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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- glue film
- wrapping layer
- metal flange
- solar powered
- neck
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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
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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
- B32B1/00—Layered products having a general shape other than plane
- 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
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 present invention relates to solar powered aircraft main spar field more particularly to a kind of connections for solar powered aircraft main spar
Structure and preparation method.
Background technique
Due to the superior material property of carbon fiber, it has been widely used among the design of solar powered aircraft main wing girder construction.
But carbon fibre composite category fragile material, it is easily damaged after structure connects stress, for composite structure
Connection, the especially connection between primary load bearing composite structure are a greatly challenges.And by metal material and composite material knot
Connectivity problem can be efficiently solved by closing use, but the combination of mesh first two material is mostly based on secondary bonding, with group
For the carbon fiber round tube of solar powered aircraft main spar: being fabricated first to carbon fiber round tube, then to metal method
Orchid is fabricated, and finally utilizes structure glue by metal flange and carbon fiber round tube bonding forming.This process lacks
Point includes:, the splicing high to the requirement on machining accuracy of metal flange and round tube due to that can not apply pressure to metal flange
Interface easily generates starved phenomenon；Second bonding uses normal temperature cure, and the curing time of general such epoxy resin structural adhesive is extremely
A few hours are needed less, and in the curing process since easily there is a situation where upper end starveds for the mobility of structure glue, and starved is necessarily led
Cause adhesive strength decline.
Summary of the invention
Technical problem solved by the present invention is compared with the prior art, providing a kind of for solar powered aircraft main spar
Connection structure and preparation method, metal flange is effectively connect with the carbon fiber pipe of solar powered aircraft main spar,
So that the quality for the solar powered aircraft main spar formed that is connected to each other by carbon fiber pipe by metal flange obtains very
Good guarantee.
The object of the invention is achieved by the following technical programs: a kind of connection knot for solar powered aircraft main spar
Structure, comprising: the first glue film, the first wrapping layer, the second glue film, metal flange, third glue film and the second wrapping layer；Wherein, described
First glue film packet is set to the carbon fiber pipe outer surface of solar powered aircraft main spar；The first wrapping layer packet is set to first glue
Film outer surface；The second glue film packet is set to first wrapping layer outer surface；The metal flange is sheathed on second glue
Film outer surface；The third glue film packet is set to the outer surface of the metal flange neck；The second wrapping layer packet is set to described
Third glue film outer surface, wherein the number of plies of second wrapping layer is several layers, and each layer of second wrapping layer is the bottom of close to
The alignment of portion one end, the other end far from bottom stagger successively arrangement.
In the above-mentioned connection structure for solar powered aircraft main spar, the neck of the metal flange offers several
Hole, several holes are uniformly distributed along the axial and circumferential of the neck.
In the above-mentioned connection structure for solar powered aircraft main spar, the spacing in axially adjacent hole is 10mm-15mm；Edge
The spacing in circumferentially-adjacent hole is 20mm-25mm.
In the above-mentioned connection structure for solar powered aircraft main spar, the axis of the neck of the metal flange along the neck
To offering slot.
In the above-mentioned connection structure for solar powered aircraft main spar, the quantity of the slot be it is multiple, multiple slots are described in
The circumferential direction of neck is uniformly distributed.
In the above-mentioned connection structure for solar powered aircraft main spar, the metal flange is titanium alloy material；Described
One wrapping layer and second wrapping layer are carbon fibre fabric prepreg.
In the above-mentioned connection structure for solar powered aircraft main spar, the number of plies of second wrapping layer is at least 3 layers, institute
It is 10mm-20mm that each layer of the second wrapping layer, which is stated, far from the distance that the other end of bottom is staggered.
A kind of preparation method of the connection structure for solar powered aircraft main spar, the described method comprises the following steps:
Step 1: the first glue film is laid in the outer surface of the carbon fiber pipe of solar powered aircraft main spar；
Step 2: the outer surface of the first glue film in step 1 is laid with the first wrapping layer；
Step 3: the outer surface of the first wrapping layer in step 2 is laid with the second glue film；
Step 4: the outer surface of the second glue film in step 3 covers upper metal flange, wherein the neck of the metal flange
Diameter of the interior circular diameter in portion equal to the outer surface for forming product after step 3；
Step 5: being arranged bag outside the metal flange in step 4, then makes to air is taken out in bag for vacuum, etc.
