CN115214872A - Composite material main bearing beam and integral forming method - Google Patents

Composite material main bearing beam and integral forming method Download PDF

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Publication number
CN115214872A
CN115214872A CN202210691660.3A CN202210691660A CN115214872A CN 115214872 A CN115214872 A CN 115214872A CN 202210691660 A CN202210691660 A CN 202210691660A CN 115214872 A CN115214872 A CN 115214872A
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China
Prior art keywords
wing box
joint
weight
wing
core layer
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Pending
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CN202210691660.3A
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Chinese (zh)
Inventor
韩蕾
王冬
孙辉
王欣怡
龚文化
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Aerospace Research Institute of Materials and Processing Technology
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Aerospace Research Institute of Materials and Processing Technology
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Priority to CN202210691660.3A priority Critical patent/CN115214872A/en
Publication of CN115214872A publication Critical patent/CN115214872A/en
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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C3/00—Wings
    • B64C3/18—Spars; Ribs; Stringers
    • B64C3/182—Stringers, longerons
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
    • B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
    • B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
    • B29C70/06—Fibrous reinforcements only
    • B29C70/10—Fibrous reinforcements only characterised by the structure of fibrous reinforcements, e.g. hollow fibres
    • B29C70/16—Fibrous reinforcements only characterised by the structure of fibrous reinforcements, e.g. hollow fibres using fibres of substantial or continuous length
    • B29C70/22—Fibrous reinforcements only characterised by the structure of fibrous reinforcements, e.g. hollow fibres using fibres of substantial or continuous length oriented in at least two directions forming a two dimensional structure
    • B29C70/228—Fibrous reinforcements only characterised by the structure of fibrous reinforcements, e.g. hollow fibres using fibres of substantial or continuous length oriented in at least two directions forming a two dimensional structure the structure being stacked in parallel layers with fibres of adjacent layers crossing at substantial angles
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
    • B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
    • B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
    • B29C70/28—Shaping operations therefor
    • B29C70/30—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core
    • B29C70/34—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression, i.e. combined with compressing after the lay-up operation
    • B29C70/342—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression, i.e. combined with compressing after the lay-up operation using isostatic pressure
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
    • B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
    • B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
    • B29C70/28—Shaping operations therefor
    • B29C70/40—Shaping or impregnating by compression not applied
    • B29C70/42—Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles
    • B29C70/44—Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles using isostatic pressure, e.g. pressure difference-moulding, vacuum bag-moulding, autoclave-moulding or expanding rubber-moulding
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
    • B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
    • B29C70/68—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts by incorporating or moulding on preformed parts, e.g. inserts or layers, e.g. foam blocks
    • B29C70/84—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts by incorporating or moulding on preformed parts, e.g. inserts or layers, e.g. foam blocks by moulding material on preformed parts to be joined
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
    • B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
    • B29C70/88—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts characterised primarily by possessing specific properties, e.g. electrically conductive or locally reinforced
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/40—Weight reduction

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Composite Materials (AREA)
  • Textile Engineering (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Moulding By Coating Moulds (AREA)

Abstract

The invention relates to a composite material main bearing beam and an integral forming method. The composite material main bearing beam comprises a main joint, a wing box tap joint, a wing box reinforcing sheet, a weight reduction core layer and a skin, wherein the main joint is fixedly connected with the wing box tap joint and the wing box reinforcing sheet, and the weight reduction core layer is fixedly connected with the wing box tap joint; the skin is prepared by coating fiber prepreg outside the main joint, the wing box tap joint, the wing box reinforcing sheet and the weight-reducing core layer according to the loading condition of the airfoil and adopting an integral compression molding process, wherein the main joint, the wing box tap joint and the wing box reinforcing sheet are metal pieces. According to the invention, the carbon fiber can be filled in the metal joint, and the filled carbon fiber and the outer skin are combined to form a plurality of T-shaped beam structures, so that the rigidity and the strength of the main bearing beam are effectively improved; the invention positions and combines the plurality of weight-reducing core layers and the rectangular beam consisting of the carbon fibers with metal parts in various forms such as butt joint, cementing, screw joint and the like, and integrally solidifies and forms, thereby further improving the reliability of the structure.

