Method for repairing structural part of aircraft with protective defects and aircraft structural part repaired thereby
By introducing isolation devices on the aircraft structural parts, the protection defects caused by paint damage are solved, the risks of corrosion and lightning strikes are reduced, and the impact of paint drying time on the production cycle is avoided.
Patent Information
- Application Number
- CN202411189621.9
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Priority Date
- 2023-08-31
- Filing Date
- 2024-08-28
- Publication Date
- 2025-02-28
AI Technical Summary
The paint of aircraft structural parts is easily damaged during the installation of fastening components, resulting in protective defects, increasing the risk of corrosion and lightning strikes, and a long drying time is required during the repair process, which affects the production cycle.
Introduce an isolation device to position around the orifice of the fastening device to avoid direct contact between the fastening device and the structural parts, and apply a paint coating on the isolation device to ensure that the drying time of the paint does not affect the installation of the fastening device.
The isolation barrier between the fastening device and structural parts is effectively restored, the risk of electrochemical corrosion is reduced, the electromagnetic hazard requirements are met, and the production cycle is avoided due to the paint drying time.
Smart Images
Figure CN119525912A_ABST
Abstract
Description
Technical Field
[0001] The present application relates to a method for repairing a protection defect, such as a paint defect, present in a region of a structural part of an aircraft, said region of the structural part having, in the vicinity of said protection defect, an orifice for receiving a fastening device. The present application also relates to an aircraft structural part subjected to such a repair method. Background Art
[0002] As is known, elements of structures of an aircraft, such as a wing or a central wing box of an aircraft, made of metal (for example an aluminium alloy), which form an interface with a fuel tank of said aircraft, are subjected to surface treatment and subsequently covered with at least one paint coat (anodised coating, primer coating and topcoat coating) in order to provide them with active protection against corrosion and an isolation function against the risk of sparking in the boundary area with the fuel tank.
[0003] The different parts of the structure of an aircraft are assembled to one another using fastening elements in order to form the structure of the aircraft.
[0004] During the manufacture of metal or aluminum alloy structural parts of an aircraft, or more generally during the assembly of the different parts that form the structure of an aircraft, damage may occur to the paint of the structural parts, such as scratches or impacts, caused in particular by drilling tools, or during operations of installing / removing (temporary or permanent) fasteners. This damage must be repaired so that the entire structure is protected, in particular against corrosion.
[0005] In order to repair such damage to the paint, the damaged surface of the structural part of the aircraft is first sanded in order to smooth the surface of the structural part and ensure that the surface is free of cracks, and then a surface treatment may be applied to the sanded surface in order to allow improved adhesion of the paint. A first primer coat (containing a corrosion inhibitor) is applied to the treated surface in order to provide the structural part with anti-corrosion properties, and finally, optionally, a second topcoat coat is applied to the surface of the damaged area of the structural part. This protection system ensures protection against corrosion and the isolation function for the fastening elements required in the event of sparks during a lightning strike. Once these steps have been carried out, the protection defect is no longer visible and the repair is complete.
[0006] However, drying times must be observed between the surface treatment and the primer coating, and then between the primer coating and the topcoat coating. Drying times must also be observed after the last coat of paint has been applied in order to be able to subsequently mount the fasteners on the structure of the aircraft. These drying times are necessary for the polymerization of the resins contained in the paint applied to the structure. However, these drying times are relatively long (more than 24 hours). If these drying times are not observed, the anti-corrosion and anti-lightning properties of the structure of the aircraft cannot be guaranteed.
[0007] When damage to the paint of the structural part occurs in the vicinity of the fastening element, there is a risk of cracking or crazing of the paint if the fastening element is mounted on a paint without observing drying time, which results in the risk that the metal material of the fastening element will come into contact with the metal material forming the substrate of the structural part. This results in an increased risk of galvanic corrosion between the fastening element and the substrate of the structural part and non-compliance with the electromagnetic hazard (EMH) requirements. The isolation barrier between the structural part of the aircraft and the fastening element, which is usually created by a different paint coating, cannot be guaranteed.
[0008] Furthermore, the repair of paint damage to structures is often carried out in limited areas with limited access, which makes the implementation of accelerated drying measures complicated or even impossible. Drying at ambient temperature requires dozens of hours, which has a significant impact on production cycles.
