CN112606426A - Curing furnace forming process for full-length composite wing beam - Google Patents
Curing furnace forming process for full-length composite wing beam Download PDFInfo
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- CN112606426A CN112606426A CN202011438566.4A CN202011438566A CN112606426A CN 112606426 A CN112606426 A CN 112606426A CN 202011438566 A CN202011438566 A CN 202011438566A CN 112606426 A CN112606426 A CN 112606426A
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/28—Shaping operations therefor
- B29C70/30—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core
- B29C70/34—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression, i.e. combined with compressing after the lay-up operation
- B29C70/342—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression, i.e. combined with compressing after the lay-up operation using isostatic pressure
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/06—Fibrous reinforcements only
- B29C70/10—Fibrous reinforcements only characterised by the structure of fibrous reinforcements, e.g. hollow fibres
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/28—Shaping operations therefor
- B29C70/54—Component parts, details or accessories; Auxiliary operations, e.g. feeding or storage of prepregs or SMC after impregnation or during ageing
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29L—INDEXING SCHEME ASSOCIATED WITH SUBCLASS B29C, RELATING TO PARTICULAR ARTICLES
- B29L2031/00—Other particular articles
- B29L2031/30—Vehicles, e.g. ships or aircraft, or body parts thereof
- B29L2031/3076—Aircrafts
- B29L2031/3085—Wings
Abstract
The invention discloses a curing furnace molding process of a full-length composite material wing spar, which comprises the following steps: s1, selecting a tool, wherein the tool is provided with a paving area and an unpaved area; cutting required woven cloth and forming auxiliary materials according to the requirements of the wing beam; s2, paving and pasting a first preset angle woven cloth in a web plate structure area, paving and pasting a second preset angle woven cloth in a marginal strip area respectively, and vacuumizing and cold compacting; s3, repeating the step S2 to complete the number of the layers of the spread and pasted woven cloth to one third and two thirds of the total number of the layers of the multi-layer woven cloth of the wing beam, carrying out hot compaction on the spread and pasted woven cloth, and repeating the step S2 to complete the spreading and pasting to obtain a solid laminated piece; s4, making a bag on the tool, performing vacuum leakage inspection, vacuumizing the spar workpiece and keeping for at least 2 hours; and S5, carrying out vacuum leakage detection again, curing the spar parts qualified by detection, and stripping the auxiliary materials to obtain the full-length composite material spar.
Description
Technical Field
The invention relates to the technical field of aviation equipment, in particular to a curing oven molding process of a full-length composite material wing spar.
Background
At present, the design and forming technology of domestic aviation composite materials mainly focuses on secondary bearing components with small size, thin thickness and simple structure, and the forming process of the curing oven is also mainly applied to forming components with small stress, such as a cover, a fairing, a spoiler and the like. The wing beam is used as a main bearing component in the wing, the design and forming process requirements are very strict, and the autoclave forming process is basically used as a preferred forming scheme in China. At present, a curing furnace process is not adopted to mold a large-size composite material wing beam, so that the problems of high production and manufacturing difficulty, long period, high requirement on facility equipment, high production cost and the like of the existing composite material wing beam are solved.
Disclosure of Invention
Aiming at the problems of high production and manufacturing difficulty, long period, high requirement on facility equipment, high production cost and the like of the existing airplane composite wing beam, the invention provides a curing oven molding process of a full-length composite wing beam through the research on the design molding process of the composite wing beam, and the molding process has the advantages of short production period, low manufacturing cost and stable and reliable quality.
