Carbon-glass hybrid composite skin for an aircraft and process for the production thereof

Abstract

The application provides a carbon-glass hybrid composite material skin for an aircraft and a preparation process thereof, and relates to the technical field of aircraft accessories.A carbon-glass hybrid composite material skin for an aircraft, the material of the skin is a mixture of carbon fibers, glass fibers and resin, and the skin is an integrally formed structure; the mass percentage of the mixture of carbon fibers and glass fibers ranges from 60% to 70%, and the mass percentage of the resin ranges from 40% to 30%; the skin is provided with a non-thickened area and a thickened area.The application solves the technical problems of the prior art, such as the heavy weight of the aircraft skin made of metal material, the complex structure design, the poor impact resistance and stability of the product during use.The material of the skin is a mixture of carbon fibers, glass fibers and resin, and the skin is integrally formed, so that the weight is light and the overall structural strength is improved.

Description

Technical Field

[0001] This invention relates to the technical field of aircraft components, and in particular to a carbon-glass hybrid composite material skin for aircraft and its preparation process. Background Technology

[0002] Aircraft skin refers to the shaped components that surround the aircraft's frame structure and are fixed to the frame with adhesives or rivets, forming the aircraft's aerodynamic shape. The skin structure formed by the aircraft skin and the frame has high load-bearing capacity and rigidity, is lightweight, and plays a role in bearing and transmitting aerodynamic loads.

[0003] In the aforementioned prior art, in order to improve load-bearing capacity and rigidity, aircraft skin is mostly made of metal materials, and the thickness of the skin is the same at all locations.

[0004] However, the heavy weight of the metal aircraft skin contributes to the overall weight of the aircraft; furthermore, the complex structural design of the aircraft skin requires the fabrication of multiple plates and the use of multiple connecting structural components for fixation, resulting in a complex manufacturing process; the uniform thickness of the skin throughout its structure leads to poor load-bearing capacity at the connection points. In use, the product exhibits poor impact resistance and stability. Summary of the Invention

[0005] The purpose of this invention is to provide a carbon-glass hybrid composite material skin for aircraft, in order to solve the technical problems existing in the prior art, such as heavy weight of metal aircraft skin, complex structural design, poor load-bearing capacity at the connection points of connecting structural components, poor impact resistance and poor stability of the product during use.

[0006] The present invention also provides a process for preparing carbon-glass hybrid composite material skin for aircraft, in order to solve the technical problem of complex processing technology of metal materials for aircraft skin in the prior art.

[0007] The present invention provides a carbon-glass hybrid composite material skin for aircraft, wherein the skin material is a mixture of carbon fiber, glass fiber and resin, and the skin is a one-piece molded structure;

[0008] The mass percentage of the carbon fiber and glass fiber mixture ranges from 60% to 70%, and the mass percentage of the resin ranges from 40% to 30%.

[0009] The ratio of carbon fiber to glass fiber ranges from 1:2 to 1:10.

[0010] The skin has a non-thickened area and a thickened area. The thickness of the thickened area is greater than that of the non-thickened area, and the difference in thickness ranges from 1mm to 10mm.

[0011] Furthermore, the non-thickened area consists of, from the inside out, layers of: carbon fiber layer and glass fiber layer.

[0012] Furthermore, the thickened areas include the thickened area of ​​the projectile mating surface and the thickened area of ​​the flap flange;

[0013] The number of layers in the thickened area of ​​the projectile mating surface, from the inside to the outside, includes: a carbon fiber layer and a glass fiber layer; the glass fiber layer is provided with a glass fiber thickening layer, the glass fiber thickening layer is provided with a glass fiber filling layer inside, and the glass fiber thickening layer is provided with a glass fiber sacrificial layer outside.

[0014] The layers of the thickened area of ​​the cap flange, from the inside out, include: a carbon-glass hybrid layer, a carbon fiber layer, and a glass fiber layer; the ratio of carbon fiber to glass fiber in the carbon-glass hybrid layer is 1:1.

[0015] Furthermore, the wall thickness of the skin ranges from 2mm to 20mm.

[0016] Furthermore, the skin is an irregular curved surface structure with an M-shaped cross-section.

[0017] Furthermore, bolt mounting holes and cover holes are provided on the arc-shaped plate in the middle of the skin;

[0018] Connect the cap to the cap hole;

[0019] The curved plate has side plates connected to both sides via a base plate, and the base plate has clearance holes for eye bolts.