[0009] The present invention aims to overcome these disadvantages. Summary of the invention
[0010] The present invention aims to propose a solution which, by introducing isolating means, makes it possible to restore the isolating barrier between the metal material forming the means for fastening two aircraft structural parts to each other and the metal material forming said aircraft structural parts.
[0011] To this end, the invention relates to a method for repairing a region of a metal structural part of an aircraft having a protection defect, said region of the structural part having a first orifice for receiving a metal fastening device, said first orifice having a first diameter. The repair method comprises:
[0012] - a sanding step during which the paint in said area of the aircraft structural part is sanded.
[0013] According to the present invention, the repair method further comprises:
[0014] - a step of positioning an isolating device, during which an isolating device is positioned around said first orifice on said polished area of the aircraft structural part, said isolating device comprising a second orifice having a second diameter substantially equal to the first diameter of the first orifice, said isolating device being positioned around the first orifice on said polished area of the aircraft structural part such that said first orifice and said second orifice are coaxial,
[0015] - a step of applying at least one paint coat, during which a primer coat is applied to the sanded areas of the aircraft structural part, said primer coat being applied to said insulation means.
[0016] Advantageously, the introduction of the isolating means makes it possible to avoid direct contact between the fastening means (made of metal material) intended to be inserted into the first orifice and the aircraft structural part (also made of metal material, such as aluminium alloy).
[0017] According to one feature, the repair method comprises, before the step of positioning the isolating device, a step of inserting the fastening device, during which step the fastening device is inserted into the second orifice of the isolating device, and during the step of positioning the isolating device, the fastening device is inserted into the first orifice, and during the step of applying a paint coating, said paint coating is applied to said fastening device.
[0018] Advantageously, the fastening means are mounted so that they are in direct contact with the insulating means and no longer in contact with regions of the aircraft structure part, which makes it possible to avoid contact between the metal material of the fastening means and the metal material forming the aircraft structure.
[0019] According to another feature, the fastening device is in the form of a shank having a first end and a second end and a head at the first end, the fastening device being inserted into the first and second apertures through the second end of the shank, the outer diameter of the isolating device being greater than the diameter of the head of the fastening device.
[0020] According to this feature, during the step of inserting the fastening means, the head of said fastening means bears against the isolating means.
[0021] According to another feature, the repair method comprises, before the step of applying the paint coating, a step of inserting fastening means, during which the fastening means are inserted into the first orifice and into the second orifice of the isolation means, and during the step of applying the paint coating, said paint coating is applied to said fastening means.
[0022] According to one feature, the repair method comprises: a step of positioning a masking device before the step of applying a paint coating, during which the masking device is positioned in the first orifice and the second orifice; and a step of removing the masking device after the step of applying a paint coating, during which the masking device is removed from the first orifice and the second orifice; and a step of inserting a fastening device, during which the fastening device is inserted into the first orifice and the second orifice.
[0023] According to one feature, during the step of inserting the fastening means, the head of said fastening means bears against the paint coating.
[0024] Advantageously, the introduction of the insulating means makes it possible to eliminate the paint drying time before the fastening means are installed. The insulating means makes it possible to avoid direct contact between the metal material of the fastening means and the metal material forming the aircraft structure, even in the event of cracks or crazing of the paint applied between the head of the fastening means and the insulating means.
[0025] According to another feature, the repair method comprises, after the grinding step, a surface treatment step during which a chemical conversion coating or a reactive primer is applied to at least a portion of the ground area of the aircraft structural part.
[0026] The invention also relates to an aircraft structural component, comprising a first metal structural part and a second structural part, the first structural part having at least a first aperture for receiving a fastening device, the first aperture having a first diameter, the first structural part and the second structural part being fastened to each other by means of a metal fastening device inserted into the first aperture.
[0027] According to the present invention, the aircraft structure also includes an isolation device, which has a second opening, a second diameter of the second opening is substantially equal to the first diameter of the first opening, the isolation device is positioned around the first opening on the first structural part so that the first opening and the second opening are coaxial, and a fastening device is inserted into the second opening and the first opening.
[0028] The invention therefore relates to an aircraft metal structural part having a region repaired by means of the repair method according to the invention.