In order to achieve the purpose, the invention provides a curing oven molding process of a full-length composite material wing spar, which comprises the following steps:
s1, selecting a full-length composite material spar laying tool, wherein the full-length composite material spar laying tool is provided with a laying area and a non-laying area, the laying area comprises a web plate structure area arranged on the laying area and flange strip areas respectively arranged on two sides of the web plate structure area, and the non-laying area refers to an area where no part is formed outside the web plate structure area and the flange strip areas; cutting required woven cloth and forming auxiliary materials according to the design requirements of the full-length composite material wing beam;
s2, paving and pasting a layer of cut first preset angle woven cloth in the web plate structure area, paving and pasting a layer of cut second preset angle woven cloth in the edge strip area respectively to form a complete woven cloth layer, and paving a vacuumizing cold-pressing practical forming auxiliary material on the complete woven cloth layer for vacuumizing and air-cooling compaction;
s3, in the laying and pasting process, repeating the step S2 to finish the layer number of the woven cloth paved and pasted to one third of the total layer number of the multi-layer woven cloth of the full-length composite wing spar (the layer number is an integer), then carrying out hot compaction on the woven cloth paved and pasted to one third of the total layer number, then continuing repeating the step S2 to finish the layer number of the woven cloth paved and pasted to two thirds of the total layer number of the multi-layer woven cloth of the full-length composite wing spar, carrying out hot compaction on the woven cloth paved and pasted to two thirds of the total layer number, then repeating the step S2 to finish the whole paving of the multi-layer woven cloth of the full-length composite wing spar, and obtaining a solid laminated part;
s4, placing breathable materials and vacuum bag films in sequence from bottom to top in the allowance area of the step S3 of the full-length composite material spar laying tool, sealing the space between the vacuum bag films and the full-length composite material spar laying tool, arranging thermocouples on a forming tool according to requirements to manufacture a spar workpiece, then carrying out vacuum leakage inspection on the spar workpiece, vacuumizing the spar workpiece under the condition of no leakage and keeping the spar for at least 2 hours;
and S5, performing vacuum leakage detection on the spar part subjected to the step S4 again, placing the spar part qualified after the vacuum leakage detection is performed again into a curing furnace for curing under the condition of no leakage, maintaining the vacuum degree until the curing is finished and cooling the tool to room temperature for pressure relief during curing, and taking out the stripping auxiliary material from the curing furnace to obtain the full-length composite spar.
Preferably, the total number of the woven cloth paving layers of the full-length composite material wing beam is 224.
Preferably, the cross section of the full-length composite spar is "[", and the auxiliary molding materials in the step S1 include a sealing tape, a tetrafluoroethylene fabric, a perforated barrier film, a non-woven fabric, a thick airfelt, a glass fiber tow, an air-permeable fabric, a thermocouple, and a mold release agent.
Preferably, in steps S2 and S3, full length combined material wing spar is equipped with the corner, to crossing over the cloth is woven to all second preset angles of corner, from the regional direction cut of one side direction border at the regional place of web structure at the corner, every paving at the corner and pasting the one deck the cloth is woven to the second preset angle, the cloth is woven to the one deck second preset angle of paving more again, and the cut of the cloth is woven to the second preset angle that increases and the cut of the cloth is woven to the second preset angle of original paving and is staggered the setting.
Preferably, the distance between the added shearing mouth of the second woven fabric with the second preset angle and the shearing mouth of the second woven fabric with the second preset angle which is originally paved and pasted is 4-8 mm, and the optimal distance is 5 mm. The range of the second preset angle woven cloth paving is 50-80 mm at two sides of the corner. Wherein, the optimal range of the second preset angle woven cloth paving is 60mm at two sides of the corner.
Preferably, in step S2, the carbon fiber arrangement angle of the woven cloth with the first preset angle is ± 45 °, that is, the carbon fiber arrangement angle is +45 ° or-45 °, the carbon fiber arrangement angle of the woven cloth with the first preset angle is 0 °, which is also called unidirectional cloth or unidirectional tape.
Preferably, the conditions for laying each complete woven fabric layer in step S2 are as follows: the paving temperature is 18-22 ℃, and the environment humidity is less than 65%, wherein the controllable environment refers to the range of the adjustable temperature of 18-22 ℃ and the range of the adjustable humidity of less than 65% when the environment changes. The vacuum pumping and air cooling compaction conditions in the step S2 are as follows: the vacuum degree is not less than 0.08Mpa, i.e. the vacuum degree is equal to 0.08Mpa, or more than 0.08Mpa and less than 0.1Mpa, the vacuum pumping time is at least 15min, and the temperature and the environmental humidity are the same as those when the woven cloth is spread and pasted. And in the step S2, the vacuumizing cold-pressing practical forming material auxiliary material comprises tetrafluoroethylene cloth, a porous isolating membrane and a breathable fabric, wherein the tetrafluoroethylene cloth, the porous isolating membrane and the breathable fabric are sequentially arranged from bottom to top on the complete woven cloth layer.
