A method for designing a dunnage thickness and a method for forming an aircraft skin

Abstract

The application relates to the technical field of aircraft skin preparation, and particularly discloses a cushion thickness design method and an aircraft skin forming method, which comprises the following steps: calculating the maximum strain of the aircraft skin after forming; calculating the maximum stress suffered by the aircraft skin during forming; calculating the maximum strain of the cushion generated during the forming process of the aircraft skin; and calculating the first thickness size of the cushion before compression. According to the method, the thickness design of the cushion can be completed in the early design process, the trial-and-error times of the cushion thickness are reduced, or the trial-and-error process is avoided, so that the design of the cushion thickness is time-saving, labor-saving, low in cost and high in efficiency.

Description

Technical Field

[0001] This invention relates to the field of aircraft skin manufacturing technology, and in particular to a method for designing padding thickness and a method for forming aircraft skin. Background Technology

[0002] Pad-assisted forming is one of the important forming methods for large-size variable-thickness single-curved skins of aircraft. The main forming process is to complete the variable-thickness processing of the aircraft skin in a flat state, use pads to fill the steps, and fix the aircraft skin and pads with clips to form an integral structure without thickness change. Then, this structure is formed by bending through a skin bending machine.

[0003] Considering cost or convenience, the padding material is generally different from the aircraft skin material. At the same time, the padding will undergo extrusion deformation during the forming process, so the thickness of the padding before compression is different from the thickness after compression. Currently, the thickness of the padding before compression is mainly determined by trial and error. Thickness compensation is made in different areas based on experience. Then, during the forming process, the uniformity of deformation in different areas of the aircraft skin is measured, and the padding thickness is increased or decreased in the corresponding areas to finally complete the padding design. This padding thickness design is time-consuming, laborious, costly and inefficient. Summary of the Invention

[0004] The purpose of this invention is to provide a method for designing padding thickness, so as to solve the problems of time-consuming, laborious, costly and inefficient padding thickness design.

[0005] On one hand, the present invention provides a method for designing the thickness of a padding material, the method comprising the following steps:

[0006] S01. Calculate the maximum strain after the aircraft skin is formed;

[0007] S02. Calculate the maximum stress that the aircraft skin experiences during the forming process;

[0008] S03. Calculate the maximum strain generated by the padding material during the aircraft skin forming process;

[0009] S04. Calculate the first thickness of the padding material before compression.

[0010] As a preferred embodiment of the aforementioned padding thickness design method, step S001 is further included before step S01, which involves testing the stress-strain curves of the aircraft skin material and the padding material.

[0011] As a preferred embodiment of the aforementioned padding thickness design method, in step S001, the stress-strain curve of the aircraft skin material is calculated through a tensile test.

[0012] As a preferred embodiment of the aforementioned padding thickness design method, in step S001, the stress-strain curve of the padding material is calculated through a tensile test.

[0013] As a preferred embodiment of the aforementioned padding thickness design method, in step S02, the maximum stress that can cause the aircraft skin to produce maximum strain is calculated based on the stress-strain curve of the aircraft skin material calculated from the tensile test.

[0014] As a preferred embodiment of the aforementioned padding thickness design method, in step S03, it is assumed that the maximum stress borne by the padding is equal to the maximum stress borne by the aircraft skin, and the maximum strain of the padding is calculated based on the stress-strain curve of the padding material calculated by tensile test.

[0015] As a preferred embodiment of the pad thickness design method, in step S04, the first thickness dimension of the pad after compression is calculated based on the first skin thickness dimension at the thickest part of the aircraft skin and the distance between the two rollers, and then the second thickness dimension of the pad before compression is calculated in combination with the maximum strain of the pad.

[0016] As a preferred embodiment of the pad thickness design method, in step S04, the third thickness of the pad after compression is calculated based on the second skin thickness dimension at the thinnest point of the aircraft skin and the distance between the two rollers, and then the fourth thickness of the pad before compression is calculated in combination with the maximum strain of the pad.

[0017] As a preferred embodiment of the aforementioned padding thickness design method, in step S01, the length of the convex surface after deformation is calculated under the premise that the neutral layer length of the aircraft skin remains unchanged, and then the maximum strain after the aircraft skin is formed is calculated.

