CN209775592U - Overall forming tool for composite fuselage skin - Google Patents
Overall forming tool for composite fuselage skin Download PDFInfo
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- CN209775592U CN209775592U CN201920418427.1U CN201920418427U CN209775592U CN 209775592 U CN209775592 U CN 209775592U CN 201920418427 U CN201920418427 U CN 201920418427U CN 209775592 U CN209775592 U CN 209775592U
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Abstract
The utility model discloses a composite fuselage skin integral molding tool, relating to the technical field of unmanned aerial vehicle manufacturing, comprising a frame and a tool template; mounting grooves are formed in two opposite surfaces of the frame, the two mounting grooves are arranged oppositely, and two ends of the tooling template are fixedly clamped on the two mounting grooves; the frame upper surface still installs two sets of loose pieces of dismantling, and two sets of loose pieces of dismantling are located the both sides of frock template respectively, and the medial surface of two sets of loose pieces of dismantling forms a continuous face with the working face of frock template, and two sets of loose pieces of dismantling constitute the curing chamber with the frock template. The utility model discloses with integrated configuration's form design profile, reduced the frock and made the degree of difficulty, realized the technological requirement of big curvature frock appearance to after product solidification shaping, through dismantling the loose piece one by one, realized the technological requirement of product drawing of patterns.
Description
Technical Field
The utility model relates to an unmanned aerial vehicle makes technical field, particularly, relates to a combined material fuselage covering integral forming frock.
Background
The advanced composite material has the advantages of light weight, high strength, high specific stiffness, strong designability, good fatigue fracture resistance, corrosion resistance, good dimensional stability and the like; simultaneously, adopt the integrated manufacturing technology can reduce part and fastener quantity by a wide margin, reduce structure weight, reduction in production cost, consequently, most unmanned aerial vehicle fuselage covering all adopts advanced combined material to make at present.
At present, a mold used in the integral molding of a fuselage skin is complex, and co-curing and secondary bonding are adopted for an integral honeycomb sandwich structure process method; however, in the co-curing process, too good resin fluidity can cause too much glue flow to cause poor glue of the panel, too high molding pressure can cause the upper panel to be sunken, the honeycomb to be unstable and slide, and the pressure is small to cause panel defects and weak glue joint, so that the yield of the fuselage skin is difficult to control.
Therefore, the composite material fuselage skin integral forming tool capable of solving the problems is provided.
Disclosure of Invention
An object of the utility model is to provide a combined material fuselage covering integrated into one piece frock to solve above-mentioned problem.
For realizing the purpose of the utility model, the technical proposal adopted is that: a composite material fuselage skin integral molding tool comprises a frame and a tool template; mounting grooves are formed in two opposite surfaces of the frame, the two mounting grooves are arranged oppositely, and two ends of the tooling template are fixedly clamped on the two mounting grooves; the frame upper surface still installs two sets of loose pieces of dismantling, and two sets of loose pieces of dismantling are located the both sides of frock template respectively, and the medial surface of two sets of loose pieces of dismantling forms a continuous face with the working face of frock template, and two sets of loose pieces of dismantling constitute the curing chamber with the frock template.
Furthermore, each group of detachable loose pieces comprises a plurality of detachable loose pieces which are sequentially and closely arranged.
Furthermore, the detachable loose piece comprises a horizontal part and an arc-shaped part vertically fixed on the horizontal part, and the inner surface of the arc-shaped part and the inner surface of the tooling template are spliced to form a continuous surface.
Furthermore, each arc-shaped part of each detachable loose piece is provided with a positioning hole.
Further, the both sides of frock template all are provided with along the platform, and two are all fixed at the top surface of frame along the platform.
Further, the tooling template and the edge platform are integrally formed.
Furthermore, two groups of detachable loose pieces are respectively arranged on the two edge platforms.
furthermore, a plurality of rotating wheels are arranged at the bottom of the frame.
Furthermore, a plurality of lifting rings are further arranged on the frame.
Furthermore, detachable locators are further installed at two ends of the tooling template.
