Composite material integral co-curing fuselage and preparation method thereof

Abstract

The invention relates to a composite material integral co-curing machine body and a preparation method thereof, and the method comprises the following steps: according to the structural design of the machine body, paving parts such as a bulkhead, a beam and a long honing on corresponding molds; then the mold is assembled to lay the skin; and finally, the mold is closed, the mold is put into a drying oven or an autoclave to be cured, and all the parts adhere to one another in the one-time curing process to form a continuous overall structure. Compared with a traditional mode that all parts are independently formed and then assembled, the fuselage structure manufactured through the method has the advantages that use of connecting pieces for riveting, screw joint, glue joint and the like is reduced, the weight of the fuselage structure is reduced, and fuel consumption is reduced; meanwhile, all the parts form a continuous structure in the co-curing process, strength loss caused by secondary connection is avoided, and the overall strength and bearing capacity of the fuselage are improved; in addition, according to the method, the manufacturing technological process is simplified, the machining cost is reduced, and the production efficiency is improved.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of aviation manufacturing, and particularly relates to a composite material integrally co-cured fuselage and a preparation method. BACKGROUND

[0002] In the field of aviation manufacturing, the performance index of the aircraft fuselage is crucial to the whole machine. The traditional aircraft fuselage is mostly made of metal materials, and various components are connected through riveting, screwing, gluing and the like, which has problems such as heavy weight, complex structure, complicated assembly process and the like. Although composite materials have been applied to fuselage manufacturing, the existing technology still has deficiencies in reducing weight and improving strength, and the method of separately forming and assembling various components not only increases the weight of connecting parts, but also affects the overall structural strength due to stress concentration at the connecting points. Meanwhile, the manufacturing process is relatively complex, the production efficiency is low, and the cost is high. Therefore, there is an urgent need for a new type of composite material integrally co-cured fuselage structure and preparation method to solve the above problems. SUMMARY

[0003] In view of this, the application discloses a composite material integrally co-cured fuselage and a preparation method, which are used to solve the existing problems.

[0004] The application adopts the following technical scheme: A preparation method of a composite material integrally co-cured fuselage, the method comprising the following steps: S1. Prepolymer preparation: layer laying and pasting are performed on a bulkhead mold, a beam mold and a stringer mold, and vacuum compression and pre-curing are performed to obtain a bulkhead prepolymer, a beam prepolymer and a stringer prepolymer; S2. Combined mold: the bulkhead mold, the beam mold, the stringer mold and a core mold are combined, and are penetrated together through a long rod to obtain a combined mold; S3. The bulkhead prepolymer, the beam prepolymer and the stringer prepolymer are assembled with the combined mold to obtain a fuselage skin inner mold; S4. Fuselage skin preparation: layer laying or winding is performed on the fuselage skin inner mold, and vacuum compression is performed to obtain a fuselage skin; S5. Closing mold: the fuselage skin is shaped to obtain a fuselage prototype; S6. Co-curing forming: the fuselage prototype is heated and cured, and the various components are adhered to each other to obtain a fuselage continuous integral structure; S7. Mold removal: a maintenance port cover is arranged on the skin of the fuselage continuous integral structure, the bulkhead mold, the beam mold and the stringer mold are removed through the maintenance port cover, and the composite material integrally co-cured fuselage is obtained.

[0005] Further, the structure of the bulkhead prepolymer, the beam prepolymer and the stringer prepolymer in step S1 is a composite material laminated structure or a sandwich structure.

[0006] Further, the shape of the frame pre-polymer, the beam pre-polymer and the longeron pre-polymer in step S1 includes C-shaped, T-shaped, Z-shaped, Y-shaped or Ω-shaped.

[0007] Further, the type of the core mold in step S2 includes a soluble core mold, a silica gel core mold, a bladder core mold or a spliced and detachable metal core mold.

[0008] Further, the non-circular section at the connection between the longeron and the frame plate is used to position the molds, and the positioning further includes using stepped sections to lock the molds.

