CN106882360A - A kind of attachment structure and preparation method for solar powered aircraft main spar - Google Patents
A kind of attachment structure and preparation method for solar powered aircraft main spar Download PDFInfo
- Publication number
- CN106882360A CN106882360A CN201610945625.4A CN201610945625A CN106882360A CN 106882360 A CN106882360 A CN 106882360A CN 201610945625 A CN201610945625 A CN 201610945625A CN 106882360 A CN106882360 A CN 106882360A
- Authority
- CN
- China
- Prior art keywords
- integument
- glued membrane
- solar powered
- metal flange
- main spar
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
- 238000002360 preparation method Methods 0.000 title claims abstract description 13
- 239000012528 membrane Substances 0.000 claims abstract description 104
- 239000002184 metal Substances 0.000 claims abstract description 91
- 229910052751 metal Inorganic materials 0.000 claims abstract description 91
- 229920000049 Carbon (fiber) Polymers 0.000 claims abstract description 30
- 239000004917 carbon fiber Substances 0.000 claims abstract description 30
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 claims abstract description 29
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 claims description 13
- 229910052799 carbon Inorganic materials 0.000 claims description 13
- 239000000835 fiber Substances 0.000 claims description 10
- 239000004744 fabric Substances 0.000 claims description 9
- 238000000034 method Methods 0.000 claims description 9
- 229910001069 Ti alloy Inorganic materials 0.000 claims description 7
- 239000000956 alloy Substances 0.000 claims description 7
- 239000002131 composite material Substances 0.000 abstract description 7
- 239000007769 metal material Substances 0.000 abstract description 5
- 239000010410 layer Substances 0.000 description 19
- 239000000463 material Substances 0.000 description 16
- 238000002386 leaching Methods 0.000 description 5
- 239000003292 glue Substances 0.000 description 4
- 230000001965 increasing effect Effects 0.000 description 4
- 230000000694 effects Effects 0.000 description 3
- 238000001125 extrusion Methods 0.000 description 3
- 239000000203 mixture Substances 0.000 description 3
- 239000000853 adhesive Substances 0.000 description 2
- 230000001070 adhesive effect Effects 0.000 description 2
- 230000002708 enhancing effect Effects 0.000 description 2
- PCHJSUWPFVWCPO-UHFFFAOYSA-N gold Chemical compound [Au] PCHJSUWPFVWCPO-UHFFFAOYSA-N 0.000 description 2
- 239000010931 gold Substances 0.000 description 2
- 229910052737 gold Inorganic materials 0.000 description 2
- 238000004519 manufacturing process Methods 0.000 description 2
- 239000002905 metal composite material Substances 0.000 description 2
- 241000233855 Orchidaceae Species 0.000 description 1
- RTAQQCXQSZGOHL-UHFFFAOYSA-N Titanium Chemical compound [Ti] RTAQQCXQSZGOHL-UHFFFAOYSA-N 0.000 description 1
- 230000007423 decrease Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 239000003822 epoxy resin Substances 0.000 description 1
- 239000011229 interlayer Substances 0.000 description 1
- 238000003754 machining Methods 0.000 description 1
- 229920000647 polyepoxide Polymers 0.000 description 1
- 238000005096 rolling process Methods 0.000 description 1
- 238000007711 solidification Methods 0.000 description 1
- 230000008023 solidification Effects 0.000 description 1
- 239000010936 titanium Substances 0.000 description 1
- 229910052719 titanium Inorganic materials 0.000 description 1
Classifications
-
- 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 attachment structure and preparation method for solar powered aircraft main spar, wherein, the attachment structure includes：First glued membrane, the first integument, the second glued membrane, metal flange, the 3rd glued membrane and the second integument；Wherein, first glued membrane is used to wrap located at carbon fiber pipe；The first integument bag is located at first glued membrane；The second glued membrane bag is located at first integument；The metal flange is sheathed on second glued membrane；Neck of the 3rd glued membrane bag located at the metal flange；The second integument bag is located at the 3rd glued membrane.Attachment structure for solar powered aircraft main spar of the invention constitutes entirety by by the first glued membrane, the first integument, the second glued membrane, metal flange, the 3rd glued membrane and the second integument, enable that metal material is effectively connected with composite, enhance bonding strength, it is ensured 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 connection for solar powered aircraft main spar
Structure and preparation method.