[0029] According to one feature, the fastening device is in the form of a shank having a first end and a second end and a head at the first end, the fastening device being inserted into the first and second orifices through the second end of the shank, the outer diameter of the isolating device being greater than the diameter of the head of the fastening device, the first structural part being covered with a primer coating. According to this feature, the head of the fastening device bears against the isolating device, and the isolating device and the fastening device are covered with the paint coating.
[0030] According to another feature, the head of the fastening device bears against the paint coating.
[0031] Advantageously, the fastening means are mounted after the paint coating has been applied and are therefore in direct contact with the paint coating of the aircraft structural part or (if the drying time of the paint coating cannot be observed and, during the placement of said fastening means, said paint coating has been forced out of the contact zone between the fastening means and the aircraft structural part) with the insulating means. In any case, the fastening means are not in direct contact with a zone of the aircraft structural part, which makes it possible to avoid contact between the metal material of the fastening means and the metal material forming the aircraft structure.
[0032] According to another feature, the isolation means takes the form of a pressure-sensitive adhesive coated on the support material.
[0033] Advantageously, the insulating device can thus be bonded to the ground area of the aircraft structural part and compressed by the fastening device when the fastening device is placed. This makes it possible to ensure that the insulating device forms an insulating barrier between the fastening device and the area of the aircraft structural part.
[0034] According to another feature, the insulating means are made of vinyl, polyurethane or epoxy resin.
[0035] According to another feature, the isolation device includes a corrosion inhibitor.
[0036] According to another feature, the thickness of the isolation device is between 30 microns and 200 microns. BRIEF DESCRIPTION OF THE DRAWINGS
[0037] Further features and advantages will become apparent from reading the following description of the invention, given by way of example only, with reference to the accompanying drawings, in which:
[0038] [ Figure 1 ] is a top view of a region of a structural part of an aircraft having a protection defect,
[0039] [ Figure 2 ] is taken along line II Figure 1 A cross-sectional view of
[0040] [ Figure 3 ] is a flow chart of the steps of a method for repairing an area of an aircraft structural part having a protection defect, which illustrates one embodiment of the present invention,
[0041] [ Figure 4 ] is a top view of a region of a structural part of an aircraft during the steps of a repair method according to one embodiment of the invention,
[0042] [ Figure 5 ] is taken along line IV-IV Figure 4 A cross-sectional view of
[0043] [ Figure 6 ] is a top view of a region of a structural part of an aircraft during a further step of a repair method according to one embodiment of the invention,
[0044] [ Figure 7 ] is taken along line VI-VI Figure 6 A cross-sectional view of
[0045] [ Figure 8 ] is a top view of a region of a structural part of an aircraft during a further step of a repair method according to one embodiment of the invention,
[0046] [ Fig. 9 ] is taken along line VIII-VIII Figure 8 A cross-sectional view of
[0047] [ Fig.10 ] is a top view of a region of a structural part of an aircraft during a further step of a repair method according to one embodiment of the invention,
[0048] [ Fig.11 ] is taken along line XX Fig.10 A cross-sectional view of
[0049] [ Fig.12 ] is a top view of a region of a structural part of an aircraft during a step of a repair method according to another embodiment of the invention,
[0050] [ Fig.13 ] is taken along line XII-XII Fig.12 A cross-sectional view of
[0051] [ Fig.14 ] is a top view of a region of a structural part of an aircraft during a further step of a repair method according to another embodiment of the invention,
[0052] [ Fig.15 ] is taken along line XIV-XIV Fig.14 A cross-sectional view of
[0053] [ Fig.16 ] is a cross-sectional view of a region of a structural part of an aircraft during the steps of a repair method according to one embodiment of the invention,
[0054] [ Fig.17 ] is a side view of the aircraft. DETAILED DESCRIPTION
[0055] like Fig.17 As shown, an aircraft 8 comprises a structure 10. This structure 10 comprises a plurality of structural parts 12, 14, which are painted and then assembled to each other using fastening means. The structural parts 12, 14 may be, but are not limited to, elements of a structure of the aircraft, such as a wing or a center wing box of the aircraft, which forms an interface with a fuel tank of the aircraft. During the assembly of these aircraft structural parts 12, 14, defects may occur in the paint of one of the structural parts 12, such as scratches or impacts. Such damage to the paint of the structural part may result in a defect in the protection of said structural part.