The utility model has the advantages that,
1. The molded surface is designed in a combined structure mode, the manufacturing difficulty is reduced, the technological requirement of the appearance of a large-curvature tool is met, and the technological requirement that the product can be demoulded is met by disassembling the detachable loose pieces one by one after the product is solidified and molded; meanwhile, the composite material fuselage skin integral forming tool can be used for manufacturing products and honeycomb positioning plates, and the cost for manufacturing the honeycomb positioning plates by independently opening the molds is reduced.
2. According to the special structural form of the fuselage skin, the product manufacturing is realized by adopting a mode of combining the integral forming tool and the process of the composite fuselage skin; meanwhile, a curing molding process is adopted, and the integrally molded product is a fuselage skin with a honeycomb sandwich structure, so that the structure of the fuselage skin is firmer; because the outer skin formed in advance does not contain honeycombs, the outer skin can be cured by using larger forming pressure, and the forming quality of the outer skin is ensured.
3. The honeycomb positioning plate and the outer skin are molded by using the same set of tool, so that the honeycomb positioning plate and the outer skin have the same thermal expansion amount; the accuracy of the laying position of the honeycomb on the outer skin can be ensured by adopting the honeycomb positioning plate to position the honeycomb; and under the condition of ensuring the quality of the honeycomb, the molding quality of the fuselage skin can be ensured by adopting the optimized curing parameters.
Drawings
Fig. 1 is a front view of the composite fuselage skin integral molding tool provided by the present invention;
Fig. 2 is a left side view of the composite fuselage skin integral molding tool provided by the present invention;
Fig. 3 is a top view of the composite fuselage skin integral molding tool provided by the present invention;
Fig. 4 is an exploded view of the composite fuselage skin integral molding tool provided by the present invention.
Reference numbers and corresponding part names in the drawings:
1. The device comprises a frame, 2, a tooling template, 3, a mounting groove, 4, a detachable loose piece, 5, a curing cavity, 6, an edge platform, 7, wheels, 8, a lifting ring, 9 and a positioner;
41. Horizontal portion, 42, arc portion, 43, locating hole.
Detailed Description
The present invention will be described in further detail below with reference to specific embodiments and with reference to the accompanying drawings.
Fig. 1 to 4 illustrate an integral forming tool for a composite fuselage skin provided by the present invention, which comprises a frame 1 and a tool template 2; two opposite surfaces of the frame 1 are respectively provided with a mounting groove 3, the two mounting grooves 3 are oppositely arranged, and two ends of the tooling template 2 are fixedly clamped on the two mounting grooves 3; still install two sets of loose pieces 4 of dismantling on the frame 1 upper surface, two sets of loose pieces 4 of dismantling are located the both sides of frock template 2 respectively, and two sets of medial surfaces that can dismantle loose piece 4 form a continuous face with the working face shape of frock template 2, and two sets of loose pieces 4 of dismantling and frock template 2 constitute solidification chamber 5.
The frame 1 is a rectangular frame 1, the mounting grooves 3 are U-shaped, the cross sections of the tooling template 2 are also U-shaped, so that the tooling template 2 is matched with the mounting grooves 3, and the two ends of the tooling template 2 can be stably supported in the two mounting grooves 3; frock template 2 and frame 1 welded fastening make the installation of frock template 2 more firm, and the high-end of frock template 2 and the top surface parallel and level of frame 1. Two sets of loose pieces 4 of dismantling all pass through the fix with screw on frame 1, and two sets of loose pieces 4 of dismantling are arranged on frame 1 symmetrically, make two sets of loose pieces 4 of dismantling be located the both sides of frock template 2 respectively, and two sets of lower extreme that can dismantle the internal surface of loose piece 4 and two side butt joints of frock template 2, make two sets of loose pieces 4 of dismantling constitute a concave solidification chamber 5 with frock template 2 jointly, and the upper end opening of solidification chamber 5 is two sets of clearances of dismantling between the upper end of loose piece 4.