After required time, bag is removed；
Step 6: the outer surface of the neck of the metal flange in step 5 is laid with third glue film；
Step 7: the outer surface of the third glue film in step 6 is laid with the second wrapping layer, the layer of second wrapping layer
Number is several layers, and each layer of second wrapping layer is aligned close to bottom end, and the other end far from bottom staggers successively row
Cloth；
Step 8: the product formed after step 7 is solidified in curing apparatus.
In the preparation method of the above-mentioned connection structure for solar powered aircraft main spar, first wrapping layer and described
Two wrapping layers are carbon fibre fabric prepreg；The curing apparatus is autoclave or baking oven.
In the preparation method of the above-mentioned connection structure for solar powered aircraft main spar, the neck of the metal flange is opened up
There are several equally distributed holes of axial and circumferential along the neck and the axial slot along the neck.
Compared with prior art, the present invention has the following advantages:
(1), the present invention is by by the first glue film, the first wrapping layer, the second glue film, metal flange, third glue film and second
Wrapping layer constitutes entirety, and metal flange is effectively connect with the carbon fiber pipe of solar powered aircraft main spar, is enhanced
Bonding strength, and each Carbon fibe pipe is connected to each other by metal flange and forms the main spar of solar powered aircraft, so that too
The fastness of positive energy aircraft main spar is good, and quality is guaranteed；
(2), the present invention uses structural material of the titanium alloy material as metal flange, enhances bonding strength；
(3), the present invention is by offering hole and slot in metal flange, increase the composite material of wrapping layer and glue film composition with
The glue-joint strength of metal flange；
(4), second wrapping layer one end alignment in the present invention and the other end stagger successively arrangement, reduce and wrap up second
The stress of layer is concentrated.
Detailed description of the invention
Fig. 1 shows the structural representation of the connection structure provided in an embodiment of the present invention for solar powered aircraft main spar
Figure；
Fig. 2 shows the connection structure part-structures for solar powered aircraft main spar that bright embodiment of the invention provides
Schematic diagram；
Fig. 3 is shown in the connection structure for solar powered aircraft main spar that bright embodiment of the invention provides, metal
The structural schematic diagram of flange.
Specific embodiment
The present invention is described in further detail below in conjunction with the accompanying drawings:
Fig. 1 shows the structural representation of the connection structure provided in an embodiment of the present invention for solar powered aircraft main spar
Figure.Fig. 2 shows the connection structure part-structure signals for solar powered aircraft main spar that bright embodiment of the invention provides
Figure.As depicted in figs. 1 and 2, which includes: carbon fiber pipe 1, the first glue film 2, the first wrapping layer 3, the second glue film 4, gold
Belong to flange 5, third glue film 6 and the second wrapping layer 7.When it is implemented, carbon fiber pipe 1 is round tube, carbon fiber one-way band is presoaked
Material, which is twined, to be laid with and is solidified using pipe rolling technique.First wrapping layer 3 and the second wrapping layer 7 use carbon fibre fabric prepreg.
The material of metal flange 5 is titanium alloy material, is enhanced and carbon fiber bonding strength.The material of first glue film 2 and the second glue film 4
It is J-272 glue film.Wherein,
First glue film 2 packet is set to carbon fiber pipe 1.Specifically, the first glue film 2 is closely layed in the outer surface of carbon fiber pipe 1,
And it is bonded together by the viscous force of the first glue film 2 with carbon fiber pipe 1.
First wrapping layer 3 packet is set to the first glue film 2.Specifically, the first wrapping layer 3 is close using carbon fibre fabric prepreg
It is layed in the outer surface of the first glue film 2, the first wrapping layer 3 is made by the viscous force of the first wrapping layer 3 and the viscous force of the first glue film 2
It combines closely with the first glue film 2.
Second glue film 4 packet is set to the first wrapping layer 3.Specifically, the second glue film 4 is closely layed in the outer of the first wrapping layer 3
Surface makes the first wrapping layer 3 and the second glue film 4 closely knot by the viscous force of the first wrapping layer 3 and the viscous force of the second glue film 4
It closes.
Metal flange 5 is sheathed on the second glue film 4.Specifically, metal flange 5, which is set in, is covered with the first glue film 2, first package
On the carbon fiber pipe 1 of layer 3 and the second glue film 4, it is preferred that the left end of metal flange 5 is flushed with the left end of carbon fiber pipe 1.Metal
The internal diameter of flange 5 is consistent with the outer diameter of carbon fiber pipe 1 for being covered with the first glue film 2, the first wrapping layer 3 and the second glue film 4, thus
So that metal flange 5 and the second glue film 4 are closely in contact.