Background technology
Due to the superior material property of carbon fiber, it has been widely used among solar powered aircraft main spar structure design.
But carbon fibre composite belongs to fragile material, easily damaged after structure connection stress, for composite structure
Connection, the particularly connection between primary load bearing composite structure are a greatly challenges.And by metal material and composite knot
Closing use can efficiently solve connectivity problem, but the combination of mesh first two material is most based on secondary bonding, with group
Into as a example by the carbon fiber pipe of solar powered aircraft main spar：Manufacture is processed to carbon fiber pipe first, then to metal method
Orchid is processed manufacture, finally using structure glue by metal flange and carbon fiber pipe bonding forming.This process lack
Point includes：It is due to pressure cannot be applied to metal flange therefore high to the requirement on machining accuracy of metal flange and pipe, it is glued
Interface easily produces starved phenomenon；Second bonding uses normal temperature cure, and the hardening time of general such epoxy resin structural adhesive is extremely
A few hours are needed less, and in the curing process because the mobility of structure glue easily occurs the situation of upper end starved, and starved is necessarily led
Adhesive strength is caused to decline.
The content of the invention
Present invention solves the technical problem that being：Compared to prior art, there is provided one kind is used for solar powered aircraft main spar
Attachment structure and preparation method so that metal flange can effectively be connected with the carbon fiber pipe of solar powered aircraft main spar,
And then cause that the quality of the solar powered aircraft main spar for being connected to each other by metal flange by carbon fiber pipe and being constituted is obtained 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, including：First glued membrane, the first integument, the second glued membrane, metal flange, the 3rd glued membrane and the second integument；Wherein, it is described
First glued membrane bag is located at the carbon fiber pipe outer surface of solar powered aircraft main spar；The first integument bag is located at first glue
Film outer surface；The second glued membrane bag is located at the first integument outer surface；The metal flange is sheathed on second glue
Film outer surface；The 3rd glued membrane bag is located at the outer surface of the metal flange neck；The second integument bag is located at described
3rd glued membrane outer surface, wherein, if the number of plies of second integument is dried layer, each layer of second integument is the bottom of near
Portion one end is alignd, and the other end away from bottom staggers arrangement successively.
In the above-mentioned attachment 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 attachment 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 attachment structure for solar powered aircraft main spar, the neck of the metal flange along the neck axle
To offering groove.
In the above-mentioned attachment structure for solar powered aircraft main spar, the quantity of the groove is multiple, and multiple grooves are along described
The circumference of neck is uniformly distributed.
In the above-mentioned attachment structure for solar powered aircraft main spar, the metal flange is titanium alloy material；Described
One integument and second integument are carbon fibre fabric prepreg..
In the above-mentioned attachment structure for solar powered aircraft main spar, the number of plies of second integument is at least 3 layers, institute
Each layer for stating the second integument is 10mm-20mm away from the distance that the other end of bottom staggers.
A kind of preparation method of attachment structure for solar powered aircraft main spar, the described method comprises the following steps：
Step one：The first glued membrane is laid in the outer surface of the carbon fiber pipe of solar powered aircraft main spar；
Step 2：Lay the first integument in the outer surface of the first glued membrane in step one；
Step 3：Lay the second glued membrane in the outer surface of the first integument in step 2；
Step 4：Metal flange on the outer surface set of the second glued membrane in step 3, wherein, the neck of the metal flange
Diameter of the interior circular diameter in portion equal to the outer surface that product is formed after step 3；
Step 5：Bag is arranged outside the metal flange in step 4, then causes to be vacuum to taking out air in bag, etc.
After the time required to, bag is removed；
Step 6：Lay the 3rd glued membrane in the outer surface of the neck of the metal flange in step 5；
Step 7：Lay the second integument, the layer of second integument in the outer surface of the 3rd glued membrane in step 6
If number is dried layer, near bottom end alignment, the other end away from bottom staggers row successively for each layer of second integument
Cloth；
Step 8：The product that will be formed after step 7 solidifies in curing apparatus.