Description

Composite material main bearing beam and integral forming method
Technical Field
The invention relates to a composite material main bearing beam and an integral forming method, and belongs to the technical field of composite materials.
Background
Aircraft weight is closely related to performance and economy, and reducing aircraft structural weight is one of the main goals in development work. The composite material airfoil has excellent mechanical properties such as high specific stiffness and specific strength, good fatigue resistance and corrosion resistance, and the like, is widely applied to aviation and aerospace structures, and the application range of the composite material airfoil is expanded from an initial secondary load-bearing structure to a current main load-bearing structure.
The main bearing wing surface mainly has the function of generating lift force, the main bearing beam in the wing surface plays a decisive role in the rigidity strength of the wing surface, the quality of the main bearing beam directly influences the stability and the reliability of the flying process of an aircraft, and how to prepare the composite material main bearing beam member with a reliable structure becomes a hotspot direction of research.
Disclosure of Invention
The invention aims to overcome the defects of the prior art and provides a light-weight high-strength composite material main carrier beam and an integral forming method of the carrier beam.
The technical scheme adopted by the invention is as follows:
in a first aspect, the invention provides a composite material main carrier beam, which comprises a main joint, a wing box tap joint, a wing box reinforcing sheet, a weight-reducing core layer and a skin, wherein the main joint is fixedly connected with the wing box tap joint and the wing box reinforcing sheet, and the weight-reducing core layer is fixedly connected with the wing box tap joint; the main joint, the wing box tap joint, the wing box reinforcing sheet and the weight-reducing core layer are coated on the outer portion of the skin.
Preferably, the skin is formed by coating fiber prepregs outside the main joint, the wing box taps, the wing box reinforcing sheets and the weight-reducing core layer according to the loading condition of the airfoil and integrally molding and curing.
Preferably, the fiber prepreg is a carbon fiber prepreg comprising carbon fibers and resin, the carbon fibers are unidirectional T300, T700, T800 or T1000 carbon fibers, and the resin is a medium-temperature or high-temperature curing epoxy resin. The fibers may also be other suitable fibers such as glass fibers or quartz fibers.
Preferably, the main joint, the wing box taps, the wing box reinforcing sheet are metal pieces; the weight-reducing core layer is made of rigid polyurethane foam, phenolic resin foam or PMI foam.
Preferably, the wing box tap is mounted in a corresponding slot of the main joint, the wing box reinforcing plate is mounted in the wing box tap and a slot reserved in the weight-reducing core layer, and one end of the weight-reducing core layer is mounted in a slot reserved in the wing box tap.
In a second aspect, the present invention provides a method for integrally forming a main load-bearing beam made of a composite material, comprising the following steps:
assembling the wing box tap with a weight-reducing core layer, wherein one end of the weight-reducing core layer is arranged in a groove reserved in the wing box tap;
assembling the wing box reinforcing sheet with a wing box tap and a weight-reducing core layer, wherein the wing box reinforcing sheet is arranged in a groove reserved in the wing box tap and the weight-reducing core layer;
coating a fiber prepreg on an assembly of the wing box tap, the wing box reinforcing sheet and the weight-reducing core layer, and performing pre-curing treatment;
assembling a pre-cured wing box tap joint, a wing box reinforcing sheet and a weight-reducing core layer combined piece with a main joint;
and for the assembly of the main joint, the wing box tap joint, the wing box reinforcing sheet and the weight-reducing core layer, fiber prepregs with different thicknesses are coated at different positions according to the loading condition of the airfoil, and the whole body is molded, cured and molded.
Preferably, the main joint, the wing box taps and the wing box stiffener are grit blasted prior to assembly of the main joint, the wing box taps and the wing box stiffener; the sand blasting treatment adopts 20-40 mesh quartz sand, and the air pressure is 0.4-0.7 MPa.
Preferably, before the weight-reducing core layer is assembled with the wing box tap, the foam of the weight-reducing core layer is subjected to drying treatment; the drying treatment is drying in an oven at 130 ℃ +/-5 ℃ for 3 hours.
