A manufacturing method of a carbon fiber wing main load-bearing C-shaped long beam

Abstract

The present application relates to a kind of carbon fiber wing main force C type long beam manufacturing method, comprising: S1. the Invar alloy steel with small thermal expansion coefficient is selected, and forming mould is made by five-axis numerical control milling equipment;S2. carbon fiber woven fabric prepreg and carbon fiber unidirectional tape prepreg of medium temperature system are alternately laid and stacked on the forming mould of S1, and the wing long beam of entire C type structure is made;S3. the wing long beam of S2 laying and stacking is bagged, and is cured into shape in curing oven.Through the method, the problem of manufacturing difficulty of full composite wing main force C type long beam is solved, and by using carbon fiber woven fabric prepreg and carbon fiber unidirectional tape prepreg of medium temperature system, the face profile of curing and forming long beam and the deformation amount of effective control are guaranteed by using Invar alloy steel with little deformation with temperature variation, the quality of long beam is improved, and the energy consumption cost of the same temperature hot press tank process is reduced.

Description

TECHNICAL FIELD

[0001] The present application belongs to the technical field of aircraft manufacturing, and particularly relates to a manufacturing method of a carbon fiber wing main load-bearing C-shaped long beam. BACKGROUND

[0002] At present, the wing main beam is a main load-bearing part of the wing. The air dynamic force of the aircraft in flight, the inertial force generated in maneuvering flight, and the impact force of the landing gear in landing make the main beam bear huge bending and shearing force. Therefore, the manufacturing quality of the wing main beam part directly affects the flight performance and personnel safety of the whole aircraft.

[0003] With the rapid development of global general aviation, the types of composite aircraft are increasing, and the materials of the wing beam are also increasingly changing from metal materials to composite materials. However, the applicant finds that due to the long length and thick thickness of the wing beam, these long and thick wing beams become the difficulty and key point in the whole composite aircraft manufacturing process. During manufacturing, multiple curing forming is often used. If the manufacturing process method is incorrect, defects such as part distortion, internal porosity and delamination of the part will be generated. These problems will affect the strength of the part, cause the main beam to fatigue under the action of the alternating stress generated by the wing vibration, expand the defects, and even cause the wing beam to break, resulting in aircraft damage and casualties. SUMMARY

[0004] To solve the above problems in the prior art, the present application provides a manufacturing method of a carbon fiber wing main load-bearing C-shaped long beam.

[0005] To solve the above technical problems, the present application adopts the following technical scheme:

[0006] The present application provides a manufacturing method of a carbon fiber wing main load-bearing C-shaped long beam, comprising:

[0007] S1. Wing beam forming die manufacturing: an Invar alloy steel with a small thermal expansion coefficient is selected to manufacture a forming die through a five-axis numerical control milling equipment;

[0008] S2. Wing beam laying and stacking: a carbon fiber woven fabric prepreg and a carbon fiber unidirectional tape prepreg of a medium temperature system are laid and stacked on the forming die obtained in S1 to manufacture a wing long beam with a whole C-shaped structure. The thickest area is 25mm thick. All the layups have station points in the spanwise direction to form a state of thick in the middle and thin at both ends. Moreover, the carbon fiber woven fabric prepreg and the carbon fiber unidirectional tape prepreg of the medium temperature system are laid in an alternating manner.

[0009] S3. Wing beam curing forming: the wing long beam laid and stacked in S2 is bagged and pushed into a curing furnace for curing forming. In this process, the vacuum degree of the solid laminate is not less than 0.08Mpa, and the leakage rate in the vacuum bag is less than 6.77Kpa / 5min.

[0010] Further, in S2, each layer of the laid carbon fiber woven fabric prepreg covers the upper and lower edge strips and the beam web area of the forming mold, the carbon fiber unidirectional tape prepreg is laid from the upper and lower edge strip area of the forming mold to the beam web area, and stops at the end point of the beam web area, and each three layers are progressively laid or reduced, and each two layers of carbon fiber woven fabric prepreg are sandwiched by three layers of carbon fiber unidirectional tape prepreg.

[0011] Further, in S2, after laying the first layer of material, pre-compaction is performed, and then pre-compaction is performed once every three layers of the whole layer, and the vacuum pressure is required to be greater than 0.08 MPa and the time is required to be greater than 15 min, and then hot compaction is performed once when the layer is laid to 45-50 layers, the temperature is 30-35℃, the vacuum pressure and the time are consistent with the pre-compaction, and the whole process of hot compaction is not more than twice.

