Co-curing and co-bonding manufacturing method of aircraft composite control surface

Abstract

The invention discloses a co-curing and co-gluing manufacturing method of an aircraft composite control surface, which comprises the following steps: firstly, manufacturing a composite frame assembly by co-curing, and then combining and molding a skin and a honeycomb core on the composite frame assembly by co-gluing instead of the traditional secondary gluing molding; when the composite material frame assembly is manufactured in a co-curing mode, the skin fake piece is used as filler, and the manufacturing cost is saved; the forming mode of co-curing and co-bonding is beneficial to product structure stability and strength improvement, when the composite material frame assembly is co-cured, a plurality of ribs and beams are accurately positioned, and overall demolding is facilitated after co-curing is completed by means of a skin fake piece.

Description

Technical Field

[0001] The invention relates to a method for manufacturing a composite aircraft rudder surface, in particular to a manufacturing method in the form of co-curing and co-bonding molding. Background Art

[0002] Some supersonic drones use honeycomb sandwich composite structures, and set up multiple ribs and main beams to adopt co-curing connection method to improve structural strength.

[0003] Generally, for the composite control surface of an aircraft, one structure is as follows: Figure 1 As shown in the figure, the composite frame assembly is composed of beams and upper ribs, middle ribs, lower ribs, and twist strip filling blocks, and the composite rudder assembly is composed of skin, honeycomb core, composite frame assembly, and structural adhesive film layer, which are formed by co-bonding and secondary bonding. However, the composite rudder used for this supersonic UAV must be designed to meet the large bending, shearing, and torsion forces on the ribs and beams during supersonic flight. The rigidity and strength of the beam-rib connection in the general co-bonding or secondary bonding molding form cannot meet such load requirements. In addition, the temperature resistance requirements of the aircraft surface are quite high during supersonic flight, and there are the following shortcomings:

[0004] 1. Since the aircraft composite rudder surface adopts a supercritical wing shape and a hyperbolic structure, the angle between the upper and lower rib edges of the box-shaped rib and the middle web is less than 90°, which makes mold forming difficult; parts cannot be demoulded at some closed angle positions, and the mold needs to be designed into a detachable module for combination;

[0005] 2. Due to the small space between ribs, it is also difficult to demould the ribs using a combined mold;

[0006] 3. Because the composite rudder contains multiple ribs and beams, it is difficult to locate their positions. Summary of the invention

[0007] Purpose of the invention: The purpose of the present invention is to provide a method for manufacturing aircraft composite control surfaces, which is formed by co-curing composite frame components and then co-bonding composite skins and honeycomb cores.

[0008] Technical solution: A method for co-curing and co-bonding a composite aircraft rudder surface, comprising the following steps:

[0009] Step 100, co-curing to manufacture a composite frame assembly, includes:

[0010] Step 101, cutting and marking the prepreg materials of the beam, upper rib, middle rib, lower rib, twist strip filling block and skin according to the design size;

[0011] Step 102, respectively completing the laying and compacting of the respective prepregs on the forming molds of the beam, the upper rib, the middle rib, and the lower rib, and completing the twisting of the prepregs of the twist strip filling block into twisted yarns;

[0012] Step 103, moving the laid beams, upper ribs, middle ribs, lower ribs, and twist strip filling blocks into a fixed position in a closed combination mold and pre-compacting them;

[0013] Step 104, pre-compacting the assembled and positioned composite frame assembly preform, trimming the surface of the preform, packaging and bagging the preform, and then sending it into an autoclave for curing and molding;

[0014] Step 105, demoulding and taking out the composite frame assembly;

[0015] Step 200, co-bonding and manufacturing a composite rudder surface assembly, includes:

[0016] Step 201, manufacturing a honeycomb core according to the design size;

[0017] Step 202, completing the laying and compacting of the prepreg on the forming mold of the skin;

[0018] Step 203, moving the laid skin, the manufactured honeycomb core and the composite frame assembly into a closed assembly mold and bonding them using a structural adhesive film layer;

[0019] Step 204, packaging and bagging the assembled and positioned composite rudder surface assembly preforms, and then sending them into an autoclave for curing and molding;

[0020] Step 205, demoulding and taking out the composite rudder surface assembly.

[0021] Furthermore, the material of the prepreg of the beam, the upper rib, the middle rib, the lower rib, the twist strip filling block and the skin is T800 phthalonitrile carbon fiber unidirectional tape.

[0022] Furthermore, the material of the honeycomb core is a glass fiber honeycomb core impregnated with phthalonitrile resin.

