CN117774376A - Forming method of large-size aircraft central wing skeleton co-cementing autoclave - Google Patents
Forming method of large-size aircraft central wing skeleton co-cementing autoclave Download PDFInfo
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- CN117774376A CN117774376A CN202311682328.1A CN202311682328A CN117774376A CN 117774376 A CN117774376 A CN 117774376A CN 202311682328 A CN202311682328 A CN 202311682328A CN 117774376 A CN117774376 A CN 117774376A
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- 238000000034 method Methods 0.000 title claims abstract description 32
- 229920001971 elastomer Polymers 0.000 claims abstract description 37
- 238000000465 moulding Methods 0.000 claims abstract description 12
- 239000002131 composite material Substances 0.000 claims description 18
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 claims description 9
- 229910052799 carbon Inorganic materials 0.000 claims description 9
- 229920000049 Carbon (fiber) Polymers 0.000 claims description 7
- 239000002313 adhesive film Substances 0.000 claims description 7
- 239000004917 carbon fiber Substances 0.000 claims description 7
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 claims description 7
- 238000010438 heat treatment Methods 0.000 claims description 6
- 238000004806 packaging method and process Methods 0.000 claims description 6
- 238000007711 solidification Methods 0.000 claims description 6
- 230000008023 solidification Effects 0.000 claims description 6
- 238000005538 encapsulation Methods 0.000 claims description 3
- 239000004744 fabric Substances 0.000 claims description 3
- 229920002379 silicone rubber Polymers 0.000 claims description 3
- 238000004026 adhesive bonding Methods 0.000 claims description 2
- 238000004519 manufacturing process Methods 0.000 abstract description 6
- 239000010410 layer Substances 0.000 description 8
- 230000007797 corrosion Effects 0.000 description 2
- 238000005260 corrosion Methods 0.000 description 2
- 238000013461 design Methods 0.000 description 2
- 239000003292 glue Substances 0.000 description 2
- 239000011157 advanced composite material Substances 0.000 description 1
- 230000004075 alteration Effects 0.000 description 1
- 238000006664 bond formation reaction Methods 0.000 description 1
- 238000004891 communication Methods 0.000 description 1
- 238000009826 distribution Methods 0.000 description 1
- 239000011229 interlayer Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000012545 processing Methods 0.000 description 1
- 238000011160 research Methods 0.000 description 1
- 238000010008 shearing Methods 0.000 description 1
- 230000035939 shock Effects 0.000 description 1
- 230000003313 weakening effect Effects 0.000 description 1
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Abstract
The invention discloses a method for forming a large-size aircraft central wing skeleton co-cementing autoclave, and relates to the technical field of aerospace. The wing framework is formed by co-cementing and curing a plurality of I-shaped framework beams which are parallel to each other and a plurality of single/double-sided framework ribs which are perpendicular to the framework beams, the I-shaped framework beams are completed through curing, and the framework ribs are preformed. Paving and curing the I-shaped skeleton beam in part forming; the skeleton rib is laid on the rubber soft mould. The preformed skeleton rib is positioned and combined with a plurality of skeleton beams through the positioning holes and solidified. The molding method has the advantages of easy demolding and low cost, not only can ensure the molding precision and nondestructive quality of the framework, but also can reduce the manufacturing cost and the manufacturing period of the parts.
Description
Technical Field
The invention relates to the technical field of advanced composite material manufacturing, in particular to a method for forming a large-size aircraft central wing skeleton co-cementing autoclave.
Background
Unmanned aerial vehicles gradually become important research hot spots in the aviation field in recent years, and carbon fiber composite materials have the characteristics of high specific strength, high specific modulus, fatigue resistance, corrosion resistance and the like, and are widely applied to the aerospace field. The lightweight design can effectively improve unmanned aerial vehicle's flight performance.
