CN118219591A - Repair method for aircraft fairing - Google Patents
Repair method for aircraft fairing Download PDFInfo
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- CN118219591A CN118219591A CN202410506605.1A CN202410506605A CN118219591A CN 118219591 A CN118219591 A CN 118219591A CN 202410506605 A CN202410506605 A CN 202410506605A CN 118219591 A CN118219591 A CN 118219591A
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- Prior art keywords
- repair
- repairing
- paving
- honeycomb
- area
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- 238000000034 method Methods 0.000 title claims abstract description 25
- 239000000835 fiber Substances 0.000 claims abstract description 34
- 239000012528 membrane Substances 0.000 claims abstract description 32
- 238000000926 separation method Methods 0.000 claims abstract description 32
- 238000010438 heat treatment Methods 0.000 claims abstract description 28
- 229920000742 Cotton Polymers 0.000 claims abstract description 24
- 239000002313 adhesive film Substances 0.000 claims abstract description 21
- 238000005498 polishing Methods 0.000 claims abstract description 8
- 239000010410 layer Substances 0.000 claims description 27
- 239000012790 adhesive layer Substances 0.000 claims description 11
- 230000000694 effects Effects 0.000 abstract description 10
- 239000000463 material Substances 0.000 abstract description 7
- 239000003292 glue Substances 0.000 abstract description 4
- 239000011347 resin Substances 0.000 description 8
- 229920005989 resin Polymers 0.000 description 8
- 238000009423 ventilation Methods 0.000 description 5
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 238000006467 substitution reaction Methods 0.000 description 2
- 241001133184 Colletotrichum agaves Species 0.000 description 1
- 239000002390 adhesive tape Substances 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000004140 cleaning Methods 0.000 description 1
- 238000004891 communication Methods 0.000 description 1
- 238000005056 compaction Methods 0.000 description 1
- 238000010276 construction Methods 0.000 description 1
- 230000032798 delamination Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 238000002474 experimental method Methods 0.000 description 1
- 239000003365 glass fiber Substances 0.000 description 1
- 238000007373 indentation Methods 0.000 description 1
- 238000012423 maintenance Methods 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 239000003973 paint Substances 0.000 description 1
- 238000007711 solidification Methods 0.000 description 1
- 230000008023 solidification Effects 0.000 description 1
- 241000894007 species Species 0.000 description 1
- 238000005507 spraying Methods 0.000 description 1
Landscapes
- Laminated Bodies (AREA)
Abstract
The invention relates to the technical field of aircraft repair, and discloses a repair method of an aircraft fairing, wherein the fairing comprises a honeycomb plate and skins arranged at the upper end and the lower end of the honeycomb plate, and a vacuum bag is adopted, and the repair method comprises the following steps: s1, removing the honeycomb plate of the damaged part and skins at the upper end and the lower end of the honeycomb plate, wherein the removed area of the honeycomb plate is marked as a repair area; s2, polishing skins at the upper end and the lower end of the repairing area, and polishing repairing steps respectively; s3, preparing a fiber layer, a glue film and a honeycomb insert; s4, mounting the honeycomb insert in a repairing area, and sequentially paving adhesive films and fiber layers on the repairing steps on the upper side and the lower side; s5, sequentially paving a first non-porous separation membrane, a mold backboard, a first heating blanket and first breathable cotton on a fiber paving layer positioned at the upper end of the repair area; s6, sequentially paving a hole separation membrane, a suction layer, a second non-porous separation membrane, a second heating blanket and second breathable cotton on the fiber paving layer positioned at the lower end of the repair area; and S7, covering vacuum bags above the first air-permeable cotton and below the second air-permeable cotton respectively, and curing. The repair method of the aircraft fairing provided by the invention can reduce paving auxiliary materials, simplify paving steps and ensure the forming effect of the surface of the skin.
Description
Technical Field
The invention relates to the technical field of aircraft repair, in particular to a repair method of an aircraft fairing.