When the fuselage skin or the honeycomb positioning plate needs to be produced, the two groups of detachable loose pieces 4 can be fixed on the frame 1 respectively, then materials for producing the fuselage skin or the honeycomb positioning plate are laid in the curing cavity 5, after the materials are laid, the composite fuselage skin integral forming tool is sent into the hot pressing tank for curing, so that the materials laid in the curing cavity 5 are cured, the materials in the curing cavity 5 are formed into the fuselage skin or the honeycomb positioning sample plate, the composite fuselage skin integral forming tool is taken out of the hot pressing tank, finally the two groups of detachable loose pieces 4 are detached from the frame 1, the fuselage skin or the honeycomb positioning sample plate produced by curing and forming is taken out of the curing cavity 5, and the taken-out honeycomb positioning sample plate is subjected to secondary processing to form the honeycomb positioning plate.
Each group of detachable loose pieces 4 comprises a plurality of detachable loose pieces 4 which are sequentially and closely arranged, so that the tension of the material can be uniformly dispersed through the plurality of detachable loose pieces 4 in the process of curing the material in the curing cavity 5 in the production process of the body skin or the honeycomb positioning plate, thereby preventing the body skin or the honeycomb positioning plate from deforming in the production process; the detachable loose piece 4 comprises a horizontal part 41 and an arc-shaped part 42 vertically fixed on the horizontal part 41, the horizontal part 41 and the arc-shaped part 42 are integrally arranged, and the inner surface of the arc-shaped part 42 is spliced with the inner surface of the tooling template 2 to form a continuous surface; the vertical part is fixed on the frame 1 through a screw, so that the installation of the detachable loose piece 4 is more convenient, and the installation of the detachable loose piece 4 is more stable.
Each arc-shaped part 42 of the detachable loose piece 4 is provided with a positioning hole 43, so that the positioning during the placement of the outer skin material piece, the inner skin material piece and the honeycomb positioning plate is facilitated, and the size of the produced fuselage skin is more accurate.
The two sides of the tooling template 2 are both provided with edge platforms 6, the two edge platforms 6 are both fixed on the top surface of the frame 1, the edge platforms 6 are in a long strip shape, the length of the edge platforms 6 is equal to that of the tooling template 2, and the two ends of the edge platforms 6 are respectively aligned with the two ends of the tooling template 2; in the installation process of the tooling template 2, the two sides of the installation template can be fixed by fixing the platform 6, so that the installation of the tooling template 2 is more stable. Frock template 2 with along platform 6 be integrated into one piece, make frock template 2 more firm with being connected along platform 6, not only make things convenient for the installation to frock template 2, and make the both sides of frock template 2 can be through along platform 6 fixed and more firm to the deformation volume of the fuselage covering and the honeycomb locating plate that make to produce is littleer.
Two sets of can dismantle loose piece 4 and install respectively on two platform 6 along, it is more convenient not only to make the installation of dismantling loose piece 4, and makes the internal surface that can dismantle loose piece 4 and frock template 2's butt joint department clearance littleer, makes the final product precision of processing out higher.
A plurality of rotating wheels 7 are also arranged at the bottom of the frame 1; the wheels 7 are at least 4, and are uniformly distributed on the bottom surface of the frame 1, so that when materials in the curing cavity 5 need to be cured, the composite material fuselage skin integral forming tool can be quickly conveyed into the hot pressing tank, and after the materials in the curing cavity 5 are cured, the composite material fuselage skin integral forming tool can be quickly pushed out of the hydraulic tank. Still be provided with a plurality of rings 8 on the frame 1, a plurality of rings 8 are located two relative lateral walls of frame 1 respectively, conveniently to the transport of combined material fuselage skin integrated into one piece frock.
Detachable locator 9 is still installed at the both ends of frock template 2, locator 9 can be for infrared locator 9 or directly adopt the pin location, makes at production fuselage covering in-process, can fix a position skin tablet and interior skin tablet, makes the size of the fuselage covering of production department more accurate.
