CN113320244A - Impact-resistant fiber metal hybrid structure for aviation and manufacturing method thereof - Google Patents

Impact-resistant fiber metal hybrid structure for aviation and manufacturing method thereof Download PDF

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CN113320244A
CN113320244A CN202110637069.5A CN202110637069A CN113320244A CN 113320244 A CN113320244 A CN 113320244A CN 202110637069 A CN202110637069 A CN 202110637069A CN 113320244 A CN113320244 A CN 113320244A
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resin
titanium alloy
manufacturing
spraying
metal
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CN113320244B (en
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彭华新
叶竟
王欢
高岩
刘成
董家乐
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Zhejiang University ZJU
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Zhejiang University ZJU
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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B15/00—Layered products comprising a layer of metal
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B15/00—Layered products comprising a layer of metal
    • B32B15/14—Layered products comprising a layer of metal next to a fibrous or filamentary layer
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B37/00—Methods or apparatus for laminating, e.g. by curing or by ultrasonic bonding
    • B32B37/06—Methods or apparatus for laminating, e.g. by curing or by ultrasonic bonding characterised by the heating method
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B37/00—Methods or apparatus for laminating, e.g. by curing or by ultrasonic bonding
    • B32B37/10—Methods or apparatus for laminating, e.g. by curing or by ultrasonic bonding characterised by the pressing technique, e.g. using action of vacuum or fluid pressure
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B37/00—Methods or apparatus for laminating, e.g. by curing or by ultrasonic bonding
    • B32B37/12—Methods or apparatus for laminating, e.g. by curing or by ultrasonic bonding characterised by using adhesives
    • B32B37/1284—Application of adhesive
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B38/00—Ancillary operations in connection with laminating processes
    • B32B38/0012—Mechanical treatment, e.g. roughening, deforming, stretching
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B38/00—Ancillary operations in connection with laminating processes
    • B32B38/0036—Heat treatment
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B38/00—Ancillary operations in connection with laminating processes
    • B32B38/16—Drying; Softening; Cleaning
    • B32B38/164—Drying
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B5/00—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts
    • B32B5/02—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts characterised by structural features of a fibrous or filamentary layer
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B5/00—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts
    • B32B5/02—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts characterised by structural features of a fibrous or filamentary layer
    • B32B5/12—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts characterised by structural features of a fibrous or filamentary layer characterised by the relative arrangement of fibres or filaments of different layers, e.g. the fibres or filaments being parallel or perpendicular to each other
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B7/00—Layered products characterised by the relation between layers; Layered products characterised by the relative orientation of features between layers, or by the relative values of a measurable parameter between layers, i.e. products comprising layers having different physical, chemical or physicochemical properties; Layered products characterised by the interconnection of layers
    • B32B7/04—Interconnection of layers
    • B32B7/12—Interconnection of layers using interposed adhesives or interposed materials with bonding properties
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B38/00—Ancillary operations in connection with laminating processes
    • B32B38/0036—Heat treatment
    • B32B2038/0048—Annealing, relaxing
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B38/00—Ancillary operations in connection with laminating processes
    • B32B38/16—Drying; Softening; Cleaning
    • B32B38/164—Drying
    • B32B2038/168—Removing solvent
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B2262/00—Composition or structural features of fibres which form a fibrous or filamentary layer or are present as additives
    • B32B2262/10—Inorganic fibres
    • B32B2262/106—Carbon fibres, e.g. graphite fibres
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B2307/00—Properties of the layers or laminate
    • B32B2307/50—Properties of the layers or laminate having particular mechanical properties
    • B32B2307/56—Damping, energy absorption
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B2307/00—Properties of the layers or laminate
    • B32B2307/70—Other properties
    • B32B2307/718—Weight, e.g. weight per square meter

Abstract

The invention discloses an impact-resistant fiber metal hybrid structure for aviation and a manufacturing method thereof. The method uses a metal surface etching combined resin spraying process to obtain a firm and light interface layer of the metal and resin-based fiber composite material, so that the resin is fully contacted and wetted with the metal surface, and the thickness of the interface layer is uniform, light and high in strength. Compared with the existing methods of glue film + mechanical polishing, spraying + anodic oxidation and the like, the fiber metal hybrid structural member manufactured by the method has the advantages that the interface bonding force of the hybrid structure prepared by the spraying + surface etching process reaches 59.5MPa, and the fiber metal hybrid structural member has the best anti-layering capability; under the 60J energy impact, the deformation amount of the sample with the same structure prepared by the method is the least, the structure is kept the most intact, and the method is proved to have higher impact resistance.