[0012] Further, in S3, a three-step curing curve is used: a holding point is set at 80℃, holding for 4-5 hours; a holding point is set at 100℃, holding for 1-2 hours; a holding point is set at 130℃, holding for at least 2 hours; the maximum heating rate should not exceed 0.8℃ / min, the minimum heating rate should not be lower than 0.2℃ / min, the heating rate from room temperature to 50℃ is not limited, and the cooling rate should not exceed 1℃ / min.

[0013] Further, the forming mold made in S1 has a thickness of 10-15 mm.

[0014] The present application has the following beneficial effects:

[0015] By adopting the above technical scheme, the problem of manufacturing the full composite wing main load-bearing C-shaped long beam can be solved, and by using the carbon fiber woven fabric prepreg and the carbon fiber unidirectional tape prepreg of the medium temperature system and the Invar alloy steel with little deformation with temperature, on the one hand, the face contour and the deformation amount of the carbon fiber wing main load-bearing C-shaped long beam can be greatly guaranteed and controlled, thereby improving the quality of the carbon fiber wing main load-bearing C-shaped long beam; on the other hand, the energy consumption cost can be significantly reduced compared with the same temperature hot pressing tank process. BRIEF DESCRIPTION OF DRAWINGS

[0016] Figure 1 is a flowchart of the manufacturing method of the carbon fiber wing main load-bearing C-shaped long beam according to the present application;

[0017] Figure 2 is a structural schematic diagram of the wing spar forming mold in the manufacturing method of the carbon fiber wing main load-bearing C-shaped long beam according to the present application. DETAILED DESCRIPTION

[0018] In order to make the purpose, technical scheme and advantages of the present application more clear, the present application is further described in detail below with reference to the drawings and examples. It should be understood that the specific examples described herein are only used to explain the present application and not to limit the present application.

[0019] As shown in Figure 1 The manufacturing method of the carbon fiber wing main load-bearing C-shaped long beam of the present application specifically comprises the following steps:

[0020] Step S1. Manufacture of beam forming die: select Invar alloy steel with small thermal expansion coefficient (such as: Invar alloy, also known as non-expansion steel, with an average expansion coefficient of 1.5x10-6℃) to make a forming die by five-axis numerical control milling equipment; wherein the thickness of the forming die is 10mm-15mm, and the thickness is uniform, which can ensure the temperature uniformity of the die during heating process.

[0021] As shown in Figure 2 The outer surfaces of the entire wing beam upper and lower edge strip area 1, beam web area 2 and fillet area 4 together form a die surface, and the tolerance of the tool forming surface profile is ±0.1mm; the roughness of the tool forming surface is 0.5μm, and the roughness of the non-forming surface is 3.2μm.

[0022] Step S2. Beam laying and stacking: carbon fiber woven fabric prepreg and carbon fiber unidirectional tape prepreg of medium temperature system are laid and stacked on the forming die made in S1 to make a wing long beam with C-shaped structure, the thickest area is 25mm thick, all layers have station points along the spanwise direction, forming a state of thick in the middle and thin at both ends, different station points can be directly marked on the tool allowance area, and the carbon fiber woven fabric prepreg and carbon fiber unidirectional tape prepreg of medium temperature system are laid and stacked in an alternating manner; wherein the temperature during the entire laying process should be controlled at 18℃-26℃, and the relative humidity is ≤65%; the carbon fiber woven fabric prepreg and carbon fiber unidirectional tape prepreg of medium temperature system can use the carbon fiber woven fabric prepreg and carbon fiber unidirectional tape prepreg of medium temperature system of Toray Company, such as: T700S-12K-50C / #2510 and T700G-12K-31E / #2510, the curing temperature of the material is 130℃.

[0023] Step S3. Beam curing forming: the wing long beam laid and stacked in S2 is bagged (the bagging material selects a material resistant to temperature above 150℃), and is pushed into a curing oven for curing forming; in this process, the vacuum degree of the solid laminate is not less than 0.08Mpa, and the leakage rate in the vacuum bag is less than 6.77Kpa / 5min.

[0024] The method solves the problem of manufacturing the full composite wing main load-bearing C-shaped long beam, and by using the carbon fiber woven fabric prepreg and the carbon fiber unidirectional tape prepreg of the medium temperature system and the Invar alloy steel with little deformation with temperature, the face profile of the carbon fiber wing main load-bearing C-shaped long beam is greatly ensured, the deformation is effectively controlled, the quality of the carbon fiber wing main load-bearing C-shaped long beam is improved, the high precision requirement of the aircraft manufacturing is met, and the energy consumption cost of curing is saved by 70% compared with the same temperature autoclave process.