Still engrave the honeycomb line on the frock template 2, make can cut according to the honeycomb line when processing into the honeycomb locating plate to honeycomb location model, make the honeycomb locating plate size of processing out more accurate.
when the composite fuselage skin needs to be machined, the machining method specifically comprises the following steps:
(1) Manufacturing a honeycomb positioning sample plate: laying the prepreg in the curing cavity 5, and sending the composite material fuselage skin integral molding tool with the laid prepreg into a hot pressing tank to cure the laid prepreg; and after the prepreg is cured, taking the composite material fuselage skin integral molding tool out of the hot pressing tank, and taking the cured prepreg down from the composite material fuselage skin integral molding tool to form a honeycomb positioning sample plate.
firstly, fixing a plurality of detachable loose pieces 4 on two edge tables 6 through screws respectively, so that the inner surfaces of the detachable loose pieces 4 and the inner surface of a tooling template 2 jointly form a curing cavity 5, and meanwhile, installing locators 9 at two ends of the tooling template 2; then, the prepreg is laid in the curing cavity 5, and the prepreg is not less than 5 layers, so that the honeycomb positioning sample plate formed after curing has certain rigidity; after the prepreg is laid, the composite material fuselage skin integral forming tool with the laid prepreg is sent into a hot pressing tank, so that the prepreg laid in the curing cavity 5 is cured, and a honeycomb positioning sample plate is formed; after the prepreg is completely cured, taking out the composite material fuselage skin integral forming tool from the hot pressing tank, and marking a positioning hole 43 on the honeycomb positioning sample plate through two positioners 9; finally, the plurality of removable loose pieces 4 are removed from the edge table 6 and the formed honeycomb locator card is removed from the curing chamber 5.
(2) Preparing parts before manufacturing: coating a release agent in the curing cavity 5, and cutting an outer skin material sheet and an inner skin material sheet; cutting the honeycomb positioning sample plate prepared in the step 1 to form a honeycomb positioning plate; and sticking demolding cloth on the surface of the honeycomb positioning plate.
Cutting the outer skin material sheet and the inner skin material sheet according to the blanking layout, cutting the honeycomb positioning sample plate prepared in the step 1 according to a honeycomb line on the honeycomb positioning sample plate by using a laser cutting machine, and cutting positioning holes 43 marked on a honeycomb positioning plate according to a positioner 9 and simultaneously cutting the positioning holes 43 on the honeycomb positioning plate in the cutting process; after the honeycomb positioning sample plate is cut, the honeycomb positioning sample plate is polished, so that the formed honeycomb positioning plate is smoother in appearance and more accurate in size; the demolding cloth is adhered to the surface of the honeycomb positioning plate, so that the honeycomb positioning plate can be conveniently taken out at a later stage; the release cloth can be directly replaced by release agent; the honeycomb positioning plate manufactured in the steps 1 and 2 only needs to be processed and produced when the fuselage skin is produced for the first time, and can be reused in the subsequent processing and production processes.
(3) forming an outer skin: the outer skin sheets prepared in the step 2 are laid in a curing cavity 5, and the composite material fuselage skin integral forming tool with the outer skin sheets laid is conveyed into a hot pressing tank to cure the laid outer skin sheets; and taking the cured outer skin material sheet down from the composite material fuselage skin integral molding tool to obtain the outer skin.
The outer skin sheets in the step 2 are laid in a curing cavity 5 coated with a release agent, and after the outer skin sheets are laid, the composite fuselage skin integral forming tool with the outer skin sheets laid is conveyed into a hot pressing tank to be cured, so that the outer skin sheets are cured in the hot pressing tank to form outer skins; after the outer skin is cured and formed, the composite material fuselage skin integral forming tool comprising the outer skin is taken out from the hot pressing tank, and after the composite material fuselage skin integral forming tool is taken out, the outer skin is stored in the curing cavity 5.