[0056] Figure 1 and Figure 2 Aircraft structural parts 12, 14 are shown which are intended to be assembled with each other. The aircraft structural part 12 comprises a metal body 16, for example an aluminum alloy body, which is covered with an aluminum oxide coating 18 after having undergone an anodizing treatment (a specific surface treatment of aluminum, which involves the use of electrolysis to produce a very strong, uniform aluminum oxide coating). For example, the thickness of the aluminum oxide coating 18 is between 1 micrometer and 5 micrometers. This aircraft structural part 12 is also covered with at least one paint coating 20. For example, the aircraft structural part 12 can be as Figure 2 As shown, the structural part 12 is covered with a single coat of paint 20, or with a first coat of paint, called primer, and a second coat of paint, called topcoat. The application of one or more coats of paint to the structural part of the aircraft depends on the location of the structural part on the aircraft. Depending on the location of the structural part on the aircraft, the structural part is subjected to various conditions that may cause corrosion to occur and therefore requires a greater or lesser degree of corrosion protection. The second topcoat is applied only to certain structural parts, depending on the location of these structural parts on the aircraft. In any case, the structural part is covered with at least one protective coat of paint. The single coat of paint 20 or the first coat of paint generally contains a corrosion inhibitor. The thickness of the paint coat 20 is, for example, about 25 microns. The aircraft structural part 12 also includes an orifice 22 for receiving a fastening device. This orifice 22 is along Figure 2 The axis represented by X in the figure is extended. Figure 2 As shown, the aluminum oxide coating 18 or the paint coating 20 extends in a plane orthogonal to the axis X. The orifice 22 passes through the aircraft structural part 12, i.e. the orifice 22 appears on both sides of the aircraft structural part 12. Likewise, the aircraft structural part 14 comprises an orifice 24 coaxial with the orifice 22 of the aircraft structural part 12, so that fastening means can be inserted into said orifices 22, 24 in order to fasten the aircraft structural parts 12, 14 to each other.
[0057] The aircraft structural part 12 has a defect 26 that penetrates the paint coating 20 and the aluminum oxide coating 18 and may penetrate into the thickness of the body 16. In the location of this defect 26, the body 16 is no longer protected by the aluminum oxide or the paint, and therefore, the body 16 does not have optimal corrosion resistance and lightning strike resistance characteristics in the location of the defect 26. Of course, not all defects 26 necessarily reach the body 16. Some defects 26 may penetrate only a portion of the paint coating 20, or penetrate the paint coating 20 and only a portion of the aluminum oxide coating 18. Damage that penetrates at least a portion of the paint coating 20 is considered a defect.
[0058] This defect 26 is located near the orifice 22 for receiving the fastener. "Near" means that, for the diameter D of the fastener, the defect 26 is located in a region centered on the orifice 22 for receiving the fastener, the diameter of which is less than or equal to five times the diameter D of the fastener. Figure 1 The region 28 shown shows the region that is believed to be near the fastening device. This region 28 contains the defect 26. Therefore, Figure 1 The region 30 shown, centered around the defect 26 and including the aperture 22 for receiving the fastening device, is considered to be a damaged region of the aircraft structural part 12 that must be repaired.
[0059] Figure 3 A flow chart showing the different steps of a method for repairing a damaged region 30 of an aircraft structural part 12 is shown.
[0060] The method comprises a first step E01 of sanding the damaged area 30 in order to smooth out (ie reduce) the defect 26 and thus remove a portion of the paint from the damaged area 30. In particular, the paint in the damaged area 30 located around the defect 26 is sanded. Figure 4 and Figure 5 The aircraft structural part 12 is shown after this grinding step E01. In the sanded damaged area 30', a portion of the paint coating 20 and a portion of the aluminum oxide coating 18 have been removed, so that the body 16 of the metal or aluminum aircraft structural part is exposed (i.e., bare) in this area 30'. In particular, the paint coating 20 and the aluminum oxide coating 18 located between the defect 26 and the orifice 22 for receiving the fastening means have been removed. Therefore, the defect 26 is no longer visible.
