CN106882360B - A kind of connection structure and preparation method for solar powered aircraft main spar - Google Patents
A kind of connection structure and preparation method for solar powered aircraft main spar Download PDFInfo
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- CN106882360B CN106882360B CN201610945625.4A CN201610945625A CN106882360B CN 106882360 B CN106882360 B CN 106882360B CN 201610945625 A CN201610945625 A CN 201610945625A CN 106882360 B CN106882360 B CN 106882360B
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- wrapping layer
- metal flange
- adhesive film
- neck
- carbon fiber
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- 238000002360 preparation method Methods 0.000 title abstract description 8
- 239000002184 metal Substances 0.000 claims abstract description 96
- 229910052751 metal Inorganic materials 0.000 claims abstract description 96
- 229920000049 Carbon (fiber) Polymers 0.000 claims abstract description 46
- 239000004917 carbon fiber Substances 0.000 claims abstract description 46
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 claims abstract description 46
- 239000003292 glue Substances 0.000 claims abstract description 39
- 239000002313 adhesive film Substances 0.000 claims description 69
- 238000000034 method Methods 0.000 claims description 11
- 229910001069 Ti alloy Inorganic materials 0.000 claims description 8
- 239000000956 alloy Substances 0.000 claims description 8
- 239000004744 fabric Substances 0.000 claims description 8
- 238000004519 manufacturing process Methods 0.000 claims description 5
- 238000005086 pumping Methods 0.000 claims description 3
- 239000002131 composite material Substances 0.000 abstract description 9
- 239000007769 metal material Substances 0.000 abstract description 5
- 239000010410 layer Substances 0.000 description 113
- 239000000463 material Substances 0.000 description 12
- 239000000853 adhesive Substances 0.000 description 11
- 230000001070 adhesive effect Effects 0.000 description 11
- 230000000694 effects Effects 0.000 description 3
- 238000010008 shearing Methods 0.000 description 3
- 239000002390 adhesive tape Substances 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 239000003822 epoxy resin Substances 0.000 description 1
- 239000011229 interlayer Substances 0.000 description 1
- 238000000465 moulding Methods 0.000 description 1
- 229920000647 polyepoxide Polymers 0.000 description 1
- 238000003825 pressing Methods 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
- 238000009827 uniform distribution Methods 0.000 description 1
Classifications
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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
- B32—LAYERED PRODUCTS
- B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
- B32B1/00—Layered products having a non-planar shape
- B32B1/08—Tubular products
-
- 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
- B32B33/00—Layered products characterised by particular properties or particular surface features, e.g. particular surface coatings; Layered products designed for particular purposes not covered by another single class
-
- 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
- B32B2260/00—Layered product comprising an impregnated, embedded, or bonded layer wherein the layer comprises an impregnation, embedding, or binder material
- B32B2260/02—Composition of the impregnated, bonded or embedded layer
- B32B2260/021—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
- 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
- B32B2307/558—Impact strength, toughness
-
- 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
- B32B2605/00—Vehicles
- B32B2605/18—Aircraft
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- Aviation & Aerospace Engineering (AREA)
- Moulding By Coating Moulds (AREA)
- Laminated Bodies (AREA)
Abstract
The invention discloses a kind of connection structures and preparation method for solar powered aircraft main spar, wherein the connection structure includes: the first glue film, the first wrapping layer, the second glue film, metal flange, third glue film and the second wrapping layer；Wherein, first glue film is set to carbon fiber pipe for wrapping；The first wrapping layer packet is set to first glue film；The second glue film packet is set to first wrapping layer；The metal flange is sheathed on second glue film；The third glue film packet is set to the neck of the metal flange；The second wrapping layer packet is set to the third glue film.Connection structure for solar powered aircraft main spar of the invention is whole by constituting the first glue film, the first wrapping layer, the second glue film, metal flange, third glue film and the second wrapping layer, metal material is effectively connect with composite material, bonding strength is enhanced, ensure that the quality of solar powered aircraft main spar.
Description
Technical Field
The invention relates to the field of main wing spars of solar airplanes, in particular to a connecting structure for the main wing spars of the solar airplanes and a preparation method of the connecting structure.