Preferably, the pre-curing treatment comprises: sequentially wrapping the assembly coated with the prepreg by using demolding cloth and an air-permeable felt, and putting into a vacuum bag; horizontally putting a vacuum bag into an oven, tearing off a protective film on the surface of a sealing adhesive tape of the vacuum bag, inserting a vacuum joint, and sealing the vacuum bag; and opening a vacuum pump, ensuring that the vacuum degree in the vacuum bag is less than-0.095 MPa, setting the temperature of the oven to 35 ℃, starting timing when the temperature of the oven reaches 35 ℃, keeping the temperature for 30min, then closing the power supply of the oven, and taking out the assembly.
Optionally, the fiber prepregs with different thicknesses are coated at different positions according to the loading condition of the airfoil, and the method comprises the following steps:
integrally covering 7mm areas of the upper wing surface and the lower wing surface: sequentially paving prepreg in the directions of [ +45/0/0/-45], [0/0/0/90], [90/0/0 ], [ -45/0/0/+45] with the main bearing Liang Qianyuan as the 0-degree direction, wherein the thickness of the prepreg is 7mm,
integrally covering 5mm areas of the upper wing surface and the lower wing surface: sequentially paving prepreg in the directions of [ +45/0/0/-45], [ -45/0/0/+45] and [ + 45/0/+45] with the main bearing Liang Qianyuan as the 0-degree direction, wherein the thickness of the prepreg is 5mm;
integrally paving 4mm areas and joint areas of the upper and lower wing surfaces: sequentially paving prepreg in the directions of [0/0/0/90] and [90/0/0/0] in a direction of 0 degree by taking the main bearing Liang Qianyuan as a direction, wherein the thickness of the prepreg is 4mm;
wholly lay upper and lower airfoil 3mm region and joint area: and (3) sequentially paving prepregs in the directions of [ +45/0/0/-45] [0/0/0/90] [ +45/0/0/-45] [ -45/0/0/+45] [90/0/0 ] [ -45/0/0/+45] with the main bearing Liang Qianyuan as the 0-degree direction, wherein the thickness of the prepregs is 3mm.
Compared with the prior art, the invention has the beneficial effects that:
1) A composite material airfoil layering method of a main joint and wing box tap combined structure is designed, the layering scheme can fill carbon fibers in a metal joint, the filled carbon fibers and an outer skin are combined to form a plurality of T-shaped beam structures (the carbon fibers in the inner part are in the Z direction, namely the vertical direction, the skin is in the XY direction, namely the horizontal direction, and the carbon fibers and the skin are overlapped together to form the T-shaped structure), and the rigidity strength of a main bearing beam is effectively improved;
2) The invention positions and combines the plurality of weight-reducing core layers and the rectangular beam consisting of the carbon fibers with metal parts in various forms such as butt joint, cementing, screw joint and the like, and integrally solidifies and forms, thereby further improving the reliability of the structure.
Drawings
FIG. 1 is a top view of a composite primary load beam.
FIG. 2 is a bottom view of a composite primary load beam.
Fig. 3 is a top view of the primary adapter.
Fig. 4 is a bottom view of the primary adapter.
Figure 5 is a top view of a wing box tap.
Figure 6 is a side view of a wing box tap.
Figure 7 is a schematic view of a wing box stiffener.
Fig. 8 is a plan view of a metal member including a main joint, a wing box tap, and a wing box reinforcing plate.
Fig. 9 is a bottom view of the metal member including the main joint, the wing box tap, and the wing box reinforcing plate.
FIG. 10 is a top view of the layup "male mold".
FIG. 11 is a bottom view of the layup "male mold".
Fig. 12 is a plan view of the weight-reducing core layer.
Fig. 13 is a bottom view of the weight-reducing core layer.
Detailed Description
In order to make the aforementioned objects, features and advantages of the present invention comprehensible, the present invention shall be described in further detail with reference to the following detailed description and accompanying drawings.