Description

Impact-resistant fiber metal hybrid structure for aviation and manufacturing method thereof
Technical Field
The invention belongs to the field of production and manufacturing of structural composite materials, and particularly relates to manufacturing of a light, high-strength and impact-resistant titanium-based fiber metal hybrid structural member for aircraft manufacturing requirements.
Background
The fiber metal hybrid structure is a composite structural member formed by mixing, laying and hot-pressing metal and resin matrix composite materials according to structural requirements. The structure is originally proposed by a scholars of the Holland Delft university, the expansion of metal fatigue cracks is well relieved through the interface bridging of metal and resin matrix composite materials, and meanwhile, the structure has good damage tolerance and high strength and modulus, so that the structure becomes a research direction for lightweight design of aerospace and vehicles.
Impact damage is one form of damage that may occur to the structure of an aircraft during service. Generally, metal materials have very good ductility and can absorb a large amount of energy in the elastic-plastic deformation stage; for most resin-based composites, the barely visible damage can greatly reduce the structural load-bearing capacity of the composite, since the material itself is very brittle and has no significant plastic deformation stage. The fiber metal hybrid structure is formed by compounding metal and resin matrix composite materials, and the impact resistance of the fiber metal hybrid structure is between the two materials; the bonding strength of the metal and the resin matrix composite material and the thickness of the adhesive layer have great relation to the impact delamination resistance and the bearing capacity of the fiber metal hybrid structure. When the glue layer is thin, the requirement on the wettability of the surface of the material is high, and if the glue layer is not wet well during bonding, a poor glue area is easy to appear, so that the resin-based composite material and metal cannot form an effective bridging effect, and when the glue film is thick, the bearing capacity of the hybrid structure is influenced. The existing titanium alloy anodic oxidation and adhesive film bonding process has the defects of not strong interface bonding force and not ideal mechanical properties due to thick adhesive layer. Therefore, the impact resistance of the fiber metal hybrid structure needs to be further improved, and the metal interface and the bonding process need to be started simultaneously. The spraying and surface etching process adopted by the invention can solve the problems and obviously improve the mechanical bearing capacity and the shock resistance of the fiber metal hybrid structure.
Disclosure of Invention
The invention aims to solve the technical problems that the novel manufacturing process and method are used for improving the bonding force of a metal and resin interface and reducing the thickness of a bonding layer under the condition of not influencing the bonding quality, so that the impact resistance and the structural bearing capacity of a fiber metal hybrid structure are improved.
The solution specifically adopted by the invention is as follows:
a method for manufacturing an impact-resistant fiber metal hybrid structure for aviation comprises the following steps:
s1, processing the titanium alloy with the cleaned surface by using a surface etching technology, and annealing the metal after etching;
s2, uniformly spraying a resin solution dissolved in an organic solvent on the surface of the titanium alloy treated by the S1, and then drying to fully volatilize the organic solvent to obtain the sized titanium alloy;
s3, alternately laying and pre-compacting the rubberized titanium alloy and the resin-based fiber prepreg according to a preset laying sequence to form a pre-compaction member, and curing and forming the pre-compaction member to form the fiber metal hybrid structure.
As a further improvement of the method of the present invention, the surface cleaning treatment includes degreasing the surface of the metal layer and removing an oxide film.
As a further improvement of the method of the present invention, the resin-based fiber prepreg is a carbon fiber prepreg or a glass fiber prepreg.
As a further improvement of the method of the invention, the surface etching technology is alkaline etching or NH4And F, etching by using solution.
As a further improvement of the method, the etched pattern on the surface of the titanium alloy is a net structure or a porous array.
As a further improvement of the process according to the invention, the organic solvent is dichloromethane or chloroform.
As a further improvement of the process of the present invention, the resin is an epoxy resin.
As a further improvement of the method, the spraying thickness of the resin solution on the glue layer on the surface of the titanium alloy is controlled to be 2-7 mg/cm2And controlling the spraying pressure to be 2-4 bar.