[0061] At the periphery of the polished damaged area 30', a smooth, gradual transition can be seen between the body 16 and the surface S of the aircraft structural part 12. This avoids abrupt angular transitions between the surface S of the aircraft structural part 12 and the bottom of the damaged area 30'. Figure 5As can be seen in FIG. 1 , at line A, the bottom of the ground damaged area 30 ′ corresponds to the paint coating 20 , while at line B, the bottom of the ground damaged area 30 ′ corresponds to the aluminum oxide coating 18 , and at line C, the bottom of the ground damaged area 30 ′ corresponds to the body 16 of the aircraft structural part 12 .
[0062] According to one embodiment, the method then comprises a step E10 of positioning insulating means 32 on the body 16 of the aircraft structural part in the region 30 ′ around the orifice 22 for receiving the fastening means. The insulating means 32 may be positioned directly on the bare body 16 of the structural part, or on the aluminum oxide coating 18, or on the paint coating 40. Figure 6 and Figure 7 The aircraft structural part 12 is shown after this step E10. The isolating means 32 is substantially cylindrical with an annular base and extends from the body 16 in a direction orthogonal to said body 16 (along the axis X). The isolating means 32 has an orifice 34 extending along the axis X, the diameter of which is substantially equal to the diameter of the orifice 22 for receiving the fastening means. The isolating means 32 is positioned in the zone 30 ′ so that the orifice 34 of said isolating means 32 and the orifice 22 for receiving the fastening means are coaxial. The thickness of the isolating means 32 (dimension along the axis X) is between 30 and 200 micrometers. The isolating means 32 thus forms a protrusion relative to the surface S of the aircraft structural part 12 outside the damaged zone 30. The isolating means 32 may take the form of a pressure-sensitive adhesive coated on a support material, for example a film made of plastic material. Only one surface of the isolating means is adhesive. The adhesive portion of the isolating device 32 is intended to be in contact with the body 16 of the aircraft structural part in the region 30' around the aperture 22 for receiving the fastening device, while the non-adhesive portion of the isolating device 32 is intended to face the head 50 of the fastening device 46. The isolating device 32 can thus be bonded to the body 16 of the aircraft structural part in the region 30' around the aperture 22 for receiving the fastening device. According to one configuration, the isolating device 32 can be made of vinyl, polyurethane or epoxy resin. According to one configuration, the isolating device 32 includes a corrosion inhibitor in order to improve the anti-corrosion properties of the aircraft structural part 12 when fastened.
[0063] According to this first embodiment, the method then comprises a step E12 of positioning the masking means 36 in the orifice 34 of the isolating means 32. The masking means 36 is substantially cylindrical with a circular base. The shape of the masking means 36 is complementary to the shape of the orifice 34 of the isolating means. The diameter of the masking means 36 is substantially equal to the diameter of the orifice 34 of the isolating means 32. Figure 7As shown, the length (dimension along axis X) of the masking device 36 can be sufficient to be inserted into the aperture 22 for receiving the fastening device. In other words, the length of the masking device 36 can be greater than the thickness of the isolation device 32. The masking device 36 is inserted into the aperture 34 of the isolation device so that the masking device is flush with the upper surface S32 of the isolation device 32. Alternatively, the masking device 36 can be inserted into the aperture 34 of the isolation device 32 so that the masking device protrudes from the upper surface S32 of the isolation device. The masking device 36 can take the form of a plug or a peg intended to be removed from the aircraft structure part 12. The masking device 36 is therefore removable. Compared with when the masking device 36 is fully inserted into the aperture 34 of the isolation device 32, the masking device 36 that partially extends to the outside of the aperture 34 of the isolation device 32 is easier to grasp.
[0064] According to this first embodiment, the method comprises a step E20 of applying a first primer coating 40 to the polished area 30 ' of the aircraft structural part. This step E20 corresponds to an anti-corrosion surface treatment. This paint coating 40 is also applied to the insulating means 32 and the masking means 36. For example, the thickness of the primer coating 40 is about 25 micrometers. In particular, the insulating means 32 is made of a material to which the primer 40 adheres.
[0065] According to this first embodiment, the method then comprises a step E22 of applying a second topcoat 42 to the first paint coat 40. For example, the thickness of the topcoat 42 is about 25 micrometers. This step E22 is optional and depends on the aircraft structural part 12, in particular on its location on the aircraft. Figure 8 and Fig. 9 The aircraft structural part 12 is shown after this step E22. Figure 8 As can be seen in FIG. 1 , once the paint coatings 40, 42 have been applied, it is no longer possible to determine the exact position of the defect 26, since the defect has been repaired. In a known manner, the position of the masking device 36 has been pre-marked so that the masking device can be easily removed. When the masking device 36 is flush with the upper surface S32 of the isolation device 32, the paint coatings 40, 42 are more evenly distributed on the isolation device than when the masking device protrudes beyond the orifice 34 of the isolation device 32.
