A large-size aircraft central wing skeleton co-bonding autoclave forming method

Abstract

This invention discloses a method for co-bonding and autoclaving a large-size aircraft center wing skeleton, relating to the field of aerospace technology. The wing skeleton is formed by co-bonding and curing several parallel I-beam skeleton beams and several single / double-sided skeleton ribs perpendicular to the skeleton beams. The I-beam skeleton beams are cured, while the skeleton ribs are pre-formed. The I-beam skeleton beams are laid and cured during part molding; the skeleton ribs are laid on a rubber soft mold. The pre-formed skeleton ribs are positioned and combined with several skeleton beams through positioning holes and then cured. This molding method has the advantages of easy demolding and low cost, ensuring not only the molding accuracy and non-destructive quality of the skeleton but also reducing the manufacturing cost and cycle time of the parts.

Description

Technical Field

[0001] This invention relates to the field of advanced composite material manufacturing technology, and in particular to a method for co-bonding autoclave molding of a large-size aircraft center wing skeleton. Background Technology

[0002] Unmanned aerial vehicles (UAVs) have gradually become an important research hotspot in the aviation field in recent years. Carbon fiber composite materials, with their characteristics of high specific strength, high specific modulus, fatigue resistance, and corrosion resistance, are widely used in the aerospace industry. Lightweight design can effectively improve the flight performance of UAVs.

[0003] As a key structural component of unmanned aerial vehicles (UAVs), the wing primarily provides lift, thus requiring excellent impact resistance. The wing frame mainly consists of trusses and ribs. Typically, the wing frame molding process employs secondary adhesive bonding or riveting, where the trusses and ribs are cured separately and then assembled using adhesive film or rivets. However, using standard parts to rivet the trusses and ribs increases the weight of the wing frame. Furthermore, composite materials suffer severe damage and weakening after machining, resulting in reduced interlaminar shear strength; therefore, adhesive bonding is the most common method for joining composite materials.

[0004] Currently, composite material molding processes are mainly divided into three bonding methods: (1) secondary bonding, (2) co-bonding, and (3) co-curing. Among them, secondary bonding is the most widely used, but it has disadvantages such as long molding cycle and many assembly times. Co-curing molding process cannot be applied to the design and manufacturing of complex wing frames. Summary of the Invention

[0005] The purpose of this invention is to provide a co-bonded autoclave molding method for a large-size aircraft center wing skeleton. This molding process has the advantages of easy demolding and low cost. It can not only ensure the molding accuracy and non-destructive quality of the skeleton, but also reduce the manufacturing cost of parts and expand the application of composite materials in the field of unmanned aerial vehicles.

[0006] To solve the above-mentioned technical problems, the present invention provides a method for co-bonding autoclave molding of a large-size aircraft center wing skeleton, comprising the following steps:

[0007] Step A: Mold design. Based on the digital model of the composite material component of the central wing frame, its structure is divided into I-beam frame beams and frame ribs. The corresponding mold laying fixtures are designed according to the structural morphology of the divided I-beam frame beams and frame ribs. Among them, the I-beam frame beams are solidified and molded using part forming fixtures, and the frame ribs are pre-pressed using rubber soft molds.

[0008] Step B: I-beam frame beam forming. Prepreg is laid on the part forming fixture between the two side baffles. After laying, it is assembled and positioned by positioning guide pins. After positioning, C-shaped prepreg of the I-beam frame beam is laid. After every 3 layers of prepreg are laid, vacuum bag process is used to compact it.

[0009] Step C: Single-sided skeleton rib forming. Lay skeleton rib prepreg on a single-sided rubber soft mold. After every 3 layers of skeleton rib prepreg are laid, vacuum bag process is required for compaction.

[0010] Step D: Double-sided skeleton rib molding. Skeleton rib prepreg is laid on both the upper and lower rubber molds. After completion, the mold is closed by positioning guide pins and then compacted using a vacuum bag process. Prepreg is laid on the side edges after the mold is closed. After every 3 layers of skeleton rib prepreg are laid, vacuum bag process is required for compaction.

[0011] Step E: Assembly, the cured I-beam frame beam and frame ribs are assembled, positioned and sealed using positioning guide pins;

[0012] Step F: Co-bonding and curing. Apply a layer of adhesive film to the bonding area of ​​the I-beam frame beam and frame ribs. Then, assemble, seal, and cure the cured I-beam frame beam and frame ribs according to the positioning grooves.

[0013] Step G: Demolding and bonding are completed. After the rubber soft mold on the skeleton ribs is removed, the central wing skeleton composite material component is obtained.

