Forming method for composite aircraft bulkhead

Abstract

The invention provides a forming method for a composite aircraft bulkhead. The forming method comprises the following steps that a combined core mold is manufactured according to the shape and size of the aircraft bulkhead; the combined core mold is pulled to penetrate through a three-dimensional knitting machine in the extending direction of the combined core mold, so that the three-dimensional knitting machine is used for knitting and wrapping fibers on the outer surface of the combined core mold to form a knitting layer; performing extrusion deformation on the braid layer according to the shape of the aircraft bulkhead to form a preformed body; the preformed body is cured; and cutting the cured preform according to the shape and size of the aircraft bulkhead to form a plurality of aircraft bulkheads. According to the scheme, the automation degree and the production efficiency can be improved, and the problem that the web is prone to wrinkling can be solved.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of automated manufacturing of composite parts, and in particular to a forming method for a composite aircraft bulkhead. BACKGROUND

[0002] Composite materials are an ideal choice for achieving lightweight aircraft fuselages due to their excellent specific stiffness, specific strength, and fatigue resistance. The manufacturing process of aircraft fuselage bulkheads, as key load-bearing components, directly affects aircraft performance and production efficiency. Currently, Z-section composite aircraft bulkheads are mainly manufactured by manually laying up prepreg to prepare a preform, and then heating and pressing the preform to solidify and form.

[0003] Therefore, there is a need to provide a forming method for a composite aircraft bulkhead to at least partially solve the above problems. SUMMARY

[0004] The forming method for a composite aircraft bulkhead provided by the present application can solve the problems of low automation, low production efficiency, and prone to wrinkles in the web. The forming method comprises the following steps: A combined core mold is made according to the shape and size of the aircraft bulkhead; The combined core mold is pulled through a three-dimensional braiding machine along the extension direction of the combined core mold, so that the three-dimensional braiding machine braids and covers fibers on the outer surface of the combined core mold to form a braided layer; The braided layer is extruded to form a preform according to the shape of the aircraft bulkhead; The preform is solidified; and The solidified preform is cut according to the shape and size of the aircraft bulkhead to form a plurality of aircraft bulkheads.

[0005] Preferably, the aircraft bulkhead comprises a web and first and second flanges connected perpendicularly to the two sides of the web, respectively, the first and second flanges extending in opposite directions, so that the cross-section of the aircraft bulkhead is approximately Z-shaped.

[0006] Preferably, the aircraft bulkhead has an arc-shaped structure along its length direction.

[0007] Preferably, the combined core mold comprises a profiled core mold in the middle, the profiled core mold comprising a first part and a second part perpendicularly intersecting with the first part, and the combined core mold further comprises four square blocks located at the outer periphery of the profiled core mold, the four square blocks being respectively located at the corners formed by the intersection of the first part and the second part, so that the cross-sectional configuration of the combined core mold is square.

[0008] Preferably, the extruding and deforming the woven layer according to the shape of the aircraft bulkhead to form a preform comprises: extracting the four square blocks located at the outer periphery of the profiled core mold from the combined core mold; and extruding the woven layer towards the four corners of the profiled core mold, so that the woven layer is attached to the outer surface of the profiled core mold.

[0009] Preferably, the cutting the solidified preform comprises: cutting the preform into four aircraft bulkheads with Z-shaped cross sections according to the shape and size of the aircraft bulkhead.

[0010] Preferably, the cross-sectional circumference of the profiled core mold is equal to the sum of the length of the Z-shaped cross section of the four aircraft bulkheads and the length of the reserved cutting section.

[0011] Preferably, the solidifying the preform comprises: placing the preform into a forming mold and solidifying by injecting resin using the RTM process.

[0012] Preferably, the resin comprises any one selected from the group consisting of unsaturated polyester resin, epoxy resin, phenolic resin, vinyl ester resin and polyurethane.

[0013] Preferably, the fiber comprises any one selected from the group consisting of carbon fiber, glass fiber, aramid fiber, ultra-high molecular weight polyethylene fiber, basalt fiber and quartz fiber.

[0014] Preferably, the three-dimensional braiding machine is a ring braiding machine.

[0015] Preferably, the pulling the combined core mold through the three-dimensional braiding machine along the extension direction of the combined core mold further comprises: making the central axis of the combined core mold perpendicular to the braiding plane and at the center of the braiding ring of the three-dimensional braiding machine.

[0016] The above-mentioned scheme adopted according to the embodiments of the present application has the following beneficial effects: (1) In the above scheme, the three-dimensional braiding machine can braid dry fibers to cover the outer surface of the combined core mold. Since the dry fibers have good conformability, they can be braided into an arc-shaped structure along the arc-shaped contour of the combined core mold, thereby avoiding wrinkles at the web bending part of the aircraft bulkhead with a Z-shaped cross section.