[0018] On the other hand, the present invention provides an aircraft skin forming method, including the padding thickness design method in any of the above-mentioned schemes.

[0019] The beneficial effects of this invention are as follows:

[0020] This invention provides a method for designing padding thickness and a method for forming aircraft skin. The method for designing padding thickness includes: calculating the maximum strain after the aircraft skin is formed; calculating the maximum stress on the aircraft skin during the forming process; calculating the maximum strain generated by the padding material during the aircraft skin forming process; and calculating the initial thickness of the padding material before compression. This method allows for the completion of padding thickness design during the design process, reducing the number of trial and error steps, and even eliminating the trial and error process altogether. This makes padding thickness design time-saving, labor-saving, cost-effective, and highly efficient. Attached Figure Description

[0021] Figure 1This is a flowchart of the padding thickness design method in an embodiment of the present invention. Detailed Implementation

[0022] The technical solution of the present invention will now be clearly and completely described with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0023] In the description of this invention, it should be noted that the terms "center," "upper," "lower," "left," "right," "vertical," "horizontal," "inner," and "outer," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing the invention and for 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 limitations on the invention. Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance. The terms "first position" and "second position" refer to two different positions. Furthermore, "above," "on top of," and "over" the first feature in relation to the second feature includes the first feature directly above and diagonally above the second feature, or simply indicates that the first feature is at a higher horizontal level than the second feature. "Below," "under," and "below" the first feature in relation to the second feature includes the first feature directly below and diagonally below the second feature, or simply indicates that the first feature is at a lower horizontal level than the second feature.

[0024] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; and they can refer to the internal connection of two components. Those skilled in the art can understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0025] Embodiments of the present invention are described in detail below. Examples of these embodiments are shown in the accompanying drawings, wherein the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The embodiments described below with reference to the accompanying drawings are exemplary and are only used to explain the present invention, and should not be construed as limiting the present invention.

[0026] like Figure 1 As shown, this embodiment provides a method for designing the thickness of a padding material, which includes the following steps:

[0027] S01. Calculate the maximum strain after the aircraft skin is formed.

[0028] S02. Calculate the maximum stress that the aircraft skin experiences during the forming process.

[0029] S03. Calculate the maximum strain generated by the padding material during the aircraft skin forming process.

[0030] S04. Calculate the first thickness of the padding material before compression.

[0031] The above method can complete the design of the padding thickness in the early stage of the computer design process, reduce the number of trial and error times for padding thickness, and even avoid the trial and error process altogether, making the design of padding thickness time-saving, labor-saving, low-cost and highly efficient.

[0032] Since each batch of materials is different, in order to ensure the accuracy of the data, in this embodiment, optionally, step S001 is included before step S01 to test the stress-strain curves of the aircraft skin material and the padding material.

[0033] Since the aircraft skin is stretched during manufacturing, in this embodiment, specifically in step S001, the stress-strain curve of the aircraft skin material is calculated through a tensile test. In this embodiment, regarding the stress-strain curve of the padding material, in step S001, the stress-strain curve of the padding material is calculated through a tensile test. Additionally, since the padding material is partially compressed during the formation of the aircraft skin, the stress-strain curve of the padding material can also be calculated through a compression test. During the calculation process, different stress-strain curves are selected based on the location of the padding material. That is, in actual production, the stress-strain curve calculated through a tensile test is used for the stretched areas of the padding material, and the stress-strain curve calculated through a compression test is used for the compressed areas of the padding material.

[0034] In step S01, the length of the convex surface after deformation is calculated, assuming the neutral layer length of the aircraft skin remains constant, and then the maximum strain after the aircraft skin is formed is calculated. In other embodiments of this example, the convex surface can also be replaced by a concave surface.

[0035] In this embodiment, in step S02, the maximum stress that can cause the aircraft skin to produce the maximum strain is calculated based on the stress-strain curve of the aircraft skin material calculated by the tensile test.

[0036] Optionally, in step S03, it is assumed that the maximum stress borne by the padding material is equal to the maximum stress borne by the aircraft skin, and the maximum strain of the padding material is calculated based on the stress-strain curve of the padding material calculated from the tensile test. In this embodiment, the aircraft skin and the padding material move up and down between the rollers, and the weight of the aircraft skin and the padding material themselves does not need to be considered.