[0009] Further, the non-circular section includes a square, a polygon or an ellipse.

[0010] Further, the size of the stepped section is 1 nm to 1 m.

[0011] Further, the structure of the fuselage skin in step S4 includes a composite laminated structure or a sandwich structure.

[0012] Further, the temperature curing in step S6 includes controlling the curing temperature, the curing time and the curing pressure according to a curing process curve, wherein the curing temperature is 20℃-200℃, the curing time is 1min-500min, and the curing pressure is 0bar-10bar.

[0013] The application also provides a composite integral co-cured fuselage prepared by the above method.

[0014] The beneficial effects of the application are as follows: Compared with the traditional method of separately forming each component and then assembling, the use of rivets, screws, adhesives and other connecting parts is reduced, thereby reducing the weight of the fuselage structure, reducing the fuel consumption of the aircraft, and improving the range and fuel economy.

[0015] The components form a continuous structure during the co-curing process, avoiding the strength loss caused by secondary connection, improving the overall strength and carrying capacity of the fuselage, reducing stress concentration, and prolonging the service life of the aircraft.

[0016] The manufacturing process is simplified, the secondary assembly and connection processes are reduced, the production cycle is shortened, the production efficiency is improved, the manufacturing cost is reduced, and it is beneficial to realize automatic and batch production. BRIEF DESCRIPTION OF DRAWINGS

[0017] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings needed to be used in the embodiments will be briefly introduced as follows. Obviously, the drawings in the following description only constitute some embodiments of the present application, and other drawings can be obtained by those skilled in the art without any creative effort on the basis of these drawings.

[0018] Figure 1 is a preparation process flowchart of the present application; Figure 2 is a partition frame structure schematic diagram of the present application; Figure 3 is a beam / stringer structure schematic diagram of the present application; Figure 4 is a skin structure schematic diagram of the present application; Figure 5 is a combined mold schematic diagram of the present application; Figure 6 is a skin layering schematic diagram of the present application; Figure 7 is a closing mold schematic diagram of the present application; Figure 8 is a frame plate layering process schematic diagram of the present application; Figure 9 is a beam / stringer layering process schematic diagram of the present application; Figure 10 is a skin layering schematic diagram of the present application. DETAILED DESCRIPTION

[0019] In order to better understand the technical solutions of the present application, the embodiments of the present application will be described in detail below with reference to the drawings.

[0020] It should be clear that the described embodiments are only some of the embodiments of the present application, but not all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without any creative effort belong to the scope of protection of the present application.

[0021] As shown in Figure 1 , the present application provides a preparation method of a composite material overall co-cured fuselage, the method comprising the following steps: S1. Preparation of prepolymer: layering and pasting on the partition frame mold, beam mold and stringer mold, and vacuumizing and compacting, pre-solidifying to obtain partition frame prepolymer, beam prepolymer and stringer prepolymer; S2. Combined mold: combining the partition frame mold, beam mold, stringer mold and core mold, and penetrating together through the long rod to obtain a combined mold, as shown in Figure 5 ; S3. Assembling the frame pre-polymer, beam pre-polymer, stringer pre-polymer with the combined mold to obtain the inner mold of the fuselage skin, as shown in Figure 4 ; S4. Preparation of the fuselage skin: placing the inner mold of the fuselage skin on a rotatable support, performing laying or winding on the inner mold of the fuselage skin, and vacuumizing and compacting to obtain the fuselage skin, as shown in Figure 6 ; S5. Closing the mold: moving the fuselage skin to the outer mold of the fuselage skin to shape to obtain the fuselage prototype, as shown in Figure 7 (1) and (2); S6. Co-curing forming: moving the fuselage prototype to an oven or autoclave for temperature rising and curing, and adhering each part to each other to obtain the continuous overall structure of the fuselage; S7. Mold removal: opening the maintenance port cover of the equipment, load or fuel tank on the skin of the continuous overall structure of the fuselage through the maintenance port cover to remove the mold to obtain the overall co-cured fuselage.