[0017] (2) The combined core mold with a roughly square cross section is used as the internal support body, which solves the problem that the aircraft bulkhead preform with a Z-shaped cross section is difficult to braid in the prior art. The above scheme can produce four aircraft bulkheads at the same time through one braiding, greatly improving the production efficiency and greatly reducing the material waste caused by the production of a single bulkhead by a single core mold.

[0018] (3) The three-dimensional preform is cut after being cured, which avoids the problem of loose fibers caused by a large amount of cutting of the braided layer, reduces the labor cost, greatly retains the overall mechanical properties of the fiber preform, and improves the overall mechanical properties of the product. BRIEF DESCRIPTION OF DRAWINGS

[0019] In order to better understand the above and other objects, features, advantages and functions of the present application, reference can be made to the preferred embodiments shown in the accompanying drawings. The same or similar reference signs in the drawings refer to the same or similar components. Those skilled in the art should understand that the drawings are intended to illustrate the preferred embodiments of the present application, and have no limiting effect on the scope of the present application, and the components in the drawings are not drawn to scale.

[0020] Figure 1 is a structural schematic view of an aircraft bulkhead according to a preferred embodiment of the present application; Figure 2 is a cross-sectional schematic view of the aircraft bulkhead shown in Figure 1 Figure 3 is a structural schematic view of a combined core mold according to a preferred embodiment of the present application; Figure 4 is a cross-sectional schematic view of the combined core mold and the braided layer wrapped outside it shown in Figure 3 Figure 5 is a schematic view of a preform after cutting according to a preferred embodiment of the present application; Figure 6 is a cross-sectional schematic view of the preform after cutting shown in Figure 5 DETAILED DESCRIPTION

[0021] ​​​Below, the forming method for composite aircraft partition frame according to the preferred embodiment of the present application will be described in detail with reference to the drawings. It can be understood that the following given is only according to the preferred embodiment of the present application, and the person skilled in the art can think of other ways that can achieve the present application on the basis of the preferred embodiment, and other ways also fall within the scope of the present application.

[0022] First of all, it should be noted that the direction terms and position terms in the present application should be understood as relative direction and position, rather than absolute direction and position.

[0023] The following refers to Figures 1-6 The forming method for composite aircraft partition frame according to the preferred embodiment of the present application will be described in detail.

[0024] Figure 1 An example of the composite aircraft partition frame 10 to be formed by the forming method provided by the present application is shown, Figure 2 A cross section of the aircraft partition frame 10 is shown. As Figure 1 shown, the aircraft partition frame 10 includes a web 11 and a first flange 14 and a second flange 15 connected perpendicularly to two sides of the web 11 respectively, the first flange 14 and the second flange 15 extend in opposite directions respectively, so that the cross section of the aircraft partition frame 10 is approximately Z-shaped. Preferably, there is a circular arc transition area 12 between the first flange 14 and the web 11, and there is a circular arc transition area 13 between the second flange 15 and the web 11. Continuing to refer to Figure 1 , the aircraft partition frame 10 has an arc-shaped structure along its length direction.

[0025] The forming method for the aircraft partition frame 10 according to the preferred embodiment of the present application includes the following steps.

[0026] Step 1: Make a combined core mold 30 according to the shape and size of the aircraft partition frame 10.

[0027] The "core mold" mentioned in the present application refers to a central mold used to support the woven fibers (described below) and give the woven fibers a specific shape and size during the forming process of the aircraft bulkhead. The "combined core mold" refers to the overall core mold formed by the combination of multiple core molds. The combined core mold 30 serves as the internal support mold of the aircraft bulkhead 10 during the forming process, which can be manufactured according to the outer shape and geometric dimensions of the aircraft bulkhead 10. Preferably, the combined core mold 30 has the same extension length as the aircraft bulkhead 10, and has the same arc structure along its extension length as the curvature of the aircraft bulkhead 10. In particular, the combined core mold 30 is configured to have an arc structure with equal cross sections in the length direction. The cross-sectional size of the combined core mold 30 is also closely related to the cross-sectional size of the aircraft bulkhead to be formed. In particular, the cross-sectional size of the combined core mold 30 is designed such that the woven fibers wrapped outside the combined core mold 30 can be cut into four aircraft bulkheads after subsequent preforming and curing steps.