As a further improvement of the method, the prepressing component is formed by laying and prepressing M layers of rubberized titanium alloy and N layers of resin-based fiber prepregs according to the laying sequence of [ Ti/FRP/Ti/FRP/Ti …/FRP/Ti ], wherein Ti represents the titanium alloy, and FRP represents one or more layers of resin-based prepreg fibers; m is more than or equal to 2, and N is more than or equal to 1.
As a further improvement of the method, the curing and forming process is die pressing or hot press curing, and the used equipment is an autoclave or a vulcanizing press.
In another aspect, the present invention provides an aerospace impact-resistant fiber metal hybrid structure made by the method of any one of the above aspects.
By the scheme, the invention at least has the following advantages: on one hand, the metal etching scheme used by the invention can effectively improve the bonding force of the metal and resin interface and has better delamination resistance. On the other hand, the prepared adhesive is used for surface spraying, so that the surface to be adhered can be completely covered, the wetting is good, and the thickness is thin and uniform. Compared with the traditional process, the fiber metal hybrid structure manufactured by the glue spraying and surface etching process has better structural strength, better impact resistance and delamination resistance and higher damage tolerance.
The foregoing description is only an overview of the technical solutions of the present invention, and in order to clearly understand the technical solutions of the present invention and to implement the technical solutions according to the contents of the description, the following is a preferred embodiment of the present invention, and the following detailed description is given with reference to the accompanying drawings.
Drawings
Fig. 1 is a schematic diagram (left diagram) and an autoclave diagram (right diagram) of the structure of the fiber metal hybrid laminate when M is 3 and N is 4;
FIG. 2 is a diagram showing the structure of the etched titanium alloy surface;
FIG. 3 is a graph of interlaminar shear performance of fiber metal hybrid structures prepared by different fabrication processes;
FIG. 4 is a graph of the mechanical response of a hybrid fiber metal structure made by different processes at 60J energy impact;
fig. 5 is an ultrasonic scan of the sample after a 60J low velocity impact (deformation zone boundaries are shown by black dots).
Detailed Description
The following detailed description of embodiments of the present invention is provided in connection with the accompanying drawings and examples. The following examples are intended to illustrate the invention but are not intended to limit the scope of the invention.
In the invention, a method for manufacturing an impact-resistant fiber metal hybrid structure for aviation is provided, which comprises the following steps:
(1) and (3) carrying out surface cleaning treatment on the titanium alloy, namely, carrying out surface oil removal and oxide film removal to obtain a clean titanium surface.
(2) And etching the surface of the titanium alloy subjected to surface cleaning treatment by using a surface etching technology under a certain condition, and annealing the metal after etching. The etching technique may be alkali etching or NH4F, etching by solution, wherein the etching pattern is in a net structure or a porous array, and the pattern formed by surface etching can provide an interface for easy combination of the resin。
(3) The resin is dissolved in an organic solvent according to a certain proportion by using the organic solvent to form a resin solution. The organic solvent used may be dichloromethane or chloroform, and the resin used may be an epoxy resin.
(4) And (3) uniformly spraying the resin solution in the step (3) on the surface of the titanium alloy treated in the step (2) under certain air pressure by using a spray gun. The parameters in the resin solution spraying process can be adjusted according to the actual conditions, and the spraying thickness of the glue layer can be controlled to be 2-7 mg/cm2The spraying pressure can be controlled to be 2-4 bar.
(5) And (4) putting the titanium alloy sprayed in the step (4) into an oven for airing, and fully volatilizing the organic solvent to obtain the rubberized titanium alloy.
(6) And (5) alternately laying the rubberized titanium alloy and the resin-based fiber prepreg in the step (5) according to a preset laying sequence and pre-compacting to form a pre-compaction component. The ply sequence may be expressed as [ Ti/FRP/Ti …/FRP/Ti ], Ti representing a titanium alloy and FRP representing one or more layers of resin-based fibre prepreg. Therefore, if the number of titanium alloy layers is M and the number of resin-based fiber prepreg layers is N, M is more than or equal to 2 and N is more than or equal to 1.
(7) And (4) curing and forming the pre-pressed member in the step (6) to obtain the impact-resistant fiber metal hybrid structure for aviation. The curing and forming process is die pressing or hot pressing curing, the used equipment can be an autoclave or a flat vulcanizing machine, and the specific curing temperature and pressure can be adjusted according to the curing temperature and pressure required by the adopted resin.