In the preparation method of the above-mentioned attachment structure for solar powered aircraft main spar, first integument and described
Two integuments are carbon fibre fabric prepreg；The curing apparatus are autoclave or baking oven.
In the preparation method of the above-mentioned attachment structure for solar powered aircraft main spar, the neck of the metal flange is opened up
Have along equally distributed several holes of axial and circumferential of the neck and the axial groove along the neck.
The present invention has the advantages that compared with prior art：
(1), the present invention is by by the first glued membrane, the first integument, the second glued membrane, metal flange, the 3rd glued membrane and second
Integument constitutes overall so that metal flange can effectively be connected with the carbon fiber pipe of solar powered aircraft main spar, enhance
Bonding strength, and each Carbon fibe pipe by metal flange be connected to each other composition solar powered aircraft main spar so that too
The fastness of positive energy aircraft main spar is good, and quality is guaranteed；
(2), the present invention strengthens bonding strength using titanium alloy material as the structural material of metal flange；
(3), the present invention offers hole and groove by metal flange, increase the composite of integument and glued membrane composition with
The glue-joint strength of metal flange；
(4) alignment of second integument one end and the other end, in the present invention is staggered arrangement successively, is reduced to the second parcel
The stress concentration of layer.
Brief description of the drawings
Fig. 1 shows the structural representation of the attachment structure for solar powered aircraft main spar provided in an embodiment of the present invention
Figure；
Fig. 2 shows the attachment structure part-structure for solar powered aircraft main spar that bright embodiment of the invention is provided
Schematic diagram；
Fig. 3 shown in the attachment structure for solar powered aircraft main spar that bright embodiment of the invention is provided, metal
The structural representation 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 attachment structure for solar powered aircraft main spar provided in an embodiment of the present invention
Figure.Fig. 2 shows that the attachment structure part-structure for solar powered aircraft main spar that bright embodiment of the invention is provided is illustrated
Figure.As depicted in figs. 1 and 2, the attachment structure includes：Carbon fiber pipe 1, the first glued membrane 2, the first integument 3, the second glued membrane 4, gold
Category flange 5, the 3rd glued membrane 6 and the second integument 7.During specific implementation, carbon fiber pipe 1 is pipe, by the preimpregnation of carbon fiber one-way band
Material to be twined laid using pipe rolling technique and solidify and formed.First integument 3 and the second integument 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 the first glued membrane 2 and the second glued membrane 4
It is J-272 glued membranes.Wherein,
The bag of first glued membrane 2 is located at carbon fiber pipe 1.Specifically, the first glued membrane 2 is closely layed in the outer surface of carbon fiber pipe 1,
And the viscous force and carbon fiber pipe 1 by the first glued membrane 2 are bonded together.
The bag of first integument 3 is located at the first glued membrane 2.Specifically, the first integument 3 is tight using carbon fibre fabric prepreg
The outer surface of the first glued membrane 2 is layed in, the first integument 3 is caused by the viscous force of the first integument 3 and the viscous force of the first glued membrane 2
Combined closely with the first glued membrane 2.
The bag of second glued membrane 4 is located at the first integument 3.Specifically, the second glued membrane 4 is closely layed in the outer of the first integument 3
Surface, the first integument 3 and the second glued membrane 4 closely knot are caused by the viscous force of the first integument 3 and the viscous force of the second glued membrane 4
Close.
Metal flange 5 is sheathed on the second glued membrane 4.Specifically, metal flange 5 is set in is covered with first the 2, first parcel of glued membrane
On the carbon fiber pipe 1 of the glued membrane 4 of layer 3 and second, 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 external diameter of the carbon fiber pipe 1 for being covered with the first glued membrane 2, the first integument 3 and the second glued membrane 4, so that
So that metal flange 5 is closely in contact with the second glued membrane 4.
Neck 51 of the bag of 3rd glued membrane 6 located at metal flange 5.Specifically, the 3rd glued membrane 6 is closely layed in metal flange 5
Neck 51 outer surface, metal flange 5 is closely connected with the 3rd glued membrane 6 by the viscous force of the 3rd glued membrane 6.