Background
To reduce the weight of an aircraft, aircraft fairing parts are typically designed as honeycomb sandwich structures, and if both sides of the structure are damaged, a double-sided repair method is generally adopted during maintenance.
In the prior art, when auxiliary materials are paved on the upper and lower side skins, the upper and lower sides are sequentially paved with the porous stripping film, the glass fiber adhesive tape and the non-porous separating film, so that the curing integrity and operability are ensured, but the paving steps are complicated, the paving auxiliary materials are too many, the surface forming effect is ignored, the follow-up paint spraying work is not facilitated, and the good pneumatic appearance cannot be maintained.
Disclosure of Invention
The purpose of the invention is that: the repair method of the aircraft fairing can reduce paving auxiliary materials, simplify paving steps and ensure the forming effect of the surface of the skin.
In order to achieve the above object, the present invention provides a repair method for an aircraft fairing, where the fairing includes a honeycomb panel and skins disposed at upper and lower ends of the honeycomb panel, and a vacuum bag is used, and the repair method includes the following steps:
S1, removing the honeycomb plate of the damaged part and the skins at the upper end and the lower end of the honeycomb plate, wherein the removed area of the honeycomb plate is marked as a repair area;
S2, polishing the skins at the upper end and the lower end of the repair area, and polishing repair steps respectively;
s3, preparing a fiber layer, a glue film and a honeycomb insert;
s4, installing the honeycomb plug in the repair area, and sequentially paving the adhesive film and the fiber paving layer on the repair steps at the upper end and the lower end of the honeycomb plug;
S5, sequentially paving a first non-porous separation membrane, a mold backboard, a first heating blanket and first breathable cotton on the fiber paving layer positioned at the upper end of the repair area;
S6, sequentially paving a hole separation membrane, a glue absorbing layer, a second non-hole separation membrane, a second heating blanket and second breathable cotton on the fiber paving layer positioned at the lower end of the repair area;
And S7, respectively covering the vacuum bag above the first air-permeable cotton and below the second air-permeable cotton, and heating, pressurizing and curing.
Further, in S4, when the repair steps at the upper end of the honeycomb plug are laid, the fiber lay and the adhesive film lay are heated and compacted by using a heat gun.
Further, the heating temperature of the heat gun is 60-70 ℃.
Further, in S5, S6, a plurality of thermocouples are disposed at intervals on the skin around the repair area, and the thermocouples are all covered in the vacuum bag.
Further, in S7, during heating, pressurizing and curing, the pressure in the vacuum bag at the lower side is 21.5-22.5 in Hg, and the pressure in the vacuum bag at the upper side is 19.5-20.5 in Hg.
Further, after S2, cleaning and dehumidifying the repair area and the repair step.
Further, the first non-porous separation membrane is structurally identical to the second non-porous separation membrane.
Compared with the prior art, the repair method of the aircraft fairing has the beneficial effects that: sequentially paving a first non-porous separation membrane, a mold back plate, a first heating blanket and first breathable cotton on a fiber paving layer positioned at the upper end of the repair area; the hole separating film, the suction adhesive layer, the second non-hole separating film, the second heating blanket and the second ventilation cotton are sequentially paved on the fiber paved layer at the lower end of the repairing area, the paving of the hole separating film and the suction adhesive layer is reduced at the upper end of the repairing area, paving auxiliary materials are reduced, paving steps are simplified, and when the repairing area is heated, pressurized and solidified, the hole separating film and the suction adhesive layer are not paved on the fiber paved layer at the upper end of the repairing area, so that resin overflowed after the adhesive film at the position is softened in the solidifying process is not sucked, but is partially diffused upwards to the surface of the skin, and the rest of the resin is diffused downwards to the honeycomb insert block, so that the surface of the skin at the upper side is smoother, and the forming effect of the surface of the skin is ensured.