The wing is used as a key structural member of the unmanned aerial vehicle and mainly provides lifting force for the unmanned aerial vehicle, so that the wing needs to have good shock resistance. The wing skeleton mainly comprises trusses and ribs. In general, the wing skeleton forming process adopts a secondary cementing or riveting mode, namely, the beam and the rib are respectively solidified and formed, and then the beam and the rib are assembled and connected by adopting an adhesive film or rivets. But riveting the spar to the rib using standard components increases the weight of the wing skeleton. In addition, the composite material is subjected to mechanical processing to generate serious damage and weakening, and the interlayer shearing performance of the composite material is reduced, so that the connection mode of the composite material is mainly gluing.
At present, the composite material forming process is mainly divided into three cementing forms: (1) secondary cementing (2) co-cementing (3) co-curing. Among them, the secondary bonding mode is most widely used, but has the disadvantages of long molding cycle, multiple assembly times and the like. The co-curing molding process cannot be applied to complex wing skeleton design and manufacture.
Disclosure of Invention
The invention aims to provide a molding method of a large-size aircraft center wing skeleton co-glued autoclave, which has the advantages of easy demolding and low cost, can ensure the molding precision and nondestructive quality of the skeleton, simultaneously reduces the manufacturing cost of parts, and expands the application of composite materials in the field of unmanned aerial vehicles.
In order to solve the technical problems, the invention provides a method for forming a large-size airplane central wing skeleton co-cementing autoclave, which comprises the following steps:
step A: dividing the structure of the central wing framework composite material component into an I-shaped framework beam and framework ribs according to the number of the central wing framework composite material component, and designing a corresponding dividing and paving tool according to the structural forms of the divided I-shaped framework beam and framework ribs; the I-shaped skeleton beam is solidified and formed by adopting a part forming tool, and the skeleton rib is preformed by adopting a rubber soft mold in a prepressing manner;
and (B) step (B): forming an I-shaped skeleton beam, paving prepreg on a part forming die between two side baffles, carrying out combined positioning through positioning guide pins after paving, paving C-shaped prepreg of the I-shaped skeleton beam after positioning, and compacting by adopting a vacuum bag process after paving every 3 layers of prepregs;
step C: forming single-sided framework ribs, paving framework rib prepregs on a single-sided rubber soft mold, and compacting each 3 layers of framework rib prepregs by adopting a vacuum bag process after paving;
step D: forming double-sided skeleton ribs, paving skeleton rib prepregs on an upper rubber mold body and a lower rubber mold body, closing the mold through positioning guide pins, and compacting by adopting a vacuum bag process; paving prepreg on the side edges after die assembly, wherein the prepreg is compacted by adopting a vacuum bag process after the 3 layers of framework rib prepregs are paved;
step E: combining, namely combining and positioning the cured I-shaped skeleton beam and the skeleton rib through positioning guide pins, and packaging;
step F: co-cementing and solidifying, namely paving a layer of adhesive film on the cementing area of the I-shaped framework beam and the framework rib, and carrying out die assembly, encapsulation and solidification on the I-shaped framework beam and the framework rib after solidification according to the positioning groove;
step G: and (5) demolding and molding, and removing the rubber soft mold on the framework rib after the co-cementing is finished, so as to obtain the central wing framework composite material component.
Preferably, in the step of forming the I-shaped skeleton beam, the I-shaped skeleton beam is divided into three parts for paving; firstly, paving C-shaped prepreg on the lower side of an I-shaped skeleton beam on a part forming tool; secondly, paving the two side edge strips on the side baffle plates, positioning and combining according to the positioning guide pins after the paving is finished, and filling carbon twisted wires in the triangular gaps on the two sides; thirdly, after filling the carbon twisted wires, paving the upper C-shaped prepreg, and packaging and solidifying the part on a part forming tool after paving.