Background
Due to the excellent material properties of carbon fiber, the carbon fiber is widely applied to the structural design of the main wing beam of the solar airplane. However, carbon fiber composite materials are brittle materials, and are easily damaged after structural connection is stressed, so that great challenges are provided for connection of composite material structures, particularly connection between main load-bearing composite material structures. The connection problem can be effectively solved by combining the metal material and the composite material, but the combination mode of the two materials at present mainly adopts secondary bonding, and takes the carbon fiber round tube forming the main wing spar of the solar airplane as an example: firstly, processing and manufacturing a carbon fiber round pipe, then processing and manufacturing a metal flange, and finally bonding and molding the metal flange and the carbon fiber round pipe by using structural adhesive. Disadvantages of this process include: because pressure cannot be applied to the metal flange, the requirement on the processing precision of the metal flange and the circular tube is extremely high, and the glue shortage phenomenon is easily generated on a glue joint interface; the secondary bonding adopts normal temperature curing, the curing time of the common epoxy resin structural adhesive at least needs several hours, and the condition of upper-end adhesive shortage is easy to occur due to the fluidity of the structural adhesive in the curing process, and the adhesive strength is inevitably reduced due to the adhesive shortage.
Disclosure of Invention
The technical problem solved by the invention is as follows: compared with the prior art, the connection structure for the main wing beam of the solar airplane and the preparation method are provided, so that the metal flange can be effectively connected with the carbon fiber pipe of the main wing beam of the solar airplane, and the quality of the main wing beam of the solar airplane formed by connecting the carbon fiber pipe with each other through the metal flange is well guaranteed.
The purpose of the invention is realized by the following technical scheme: a connection structure for a solar aircraft main spar, comprising: the adhesive tape comprises a first adhesive film, a first wrapping layer, a second adhesive film, a metal flange, a third adhesive film and a second wrapping layer; the first adhesive film is wrapped on the outer surface of the carbon fiber pipe of the main wing beam of the solar airplane; the first wrapping layer is wrapped on the outer surface of the first adhesive film; the second adhesive film is wrapped on the outer surface of the first wrapping layer; the metal flange is sleeved on the outer surface of the second adhesive film; the third adhesive film is wrapped on the outer surface of the neck of the metal flange; the second wrapping layer is wrapped on the outer surface of the third adhesive film, the number of layers of the second wrapping layer is a plurality of layers, each layer of the second wrapping layer is aligned near one end of the bottom, and the other end, far away from the bottom, of the second wrapping layer is arranged in a staggered mode in sequence.
In the connecting structure for the main wing beam of the solar airplane, the neck of the metal flange is provided with a plurality of holes, and the holes are uniformly distributed along the axial direction and the circumferential direction of the neck.
In the connecting structure for the main wing beam of the solar airplane, the distance between adjacent holes along the axial direction is 10-15 mm; the distance between adjacent holes along the circumferential direction is 20mm-25 mm.
In the connecting structure for the main wing spar of the solar airplane, the neck of the metal flange is provided with the groove along the axial direction of the neck.
In the above connection structure for a main spar of a solar aircraft, the number of the grooves is multiple, and the grooves are uniformly distributed along the circumferential direction of the neck.
In the connecting structure for the main wing beam of the solar airplane, the metal flange is made of a titanium alloy material; the first wrapping layer and the second wrapping layer are both carbon fiber fabric prepreg. .
In the connecting structure for the main wing beam of the solar airplane, the number of the second wrapping layers is at least 3, and the staggered distance of the other ends, far away from the bottom, of the layers of the second wrapping layers is 10-20 mm.