The packet of third glue film 6 is set to the neck 51 of metal flange 5.Specifically, third glue film 6 is closely layed in metal flange 5
Neck 51 outer surface, metal flange 5 is closely connect with third glue film 6 by the viscous force of third glue film 6.
Second wrapping layer 7 packet is set to third glue film 6, wherein the number of plies of the second wrapping layer 7 is several layers, the second wrapping layer 7
Each layer be aligned close to 52 one end of bottom, stagger successively arrangement far from the other end of bottom 52.Specifically, the second wrapping layer 7
It is closely layed in the outer surface of third glue film 6, is closely connected the second wrapping layer 7 with third glue film 6 by the viscous force of third glue film 6
It connects.The number of plies of second wrapping layer 7 is multilayer, wherein each layer is aligned in one end close to bottom 52, and the other end staggers successively arrangement
It is stepped.It concentrates to reduce metal flange 5 to the stress of the second wrapping layer 7, is protected in the case where ensure that bonding strength
The second wrapping layer 7 is protected.Further, the number of plies of the second wrapping layer 7 is at least 3 layers, and each layer of the second wrapping layer 7 is far from bottom
The distance that 52 other end is staggered is 10mm-20mm, so that protecting the second wrapping layer 7 in the case where ensure that bonding strength
Effect is more obvious.
The present embodiment is by wrapping the first glue film, the first wrapping layer, the second glue film, metal flange, third glue film and second
Covering layer constitutes entirety, and metal material is effectively connect with composite material, enhances bonding strength；Using titanium alloy material
Expect the structural material as metal flange, enhances bonding strength；It is staggered successively by the alignment of second wrapping layer one end and the other end
Arrangement reduces metal flange and concentrates to the stress of the second wrapping layer, effectively protects the second wrapping layer, and it is strong to increase connection
Degree.
Fig. 3 is shown in the connection structure for solar powered aircraft main spar that bright embodiment of the invention provides, metal
The structural schematic diagram of flange.As shown in figure 3, the neck 51 of metal flange 5 offers several holes, several holes are along neck 51
Axial and circumferential are uniformly distributed.Specifically, several holes are uniformly distributed along the axial and circumferential of neck 51, due to the second glue film 4
Thickness with third glue film 6 is than relatively thin, so that the first wrapping layer 3 is partially submerged into phase together with the corresponding aperture of second glue film 4
The hole of corresponding neck 51 increases shear strength between metal flange 5 and the first wrapping layer 3.Due to the thickness ratio of third glue film 6
It is relatively thin, so that the hole that is partially submerged into corresponding neck 51 of second wrapping layer 7 together with the corresponding aperture of third glue film 6, increases
Add shear strength between metal flange 5 and the second wrapping layer 7.
In above-mentioned implementation, the spacing in axially adjacent hole is 10mm-15mm；The spacing in circumferentially adjacent hole is 20mm-
25mm.So that increasing between metal flange 5 and the first wrapping layer 3 effect of shear strength and increase metal flange 5 and the
The effect of shear strength is more significant between two wrapping layers 7.
In above-mentioned implementation, the neck 51 of metal flange 5 offers slot 53 along the axial direction of neck 51.It can will be golden using slot 53
The extra leaching material belonged between the extra leaching material and metal flange 5 and the second wrapping layer 7 between flange 5 and carbon fiber pipe 1 squeezes out, and keeps away
Exempt from occur gap between metal flange 5 and the first wrapping layer 3, the second wrapping layer 7, increases interlaminar strength.
Further, the quantity of slot 53 is multiple, and multiple slots 53 are uniformly distributed along the circumferential direction of neck 51.By being uniformly distributed
Further enhance the bonding strength of metal flange 5 and the first wrapping layer 3, the second wrapping layer 7.
Connection structure for solar powered aircraft main spar of the invention is by by the first glue film, the first wrapping layer, second
Glue film, metal flange, third glue film and the second wrapping layer constitute entirety, and metal material is effectively connected with composite material
It connects, enhances bonding strength, further, form solar energy since each Carbon fibe pipe is connected to each other by metal flange and fly
The main spar of machine, so that the fastness of solar powered aircraft main spar is good, quality is guaranteed very well；Using titanium alloy material conduct
The structural material of metal flange enhances bonding strength；Arrangement is staggered successively by one end alignment of the second wrapping layer and the other end
It reduces metal flange to concentrate the stress of the second wrapping layer, effectively protects the second wrapping layer, increase bonding strength.
