CN113146142A - Device and method for rapidly repairing cracks on hole edge of metal skin of airplane by using composite material - Google Patents
Device and method for rapidly repairing cracks on hole edge of metal skin of airplane by using composite material Download PDFInfo
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- CN113146142A CN113146142A CN202110397647.2A CN202110397647A CN113146142A CN 113146142 A CN113146142 A CN 113146142A CN 202110397647 A CN202110397647 A CN 202110397647A CN 113146142 A CN113146142 A CN 113146142A
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- 238000000034 method Methods 0.000 title claims abstract description 43
- 239000002131 composite material Substances 0.000 title claims abstract description 35
- 229910052751 metal Inorganic materials 0.000 title claims abstract description 27
- 239000002184 metal Substances 0.000 title claims abstract description 27
- 238000007789 sealing Methods 0.000 claims abstract description 9
- 238000001514 detection method Methods 0.000 claims abstract description 5
- 239000004744 fabric Substances 0.000 claims description 25
- 239000000853 adhesive Substances 0.000 claims description 18
- 230000001070 adhesive effect Effects 0.000 claims description 18
- 239000010410 layer Substances 0.000 claims description 15
- CSCPPACGZOOCGX-UHFFFAOYSA-N Acetone Chemical compound CC(C)=O CSCPPACGZOOCGX-UHFFFAOYSA-N 0.000 claims description 12
- 239000000835 fiber Substances 0.000 claims description 11
- NBIIXXVUZAFLBC-UHFFFAOYSA-N Phosphoric acid Chemical compound OP(O)(O)=O NBIIXXVUZAFLBC-UHFFFAOYSA-N 0.000 claims description 10
- 238000007743 anodising Methods 0.000 claims description 9
- XLYOFNOQVPJJNP-UHFFFAOYSA-N water Substances O XLYOFNOQVPJJNP-UHFFFAOYSA-N 0.000 claims description 9
- 230000000694 effects Effects 0.000 claims description 7
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 claims description 7
- 229920000049 Carbon (fiber) Polymers 0.000 claims description 6
- 239000004917 carbon fiber Substances 0.000 claims description 6
- 239000003365 glass fiber Substances 0.000 claims description 6
- 239000000463 material Substances 0.000 claims description 6
- 239000012790 adhesive layer Substances 0.000 claims description 5
- 229910000147 aluminium phosphate Inorganic materials 0.000 claims description 5
- 238000001035 drying Methods 0.000 claims description 5
- 239000003973 paint Substances 0.000 claims description 5
- 239000007787 solid Substances 0.000 claims description 5
- 238000004381 surface treatment Methods 0.000 claims description 5
- 238000005498 polishing Methods 0.000 claims description 4
- 239000000565 sealant Substances 0.000 claims description 4
- 239000000126 substance Substances 0.000 claims description 4
- 238000007664 blowing Methods 0.000 claims description 3
- 230000007704 transition Effects 0.000 claims description 3
- 239000002966 varnish Substances 0.000 claims description 3
- 238000011179 visual inspection Methods 0.000 claims description 3
- 238000004140 cleaning Methods 0.000 claims description 2
- 238000005507 spraying Methods 0.000 claims description 2
- 239000012528 membrane Substances 0.000 claims 2
- 239000011248 coating agent Substances 0.000 claims 1
- 238000000576 coating method Methods 0.000 claims 1
- 239000012945 sealing adhesive Substances 0.000 claims 1
- 238000009736 wetting Methods 0.000 claims 1
- 238000005553 drilling Methods 0.000 description 4
- 230000007774 longterm Effects 0.000 description 4
- 230000003014 reinforcing effect Effects 0.000 description 4
- 239000003292 glue Substances 0.000 description 3
- 229910000838 Al alloy Inorganic materials 0.000 description 2
- 244000137852 Petrea volubilis Species 0.000 description 2
- 239000002390 adhesive tape Substances 0.000 description 2
- 230000005540 biological transmission Effects 0.000 description 2
- 238000007689 inspection Methods 0.000 description 2
- 238000002955 isolation Methods 0.000 description 2
- 238000004519 manufacturing process Methods 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000005520 cutting process Methods 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 238000011065 in-situ storage Methods 0.000 description 1
- 230000000977 initiatory effect Effects 0.000 description 1
- 238000009434 installation Methods 0.000 description 1
- 238000012423 maintenance Methods 0.000 description 1
- 230000013011 mating Effects 0.000 description 1
- 238000010422 painting Methods 0.000 description 1
- 238000005086 pumping Methods 0.000 description 1
- 230000002787 reinforcement Effects 0.000 description 1
- 238000002791 soaking Methods 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
- 239000002759 woven fabric Substances 0.000 description 1
Images
Classifications
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
- B23P—METAL-WORKING NOT OTHERWISE PROVIDED FOR; COMBINED OPERATIONS; UNIVERSAL MACHINE TOOLS
- B23P6/00—Restoring or reconditioning objects
- B23P6/04—Repairing fractures or cracked metal parts or products, e.g. castings
Abstract
The invention discloses a device and a method for rapidly repairing cracks on the edge of a hole of a metal skin of an airplane by using a composite material, wherein the method comprises the following steps: the method comprises the steps of flaw detection of the cracks on the edge of the hole, designing and repairing a scheme, processing the surface of the crack part, sealing and fastening the periphery of the nail, repairing the cracks on the edge of the hole, pressurizing and curing and processing the surface of the airplane.