The composite material main bearing Liang Ru of the invention is shown in fig. 1 and fig. 2 and comprises a main joint, a wing box tap, a wing box reinforcing sheet and a weight-reducing core layer, and the outer part of the composite material main bearing is coated with a skin (the skin is not shown in fig. 1 and fig. 2). The main joint is fixedly connected with the wing box tap and the wing box reinforcing sheet, and the weight reducing core layer is fixedly connected with the wing box tap. The covering is formed by coating fiber prepreg on the outer portions of the main joint, the wing box tap joint, the wing box reinforcing sheet and the weight-reducing core layer according to the loading condition of the wing surface and integrally molding and curing.
The weight-reducing core layer is made of foam materials, such as rigid polyurethane foam, phenolic resin foam or PMI foam.
The composite main load-bearing beam can also comprise a front edge web and a rear edge web, as shown in fig. 1 and 2, the front edge web and the rear edge web are positioned on the front edge and the rear edge of the main load-bearing beam and are used for transmitting Z-direction (vertical direction) load, and the composite main load-bearing beam is made of carbon fiber resin matrix composite materials and other materials. The main load beam may also be provided with webs, i.e. panels between the foam of the weight reducing core layer connecting the upper and lower composite airfoils (see figure 10).
The main joint, the wing box tap joint and the wing box reinforcing sheet are metal pieces made of metal materials such as titanium alloy or stainless steel. The main joint is used for connecting the main bearing beam and a folding and unfolding mechanism of the aircraft.
In one embodiment of the invention, the main joint is as shown in figures 3 and 4, the wing box tap is as shown in figures 5 and 6 and the wing box stiffener is as shown in figure 7. Fig. 8 and 9 are structural views of a metal fitting formed by assembling a main joint, a wing box tap, and a wing box reinforcing sheet. The wing box taps are mounted in corresponding slots in the main joint and are fixedly connected with the main joint by means of screws and the like. The wing box reinforcing plate is arranged in a groove reserved in the wing box tap. The wing box tap is provided with a slot into which one end of the weight reducing core layer is inserted. In this embodiment there are 4 wing box taps, and in other embodiments other numbers of wing box taps may be provided.
The integral forming method of the composite material main bearing beam comprises the steps of firstly combining the wing box tap, the wing box reinforcing sheet and the weight-reducing core layer, then coating carbon fiber prepreg, pre-curing, then combining with the main joint to form a layering male mold, as shown in figures 10 and 11, then coating carbon fiber prepreg with different thicknesses at different positions of the male mold according to the loading condition of an airfoil surface, and finally carrying out integral mould pressing, curing and forming.
The carbon fiber prepreg comprises carbon fibers and resin, wherein the carbon fibers are unidirectional T300, T700, T800 or T1000 carbon fibers, and the resin is medium-temperature or high-temperature curing epoxy resin.
In one embodiment of the invention, a method for integrally forming a composite main load-bearing beam is provided, which comprises the following steps:
assembling the wing box tap with a weight-reducing core layer, wherein one end of the weight-reducing core layer is arranged in a groove reserved in the wing box tap;
assembling a wing box reinforcing sheet with a wing box tap joint and a weight-reducing core layer, wherein the wing box reinforcing sheet is arranged in a groove reserved in the wing box tap joint and the weight-reducing core layer;
coating a fiber prepreg on an assembly of the wing box tap, the wing box reinforcing sheet and the weight-reducing core layer, and performing pre-curing treatment;
assembling the pre-cured wing box tap joint, the wing box reinforcing sheet and the weight-reducing core layer assembly with the main joint;
and for the combined parts of the main joint, the wing box tap joint, the wing box reinforcing sheet and the weight-reducing core layer, fiber prepregs with different thicknesses are coated at different positions according to the loading condition of the airfoil, and finally, the whole body is molded, cured and molded.
In one embodiment of the invention, the main joint, the wing box tap and the wing box reinforcing sheet are subjected to sand blasting treatment before the main joint, the wing box tap and the wing box reinforcing sheet are assembled, so that the roughness of the surface of the metal piece is improved, and the bonding strength of the metal piece is improved. Further preferably, the sand blasting is performed by using 20-40 mesh quartz sand with the air pressure of 0.4-0.7 MPa.