[0023] Furthermore, in step 102 and step 202, every 1 to 3 layers of prepreg vacuum bags are pre-compacted during laying.

[0024] Furthermore, in step 102 and step 202, the angle of the plies is 0°, 90°, or ±45°, and the plies are laid in a staggered manner.

[0025] Furthermore, in step 104 and step 204, the air tightness of the packaging bag is tested after packaging.

[0026] Furthermore, in step 104, the beam, the upper rib, the middle rib and the lower rib are raised in a closed combination mold by using a skin dummy.

[0027] Furthermore, in step 105, the demoulding tool is used to assist in demoulding by bolt pulling. After demoulding, the composite frame component is trimmed to remove excess burrs and is wiped with acetone to remove dirt, and then air-dried and sealed for later use.

[0028] Beneficial effects: The advantages of the present invention are: first, the composite frame assembly is manufactured by co-curing, and then the skin and the honeycomb core are bonded together on the composite frame assembly by co-bonding, replacing the traditional secondary bonding molding; when the composite frame assembly is co-cured and manufactured, the skin dummy is used as a filler to save manufacturing costs; the co-curing and then co-bonding molding method is beneficial to the stability of the product structure and the improvement of strength. When the composite frame assembly is co-cured, multiple ribs and beams are accurately positioned, and the overall demoulding is facilitated after the co-curing is completed with the help of the skin dummy. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] Figure 1 This is a schematic diagram of the structure of the composite rudder surface assembly. DETAILED DESCRIPTION

[0030] The present invention is further explained below in conjunction with the accompanying drawings and specific embodiments.

[0031] An aircraft composite rudder surface, the structure of which is as shown in the attached Figure 1 As shown, a composite frame assembly is formed by a beam 1 and an upper rib 2, a middle rib 3 (two middle ribs in the figure are respectively denoted as middle rib 31 and middle rib 32), a lower rib 4, and a twist strip filling block 5; a composite rudder surface assembly is formed by a skin 6, a honeycomb core 7 (three honeycomb cores in the figure are respectively denoted as honeycomb core 71, honeycomb core 72, and honeycomb core 73), a composite frame assembly, and a structural adhesive film layer; a composite rudder surface is manufactured by a co-curing and then co-bonding manufacturing method, and the manufacturing process includes: step 100, co-curing to manufacture a composite frame assembly; step 200, co-bonding to manufacture a composite rudder surface assembly.

[0032] Step 100 , co-curing to manufacture a composite frame assembly, specifically includes the following steps 101 to 105 .

[0033] Step 101, according to the design size, cut and mark the prepreg layers for making beams, upper ribs, middle ribs, lower ribs, twist strip filling blocks, and skins; the material of the prepreg is T800 phthalonitrile carbon fiber unidirectional tape.

[0034] Step 102, complete the laying and compaction of each layer of prepreg on the beam forming mold, the laying information is: (45 / -45 / 0 / 90 / 45 / -45 / 0 / 90)s, when laying, every 1 to 3 layers of prepreg vacuum bag are pre-compacted, the laying angle is 0° or 90° or ±45°, and the prepregs are laid in a staggered manner, and the prepreg splicing positions are staggered.

[0035] The production of upper ribs, middle ribs and lower ribs refers to the production of beams and is completed on their respective forming molds.

[0036] The prepreg is rolled into twisted yarns to make twisted strip filling blocks.

[0037] Step 103, move the laid beams, upper ribs, middle ribs, lower ribs, and twist strip filling blocks into the closed combination mold, and use skin dummies to artificially raise the beams, upper ribs, middle ribs, and lower ribs in the mold to achieve the co-cured composite frame assembly The shape is consistent with the design shape, and pre-compacted after fixing the position.

[0038] Step 104, pre-compact the assembled and positioned composite frame assembly preforms, trim the surfaces of the preforms, and package them into bags. After packaging and bagging, test the air tightness of the packaging bags, and then send them into an autoclave for curing and molding.

[0039] Step 105, demoulding and taking out the composite frame assembly, using a demoulding tool to assist in demoulding by bolt pulling, trimming the composite frame assembly after demoulding to remove excess burrs, and wiping with acetone to remove dirt, air-drying, and sealing for later use.

[0040] Step 200, co-bonding to manufacture a composite rudder assembly, specifically includes the following steps 201 to 205.

[0041] Step 201 : manufacturing a honeycomb core according to the design size, wherein the material of the honeycomb core is a glass fiber honeycomb core impregnated with phthalonitrile resin.