Description
Technical Field
The invention relates to a device and a method for repairing cracks on the edge of a metal skin hole of an airplane, in particular to a device and a method for rapidly repairing cracks on the edge of the metal skin hole of the airplane by using a composite material.
Background
In the service process of an airplane, the skin of the airplane plays an important role in maintaining the aerodynamic appearance and structure of the airplane, the actual effect of the structure of the skin of the airplane can bring serious consequences, at present, the skin and the airplane body are combined through fasteners such as rivets, screws and the like, after long-term flight, cracks are easy to generate at the hole edges due to the process of drilling and riveting and the stress of the skin, the cracks are required to be maintained to ensure normal flight, for the cracks with the length less than 5mm, crack stopping holes are generally drilled at the positions about 2mm away from the tips of the cracks to maintain, and when the cracks with the length exceeding the length are required, maintenance mechanical connection repair or composite material glue joint repair is required.
The mechanical connection repair is to rivet a metal reinforcing sheet on the skin to repair the original damaged structure after the original damaged structure is cut off, the shape and the position of the reinforcing sheet can be determined according to the crack condition, for the cracks at the edge of the raised head nail hole of the special-shaped surface, the original structure of the special-shaped surface is damaged by the drilling and riveting reinforcement after the cutting, the transmission efficiency of the reinforcing sheet is low, the manufacturing and installation operations of the reinforcing sheet need to be complained and debugged, the process is complex, the repair efficiency is low, and the effect is limited.
The composite material glue joint repairing method can be divided into two methods, one method is dry repairing, namely, the prepared pre-cured patch is glued and applied to the surface of a repairing area, but the method is not suitable for the cracks of the hole edge of the special-shaped surface nail, and firstly, the pre-cured patch is of a rigid structure, a special mould is required for manufacturing, a deformation error still exists, and the application effect is not good; secondly, the original raised head nail needs to be removed, and the process of drilling the nail can cause the uncontrollable further initiation and propagation of the crack.
The other method is wet repair, namely, the structural adhesive is used for soaking fibers to prepare a prepreg which is then attached to the surface of a metal structure, the method is particularly suitable for repairing the special-shaped surface, however, due to the raised head and the rivet, if the prepreg is directly attached, the nail can be wrapped, a blind nail is caused, and the nondestructive detection of the structure at the later stage is not facilitated; secondly, the shape of the wrapping area is complex, so that dead corners are easily caused, the bonding quality is reduced, and the transmission of the patch is not facilitated; thirdly, the repair part is difficult to remove after long-term use and failure, so a proper device and a proper method are designed to solve the problem caused by the existence of the raised head rivet and the crack of the nail hole edge of the raised head of the special profile surface repaired by the composite material wet method.
Disclosure of Invention
Technical problem to be solved by the invention
The invention aims to provide a device and a method for rapidly repairing cracks on the hole edge of an airplane metal skin by a composite material, which aim to solve the problems in the background technology, and the device and the method are suitable for solving the problems caused by the existence of a raised head rivet and the problem caused by the wet repairing of cracks on the hole edge of the raised head rivet with a special profile of the composite material.
Technical scheme
In order to achieve the purpose, the invention provides the following technical scheme: the composite material fast repairing device and method for airplane metal skin hole edge crack includes the steps of 1: flaw detection is carried out on the cracks at the hole edges; step 2: designing a repair scheme; and step 3: processing the surface of the crack part; and 4, step 4: sealing and fastening the periphery of the nail; and 5: repairing cracks at the hole edge and pressurizing and curing; and 7: and (4) treating the surface of the airplane.
Preferably, in step 1, the crack of the edge of the rivet or screw is determined by dye-penetrant inspection or visual inspection, and then the damage range and the crack length are measured by a measuring tape, and the data is recorded.