[0066] According to this first embodiment, the method then comprises a step E30 of removing the masking means 36 from the orifice 34 of the isolating means 32 and, if applicable, from the orifice 22 for receiving the fastening means. The masking means 36 covered with the paint coating 40, 42 is removed from the orifices 34, 22 in order to leave a free space for the fastening means. Since the diameter of the masking means 36 is equal to the diameter of the orifice of the isolating means 32, only the paint coating 40, 42 on the masking means 36 is removed, which leaves the orifice 34 of the isolating means 32 free and the upper surface S32 of the isolating means 32 covered with the paint coating 40, 42.
[0067] According to this first embodiment, the method then comprises a step E40 of inserting the fastening means 46 into the orifice 34 of the insulating means 32, into the orifice 22 for receiving the fastening means and into the orifice 24 of the aircraft structural part 14. This step E40 may be performed before the drying time of the paint coating 42 has elapsed. Fig.10 and Fig.11 The aircraft structural part 12, 14 is shown after this step E40. The fastening device 46 takes the form of a shank 48 having a first end 48a and a second end 48b and a head 50 located at the first end 48a of the shank 48. The diameter of the head 50 is greater than the diameter of the shank 48. The diameter of the shank 48 is substantially equal to the diameter of the orifice 34 of the isolating device 32. The fastening device 46 is inserted into the orifice 34 through the second end 48b of the shank 48, then into the orifice 22, and then into the orifice 24. The outer diameter of the isolating device 32 is greater than the diameter of the head 50. During the step E40 of inserting the fastening device 46, the head 50 bears against the paint coating 42. More specifically, when the fastening device 46 is fastened, the head 50 comes into contact with the paint coating 42 during the drying of the paint coating 42, which may cause cracking or crazing, or cause the paint to move out of the area of the head 50 facing the fastening device 46 and move above the upper surface S32 of the isolation device. Fig.11The fastener 46 (as shown) may remain between the head 50 and the isolating device 32, but not necessarily over the entire surface between the fastener 46 and the isolating device 32. The same applies to the paint coating 40, only a portion of which may remain between the head 50 of the fastener 46 and the isolating device 32. In any case, the fastener 46 made of a metallic material is in contact with the paint coating 40, 42, or in direct contact with the isolating device 32 (in the case where all the paint has been expelled from the area between the head 50 and the upper surface S32 of the isolating device 32, or in the case where cracking has occurred over the entire thickness of the paint coating 40, 42). Because the paint coating 40, 42 and the isolating device 32 are made of non-metallic materials, the risk of galvanic corrosion between the fastener 46 and the aircraft structural part 12 is greatly reduced, and the electromagnetic requirements are met.
[0068] According to one feature, the fastening means 46 comprises a washer (not shown in the figures) intended to be arranged under the head 50 of the fastening means and to bear against the paint coating 42 during the step E40 of inserting the fastening means 46. The outer diameter of the isolating means 32 is greater than the diameter of the washer. During the fastening of the fastening means 46, a portion of the paint 42 may remain between the washer and the isolating means 32, but not necessarily over the entire surface between the fastening means 46 and the isolating means 32. The same applies to the paint coating 40, only a portion of which may remain between the washer and the isolating means 32. In any case, the fastening means 46 made of metal material will be in contact with the paint coatings 40, 42, or in direct contact with the isolating means 32 (in the case where all the paint has been expelled from the area between the washer and the upper surface S32 of the isolating means 32, or in the case where cracking has occurred over the entire thickness of the paint coatings 40, 42).