Drawings
FIG. 1 is a flow chart of a method of repairing an aircraft fairing in accordance with an embodiment of the invention;
FIG. 2 is an overall block diagram of a method of repairing an aircraft fairing in accordance with an embodiment of the invention;
In the figure, 1, a fiber layer;
2. an adhesive film;
3. a honeycomb plug;
4. a first non-porous separation membrane;
5. A mold back plate;
6. A first heating blanket;
7. A first air-permeable cotton;
8. a thermocouple;
9. a vacuum bag;
10. A porous separation membrane;
11. a glue sucking layer;
12. a second non-porous separation membrane;
13. A second heating blanket;
14. And second breathable cotton.
Detailed Description
The following describes in further detail the embodiments of the present invention with reference to the drawings and examples. The following examples are illustrative of the invention and are not intended to limit the scope of the invention.
In the description of the present invention, terms such as "upper", "lower", "left", "right", "front", "rear", "inner", "outer", "transverse", "longitudinal", and the like are used for convenience in describing the present invention and simplifying the description only, and are not intended to limit the present invention to the specific orientations or configurations and operations of the indicated devices, elements or components, and thus should not be construed as limiting the present invention. The specific meaning of these terms in the present invention will be understood by those of ordinary skill in the art according to the specific circumstances.
In the description of the present invention, the terms "provided," "disposed," "connected," and "disposed" are to be construed broadly, and may be, for example, fixedly connected, detachably connected, or of unitary construction; may be a mechanical connection, or an electrical connection; may be directly connected, or indirectly connected through intervening media, or may be in internal communication between two devices, elements, or components. The specific meaning of the above terms in the present invention can be understood by those of ordinary skill in the art according to the specific circumstances.
Furthermore, the terms "first," "second," and the like, are used primarily to distinguish between different devices, elements, or components (the particular species and configurations may be the same or different), and are not used to indicate or imply the relative importance and number of devices, elements, or components indicated. Unless otherwise indicated, the meaning of "a plurality" is two or more.
The technical scheme of the invention is further described below with reference to the embodiment and the attached drawings.
As shown in fig. 1, in a repair method for an aircraft fairing according to an embodiment of the present invention, the fairing includes a honeycomb panel and skins disposed at upper and lower ends of the honeycomb panel, and a vacuum bag 9 is used, including the steps of:
S1, removing the honeycomb plate of the damaged part and skins at the upper end and the lower end of the honeycomb plate, wherein the removed area of the honeycomb plate is marked as a repair area;
S2, polishing skins at the upper end and the lower end of the repairing area, and polishing repairing steps respectively;
s3, preparing a fiber layer 1, an adhesive film 2 and a honeycomb insert 3;
S4, mounting the honeycomb plug 3 in a repairing area, and sequentially paving adhesive films 2 and fiber layers 1 on repairing steps at the upper end and the lower end of the honeycomb plug 3;
s5, sequentially paving a first non-porous separation membrane 4, a die backboard 5, a first heating blanket 6 and first breathable cotton 7 on the fiber paving layer 1 positioned at the upper end of the repair area;
S6, sequentially paving a hole separation membrane 10, a suction adhesive layer 11, a second non-porous separation membrane 12, a second heating blanket 13 and second breathable cotton 14 on the fiber pavement 1 positioned at the lower end of the repair area;
and S7, covering the vacuum bags 9 above the first air-permeable cotton 7 and below the second air-permeable cotton 14 respectively, and performing heating, pressurizing and curing.
Based on the technical scheme, a first non-porous separation membrane 4, a die back plate 5, a first heating blanket 6 and first ventilation cotton 7 are sequentially paved on a fiber paving layer 1 positioned at the upper end of a repair area; the hole separation membrane 10, the suction adhesive layer 11, the second non-porous separation membrane 12, the second heating blanket 13 and the second ventilation cotton 14 are sequentially paved on the fiber paved layer 1 positioned at the lower end of the repairing area, the paving of the hole separation membrane 10 and the suction adhesive layer 11 is reduced at the upper end of the repairing area, paving auxiliary materials are reduced, paving steps are simplified, and when the repairing area is heated, pressurized and solidified, the hole separation membrane 10 and the suction adhesive layer 11 are not paved on the fiber paved layer 1 positioned at the upper end of the repairing area, so that resin overflowed after softening the adhesive film 2 at the position in the solidifying process is not sucked, but is partially diffused upwards to the surface of the skin, and the rest of the resin is diffused downwards to the honeycomb inserting block 3, so that the surface of the skin positioned at the upper side is smoother, and the forming effect of the surface of the skin is ensured.