Preferably, in step 2, according to the recorded data, the material of the crack occurrence part and the material of the repaired composite material, the size of the composite material patch is designed by using a finite element method, the composite material adopts plain woven fabric carbon fiber and is soaked with high-strength structural adhesive, the paving direction is radial, a layer of glass fiber fabric is paved between the carbon fiber fabric and the metal, and the glass fiber fabric needs to be soaked by the same structural adhesive.
Preferably, in step 3, according to the area required by the composite material patching, the paint layer around the crack is polished and removed by using coarse sand paper and is cleaned by using water and acetone, then the patching area is polished by replacing fine sand paper, the polished area is brushed clean by using a brush, and the polished area is cleaned by using water and acetone.
Preferably, in step 4, structural glue is coated at the positions of the nails of the cracked metal plate for fastening, so that the nails are prevented from falling off after repair and long-term use, and then quick-drying type sealant is coated at the positions of the nails and cracks.
Preferably, in step 5, phosphoric acid anodizing adhesive is coated to perform chemical treatment on the adhesive surface, the adhesive layer is removed after the chemical treatment is completed, the surface is cleaned by water, the drying is performed by hot air blowing, the anodizing effect of the surface treatment area is observed by using an off-normal light lens, if the surface treatment area fails, polishing is performed again, then phosphoric acid anodizing treatment is performed, after the success, structural adhesive is firstly used for infiltrating the fiber fabric aiming at the raised head nail and the special-shaped surface, the layers are laid according to a repair scheme, a telescopic type hole digger is used for selecting the proper diameter of a rod of the hole digger, hole positions are reserved on the fiber fabric, the fiber fabric with the specified thickness is attached to the metal surface for attaching, after the completion, the hole positions are adjusted by the hole digger, a demolding cloth, a suction adhesive tape and an isolation film are added on the surface of the composite material, then a deformable flexible film is added on the layer to ensure the uniform pressure, an electric blanket, a breathable felt and a bag is laid, and then vacuum curing is performed.
Preferably, in step 7, after the curing is completed, the surface vacuum bag and the reverse side sealing tape are removed, the repaired edge is polished to ensure smooth transition, a layer of varnish is coated, and then the follow-up operations such as paint spraying and the like are performed according to the requirement of the repaired part of the airplane.
Preferably, the perforator comprises a first section of step, the first section of step is in a conical shape, the first section of step is in a solid structure, the outer wall of the first section of step is in sliding connection with a telescopic matching type step, the outer wall of the telescopic matching type step is in sliding connection with a second section of step, and the second section of step is in a closed cylindrical structure.
Advantageous effects
Compared with the prior art, the invention has the beneficial effects that: the operation method can be used for repairing in situ on the airplane, is simple and easy to operate, has high speed, effectively slows down crack propagation, does not need to drill and rivet the airplane body and the repaired composite material, does not need to remove original nails, avoids secondary damage to the airplane body and the cracks, effectively strengthens the hole edges due to local concentration of fibers at the hole edges, further improves the bearing capacity of the structure and prolongs the flight life of the airplane.
Drawings
FIG. 1 is a schematic flow-through of the process of the present invention;
fig. 2 is a schematic structural diagram of the perforator of the present invention.
In the figure: 1. a perforator; 11. a first stage step; 12. a scalable mating type step; 13. and a second step section.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. The components of embodiments of the present invention generally described and illustrated in the figures herein may be arranged and designed in a wide variety of different configurations. Thus, the following detailed description of the embodiments of the present invention, presented in the figures, is not intended to limit the scope of the invention, as claimed, but is merely representative of selected embodiments of the invention. All other embodiments, which can be derived by a person skilled in the art from the embodiments of the invention without making creative efforts, shall fall within the protection scope of the invention.
In an embodiment, referring to fig. 1-2, the present invention provides a technical solution: the composite material fast repairing device and method for airplane metal skin hole edge crack includes the steps of 1: flaw detection is carried out on the cracks at the hole edges; step 2: designing a repair scheme; and step 3: processing the surface of the crack part; and 4, step 4: sealing and fastening the periphery of the nail; and 5: repairing cracks at the hole edge and pressurizing and curing; and 7: and (4) treating the surface of the airplane.
Referring to fig. 1, in step 1, the edge crack of the rivet or screw is determined by dye-penetrant inspection or visual inspection, and then the damage range and the crack length are measured by a measuring tape, and the data is recorded.