A method of making a connection structure for a solar aircraft main spar, the method comprising the steps of:
the method comprises the following steps: paving a first adhesive film on the outer surface of a carbon fiber pipe of a main wing beam of the solar airplane;
step two: paving a first wrapping layer on the outer surface of the first adhesive film in the first step;
step three: paving a second adhesive film on the outer surface of the first wrapping layer in the step two;
step four: sleeving a metal flange on the outer surface of the second adhesive film in the third step, wherein the diameter of the inner circle of the neck part of the metal flange is equal to the diameter of the outer surface of the product formed in the third step;
step five: sleeving a bag outside the metal flange in the fourth step, pumping air into the bag to ensure that the bag is vacuum, and removing the bag after waiting for the required time;
step six: paving a third adhesive film on the outer surface of the neck of the metal flange in the fifth step;
step seven: laying a second wrapping layer on the outer surface of the third adhesive film in the sixth step, wherein the number of the second wrapping layer is a plurality of layers, the layers of the second wrapping layer are aligned at one end close to the bottom, and the other ends far away from the bottom are sequentially arranged in a staggered manner;
step eight: and curing the product formed after the seventh step in a curing device.
In the preparation method of the connecting structure for the main wing beam of the solar airplane, the first wrapping layer and the second wrapping layer are both carbon fiber fabric prepreg; the curing equipment is an autoclave or an oven.
In the above method for manufacturing the connecting structure for the main wing spar of the solar airplane, the neck of the metal flange is provided with a plurality of holes uniformly distributed along the axial direction and the circumferential direction of the neck and a groove along the axial direction of the neck.
Compared with the prior art, the invention has the following beneficial effects:
(1) according to the solar airplane main wing beam, the first adhesive film, the first wrapping layer, the second adhesive film, the metal flange, the third adhesive film and the second wrapping layer form a whole, so that the metal flange and the carbon fiber pipe of the solar airplane main wing beam can be effectively connected, the connection strength is enhanced, and the carbon fiber pipes are mutually connected through the metal flange to form the solar airplane main wing beam, so that the firmness of the solar airplane main wing beam is good, and the quality is ensured;
(2) the titanium alloy material is adopted as the structural material of the metal flange, so that the connection strength is enhanced;
(3) according to the invention, the metal flange is provided with the holes and the grooves, so that the bonding strength between the composite material consisting of the wrapping layer and the adhesive film and the metal flange is increased;
(4) one end of the second wrapping layer is aligned, and the other end of the second wrapping layer is arranged in a staggered mode, so that stress concentration on the second wrapping layer is reduced.
Drawings
FIG. 1 illustrates a schematic structural view of a connection structure for a main spar of a solar aircraft provided by an embodiment of the invention;
FIG. 2 illustrates a schematic structural view of a connecting structure portion for a main spar of a solar aircraft provided by an illustrative embodiment of the invention;
fig. 3 shows a schematic structural view of a metal flange in a connecting structure for a main spar of a solar aircraft according to an embodiment of the invention.
Detailed Description
The invention is described in further detail below with reference to the accompanying drawings:
fig. 1 shows a schematic structural diagram of a connecting structure for a main spar of a solar aircraft according to an embodiment of the invention. Fig. 2 shows a schematic structural view of a connecting structure part for a main spar of a solar airplane according to an embodiment of the invention. As shown in fig. 1 and 2, the connection structure includes: the carbon fiber tube comprises a carbon fiber tube 1, a first glue film 2, a first wrapping layer 3, a second glue film 4, a metal flange 5, a third glue film 6 and a second wrapping layer 7. In specific implementation, the carbon fiber pipe 1 is a circular pipe, and the carbon fiber unidirectional tape prepreg is wound and laid by using a pipe coiling process and cured. The first wrapping layer 3 and the second wrapping layer 7 adopt carbon fiber fabric prepreg. The metal flange 5 is made of titanium alloy material, so that the connection strength with carbon fiber is enhanced. The first adhesive film 2 and the second adhesive film 4 are made of J-272 adhesive films. Wherein,
the first glue film 2 is wrapped on the carbon fiber tube 1. Specifically, the first adhesive film 2 is tightly laid on the outer surface of the carbon fiber tube 1 and is bonded with the carbon fiber tube 1 through the adhesive force of the first adhesive film 2.
The first wrapping layer 3 is wrapped on the first adhesive film 2. Specifically, the first wrapping layer 3 is tightly paved on the outer surface of the first adhesive film 2 by adopting carbon fiber fabric prepreg, and the first wrapping layer 3 is tightly combined with the first adhesive film 2 by the adhesion force of the first wrapping layer 3 and the adhesion force of the first adhesive film 2.