[0028] In one embodiment, as shown in Figure 3 and Figure 4 The combined core mold 30 includes a profiled core mold 31 in the middle and four square blocks 32 at the outer periphery of the profiled core mold 31. The profiled core mold 31 includes a first portion and a second portion perpendicular to the first portion, both of which have a certain thickness. The four square blocks 32 are respectively located at the four corners formed by the intersection of the first portion and the second portion, and together with the profiled core mold 31 form the combined core mold 30 with a generally square cross section and an arc shape in the length direction. The square structure of the combined core mold 30 is suitable for three-dimensional weaving as an internal support during the subsequent three-dimensional weaving step, thereby solving the problem of difficulty in weaving the traditional Z-shaped cross-section aircraft bulkhead. The cross-sectional perimeter of the profiled core mold 31 is equal to the sum of the length of the Z-shaped cross-section of the four aircraft bulkheads 10 and the length of the reserved cutting section. That is, the outer periphery of the cross section of the profiled core mold 31 is a closed shape formed by connecting the Z-shaped cross-section of the four aircraft bulkheads 10 and the reserved cutting section.

[0029] Preferably, the outer surface of the combined core mold 30 is smooth and flat, so as to ensure that the fibers will not be damaged during the subsequent formation of the woven layer 33.

[0030] Step 2: pulling the combined core mold 30 in the extension direction of the combined core mold 30 through the three-dimensional weaving machine, so that the three-dimensional weaving machine weaves and wraps fibers on the outer surface of the combined core mold 30 to form a woven layer 33.

[0031] In this step, the combined core mold 30 can be clamped by a core mold pulling device and pulled along the extension direction of the combined core mold 30, so that the combined core mold 30 passes through the three-dimensional braiding machine. The three-dimensional braiding machine can be a ring braiding machine. During the pulling process, the central axis of the combined core mold 30 is perpendicular to the braiding plane and is at the center of the braiding ring of the three-dimensional braiding machine, and the combined core mold 30 is controlled to move stably and regularly at a pulling speed that meets the braiding angle of the established process. The three-dimensional braiding machine braids dry fibers to cover the outer surface of the combined core mold 30 and repeatedly braids according to the established process to form a three-dimensional braided layer 33 with a certain thickness. It can be understood that the three-dimensional braided layer 33 has a closed shape in the circumferential direction. Preferably, the fibers can be any one selected from the following: carbon fibers, glass fibers, aramid fibers, ultra-high molecular polyethylene fibers, basalt fibers, and quartz fibers.

[0032] Since the dry fibers have good conformability, the above scheme can braided a braided layer with an arc structure along the arc profile of the combined core mold, which can avoid wrinkles at the bending part of the web of the subsequently formed aircraft bulkhead with a Z-shaped cross section.

[0033] Step 3: Extruding and deforming the braided layer 33 according to the shape of the aircraft bulkhead to form a preform.

[0034] After the braiding step of the above step 2 is completed, the three-dimensional braided layer 33 formed by wrapping the combined core mold 30 can be taken out of the three-dimensional braiding machine, the four square blocks 32 wrapped around the outer periphery of the profiled core mold 31 can be extracted from the combined core mold 30, and the three-dimensional braided layer 33 can be extruded towards the vacant square block area so that the braided layer 33 is attached to the outer surface of the intermediate profiled core mold 31. The braided layer 33 formed in this way is called a preform, which has a shape matching the outer periphery of the profiled core mold 31.

[0035] Step 4: curing the preform.

[0036] The preform formed in step 3 is placed in a molding mold together with the profiled core mold 31 inside, and a resin is injected into the molding mold by using an RTM (resin transfer molding) process, so that the preform is cured and formed. Preferably, the material of the profiled core mold 31 is configured to have a low thermal deformation amount within the curing temperature range, so that the deformation caused by temperature change can be avoided to affect the molding quality. Preferably, the resin can be any one selected from the following: unsaturated polyester resin, epoxy resin, phenolic resin, vinyl ester resin, and polyurethane, which can be selected by those skilled in the art according to actual needs.

[0037] Step 5: cutting the cured preform to form an aircraft bulkhead 10 with a predetermined shape and size.

[0038] The preform is taken out of the forming mold, and is cut into four aircraft partition frames with cross sections as shown in Figure 5 by a cutting device according to the size of the aircraft partition frame 10. Figure 6 The inner-arc aircraft partition frames 51 and 52 are mirror-symmetrical structures, and have the same size. Figure 6 The cross-sectional structures 61 and 62 of the inner-arc aircraft partition frames 51 and 52 are shown. The outer-arc aircraft partition frames 53 and 54 are also mirror-symmetrical structures, and have the same size. Figure 6 The cross-sectional structures 63 and 64 of the outer-arc aircraft partition frames 53 and 54 are shown. Therefore, the forming method according to the present application can produce two pairs of aircraft partition frames with different sizes, i.e., one pair of inner-arc aircraft partition frames 51 and 52 and one pair of outer-arc aircraft partition frames 53 and 54, by using the combined core mold structure, which greatly accelerates the production efficiency and greatly reduces the material waste caused by the single core mold producing a single aircraft partition frame.