In the embodiment, referring to fig. 1, in step 2, according to the recorded data, the material of the crack occurrence portion and the material of the repaired composite material, the size of the composite material patch is designed by using a finite element method, the composite material adopts plain woven carbon fiber CF3031, the J-349 high-strength structural adhesive is soaked, the paving direction is radial, in order to prevent the metal and the carbon fiber part from being corroded due to the potential difference, a layer of glass fiber fabric is paved between the carbon fiber fabric and the metal, and the glass fiber fabric needs to be soaked by the same structural adhesive.
In step 3, according to the area required for the composite material to be attached, the paint layer around the crack is removed by polishing with coarse sandpaper and cleaning with water and acetone, then the attached area is polished with 100-mesh or 80-mesh fine sandpaper, the polished area is brushed clean with a brush, and the cleaned area is cleaned with water and acetone.
In step 4, the nails are prevented from dropping after repair and after long-term use, and then quick-drying sealant, such as hangao 413, is coated on the nails and crack gaps, and for the accessible parts, a sealing tape can be pasted on the back side, and the front side does not need to be coated with the sealant.
In the embodiment, referring to fig. 1-2, in step 5, AC-130 paste phosphoric acid anodizing adhesive is coated to chemically treat the adhesive surface, after that, the adhesive layer is removed, and is cleaned with water, dried by hot air blowing, the anodizing effect of the surface treatment area is observed by using an off-normal light lens, if the surface treatment area fails, the anodizing treatment is performed by polishing again, after the success, the fiber fabric is infiltrated by structural adhesive according to the raised head nail and the special-shaped surface, and is layered according to the repair scheme, a telescopic type hole drawer is used, a proper hole drawer rod diameter (0.1 mm larger than the nail hole diameter) is selected, a hole position is reserved on the fiber fabric, the fiber fabric with a specified thickness is attached to the metal surface fabric, after that, the hole position is adjusted by the hole drawer, a layer of demoulding cloth, an adhesive tape and an isolation film is added on the surface of the composite material, and then a layer of a deformable flexible film is added to ensure uniform pressure, laying electric blanket, air-permeable felt, sealing bag, and vacuum-pumping for curing.
In step 7, referring to fig. 1, after the curing is completed, the surface vacuum bag and the reverse side sealing tape are removed, the repaired edge is polished to ensure smooth transition, a layer of varnish is applied, and the subsequent operations such as painting are performed according to the requirement of the repaired part of the airplane.
In an embodiment, referring to fig. 2, the hole digger comprises a first section of step, the first section of step is in a conical shape and is inserted into a wet prepreg conveniently, the first section of step is in a solid structure, the solid structure is not easy to deform, the hole digger can be lengthened by self weight, the outer wall of the first section of step is connected with a telescopic matching type step in a sliding mode, according to the size of a conventional raised head nail of an airplane, 3mm is taken as the minimum diameter, one step is designed at intervals of 1mm, the minimum step height is 10mm, the step height is increased by 2mm along with the increase of the number of steps, the outer wall of the telescopic matching type step is connected with a second section of step in a sliding mode, the second section of step is in a closed cylindrical structure, except that the first section of step is in an aluminum alloy solid structure, the rest are aluminum alloy hollow pipes, each step part is matched by a miniature raised semi-circular column, one side of the second section of step is provided with a rotating shaft which can be rotated to deepen the drilling depth of the hole guiding device, the diameter of the step can be freely selected according to the actual size of the raised head nail through stretching and compressing, and then the hole guiding device is matched with a conical drill bit for use, so that the hole guiding positioning can be realized.
The above description is only for the specific embodiments of the present invention, but the scope of the present invention is not limited thereto, and any person skilled in the art can easily conceive of the changes or substitutions within the technical scope of the present invention.
Claims (8)
1. The method for quickly repairing the cracks on the hole edge of the metal skin of the airplane by using the composite material is characterized by comprising the following steps of: step 1: flaw detection is carried out on the cracks at the hole edges; step 2: designing a repair scheme; and step 3: processing the surface of the crack part; and 4, step 4: sealing and fastening the periphery of the nail; and 5: repairing cracks at the hole edge and pressurizing and curing; and 7: and (4) treating the surface of the airplane.
2. The method for rapidly repairing cracks on the hole edge of the metal skin of the airplane by using the composite material as claimed in claim 1, wherein the method comprises the following steps: in step 1, the edge cracks of the rivet or the screw are determined by adopting a dye check method or a visual inspection method, then the damage range and the crack length are measured by using a measuring tape, and data are recorded.
