Mold for molding composite material aircraft fuselage skin autoclave
By designing a mold with a composite structural functional layer and adjusting the shear internal stress distribution, the deformation and residual stress problems in the autoclave molding process of composite aircraft fuselage skins were solved, the molding accuracy was improved, and the production cost was reduced.
Patent Information
- Application Number
- CN202422681066.3
- Authority / Receiving Office
- CN · China
- Patent Type
- Utility models(China)
- Current Assignee / Owner
- Filing Date
- 2024-11-05
- Publication Date
- 2025-09-19
- Estimated Expiration
- 2034-11-05
AI Technical Summary
In the existing technology, during the autoclave molding process of composite aircraft fuselage skins, the deformation caused by the chemical shrinkage of the resin matrix and the thermal expansion and contraction of the mold does not match, resulting in asymmetric shear stress, affecting molding accuracy and residual stress, and increasing manufacturing costs.
A mold with a composite structural functional layer is designed, including a composite template layer of fiber composite material interlayer and silicone rubber interlayer. The shear internal stress distribution is adjusted to offset the shear stress in the truss bonding area, and the unbonded area is allowed to expand and contract freely, thereby reducing the curing residual stress.
The symmetrical distribution of internal stress on the upper and lower surfaces of the composite material skin is achieved, which reduces curing deformation, improves molding accuracy, and reduces production costs.
Smart Images
Figure CN223354693U_ABST
Abstract
Description
Technical Field
[0001] The utility model relates to a carbon fiber composite material forming device, in particular to a mold for composite material aircraft fuselage skin autoclave forming. Background Art
[0002] Autoclave molding technology has been widely used in the manufacture of composite aerospace components. However, due to chemical shrinkage of the resin matrix and thermal expansion and contraction of the part and mold, mold deformation and part deformation during the composite curing process can be mismatched. This can lead to shear stresses at the bonded surfaces and asymmetric stresses between the upper and lower layers of the component, causing deformation during curing. This is particularly true for composite aircraft fuselage skins with stringer structures on their upper surfaces. Shear stresses also occur at the bonded surfaces between the stringers and the skin, causing more complex deformation and generating residual stresses. Deformation of composite skins currently used in aircraft fuselages hinders assembly. Existing solutions involve precalculating the skin's curing deformation and then compensating for this deformation by reshaping the mold surface through reverse engineering. While this technique can ensure that the skin meets dimensional requirements after deformation, it requires repeated modification and adjustment of the mold surface dimensions. Changes in skin thickness and structure require remanufacturing the mold, significantly increasing manufacturing costs. Furthermore, the high residual stresses in the skin after deformation can also affect its service life. Summary of the Invention
[0003] In response to the above technical problems, the utility model provides a mold for hot-pressing the fuselage skin of composite aircraft. By having a composite structural functional layer, it can improve the curing deformation of the carbon fiber composite material skin caused by the asymmetric shear stress between the upper and lower surfaces, reduce the curing residual stress, and improve the molding accuracy.
[0004] The purpose of the present utility model is achieved as follows: A mold for composite aircraft fuselage skin autoclave molding comprises a metal base plate, a composite template layer disposed on the metal base plate, the composite template layer comprising fiber composite material interlayers and silicone rubber interlayers arranged parallel and staggered, and clamps disposed on the periphery of the metal base plate and the composite template layer.
[0005] Furthermore, the distribution position of the fiber composite material interlayer in the composite template layer corresponds to the position of the bonding area of the inverted "T"-shaped composite material truss on the top surface of the skin.
[0006] Furthermore, the fiber laying directions of the top composite material plate in the fiber composite sandwich and the bottom composite material plate in the inverted "T"-shaped composite material truss correspond to the same direction.
[0007] Furthermore, the longitudinal ribs of the inverted "T"-shaped composite material truss are perpendicular to the plane of the skin.
[0008] Furthermore, the height of the clamp is flush with the top surface of the composite template layer.
