CN102092135A - Method for improving rigidity of wing surface structure of composite material - Google Patents
Method for improving rigidity of wing surface structure of composite material Download PDFInfo
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- CN102092135A CN102092135A CN2010105832171A CN201010583217A CN102092135A CN 102092135 A CN102092135 A CN 102092135A CN 2010105832171 A CN2010105832171 A CN 2010105832171A CN 201010583217 A CN201010583217 A CN 201010583217A CN 102092135 A CN102092135 A CN 102092135A
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Abstract
The invention relates to a method for improving the rigidity of a wing surface structure of a common modulus fibre reinforced resin based composite material by using high modulus fibres, belonging to the technical field of fibre reinforced resin based composite materials. The method for improving the rigidity of the wing surface structure of the composite material is realized by placing the high modulus fibres into the originally designed spreading layer of the wing surface structure of the composite material through continuously manually or mechanically spreading according to a certain distance. Without the limit of the original resin system and reinforcement fibres of the wing surface structure of the composite material, the improvement on the rigidity of the wing surface structure of the composite material can be realized with lower cost on the basis of unchanging the original process and process parameters of the composite material structure. In the method, a thin-walled wing surface structure of a thermoset composite material formed by adopting an autoclave process or resin transfer molding process or resin membrane impregnating process or mould pressing process can be adopted, and a wing surface structure of a thermoplastic composite material formed by adopting a mould pressing process or thermal deformation process or membrane forming process can also be adopted.
Description
Technical field
The invention belongs to the fiber-reinforced resin matrix compound material applied technical field, relate to a kind of method that improves composite airfoil structure rigidity.
Background technology
The composite airfoil structure is as one of integrally-built important component part of aerospace flight vehicle, and its effect mainly provides or produce lift (as the wing aerofoil of fixed wing aircraft, the rotor aerofoil of helicopter etc.), flight attitude is adjusted power (as vertical tail aerofoil, the tailplane aerofoil of fixed wing aircraft) or equilibrant force (as the tailplane aerofoil of fixed wing aircraft, the missile wing aerofoil of guided missile etc.).Therefore, airfoil structure generally is designed to satisfy the face or the chip architecture of aerodynamic configuration requirement.In order to ensure the security of aerospace flight vehicle and the accuracy of executing the task, when the design airfoil structure, need satisfy the strength and stiffness index simultaneously.
Composite provides more wide design space for Flight Vehicle Design person.Composite has series of advantages such as specific strength height, specific modulus height, designability are strong, excelling in fatigue property, and therefore, composite is subjected to the favor of aerospace field day by day, becomes one of aircraft main material of new generation.Under identical weight, the polymer matrix composites airfoil structure has than the better pneumatic bearing capacity of metal material airfoil structure, can effectively increase flying height and flight time, is a main direction of airfoil structure development.
But, when the design airfoil structure, need the higher rigidity of structure, but existing polymer matrix composites material and rigidity thereof can not satisfy the demand of design along with the further lifting of performances such as aerospace flight vehicle mobility, voyage.Its reason mainly is under the requirement of composite structure economic index, and the airfoil structure polymer matrix composites generally generally adopt common modulus, and (fiber of 200GPa～280GPa) is as strengthening body.(280GPa～350GPa) when strengthening body, composite structure will face the too high and awkward condition of " can not using " of cost and when total adopts high modulus fibre.In order to alleviate this contradiction, the designer tends to adopt and increases the design safety that method such as structural thickness comes the compensate for stiffness deficiency to bring, but this way can directly cause the weight of aerospace flight vehicle to increase, and then can reduce its mobility, even can increase its life cycle cost (particularly for civilian commercial aircraft).
Summary of the invention
The objective of the invention is: propose a kind of shortcoming that can overcome existing standard modulus composite airfoil structure rigidity of structure deficiency, can overcome again and adopt medium and above modulus carbon fiber to prepare the too high shortcoming of composite airfoil structure cost, and make composite when the preparation airfoil structure, can take into account high rigidity and the low-cost a kind of method that improves composite airfoil structure rigidity that requires simultaneously.
