The Young's modulus of the mouth cover material is multiplied by the load transfer ratio to obtain the Young's modulus of the load-bearing mouth cover material after the ratio is reduced.

5. The design method of an aircraft wing load-bearing cover according to claim 1, characterized in that: The design requirement is that the safety margin of the wing box structure and the flap structure body and connections must be no less than 1.

0.

6. The design method of an aircraft wing load-bearing cover according to claim 1, characterized in that: S4 also includes: The overall finite element model of the wing structure is brought into the solution to obtain the stress and strain state of the overall finite element model of the wing structure; the stress and strain state is affected by the flap installed in the flap opening area; The method further comprises: Based on the stress and strain state of the overall finite element model of the wing structure, check whether the wing structure meets the design requirements.

7. A design device for an aircraft wing load-bearing cover, characterized in that: Used to perform the method according to any one of claims 1 to 6.

8. A computer storage medium, characterized in that include: memory and processor; The memory is configured to store executable instructions; The processor is configured to implement the method according to any one of claims 1 to 6 when executing the executable instructions stored in the memory.

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