Preferably, in S4, when the repair steps at the upper end of the honeycomb plug 3 are laid, the fiber lay-up 1 and the adhesive film 2 are both heated and compacted by using a heat gun after being laid. The repairing ladder at the upper end of the honeycomb inserting block 3 is paved, redundant resin after the adhesive film 2 is softened can be uniformly distributed on the surface of the skin positioned at the upper side after being compacted, and the smoothness of the skin is improved.
More preferably, the heating temperature of the heat gun is 60-70 ℃. The heating temperature of the hot air gun is preferably 65 ℃, when the heating temperature of the hot air gun is less than 60 ℃, the softening effect of the adhesive film 2 is poor, even if the compaction operation is carried out, a large gap is likely to exist between the fiber layering layer 1 and the adhesive film 2, the porosity is difficult to control, and the problem of insufficient structural strength after repair is easily caused due to high porosity; when the heating temperature of the hot air gun is higher than 70 ℃, the fiber paving layer 1 is easy to solidify in advance, and the hot air gun is operated, so that solidification is uneven, the adhesive film 2 is excessively softened and is unfavorable for subsequent paving operation, so that when the heating temperature of the hot air gun is 60-70 ℃, the porosity can be reduced, the bonding degree between the fiber paving layer 1 and the adhesive film 2 is ensured, the fiber paving layer 1 is prevented from solidifying in advance, the adhesive film 2 is prevented from being excessively softened, the subsequent paving operation is influenced, and the paving effect is good.
Preferably, in S5, S6, a plurality of thermocouples 8 are arranged at intervals on the skin around the repair area, and the thermocouples 8 are each covered in a vacuum bag 9. The thermocouples 8 are arranged around the repairing area, and as the heating blankets are paved above and below the repairing area, the repairing area can be heated in all directions during the curing operation, so that the fiber pavement 1 is cured more uniformly, the adhesive film 2 is softened uniformly, and the repairing effect is enhanced.
Preferably, in S7, the pressure in the vacuum bag 9 located at the lower side is 21.5-22.5 in Hg and the pressure in the vacuum bag 9 located at the upper side is 19.5-20.5 in Hg during the heat-pressure curing. The pressure difference in the vacuum bags 9 on the upper side and the lower side is controlled to be 1-3 in Hg, so that the exhaust effect of gas generated by the adhesive film 2 on the lower side and overflowing resin can be ensured, the possibility of indentation of the skin on the upper side is reduced, the pressure difference is obtained by actual production and experiments, and if the pressure difference is larger, the lower side is required to reach higher pressure or the pressure on the upper side is lower. The lower side reaches higher pressure and needs stronger vacuum air source equipment, and is difficult to stably reach, and the lower pressure of the upper side can lead to the fact that the mould backplate 5 can not laminate completely, leads to bonding failure or delamination easily.
Preferably, after S2, the repair area and the repair steps are cleaned and dehumidified.
Preferably, the first non-porous separation membrane 4 is identical in structure to the second non-porous separation membrane 12.
In summary, the embodiment of the invention provides a repair method for an aircraft fairing, which comprises the steps of sequentially paving a first non-porous separation membrane 4, a mold back plate 5, a first heating blanket 6 and first ventilation cotton 7 on a fiber layer 1 positioned at the upper end of a repair area; the hole separation membrane 10, the suction adhesive layer 11, the second non-porous separation membrane 12, the second heating blanket 13 and the second ventilation cotton 14 are sequentially paved on the fiber paved layer 1 positioned at the lower end of the repairing area, the paving of the hole separation membrane 10 and the suction adhesive layer 11 is reduced at the upper end of the repairing area, paving auxiliary materials are reduced, paving steps are simplified, and when the repairing area is heated, pressurized and solidified, the hole separation membrane 10 and the suction adhesive layer 11 are not paved on the fiber paved layer 1 positioned at the upper end of the repairing area, so that resin overflowed after softening the adhesive film 2at the position in the solidifying process is not sucked, but is partially diffused upwards to the surface of the skin, and the rest of the resin is diffused downwards to the honeycomb inserting block 3, so that the surface of the skin positioned at the upper side is smoother, and the forming effect of the surface of the skin is ensured.