Preferably, the curing parameters of the i-beam: the room temperature is heated to 80 ℃, the heating rate is less than or equal to 2 ℃/min, the temperature is kept for 30min, the temperature difference is 80+/-3 ℃, the pressure is increased to 0.6Mpa after the temperature is kept, the temperature is heated to 100 ℃ at the speed of 1.5 ℃/min, the temperature is heated to 125 ℃ at the speed of 0.8 ℃/min, the temperature is kept for 90min, and the temperature deviation is 125+/-6 ℃.
Preferably, in the step of forming the skeleton rib, the skeleton rib is divided into two types, one type is a single-sided box shape, and the other type is a double-sided box shape;
paving a single-sided box-shaped framework rib on the single-sided rubber soft mold to paste framework rib prepreg;
the two-sided box-shaped framework rib is divided into three parts for paving, firstly, prepregs of the framework ribs on the two sides are respectively paved on corresponding rubber soft molds, and the rubber soft molds on the two sides are combined according to the positioning holes after the paving is finished; secondly, filling carbon twisted wires in triangular gaps on two sides; thirdly, paving the edge strips after filling.
Preferably, in the co-cementing curing step, the pressure of the process parameters is 0.3-0.8MPa, the curing temperature is 125+/-6 ℃, the curing time is 120-180min, and the pressure is gradually increased from 0.1MPa to the applied pressure when the pressure is applied.
Preferably, in the co-glue curing step, the curing parameters: the room temperature is heated to 80 ℃, the heating rate is less than or equal to 2 ℃/min, the temperature is kept for 30min, the temperature difference is 80+/-3 ℃, the pressure is increased to 0.5Mpa after the temperature is kept, the temperature is heated to 100 ℃ at the speed of 1.25 ℃/min, the temperature is heated to 125 ℃ at the speed of 0.8 ℃/min, the temperature is kept for 90min, and the temperature deviation is 125+/-6 ℃.
Preferably, the C-shaped prepreg and the framework rib prepreg of the I-shaped framework beam are all in a carbon fiber fabric, carbon fiber unidirectional tape and adhesive film combined mode.
Preferably, the rubber soft mold is made of flexible high-temperature-resistant rubber, the shape of the rubber soft mold is a box-shaped structure consistent with the skeleton rib, the rubber soft mold is made of silicon rubber, and the lowest temperature resistance is 190 ℃, and the elongation is more than or equal to 300%.
Compared with the prior art, the invention aims to complete the manufacturing process of the central wing framework of the co-bonded composite material by adopting a rubber soft film and part forming tool and assisting an autoclave forming mode, thereby realizing the following requirements:
1. the part forming period is shortened, and the part can entering and assembling times are reduced;
2. the weight increase of parts caused by riveting and fixing the standard parts is avoided;
3. the problem of assembly tolerance distribution is solved;
4. the problem of contact corrosion between the standard component and the composite material is avoided.
Drawings
FIG. 1 is a schematic view of an I-beam for use in the present invention;
FIG. 2 is a schematic view of a double-sided skeletal rib structure provided by the present invention;
FIG. 3 is a schematic view of the double-sided skeletal rib molding provided by the invention;
FIG. 4 is a schematic illustration of the co-bonding formation of an I-shaped framework beam and framework ribs provided by the invention;
fig. 5 is a schematic structural view of a central wing framework composite member provided by the present invention.
Detailed Description
The invention is described in further detail below with reference to the attached drawings and specific examples. Advantages and features of the invention will become more apparent from the following description and from the claims. It should be noted that the drawings are in a very simplified form and are all to a non-precise scale, merely for convenience and clarity in aiding in the description of embodiments of the invention.
In the description of the present invention, it should be understood that the terms "center", "longitudinal", "lateral", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc. indicate orientations or positional relationships based on the orientations or positional relationships shown in the drawings, are merely for convenience in describing the present invention and simplifying the description, and do not indicate or imply that the devices or elements referred to must have a specific orientation, be configured and operated in a specific orientation, and thus should not be construed as limiting the present invention.