[0014] Preferably, in the I-beam frame beam forming step, the I-beam frame beam is laid in three parts: First, the lower C-shaped prepreg of the I-beam frame beam is laid on the part forming fixture; Second, the two side edge strips are laid on the side baffles. After laying, they are positioned and assembled according to the positioning guide pins, and carbon twisted wire is filled into the triangular gaps on both sides; Third, after the carbon twisted wire is filled, the upper C-shaped prepreg is laid. After laying, it is sealed and cured on the part forming fixture.

[0015] Preferred curing parameters for the I-beam frame beam: heating from room temperature to 80℃, heating rate ≤2℃ / min, holding for 30min, holding temperature difference 80±3℃, pressurizing to 0.6Mpa after holding, heating to 100℃ at a rate of 1.5℃ / min, then heating to 125℃ at a rate of 0.8℃ / min, holding for 90min, holding temperature deviation 125±6℃.

[0016] Preferably, in the skeleton rib forming step, the skeleton ribs are divided into two types: one type is single-sided box-shaped, and the other type is double-sided box-shaped.

[0017] Single-sided box-shaped skeleton ribs are laid on a single-sided rubber soft mold with skeleton rib prepreg;

[0018] The double-sided box-shaped skeleton ribs are laid in three parts. First, the prepreg of the skeleton ribs on both sides is laid on the corresponding rubber soft molds. After the laying is completed, the rubber soft molds on both sides are combined according to the positioning holes. Second, carbon twisted wire is filled into the triangular gaps on both sides. Third, the edge strip is laid after the filling is completed.

[0019] Preferably, in the co-bonding curing step, the process parameters are: pressure 0.3-0.8 MPa, curing temperature 125±6℃, curing time 120-180 min, and pressure gradually increased from 0.1 MPa to the applied pressure.

[0020] Preferably, in the co-bonding curing step, the curing parameters are as follows: room temperature to 80℃, heating rate ≤2℃ / min, holding for 30min, holding temperature difference 80±3℃, pressurizing to 0.5Mpa after holding, heating to 100℃ at a rate of 1.25℃ / min, heating to 125℃ at a rate of 0.8℃ / min, holding for 90min, and holding temperature deviation 125±6℃.

[0021] Preferably, the C-shaped prepreg of the I-beam frame beam and the prepreg of the frame ribs are both made of carbon fiber fabric, carbon fiber unidirectional tape and film combination.

[0022] Preferably, the rubber mold is made of flexible high-temperature resistant rubber material, and its shape is a box-shaped structure consistent with the skeleton ribs. The material is silicone rubber, with a minimum temperature resistance of 190℃ and an elongation of ≥300%.

[0023] Compared with the prior art, the purpose of this invention is to complete the manufacturing process of the central wing frame of the co-bonded composite material by using a rubber soft film, part forming tooling, and assisted autoclave forming, thereby achieving the following requirements:

[0024] 1. Shorten the parts forming cycle and reduce the number of parts entering the tank and assembling;

[0025] 2. Avoid riveting standard parts to increase their weight;

[0026] 3. Solve the problem of assembly tolerance allocation;

[0027] 4. Avoid contact corrosion between standard parts and composite materials. Attached Figure Description

[0028] Figure 1 This is a schematic diagram of the forming process of the I-beam frame beam provided by the present invention;

[0029] Figure 2 This is a schematic diagram of the double-sided skeleton rib structure provided by the present invention;

[0030] Figure 3This is a schematic diagram of the double-sided skeleton rib forming provided by the present invention;

[0031] Figure 4 This is a schematic diagram of the co-bonding molding of the I-beam frame beam and the frame ribs provided by the present invention;

[0032] Figure 5 This is a structural schematic diagram of the composite material component of the central wing frame provided by the present invention. Detailed Implementation

[0033] The present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. The advantages and features of the present invention will become clearer from the following description and claims. It should be noted that the drawings are all in a very simplified form and use non-precise proportions, and are only used to facilitate and clarify the illustration of the embodiments of the present invention.

[0034] In the description of this 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 the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this invention.

[0035] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; and they can refer to the internal connection of two components. Those skilled in the art will understand the specific meaning of the above terms in this invention based on the specific circumstances. Example

[0036] This invention provides a method for co-bonding autoclave molding of a large-size aircraft center wing skeleton. Please refer to [link / reference]. Figure 1-5 It includes the following steps:

[0037] Step A: Mold design. Based on the digital model of the composite material component of the central wing frame, its structure is divided into I-beam frame beams and frame ribs. The corresponding mold laying fixtures are designed according to the structural morphology of the divided I-beam frame beams and frame ribs. Among them, the I-beam frame beams are solidified and molded using part forming fixtures, and the frame ribs are pre-pressed using rubber soft molds.