Technical scheme of the present invention is:
(1) selects fiber according to fiber modulus: based on the fortifying fibre modulus＜280GPa of composite airfoil structure, selection is than the high at least one grade fiber of composite airfoil structure fortifying fibre modulus, is that carbon fiber or the modulus of 280GPa～650GPa is that 350GPa～400GPa boron fibre or modulus are 350GPa～450GPa silicon carbide fibre comprising modulus;
(2) spacing of selected fiber is set: when adopting three kinds of different fibers, interfibrous spacing is, the boron fibre spacing is 10 pieces/centimetre～100 pieces/centimetre, and the silicon carbide fibre spacing is 10 pieces/centimetre～100 pieces/centimetre, the spacing of carbon fiber be 1 the bundle/centimetre～10 the bundle/centimetre;
(3) lay of selected fiber: with the fiber selected in the step (1) set by step the spacing in (2) by hand or the direct lay of method placed continuously of machinery to the fiber lay down laminar surface of airfoil structure, and the resin solution that is complementary with resin matrix at its surperficial spraying or brushing and airfoil structure, quantity for spray is 2g/m
2～10g/m
2, the position for the treatment of permanent domicile selection fiber after solvent volatilizees fully in the resin solution; The fiber lay down layer of preparing the airfoil structure that is stained with selected fiber is designated as high-modulus shop layer, and substituting former airfoil structure shop number of plies amount with high-modulus shop layer segment is the fiber lay down layer of n＞3, and it substitutes 1/n～2/3 of amount for fibrillation shop number of plies amount;
(4) composite airfoil structure moulding is spread layer with the high-modulus of preparation in the step (3) and airfoil structure remains the initial shop layer design laying that fibrillation shop layer is pressed airfoil structure; When fibrillation shop layer is prepreg or preimpregnation cloth,, and when fibrillation shop layer is fabric, adopt infiltration moulding of resin transfer molding or resin molding or vacuum bag moulding with autoclave molding or compression molding or moulding of thermal change type or diaphragm forming; The heating rate of above-mentioned moulding process is＜2 ℃/min.
Described airfoil structure resin matrix is thermosetting resin or thermoplastic resin or thermosetting resin and the mixing of thermoplastic resin, thermosetting resin: epoxy resin, bimaleimide resin, thermoset polyimide resin, benzoxazine colophony, unsaturated polyester resin, phenolic resins or cyanate ester resin, thermoplastic resin: TPI, PEI, polyphenylene sulfide, polyether-ether-ketone, polyether-ketone, polyether sulfone, polyethylene or High molecular weight polyethylene, aforementioned hot thermosetting resin and thermoplastic resin ratio are 1: 20～1: 2 mixture.
The fiber lay down layer of airfoil structure recited above is 1) fabric, comprise unidirectional or plain weave or satin weave or twills, 2) unidirectional pre-immersion material or above-mentioned 1) described in the preimpregnation cloth of fabric.
Above the quality percentage composition of the resin solution used in (3) be 1%～20%.
When prepared layer Central Plains, high-modulus shop fiber lay down layer is prepreg or preimpregnation cloth, do not spray or brush resin solution.
Ply sequence behind the layer of the alternative fibrillation of layer shop, described high-modulus shop is one of following: 1) when airfoil structure has only single face as working face, high-modulus shop layer is laid in the position, top layer of composite airfoil structure continuously, and the high modulus fibre warp-wise in the layer of high-modulus shop is along the aerofoil length direction; 2) two-sided during as working face when airfoil structure, high-modulus shop layer is laid in the middle part of composite airfoil structure continuously, and the high modulus fibre warp-wise in the layer of high-modulus shop is along the aerofoil length direction.When fibrillation shop layer is prepreg or preimpregnation cloth, the forming temperature of described autoclave molding method is 80 ℃～300 ℃, pressure is 0.20MPa～0.60MPa, described mould pressing process forming temperature is 80 ℃～300 ℃, briquetting pressure is 0.20MPa～1.0MPa, described thermal change type technological forming type is 80 ℃～300 ℃, briquetting pressure is 0.20MPa～0.40MPa, described diaphragm forming technological forming temperature is 80 ℃～300 ℃, briquetting pressure is 0.20MPa～0.70MPa, and when fibrillation shop layer is fabric, the injection temperature of described resin transfer molding technology is 30 ℃～300 ℃, injection pressure is 0.05MPa～0.70MPa, and the forming temperature of described resin molding Infiltration Technics is 80 ℃～300 ℃, 80 ℃～300 ℃ of the moulding of described vacuum bag moulding process.