The second glue film 4 is wrapped on the first wrapping layer 3. Specifically, the second glue film 4 is tightly paved on the outer surface of the first wrapping layer 3, and the first wrapping layer 3 and the second glue film 4 are tightly combined through the adhesive force of the first wrapping layer 3 and the adhesive force of the second glue film 4.
The metal flange 5 is sleeved on the second adhesive film 4. Specifically, the metal flange 5 is sleeved on the carbon fiber tube 1 paved with the first adhesive film 2, the first wrapping layer 3 and the second adhesive film 4, and preferably, the left end of the metal flange 5 is flush with the left end of the carbon fiber tube 1. The inner diameter of the metal flange 5 is consistent with the outer diameter of the carbon fiber pipe 1 paved with the first glue film 2, the first wrapping layer 3 and the second glue film 4, so that the metal flange 5 is in close contact with the second glue film 4.
The third glue film 6 is wrapped around the neck 51 of the metal flange 5. Specifically, the third adhesive film 6 is tightly laid on the outer surface of the neck 51 of the metal flange 5, and the metal flange 5 is tightly connected with the third adhesive film 6 by the adhesive force of the third adhesive film 6.
The second wrapping layer 7 is wrapped on the third adhesive film 6, wherein the number of layers of the second wrapping layer 7 is a plurality of layers, each layer of the second wrapping layer 7 is aligned at one end close to the bottom 52, and the other end far away from the bottom 52 is sequentially arranged in a staggered manner. Specifically, the second wrapping layer 7 is tightly paved on the outer surface of the third glue film 6, and the second wrapping layer 7 is tightly connected with the third glue film 6 through the adhesive force of the third glue film 6. The number of layers of the second wrapping layer 7 is multiple, wherein, the layers are aligned at one end close to the bottom 52, and the other ends are arranged in a staggered way in sequence to form a ladder shape. Thereby reduced metal flange 5 to the stress concentration of second parcel layer 7, protected second parcel layer 7 under the circumstances of having guaranteed joint strength. Furthermore, the number of layers of the second wrapping layer 7 is at least 3, and the staggered distance of the other ends of the layers of the second wrapping layer 7, which are far away from the bottom 52, is 10-20 mm, so that the effect of protecting the second wrapping layer 7 under the condition of ensuring the connection strength is more obvious.
In the embodiment, the first adhesive film, the first wrapping layer, the second adhesive film, the metal flange, the third adhesive film and the second wrapping layer form a whole, so that the metal material and the composite material can be effectively connected, and the connection strength is enhanced; the titanium alloy material is adopted as the structural material of the metal flange, so that the connection strength is enhanced; through second wrapping up the layer one end and align and the other end staggers in proper order and arranges and has reduced the stress concentration of metal flange to the second wrapping up layer, the effectual second wrapping up layer that has protected has increased joint strength.
Fig. 3 shows a schematic structural view of a metal flange in a connecting structure for a main spar of a solar aircraft according to an embodiment of the invention. As shown in fig. 3, the neck portion 51 of the metal flange 5 is provided with a plurality of holes, and the plurality of holes are uniformly distributed along the axial direction and the circumferential direction of the neck portion 51. Specifically, a plurality of holes are uniformly distributed along the axial direction and the circumferential direction of the neck 51, and the thicknesses of the second adhesive film 4 and the third adhesive film 6 are thinner, so that the first wrapping layer 3 and the part of the corresponding hole of the second adhesive film 4 are embedded into the corresponding hole of the neck 51, and the shearing strength between the metal flange 5 and the first wrapping layer 3 is increased. Due to the relatively thin thickness of the third glue film 6, the second wrapping layer 7 is embedded in the hole of the corresponding neck 51 together with the corresponding hole of the third glue film 6, and the shear strength between the metal flange 5 and the second wrapping layer 7 is increased.
In the implementation, the distance between the adjacent holes along the axial direction is 10mm-15 mm; the distance between adjacent holes along the circumferential direction is 20mm-25 mm. So that the effect of increasing the shear strength between the metal flange 5 and the first wrapping layer 3 and the effect of increasing the shear strength between the metal flange 5 and the second wrapping layer 7 are more remarkable.