[0009] Compared to the prior art, the present invention regulates the shear stress distribution between the composite aircraft fuselage skin's bottom surface and the mold during autoclave curing by designing a composite formwork layer. This offsets the shear stress inflicted by the composite stringers on the top surface of the composite skin. The fiber composite interlayer in the composite formwork layer is positioned to coincide with the bonded area between the skin and stringers, and the fiber layup orientation of the top composite sheet of the fiber composite interlayer aligns with the fiber layup orientation of the bottom composite sheet of the inverted "T"-shaped composite stringers, thereby offsetting the curing shear stress. The top surface of the skin not bonded to the stringers is free to expand and contract during curing. Therefore, a silicone rubber interlayer is designed in the corresponding position of the composite formwork layer. Due to its low modulus, silicone rubber also does not affect the expansion and contraction of the composite skin. This ultimately achieves symmetrical internal stress distribution between the top and bottom surfaces of the composite skin during curing, reducing curing residual stresses and thus minimizing autoclave curing deformation of the composite skin, improving molding precision, and significantly reducing the quality and production cost of the aircraft fuselage skin. BRIEF DESCRIPTION OF THE DRAWINGS
[0010] Figure 1 This is a planar structural diagram of the mold of the utility model;
[0011] Figure 2 This is a side structural diagram of the utility model in use;
[0012] Figure 3 for Figure 2 A schematic diagram of the enlarged cross-section structure at point A;
[0013] Figure 4 This is a three-dimensional structural diagram of the utility model in use;
[0014] In the figure: 1-metal base plate, 2-clamp, 3-composite template layer, 31-fiber composite material interlayer, 32-silicone rubber interlayer; 4-skin, 5-inverted "T"-shaped composite material purlin, 51-bottom composite material plate; 6-top composite material plate. DETAILED DESCRIPTION
[0015] The following structural drawings and embodiments further illustrate the present invention. Figures 1 to 4 A mold for composite aircraft fuselage skin autoclave molding, comprising a metal base plate 1, a composite template layer 3 is arranged on the metal base plate 1, and the composite template layer 3 is composed of a fiber composite material interlayer 31 and a silicone rubber interlayer 32 arranged parallel to each other (such as Figure 1 and Figure 2As shown). The distribution position of the fiber composite sandwich 31 in the composite template layer 3 corresponds to the bonding area position of the inverted "T" shaped composite truss 5 on the top surface of the skin 4. The fiber ply directions of the top composite plate 6 in the fiber composite sandwich 31 and the bottom composite plate 51 in the inverted "T" shaped composite truss 5 are the same (as shown). Figure 3 As shown). The longitudinal ribs of the inverted "T" shaped composite truss 5 are perpendicular to the plane of the skin 4. The outer periphery of the metal base plate 1 and the composite formwork layer 3 is provided with a clamp 2, and the height of the clamp 2 is flush with the top surface of the composite formwork layer 3. The inverted "T" shaped composite truss 5 and the top surface of the composite skin 4 can be cured and bonded by autoclave molding technology. The composite formwork layer 3 and the metal base plate 1 are bonded and fixed by an adhesive. The distribution position of the silicone rubber interlayer 32 corresponds one-to-one to the position of the area on the top surface of the skin 4 that is not bonded to the truss (as shown). Figure 4 shown).
[0016] Example: An implementation case of the present invention is listed below. A mold for hot-pressing the fuselage skin of a composite aircraft is composed of a metal base plate 1, a composite template layer 3 and a clamp 2. The overall size of the mold is 1200×1000×50 mm, wherein the thickness of the metal base plate 1 is 49.2 mm, and the thickness of the composite template layer 3 is 0.8 mm. The metal base plate 1 is made of an alloy steel material with fast thermal conductivity and not easy to deform. The composite template layer 3 is bonded to the metal base plate 1 with a high-temperature resistant adhesive. The composite template layer 3 is composed of a fiber composite material interlayer 31 with a total thickness of 0.8 mm and a silicone rubber interlayer 32 with a thickness of 0.8 mm, which are arranged alternately with each other, wherein the fiber composite material interlayer 31 is formed by curing four layers of carbon fiber epoxy resin-based composite material with a thickness of 0.2 mm. The fiber angles are 0°, 90°, 90°, and 0°, respectively. There are four fiber composite interlayers 31, with a width of 62 mm and a spacing of 150 mm. The width and spacing of the fiber composite interlayers 31 are aligned with the distribution of the bonding area between the inverted "T"-shaped composite stringers 5 and the skin 4. Furthermore, the fiber layup orientation of the top composite sheet 6 of the fiber composite interlayer 31 and the bottom composite sheet 51 of the inverted "T"-shaped composite stringers 5 align, both adopting a 0° layup structure. This ensures that the shear internal stresses imposed on the skin by the fiber composite interlayer 31 and the inverted "T"-shaped composite stringers 5 during curing are symmetrical and equal in the upper and lower parts. The areas of the skin 4 not bonded to the stringers are free to expand and contract, so a silicone rubber interlayer 32 is designed in the corresponding position of the composite formwork layer 3 of the mold. This is because silicone rubber has a low modulus and does not restrict the curing shrinkage of the skin 4 during curing. Ultimately, a symmetrical distribution of internal stresses on the upper and lower surfaces of the composite skin is achieved during curing. In addition, an integral clamp 2 is installed around the mold's periphery, its height aligned with the top surface of the composite formwork layer 3. This protects the composite formwork layer 3 and increases the mold's service life. Tests have shown that the maximum curing deformation of the composite skin molded using this invention is reduced from 2.3 mm to 0.7 mm, a 70% decrease. Autoclave curing simulations show that the composite skin's internal stress during curing is reduced from 56 MPa to 37 MPa, a 34% decrease.
Claims
1. A mold for composite aircraft fuselage skin autoclave molding, comprising a metal base plate, characterized in that: A composite template layer is arranged on the metal bottom plate. The composite template layer is composed of fiber composite material interlayers and silicone rubber interlayers arranged parallel and staggered. Clamps are arranged on the periphery of the metal bottom plate and the composite template layer.
2. The mold for composite aircraft fuselage skin autoclave molding according to claim 1, characterized in that: The distribution position of the fiber composite material interlayer in the composite template layer corresponds to the position of the bonding area of the inverted "T"-shaped composite material truss on the top surface of the skin.
3. The mold for composite aircraft fuselage skin autoclave molding according to claim 1, characterized in that: The fiber laying directions of the top composite material plate in the fiber composite material sandwich and the bottom composite material plate in the inverted "T"-shaped composite material truss correspond to the same direction.
4. The mold for composite aircraft fuselage skin autoclave molding according to claim 2 or 3, characterized in that: The longitudinal ribs of the inverted "T"-shaped composite material truss are perpendicular to the plane of the skin.
5. The mold for composite aircraft fuselage skin autoclave molding according to claim 1, characterized in that: The height of the clamp is flush with the top surface of the composite template layer.