[0039] The forming method according to the preferred embodiment of the present application uses the combined core mold with a roughly square cross section as the internal support body, which solves the problem that the aircraft partition frame preform with a Z-shaped cross section is difficult to weave, and the above scheme can finally produce four aircraft partition frames at one time, which greatly accelerates the production efficiency and greatly reduces the material waste caused by the single core mold producing a single partition frame. In addition, the three-dimensional weaving machine can weave the dry fibers to cover the outer surface of the combined core mold. Since the dry fibers have good conformability, they can be woven into an arc-shaped structure along the arc-shaped contour of the combined core mold, which can avoid the wrinkles at the web bending part of the aircraft partition frame with a Z-shaped cross section.

[0040] The forming method according to the preferred embodiment of the present application uses the RTM process to infiltrate the resin into the dry fibers in the preform and to solidify, which can not only preserve the mechanical properties of the dry fibers and improve the comprehensive mechanical properties of the aircraft partition frame, but also can use the cutting device to divide the preform into four aircraft partition frames after solidification, thereby improving the production efficiency of the aircraft partition frame.

[0041] The above description of various embodiments of the present application is provided to a person of ordinary skill in the relevant art for descriptive purposes. The present application is not intended to be exclusive or limited to a single disclosed embodiment. As such, a person of ordinary skill in the art having the benefit of the above teachings will appreciate and understand other embodiments of the present application. Accordingly, although specific embodiments have been described, a person of ordinary skill in the art will appreciate or readily be able to develop other embodiments without undue experimentation. The present application is intended to include all such alternatives, modifications and variations as falling within the spirit and scope of the present application as described above.

Claims

1. A method of forming a composite aircraft bulkhead, characterized by, The forming method comprises the following steps: manufacturing a combined core mold according to the shape and size of the aircraft partition frame; pulling the combined core mold through a three-dimensional braiding machine along the extension direction of the combined core mold, so that the three-dimensional braiding machine braids and covers fibers on the outer surface of the combined core mold to form a braided layer; extruding the braided layer according to the shape of the aircraft partition frame to form a preformed body; solidifying the preformed body; and cutting the solidified preformed body according to the shape and size of the aircraft partition frame to form a plurality of aircraft partition frames.

2. The molding method according to claim 1, characterized by, The aircraft partition frame comprises a web plate and a first flange and a second flange respectively perpendicularly connected on both sides of the web plate, the first flange and the second flange respectively extend in opposite directions, so that the cross section of the aircraft partition frame is approximately Z-shaped.

3. The molding method according to claim 2, characterized by, The aircraft partition frame has an arc-shaped structure along its length direction.

4. The molding method according to claim 2, characterized by, The combined core mold comprises a profiled core mold in the middle, the profiled core mold comprises a first part and a second part perpendicularly intersecting with the first part, and the combined core mold further comprises four square blocks on the outer periphery of the profiled core mold, the four square blocks are respectively located at the corners formed by the intersection of the first part and the second part, so that the cross section of the combined core mold is square.

5. The molding method according to claim 4, characterized by The extruding the braided layer according to the shape of the aircraft partition frame to form a preformed body comprises: extracting the four square blocks placed on the outer periphery of the profiled core mold from the combined core mold; and extruding the braided layer towards the four corners of the profiled core mold, so that the braided layer is attached to the outer surface of the profiled core mold.

6. The molding method according to claim 5, characterized by The cutting the solidified preformed body comprises: cutting the preformed body into four aircraft partition frames with Z-shaped cross sections according to the shape and size of the aircraft partition frame.

7. The molding method according to claim 5 or 6, characterized by, The cross section circumference of the profiled core mold is equal to the sum of the length of the Z-shaped cross section of the four aircraft partition frames and the length of the reserved cutting section.

8. The molding method according to claim 1, characterized by, The solidifying the preformed body comprises: putting the preformed body into a forming mold and using an RTM process to inject resin for solidification forming.

9. The molding method according to claim 8, characterized by, The resin comprises any one selected from the following: unsaturated polyester resin, epoxy resin, phenolic resin, vinyl ester resin and polyurethane.

10. The molding method according to claim 1, characterized by, The fibers comprise any one selected from the following: carbon fiber, glass fiber, aramid fiber, ultrahigh molecular polyethylene fiber, basalt fiber and quartz fiber.

11. The molding method according to claim 1, characterized by, The three-dimensional braiding machine is a ring braiding machine.

12. The molding method according to claim 11, characterized by, The pulling the combined core mold through a three-dimensional braiding machine along the extension direction of the combined core mold further comprises: making the central axis of the combined core mold perpendicular to the braiding plane and at the center of the braiding ring of the three-dimensional braiding machine.

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