In the above embodiment, the neck portion 51 of the metal flange 5 is provided with the groove 53 along the axial direction of the neck portion 51. Utilize groove 53 can extrude the unnecessary material that soaks between metal flange 5 and the carbon fiber pipe 1 and the unnecessary material that soaks between metal flange 5 and the second wrapping layer 7, avoid appearing the space between metal flange 5 and first wrapping layer 3, the second wrapping layer 7, increase interlaminar strength.
Further, the number of the grooves 53 is plural, and the plural grooves 53 are uniformly distributed along the circumferential direction of the neck portion 51. The connection strength of the metal flange 5 and the first wrapping layer 3 and the second wrapping layer 7 is further enhanced through uniform distribution.
According to the connecting structure for the main wing beam of the solar airplane, the first adhesive film, the first wrapping layer, the second adhesive film, the metal flange, the third adhesive film and the second wrapping layer form a whole, so that a metal material and a composite material can be effectively connected, the connecting strength is enhanced, further, the main wing beam of the solar airplane is formed by connecting the carbon fiber pipes through the metal flange, the firmness of the main wing beam of the solar airplane is good, and the quality is well guaranteed; the titanium alloy material is adopted as the structural material of the metal flange, so that the connection strength is enhanced; the one end through the second parcel layer aligns and the other end staggers in proper order arranges and has reduced the stress concentration of metal flange to the second parcel layer, and the effectual second parcel layer that has protected has increased joint strength.
The invention also provides a preparation method of the connecting structure for the main wing spar of the solar airplane, which comprises the following steps:
the method comprises the following steps: paving a first adhesive film 2 on the outer surface of a carbon fiber pipe 1 of a main wing beam of the solar airplane;
step two: paving a first wrapping layer 3 on the outer surface of the first adhesive film 2 in the first step;
step three: paving a second adhesive film 4 on the outer surface of the first wrapping layer 3 in the step two;
step four: sleeving a metal flange 5 on the outer surface of the second adhesive film 4 in the third step, wherein the inner circle diameter of the neck part 51 of the metal flange 5 is equal to the diameter of the outer surface of the product formed in the third step;
step five: sleeving a bag outside the metal flange 5 in the fourth step, pumping air into the bag to ensure that the bag is vacuum, and removing the bag after waiting for the required time;
step six: paving a third glue film 6 on the outer surface of the neck part 51 of the metal flange 5 in the fifth step;
step seven: laying a second wrapping layer 7 on the outer surface of the third glue film 6 in the sixth step;
step eight: and curing the product formed after the seventh step in a curing device.
In the second step, the first wrapping layer 3 adopts carbon fiber fabric prepreg.
In the fourth step, the metal flange 5 is made of a titanium alloy material, the metal flange 5 is sleeved on the carbon fiber tube 1 paved with the first adhesive film 2, the first wrapping layer 3 and the second adhesive film 4, and preferably, the left end of the metal flange 5 is flush with the left end of the carbon fiber tube 1. The inner diameter of the metal flange 5 is consistent with the outer diameter of the carbon fiber pipe 1 paved with the first glue film 2, the first wrapping layer 3 and the second glue film 4, so that the metal flange 5 is in close contact with the second glue film 4.
In the fifth step, the connection strength between the metal flange 5 and the carbon fiber tube 1 paved with the first adhesive film 2, the first wrapping layer 3 and the second adhesive film 4 is enhanced by using the vacuumized bag. The neck 51 of the metal flange 5 is provided with a plurality of holes uniformly distributed along the axial direction and the circumferential direction of the neck 51, and the neck 51 of the metal flange 5 is further provided with a groove along the axial direction of the neck 51. So that during the pressing with the vacuum bag, the first wrapping 3 with the corresponding hole portion of the second glue film 4 is embedded in the corresponding hole of the neck 51, increasing the shear strength between the metal flange 5 and the first wrapping 3. Utilize the groove can extrude the unnecessary material that soaks between metal flange 5 and the carbon fiber pipe 1, avoid appearing the space between metal flange 5 and the first wrapping layer 3, increase interlaminar strength.