[0020] The present invention provides a process for preparing a carbon-glass hybrid composite material skin for aircraft, comprising the following steps:

[0021] Step 1: Preparing materials

[0022] Prepare carbon fiber prepreg and glass fiber prepreg in a 1:2 ratio;

[0023] Step 2: Apply release agent and base coat to the surface of the mold.

[0024] Clean the surface of the mold, and then apply a release agent to the surface of the mold;

[0025] Dissolve IS1301 resin in acetone, stir thoroughly to form a primer, apply it to the surface of the mold, and then let it dry.

[0026] After drying, use a feeding machine to feed the materials;

[0027] Step 3: Lay a fiber layer on the surface of the skin mold.

[0028] The laying sequence is as follows: prepreg is laid on the surface of the skin mold in the following order: non-thickened area, thickened area of ​​the projectile mating surface, and thickened area of ​​the cap flange.

[0029] First, lay the non-thickened area: with the skin length as the axial direction at 0°, use the 0°, 90°, 45°, -45°, 0° cycle to lay carbon fiber prepreg and glass fiber prepreg on the surface of the skin mold in sequence, forming carbon fiber layer and glass fiber layer on the surface of the skin mold in sequence.

[0030] Secondly, the thickened area of ​​the projectile mating surface is laid: a thickened glass fiber layer is formed inside the glass fiber layer; a glass fiber filling area is formed inside the thickened glass fiber layer, which is filled with glass fiber prepreg to form a glass fiber filling layer; a glass fiber sacrificial area is formed outside the thickened glass fiber layer, which is filled with glass fiber prepreg to form a glass fiber sacrificial layer.

[0031] Next, lay the thickened area of ​​the cap flange: first lay the carbon-glass hybrid layer of the thickened area, a total of 10 layers, in which the odd-numbered layers from 1 to 10 are carbon fiber layers, and the even-numbered layers from 1 to 10 are glass fiber layers, with the ratio of carbon fiber prepreg to glass fiber prepreg being 1:1; then lay the carbon fiber layer and the glass fiber layer on top of the carbon-glass hybrid layer in sequence;

[0032] Finally, attach the cap to the cap hole and secure the side of the cap with multiple pins and threaded sleeves.

[0033] Step 4: Skin curing

[0034] After the prepreg is laid on the surface of the skin mold, the skin is cured using the following dry molding method:

[0035] ① Lay a layer of release cloth on the surface of the prepreg;

[0036] ②Wrap filter paper around the surface of the material sheet and secure it with tape;

[0037] ③ Cover the surface of the sheet with a non-porous release film and secure it with tape;

[0038] ④ Cover the surface of the material sheet with breathable felt and secure it with tape;

[0039] ⑤ Cover the surface of the material with a vacuum bag film, and then perform vacuuming at room temperature;

[0040] ⑥ Transfer the mold to the autoclave and connect a thermocouple to the surface of the mold;

[0041] The molds were then heated and pressurized for curing.

[0042] Step 5: Demolding of the skin

[0043] After curing, remove the skin blank and remove the vacuum film, sealing strip, and auxiliary materials on the mold surface;

[0044] Step Six: Processing with a Molding Skin Machine

[0045] First, process the outer contour of the carbon-glass composite skin of the aircraft;

[0046] Secondly, process the operating hole for the cover;

[0047] Finally, the skin and frame are assembled, and the assembled skin and frame are placed together on a machine tool. The thickened area of ​​the projectile mating surface is CNC machined to make the thickened area and the non-thickened area of ​​the projectile mating surface smoothly transition; thus, the product processing is completed.

[0048] This invention also provides a process for preparing a carbon-glass hybrid composite skin for aircraft, comprising the following steps:

[0049] Step 1: Preparing materials

[0050] Carbon fiber and glass fiber are prepared in a 1:2 ratio;

[0051] Step 2: Apply release agent and base coat to the surface of the mold.

[0052] Clean the surface of the mold, and then apply a release agent to the surface of the mold;

[0053] Dissolve IS1301 resin in acetone, stir thoroughly to form a primer, apply it to the surface of the mold, and then let it dry.

[0054] After drying, use a feeding machine to feed the materials;

[0055] Step 3: Lay a fiber layer on the surface of the skin mold.

[0056] The laying sequence is as follows: dry fibers are laid on the surface of the skin mold in the following order: non-thickened area, thickened area of ​​the projectile mating surface, and thickened area of ​​the cap flange.

[0057] First, lay the non-thickened area: with the skin length as the axial direction at 0°, use the cyclic sequence of 0°, 90°, 45°, -45°, 0° to lay carbon fiber and glass fiber on the surface of the skin mold in sequence, forming carbon fiber layer and glass fiber layer on the surface of the skin mold in sequence.