Advantage of the present invention is: the present invention has following advantage:
The first, use this method both can improve composite airfoil structure rigidity, do not make the cost of airfoil structure significantly improve again, the through engineering approaches cost is low.
The second, this method does not change original parameter of original airfoil structure composite material process planning, need not to carry out the adjustment of technological parameter, and the technology versatility is good.
The 3rd, this method not only is suitable for the same with thermosetting compound material airfoil structure, and is applicable to the thermoplastic composite airfoil structure, is not subjected to the material selectional restriction of airfoil structure, and material adaptability is strong.
The specific embodiment
Below the present invention is described in further details.The present invention proposes a kind of method that improves composite airfoil structure rigidity, it is characterized in that, the operating procedure of this method is as follows:
(1) selects fiber according to fiber modulus: based on the fortifying fibre modulus＜280GPa of composite airfoil structure, selection is than the high at least one grade fiber of composite airfoil structure fortifying fibre modulus, is that carbon fiber or the modulus of 280GPa～650GPa is that 350GPa～400GPa boron fibre or modulus are 350GPa～450GPa silicon carbide fibre comprising modulus;
(2) spacing of selected fiber is set: when adopting three kinds of different fibers, interfibrous spacing is, the boron fibre spacing is 10 pieces/centimetre～100 pieces/centimetre, and the silicon carbide fibre spacing is 10 pieces/centimetre～100 pieces/centimetre, the spacing of carbon fiber be 1 the bundle/centimetre～10 the bundle/centimetre;
(3) lay of selected fiber: with the fiber selected in the step (1) set by step the spacing in (2) by hand or the direct lay of method placed continuously of machinery to the fiber lay down laminar surface of airfoil structure, and the resin solution that is complementary with resin matrix at its surperficial spraying or brushing and airfoil structure, quantity for spray is 2g/m
2～10g/m
2, the position for the treatment of permanent domicile selection fiber after solvent volatilizees fully in the resin solution; The fiber lay down layer of preparing the airfoil structure that is stained with selected fiber is designated as high-modulus shop layer, and substituting former airfoil structure shop number of plies amount with high-modulus shop layer segment is the fiber lay down layer of n＞3, and it substitutes 1/n～2/3 of amount for fibrillation shop number of plies amount;
(4) composite airfoil structure moulding is spread layer with the high-modulus of preparation in the step (3) and airfoil structure remains the initial shop layer design laying that fibrillation shop layer is pressed airfoil structure; When fibrillation shop layer is prepreg or preimpregnation cloth,, and when fibrillation shop layer is fabric, adopt infiltration moulding of resin transfer molding or resin molding or vacuum bag moulding with autoclave molding or compression molding or moulding of thermal change type or diaphragm forming; The heating rate of above-mentioned moulding process is＜2 ℃/mi n.
Described airfoil structure resin matrix is thermosetting resin or thermoplastic resin or thermosetting resin and the mixing of thermoplastic resin, thermosetting resin: epoxy resin, bimaleimide resin, thermoset polyimide resin, benzoxazine colophony, unsaturated polyester resin, phenolic resins or cyanate ester resin, thermoplastic resin: TPI, PEI, polyphenylene sulfide, polyether-ether-ketone, polyether-ketone, polyether sulfone, polyethylene or High molecular weight polyethylene, aforementioned hot thermosetting resin and thermoplastic resin ratio are 1: 20～1: 2 mixture.
The fiber lay down layer of airfoil structure recited above is 1) fabric, comprise unidirectional or plain weave or satin weave or twills, 2) unidirectional pre-immersion material or above-mentioned 1) described in the preimpregnation cloth of fabric.
Above the quality percentage composition of the resin solution used in (3) be 1%～20%.
When prepared layer Central Plains, high-modulus shop fiber lay down layer is prepreg or preimpregnation cloth, do not spray or brush resin solution.
Ply sequence behind the layer of the alternative fibrillation of layer shop, described high-modulus shop is one of following: 1) when airfoil structure has only single face as working face, high-modulus shop layer is laid in the position, top layer of composite airfoil structure continuously, and the high modulus fibre warp-wise in the layer of high-modulus shop is along the aerofoil length direction; 2) two-sided during as working face when airfoil structure, high-modulus shop layer is laid in the middle part of composite airfoil structure continuously, and the high modulus fibre warp-wise in the layer of high-modulus shop is along the aerofoil length direction.