The present invention also provides a kind of preparation method of connection structure for solar powered aircraft main spar, this method includes
Following steps:
Step 1: the first glue film 2 is laid in the outer surface of the carbon fiber pipe 1 of solar powered aircraft main spar；
Step 2: the outer surface of the first glue film 2 in step 1 is laid with the first wrapping layer 3；
Step 3: the outer surface of the first wrapping layer 3 in step 2 is laid with the second glue film 4；
Step 4: the outer surface of the second glue film 4 in step 3 covers upper metal flange 5, wherein the neck of metal flange 5
Diameter of the interior circular diameter in portion 51 equal to the outer surface for forming product after step 3；
Step 5: being arranged bag outside the metal flange 5 in step 4, then makes to air is taken out in bag for vacuum,
After the time required to waiting, bag is removed；
Step 6: the outer surface of the neck 51 of the metal flange 5 in step 5 is laid with third glue film 6；
Step 7: the outer surface of the third glue film 6 in step 6 is laid with the second wrapping layer 7；
Step 8: the product formed after step 7 is solidified in curing apparatus.
In step 2, the first wrapping layer 3 uses carbon fibre fabric prepreg.
In step 4, metal flange 5 uses titanium alloy material, and metal flange 5, which is set in, is covered with the first glue film 2, first
On the carbon fiber pipe 1 of wrapping layer 3 and the second glue film 4, it is preferred that the left end of metal flange 5 is flushed with the left end of carbon fiber pipe 1.
The internal diameter of metal flange 5 is consistent with the outer diameter of carbon fiber pipe 1 for being covered with the first glue film 2, the first wrapping layer 3 and the second glue film 4,
So that metal flange 5 and the second glue film 4 are closely in contact.
In step 5, makes metal flange 5 using the bag vacuumized and be covered with the first glue film 2, the first wrapping layer 3 and the
Bonding strength enhancing between the carbon fiber pipe 1 of two glue films 4.The neck 51 of metal flange 5 offer along neck 51 axial direction and
Several circumferential equally distributed holes, and the axial slot along neck 51 is also provided in the neck of metal flange 5 51.To
Using vacuum bag compress during so that the first wrapping layer 3 together with the second glue film 4 corresponding aperture be partially submerged into it is corresponding
Neck 51 hole, increase shear strength between metal flange 5 and the first wrapping layer 3.It can be by metal flange 5 and carbon using slot
Extra leaching material between fibre pipe 1 squeezes out, and avoids occurring gap between metal flange 5 and the first wrapping layer 3, and it is strong to increase interlayer
Degree.
In step 7, the second wrapping layer 7 uses carbon fibre fabric prepreg.The number of plies of second wrapping layer 7 is several layers,
Second wrapping layer 7 forms step structure.Specifically, the number of plies of the second wrapping layer 7 is multilayer, wherein every layer of one end alignment,
The other end is stepped.It is concentrated to reduce metal flange 5 to the stress of the second wrapping layer 7, protects the second wrapping layer 7.
In step 8, solidified in autoclave or baking oven, it, also can be to being formed after step 7 during cured
Product play the role of a pressure so that the second wrapping layer 7 together with third glue film 6 corresponding aperture be partially submerged into it is corresponding
Neck 51 hole, increase shear strength between metal flange 5 and the second wrapping layer 7, while also again increasing metal flange 5
The shear strength between the first wrapping layer 3, moreover, can be by the extra leaching material between metal flange 5 and carbon fiber pipe 1 using slot
Extra leaching material between metal flange 5 and the second wrapping layer 7 squeezes out, and metal flange 5 and the first wrapping layer 3, second is avoided to wrap
Occur gap between covering layer 7, increases interlaminar strength.
The preparation method of connection structure for solar powered aircraft main spar of the invention is by wrapping the first glue film, first
Covering layer, the second glue film, metal flange, third glue film and the second wrapping layer constitute whole, enable metal material and composite material
Enough effective connections, enhance bonding strength, further, since each Carbon fibe pipe is connected to each other group by metal flange
At the main spar of solar powered aircraft, so that the fastness of solar powered aircraft main spar is good, quality is guaranteed very well；It is closed using titanium
Structural material of the golden material as metal flange enhances bonding strength；By the alignment of one end of the second wrapping layer and the other end according to
The secondary arrangement that is staggered reduces metal flange and concentrates to the stress of the second wrapping layer, effectively protects the second wrapping layer.