[0037] However, when the aircraft skin and padding move horizontally, the maximum stress on the padding can be obtained by subtracting the stress generated by the weight of the aircraft skin and / or padding from the maximum stress on the aircraft skin. For example, when three sets of rollers are arranged side-by-side, each set has two rollers that can clamp and apply pressure to the aircraft skin and padding. The maximum stress on the padding is obtained by subtracting the stress generated on the middle roller by the weight of the aircraft skin and / or padding between the two side rollers from the maximum stress on the aircraft skin. In other words, taking the padding on the upper side of the aircraft skin as an example, the maximum stress on the padding is A = (B·CD) / E, where B is the maximum stress on the aircraft skin, C is the contact area between the aircraft skin and the rollers; D is the weight of the aircraft skin and padding between the two side rollers; and E is the contact area between the padding and the rollers.

[0038] It should be noted that the conversion between weight and mass is well known to those skilled in the art, so the calculation methods for the weight of aircraft skin and padding will not be elaborated here.

[0039] In this embodiment, optionally, in step S04, the first thickness of the padding material after compression is calculated based on the first skin thickness at the thickest point of the aircraft skin and the distance between the two rollers. Then, the second thickness of the padding material before compression is calculated by combining the maximum strain of the padding material at this point. Using the above method, the thickness of the padding material before compression at the thickest point of the aircraft skin can be calculated.

[0040] Further, in step S04, the third thickness of the padding material after compression is calculated based on the second thickness dimension at the thinnest point of the aircraft skin and the distance between the two rollers. Then, the fourth thickness dimension of the padding material before compression is calculated by combining the maximum strain of the padding material at this point. Using the above method, the thickness of the padding material before compression at the thinnest point of the aircraft skin can be calculated.

[0041] By selecting points at different thicknesses of the aircraft skin and repeating the above method, the thickness of the padding material at all thicknesses of the aircraft skin can be determined.

[0042] This embodiment also provides an aircraft skin forming method, including the padding thickness design method in the above scheme.

[0043] Obviously, the above embodiments of the present invention are merely examples for clearly illustrating the present invention, and are not intended to limit the implementation of the present invention. Those skilled in the art can make other variations or modifications based on the above description. It is neither necessary nor possible to exhaustively describe all embodiments here. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of the present invention should be included within the scope of protection of the claims of the present invention.

Claims

1. A method of dunnage thickness design, characterized by, The method comprises the following steps: S01, calculating the maximum strain of the aircraft skin after forming; S02, calculating the maximum stress of the aircraft skin during forming; S03, calculating the maximum strain of the gasket during forming of the aircraft skin; S04, calculating the thickness of the gasket before compression; Before step S01, step S001 is further included, which tests the stress-strain curve of the aircraft skin material and the gasket material; the stress-strain curve of the aircraft skin material is calculated through a tensile test; the stress-strain curve of the gasket material is calculated through a tensile test; In step S02, the maximum stress that can cause the aircraft skin to generate the maximum strain is calculated according to the stress-strain curve of the aircraft skin material calculated through the tensile test; In step S03, the maximum strain of the gasket is calculated based on the stress-strain curve of the gasket material calculated through the tensile test, assuming that the maximum stress of the gasket is equal to the maximum stress of the aircraft skin; In step S04, the first thickness size of the gasket after compression is calculated according to the first skin thickness size at the thickest part of the aircraft skin and the distance between the two rollers, and then the second thickness size of the gasket before compression at this part is calculated in combination with the maximum strain of the gasket at this part; In step S04, the third thickness size of the gasket after compression is calculated according to the second skin thickness size at the thinnest part of the aircraft skin and the distance between the two rollers, and then the fourth thickness size of the gasket before compression at this part is calculated in combination with the maximum strain of the gasket at this part.

2. The method of mattress thickness design according to claim 1, wherein, In step S01, the length of the convex surface after deformation is calculated under the premise that the neutral layer length of the aircraft skin remains unchanged, and then the maximum strain of the aircraft skin after forming is calculated.

3. A method of forming an aircraft skin, characterized in that, The method comprises the thickness design method of the gasket according to any one of claims 1-2. The method comprises the thickness design method of the gasket according to any one of claims 1-2.

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