Preferably, the hot pressing conditions for hot compaction in step S3 are: hot compacting temperature is 40 + -2 deg.C, vacuum degree is not less than 0.0745Mpa, i.e. vacuum degree is equal to 0.0745Mpa, or greater than 0.0745Mpa and less than 0.1Mpa, hot pressing time is at least 15 min; in the step S3, after 3 to 5 complete woven fabric layers are laid and attached, an air guide material needs to be placed in the allowance area of the part to be formed, and the air guide material is connected with the non-laid area of the full-length composite material spar laying tool; and then continues to repeat the step S2.
Preferably, the vacuum leakage check conditions in step S4 are: when the vacuum is pumped, the vacuum degree of the solid laminated piece is not lower than 0.08Mpa, and the reading of a vacuum meter is not reduced by more than 6.77Kpa within 5 min; otherwise, if the condition is not met, the vacuum leakage is judged. The breathable materials in the step S4 comprise thick breathable felt, glass fiber tows, a porous isolating membrane and breathable fabric. The method comprises the steps of laying glass fiber tows, namely electronic yarns, on a solid laminated piece, laying a plurality of glass fiber tows in an area with more layers of woven cloth, then arranging a porous isolating membrane on the glass fiber tows, then laying breathable fabric on the porous isolating membrane, placing a circle of 2 layers of breathable felts at the joint of the allowance area of a full-length composite material wing spar laying tool and the breathable fabric, namely arranging a layer of breathable felt on the upper side and the lower side of the edge of the breathable fabric, which is in contact with the full-length composite material wing spar laying tool, then laying a layer of vacuum bag film on the outer side of the breathable felts and the breathable fabric, and sealing the vacuum bag film on the full-length composite material wing spar laying tool by using a sealing adhesive tape to obtain the vacuum bag of the solid laminated piece.
Preferably, the vacuum leakage check conditions in step S5 are: when the vacuum is pumped, the vacuum degree of the solid laminate is not lower than 0.08MPa, and the reading descending rate of a vacuum meter should not exceed 6.77Kpa/5 min. A bagging solid laminate that does not meet this condition cannot be placed into a curing oven. The curing process in the curing oven in step S5 is as follows:
s51, keeping the vacuum degree of the curing furnace at 0.074MPa within half an hour when curing is started, raising the temperature in the curing furnace from room temperature (20 +/-5 ℃) to 80 +/-5 ℃ within 1.6h when curing is started, keeping the temperature raising rate at 0.2-0.8 ℃/min, and then keeping the temperature for 4-5 h;
s52, after the step S51 is completed, the temperature of the curing furnace is raised to 100 +/-5 ℃ within 1 hour, the temperature raising rate is 0.2-0.8 ℃/min, and the temperature is kept for at least 1 hour;
s53, after the step S52 is completed, the temperature of the curing furnace is raised to 130 +/-5 ℃ within 1.4h, the temperature raising rate is 0.2-0.8 ℃/min, and the temperature is kept for at least 2 hours;
s54, after the step S53 is completed, pressure is released to the same pressure inside and outside the curing furnace at 16.5h, the temperature of the curing furnace is reduced, the temperature reduction rate is less than or equal to 3 ℃/min, when the temperature of the curing furnace is reduced to 60 ℃ or below, the vacuum degree in the curing furnace is removed, and the solid laminated part is taken out of the curing furnace, so that the large-size full-length composite material wing beam is prepared.
Compared with the prior art, its beneficial effect lies in:
according to the curing furnace forming process of the full-length composite material wing spar, the woven cloth layers are laid on the full-length composite material wing spar in the full-length composite material wing spar laying tool, vacuumizing, air-cooling and hot compacting are performed to compact the woven cloth layers, the integrated entity laminating part is manufactured at one time, vacuumizing and curing are performed through the steps of bag making, curing of the curing furnace and the like, and the full-length composite material wing spar is manufactured, so that the using number of tool molds and the using times of facility equipment are reduced compared with the existing forming processes of glue joint co-curing, secondary glue joint and the like, the forming efficiency is higher, the production period is shortened by about 30% -40%, and the production cost is reduced by about 40% -50%; meanwhile, the integrated manufacturing process further stabilizes and improves the quality of the full-length composite material wing beam, and avoids the phenomena of insufficient bonding strength or flaws and the like possibly caused by the gluing and forming process of related parts.