In the fiber metal hybrid structure, the structure can be divided into a metal layer, an adhesive layer and a resin-based composite material layer. Wherein the metal layer is a titanium alloy body; the resin-based composite material layer is a composite material manufactured by prepreg, such as a carbon fiber composite material or a glass fiber composite material; the bonding layer is a resin-rich area formed by bonding the titanium alloy and the resin-based composite material.
In order to further demonstrate the specific technical effects of the present invention, as shown in fig. 1, the specific implementation of the above-mentioned manufacturing method and the properties of the resulting fiber metal hybrid structure will now be described in detail by several embodiments.
Example 1:
a method for manufacturing an impact-resistant fiber metal hybrid structure for aviation, comprising the steps of:
(1) degreasing the surface of the layered titanium alloy by using a surfactant, and using HF and HNO3Mixed acid is used for removing an oxidation film;
(2) putting the titanium alloy into an electrolytic bath by using ammonium fluoride electrolyte with the concentration of 1% for reacting for 5min, taking out the titanium alloy after etching is finished, cleaning the surface of the titanium alloy, and then annealing the titanium alloy;
(3) dissolving epoxy resin glue with a curing agent into dichloromethane by using dichloromethane as an organic solvent, wherein the mass fraction of the epoxy resin glue is 10%, and forming a resin solution;
(4) uniformly spraying the resin solution in the step (3) on the surface of the titanium alloy treated in the step (2) by using a spray gun under the air pressure of 2-4 bar, wherein the spraying thickness of the glue layer is 4.5mg/cm2；
(5) Putting the titanium alloy sprayed in the step (4) in a 50 ℃ oven for 40min, and fully volatilizing the organic solvent to obtain the rubberized titanium alloy;
(6) laying the titanium alloy and the carbon fiber prepreg in the step (5) according to a laying sequence of [ Ti/0/90/Ti/90/0/Ti ] and pre-compacting, wherein the specific laying structure schematic diagram is shown in figure 1, and two layers of carbon fiber prepregs are laid between two adjacent layers of titanium alloy according to laying angles of 0 degree and 90 degrees respectively;
(7) and (5) placing the prepressing component in the step (6) into an autoclave for curing molding according to the curing temperature and pressure required by the resin.
Example 2:
a preparation method for improving interlayer strength of a fiber metal laminated plate comprises the following steps:
(1) degreasing the surface of the layered titanium alloy by using a surfactant, and using HF and HNO3Mixed acid is used for removing an oxidation film;
(2) performing alkali etching on the surface of the titanium alloy by using NaOH with the concentration of 2M at the temperature of 80 ℃ and under the voltage of 30V for 40min, and then annealing the titanium alloy;
(3) dissolving epoxy resin glue with a curing agent in chloroform by using chloroform as an organic solvent, wherein the mass fraction of the epoxy resin glue is 7 percent, and forming a resin solution;
(4) uniformly spraying the resin solution in the step (3) on the surface of the titanium alloy treated in the step (2) by using a spray gun under the air pressure of 2-4 bar, wherein the spraying thickness of the glue layer is 4.5mg/cm2；
(5) Putting the titanium alloy sprayed in the step (4) in a 50 ℃ oven for 40min, and fully volatilizing the organic solvent to obtain the rubberized titanium alloy;
(6) laying the titanium alloy and the carbon fiber prepreg in the step (5) according to a laying sequence of [ Ti/0/90/Ti/90/0/Ti ] and pre-compacting, wherein the specific laying structure schematic diagram is shown in figure 1, and two layers of carbon fiber prepregs are laid between two adjacent layers of titanium alloy according to laying angles of 0 degree and 90 degrees respectively;
(7) and (5) placing the prepressing component in the step (6) into an autoclave for curing molding according to the curing temperature and pressure required by the resin.
Example 3:
a preparation method for improving interlayer strength of a fiber metal laminated plate comprises the following steps:
(1) degreasing the surface of the layered titanium alloy by using a surfactant, and using HF and HNO3Mixed acid is used for removing an oxidation film;
(2) putting the titanium alloy into an electrolytic bath by using ammonium fluoride electrolyte with the concentration of 1% for reacting for 5min, taking out the titanium alloy after etching is finished, cleaning the surface of the titanium alloy, and then annealing the titanium alloy;
(3) dissolving epoxy resin glue with a curing agent in chloroform by using chloroform as an organic solvent, wherein the mass fraction of the epoxy resin glue is 7 percent, and forming a resin solution;