In the description of the present invention, it should be noted that, unless explicitly specified and limited otherwise, the terms "mounted," "connected," and "connected" are to be construed broadly, and may be either fixedly connected, detachably connected, or integrally connected, for example; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can be communication between two elements. The specific meaning of the above terms in the present invention can be understood by those of ordinary skill in the art in a specific case.
Examples
The invention provides a method for forming a large-size airplane central wing skeleton co-bonding autoclave, referring to fig. 1-5, comprising the following steps:
step A: dividing the structure of the central wing framework composite material component into an I-shaped framework beam and framework ribs according to the number of the central wing framework composite material component, and designing a corresponding dividing and paving tool according to the structural forms of the divided I-shaped framework beam and framework ribs; the I-shaped skeleton beam is solidified and formed by adopting a part forming tool, and the skeleton rib is preformed by adopting a rubber soft mold in a prepressing manner;
and (B) step (B): forming an I-shaped skeleton beam, paving prepreg on a part forming die between two side baffles, carrying out combined positioning through positioning guide pins after paving, paving C-shaped prepreg of the I-shaped skeleton beam after positioning, and compacting by adopting a vacuum bag process after paving every 3 layers of prepregs;
step C: forming single-sided framework ribs, paving framework rib prepregs on a single-sided rubber soft mold, and compacting each 3 layers of framework rib prepregs by adopting a vacuum bag process after paving;
step D: forming double-sided skeleton ribs, paving skeleton rib prepregs on an upper rubber mold body and a lower rubber mold body, closing the mold through positioning guide pins, and compacting by adopting a vacuum bag process; paving prepreg on the side edges after die assembly, wherein the prepreg is compacted by adopting a vacuum bag process after the 3 layers of framework rib prepregs are paved;
step E: combining, namely combining and positioning the cured I-shaped skeleton beam and the skeleton rib through positioning guide pins, and packaging;
step F: co-cementing and solidifying, namely paving a layer of adhesive film on the cementing area of the I-shaped framework beam and the framework rib, and carrying out die assembly, encapsulation and solidification on the I-shaped framework beam and the framework rib after solidification according to the positioning groove;
step G: and (5) demolding and molding, and removing the rubber soft mold on the framework rib after the co-cementing is finished, so as to obtain the central wing framework composite material component.
Specifically, in the step of forming the I-shaped skeleton beam, the I-shaped skeleton beam is divided into three parts for paving; firstly, paving C-shaped prepreg on the lower side of an I-shaped skeleton beam on a part forming tool; secondly, paving the two side edge strips on the side baffle plates, positioning and combining according to the positioning guide pins after the paving is finished, and filling carbon twisted wires in the triangular gaps on the two sides; thirdly, after filling the carbon twisted wires, paving the upper C-shaped prepreg, and packaging and solidifying the part on a part forming tool after paving.
Specifically, curing parameters of the i-beam: the room temperature is heated to 80 ℃, the heating rate is less than or equal to 2 ℃/min, the temperature is kept for 30min, the temperature difference is 80+/-3 ℃, the pressure is increased to 0.6Mpa after the temperature is kept, the temperature is heated to 100 ℃ at the speed of 1.5 ℃/min, the temperature is heated to 125 ℃ at the speed of 0.8 ℃/min, the temperature is kept for 90min, and the temperature deviation is 125+/-6 ℃.
Specifically, in the skeleton rib forming step, the skeleton ribs are divided into two types, one type is a single-sided box shape, and the other type is a double-sided box shape; and paving the single-sided box-shaped framework rib on the single-sided rubber soft mold to paste the framework rib prepreg. The two-sided box-shaped framework rib is divided into three parts for paving, firstly, prepregs of the framework ribs on the two sides are respectively paved on corresponding rubber soft molds, and the rubber soft molds on the two sides are combined according to the positioning holes after the paving is finished; secondly, filling carbon twisted wires in triangular gaps on two sides; thirdly, paving the edge strips after filling.