In one embodiment of the invention, the foam of the weight reducing core layer is dried prior to assembly of the weight reducing core layer with the wing box taps. Further preferably, the drying treatment is drying in an oven at 130 ℃ ± 5 ℃ for 3 hours.
In one embodiment of the present invention, the pre-curing process is performed by placing the assembly coated with the prepreg in a vacuum bag and then placing the assembly in an oven for pre-curing. Further preferably, the assembly coated with the prepreg is sequentially wrapped with a release cloth and an air felt and placed in a vacuum bag; putting the vacuum bag into an oven in a horizontal state, tearing off the surface protective film of the sealing adhesive tape of the vacuum bag, inserting a vacuum joint, and sealing the vacuum bag; opening the vacuum pump, ensuring that the vacuum degree in the vacuum bag is less than-0.095 MPa, and vacuumizing for 30min at room temperature; if the room temperature is lower than 25 ℃, vacuumizing and pre-compacting in an oven, setting the temperature of the oven to be 35 ℃, starting timing when the temperature of the oven reaches 35 ℃, keeping the temperature for 30min, then closing the power supply of the oven, and taking out the assembly.
In an embodiment of the invention, the wrapping of the fiber prepregs with different thicknesses at different positions according to the loading condition of the airfoil comprises the following steps:
integrally covering 7mm areas of the upper wing surface and the lower wing surface: sequentially paving prepreg in the directions of [ +45/0/0/-45], [0/0/0/90], [90/0/0 ], [ -45/0/0/+45] with the main bearing Liang Qianyuan as the 0-degree direction, wherein the thickness of the prepreg is 7mm,
integrally covering 5mm areas of the upper wing surface and the lower wing surface: sequentially paving prepreg in the directions of [ +45/0/0/-45], [ -45/0/0/+45] and [ + 45/0/+45] with the main bearing Liang Qianyuan as the 0-degree direction, wherein the thickness of the prepreg is 5mm;
integrally paving 4mm areas and joint areas of the upper and lower wing surfaces: sequentially paving prepreg in the directions of [0/0/0/90] and [90/0/0/0] in a direction of 0 degree by taking the main bearing Liang Qianyuan as a direction, wherein the thickness of the prepreg is 4mm;
wholly lay upper and lower airfoil 3mm region and joint area: and (3) sequentially paving prepregs in the directions of [ +45/0/0/-45] [0/0/0/90] [ +45/0/0/-45] [ -45/0/0/+45] [90/0/0 ] [ -45/0/0/+45] with the main bearing Liang Qianyuan as the 0-degree direction, wherein the thickness of the prepregs is 3mm.
In one embodiment of the invention, a method for integrally forming a composite main load-bearing beam is provided, which comprises the following steps:
in a first step, the main joint, the wing box taps and the wing box stiffener are grit blasted. The sandblasting is to improve the roughness of the surface of the metal member and to improve the bonding strength of the metal member. The quartz sand used for sand blasting is 20-40 meshes, and the air pressure is 0.4-0.7 MPa.
The second step is that: after drying the foam of the weight-reducing core layer of the main load-bearing beam in an oven at 130 +/-5 ℃ for 3 hours, cleaning the surface debris by using a blower. The dried main beam foam is sealed and stored immediately, is forbidden to be exposed and stored in the air, and can be taken out when in use. In this embodiment, the weight-reduced core layer, i.e., the foam, is shown in fig. 12 and 13, and includes 4 strip-shaped foams, and a strip-shaped gap is provided between the 4 strip-shaped foams at one end, so as to facilitate insertion into a groove reserved on the tap of the wing box.
The third step: and (4) butting each piece of foam with the corresponding wing box tap joint for trial assembly, and inserting the foam into a groove reserved in the wing box tap joint. The foam is trimmed according to the condition and can be polished by 120-mesh abrasive paper, so that the foam is completely matched with the insertion end face of the tap joint of the wing box, the gap is smaller than 1mm, and enough material spreading gaps are reserved between webs of the foam.
The fourth step: and uniformly sticking a layer of J-47A adhesive film on the wing box tap and the foam bonding area, and butting and assembling.