[0069] According to a second embodiment, after the sanding step E01 and before the step E20 of applying the first coat of paint, the method comprises, as a variant of steps E10 and E12, a step E40' of inserting the fastening means 46 into the orifice 34 of the insulating means 32. The head 50 of the fastening means 46 bears against the upper surface S32 of the insulating means 32. According to this second embodiment, the method then comprises a step E10' of positioning the insulating means 32 together with the fastening means 46 on the body 16 of the aircraft structural part in the region 30' around the orifice 22 for receiving the fastening means. During this step E10', the insulating means 32 is positioned so that the fastening means 46 is inserted into the orifice 22 for receiving the fastening means and into the orifice 24 of the aircraft structural part 14. Fig.12 and Fig.13The aircraft structure part 12 is shown after this step E10'. The fastening means 46 made of metal are in direct contact with the insulating means 32. The risk of galvanic corrosion between the fastening means 46 and the aircraft structure part 12 is then greatly reduced.
[0070] According to this second embodiment, a paint coating 40 is applied to the isolating means 32 and the fastening means 46. The fastening means are positioned before the paint coating is applied. In particular, the isolating means 32 and the fastening means 46 are made of a material to which the primer 40 adheres. Fig.14 and Fig.15 The aircraft structure part 12 is shown after step E22 according to this second embodiment. Fig.14 As can be seen in FIG. 4 , once the paint coatings 40 , 42 have been applied, the exact location of the defect 26 can no longer be determined, since the defect has been repaired.
[0071] Advantageously, in this second embodiment, the fastening means are painted and are therefore not visible after application of the paint coating 40 , 42 , in contrast to the first embodiment described.
[0072] As a variant of this second embodiment, after the step E10 of positioning the isolating means and before the step E20 of applying the first paint coat 40, the method comprises, as a variant of steps E10' and E12, a step E40'' of inserting the fastening means 46 into the orifices 34 of the isolating means 32 and into the orifices 22 for receiving the fastening means 46. Thus, the isolating means 32 are arranged on the structural part before the insertion of the fastening means 46, and not simultaneously with the fastening means 46, as in the second embodiment.
[0073] like Fig.11 and Fig.15 As shown, after the different steps of the repair method, there is a height difference (indicated by H, along the axis X) between the upper surface S42 of the second paint coating 42 on the aircraft structural part 12 outside the damaged area 30 and the upper surface on the fastening means 46 or in the vicinity of said fastening means 46. The upper surface S42 is not linear. This height difference H is not visible to the naked eye, or is extremely small. Therefore, the repair of the paint defect 26 is not visible from the outside after the repair (only the presence of the isolation means 32 indicates the presence of the protective defect that has been subjected to the repair method according to the invention).
[0074] According to another embodiment, after the sanding step E01 and before the step E20 of applying the first paint coating, the method comprises a step E50 of surface treatment of the sanded area 30' of the aircraft structure part, during which a surface treatment 52, such as a chemical conversion coating or a reactive primer, is applied to a portion of the sanded area 30' of the aircraft structure part. Such a surface treatment 52 allows the adhesion of the primer coating to be improved. The surface treatment can be applied in a bath or locally. Fig.16 The aircraft structural part 12 is shown after this step E50. In particular, the surface treatment 52 is not applied on the pre-existing primer coating 20, but only on the pre-existing aluminum oxide coating 18 or on the bare body 16 of the aircraft structural part 12. Preferably, the (local or bath) surface treatment is applied only to the bare body 16 of the aircraft structural part 12. Overlapping surface treatments are preferably avoided. According to one embodiment, during the grinding step E01, the aluminum oxide coating 18 is completely removed, since it is thin (approximately between 1 micrometer and 5 micrometers), which makes it possible to apply the surface treatment 52 only to the bare body 16 of the aircraft structural part 12. For example, as Fig.16 As shown, the surface treatment 52 is applied at line B and line C instead of line A. On the primer coating 20, this surface treatment 52 is not necessary. The drying time of this surface treatment 52 is between 30 minutes and 1 hour. For example, the thickness of this surface treatment 52 is less than or equal to 1 micron. The thickness of the aluminum oxide coating 18 after the bath surface treatment can be about 5 microns, and the thickness after the local surface treatment (such as chemical conversion) can be about 1 micron.
[0075] Whatever the embodiment, the presence of the isolating means 32 between the fastening means 46 and the aircraft structural part 12 of the aircraft makes it possible to create an isolating barrier between the metal material forming the fastening means 46 and the metal material forming the aircraft structural part, which makes it possible to reduce the risk of galvanic corrosion when repairing defects in the vicinity of the fastening means and to ensure that electromagnetic requirements are met along the entire length of the aircraft structural part.