When fibrillation shop layer is prepreg or preimpregnation cloth, the forming temperature of described autoclave molding method is 80 ℃～300 ℃, pressure is 0.20MPa～0.60MPa, described mould pressing process forming temperature is 80 ℃～300 ℃, briquetting pressure is 0.20MPa～1.0MPa, described thermal change type technological forming type is 80 ℃～300 ℃, briquetting pressure is 0.20MPa～0.40MPa, described diaphragm forming technological forming temperature is 80 ℃～300 ℃, briquetting pressure is 0.20MPa～0.70MPa, and when fibrillation shop layer is fabric, the injection temperature of described resin transfer molding technology is 30 ℃～300 ℃, injection pressure is 0.05MPa～0.70MPa, and the forming temperature of described resin molding Infiltration Technics is 80 ℃～300 ℃, 80 ℃～300 ℃ of the moulding of described vacuum bag moulding process.
Embodiment 1: fixed wing aircraft wing aerofoil increases just.
Fixed wing aircraft wing airfoil structure was used the preparation of T300 (modulus is 230GPa) carbon fiber epoxy prepreg fiber lay down layer originally, selecting boron fibre (modulus is 394GPa) is high modulus fibre, adopt manual method of placing to be placed into T300 carbon fiber epoxy prepreg surface boron fibre by 10 pieces/centimetre spacing, obtain being stained with the T300 carbon fiber epoxy prepreg of boron fibre (modulus is 394GPa), be designated as high-modulus shop layer; The two-layer back of outermost with high-modulus shop layer replacement fixed wing aircraft wing aerofoil fibrillation shop layer forms new airfoil structure fiber lay down layer then, adopt the original autoclave forming process moulding of fixed wing aircraft wing, be specially and under pressure 0.50MPa～0.60MPa, be warming up to 180 ℃, and after keeping 2h, heating rate is＜2 ℃/mi n, can obtain increasing firm fixed wing aircraft wing aerofoil after reducing to room temperature then.Fixed wing aircraft wing aerofoil ratio of rigidity after increasing had just improved 10% originally.
Embodiment 2: fixed wing aircraft wing aerofoil increases just.
This embodiment and embodiment 1 different are to select silicon carbide fibre, and (modulus is 434GPa～448GPa) be high modulus fibre, and other is identical with embodiment 1.Fixed wing aircraft wing aerofoil ratio of rigidity after increasing had just improved 12% originally.
Embodiment 3: fixed wing aircraft wing aerofoil increases just
This embodiment is that with the different of embodiment 1 forming method of fixed wing aircraft wing airfoil structure is a mould pressing process, be specially and under pressure 0.50MPa～0.70MPa, be warming up to 180 ℃, and after keeping 2h, heating rate is＜2 ℃/mi n, can obtain increasing firm fixed wing aircraft wing aerofoil after reducing to room temperature then, the fixed wing aircraft wing aerofoil ratio of rigidity that other is identical with embodiment 1 after increasing had just improved 10% originally.
Embodiment 4: fixed wing aircraft empennage aerofoil increases just
It is high modulus fibre that this embodiment and embodiment 1 different are to select T1000 (modulus is 294GPa) carbon fiber, with the T1000 carbon fiber adopt the pre-dipping machine of hot melt by 1 bundle/centimetre spacing be placed into T300 carbon fiber epoxy prepreg surface automatically, obtain being stained with the T300 carbon fiber epoxy prepreg of boron fibre (modulus is 394GPa), be designated as high-modulus shop layer; Form new airfoil structure fiber lay down layer with behind four layers of the outermosts of high-modulus shop layer replacement fixed wing aircraft empennage aerofoil fibrillation shop layer then, adopt the original autoclave forming process moulding of fixed wing aircraft empennage aerofoil, other is identical with embodiment 1.Obtain increasing firm fixed wing aircraft empennage aerofoil ratio of rigidity and improved 8% originally.
Embodiment 5: fixed wing aircraft empennage aerofoil increases just
This embodiment 5 and embodiment 4 different are to select silicon carbide fibre, and (modulus is 434GPa～448GPa) be high modulus fibre, and the spacing of fiber is 8 pieces/centimetre, and other is identical with embodiment 4.Fixed wing aircraft empennage aerofoil ratio of rigidity after increasing had just improved 14% originally.