The fifth step: and (3) butting and trial-assembling the wing box reinforcing sheet, the wing box tap and the foam, arranging the wing box reinforcing sheet in the wing box tap and the reserved groove of the foam machine, locally finishing the wing box reinforcing sheet, and enabling the wing box reinforcing sheet to be flush with the butt joint end faces of the wing box tap and the foam.
And a sixth step: uniformly sticking a layer of J-47A adhesive film on the bonding area of the wing box reinforcing sheet, and butting and assembling with the wing box tap and the foam.
The seventh step: the 4 wing box taps, the wing box reinforcing sheet and the foam assembly are respectively coated with a layer of prepreg in the direction of [90/-45/0/+45], and the prepreg are sequentially butted at the web corners of the upper wing surface and the lower wing surface of each foam.
Eighth step: and (3) sequentially wrapping the assembly covered with the prepreg by using demolding cloth and an air-permeable felt, and putting the assembly into a vacuum bag prepared in advance.
And putting the vacuum bag into an oven in a horizontal state, tearing off the surface protective film of the sealing adhesive tape of the vacuum bag, inserting a vacuum joint, and sealing the vacuum bag.
And opening the vacuum pump to ensure that the vacuum degree in the vacuum bag is less than-0.095 MPa, and vacuumizing for 30min at room temperature. If the room temperature is lower than 25 ℃, vacuumizing and pre-compacting in an oven, setting the temperature of the oven to be 35 ℃, starting timing when the temperature of the oven reaches 35 ℃, keeping the temperature for 30min, then closing the power supply of the oven, and taking out the assembly.
The ninth step: a layer of J-47A adhesive film is uniformly adhered to the bonding area of the main joint, and the tap joint of the wing box is placed at the corresponding position of the main joint to ensure that the contact surface of the wing box and the main joint is completely attached, and the gap is less than 1mm.
The tenth step: and removing prepreg at the mounting positions of the wing box tap and the main joint corresponding to the screws, uniformly brushing a layer of anaerobic adhesive on the threaded areas of the screws, putting the screws into the corresponding mounting positions, and screwing down the screws to ensure that the screws are flush with the surfaces of the wing box tap. And filling a proper amount of prepreg in the screw hole to enable the surface of the screw to be flush.
The eleventh step: the 7mm area of the upper wing surface and the lower wing surface is integrally paved. The prepreg in the directions of [ +45/0/0/-45], [0/0/0/90], [90/0/0 ], [ -45/0/0/+45] is sequentially laid on the main bearing Liang Qianyuan in the 0-degree direction, and the trailing edges of the upper airfoil web corners are sequentially butted. The position of the region with a prepreg thickness of 7mm,7mm in this step is shown in FIG. 11.
The twelfth step: the 5mm area of the upper wing surface and the lower wing surface is integrally paved. Prepreg materials in the directions of [ +45/0/0/-45], [ -45/0/0/+45] and [ + 45/0/+45] are sequentially laid on the main bearing Liang Qianyuan as the 0-degree direction, and are sequentially butted at the trailing edges of the corners of the lower airfoil web respectively. The position of the region having a prepreg thickness of 5mm,5mm in this step is shown in FIG. 11.
The thirteenth step: the 4mm area and the joint area of the upper wing surface and the lower wing surface are integrally paved. And (3) sequentially paving prepreg in the directions of [0/0/0/90] and [90/0/0/0] by taking the main bearing Liang Qianyuan as the 0-degree direction, and sequentially butting the trailing edges of the web plates at the upper wing surfaces respectively. The positions of the prepreg in the present step in the region of 4mm,4mm in thickness and the joint region are shown in fig. 11.
A fourteenth step of: and (3) sequentially wrapping the component paved with the prepreg by using demolding cloth and an air-permeable felt, and putting the component into a vacuum bag prepared in advance.
And putting the vacuum bag into an oven in a horizontal state, tearing off the surface protective film of the sealing adhesive tape of the vacuum bag, inserting a vacuum joint, and sealing the vacuum bag.