(4) uniformly spraying the resin solution in the step (3) on the surface of the titanium alloy treated in the step (2) by using a spray gun under the air pressure of 2-4 bar, wherein the spraying thickness of the glue layer is 4.5mg/cm2；
(5) Putting the titanium alloy sprayed in the step (4) in a 50 ℃ oven for 40min, and fully volatilizing the organic solvent to obtain the rubberized titanium alloy;
(6) laying the titanium alloy and the carbon fiber prepreg in the step (5) according to a laying sequence of [ Ti/0/90/Ti/90/0/Ti ] and pre-compacting, wherein the specific laying structure schematic diagram is shown in figure 1, and two layers of carbon fiber prepregs are laid between two adjacent layers of titanium alloy according to laying angles of 0 degree and 90 degrees respectively;
(7) and (4) placing the prepressing component in the step (6) into a flat vulcanizing machine for compression molding according to the curing temperature and pressure required by the resin.
The fiber metal hybrid structure with high bearing capacity and good impact resistance can be manufactured by the spraying and surface etching processes in the three embodiments, and the specific technical effect of the invention is shown by taking the preparation process of the spraying and surface etching processes in embodiment 1 as an example.
In embodiment 1, the surface of the titanium alloy is etched, the shape of the etched metal surface is as shown in fig. 2, and the metal surface has a porous structure with staggered layers, so that the bonding force with resin can be effectively improved. In example 1, the etched titanium alloy was coated by a spray coating process, and a hybrid laminate was prepared according to the stacking order of [ Ti/0/90/Ti/90/0/Ti ]. Meanwhile, in order to compare the effect difference of other methods in the prior art, three existing processes of glue film + mechanical polishing, spraying + anodic oxidation are used for manufacturing the fiber metal hybrid structure with the same thickness and the same layer, and a short beam shearing test method is used for comparing the interlayer bearing capacity of the laminates. FIG. 3 is a graph of the interlaminar shear strength of hybrid structural laminates from example 1 and three other different methods of manufacture. It can be seen that the laminates made using the adhesive film bonding process have much less load bearing capacity than the shear strength of the samples using spray coating + mechanical grinding due to the thicker adhesive film thickness; the shear strength of the laminated plate with the hybrid structure manufactured by the spraying and surface etching processes is higher than that of the laminated plate manufactured by the spraying and bonding processes and the titanium alloy by mechanical grinding and anodic oxidation, and reaches 59.5 MPa. This result demonstrates that the hybrid fiber metal structure produced by spray coating + surface etching process is superior to other manufacturing schemes in terms of interlaminar load-bearing capacity.
In order to demonstrate the advantage of the invention in terms of impact resistance of the fiber metal hybrid structure, the fiber metal hybrid structure samples manufactured by spraying, mechanical grinding, spraying, anodic oxidation and spraying and surface etching in example 1 were subjected to a low-speed impact test under an energy condition of 60J, and the response of the structure bearing capacity with impact time during the impact process is shown in fig. 4. From the graph, it can be seen that the samples manufactured by spraying + mechanical grinding and spraying + anodizing have sudden load drop at 7.7 ms; the load curve of the sample manufactured by spraying and surface etching in the embodiment 1 is uniform and symmetrical, and the phenomenon of abrupt load drop does not occur. From fig. 5, it can be seen from the graph of the ultrasonic scanning impact area after the impact test of the three samples, that the deformation and delamination areas of the samples using spraying + mechanical polishing and spraying + anodization are much larger than those of the samples manufactured using the spraying + surface etching technology, which indicates that the technology used in the present invention can better maintain the structural integrity, has high damage tolerance and has more excellent impact resistance in the impact test.
In summary, the manufacturing method of the invention uses the metal surface etching combined with the epoxy resin spraying process to obtain the firm and light interface layer of the metal and resin-based fiber composite material, which not only makes the resin fully contact and wet with the metal surface, but also makes the thickness of the interface layer uniform, light and high in strength. The fiber metal hybrid structural member manufactured by the method has higher delamination resistance and impact resistance.
The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention, it should be noted that, for those skilled in the art, many modifications and variations can be made without departing from the technical principle of the present invention, and these modifications and variations should also be regarded as the protection scope of the present invention.