[0038] Step B: I-beam frame beam forming. Prepreg is laid on the part forming fixture between the two side baffles. After laying, it is assembled and positioned by positioning guide pins. After positioning, C-shaped prepreg of the I-beam frame beam is laid. After every 3 layers of prepreg are laid, vacuum bag process is used to compact it.

[0039] Step C: Single-sided skeleton rib forming. Lay skeleton rib prepreg on a single-sided rubber soft mold. After every 3 layers of skeleton rib prepreg are laid, vacuum bag process is required for compaction.

[0040] Step D: Double-sided skeleton rib molding. Skeleton rib prepreg is laid on both the upper and lower rubber molds. After completion, the mold is closed by positioning guide pins and then compacted using a vacuum bag process. Prepreg is laid on the side edges after the mold is closed. After every 3 layers of skeleton rib prepreg are laid, vacuum bag process is required for compaction.

[0041] Step E: Assembly, the cured I-beam frame beam and frame ribs are assembled, positioned and sealed using positioning guide pins;

[0042] Step F: Co-bonding and curing. Apply a layer of adhesive film to the bonding area of ​​the I-beam frame beam and frame ribs. Then, assemble, seal, and cure the cured I-beam frame beam and frame ribs according to the positioning grooves.

[0043] Step G: Demolding and bonding are completed. After the rubber soft mold on the skeleton ribs is removed, the central wing skeleton composite material component is obtained.

[0044] Specifically, in the forming process of the I-beam frame beam, the I-beam frame beam is laid in three parts: First, the C-shaped prepreg on the lower side of the I-beam frame beam is laid on the part forming fixture; Second, the side edge strips are laid on the side baffles. After laying, they are positioned and assembled according to the positioning guide pins, and carbon twisted wire is filled into the triangular gaps on both sides; Third, after the carbon twisted wire is filled, the upper C-shaped prepreg is laid. After laying, it is sealed and cured on the part forming fixture.

[0045] Specifically, the curing parameters for the I-beam frame beam are as follows: heat from room temperature to 80℃, heating rate ≤ 2℃ / min, hold for 30min, temperature difference during holding 80±3℃, pressurize to 0.6Mpa after holding, heat to 100℃ at a rate of 1.5℃ / min, then heat to 125℃ at a rate of 0.8℃ / min, hold for 90min, and temperature deviation 125±6℃.

[0046] Specifically, in the skeleton rib forming process, the skeleton ribs are divided into two types: single-sided box-shaped and double-sided box-shaped. For single-sided box-shaped skeleton ribs, skeleton rib prepreg is laid on a single-sided rubber mold. For double-sided box-shaped skeleton ribs, the laying process is divided into three parts: First, the skeleton rib prepreg on both sides is laid on the corresponding rubber molds, and after laying, the two rubber molds are assembled according to the positioning holes; second, carbon twisted wire is filled into the triangular gaps on both sides; third, after filling, the edging strip is laid.

[0047] Specifically, in the co-bonding and curing step, the process parameters are: pressure 0.3-0.8 MPa, curing temperature 125±6℃, curing time 120-180 min. When applying pressure, gradually increase it from 0.1 MPa to the applied pressure to prevent uneven stress on the rubber mold caused by direct pressure application.

[0048] In some embodiments, during the co-bonding curing step, the curing parameters are as follows: room temperature is increased to 80°C, heating rate is ≤2°C / min, temperature is maintained for 30 min, temperature difference during maintenance is 80±3°C, pressure is applied to 0.5 MPa after maintenance, and the temperature is increased to 100°C at a rate of 1.25°C / min, then increased to 125°C at a rate of 0.8°C / min, temperature is maintained for 90 min, and temperature deviation is 125±6°C.

[0049] Specifically, the C-shaped prepreg of the I-beam frame beam and the prepreg of the frame ribs are both made of carbon fiber fabric, carbon fiber unidirectional tape and film combination.

[0050] Specifically, the rubber mold is made of flexible high-temperature resistant rubber material, and its shape is a box-shaped structure consistent with the skeleton ribs. The material is silicone rubber, with a minimum temperature resistance of 190℃ and an elongation of ≥300%.

[0051] The above description is merely a description of preferred embodiments of the present invention and is not intended to limit the scope of the present invention in any way. Any changes or modifications made by those skilled in the art based on the above disclosure shall fall within the protection scope of the claims.