[0058] Secondly, the thickened area of ​​the projectile mating surface is laid: a glass fiber thickened layer is formed inside the glass fiber layer; a glass fiber filling area is formed inside the glass fiber thickened layer, and glass fiber is used to fill it to form a glass fiber filling layer; a glass fiber sacrificial area is formed outside the glass fiber thickened layer, and glass fiber is used to fill it to form a glass fiber sacrificial layer.

[0059] Next, lay the thickened area of ​​the cap flange: first lay the carbon-glass hybrid layer of the thickened area, a total of 10 layers, in which the odd-numbered layers from 1 to 10 are carbon fiber layers, and the even-numbered layers from 1 to 10 are glass fiber layers, with a carbon fiber to glass fiber ratio of 1:1; then lay the carbon fiber layer and the glass fiber layer on top of the carbon-glass hybrid layer in sequence.

[0060] Finally, attach the cap to the cap hole and secure the side of the cap with multiple pins and threaded sleeves.

[0061] Step 4: Skin curing

[0062] After the dry fibers on the surface of the skin mold are laid, the skin is cured using the following wet molding method:

[0063] ① After the layers are laid, a release cloth is laid on the surface of the sheet, followed by the perforated membrane, the flow guide, the glue injection line, and the vacuum line. The vacuum bag membrane is sealed with a sealing strip, and then connected to the resin collector for vacuuming.

[0064] ② Weigh out the IS1101-A resin, first add 0.8% of the IS1101-B accelerator and stir, then add 1.5% of the IS1101-C curing agent and stir.

[0065] ③ Let the material sheet stand to defoam, and then apply glue;

[0066] ④ The sheet material is cured at room temperature;

[0067] Step 5: Demolding of the skin

[0068] After curing, remove the skin blank and remove the vacuum film, sealing strip, and auxiliary materials on the mold surface;

[0069] Step Six: Processing with a Molding Skin Machine

[0070] First, process the outer contour of the carbon-glass composite skin of the aircraft;

[0071] Secondly, process the operating hole for the cover;

[0072] Finally, the skin and frame are assembled, and the assembled skin and frame are placed together on a machine tool. The thickened area of ​​the projectile mating surface is CNC machined to make the thickened area and the non-thickened area of ​​the projectile mating surface smoothly transition; thus, the product processing is completed.

[0073] The present invention provides a carbon-glass hybrid composite material skin for aircraft, which has the following technical effects:

[0074] The skin material of this invention uses a mixture of carbon fiber, glass fiber, and resin, resulting in a lightweight overall structure. The skin employs a one-piece molding structure, simplifying the design and reducing the number of connecting components. The mass percentage of the carbon fiber and glass fiber mixture ranges from 60% to 70%, while the mass percentage of resin ranges from 40% to 30%. The high proportion of carbon fiber and glass fiber, with a carbon fiber to glass fiber ratio ranging from 1:2 to 1:10, enhances the overall structural strength of the skin. The skin is divided into non-thickened and thickened areas, with the thickened areas being thicker than the non-thickened areas, and this thickness difference ranging from 1mm to 10mm. This thickened area further strengthens the structural strength at stress points. In use, the product exhibits strong impact resistance and high stability.

[0075] This invention provides a manufacturing process for a carbon-glass hybrid composite material skin for aircraft, comprising the following steps: material preparation, applying a release agent and primer to the surface of the mold, laying a fiber layer on the surface of the skin mold, skin curing, skin demolding treatment, and machining of the formed skin. The laying sequence on the surface of the skin mold is: non-thickened area, thickened area of ​​the projectile mating surface, and thickened area of ​​the flap flange. The manufacturing process of the skin is simple, the skin is integrally molded, saving materials and reducing the overall weight. Attached Figure Description

[0076] To more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the specific embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0077] Figure 1 A schematic diagram of the thickened area of ​​the projectile mating surface of a carbon-glass hybrid composite material skin for an aircraft, provided in an embodiment of the present invention;

[0078] Figure 2 A schematic diagram of the thickened flange area of ​​the hatch of a carbon-glass hybrid composite material skin for an aircraft, provided in an embodiment of the present invention;

[0079] Figure 3 A schematic diagram of the inner side of a carbon-glass hybrid composite material skin for an aircraft, provided as an embodiment of the present invention;

[0080] Figure 4 A schematic diagram of the outer surface of a carbon-glass hybrid composite material skin for an aircraft, provided by an embodiment of the present invention;

[0081] Figure 5This is a process flow diagram for preparing a carbon-glass hybrid composite material skin for an aircraft, provided as an embodiment of the present invention.