Embodiment described above is the present invention more preferably specific embodiment, and those skilled in the art is in this hair
The usual variations and alternatives carried out in bright technical proposal scope should be all included within the scope of the present invention.
Claims (8)
1. a kind of connection structure for solar powered aircraft main spar, characterized by comprising: carbon fiber pipe (1), the first glue film
(2), the first wrapping layer (3), the second glue film (4), metal flange (5), third glue film (6) and the second wrapping layer (7)；Wherein,
First glue film (2) packet is set to carbon fiber pipe (1) outer surface of solar powered aircraft main spar；
First wrapping layer (3) packet is set to the first glue film (2) outer surface；
Second glue film (4) packet is set to the first wrapping layer (3) outer surface；
The metal flange (5) is sheathed on the second glue film (4) outer surface；
Third glue film (6) packet is set to the outer surface of the metal flange (5) neck (51)；
Second wrapping layer (7) packet is set to third glue film (6) outer surface, wherein the number of plies of second wrapping layer (7)
For several layers, each layer of second wrapping layer (7) is aligned close to bottom (52) one end, the other end far from bottom (52) according to
The secondary arrangement that is staggered；
The neck (51) of the metal flange (5) offers several holes, axial direction and week of several holes along the neck (51)
To being uniformly distributed.
2. the connection structure according to claim 1 for solar powered aircraft main spar, it is characterised in that: axially adjacent
The spacing in hole is 10mm-15mm；The spacing in circumferentially adjacent hole is 20mm-25mm.
3. the connection structure according to claim 1 for solar powered aircraft main spar, it is characterised in that: the metal method
The neck (51) of blue (5) offers slot (53) along the axial direction of the neck (51).
4. the connection structure according to claim 3 for solar powered aircraft main spar, it is characterised in that: the slot (53)
Quantity be it is multiple, multiple slots (53) are uniformly distributed along the circumferential direction of the neck (51).
5. the connection structure according to claim 1 for solar powered aircraft main spar, it is characterised in that: the metal method
Blue (5) are titanium alloy material；First wrapping layer (3) and second wrapping layer (7) are carbon fibre fabric prepreg.
6. the connection structure according to claim 1 for solar powered aircraft main spar, it is characterised in that: second packet
The number of plies of covering layer (7) is at least 3 layers, the distance that each layer of second wrapping layer (7) is staggered far from the other end of bottom (52)
For 10mm-20mm.
7. a kind of preparation method of the connection structure for solar powered aircraft main spar, is characterized in that, the method includes following
Step:
Step 1: the first glue film (2) are laid in the outer surface of the carbon fiber pipe (1) of solar powered aircraft main spar；
Step 2: the outer surface of the first glue film (2) in step 1 is laid with the first wrapping layer (3)；
Step 3: the outer surface of the first wrapping layer (3) in step 2 is laid with the second glue film (4)；
Step 4: the outer surface of the second glue film (4) in step 3 puts on metal flange (5), wherein the metal flange
(5) diameter of the interior circular diameter of neck (51) equal to the outer surface for forming product after step 3；
Step 5: being arranged bag outside the metal flange (5) in step 4, then makes to air is taken out in bag for vacuum, etc.
After required time, bag is removed；
Step 6: the outer surface of the neck (51) of the metal flange (5) in step 5 is laid with third glue film (6)；
Step 7: the outer surface of the third glue film (6) in step 6 is laid with the second wrapping layer (7), wherein second package
The number of plies of layer (7) is several layers, and each layer of second wrapping layer (7) is aligned close to bottom (52) one end, far from bottom
(52) the other end staggers successively arrangement；
Step 8: the product formed after step 7 is solidified in curing apparatus；
Wherein, it is equally distributed to offer the axial and circumferential along the neck (51) for the neck (51) of the metal flange (5)
The slot in several holes and the axial direction along the neck (51).
8. the preparation method of the connection structure according to claim 7 for solar powered aircraft main spar, it is characterised in that:
First wrapping layer (3) and second wrapping layer (7) are carbon fibre fabric prepreg；The curing apparatus is autoclave
Or baking oven.
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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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