Drawings
FIG. 1 is a first perspective view of a full length composite spar in accordance with an embodiment of the present invention.
FIG. 2 is a second perspective view of a full length composite spar in accordance with embodiments of the present invention.
FIG. 3 is a cross-sectional layup schematic view of the full length composite spar of FIG. 1 along section line A-A.
FIG. 4 is a schematic diagram of a bagging process for a full length composite spar.
FIG. 5 is a curing process diagram of a curing oven.
In the figure, 1, a full length composite spar; 2. a full-length composite material wing spar laying tool; 4. sealing the adhesive tape; 5. the web plate structure area is correspondingly paved with the woven cloth; 6. the edge strip area is correspondingly paved with the adhered woven cloth; 7. an air-permeable felt; 8. a release agent; 9. a solid laminate; 10. glass fiber tows; 11. a perforated barrier film; 12. a breathable fabric; 13. and (4) vacuum bag film.
Detailed Description
The following detailed description of embodiments of the present invention is provided in connection with the accompanying drawings and examples. The following examples are intended to illustrate the invention but are not intended to limit the scope of the invention.
The embodiment of the invention provides a curing furnace molding process of a full-length composite material wing beam as a preferred technical scheme, which comprises the following steps:
s1, selecting a full-length composite material spar laying tool, wherein the full-length composite material spar laying tool is provided with a laying area and a non-laying area, the laying area comprises a web plate structure area arranged on the laying area and flange strip areas respectively arranged on two sides of the web plate structure area, the non-laying area refers to an area without parts outside the web plate structure area and the flange strip areas, and the parts refer to partial areas forming the full-length composite material spar; cutting required woven cloth and forming auxiliary materials according to the design requirements of the full-length composite material wing beam;
s2, firstly, paving a layer of cut first preset angle woven cloth on the web structure area of the step S1, namely weaving cloth with a carbon fiber arrangement angle of +45 degrees, then respectively paving a layer of second preset angle weaving cloth which is cut in the edge strip area, namely the woven cloth with the carbon fiber arrangement angle of 0 degree, the second woven cloth with the preset angle is lapped and stuck on the first woven cloth with the preset angle to form a complete woven cloth layer, then sequentially laying tetrafluoroethylene cloth and a porous isolating membrane (namely vacuumizing cold pressing practical forming auxiliary materials) on the complete weaved cloth layer from bottom to top, sealing a vacuum bag on the periphery of the full-length composite material wing spar tool through a sealing adhesive tape to form a closed space, and then pumping out air in the vacuum bag to form negative pressure, and vacuumizing and cold compacting the lapped woven fabric layer. The conditions when each complete woven fabric layer is laid and attached are as follows: the controllable environment with the temperature of 20 ℃ and the environmental humidity of 60 percent is paved. The laid tetrafluoroethylene cloth is more beneficial to discharging gas mixed in the prepreg in real time during vacuumizing and cold pressing, and the porous isolating film is laid between the breathable fabric and the tetrafluoroethylene cloth, so that the breathable fabric has the effects of breathability and prevention of the breathable fabric from being stuck on the woven cloth; the vacuumizing and air-cooling compaction conditions are as follows: the vacuum degree is 0.09Mpa, the vacuumizing time is 25min, and the temperature for vacuumizing, air-cooling and compacting is the same as the temperature for paving the woven cloth so as to discharge gas molecules between the woven cloth layers.
And S3, repeating the step S2 to spread and paste the woven cloth with 224 layers in total in the spreading process. Wherein, every shop pastes 4 complete woven fabrics layers of completion after, need place the air guide material in the surplus district (the region of not spreading the woven fabrics of pasting in the region of mating formation promptly) of treating the shaping part, the air guide material with the region of mating formation of full length combined material spar shop layer frock links to each other not mating formation, the air guide material is here the non-woven fabrics, and the air guide material of course also can be glass fibre silk bundle, glass fiber yarn to discharge the gas between the woven fabrics layer in the in-process of evacuation cold compaction or evacuation hot compaction, improve the quality of full length combined material spar.