The bag of second integument 7 is located at the 3rd glued membrane 6, wherein, if the number of plies of the second integument 7 is dried layer, the second integument 7
Each layer near the alignment of the one end of bottom 52, the other end away from bottom 52 staggers arrangement successively.Specifically, the second integument 7
The outer surface of the 3rd glued membrane 6 is closely layed in, is closely connected the second integument 7 with the 3rd glued membrane 6 by the viscous force of the 3rd glued membrane 6
Connect.The number of plies of the second integument 7 is multilayer, wherein, each layer aligns in the one end near bottom 52, and the other end staggers arrangement successively
It is stepped.So as to reduce 5 pairs of stress concentrations of the second integument 7 of metal flange, it is ensured that protected in the case of bonding strength
The second integument 7 is protected.Further, the number of plies of the second integument 7 is at least 3 layers, and each layer of the second integument 7 is away from bottom
The distance that 52 other end staggers is 10mm-20mm so that the second integument 7 is protected in the case of ensure that bonding strength
Effect is more obvious.
The present embodiment is wrapped by by the first glued membrane, the first integument, the second glued membrane, metal flange, the 3rd glued membrane and second
Covering layer constitutes overall so that metal material can effectively be connected with composite, enhance bonding strength；Using titanium alloy material
Expect, as the structural material of metal flange, to strengthen bonding strength；Alignd by second integument one end and the other end staggers successively
Arrangement reduces stress concentration of the metal flange to the second integument, is effectively protected the second integument, increased connection strong
Degree.
Fig. 3 shown in the attachment structure for solar powered aircraft main spar that bright embodiment of the invention is provided, metal
The structural representation 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 glued membrane 4
With the thickness of the 3rd glued membrane 6 than relatively thin, so that the first integument 3 is partially submerged into phase together with the corresponding aperture of the second glued membrane 4
The hole of corresponding neck 51, increases shear strength between the integument 3 of metal flange 5 and first.Due to the thickness ratio of the 3rd glued membrane 6
It is relatively thin, so that the hole that is partially submerged into corresponding neck 51 of second integument 7 together with the corresponding aperture of the 3rd glued membrane 6, increases
Plus shear strength between the integument 7 of metal flange 5 and second.
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 the effect of shear strength and increase metal flange 5 and the between the integument 3 of metal flange 5 and first
The effect of shear strength is more significantly between two integuments 7.
In above-mentioned implementation, the neck 51 of metal flange 5 offers groove 53 along the axial direction of neck 51.Can be by gold using groove 53
Unnecessary leaching material extrusion between unnecessary leaching material and the integument 7 of metal flange 5 and second between category flange 5 and carbon fiber pipe 1, keeps away
Exempt from occur space between the integument 3 of metal flange 5 and first, the second integument 7, increase interlaminar strength.
Further, the quantity of groove 53 is multiple, and multiple grooves 53 are uniformly distributed along the circumference of neck 51.By being uniformly distributed
Further enhancing the integument 3 of metal flange 5 and first, the bonding strength of the second integument 7.
Attachment structure for solar powered aircraft main spar of the invention is by by the first glued membrane, the first integument, second
Glued membrane, metal flange, the 3rd glued membrane and the second integument constitute overall so that metal material can effectively connect with composite
Connect, enhance bonding strength, further, due to each Carbon fibe pipe by metal flange be connected to each other composition solar energy fly
The main spar of machine so that the fastness of solar powered aircraft main spar is good, and quality is ensured very well；Using titanium alloy material conduct
The structural material of metal flange, strengthens bonding strength；Alignd by one end of the second integument and the other end staggers arrangement successively
Stress concentration of the metal flange to the second integument is reduced, the second integument is effectively protected, bonding strength is increased.
Present invention also offers a kind of preparation method of the attachment structure for solar powered aircraft main spar, the method includes
Following steps：
Step one：The first glued membrane 2 is laid in the outer surface of the carbon fiber pipe 1 of solar powered aircraft main spar；
Step 2：Lay the first integument 3 in the outer surface of the first glued membrane 2 in step one；
Step 3：Lay the second glued membrane 4 in the outer surface of the first integument 3 in step 2；
Step 4：Metal flange 5 on the outer surface set of the second glued membrane 4 in step 3, wherein, the neck of metal flange 5
Diameter of the interior circular diameter in portion 51 equal to the outer surface that product is formed after step 3；
Step 5：Bag is arranged outside the metal flange 5 in step 4, then causes to be vacuum to taking out air in bag,
After the time required to waiting, bag is removed；
Step 6：Lay the 3rd glued membrane 6 in the outer surface of the neck 51 of the metal flange 5 in step 5；
Step 7：Lay the second integument 7 in the outer surface of the 3rd glued membrane 6 in step 6；
Step 8：The product that will be formed after step 7 solidifies in curing apparatus.