[0082] Icons: 101-Carbon fiber layer; 102-Glass fiber layer; 103-Glass fiber thickened layer; 104-Glass fiber filler layer; 105-Glass fiber sacrificial layer; 106-Carbon-glass hybrid layer; 107-Curved plate; 108-Bolt mounting hole; 109-Cap hole; 110-Cap; 111-Base plate; 112-Side plate; 113-Eyelash screw clearance hole. Detailed Implementation

[0083] The technical solution of the present invention will now be clearly and completely described with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0084] In the description of this invention, it should be noted that the terms "center," "upper," "lower," "left," "right," "vertical," "horizontal," "inner," and "outer," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing the invention and for 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 the invention. Furthermore, the terms "first," "second," and "third" are used for descriptive purposes only and should not be construed as indicating or implying relative importance.

[0085] 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 can understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0086] like Figures 1-4 As shown, the present invention provides a carbon-glass hybrid composite material skin for aircraft, the material of which is a mixture of carbon fiber, glass fiber and resin, and the skin is an integrally molded structure;

[0087] The mass percentage of the carbon fiber and glass fiber mixture ranges from 60% to 70%, and the mass percentage of the resin ranges from 40% to 30%.

[0088] The ratio of carbon fiber to glass fiber ranges from 1:2 to 1:10.

[0089] The skin has a non-thickened area and a thickened area. The thickness of the thickened area is greater than that of the non-thickened area, and the difference in thickness ranges from 1mm to 10mm.

[0090] In this embodiment, the skin is mainly applied to the leeward side of the aircraft. The skin is made of a carbon-glass hybrid composite material and has a one-piece molded structural design. The use of carbon fiber improves the structural strength of the skin, reduces weight, and lowers costs.

[0091] The thickened skin area is designed to ensure a tighter fit with the sides of the projectile, resulting in higher precision and better structural strength.

[0092] Furthermore, the non-thickened area consists of the following layers from the inside out: carbon fiber layer 101 and glass fiber layer 102.

[0093] In this embodiment, the non-thickened area uses a composite material of carbon fiber and glass fiber, and the carbon fiber improves the overall structural strength.

[0094] Furthermore, the thickened areas include the thickened area of ​​the projectile mating surface and the thickened area of ​​the flap flange;

[0095] The thickened area of ​​the projectile mating surface comprises, from the inside out, a carbon fiber layer 101 and a glass fiber layer 102; the glass fiber layer 102 is provided with a glass fiber thickening layer 103, the glass fiber thickening layer 103 has a glass fiber filling layer 104 inside, and a glass fiber sacrificial layer 105 outside. Figure 1 As shown, the thickened area of ​​the projectile mating surface is laid on top of the non-thickened area to improve the structural strength of the thickened area. The outermost glass fiber sacrificial layer 105 is provided to facilitate the processing of excess material.

[0096] The layers of the thickened flange area of ​​the cap, from the inside out, include: a carbon-glass hybrid layer 106, a carbon fiber layer 101, and a glass fiber layer 102; the carbon fiber to glass fiber ratio in the carbon-glass hybrid layer 106 is 1:1. Figure 2 As shown, the thickened area of ​​the cap flange is laid on top of the non-thickened area to improve the structural strength of the thickened area of ​​the cap flange. A carbon-glass hybrid layer 106 is provided at the bottom of the carbon fiber layer 101 to improve the structural strength of the cap.

[0097] Furthermore, the wall thickness of the skin ranges from 2mm to 20mm.

[0098] In this embodiment, when the skin is applied to the fuselage, since the length of the fuselage is between 2m and 8m and the width is between 0.5m and 2m, the thickness of the skin wall is between 3mm and 20mm.

[0099] When skin is applied to the wing, since the length and width of the wing are between 0.3m and 5m, the skin thickness is between 2mm and 15mm.

[0100] Generally, this type of skin is applied to the windward side of the fuselage.

[0101] Furthermore, the skin is an irregular curved surface structure with an M-shaped cross-section.

[0102] In existing technology, the skin structure is a split structure, namely: a central arc-shaped plate, with a side plate connected to each of the two sides of the arc-shaped plate, and the arc-shaped plate and the side plates are fixed together by connecting structural members. Furthermore, in actual manufacturing, due to the considerable length of the arc-shaped plate and side plates, they are often connected in segments, with each segment connected by connecting structural members. After being assembled into the arc-shaped plate and side plates, the arc-shaped plate and the side plates are then fixed together by connecting structural members. This not only reduces the overall structural strength of the skin and increases its overall weight, but also reduces work efficiency.