[0042] Step 202, complete the laying and compaction of each layer of prepreg on the forming mold of the skin, the laying information is: (45 / -45 / 0 / 90 / 45 / -45 / 0 / 0 / -45 / 45 / 90 / 0 / -45 / 45), when laying, every 1 to 3 layers of prepreg vacuum bag are pre-compacted, the laying angle is 0° or 90° or ±45°, and the prepregs are laid in an alternating manner, and the prepreg splicing positions are staggered.

[0043] Step 203, moving the laid skin, the manufactured honeycomb core and the cleaned and trimmed composite frame assembly into a closed assembly mold, and bonding them together using a high temperature resistant structural adhesive film layer.

[0044] Step 204, the assembled and positioned composite rudder surface assembly preform is packaged and bagged, and after packaging and bagging, the air tightness of the packaging bag is tested, and then the bag is sent to a hot autoclave for curing and molding.

[0045] Step 205, demoulding and taking out the composite rudder surface assembly.

[0046] The present application first manufactures the composite frame assembly by co-curing, and then combines the skin and honeycomb core on the composite frame assembly by co-gluing. When the composite frame assembly is co-cured, the skin dummy is used as a filler, which can save manufacturing costs while ensuring the shape of the product; the co-curing and then co-gluing molding method is conducive to the stability of the product structure and the improvement of strength. When the composite frame assembly is co-cured, multiple ribs and beams are accurately positioned, and the overall demoulding is facilitated after the co-curing is completed with the help of the skin dummy.

Claims

1. A method for co-curing and co-bonding aircraft composite rudder surfaces, characterized in that The following steps are involved: Step 100, co-curing to manufacture a composite frame assembly, includes: Step 101, cutting and marking the prepreg materials of the beam, upper rib, middle rib, lower rib, twist strip filling block and skin according to the design size; Step 102, respectively completing the laying and compacting of the respective prepregs on the forming molds of the beam, the upper rib, the middle rib, and the lower rib, and completing the twisting of the prepregs of the twist strip filling block into twisted yarns; Step 103, moving the laid beams, upper ribs, middle ribs, lower ribs, and twist strip filling blocks into a fixed position in a closed combination mold and pre-compacting them; Step 104, pre-compacting the assembled and positioned composite frame assembly preform, trimming the surface of the preform, packaging and bagging the preform, and then sending it into an autoclave for curing and molding; Step 105, demoulding and taking out the composite frame assembly; Step 200, co-bonding and manufacturing a composite rudder surface assembly, includes: Step 201, manufacturing a honeycomb core according to the design size; Step 202, completing the laying and compacting of the prepreg on the forming mold of the skin; Step 203, moving the laid skin, the manufactured honeycomb core and the composite frame assembly into a closed assembly mold and bonding them using a structural adhesive film layer; Step 204, packaging and bagging the assembled and positioned composite rudder surface assembly preforms, and then sending them into an autoclave for curing and molding; Step 205, demoulding and taking out the composite rudder surface assembly.

2. The method for manufacturing a composite aircraft control surface by co-curing and co-bonding according to claim 1, characterized in that: The material of the prepreg of beam, upper rib, middle rib, lower rib, twist strip filling block and skin is T800 phthalonitrile carbon fiber unidirectional tape.

3. The method for manufacturing a composite aircraft control surface by co-curing and co-bonding according to claim 1, characterized in that: The material of the honeycomb core is a glass fiber honeycomb core impregnated with phthalonitrile resin.

4. The method for manufacturing a composite aircraft control surface by co-curing and co-bonding according to claim 1, characterized in that: In step 102 and step 202, every 1 to 3 layers of prepreg vacuum bags are pre-compacted during laying.

5. The method for manufacturing a composite aircraft control surface by co-curing and co-bonding according to claim 1, characterized in that: In step 102 and step 202, the angle of the plies is 0°, 90°, or ±45°, and the plies are laid in a staggered manner.

6. The method for manufacturing a composite aircraft control surface by co-curing and co-bonding according to claim 1, characterized in that: In step 104 and step 204, the air tightness of the packaging bag is tested after packaging.

7. The method for manufacturing a composite aircraft control surface by co-curing and co-bonding according to claim 1, characterized in that: In step 104, the beam, upper rib, middle rib and lower rib are raised in a closed assembly mold by using a skin dummy.

8. The method for manufacturing a composite aircraft control surface by co-curing and co-bonding according to claim 1, characterized in that: In step 105, the demoulding tool performs bolt pulling to assist demoulding. After demoulding, the composite frame component is trimmed to remove excess burrs and wiped with acetone to remove dirt, and then air-dried and sealed for standby use.

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