In step 2, the first integument 3 uses carbon fibre fabric prepreg.
In step 4, metal flange 5 uses titanium alloy material, metal flange 5 to be set in and be covered with the first glued membrane 2, first
On the carbon fiber pipe 1 of the glued membrane 4 of integument 3 and second, 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 external diameter of the carbon fiber pipe 1 for being covered with the first glued membrane 2, the first integument 3 and the second glued membrane 4,
So that metal flange 5 is closely in contact with the second glued membrane 4.
In step 5, using the bag for vacuumizing so that metal flange 5 be covered with the first glued membrane 2, the first integument 3 and
Bonding strength enhancing between the carbon fiber pipe 1 of two glued membranes 4.The neck 51 of metal flange 5 offer along neck 51 axial direction and
Circumferential equally distributed several holes, and neck 51 in metal flange 5 is further opened with along the axial groove of neck 51.So as to
During being compressed using vacuum bag so that the first integument 3 is corresponding together with being partially submerged into for the corresponding aperture of the second glued membrane 4
Neck 51 hole, increase the integument 3 of metal flange 5 and first between shear strength.Can be by metal flange 5 and carbon using groove
Unnecessary leaching material extrusion between fibre pipe 1, it is to avoid occur space between the integument 3 of metal flange 5 and first, increase interlayer strong
Degree.
In step 7, the second integument 7 uses carbon fibre fabric prepreg.If the number of plies of the second integument 7 is dried layer,
Second integument 7 forms step structure.Specifically, the number of plies of the second integument 7 is multilayer, wherein, every layer of one end alignment,
The other end is stepped.So as to reduce 5 pairs of stress concentrations of the second integument 7 of metal flange, the second integument 7 is protected.
In step 8, solidified in autoclave or baking oven, during solidification, also can be to being formed after step 7
Product play the role of individual pressure so that the second integument 7 is corresponding together with being partially submerged into for the corresponding aperture of the 3rd glued membrane 6
Neck 51 hole, increase the integument 7 of metal flange 5 and second between shear strength, while also again increasing metal flange 5
The shear strength between the first integument 3, and, can be by the unnecessary leaching material between metal flange 5 and carbon fiber pipe 1 using groove
And the unnecessary leaching material extrusion between the integument 7 of metal flange 5 and second, it is to avoid the integument 3, second of metal flange 5 and first is wrapped
Occur space between covering layer 7, increase interlaminar strength.
The preparation method of the attachment structure for solar powered aircraft main spar of the invention is wrapped by by the first glued membrane, first
Covering layer, the second glued membrane, metal flange, the 3rd glued membrane and the second integument constitute overall so that metal material and composite energy
Enough effective connections, enhance bonding strength, further, because each Carbon fibe pipe is connected to each other group by metal flange
Into the main spar of solar powered aircraft so that the fastness of solar powered aircraft main spar is good, and quality is ensured very well；Closed using titanium
Golden material strengthens bonding strength as the structural material of metal flange；Alignd by one end of the second integument and the other end according to
Secondary arrangement of staggering reduces stress concentration of the metal flange to the second integument, is effectively protected the second integument.
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 the range of bright technical scheme all should be comprising within the scope of the present invention.