[0103] In this embodiment, the skin adopts an M-shaped cross-section and is integrally molded, reducing connecting structural components, resulting in good structural strength of the entire skin, saving raw materials, reducing overall weight, and improving work efficiency. Furthermore, this application connects a base plate 111 to each side of the arc-shaped plate 107, and the base plate 111 improves the support force at the bottom of the arc-shaped plate 107.

[0104] like Figure 3 , Figure 4 As shown, further, the arc-shaped plate 107 in the middle of the skin is provided with bolt mounting holes 108 and cover holes 109.

[0105] Connect the cap 110 at cap hole 109;

[0106] The two sides of the arc plate 107 are respectively connected to the side plate 112 through the base plate 111, and the base plate 111 is provided with eye bolt clearance holes 113.

[0107] Bolt mounting holes 108 are used to connect the projectile. The structure with cap holes 109 is to facilitate the operator's hand reaching inside the skin to install parts. The structure with caps 110 is to seal cap holes 109 after operation, preventing foreign objects from entering the skin. A base plate 111 is connected to each side of the curved plate 107 to increase the support strength on both sides of the curved plate 107. Eye bolt clearance holes 113 are provided on the base plates 111 to protect the sides of the skin from impacts.

[0108] like Figure 5 As shown, the present invention provides a process for preparing a carbon-glass hybrid composite material skin for aircraft, comprising the following steps:

[0109] Step 1: Preparing materials

[0110] Prepare carbon fiber prepreg and glass fiber prepreg in a 1:2 ratio;

[0111] Step 2: Apply release agent and base coat to the surface of the mold.

[0112] Clean the surface of the mold, and then apply a release agent to the surface of the mold;

[0113] Dissolve IS1301 resin in acetone, stir thoroughly to form a primer, apply it to the surface of the mold, and then let it dry.

[0114] After drying, use a feeding machine to feed the materials;

[0115] Step 3: Lay a fiber layer on the surface of the skin mold.

[0116] The laying sequence is as follows: prepreg is laid on the surface of the skin mold in the following order: non-thickened area, thickened area of ​​the projectile mating surface, and thickened area of ​​the cap flange.

[0117] First, lay the non-thickened area: with the skin length as the axial direction at 0°, in the cyclical sequence of 0°, 90°, 45°, -45°, 0°, carbon fiber prepreg and glass fiber prepreg are laid on the surface of the skin mold in sequence, forming carbon fiber layer 101 and glass fiber layer 102 on the surface of the skin mold in sequence.

[0118] Secondly, the thickened area of ​​the projectile mating surface is laid: a glass fiber thickened layer 103 is formed inside the glass fiber layer 102; a glass fiber filling area is formed inside the glass fiber thickened layer 103, which is filled with glass fiber prepreg to form a glass fiber filling layer 104; a glass fiber sacrificial area is formed outside the glass fiber thickened layer 103, which is filled with glass fiber prepreg to form a glass fiber sacrificial layer 105.

[0119] Next, lay the thickened area of ​​the cap flange: First, lay the carbon-glass hybrid layer 106 of the thickened area, a total of 10 layers, in which the layup pattern is as follows: the odd-numbered layers from layer 1 to layer 10, that is, layers 1, 3, 5, 7, and 9 are carbon fiber layers 101, and the even-numbered layers from layer 1 to layer 10, that is, layers 2, 4, 6, 8, and 10 are glass fiber layers 102. The ratio of carbon fiber prepreg to glass fiber prepreg is 1:1; then lay the carbon fiber layer 101 and the glass fiber layer 102 on the top of the carbon-glass hybrid layer 106 in sequence.

[0120] Finally, the cover 110 is connected to the cover hole 109, and the side of the cover 110 is fixed by multiple pins and screws.

[0121] Step 4: Skin curing

[0122] After the prepreg is laid on the surface of the skin mold, the skin is cured using the following dry molding method:

[0123] ① Lay a layer of release cloth on the surface of the prepreg;

[0124] ②Wrap filter paper around the surface of the material sheet and secure it with tape;

[0125] ③ Cover the surface of the sheet with a non-porous release film and secure it with tape;

[0126] ④ Cover the surface of the material sheet with breathable felt and secure it with tape;

[0127] ⑤ Cover the surface of the material with a vacuum bag film, and then perform vacuuming at room temperature, requiring a vacuum degree ≤ -0.095MPa;

[0128] ⑥ Transfer the mold to the autoclave and connect a thermocouple to the surface of the mold;

[0129] The mold was subjected to heating and pressure curing at 90℃ for 3 hours.