And then, continuously repeating the step S2 to complete the process from the number of the laid woven cloth layers to one third of the total number of the multi-layer woven cloth layers of the full-length composite material wing beam (the number of the layers is an integer), and then performing hot compaction on the laid woven cloth layers of one third of the total number of the laid woven cloth layers so as to ensure that the laid woven cloth layers are completely attached to the mold surface of the full-length composite material wing beam laying tool as far as possible, so that gas molecules among the woven cloth layers are removed as far as possible, and the quality is improved. The hot pressing conditions of the hot compaction are as follows: hot compacting at 40 deg.C under 0.0745MPa for 15 min; and then continuously repeating the step S2 to complete the spreading of the number of layers of the woven cloth to two thirds of the total number of layers of the multi-layer woven cloth of the full-length composite material wing beam, and carrying out hot compaction on the two thirds of the total number of layers of the woven cloth which is completely spread, wherein the hot compaction conditions are as follows: hot compacting at 40 deg.C under 0.0745MPa for 15 min; then repeating the step S2 to finish the paving and pasting of the multi-layer woven cloth of the full-length composite material wing beam, so as to obtain a solid laminated piece;
as an optimized technical solution of the above scheme, in the steps S2 and S3, the full-length composite spar is provided with corners (including the corners of the spar in the length direction and the corners of the edge strip regions), to weave the cloth at all second preset angles across the corners, cuts from one side of the web structure region at the corners to the edge strip region direction, lays each layer of the cloth at the corners after the cloth is woven at the second preset angles, lays and pastes a layer of the cloth at the second preset angles more, and the cut of the cloth woven at the second preset angles that is increased is staggered from the cut of the cloth woven at the second preset angles that was originally laid. The distance between the added shearing mouth of the second preset angle woven cloth and the shearing mouth of the second preset angle woven cloth which is originally paved and pasted is 6mm, and the paving and pasting of the corner woven cloth are facilitated after the shearing mouths. The optimal range that the cloth was woven to the angle of second preset angle that increases was laid and is pasted is each 60mm in both sides of corner, because the cloth material performance of weaving of corner is impaired behind the cutting, staggers the cutting of cloth is woven to the previous layer second preset angle through the cloth of weaving that increases to guarantee the performance of final full length combined material wing spar. If the paving range is too small, the performance strength and the like of the composite material can not meet the design requirements, the paving range is too large, the total weight is increased, and the possibility of interference exists in subsequent assembly.
S4, sequentially placing a breathable material and a vacuum bag film in a surplus area (namely, the surplus area of the woven cloth material laid by the full-length composite material spar laying tool) of the step S3 from bottom to top in the full-length composite material spar laying tool, wherein the breathable material in the step S4 comprises a thick breathable felt, glass fiber tows, a porous isolating film and a breathable fabric, the glass fiber tows are also called electronic yarns and are laid on the solid laminated part, the glass fiber tows are convenient for air mixed between the woven cloth layers and gas molecules generated by chemical reaction of resin during curing to be discharged, the defects of porosity, looseness and the like of the spar are reduced, and the quality of the spar is improved. The method comprises the steps of laying a plurality of glass fiber tows in an area where woven cloth is laid and adhered in a large number of layers, arranging a porous isolating film on the glass fiber tows, laying breathable fabric on the porous isolating film, placing a circle of 2 layers of breathable felts at the joint of a margin area of a full-length composite material wing spar laying tool and the breathable fabric, namely arranging a circle of breathable felts on the upper side and the lower side of the edge of the breathable fabric, which is in contact with the full-length composite material wing spar laying tool, respectively, laying a layer of vacuum bag film on the outer side of the breathable felts and the breathable fabric, sealing the vacuum bag film on the full-length composite material wing spar laying tool by using a sealing adhesive tape, manufacturing a vacuum bag containing a solid laminated part, namely a bag, and arranging thermocouples on a forming tool according to requirements to manufacture a wing spar part, wherein the wing spar part is manufactured as. Then, carrying out vacuum leakage inspection on the wing beam part, vacuumizing the wing beam part under the condition of no leakage and keeping for 3 hours so as to ensure that no air leakage phenomenon occurs in the bag making process and the vacuum bag is completely attached to the laid prepreg, and exhausting the gas mixed between the woven cloth layers as far as possible; the conditions for the vacuum leak check were: when the vacuum is pumped, the vacuum degree of the solid laminated piece is 0.08Mpa, and the reading of a vacuum meter does not drop over 6.77Kpa within 5 min; otherwise, spar parts that do not meet the vacuum leak conditions described above should be judged as vacuum leaks.