Embodiment 6: guided missile missile wing aerofoil increases just
It is high modulus fibre that this embodiment and embodiment 1 different are to select boron fibre (modulus is 394GPa), adopt manual method of placing to be placed into the original T300 carbon fibre fabric of guided missile missile wing aerofoil shop laminar surface boron fibre by 20 pieces/centimetre spacing, obtain being stained with the T 300 carbon fibers shop layer of boron fibre, be designated as high-modulus shop layer; Form new airfoil structure fiber lay down layer after replacing four layers of the outermosts of guided missile missile wing aerofoil fibrillation shop layer with high-modulus shop layer then, adopt the moulding of the original resin transfer moulding technology of guided missile missile wing, be specially under 105 ℃～110 ℃, progressively under the vacuum aided, progressively pressure is risen to 0.40MPa by 0.05MPa, after to be injected the finishing mold temperature is warming up to 200 ℃, and after keeping 2h, heating rate is＜2 ℃/min, reduce to room temperature then after, can obtain increasing firm guided missile missile wing.Other is identical with embodiment 1.Guided missile missile wing aerofoil ratio of rigidity after increasing had just improved 10% originally.
Embodiment 7: guided missile missile wing aerofoil increases just
The T800 carbon fiber shop layer that present embodiment and embodiment 6 different are to adopt 3% bimaleimide resin acetone solvent to be sprayed on and are stained with boron fibre, treat that solvent volatilizees fully after, be designated as high-modulus shop layer, other is identical with embodiment 6.Guided missile missile wing aerofoil ratio of rigidity after increasing had just improved 10% originally.
Embodiment 8: guided missile missile wing aerofoil increases just
Present embodiment adopts the moulding of resin molding Infiltration Technics with the different guided missile missile wing aerofoils that are of embodiment 6, be specially under 120 ℃～130 ℃, progressively under the vacuum aided, progressively pressure is risen to 0.40MPa by 0.05MPa, after to be injected the finishing mold temperature is warming up to 200 ℃, and after keeping 2h, heating rate is＜2 ℃/min, can obtain increasing firm guided missile missile wing after reducing to room temperature then.Other is identical with embodiment 6.Guided missile missile wing aerofoil ratio of rigidity after increasing had just improved 10% originally.
Embodiment 9: the lifting airscrew aerofoil increases just
Selecting M40 carbon fiber (modulus is 392GPa) is high modulus fibre, adopt the pre-dipping machine of hot melt to be placed into the lifting airscrew aerofoil automatically on the M40 carbon fiber and put T300 thermoplastic carbon fiber polyimides prepreg surface by 6 pieces/centimetre spacing, obtain being stained with the T300 thermoplastic carbon fiber polyimides prepreg surface of M40 carbon fiber, be designated as high-modulus shop layer; Form new airfoil structure fiber lay down layer after replacing middle six layers of lifting airscrew aerofoil fibrillation shop layer with high-modulus shop layer then, adopt the original die press technology for forming moulding of lifting airscrew aerofoil, be specially and under pressure 0.50MPa～0.60MPa, be warming up to 380 ℃～400 ℃, and after keeping 1h, heating rate is＜2 ℃/min, reduce to room temperature then, can obtain increasing firm lifting airscrew aerofoil.Lifting airscrew aerofoil ratio of rigidity after increasing had just improved 10% originally.