The foregoing is merely a preferred embodiment of the present invention, and it should be noted that modifications and substitutions can be made by those skilled in the art without departing from the technical principles of the present invention, and these modifications and substitutions should also be considered as being within the scope of the present invention.
Claims (7)
1.A method for repairing a fairing of an aircraft, said fairing comprising a honeycomb panel and skins arranged at the upper and lower ends of said honeycomb panel, a vacuum bag (9) being used, characterized in that it comprises the steps of:
S1, removing the honeycomb plate of the damaged part and the skins at the upper end and the lower end of the honeycomb plate, wherein the removed area of the honeycomb plate is marked as a repair area;
S2, polishing the skins at the upper end and the lower end of the repair area, and polishing repair steps respectively;
s3, preparing a fiber layer (1), an adhesive film (2) and a honeycomb insert (3);
S4, installing the honeycomb plug (3) in the repairing area, and sequentially paving the adhesive film (2) and the fiber paving layer (1) on the repairing steps at the upper end and the lower end of the honeycomb plug (3);
s5, sequentially paving a first non-porous separation membrane (4), a mold back plate (5), a first heating blanket (6) and first breathable cotton (7) on the fiber paving layer (1) positioned at the upper end of the repair area;
S6, sequentially paving a hole separation membrane (10), a suction adhesive layer (11), a second non-porous separation membrane (12), a second heating blanket (13) and second breathable cotton (14) on the fiber paving layer (1) positioned at the lower end of the repair area;
And S7, respectively covering the vacuum bag (9) above the first air-permeable cotton (7) and below the second air-permeable cotton (14), and performing heating, pressurizing and curing.
2. The method for repairing an aircraft fairing according to claim 1, wherein in S4, when the repair steps at the upper end of the honeycomb insert (3) are laid, the fiber mat (1) and the adhesive film (2) are heated and compacted by a heat gun after being laid.
3. The method of repairing an aircraft fairing according to claim 2, wherein the heating temperature of the heat gun is 60-70 ℃.
4. Method for repairing an aircraft fairing according to claim 1, characterized in that in S5, S6 a plurality of thermocouples (8) are arranged at intervals on the skin around the repair area, and that the thermocouples (8) are each covered in the vacuum bag (9).
5. The method for repairing an aircraft fairing according to claim 1, characterized in that in S7, the pressure in the vacuum bag (9) located on the lower side is 21.5-22.5in Hg and the pressure in the vacuum bag (9) located on the upper side is 19.5-20.5in Hg when heat and pressure are applied and cured.
6. The method for repairing an aircraft fairing according to claim 1, wherein after S2, the repair area and the repair steps are cleaned and dehumidified.
7. The aircraft fairing repair method according to claim 1, characterized in that the first non-porous separation membrane (4) is structurally identical to the second non-porous separation membrane (12).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202410506605.1A CN118219591A (en) | 2024-04-25 | 2024-04-25 | Repair method for aircraft fairing |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202410506605.1A CN118219591A (en) | 2024-04-25 | 2024-04-25 | Repair method for aircraft fairing |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN118219591A true CN118219591A (en) | 2024-06-21 |
Family
ID=91504830
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202410506605.1A Pending CN118219591A (en) | 2024-04-25 | 2024-04-25 | Repair method for aircraft fairing |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN118219591A (en) |
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2024
- 2024-04-25 CN CN202410506605.1A patent/CN118219591A/en active Pending
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