And S5, carrying out vacuum leakage detection on the spar member after the step S4 is finished again, wherein the vacuum leakage condition is the same as that in the step S4, placing the spar member qualified in the vacuum leakage detection into a curing furnace for curing under the condition of no leakage, maintaining the vacuum degree until the curing is finished, and releasing the pressure when the tool is cooled to room temperature, and taking out the stripping auxiliary material from the curing furnace to obtain the full-length composite material spar.
Specifically, the curing process in the curing furnace in step S5 is as shown in fig. 5:
s51, maintaining the vacuum degree of the curing furnace at 0.074MPa within half an hour when curing is started, increasing the temperature in the curing furnace from room temperature (25 ℃) to 85 ℃ within 1.6h when curing is started, maintaining the temperature increase rate at 0.625 ℃/min, and then keeping the temperature for 5 h;
s52, after the step S51 is completed, the temperature of the curing furnace is increased to 104 ℃ within 1 hour, the temperature increase rate is 0.32 ℃/min, and the temperature is kept for 2 hours;
s53, after the step S52 is completed, the temperature of the curing furnace is increased to 130 ℃ within 1.4h, the temperature increase rate is 0.31 ℃/min, and the heat is preserved for 3 hours;
s54, after the step S53 is completed, pressure is released to the same pressure inside and outside the curing oven within 16.5h, the temperature of the curing oven is reduced, the cooling rate is 2.8 ℃/min, when the temperature of the curing oven is reduced to 40 ℃, the vacuum degree in the curing oven is removed, the solid laminated piece is taken out from the curing oven and transferred to a stripping area, and then a vacuum bag, a breathable material and a sealing adhesive tape are peeled off, so that the large-size full-length composite material wing beam is prepared. In the curing process of the steps S51 to S54 with different temperatures and curing time, the matrix material of the carbon fiber composite material in the form of woven cloth is changed from liquid to solid material after curing, and the woven cloth laying of the wing beam ensures the temperature of each part to be consistent as far as possible through a plurality of temperature stages and corresponding time, so that the temperature step difference is reduced, the residual stress and deformation caused by thermal effect and chemical effect are avoided as far as possible, and the final performance quality of the wing beam is improved.
In summary, according to the curing furnace forming process for the full-length composite material wing spar provided by the embodiment of the invention, the full-length composite material wing spar is paved with the woven cloth layers in the full-length composite material wing spar paving tool, the vacuum pumping air cooling compaction and the hot compaction are carried out to compact all the woven cloth paving layers, gas molecules between the paving layers are removed, the quality of the wing spar is improved, an integrated entity laminating part is manufactured at one time, and the vacuum pumping and the curing are carried out through the steps of bag making, curing in the curing furnace and the like, so that the full-length composite material wing spar is manufactured, compared with the existing glue joint co-curing, secondary glue joint and other forming processes, the using number of tool molds and the using times of facility equipment are reduced, the forming efficiency is higher, the production period is shortened by about 30% -40%, and the; and the cloth is conveniently woven through the shearing openings at the corners of the wing beams, and the second preset angle of the cloth is increased to ensure the quality of the wing beams. The quality of the spar is improved by removing the gas molecules of the woven cloth layer as much as possible through the bag-making process, and then the residual stress and deformation caused by thermal and chemical effects are reduced or even avoided through the solidification in multiple temperature stages, so that the final performance quality of the spar is improved. Meanwhile, the integrated manufacturing process further stabilizes and improves the quality of the full-length composite material wing beam, and avoids the phenomena of insufficient bonding strength or flaws and the like possibly caused by the gluing and forming process of related parts.