[0022] The fuselage structure prepared by the preparation method of the application is composed of an integrated skin, frame, beam and stringer. The frame, beam and stringer serve as a load-bearing framework, and the skin is relatively thin for maintaining the aerodynamic shape. The equipment maintenance port cover is locally arranged to facilitate mold removal. Compared with the traditional method of separately forming each part and then assembling, the method for preparing the fuselage structure reduces the use of rivets, screws, adhesives and other connecting parts, reduces the weight of the fuselage structure, reduces fuel consumption, improves the range and fuel economy; the continuous structure of each part is formed in the co-curing process, avoiding the strength loss caused by secondary connection, improving the overall strength and load-carrying capacity of the fuselage, reducing stress concentration, and prolonging the service life of the aircraft; the manufacturing process is simplified, the secondary assembly and connection processes are reduced, the production cycle is shortened, the production efficiency is improved, the manufacturing cost is reduced, and batch production is facilitated.

[0023] The structural form of the frame pre-polymer, beam pre-polymer and stringer pre-polymer includes C-shaped, T-shaped, Z-shaped, Y-shaped or Ω-shaped, and other shapes are also possible, which are not limited by the application. In step S2, a non-circular section positioning frame mold is used at the connection between the long rod and the frame plate. The positioning further includes using a stepped section to hold the frame mold.

[0024] Further, the non-circular section includes a square, polygon or ellipse.

[0025] Further, the frame pre-polymer, beam pre-polymer and stringer pre-polymer in step S1 are of a composite laminated structure or a sandwich structure.

[0026] Further, the type of the core mold in step S2 includes a soluble core mold, a silica gel core mold, an air bag core mold or a spliced and detachable metal core mold.

[0027] Further, the step difference size of the step difference section is 1 nm-1 m.

[0028] Further, the structure of the fuselage skin in step S4 comprises a composite laminated structure or a sandwich structure.

[0029] Further, the step S6 curing comprises: according to the curing process curve, controlling the curing temperature, curing time and curing pressure parameters, wherein the curing temperature is 20℃-200℃, the curing time is 1min-500min, and the curing pressure is 0bar-10bar.

[0030] The following is described by specific examples.

[0031] Example 1 The structure of a certain fuselage is designed with a low-cost sandwich structure, in which the bulkhead is a C-shaped sandwich structure as shown in Figure 2 , the stringer is a T-shaped sandwich structure as shown in Figure 3 , and the skin is a laminated sandwich structure as shown in Figure 4 . The bulkhead manufacturing process and structure are shown in Figure 8 , the inside carbon fiber composite panel 2, the sandwich aviation panel 3 and the outside carbon fiber composite panel 4 are sequentially laid on the bulkhead mold 1, and the bulkhead pre-polymer is obtained by vacuum pre-curing; the stringer manufacturing process is shown in Figure 9 , the inside carbon fiber composite panels 7 and 8 are sequentially laid on the stringer mold 5, the sandwich aviation panel is laid in the middle and the mold is clamped, the carbon fiber composite panel 6 is laid on the top, and the stringer pre-polymer is obtained by vacuum pre-curing; the assembly mold process is shown in Figure 5 , the bulkhead pre-polymer, the stringer pre-polymer, the bulkhead mold, the stringer mold and the splicing metal core mold are combined together and connected by a through rod; the skin manufacturing process is shown in Figure 10 , the inside skin 10, the honeycomb core 11 and the outside skin 12 are respectively laid on the assembly mold, and the skin pre-polymer is obtained by vacuum pre-curing; finally, the mold is closed, the skin pre-polymer is moved to the fuselage skin outer mold for shaping, and is placed in a hot press tank, and the curing is completed under a curing pressure of 6bar according to a curing schedule of 80℃ / 1h+130℃ / 2h, and the integral molding is formed, and the mold is demolded after being cooled to room temperature.