[0130] Step 5: Demolding of the skin

[0131] After curing, remove the skin blank and remove the vacuum film, sealing strip, and auxiliary materials on the mold surface;

[0132] Step Six: Processing with a Molding Skin Machine

[0133] First, process the outer contour of the carbon-glass composite skin of the aircraft;

[0134] Secondly, process the operating hole for the cover;

[0135] Finally, the skin and frame are assembled, and the assembled skin and frame are placed together on a machine tool. The thickened area of ​​the projectile mating surface is CNC machined to make the thickened area and the non-thickened area of ​​the projectile mating surface smoothly transition; thus, the product processing is completed.

[0136] This invention also provides a process for preparing a carbon-glass hybrid composite skin for aircraft, comprising the following steps:

[0137] Step 1: Preparing materials

[0138] Carbon fiber and glass fiber are prepared in a 1:2 ratio;

[0139] Step 2: Apply release agent and base coat to the surface of the mold.

[0140] Clean the surface of the mold, and then apply a release agent to the surface of the mold;

[0141] Dissolve IS1301 resin in acetone, stir thoroughly to form a primer, apply it to the surface of the mold, and then let it dry.

[0142] After drying, use a feeding machine to feed the materials;

[0143] Step 3: Lay a fiber layer on the surface of the skin mold.

[0144] The laying sequence is as follows: dry fibers are laid on the surface of the skin mold in the following order: non-thickened area, thickened area of ​​the projectile mating surface, and thickened area of ​​the cap flange.

[0145] First, lay the non-thickened area: with the skin length as the axial direction at 0°, use the cyclic sequence of 0°, 90°, 45°, -45°, 0° to lay carbon fiber and glass fiber on the surface of the skin mold in sequence, and form carbon fiber layer 101 and glass fiber layer 102 on the surface of the skin mold in sequence.

[0146] Secondly, the thickened area of ​​the projectile mating surface is laid: a glass fiber thickened layer 103 is formed inside the glass fiber layer 102; a glass fiber filling area is formed inside the glass fiber thickened layer 103, and glass fiber is used to fill it to form a glass fiber filling layer 104; a glass fiber sacrificial area is formed outside the glass fiber thickened layer 103, and glass fiber is used to fill it to form a glass fiber sacrificial layer 105.

[0147] Next, lay the thickened area of ​​the cap flange: First, lay the carbon-glass hybrid layer 106 of the thickened area, a total of 10 layers. The layering method is as follows: the odd-numbered layers from layer 1 to layer 10, that is, layers 1, 3, 5, 7, and 9 are carbon fiber layers 101, and the even-numbered layers from layer 1 to layer 10, that is, layers 2, 4, 6, 8, and 10 are glass fiber layers 102. The ratio of carbon fiber to glass fiber is 1:1. Then, lay the carbon fiber layer 101 and the glass fiber layer 102 on the top of the carbon-glass hybrid layer 106 in sequence.

[0148] Finally, the cover 110 is connected to the cover hole 109, and the side of the cover 110 is fixed by multiple pins and screws.

[0149] Step 4: Skin curing

[0150] After the dry fibers are laid on the surface of the skin mold, the skin is cured using the following wet molding method:

[0151] ① After the layers are laid, a release cloth is laid on the surface of the sheet, followed by the perforated membrane, the flow guide, the glue injection line, and the vacuum line. The vacuum bag membrane is sealed with a sealing strip, and then connected to the resin collector for vacuuming.

[0152] ② Weigh out the IS1101-A resin, first add 0.8% of the IS1101-B accelerator and stir, then add 1.5% of the IS1101-C curing agent and stir.

[0153] ③ Let the material sheet stand for 10 minutes to defoam, and then pour the glue;

[0154] ④ The sheet material is cured at room temperature for at least six hours;

[0155] Step 5: Demolding of the skin

[0156] After curing, remove the skin blank and remove the vacuum film, sealing strip, and auxiliary materials on the mold surface;

[0157] Step Six: Processing with a Molding Skin Machine

[0158] First, process the outer contour of the carbon-glass composite skin of the aircraft;

[0159] Secondly, process the operating hole for the cover;

[0160] Finally, the skin and frame are assembled, and the assembled skin and frame are placed together on a machine tool. The thickened area of ​​the projectile mating surface is CNC machined to make the thickened area and the non-thickened area of ​​the projectile mating surface smoothly transition; thus, the product processing is completed.