In step seven, the second wrapping layer 7 adopts carbon fiber fabric prepreg. The number of layers of the second wrapping layer 7 is a plurality of layers, and the second wrapping layer 7 forms a step-shaped structure. Specifically, the number of the second wrapping layers 7 is multiple, wherein one end of each layer is aligned, and the other end is stepped. Thereby reducing the stress concentration of the metal flange 5 on the second wrapping layer 7 and protecting the second wrapping layer 7.
In the eighth step, the product formed after the seventh step is cured in an autoclave or an oven, and in the curing process, the product is under the action of pressure, so that the second wrapping layer 7 and the part of the corresponding hole of the third adhesive film 6 are embedded into the hole of the corresponding neck 51, the shearing strength between the metal flange 5 and the second wrapping layer 7 is increased, meanwhile, the shearing strength between the metal flange 5 and the first wrapping layer 3 is increased again, in addition, the groove can be used for extruding redundant impregnating materials between the metal flange 5 and the carbon fiber tube 1 and redundant impregnating materials between the metal flange 5 and the second wrapping layer 7, gaps are prevented from appearing between the metal flange 5 and the first wrapping layer 3 and between the metal flange 5 and the second wrapping layer 7, and the interlayer strength is increased.
According to the preparation method of the connecting structure for the main wing beam of the solar airplane, the first adhesive film, the first wrapping layer, the second adhesive film, the metal flange, the third adhesive film and the second wrapping layer form a whole, so that the metal material and the composite material can be effectively connected, the connecting strength is enhanced, and further, the main wing beam of the solar airplane is formed by connecting the carbon fiber pipes through the metal flange, so that the firmness of the main wing beam of the solar airplane is good, and the quality is well guaranteed; the titanium alloy material is adopted as the structural material of the metal flange, so that the connection strength is enhanced; one end of the second wrapping layer is aligned, and the other end of the second wrapping layer is arranged in a staggered mode in sequence, so that stress concentration of the metal flange on the second wrapping layer is reduced, and the second wrapping layer is effectively protected.
The above-described embodiments are merely preferred embodiments of the present invention, and general changes and substitutions by those skilled in the art within the technical scope of the present invention are included in the protection scope of the present invention.
Claims (8)
1. A connection structure for a solar aircraft main spar, comprising: the carbon fiber pipe comprises a carbon fiber pipe (1), a first adhesive film (2), a first wrapping layer (3), a second adhesive film (4), a metal flange (5), a third adhesive film (6) and a second wrapping layer (7); wherein,
the first adhesive film (2) is wrapped on the outer surface of the carbon fiber pipe (1) of the main wing beam of the solar airplane;
the first wrapping layer (3) is wrapped on the outer surface of the first adhesive film (2);
the second adhesive film (4) is wrapped on the outer surface of the first wrapping layer (3);
the metal flange (5) is sleeved on the outer surface of the second adhesive film (4);
the third adhesive film (6) is wrapped on the outer surface of the neck (51) of the metal flange (5);
the second wrapping layer (7) is wrapped on the outer surface of the third glue film (6), the number of layers of the second wrapping layer (7) is a plurality of layers, one end, close to the bottom (52), of each layer of the second wrapping layer (7) is aligned, and the other end, far away from the bottom (52), of each layer is sequentially arranged in a staggered mode;
the metal flange is characterized in that a plurality of holes are formed in the neck (51) of the metal flange (5), and the holes are uniformly distributed along the axial direction and the circumferential direction of the neck (51).
2. A connection structure for a solar aircraft main spar according to claim 1, wherein: the distance between the adjacent holes along the axial direction is 10mm-15 mm; the distance between adjacent holes along the circumferential direction is 20mm-25 mm.
3. A connection structure for a solar aircraft main spar according to claim 1, wherein: the neck (51) of the metal flange (5) is provided with a groove (53) along the axial direction of the neck (51).
4. A connection structure for a solar aircraft main spar according to claim 3, wherein: the number of the grooves (53) is multiple, and the multiple grooves (53) are uniformly distributed along the circumferential direction of the neck part (51).