And opening the vacuum pump to ensure that the vacuum degree in the vacuum bag is less than-0.095 MPa, and vacuumizing for 30min at room temperature. If the room temperature is lower than 25 ℃, vacuumizing and pre-compacting in an oven, setting the temperature of the oven to be 35 ℃, starting timing when the temperature of the oven reaches 35 ℃, keeping the temperature for 30min, then closing the power supply of the oven, taking out the assembly,
the fifteenth step: the 3mm area of the upper wing surface and the lower wing surface and the joint area are integrally paved. Prepreg in the direction of [ +45/0/0/-45] [0/0/0/90] [ + 45/0/0/0/0/-45 ] [ +45/0/0/-45] [ -45/0/0/+45] [90/0/0/0] [ -45/0/0/+45] is sequentially laid from the wing root to the wing tip with the main bearing Liang Qianyuan as the 0-degree direction, and the prepreg is respectively butted with the trailing edges of the upper and lower wing web corners in sequence. The position of the prepreg in this step at the area of thickness 3mm,3mm and the joint area is shown in FIG. 11.
Sixteenth, step: and (3) loading the coated airfoil surface prefabricated part into a compression molding die, and placing the die on a hot press.
Seventeenth step: heating the hot press from room temperature → 130 ℃ +/-5 ℃ → heat preservation 2h at the speed of (2-5) ° c/min → applying the pressure of 0.3MPa → continuously heating to 185 ℃ +/-5 ℃ → heat preservation 2h, and closing the hot press.
And eighteenth step: naturally cooling to the temperature of the mould below 50 ℃ to demould the product.
The position and direction of the prepreg to be laid in each of the above steps may be adjusted according to actual circumstances, and are not limited to the embodiments in the above steps.
The particular embodiments of the present invention disclosed above are illustrative only and are not intended to be limiting, since various alternatives, modifications, and variations will be apparent to those skilled in the art without departing from the spirit and scope of the invention. The invention should not be limited to the disclosure of the embodiments in the specification, but the scope of the invention is defined by the appended claims.

Claims (10)

1. The composite main load-bearing beam is characterized by comprising a main joint, a wing box tap joint, a wing box reinforcing sheet, a weight-reducing core layer and a skin, wherein the main joint is fixedly connected with the wing box tap joint and the wing box reinforcing sheet, and the weight-reducing core layer is fixedly connected with the wing box tap joint; the main joint, the wing box tap, the wing box reinforcing sheet and the weight-reducing core layer are externally coated with the skin.
2. The composite main load beam as claimed in claim 1, wherein the skin is formed by coating fiber prepreg on the outer portions of the main joint, the wing box tap, the wing box reinforcing sheet and the weight-reducing core layer according to the loading condition of the wing surface and integrally molding and curing.
3. The composite primary load beam of claim 2 wherein said fiber prepreg is a carbon fiber prepreg comprising carbon fibers and a resin, the carbon fibers being unidirectional T300, T700, T800 or T1000 carbon fibers and the resin being a medium or high temperature curing epoxy resin.
4. The composite main load beam defined in claim 1 wherein said main joint, said wing box tap, said wing box stiffener are metallic pieces; the weight-reducing core layer is made of hard polyurethane foam, phenolic resin foam or PMI foam.
5. The composite main carrier beam as defined in claim 1 wherein the wing box taps are mounted in respective slots in the main joint, the wing box stiffener is mounted in a slot reserved in the wing box tap and the weight reducing core layer, one end of the weight reducing core layer being mounted in a slot reserved in the wing box tap.
6. The integral forming method of the composite main load-bearing beam is characterized by comprising the following steps of:
assembling the wing box tap and the weight reduction core layer, wherein one end of the weight reduction core layer is arranged in a groove reserved in the wing box tap;
assembling the wing box reinforcing sheet with a wing box tap and a weight-reducing core layer, wherein the wing box reinforcing sheet is arranged in a groove reserved in the wing box tap and the weight-reducing core layer;
coating a fiber prepreg on an assembly of the wing box tap, the wing box reinforcing sheet and the weight-reducing core layer, and performing pre-curing treatment;
assembling the pre-cured wing box tap joint, the wing box reinforcing sheet and the weight-reducing core layer assembly with the main joint;
and for the combined parts of the main joint, the wing box tap joint, the wing box reinforcing sheet and the weight-reducing core layer, fiber prepregs with different thicknesses are coated at different positions according to the loading condition of the wing surface, and the whole body is molded, cured and molded.
