Method for repairing a composite aircraft panel

Abstract

The invention relates to a method for repairing an aircraft panel (2), the panel being formed of a fibrous layer (20), the method comprising: - a step of removing the damaged parts from the panel so as to form a through-opening; - a step of filling the opening with repair plies, the repair plies comprising replacement plies arranged within the opening, and outer cover plies and inner cover plies (26) arranged on either side of the replacement plies, the outer cover plies being covered by a metal protective element (30); - a step of impregnating and curing a resin; - a step of consolidating the panel by means of fastening rivets (32), each fastening rivet (32) passing through the metal protective element (30), all of the cover plies and the fibrous layer (20).

Description

Method for repairing an aircraft composite panel Technical Field

[0001] The present invention relates to a method for repairing an aircraft part made of composite material, and more particularly a composite material panel forming part of a fairing, such as a nacelle cover of a propulsion assembly or a fuselage element. Previous technique

[0002] Aircraft are currently designed with a significant number of composite material parts. These can include, for example, parts forming sections of the fuselage or parts of a propulsion nacelle.

[0003] Composite material parts are generally formed from a layering of fibrous plies (containing, for example, carbon, glass, Kevlar fibers, etc.), these fibrous plies being embedded in a resin.

[0004] When composite material parts are exposed to the airflow circulating around the aircraft, or around and through the propulsion system, they can be damaged during aircraft operation. If a composite material part is damaged, it is repaired whenever possible to restore its structural and aerodynamic properties.

[0005] Repairs to composite material parts must be carried out in such a way that the repaired part meets the original specifications, and in particular must enable the repaired part to meet the same criteria as the original part, in terms of fire resistance, lightning resistance and mechanical strength.

[0006] The known repair techniques are not entirely satisfactory. One known technique involves repairing the area by attaching a metal piece, usually called a metal doubler, to the damaged area. However, this type of repair has the disadvantage of adding weight to the repaired part. altering its aesthetic appearance and also reducing its aerodynamic qualities. Mechanical resistance is also compromised, as a metal reinforcement makes the repaired part more rigid than the original.

[0007] A second technique involves repairing the part by replacing the damaged fiber plies with new, bonded fiber plies. This type of repair has the advantage of not adding weight to the repaired part and of preserving its original mechanical strength, but it has the disadvantage of not preserving the fire resistance of the original part. Because the repair plies are bonded to the original part with resin, the resin is quickly consumed by fire, causing the repair plies to detach.

[0008] The objective of the present invention is to propose a method for repairing an aircraft part made of composite material which preserves the qualities of the original part, including with regard to fire resistance. Description of the invention

[0009] To this end, the invention relates to a method for repairing an aircraft composite panel, the panel being formed by a fibrous layer comprising fibrous plies and covered by a metallized layer, the method comprising: - a step of removing the damaged parts of the panel, the damaged parts being removed in such a way as to leave a through opening within the fibrous layer; - a step of filling the opening with fibrous repair folds, the fibrous repair folds comprising replacement folds arranged within the opening made in the fibrous layer, and comprising external cover folds and internal cover folds arranged on either side of the replacement folds, the external cover folds being covered by a metallic protective element; - a step of impregnation and hardening of a resin within the fibrous repair folds and the metallic protective element; - a consolidation step using fixing rivets, each fixing rivet passing through the metallic protective element, all the cover folds and all the fibrous folds of the fibrous layer.

[0010] Thus, the method according to the invention makes it possible to carry out a repair compatible with the fire resistance required for the original part. Indeed, the rivet fastening maintains the bond and cohesion between the sound fiber plies and the replacement fiber plies, even when the resin is consumed in the event of a fire. The fastening rivets therefore prevent the replacement plies from detaching in the event of a fire.

[0011] The process according to the invention may include one or more of the following optional features considered alone or in all possible combinations.

[0012] According to one characteristic, the fixing rivets are arranged around the opening made in the fibrous layer so as not to pass through the replacement folds.

[0013] According to one characteristic, the fixing rivets are arranged at an equidistance from the opening made in the fibrous layer.

[0014] According to one characteristic, the fixing rivets are arranged along a circle.

[0015] According to one characteristic, the fixing rivets are angularly and regularly distributed around the opening made in the fibrous layer.

[0016] According to one characteristic, the spacing between two adjacent fixing rivets is between 15 and 70 mm.

[0017] According to one characteristic, a portion of the fixing rivets, for example at least one fixing rivet out of two, is mounted without the addition of an insulating material.

[0018] According to one characteristic, the fixing rivets which are mounted without the addition of an insulating material are covered with a layer of varnish.

[0019] According to one characteristic, each fixing rivet has a fixing shank, the fixing shanks of the fixing rivets being embedded in a layer of an electrically insulating material, such as silicone.

[0020] According to one characteristic, each fixing rivet has a countersunk head.

[0021] The invention also relates to a composite material panel for aircraft, the panel comprising a repair carried out using the repair process according to one of the preceding claims. Brief description of the drawings

[0022] Figure 1 is a perspective view of an aircraft propulsion assembly, including a cowling formed by a composite material panel, the panel having been damaged by an impact from a foreign body and requiring repair.

[0023] Figure 2 is a partial cross-sectional view of the damaged area of ​​the hood visible in Figure 1.

[0024] Figure 3 is a view analogous to Figure 2, showing the panel after the implementation of a first step of the process according to the invention.

[0025] Figure 4 is a view analogous to Figure 2, showing the panel after implementation of a second step of the process according to the invention.

[0026] Figure 5 is a view analogous to Figure 2, showing the panel after implementation of the final step of the process according to the invention.

[0027] Figure 6 is a top view showing the panel after repair.

[0028] Figure 7 is a detailed cross-sectional view showing a first type of mounting for a fixing rivet.