[0076] Furthermore, this makes it possible to make the repair of the protection defects independent of the drying time of the paint applied to said structural part of the aircraft.
Claims
1. A method for repairing a region (30) of a metal structural part (12) of an aircraft having a protection defect (26), said region (30) of said structural part having a first aperture (22) for receiving a metal fastening device, said first aperture (22) having a first diameter, said repair method comprising: - a sanding step (E01) during which the paint in said area (30) of said aircraft structural part is sanded, Characterized in that the repair method further comprises: - a step (E10, E10') of positioning an isolating device (32), during which the isolating device (32) is positioned around the first orifice (22) on the ground area (30') of the aircraft structural part, the isolating device (32) comprising a second orifice (34) having a second diameter substantially equal to the first diameter of the first orifice (22), the isolating device (32) being positioned around the first orifice (22) on the ground area (30') of the aircraft structural part such that the first orifice and the second orifice (22, 34) are coaxial, - a step (E20) of applying at least one paint coating (40), during which a primer coating (40) is applied to the sanded area (30') of the aircraft structural part, said primer coating (40) being applied to the insulation device (32).
2. The repair method according to claim 1, comprising: Prior to the step (E10') of positioning the isolating device (32), a step (E40') of inserting a fastening device (46) is performed, during which the fastening device (46) is inserted into the second orifice (34) of the isolating device (32), and wherein during the step (E10') of positioning the isolating device (32), the fastening device (46) is inserted into the first orifice (22), and during the step (E20) of applying the paint coating (40), the paint coating (40) is applied to the fastening device (46).
3. The repair method according to claim 2, wherein: The fastening device (46) is in the form of a shank (48) having a first end and a second end (48a, 48b) and a head (50) located at the first end (48a), the fastening device (46) being inserted into the first and second orifices (22, 34) through the second end (48b) of the shank (48), wherein the outer diameter of the isolation device (32) is greater than the diameter of the head (50) of the fastening device (46), and wherein during the step (E40') of inserting the fastening device (46), the head (50) of the fastening device (46) bears against the isolation device (32).
4. The repair method according to claim 1, comprising: Prior to the step (E20) of applying the paint coating (40), a step (E40'') of inserting a fastening device (46), during which the fastening device (46) is inserted into the first orifice (22) and the second orifice (34) of the insulating device (32), and during the step (E20) of applying the paint coating (40), the paint coating (40) is applied to the fastening device (46).
5. The repair method according to claim 1, comprising: a step (E12) of positioning a masking device (36) before the step (E20) of applying the paint coating (40), during which the masking device (36) is positioned in the first orifice and the second orifice (22, 34); and a step (E30) of removing the masking device (36) after the step (E20) of applying the paint coating (40), during which the masking device (36) is removed from the first orifice and the second orifice (22, 34); and a step (E40) of inserting a fastening device (46), during which the fastening device (46) is inserted into the first orifice and the second orifice (22, 34).
6. The repair method according to claim 5, wherein: The fastening device (46) is in the form of a shank (48) having a first end and a second end (48a, 48b) and a head (50) located at the first end (48a), the fastening device (46) being inserted into the first and second orifices (22, 34) through the second end (48b) of the shank (48), wherein the outer diameter of the isolation device (32) is greater than the diameter of the head (50) of the fastening device (46), and wherein during the step (E40') of inserting the fastening device (46), the head (50) of the fastening device (46) bears against the paint coating (40).
7. The repair method according to any one of claims 1 to 6, comprising: After the grinding step (E01), a surface treatment step (E50) is provided during which a chemical conversion coating or a reactive primer (52) is applied to at least a portion of the ground area (30') of the aircraft structural part.
8. Repair method according to any of the preceding claims, wherein: The isolation means (32) is in the form of a pressure sensitive adhesive coated on a support material.
9. Repair method according to any of the preceding claims, wherein: The isolating device (32) is made of vinyl, polyurethane or epoxy resin.
10. Repair method according to one of the preceding claims, wherein: The isolation device (32) includes a corrosion inhibitor.
11. Repair method according to one of the preceding claims, wherein: The thickness of the isolation device (32) is between 30 micrometers and 200 micrometers.
12. An aircraft metal structural component (12) having a region (30) repaired using the repair method according to one of the preceding claims.