[0161] The present invention discloses a process for preparing a carbon-glass hybrid composite material skin for an aircraft, comprising the following steps: material preparation, applying a release agent and a primer to the surface of the mold, laying a fiber layer on the surface of the skin mold, skin curing, skin demolding treatment, and machining of the formed skin. During skin curing, the above-mentioned dry or wet molding method is used, the skin molding process is simple, and the product has high structural strength after molding.

[0162] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some or all of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the embodiments of the present invention.

Claims

1. A carbon-glass hybrid composite skin for aircraft, characterized in that, The material of the skin is a mixture of carbon fiber, glass fiber and resin, and the skin is a one-piece molded structure; The mass percentage of the mixture of carbon fiber and glass fiber ranges from 60% to 70%, and the mass percentage of the resin ranges from 40% to 30%. The ratio of the carbon fiber and the glass fiber is between 1:2 and 1:

10. The skin has a non-thickened area and a thickened area, the thickness of the thickened area is greater than the thickness of the non-thickened area, and the difference in thickness ranges from 1mm to 10mm. The thickened area includes the thickened area of ​​the projectile mating surface and the thickened area of ​​the flap flange; The number of layers in the thickened area of ​​the projectile mating surface, from the inside to the outside, includes: a carbon fiber layer (101) and a glass fiber layer (102); a glass fiber thickening layer (103) is provided inside the glass fiber layer (102), a glass fiber filling layer (104) is provided inside the glass fiber thickening layer (103), and a glass fiber sacrificial layer (105) is provided outside the glass fiber thickening layer (103). The number of layers in the thickened flap area from the inside out includes: a carbon-glass hybrid layer (106), a carbon fiber layer (101), and a glass fiber layer (102); the carbon fiber to glass fiber ratio in the carbon-glass hybrid layer (106) is 1:

1.

2. The carbon-glass hybrid composite skin for an aircraft according to claim 1, characterized in that, The non-thickened area consists of the following layers from the inside out: carbon fiber layer (101) and glass fiber layer (102).

3. The carbon-glass hybrid composite skin for an aircraft according to claim 1, characterized in that, The wall thickness of the skin ranges from 2mm to 20mm.

4. The carbon-glass hybrid composite skin for an aircraft according to claim 1, characterized in that, The skin is an irregular curved surface structure with an M-shaped cross-section.

5. A carbon-glass hybrid composite skin for an aircraft according to claim 4, characterized in that, The arc-shaped plate (107) in the middle of the skin is provided with bolt mounting holes (108) and cover holes (109). The cap (110) is connected at the cap hole (109); The two sides of the arc plate (107) are respectively connected to the side plate (112) through the base plate (111), and the base plate (111) is provided with eye bolt clearance holes (113).

6. A process for preparing a carbon-glass hybrid composite skin for an aircraft, characterized in that, Includes the following steps: Step 1: Preparing materials Prepare carbon fiber prepreg and glass fiber prepreg in a 1:2 ratio; Step 2: Apply release agent and base coat to the surface of the mold. Clean the surface of the mold, and then apply a release agent to the surface of the mold; Dissolve the resin in acetone, stir thoroughly to form a primer, apply it to the surface of the mold, and then let it dry. After drying, use a feeding machine to feed the materials; Step 3: Lay a fiber layer on the surface of the skin mold. The laying order is as follows: non-thickened area, thickened area of ​​projectile mating surface, thickened area of ​​cap flange. Prepreg is laid on the surface of the skin mold in the following order: First, lay the non-thickened area: with the skin length as the axial direction at 0°, in the cyclical order of 0°, 90°, 45°, -45°, 0°, carbon fiber prepreg and glass fiber prepreg are laid on the surface of the skin mold in sequence, forming a carbon fiber layer (101) and a glass fiber layer (102) on the surface of the skin mold in sequence. Next, the thickened area of ​​the projectile mating surface is laid: a glass fiber thickened layer (103) is formed inside the glass fiber layer (102); a glass fiber filling area is formed inside the glass fiber thickened layer (103), which is filled with glass fiber prepreg to form a glass fiber filling layer (104); a glass fiber sacrificial area is formed outside the glass fiber thickened layer (103), which is filled with glass fiber prepreg to form a glass fiber sacrificial layer (105). Next, lay the thickened area of ​​the cap flange: first lay the carbon-glass hybrid layer (106) of the thickened area, a total of 10 layers, in which the odd-numbered layers from 1 to 10 are carbon fiber layers (101), and the even-numbered layers from 1 to 10 are glass fiber layers (102), with the ratio of carbon fiber prepreg to glass fiber prepreg being 1:1; then lay the carbon fiber layer (101) and the glass fiber layer (102) on the top of the carbon-glass hybrid layer (106) in sequence. Finally, the cap (110) is connected to the cap hole, and the side of the cap (110) is fixed by multiple pins and threaded sleeves. Step 4: Skin curing After the prepreg is laid on the surface of the skin mold to form a sheet, the skin is cured using the following dry molding method: ① Lay a layer of release cloth on the surface of the sheet material; ②Wrap filter paper around the surface of the material sheet and secure it with tape; ③ Cover the surface of the sheet with a non-porous release film and secure it with tape; ④ Cover the surface of the material sheet with breathable felt and secure it with tape; ⑤ Cover the surface of the material with a vacuum bag film, and then perform vacuuming at room temperature; ⑥ Transfer the mold to the autoclave and connect a thermocouple to the surface of the mold; The molds were then heated and pressurized for curing. Step 5: Demolding of the skin After curing, remove the skin blank and remove the vacuum bag film, tape, and auxiliary materials from the mold surface; Step Six: Processing with a Molding Skin Machine First, process the outer contour of the carbon-glass composite skin of the aircraft; Next, process the cap hole (109); Finally, the skin and frame are assembled, and the assembled skin and frame are placed together on a machine tool. The thickened area of ​​the projectile mating surface is CNC machined to make the thickened area and the non-thickened area of ​​the projectile mating surface smoothly transition; thus, the product processing is completed.