Claims (7)
1. a method that improves composite airfoil structure rigidity is characterized in that, the operating procedure of this method is as follows:
(1) selects fiber according to fiber modulus: based on the fortifying fibre modulus＜280GPa of composite airfoil structure, selection is than the high at least one grade fiber of composite airfoil structure fortifying fibre modulus, is that carbon fiber or the modulus of 280GPa～650GPa is that 350GPa～400GPa boron fibre or modulus are 350GPa～450GPa silicon carbide fibre comprising modulus;
(2) spacing of selected fiber is set: when adopting three kinds of different fibers, interfibrous spacing is, the boron fibre spacing is 10 pieces/centimetre～100 pieces/centimetre, and the silicon carbide fibre spacing is 10 pieces/centimetre～100 pieces/centimetre, the spacing of carbon fiber be 1 the bundle/centimetre～10 the bundle/centimetre;
(3) lay of selected fiber: with the fiber selected in the step (1) set by step the spacing in (2) by hand or the direct lay of method placed continuously of machinery to the fiber lay down laminar surface of airfoil structure, and the resin solution that is complementary with resin matrix at its surperficial spraying or brushing and airfoil structure, quantity for spray is 2g/m
2～10g/m
2, the position for the treatment of permanent domicile selection fiber after solvent volatilizees fully in the resin solution; The fiber lay down layer of preparing the airfoil structure that is stained with selected fiber is designated as high-modulus shop layer, and substituting former airfoil structure shop number of plies amount with high-modulus shop layer segment is the fiber lay down layer of n＞3, and it substitutes 1/n～2/3 of amount for fibrillation shop number of plies amount;
(4) composite airfoil structure moulding is spread layer with the high-modulus of preparation in the step (3) and airfoil structure remains the initial shop layer design laying that fibrillation shop layer is pressed airfoil structure; When fibrillation shop layer is prepreg or preimpregnation cloth,, and when fibrillation shop layer is fabric, adopt infiltration moulding of resin transfer molding or resin molding or vacuum bag moulding with autoclave molding or compression molding or moulding of thermal change type or diaphragm forming; The heating rate of above-mentioned moulding process is＜2 ℃/mi n.
2. a kind of method that improves composite airfoil structure rigidity according to claim 1, it is characterized in that, described airfoil structure resin matrix is thermosetting resin or thermoplastic resin or thermosetting resin and the mixing of thermoplastic resin, thermosetting resin: epoxy resin, bimaleimide resin, thermoset polyimide resin benzoxazine colophony, unsaturated polyester resin, phenolic resins or cyanate ester resin, thermoplastic resin: TPI, PEI, polyphenylene sulfide, polyether-ether-ketone, polyether-ketone, polyether sulfone, polyethylene or High molecular weight polyethylene, aforementioned hot thermosetting resin and thermoplastic resin ratio are 1: 20～1: 2 mixture.
3. a kind of method that improves composite airfoil structure rigidity according to claim 1, it is characterized in that, the fiber lay down layer of described airfoil structure is 1) fabric, comprise unidirectional or plain weave or satin weave or twills, 2) unidirectional pre-immersion material or above-mentioned 1) described in the preimpregnation cloth of fabric.
4. a kind of method that improves composite airfoil structure rigidity according to claim 1 is characterized in that the quality percentage composition of described resin solution is 1%～20%.
5. a kind of method that improves composite airfoil structure rigidity according to claim 1 is characterized in that, when prepared layer Central Plains, high-modulus shop fiber lay down layer is prepreg or preimpregnation cloth, does not spray or brush resin solution.
6. a kind of method that improves composite airfoil structure rigidity according to claim 1, it is characterized in that, ply sequence behind the layer of the alternative fibrillation of layer shop, described high-modulus shop is one of following: 1) when airfoil structure has only single face as working face, high-modulus shop layer is laid in the position, top layer of composite airfoil structure continuously, and the high modulus fibre warp-wise in the layer of high-modulus shop is along the aerofoil length direction; 2) two-sided during as working face when airfoil structure, high-modulus shop layer is laid in the middle part of composite airfoil structure continuously, and the high modulus fibre warp-wise in the layer of high-modulus shop is along the aerofoil length direction.
7. a kind of method that improves composite airfoil structure rigidity according to claim 1, it is characterized in that, when fibrillation shop layer is prepreg or preimpregnation cloth, the forming temperature of described autoclave molding method is 80 ℃～300 ℃, pressure is 0.20MPa～0.60MPa, described mould pressing process forming temperature is 80 ℃～300 ℃, briquetting pressure is 0.20MPa～1.0MPa, described thermal change type technological forming type is 80 ℃～300 ℃, briquetting pressure is 0.20MPa～0.40MPa, described diaphragm forming technological forming temperature is 80 ℃～300 ℃, briquetting pressure is 0.20MPa～0.70MPa, and when fibrillation shop layer is fabric, the injection temperature of described resin transfer molding technology is 30 ℃～300 ℃, injection pressure is 0.05MPa～0.70MPa, and the forming temperature of described resin molding Infiltration Technics is 80 ℃～300 ℃, 80 ℃～300 ℃ of the moulding of described vacuum bag moulding process.
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