[0032] Example 2 The fuselage structure form adopts a traditional laminated structure design, wherein the form of the bulkhead structure is an I-shaped laminated structure, the form of the beam structure is an Ω-shaped laminated structure, and the form of the skin structure is a laminated structure. The manufacturing of the bulkhead is performed on a bulkhead silicone rubber mold, the manufacturing of the beam is performed on a PMI foam core mold, after the laying and pasting of each part on the corresponding mold is completed, vacuumizing and compaction treatment is performed, then the mold is combined, the bulkhead is filled with a balloon core mold, all frame plates are connected through a through rod, the skin is laid and pasted on the support, after the laying and pasting of the skin is completed, the skin is transferred to the closing mold and placed in an oven, curing is completed under a curing pressure of 10 bar, according to a curing system of 70℃ / 1h+120℃ / 2h+150℃ / 1h, and the skin is integrally formed, after the temperature is reduced to room temperature, demolding treatment is performed, the beam core mold is not taken out, and the remaining mold is taken out.

[0033] The above has carried on the detailed introduction to the embodiment of the present application, the principle and the implementation mode of the present application have been described in this article by applying specific examples, the above embodiment is only for helping understanding the method of the present application and its core idea, simultaneously, for the general technical personnel of the field, according to the idea of the present application, there will be changes in the specific implementation mode and the application range, and the above-mentioned description should not be understood as the limitation of the present application.

Claims

1. A method for preparing a composite material integral co-cured fuselage, characterized in that, The method includes the following steps: S1. Prepolymer preparation: lay-up on the frame mold, beam mold and stringer mold, and vacuum compact and pre-cur to obtain the frame prepolymer, beam prepolymer and stringer prepolymer; S2. Combined mold: The frame mold, beam mold, stringer mold and core mold are combined and connected together by a long rod to obtain the combined mold; S3. Assemble the prepolymer of the partition frame, the prepolymer of the beam, the prepolymer of the stringer and the combined mold to obtain the inner mold of the fuselage skin; S4. Preparation of fuselage skin: The fuselage skin is laid or wrapped on the inner mold of the fuselage skin and then vacuum-pressed to obtain the fuselage skin; S5. Mold assembly: Shape the fuselage skin to obtain the fuselage prototype; S6. Co-curing molding: The prototype of the fuselage is heated and cured, and the various parts are bonded together to obtain a continuous integral structure of the fuselage; S7. Mold Removal: A maintenance port cover is provided on the skin of the continuous integral structure of the machine body. The frame mold, beam mold and stringer mold are removed through the maintenance port cover to obtain the integral co-cured composite material machine body.

2. The method according to claim 1, characterized in that, The prepolymer of the partition frame, the prepolymer of the beam, and the prepolymer of the stringer mentioned in step S1 are composite material laminated structures or sandwich structures.

3. The method according to claim 1, characterized in that, The structural forms of the prepolymer frame, prepolymer beam, and prepolymer stringer mentioned in step S1 include: C-shaped, T-shaped, Z-shaped, Y-shaped, or Ω-shaped.

4. The method according to claim 1, characterized in that, The types of core molds mentioned in step S2 include: soluble core molds, silicone core molds, airbag core molds, or splicing and disassembling metal core molds.

5. The method according to claim 1, characterized in that, In step S2, the connection between the long rod and each mold is positioned using a non-circular cross section. The positioning also includes using a stepped cross section to hold each mold in place.

6. The method according to claim 5, characterized in that, The non-circular cross-section includes square, polygonal, or elliptical shapes.

7. The method according to claim 5, characterized in that, The step size of the step cross section is 1 nm to 1 m.

8. The method according to claim 1, characterized in that, The structure of the fuselage skin described in step S4 includes a composite material laminate structure or a sandwich structure.

9. The method according to claim 1, characterized in that, Step S6, the heating and curing process, includes controlling the curing temperature, curing time, and curing pressure parameters according to the curing process curve, wherein the curing temperature is 20℃-200℃, the curing time is 1min-500min, and the curing pressure is 0bar-10bar.

10. A composite material integral co-cured fuselage, characterized in that, The integral co-cured body is obtained by the preparation method described in any one of claims 1-9.

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