Claims (10)

1. A manufacturing method of an impact-resistant fiber metal hybrid structure for aviation is characterized by comprising the following steps: the method comprises the following steps:
s1, processing the titanium alloy with the cleaned surface by using a surface etching technology, and annealing the metal after etching;
s2, uniformly spraying a resin solution dissolved in an organic solvent on the surface of the titanium alloy treated by the S1, and then drying to fully volatilize the organic solvent to obtain the sized titanium alloy;
s3, alternately laying and pre-compacting the rubberized titanium alloy and the resin-based fiber prepreg according to a preset laying sequence to form a pre-compaction member, and curing and forming the pre-compaction member to form the fiber metal hybrid structure.
2. The manufacturing method according to claim 1, characterized in that: the surface cleaning treatment comprises degreasing the surface of the metal layer and removing an oxidation film.
3. The manufacturing method according to claim 1, characterized in that: the resin-based fiber prepreg is a carbon fiber prepreg or a glass fiber prepreg.
4. The manufacturing method according to claim 1, characterized in that: the surface etching technology is alkali etching or NH4And F, etching by using the solution, wherein the etching pattern is a net structure or a porous array.
5. The manufacturing method according to claim 1, characterized in that: the organic solvent is dichloromethane or chloroform.
6. The manufacturing method according to claim 1, characterized in that: the resin is epoxy resin.
7. The manufacturing method according to claim 1, characterized in that: the spraying thickness of the resin solution on the adhesive layer on the surface of the titanium alloy is controlled to be 2-7 mg/cm2And controlling the spraying pressure to be 2-4 bar.
8. The manufacturing method according to claim 1, characterized in that: the prepressing component is formed by laying and prepressing M layers of rubberized titanium alloy and N layers of resin-based fiber prepreg according to the paving sequence of [ Ti/FRP/Ti/FRP/Ti …/FRP/Ti ], wherein Ti represents titanium alloy, and FRP represents resin-based fiber prepreg; m is more than or equal to 2, and N is more than or equal to 1.
9. The manufacturing method according to claim 1, characterized in that: the curing molding process is die pressing or hot pressing curing, and the used equipment is an autoclave or a flat vulcanizing machine.
10. An aerospace impact-resistant fiber metal hybrid structure made by the method of any one of claims 1-9.
CN202110637069.5A 2021-06-08 2021-06-08 Impact-resistant fiber metal hybrid structure for aviation and manufacturing method thereof Active CN113320244B (en)

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CN103029376A (en) * 2012-12-21 2013-04-10 官宇寰 Metal-fiber composite laminate and manufacturing method thereof

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* Cited by examiner, † Cited by third party
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CN103029376A (en) * 2012-12-21 2013-04-10 官宇寰 Metal-fiber composite laminate and manufacturing method thereof

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