The above description is only a preferred embodiment of the present invention, and it should be noted that, for those skilled in the art, various modifications and substitutions can be made without departing from the technical principle of the present invention, and these modifications and substitutions should also be regarded as the protection scope of the present invention.
Claims (10)
1. The curing furnace molding process of the full-length composite material wing beam is characterized by comprising the following steps of:
s1, selecting a full-length composite material spar laying tool, wherein the full-length composite material spar laying tool is provided with a laying area and a non-laying area, the laying area comprises a web plate structure area arranged on the laying area and flange strip areas respectively arranged on two sides of the web plate structure area, and the non-laying area refers to an area where no part is formed outside the web plate structure area and the flange strip areas; cutting required woven cloth and forming auxiliary materials according to the design requirements of the full-length composite material wing beam;
s2, paving and pasting a layer of cut first preset angle woven cloth in the web plate structure area, paving and pasting a layer of cut second preset angle woven cloth in the edge strip area respectively to form a complete woven cloth layer, and paving a vacuumizing cold-pressing practical forming auxiliary material on the complete woven cloth layer for vacuumizing and air-cooling compaction;
s3, in the laying and pasting process, repeating the step S2 to finish the layer number of the woven cloth paved and pasted to one third of the total layer number of the multi-layer woven cloth of the full-length composite wing spar (the layer number is an integer), then carrying out hot compaction on the woven cloth paved and pasted to one third of the total layer number, then continuing repeating the step S2 to finish the layer number of the woven cloth paved and pasted to two thirds of the total layer number of the multi-layer woven cloth of the full-length composite wing spar, carrying out hot compaction on the woven cloth paved and pasted to two thirds of the total layer number, then repeating the step S2 to finish the whole paving of the multi-layer woven cloth of the full-length composite wing spar, and obtaining a solid laminated part;
s4, placing breathable materials and vacuum bag films in sequence from bottom to top in the allowance area of the step S3 of the full-length composite material spar laying tool, sealing the space between the vacuum bag films and the full-length composite material spar laying tool, arranging thermocouples on a forming tool according to requirements to manufacture a spar workpiece, then carrying out vacuum leakage inspection on the spar workpiece, vacuumizing the spar workpiece under the condition of no leakage and keeping the spar for at least 2 hours;
and S5, performing vacuum leakage detection on the spar part subjected to the step S4 again, placing the spar part qualified after the vacuum leakage detection is performed again into a curing furnace for curing under the condition of no leakage, maintaining the vacuum degree until the curing is finished and cooling the tool to room temperature for pressure relief during curing, and taking out the stripping auxiliary material from the curing furnace to obtain the full-length composite spar.
2. The curing oven molding process of a full length composite spar according to claim 1, wherein: the total number of the woven cloth paving layers of the full-length composite material wing beam is 224.
3. The curing oven molding process of a full length composite spar according to claim 1, wherein: the molding auxiliary material in the step S1 includes a sealing tape, a tetrafluoroethylene fabric, a porous isolation film, a non-woven fabric, a thick air-permeable felt, a glass fiber tow, an air-permeable fabric, a thermocouple, and a release agent.
4. The curing oven molding process of a full length composite spar according to claim 1, wherein: in steps S2 and S3, full length combined material wing spar is equipped with the corner, to strideing across all second preset angle woven cloth of corner, from the regional direction cut of one side direction border at the regional place of web structure at the corner, every paving at the corner and pasting the one deck the second preset angle woven cloth back, the one deck second preset angle woven cloth of paving more again, the cut of the second preset angle woven cloth that increases and the original second preset angle woven cloth of paving staggers the setting.
5. The curing oven molding process of a full length composite spar according to claim 4, wherein: the distance between the added shearing mouth of the second preset angle woven cloth and the shearing mouth of the second preset angle woven cloth paved and pasted originally is 4-8 mm, and the range of the added second preset angle woven cloth paved and pasted is 50-80 mm on two sides of the corner respectively.
6. The curing oven molding process of a full length composite spar according to claim 1, wherein: in the step S2, the carbon fiber arrangement angle of the woven cloth at the first preset angle is ± 45 °, and the carbon fiber arrangement angle of the woven cloth at the first preset angle is 0 °.