7. A process for preparing a carbon-glass hybrid composite skin for an aircraft, characterized in that, Includes the following steps: Step 1: Preparing materials Carbon fiber and glass fiber are prepared in a 1:2 ratio; Step 2: Apply release agent and base coat to the surface of the mold. Clean the surface of the mold, and then apply a release agent to the surface of the mold; Dissolve the resin in acetone, stir thoroughly to form a primer, apply it to the surface of the mold, and then let it dry. After drying, use a feeding machine to feed the materials; Step 3: Lay a fiber layer on the surface of the skin mold. The laying order is as follows: non-thickened area, thickened area of ​​projectile mating surface, thickened area of ​​cap flange. Dry fibers are laid on the surface of the skin mold, using the following laying sequence: First, lay the non-thickened area: with the skin length as the axial direction at 0°, carbon fiber and glass fiber are laid sequentially on the surface of the skin mold in the order of 0°, 90°, 45°, -45° and 0°, forming a carbon fiber layer (101) and a glass fiber layer (102) on the surface of the skin mold. Secondly, the thickened area of ​​the projectile mating surface is laid: a glass fiber thickened layer (103) is formed inside the glass fiber layer (102); a glass fiber filling area is formed inside the glass fiber thickened layer (103), and glass fiber is used to fill it to form a glass fiber filling layer (104); a glass fiber sacrificial area is formed outside the glass fiber thickened layer (103), and glass fiber is used to fill it to form a glass fiber sacrificial layer (105). Next, lay the thickened area of ​​the cap flange: first lay the carbon-glass hybrid layer (106) of the thickened area, a total of 10 layers, in which the odd-numbered layers from 1 to 10 are carbon fiber layers (101), and the even-numbered layers from 1 to 10 are glass fiber layers (102), with the ratio of carbon fiber to glass fiber being 1:1; then lay the carbon fiber layer (101) and the glass fiber layer (102) on the top of the carbon-glass hybrid layer (106) in sequence. Finally, the cap (110) is connected to the cap hole, and the side of the cap (110) is fixed by multiple pins and threaded sleeves. Step 4: Skin curing After the dry fibers are laid on the surface of the skin mold to form a sheet, the skin is cured using the following wet molding method: ① After the layers are laid, a release cloth is laid on the surface of the sheet, followed by the perforated membrane, the flow guide, the glue injection line, and the vacuum line. The vacuum bag membrane is sealed with a sealing strip, and then connected to the resin collector for vacuuming. ② Weigh the resin, first add 0.8% accelerator and stir, then add 1.5% curing agent and stir; ③ Let the material sheet stand to defoam, and then apply glue; ④ The sheet material is cured at room temperature; Step 5: Demolding of the skin After curing, remove the skin blank and remove the vacuum bag film, sealing strip, and auxiliary materials on the mold surface; Step Six: Processing with a Molding Skin Machine First, process the outer contour of the carbon-glass composite skin of the aircraft; Next, process the cap hole (109); Finally, the skin and frame are assembled, and the assembled skin and frame are placed together on a machine tool. The thickened area of ​​the projectile mating surface is CNC machined to make the thickened area and the non-thickened area of ​​the projectile mating surface smoothly transition; thus, the product processing is completed.

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