Claims (10)
1. a kind of attachment structure for solar powered aircraft main spar, it is characterised in that including：Carbon fiber pipe (1), the first glued membrane
(2), the first integument (3), the second glued membrane (4), metal flange (5), the 3rd glued membrane (6) and the second integument (7)；Wherein,
First glued membrane (2) bag is located at carbon fiber pipe (1) outer surface of solar powered aircraft main spar；
First integument (3) bag is located at the first glued membrane (2) outer surface；
Second glued membrane (4) bag is located at the first integument (3) outer surface；
The metal flange (5) is sheathed on the second glued membrane (4) outer surface；
Outer surface of 3rd glued membrane (6) bag located at the metal flange (5) neck (51)；
Second integument (7) bag is located at the 3rd glued membrane (6) outer surface, wherein, the number of plies of second integument (7)
If being dried layer, each layer of second integument (7) near the alignment of bottom (52) one end, away from bottom (52) the other end according to
Secondary arrangement of staggering.
2. the attachment structure for solar powered aircraft main spar according to claim 1, it is characterised in that：The metal method
The neck (51) of blue (5) offers several holes, and several holes are uniformly distributed along the axial and circumferential of the neck (51).
3. the attachment structure for solar powered aircraft main spar according to claim 2, it is characterised in that：It is axially adjacent
The spacing in hole is 10mm-15mm；The spacing in circumferentially adjacent hole is 20mm-25mm.
4. according to any described attachment structures for solar powered aircraft main spar of claim 1-2, it is characterised in that：It is described
The neck (51) of metal flange (5) offers groove (53) along the axial direction of the neck (51).
5. the attachment structure for solar powered aircraft main spar according to claim 4, it is characterised in that：The groove (53)
Quantity be multiple, multiple grooves (53) are uniformly distributed along the circumference of the neck (51).
6. the attachment structure for solar powered aircraft main spar according to claim 2, it is characterised in that：The metal method
Blue (5) are titanium alloy material；First integument (3) and second integument (7) are carbon fibre fabric prepreg.
7. the attachment structure for solar powered aircraft main spar according to claim 1, it is characterised in that：Second bag
The number of plies of covering layer (7) is at least 3 layers, the distance that each layer of second integument (7) staggers away from the other end of bottom (52)
It is 10mm-20mm.
8. a kind of preparation method of attachment structure for solar powered aircraft main spar, is characterised by, methods described includes following
Step：
Step one：The first glued membrane (2) is laid in the outer surface of the carbon fiber pipe (1) of solar powered aircraft main spar；
Step 2：Lay the first integument (3) in the outer surface of the first glued membrane (2) in step one；
Step 3：Lay the second glued membrane (4) in the outer surface of the first integument (3) in step 2；
Step 4：The outer surface of the second glued membrane (4) in step 3 puts metal flange (5), wherein, the metal flange
(5) diameter of the interior circular diameter of neck (51) equal to the outer surface that product is formed after step 3；
Step 5：Bag is arranged outside the metal flange (5) in step 4, then causes to be vacuum to taking out air in bag, etc.
After the time required to, bag is removed；
Step 6：Lay the 3rd glued membrane (6) in the outer surface of the neck (51) of the metal flange (5) in step 5；
Step 7：The second integument (7) is laid in the outer surface of the 3rd glued membrane (6) in step 6, wherein, second parcel
Layer (7) if the number of plies be dried layer, each layer of second integument (7) near the alignment of bottom (52) one end, away from bottom
(52) the other end staggers arrangement successively；
Step 8：The product that will be formed after step 7 solidifies in curing apparatus.
9. the preparation method of the attachment structure for solar powered aircraft main spar according to claim 8, it is characterised in that：
First integument (3) and second integument (7) are carbon fibre fabric prepreg；The curing apparatus are autoclave
Or baking oven.
10. the preparation method of the attachment structure for solar powered aircraft main spar according to claim 8, its feature exists
In：The neck (51) of the metal flange (5) offer axial and circumferential along the neck (51) it is equally distributed several
Hole and the axial groove along the neck (51).
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| 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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| 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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| CN106882360B CN106882360B (en) | 2019-05-24 |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110451980A (en) * | 2019-09-16 | 2019-11-15 | 南京膜材料产业技术研究院有限公司 | A kind of preparation method of the enhanced ceramic high temperature flue gas filter pipe of flange neck |
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| 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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| 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 |
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| CN110451980B (en) * | 2019-09-16 | 2021-11-05 | 江苏久朗高科技股份有限公司 | Preparation method of flange neck enhanced ceramic high-temperature flue gas filter pipe |
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|---|---|
| CN106882360B (en) | 2019-05-24 |
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