[0025] In a possible implementation, in S2, each layer of the laid carbon fiber woven fabric prepreg covers the upper and lower edge strips 1 and the beam web area 2 of the beam of the forming mold, and the carbon fiber unidirectional tape prepreg is laid from the upper and lower edge strip areas 1 of the forming mold to the beam web area 2, and stops at the end point 3 of the beam web area 2, as shown in Figure 2 ; and every three layers are progressively laid or reduced, and every two layers of the carbon fiber woven fabric prepreg are sandwiched with three layers of the carbon fiber unidirectional tape prepreg. By this method, the strength of the carbon fiber wing main load-bearing C-shaped long beam can be further improved to meet the strength requirement of the wing main load-bearing long beam in aircraft manufacturing.

[0026] In a possible implementation, to minimize the porosity of the prepared carbon fiber wing main load-bearing C-shaped long beam and further improve the strength of the carbon fiber wing main load-bearing C-shaped long beam, in S2, after laying the first layer of material, pre-compaction is performed, and then pre-compaction is performed once every three layers, the vacuum pressure is greater than 0.08 MPa, and the time is greater than 15 min, and then hot compaction is performed once when the layer is 45-50, the temperature is 30-35°C, the vacuum pressure and the time are the same as the pre-compaction, and the whole process is not more than twice.

[0027] In a possible implementation, in S3, a three-step curing curve is used: a holding point is set at 80°C, and the holding time is 4-5 hours; a holding point is set at 100°C, and the holding time is 1-2 hours; a holding point is set at 130°C, and the holding time is at least 2 hours; the maximum heating rate should not exceed 0.8°C / min, the minimum heating rate should not be lower than 0.2°C / min, the heating rate from room temperature to 50°C is not limited, and the cooling rate should not exceed 1°C / min. By using the three-step curing curve, the resin can flow fully in the curing process, and the porosity is further reduced.

[0028] The above is the preferred embodiment of the present application. It should be noted that those skilled in the art can make some improvements and refinements without departing from the principles of the present application, and these improvements and refinements are also considered within the protection scope of the present application.

Claims

1. A method for manufacturing a C-shaped long beam that serves as the main load-bearing component for a carbon fiber wing, characterized in that, include: S1. Wing Spar Forming Mold Manufacturing: Invar alloy steel with a low coefficient of thermal expansion is selected to manufacture the forming mold using a five-axis CNC milling machine. The tooling forming surface tolerance requirement of the mold is ±0.1mm; the tooling forming surface roughness requirement is 0.5μm, and the non-forming surface is 3.2μm. S2. Wing Spall Laying and Stacking: Medium-temperature carbon fiber woven prepreg and unidirectional carbon fiber tape prepreg are laid and stacked onto the molding die obtained in S1 to create the entire C-shaped wing spall. The thickest area is 25mm thick. All layers have anchor points along the spanwise direction, resulting in a thicker middle and thinner ends. The medium-temperature carbon fiber woven prepreg and unidirectional carbon fiber tape prepreg are laid alternately. Each layer of carbon fiber woven prepreg completely covers the upper and lower edge strips of the molding die and the web area of ​​the spall. The unidirectional carbon fiber tape prepreg... The prepreg is laid from the upper and lower edge strips of the wing beam of the molding die to the web area of ​​the beam. The layering is progressive or reduced every three layers. Three layers of unidirectional carbon fiber tape prepreg are sandwiched between every two layers of carbon fiber woven prepreg. After the first layer of material is laid, pre-compaction is performed. Then, pre-compaction is performed every three layers. The vacuum pressure is required to be greater than 0.08 MPa and the time is greater than 15 minutes. After laying 45 to 50 layers, hot compaction is performed at a temperature of 30℃ to 35℃. The vacuum pressure and time are the same as those for pre-compaction. The hot compaction process does not exceed two times in the entire process. S3. Wing Spall Curing: The stacked wing spalls from S2 are bagged and cured in a curing oven. During this process, the vacuum degree of the solid laminate is not less than 0.08 MPa, and the leakage rate inside the vacuum bag is less than 6.77 kPa / 5 min. A three-step curing curve is used: a heat preservation point is set at 80℃ for 4-5 hours; a heat preservation point is set at 100℃ for 1-2 hours; and a heat preservation point is set at 130℃ for at least 2 hours. The maximum heating rate should not exceed 0.8℃ / min, the minimum heating rate should not be less than 0.2℃ / min, the heating rate from room temperature to 50℃ is not limited, and the cooling rate should not exceed 1℃ / min.

2. The method according to claim 1, characterized in that, The molding die made by S1 has a thickness of 10mm to 15mm.

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