3. The method for rapidly repairing cracks on the hole edge of the metal skin of the airplane by using the composite material as claimed in claim 1, wherein the method comprises the following steps: in step 2, according to the recorded data, the material of the crack occurrence part and the material of the repaired composite material, the size of the composite material patch is designed by using a finite element method, the composite material adopts plain weave fabric carbon fiber to soak high-strength structural adhesive, the paving direction is radial, a layer of glass fiber fabric is paved between the carbon fiber fabric and metal, and the glass fiber fabric needs to be soaked by the same structural adhesive.
4. The method for rapidly repairing cracks on the hole edge of the metal skin of the airplane by using the composite material as claimed in claim 1, wherein the method comprises the following steps: in step 3, according to the area required by the composite material patching, a paint layer at the periphery of the crack is polished and removed by using coarse abrasive paper and is cleaned by using water and acetone, then the patching area is polished by replacing fine abrasive paper, the polished area is brushed clean by using a brush, and the polished area is cleaned by using water and acetone.
5. The method for rapidly repairing cracks on the hole edge of the metal skin of the airplane by using the composite material as claimed in claim 1, wherein the method comprises the following steps: in the step 4, structural adhesive is coated at the positions of the cracked metal plate nails for fastening, so that the nails are prevented from falling off after being repaired and used for a long time, and then quick-drying sealant is coated at the positions of the nails and cracks.
6. The method for rapidly repairing cracks on the hole edge of the metal skin of the airplane by using the composite material as claimed in claim 1, wherein the method comprises the following steps: in step 5, coating phosphoric acid anodizing adhesive to carry out chemical treatment on the adhesive surface, removing the adhesive layer after the chemical treatment, cleaning the adhesive layer by using water, drying the adhesive layer by using hot air blowing, observing the anodizing effect of the surface treatment area by using an off-normal light lens, polishing again if the anodic treatment fails, then carrying out phosphoric acid anodizing treatment, wetting fiber fabrics by using structural adhesive according to the raised head nails and the special-shaped surfaces, laying layers according to a repair scheme, using a retractable type hole leading device, selecting a proper diameter of a rod of the hole leading device, reserving hole positions on the fiber fabrics, attaching the fiber fabrics with a specified thickness to the metal surface fabrics, adjusting the hole positions by using the hole leading device after the hole leading device is completed, covering demolding fabrics, adhesive absorbing fabrics and isolating membranes on the surface of the composite material, then covering a layer of deformable flexible membrane to ensure uniform pressure, laying an electric blanket, a breathable felt and a sealing bag, and then vacuumizing and curing.
7. The method for rapidly repairing cracks on the hole edge of the metal skin of the airplane by using the composite material as claimed in claim 1, wherein the method comprises the following steps: in step 7, after the curing is completed, the surface vacuum bag and the reverse side sealing adhesive tape are removed, the repaired edge is polished to ensure smooth transition, a layer of varnish is coated, and then the follow-up operations such as paint spraying and the like are carried out according to the requirement of the repaired part of the airplane.
8. The composite material quick repair airplane metal skin hole edge crack device according to claim 6, characterized in that: the perforator comprises a first section step, wherein the first section step is in a conical shape, the first section step is of a solid structure, the outer wall of the first section step is in sliding connection with a telescopic matching type step, the outer wall of the telescopic matching type step is in sliding connection with a second section step, and the second section step is of a closed cylindrical structure.
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110397647.2A CN113146142A (en) | 2021-04-14 | 2021-04-14 | Device and method for rapidly repairing cracks on hole edge of metal skin of airplane by using composite material |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110397647.2A CN113146142A (en) | 2021-04-14 | 2021-04-14 | Device and method for rapidly repairing cracks on hole edge of metal skin of airplane by using composite material |
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| CN113146142A true CN113146142A (en) | 2021-07-23 |
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| CN202110397647.2A Pending CN113146142A (en) | 2021-04-14 | 2021-04-14 | Device and method for rapidly repairing cracks on hole edge of metal skin of airplane by using composite material |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114749867A (en) * | 2022-05-04 | 2022-07-15 | 国营芜湖机械厂 | Repair method suitable for high-temperature titanium alloy structure cracks |
| CN114851591A (en) * | 2022-04-12 | 2022-08-05 | 国营芜湖机械厂 | Method for repairing metal structure crack by using thermoplastic carbon fiber composite material |
| CN114888521A (en) * | 2022-05-07 | 2022-08-12 | 国营芜湖机械厂 | Composite metal fatigue crack inhibiting method |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN114749867A (en) * | 2022-05-04 | 2022-07-15 | 国营芜湖机械厂 | Repair method suitable for high-temperature titanium alloy structure cracks |
| CN114888521A (en) * | 2022-05-07 | 2022-08-12 | 国营芜湖机械厂 | Composite metal fatigue crack inhibiting method |
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