7. The curing oven molding process of a full length composite spar according to claim 1, wherein: the conditions for paving each complete woven fabric layer in the step S2 are as follows: paving and pasting the controllable environment with the temperature of 18-22 ℃ and the environmental humidity of less than 65%; the vacuum pumping and air cooling compaction conditions in the step S2 are as follows: vacuum degree not less than 0.08Mpa, and vacuumizing time at least 15 min; and in the step S2, the vacuumizing cold-pressing practical forming material auxiliary material comprises tetrafluoroethylene cloth, a porous isolating membrane and a breathable fabric, wherein the tetrafluoroethylene cloth, the porous isolating membrane and the breathable fabric are sequentially arranged from bottom to top on the complete woven cloth layer.
8. The curing oven molding process of a full length composite spar according to claim 1, wherein: the hot pressing conditions for the hot compaction in the step S3 are: hot compacting temperature is 40 + -2 deg.C, vacuum degree is not less than 0.0745Mpa, and vacuumizing and hot pressing time is at least 15 min; in the step S3, after 3 to 5 complete woven fabric layers are laid and attached, an air guide material needs to be placed in the margin area of the part to be molded, the air guide material is connected with the non-laid area of the full-length composite material spar laying tool, and then the step S2 is continuously repeated.
9. The curing oven molding process of a full length composite spar according to claim 1, wherein: the conditions for the vacuum leak check in step S4 are: when the vacuum is pumped, the vacuum degree of the solid laminated piece is not lower than 0.08Mpa, and the reading of a vacuum meter is not reduced by more than 6.77Kpa within 5 min; the breathable material in the step S4 includes a thick breathable felt, glass fiber tows, a porous isolation film, and a breathable fabric, wherein the glass fiber tows are laid on the solid laminate, a plurality of glass fiber tows are laid on the solid laminate in an area where the woven fabric is laid with a large number of layers, the porous isolation film is arranged on the glass fiber tows, a breathable fabric is laid on the porous isolation film, a circle of 2 layers of breathable felts are placed at a joint of a margin area of the full-length composite material spar laying tool and the breathable fabric, that is, the breathable fabric is provided with a layer of breathable felt on the upper side and the lower side of an edge in contact with the full-length composite material spar laying tool, a layer of vacuum bag film is laid on the outer side of the breathable felts and the breathable fabric, and the vacuum bag film is sealed on the full-length composite material spar laying tool by a sealing tape, a vacuum bag of the solid laminate is made.
10. The curing oven molding process of a full length composite spar according to claim 1, wherein the vacuum leak check conditions in step S5 are: when the vacuum is pumped, the vacuum degree of the solid laminate is not lower than 0.08Mpa, and the reading descending rate of a vacuum meter is not more than 6.77Kpa/5 min; the curing process in the curing oven in step S5 is as follows:
s51, the vacuum degree of the curing furnace needs to reach and be kept at 0.074MPa within half an hour when the curing is started, the temperature in the curing furnace is increased from room temperature (20 +/-5 ℃) to 80 +/-5 ℃ within 1.6h when the curing is started, the temperature rising rate is kept at 0.2-0.8 ℃/min, and then the temperature is kept for 4-5 h;
s52, after the step S51 is completed, the temperature of the curing furnace is raised to 100 +/-5 ℃ within 1 hour, the temperature raising rate is 0.2-0.8 ℃/min, and the temperature is kept for at least 1 hour;
s53, after the step S52 is completed, the temperature of the curing furnace is raised to 130 +/-5 ℃ within 1.4h, the temperature raising rate is 0.2-0.8 ℃/min, and the temperature is kept for at least 2 hours;
s54, after the step S53 is completed, pressure is released to the same pressure inside and outside the curing furnace at 16.5h, the temperature of the curing furnace is reduced, the temperature reduction rate is less than or equal to 3 ℃/min, when the temperature of the curing furnace is reduced to 60 ℃ or below, the vacuum degree in the curing furnace is removed, and the solid laminated part is taken out of the curing furnace, so that the large-size full-length composite material wing beam is prepared.
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| CN110978562A (en) * | 2019-11-28 | 2020-04-10 | 长春长光宇航复合材料有限公司 | Preparation method of high-temperature-resistant composite material cabin section |
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