Specifically, in the co-cementing curing step, the pressure of the technological parameters is 0.3-0.8MPa, the curing temperature is 125+/-6 ℃, the curing time is 120-180min, and when the pressure is applied, the pressure is gradually increased from 0.1MPa to the applied pressure, so that uneven stress of the rubber soft mold caused by direct pressure application is prevented.
In some embodiments, in the co-glue curing step, the curing parameters: the room temperature is heated to 80 ℃, the heating rate is less than or equal to 2 ℃/min, the temperature is kept for 30min, the temperature difference is 80+/-3 ℃, the pressure is increased to 0.5Mpa after the temperature is kept, the temperature is heated to 100 ℃ at the speed of 1.25 ℃/min, the temperature is heated to 125 ℃ at the speed of 0.8 ℃/min, the temperature is kept for 90min, and the temperature deviation is 125+/-6 ℃.
Specifically, the C-shaped prepreg and the framework rib prepreg of the I-shaped framework beam are all in a carbon fiber fabric, carbon fiber unidirectional tape and adhesive film combined mode.
Specifically, the rubber soft mold is made of flexible high-temperature-resistant rubber, the appearance of the rubber soft mold is a box-shaped structure consistent with the skeleton rib, the rubber soft mold is made of silicon rubber, and the lowest temperature resistance is 190 ℃, and the elongation is more than or equal to 300%.
The above description is only illustrative of the preferred embodiments of the present invention and is not intended to limit the scope of the present invention, and any alterations and modifications made by those skilled in the art based on the above disclosure shall fall within the scope of the appended claims.
Claims (8)
1. The molding method of the large-size airplane central wing skeleton co-cementing autoclave is characterized by comprising the following steps of:
step A: dividing the structure of the central wing framework composite material component into an I-shaped framework beam and framework ribs according to the number of the central wing framework composite material component, and designing a corresponding dividing and paving tool according to the structural forms of the divided I-shaped framework beam and framework ribs; the I-shaped skeleton beam is solidified and formed by adopting a part forming tool, and the skeleton rib is preformed by adopting a rubber soft mold in a prepressing manner;
and (B) step (B): forming an I-shaped skeleton beam, paving prepreg on a part forming die between two side baffles, carrying out combined positioning through positioning guide pins after paving, paving C-shaped prepreg of the I-shaped skeleton beam after positioning, and compacting by adopting a vacuum bag process after paving every 3 layers of prepregs;
step C: forming single-sided framework ribs, paving framework rib prepregs on a single-sided rubber soft mold, and compacting each 3 layers of framework rib prepregs by adopting a vacuum bag process after paving;
step D: forming double-sided skeleton ribs, paving skeleton rib prepregs on an upper rubber mold body and a lower rubber mold body, closing the mold through positioning guide pins, and compacting by adopting a vacuum bag process; paving prepreg on the side edges after die assembly, wherein the prepreg is compacted by adopting a vacuum bag process after the 3 layers of framework rib prepregs are paved;
step E: combining, namely combining and positioning the cured I-shaped skeleton beam and the skeleton rib through positioning guide pins, and packaging;
step F: co-cementing and solidifying, namely paving a layer of adhesive film on the cementing area of the I-shaped framework beam and the framework rib, and carrying out die assembly, encapsulation and solidification on the I-shaped framework beam and the framework rib after solidification according to the positioning groove;
step G: and (5) demolding and molding, and removing the rubber soft mold on the framework rib after the co-cementing is finished, so as to obtain the central wing framework composite material component.
2. The method for forming a large-size aircraft center wing skeleton co-bonded autoclave according to claim 1, wherein in the step of forming the i-shaped skeleton beam, the i-shaped skeleton beam is divided into three parts for paving; firstly, paving C-shaped prepreg on the lower side of an I-shaped skeleton beam on a part forming tool; secondly, paving the two side edge strips on the side baffle plates, positioning and combining according to the positioning guide pins after the paving is finished, and filling carbon twisted wires in the triangular gaps on the two sides; thirdly, after filling the carbon twisted wires, paving the upper C-shaped prepreg, and packaging and solidifying the part on a part forming tool after paving.