7. The method of claim 6 wherein the main joint, wing box tap and wing box stiffener are grit blasted prior to assembly of the main joint, wing box tap and wing box stiffener; the sand blasting treatment adopts 20-40 mesh quartz sand, and the air pressure is 0.4-0.7 MPa.
8. The method of claim 6, wherein the weight-reducing core foam is subjected to a drying process prior to assembly of the weight-reducing core with the wing box taps; the drying treatment is drying in an oven at 130 ℃ +/-5 ℃ for 3 hours.
9. The method of claim 6, wherein the pre-curing process comprises: sequentially wrapping the assembly coated with the prepreg by using demolding cloth and an air-permeable felt, and putting into a vacuum bag; horizontally putting a vacuum bag into an oven, tearing off a protective film on the surface of a sealing adhesive tape of the vacuum bag, inserting a vacuum joint, and sealing the vacuum bag; and opening a vacuum pump, ensuring that the vacuum degree in the vacuum bag is less than-0.095 MPa, setting the temperature of the oven to 35 ℃, starting timing when the temperature of the oven reaches 35 ℃, keeping the temperature for 30min, then closing the power supply of the oven, and taking out the assembly.
10. The method of claim 6, wherein the wrapping of fiber prepregs of different thicknesses at different locations according to loading of the airfoil comprises:
integrally covering 7mm areas of the upper wing surface and the lower wing surface: sequentially paving prepreg in the directions of [ +45/0/0/-45], [0/0/0/90], [90/0/0 ], [ -45/0/0/+45] with the main bearing Liang Qianyuan as the 0-degree direction, wherein the thickness of the prepreg is 7mm,
integrally covering 5mm areas of the upper wing surface and the lower wing surface: sequentially paving prepreg in the directions of [ +45/0/0/-45], [ -45/0/0/+45] and [ + 45/0/+45] with the main bearing Liang Qianyuan as the 0-degree direction, wherein the thickness of the prepreg is 5mm;
integrally paving 4mm areas and joint areas of the upper and lower wing surfaces: sequentially paving the prepreg in the directions of [0/0/0/90] and [90/0/0/0] in the direction of 0 degree by taking Liang Qianyuan as a main bearing, wherein the thickness of the prepreg is 4mm;
wholly lay upper and lower airfoil 3mm region and joint area: and (3) sequentially paving prepregs in the directions of [ +45/0/0/-45] [0/0/0/90] [ +45/0/0/-45] [ -45/0/0/+45] [90/0/0 ] [ -45/0/0/+45] with the main bearing Liang Qianyuan as the 0-degree direction, wherein the thickness of the prepregs is 3mm.
CN202210691660.3A 2022-06-17 2022-06-17 Composite material main bearing beam and integral forming method Pending CN115214872A (en)

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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN115742343A (en) * 2022-11-11 2023-03-07 航天特种材料及工艺技术研究所 Tenon-and-mortise connected composite material airfoil and forming method thereof
CN117073467A (en) * 2023-09-04 2023-11-17 北京爱思达航天科技有限公司 Light composite material flying wing capable of carrying devices and preparation method thereof
CN117073467B (en) * 2023-09-04 2024-07-26 北京爱思达航天科技有限公司 Light composite material flying wing capable of carrying devices and preparation method thereof

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN115742343A (en) * 2022-11-11 2023-03-07 航天特种材料及工艺技术研究所 Tenon-and-mortise connected composite material airfoil and forming method thereof
CN117073467A (en) * 2023-09-04 2023-11-17 北京爱思达航天科技有限公司 Light composite material flying wing capable of carrying devices and preparation method thereof
CN117073467B (en) * 2023-09-04 2024-07-26 北京爱思达航天科技有限公司 Light composite material flying wing capable of carrying devices and preparation method thereof

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