(4) Honeycomb laying: brushing a first layer of adhesive on the inner surface of the outer skin prepared in the step 3; placing the honeycomb positioning plate prepared in the step 2 on an outer skin, and projecting the honeycomb by using a laser projector to accurately position the honeycomb positioning plate; at the moment, the honeycomb is laid on the outer skin through the positioning of the honeycomb positioning plate, and finally the composite material fuselage skin integral forming tool after the honeycomb is laid is sent into a hot pressing tank for hot compaction; taking out the hot-compacted composite material fuselage skin integral forming tool from the hot pressing tank, taking down the honeycomb positioning plate from the composite material fuselage skin integral forming tool, and brushing a second layer of binder on the honeycomb region and the non-honeycomb region; and finally, conveying the composite material fuselage skin integral forming tool into the hot-pressing tank again for pre-compaction.
the first adhesive and the second adhesive in the step 4 are core adhesive films, the first adhesive is uniformly distributed on the surface of the outer skin in the first adhesive brushing and attaching process, the honeycomb positioning plate prepared in the step 2 is placed on the outer skin brushed with the first adhesive after the first adhesive is brushed and attached, and the positioning holes 43 in the honeycomb positioning plate are positioned by the two positioners 9 in the honeycomb positioning plate placing process, so that the honeycomb positioning plate is placed more accurately; after the honeycomb positioning plate is placed, the honeycombs are placed on the outer skin coated with the first adhesive one by one through the positioning of the honeycomb positioning plate, and after the honeycombs are placed, the composite material fuselage skin integral forming tool is conveyed into a hot pressing tank to be subjected to hot pressing, so that the honeycombs are bonded and fixed on the outer skin.
After the honeycomb and the outer skin are bonded and fixed, the composite material fuselage skin integral forming tool is taken out of the hot pressing tank, and the honeycomb positioning plate is not bonded and fixed with the outer skin due to the fact that the demolding cloth is adhered to the surface of the honeycomb positioning plate, so that the honeycomb positioning plate can be quickly taken down from the outer skin; after the honeycomb positioning plate is taken down, brushing a second adhesive on the surfaces of the honeycomb and the first adhesive after hot compaction, after the brushing of the second adhesive is finished, sending the composite material fuselage skin integral forming tool into a hot pressing tank for pre-compaction at the temperature of 60 +/-5 ℃, and keeping the temperature for 30 +/-5 min in the pre-compaction process so as to cure the second adhesive, thereby firmly bonding the honeycomb and the outer skin.
(5) Integral molding: and (4) taking the composite material fuselage skin integral forming tool in the step (4) out of the autoclave, laying an inner skin material sheet on the honeycomb, sending the composite material fuselage skin integral forming tool with the inner skin material sheet laid on the honeycomb into the autoclave for curing, finally taking the composite material fuselage skin integral forming tool out of the autoclave, and taking down the formed skin on the composite material fuselage skin integral forming tool.
After the pre-compaction in the step 4, taking out the composite fuselage skin integral forming tool from the hot pressing tank, after taking out the composite fuselage skin integral forming tool, laying an inner skin sheet on the honeycomb, after laying an outer wall on the inner skin sheet, sending the composite fuselage skin integral forming tool into the hot pressing tank again for curing, so that the inner skin sheet is bonded and fixed on the honeycomb through a second adhesive, the outer skin sheet, the honeycomb and the inner skin sheet are cured together to form the fuselage skin, after the fuselage skin is cured and formed, taking out the composite fuselage skin integral forming tool from the hot pressing tank, at the moment, detaching the detachable loose piece 4 from the edge table 6, and finally taking out the cured and formed fuselage skin from the curing chamber 5 to complete the production of the fuselage skin.
In the step 3, in the process of conveying the composite fuselage skin integral forming tool provided with the outer skin material sheet into the autoclave for curing, vacuumizing the autoclave to be below-0.7 Bar, and heating to 80 +/-5 ℃ at a speed of less than 1.5 ℃/min; and while raising the temperature, introducing the atmosphere when the autoclave is pressurized to 1.0Bar at a pressurizing rate of 0.5Bar/min, and continuously pressurizing to 5.0Bar +/-0.2 Bar at a pressurizing rate of 0.5 Bar/min. In the step 5, the composite fuselage skin integral forming tool paved with the inner skin material sheet is conveyed into an autoclave for curing, the autoclave is firstly vacuumized to-0.28 Bar-0.34 Bar, and the temperature is increased to 80 +/-5 ℃ at a speed of less than 1.5 ℃/min; while raising the temperature, the autoclave was vented when pressurized to 1.0Bar at a pressurization rate of 0.5Bar/min, and continued to be pressurized to 3.0Bar + -0.2 Bar at a pressurization rate of 0.5 Bar/min.