3. The method for forming a large-size aircraft center rib co-bonded autoclave of claim 2, wherein the curing parameters of the i-shaped rib frame beam: the room temperature is heated to 80 ℃, the heating rate is less than or equal to 2 ℃/min, the temperature is kept for 30min, the temperature difference is 80+/-3 ℃, the pressure is increased to 0.6Mpa after the temperature is kept, the temperature is heated to 100 ℃ at the speed of 1.5 ℃/min, the temperature is heated to 125 ℃ at the speed of 0.8 ℃/min, the temperature is kept for 90min, and the temperature deviation is 125+/-6 ℃.
4. The method for forming a large-sized aircraft center wing skeleton co-bonded autoclave according to claim 1, wherein in the skeleton rib forming step, the skeleton ribs are divided into two types, one type is a single-sided box shape and the other type is a double-sided box shape;
paving a single-sided box-shaped framework rib on the single-sided rubber soft mold to paste framework rib prepreg;
the two-sided box-shaped framework rib is divided into three parts for paving, firstly, prepregs of the framework ribs on the two sides are respectively paved on corresponding rubber soft molds, and the rubber soft molds on the two sides are combined according to the positioning holes after the paving is finished; secondly, filling carbon twisted wires in triangular gaps on two sides; thirdly, paving the edge strips after filling.
5. The method for forming a large-size aircraft central wing skeleton co-cementing autoclave according to claim 1, wherein in the co-cementing curing step, the pressure of the technological parameters is 0.3-0.8MPa, the curing temperature is 125+ -6 ℃, the curing time is 120-180min, and the pressure is gradually increased from 0.1MPa to the applied pressure when the pressure is applied.
6. A method for forming a large-size aircraft center wing skeleton co-bonded autoclave according to claim 5, wherein in the co-bonding curing step, curing parameters are as follows: the room temperature is heated to 80 ℃, the heating rate is less than or equal to 2 ℃/min, the temperature is kept for 30min, the temperature difference is 80+/-3 ℃, the pressure is increased to 0.5Mpa after the temperature is kept, the temperature is heated to 100 ℃ at the speed of 1.25 ℃/min, the temperature is heated to 125 ℃ at the speed of 0.8 ℃/min, the temperature is kept for 90min, and the temperature deviation is 125+/-6 ℃.
7. The method for forming the large-size aircraft center wing skeleton co-bonded autoclave according to claim 1, wherein the C-shaped prepreg and the skeleton rib prepreg of the I-shaped skeleton beam are in a combination form of carbon fiber fabrics, carbon fiber unidirectional tapes and adhesive films.
8. The molding method of the large-size aircraft center wing skeleton co-gluing autoclave according to claim 1, wherein the rubber soft mold is made of flexible high-temperature-resistant rubber, has a box-shaped structure consistent with skeleton ribs, is made of silicon rubber, and has the lowest temperature resistance of 190 ℃ and the elongation of more than or equal to 300%.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311682328.1A CN117774376A (en) | 2023-12-08 | 2023-12-08 | Forming method of large-size aircraft central wing skeleton co-cementing autoclave |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311682328.1A CN117774376A (en) | 2023-12-08 | 2023-12-08 | Forming method of large-size aircraft central wing skeleton co-cementing autoclave |
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| Publication Number | Publication Date |
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| CN117774376A true CN117774376A (en) | 2024-03-29 |
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|---|---|---|---|
| CN202311682328.1A Pending CN117774376A (en) | 2023-12-08 | 2023-12-08 | Forming method of large-size aircraft central wing skeleton co-cementing autoclave |
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| Country | Link |
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| CN (1) | CN117774376A (en) |
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- 2023-12-08 CN CN202311682328.1A patent/CN117774376A/en active Pending
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