5. A connection structure for a solar aircraft main spar according to claim 1, wherein: the metal flange (5) is made of a titanium alloy material; the first wrapping layer (3) and the second wrapping layer (7) are both carbon fiber fabric prepreg.
6. A connection structure for a solar aircraft main spar according to claim 1, wherein: the number of layers of the second wrapping layer (7) is at least 3, and the staggered distance of the other ends of the layers of the second wrapping layer (7) far away from the bottom (52) is 10-20 mm.
7. A method for making a connection structure for a solar aircraft main spar, characterized in that it comprises the following steps:
the method comprises the following steps: paving a first adhesive film (2) on the outer surface of a carbon fiber pipe (1) of a main wing beam of the solar airplane;
step two: paving a first wrapping layer (3) on the outer surface of the first adhesive film (2) in the first step;
step three: paving a second adhesive film (4) on the outer surface of the first wrapping layer (3) in the second step;
step four: sleeving a metal flange (5) on the outer surface of the second adhesive film (4) in the third step, wherein the inner circle diameter of a neck part (51) of the metal flange (5) is equal to the diameter of the outer surface of the product formed in the third step;
step five: sleeving a bag outside the metal flange (5) in the fourth step, pumping air into the bag to ensure that the bag is vacuum, and removing the bag after waiting for the required time;
step six: paving a third glue film (6) on the outer surface of the neck (51) of the metal flange (5) in the fifth step;
step seven: laying a second wrapping layer (7) on the outer surface of the third adhesive film (6) in the sixth step, wherein the number of the second wrapping layer (7) is multiple, each layer of the second wrapping layer (7) is aligned at one end close to the bottom (52), and the other ends far away from the bottom (52) are sequentially arranged in a staggered manner;
step eight: curing the product formed after the seventh step in a curing device;
the metal flange is characterized in that the neck part (51) of the metal flange (5) is provided with a plurality of holes which are uniformly distributed along the axial direction and the circumferential direction of the neck part (51) and a groove along the axial direction of the neck part (51).
8. Method for the production of a connection structure for a solar aircraft main spar according to claim 7, characterized in that: the first wrapping layer (3) and the second wrapping layer (7) are both carbon fiber fabric prepreg; the curing equipment is an autoclave or an oven.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201610945625.4A CN106882360B (en) | 2016-11-02 | 2016-11-02 | A kind of connection structure and preparation method for solar powered aircraft main spar |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201610945625.4A CN106882360B (en) | 2016-11-02 | 2016-11-02 | A kind of connection structure and preparation method for solar powered aircraft main spar |
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| EP1156183A1 (en) * | 2000-05-17 | 2001-11-21 | Bauer Spezialtiefbau GmbH | Double-walled drill pipe |
| CN201902713U (en) * | 2010-12-16 | 2011-07-20 | 赵欣荣 | Novel composite material pipe structure |
| CN102278342A (en) * | 2011-06-28 | 2011-12-14 | 北京航空航天大学 | Method for bonding carbon fiber pipe with metal flanges internally and externally |
| CN102815210A (en) * | 2012-08-30 | 2012-12-12 | 同济大学 | Composite-material automobile transmission shaft formed by pulling, squeezing and winding and preparation method thereof |
| CN104454944A (en) * | 2014-09-25 | 2015-03-25 | 武汉理工大学 | Ribbed woven winding carbon fiber composite transmission shaft |
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Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP1156183A1 (en) * | 2000-05-17 | 2001-11-21 | Bauer Spezialtiefbau GmbH | Double-walled drill pipe |
| CN201902713U (en) * | 2010-12-16 | 2011-07-20 | 赵欣荣 | Novel composite material pipe structure |
| CN102278342A (en) * | 2011-06-28 | 2011-12-14 | 北京航空航天大学 | Method for bonding carbon fiber pipe with metal flanges internally and externally |
| CN102815210A (en) * | 2012-08-30 | 2012-12-12 | 同济大学 | Composite-material automobile transmission shaft formed by pulling, squeezing and winding and preparation method thereof |
| CN104454944A (en) * | 2014-09-25 | 2015-03-25 | 武汉理工大学 | Ribbed woven winding carbon fiber composite transmission shaft |
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