Claims

1. A method for co-bonding autoclave molding of a large-size aircraft center wing skeleton, characterized in that, Includes the following steps: Step A: Mold design. Based on the digital model of the composite material component of the central wing frame, its structure is divided into I-beam frame beams and frame ribs. The corresponding mold laying fixtures are designed according to the structural morphology of the divided I-beam frame beams and frame ribs. Among them, the I-beam frame beams are solidified and molded using part forming fixtures, and the frame ribs are pre-pressed using rubber soft molds. Step B: I-beam frame beam forming. Prepreg is laid on the part forming fixture between the two side baffles. After laying, it is assembled and positioned by positioning guide pins. After positioning, C-shaped prepreg of the I-beam frame beam is laid. After every 3 layers of prepreg are laid, vacuum bag process is used to compact it. Step C: Single-sided skeleton rib forming. Lay skeleton rib prepreg on a single-sided rubber soft mold. After every 3 layers of skeleton rib prepreg are laid, vacuum bag process is required for compaction. Step D: Double-sided skeleton rib molding. Skeleton rib prepreg is laid on both the upper and lower rubber molds. After completion, the mold is closed by positioning guide pins and then compacted using a vacuum bag process. Prepreg is laid on the side edges after the mold is closed. After every 3 layers of skeleton rib prepreg are laid, vacuum bag process is required for compaction. Step E: Assembly, the cured I-beam frame beam and frame ribs are assembled, positioned and sealed using positioning guide pins; Step F: Co-bonding and curing. Apply a layer of adhesive film to the bonding area of ​​the I-beam frame beam and frame ribs. Then, assemble, seal, and cure the cured I-beam frame beam and frame ribs according to the positioning grooves. Step G: Demolding and bonding are completed. After the rubber soft mold on the skeleton ribs is removed, the central wing skeleton composite material component is obtained.

2. The method for co-bonding autoclave molding of a large-size aircraft center wing skeleton as described in claim 1, characterized in that, In the forming process of the I-beam frame beam, the I-beam frame beam is laid in three parts: First, the C-shaped prepreg on the lower side of the I-beam frame beam is laid on the part forming fixture; Second, the side edge strips are laid on the side baffles. After the laying is completed, they are positioned and assembled according to the positioning guide pins, and carbon twisted wire is filled into the triangular gaps on both sides; Third, after the carbon twisted wire is filled, the upper C-shaped prepreg is laid. After the laying is completed, it is sealed and cured on the part forming fixture.

3. The method for co-bonding autoclave molding of a large-size aircraft center wing skeleton as described in claim 2, characterized in that, Curing parameters for the I-beam frame: Heating from room temperature to 80℃ at a rate of ≤2℃ / min, holding for 30min, with a temperature difference of 80±3℃, pressurizing to 0.6MPa after holding, heating to 100℃ at a rate of 1.5℃ / min, then heating to 125℃ at a rate of 0.8℃ / min, holding for 90min, with a temperature deviation of 125±6℃.

4. The method for co-bonding autoclave molding of a large-size aircraft center wing skeleton as described in claim 1, characterized in that, In the skeleton rib forming step, the skeleton ribs are divided into two categories: one is single-sided box-shaped, and the other is double-sided box-shaped. Single-sided box-shaped skeleton ribs are laid on a single-sided rubber soft mold with skeleton rib prepreg; The double-sided box-shaped skeleton ribs are laid in three parts. First, the prepreg of the skeleton ribs on both sides is laid on the corresponding rubber soft molds. After the laying is completed, the rubber soft molds on both sides are combined according to the positioning holes. Second, carbon twisted wire is filled into the triangular gaps on both sides. Third, the edge strip is laid after the filling is completed.

5. The method for co-bonding autoclave molding of a large-size aircraft center wing skeleton as described in claim 1, characterized in that, In the co-bonding curing step, the process parameters are: pressure 0.3-0.8 MPa, curing temperature 125±6℃, curing time 120-180 min, and pressure gradually increased from 0.1 MPa to the applied pressure.

6. The method for co-bonding autoclave molding of a large-size aircraft center wing skeleton as described in claim 5, characterized in that, In the co-bonding curing step, the curing parameters are as follows: temperature rise from room temperature to 80℃, heating rate ≤2℃ / min, hold for 30min, temperature difference during holding 80±3℃, pressurize to 0.5Mpa after holding, and heat up to 100℃ at a rate of 1.25℃ / min, then heat up to 125℃ at a rate of 0.8℃ / min, hold for 90min, and temperature deviation 125±6℃.

7. The method for co-bonding autoclave molding of a large-size aircraft center wing skeleton as described in claim 1, characterized in that, The C-shaped prepreg of the I-beam frame beam and the prepreg of the frame ribs are both made of carbon fiber fabric, carbon fiber unidirectional tape and film combination.

8. The method for co-bonding autoclave molding of a large-size aircraft center wing skeleton as described in claim 1, characterized in that, The rubber mold is made of flexible, high-temperature resistant rubber. Its shape is a box-shaped structure consistent with the skeleton ribs. The material is silicone rubber, with a minimum temperature resistance of 190℃ and an elongation of ≥300%.

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