in the steps 3 and 5, in the curing process of the outer skin material sheet and the inner skin material sheet, firstly, the temperature in the autoclave is raised to 80 +/-5 ℃, and the autoclave is kept at the constant temperature for the first time, wherein the constant temperature time for the first time is 20-25 min; after the first constant temperature is finished, heating to 130 +/-5 ℃ at the speed of less than 1.5 ℃/min, and starting the second constant temperature, wherein the second constant temperature is 100-120 min; and after the second constant temperature is finished, cooling the interior of the autoclave at the speed of not more than 1.5 ℃/min, and finally, releasing the pressure of the autoclave after the temperature is reduced to below 60 ℃.
The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention, and various modifications and changes may be made by those skilled in the art. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (10)
1. The integral forming tool for the composite fuselage skin is characterized by comprising a frame and a tool template; mounting grooves are formed in two opposite surfaces of the frame, the two mounting grooves are arranged oppositely, and two ends of the tooling template are fixedly clamped on the two mounting grooves; the frame upper surface still installs two sets of loose pieces of dismantling, and two sets of loose pieces of dismantling are located the both sides of frock template respectively, and the medial surface of two sets of loose pieces of dismantling forms a continuous face with the working face of frock template, and two sets of loose pieces of dismantling constitute the curing chamber with the frock template.
2. The composite fuselage skin integral molding tooling of claim 1, wherein each group of the detachable loose pieces comprises a plurality of detachable loose pieces which are closely arranged in sequence.
3. The composite fuselage skin integral molding tooling of claim 1 or 2, wherein the detachable loose piece comprises a horizontal portion and an arc-shaped portion vertically fixed on the horizontal portion, and the inner surface of the arc-shaped portion and the inner surface of the tooling template are spliced to form a continuous surface.
4. The composite fuselage skin integral molding tooling of claim 3, wherein the arc-shaped portion of each detachable loose piece is provided with a positioning hole.
5. The composite fuselage skin integral molding tooling of claim 1, wherein two sides of the tooling template are provided with edge platforms, and the two edge platforms are fixed on the top surface of the frame.
6. The composite fuselage skin integral molding tooling of claim 5, wherein the tooling template is integrally molded with the edge platform.
7. The composite fuselage skin integral molding tooling as claimed in claim 5 or 6, wherein the two sets of detachable loose pieces are respectively mounted on two edge platforms.
8. The composite fuselage skin integral molding tooling of claim 1, wherein a plurality of rotating wheels are further mounted to the bottom of the frame.
9. The composite fuselage skin integral molding tooling as claimed in claim 1, wherein a plurality of lifting rings are further provided on the frame.
10. The composite fuselage skin integral molding tooling of claim 1, wherein two ends of the tooling template are further provided with detachable locators.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201920418427.1U CN209775592U (en) | 2019-03-29 | 2019-03-29 | Overall forming tool for composite fuselage skin |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201920418427.1U CN209775592U (en) | 2019-03-29 | 2019-03-29 | Overall forming tool for composite fuselage skin |
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| CN209775592U true CN209775592U (en) | 2019-12-13 |
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| CN201920418427.1U Active CN209775592U (en) | 2019-03-29 | 2019-03-29 | Overall forming tool for composite fuselage skin |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109774194A (en) * | 2019-03-29 | 2019-05-21 | 四川省新万兴碳纤维复合材料有限公司 | A kind of integrally formed tooling of composite material fuselage covering and its moulding process |
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2019
- 2019-03-29 CN CN201920418427.1U patent/CN209775592U/en active Active
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109774194A (en) * | 2019-03-29 | 2019-05-21 | 四川省新万兴碳纤维复合材料有限公司 | A kind of integrally formed tooling of composite material fuselage covering and its moulding process |
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