[0029] Figure 8 is a view analogous to figure 7, showing a second type of mounting for a fixing rivet.

[0030] Figure 9 is a diagram representing the steps of a process according to the invention. Detailed description

[0031] Figure 1 shows a nacelle 1 of an aircraft propulsion system comprising a turbofan engine. The nacelle 1 includes a front section 10 forming an air intake 10a, and a midsection 12 which includes two fan cowls 12a. enveloping a turbojet fan casing, and a rear section 14, which in this example incorporates a thrust reversal system.

[0032] The elements of the nacelle 1 forming parts of fairings, such as the fan cowls 12a, are formed by a structure comprising a panel 2 made of composite material.

[0033] In the example, the composite panel 2 forming one of the blower hoods 12a is damaged in a damaged area 2a and requires repair. The implementation of a repair method according to the invention for repairing the damaged area 2a is described below, with reference to Figures 2 to 9.

[0034] Figure 2 is a partial cross-sectional view of panel 2 of Figure 1, showing the damaged area 2a and its surroundings. As shown in Figure 2, panel 2 has a fibrous layer 20 formed by superimposed fibrous plies 200, these fibrous plies 200 comprising, for example, carbon, glass, or Kevlar fibers. In the example, the fibrous layer 20 has four superimposed fibrous plies 200. The external surface of the fibrous layer 20, i.e., the surface facing outwards from the nacelle when the panel is in its operating configuration, is covered by a metallized layer 22 for lightning protection. In the example, the metallized layer 22 is formed by an electrically conductive metal mesh. Such a mesh may comprise, for example, bronze, copper, aluminum, or an alloy containing one of these metals.

[0035] As shown in Figure 2, panel 2 is pierced through at the damaged area 2a.

[0036] Figure 3 illustrates the panel of Figure 2 after the implementation of a first step 102 of the repair process 100 according to the invention, during which the damaged parts of the panel are removed. In this first step 102, the metallized layer 22 and the fibrous plies 200 are machined to define a repair area 2b with a substantially circular outline. Thus, at the end of this first step 102, the metallized layer 22 has a circular opening 22a of diameter D1, and the fibrous layer 20 has a circular through opening 20a with a diameter D2, preferably less than to diameter Dl. Advantageously, the fibrous layer 20 is machined so as to present a flared (or "stepped") opening 20a, this opening having, at the level of the internal surface of the panel 2 a diameter equal to the diameter D2, and, at the level of the external surface of the panel 2, a diameter D3 greater than the diameter D2 (but less than the diameter Dl).

[0037] Figure 4 illustrates the panel of Figure 2 after implementation of a second step 104 of the repair process according to the invention. In this second step 104, the opening 20a in the damaged area 2a of the panel is filled with fibrous repair plies 24-28. The fibrous repair plies 24-28 comprise: - replacement fibrous folds 24, which are arranged so as to fill the opening 20a provided in the fibrous layer 20, the replacement fibrous folds 24 therefore being of dimensions compatible with their insertion into the opening 20a; - one or more internal covering folds 26, or internal folds 26, intended to cover part of the internal surface of the fibrous layer 20, these folds having a diameter D4 strictly greater than the largest diameter of the opening 20a made in the fibrous layer 20, and preferably greater than or equal to the diameter Dl; - one or more external cover folds 28, or external folds 28, intended to cover part of the external surface of the fibrous layer 20, these folds having a diameter D5 less than the diameter Dl, but greater than the largest diameter of the opening 20a made in the fibrous layer 20.

[0038] As can be seen in Figure 4, the repair fibrous plies 24-28 are arranged so as to completely fill the opening 20a, and so as to cover this opening, on each side of the panel 2, as well as part of the healthy plies, i.e. the fibrous plies 200 of the fibrous layer 20. The replacement plies 24 are thus sandwiched between the internal cover plies 26 and the external cover plies 28.

[0039] Once the repair folds 24-28 are in place, the outer cover folds 28 are covered by a metallic protective element 30 (visible in Figure 5), intended in particular to provide protection against lightning, such as the metallized layer 22. In the example, the metallic protective element 30 is formed by an electrically conductive metal mesh. Such a mesh may include, for example bronze, copper, aluminum, or an alloy containing one of these metals. As shown in Figure 5, in order to cover the outer cover folds 28, the metallic protection element 30 has a diameter D6 greater than the diameter of the outer folds 28, and greater than the diameter DI of the opening in the metallized layer 22. The metallic protection element 30 is thus arranged so as to be in electrical contact with the metallized layer 22, which ensures continuity of lightning protection.

[0040] When the repair folds 24-28 and the metal protective element 30 are in place, a third step 106 is carried out, during which a resin is impregnated and cured. During this step, all the repair folds 24-28 and the metal protective element 30 are impregnated with resin, and then this resin is cured. At the end of this third step 106, the metal protective element 30 is bonded to the external cover folds 28 by means of the cured resin.

[0041] Figure 5 illustrates the panel of Figure 2 after implementation of a fourth step 108 of the process according to the invention. The fourth step 108 is a consolidation step and is carried out after the resin has hardened. During this step 108, the repair is consolidated by means of fixing rivets 32. The fixing rivets 32 are arranged so as to pass through the metal protective element 30, the external 28 and internal 26 cover folds, and also through all the fibrous folds 200 of the fibrous layer 20. The fixing rivets 32 are further arranged around the opening 20a, so that the fixing rivets 32 do not pass through the replacement folds 24. Thus, the fixing rivets 32 bear, on the one hand, on the metal protective element 30, and, on the other hand, on the internal folds 26.The use of fixing rivets secures all the added elements during the repair process. The fixing rivets 32 ensure a strong and durable attachment of the various elements added to panel 2 as part of the repair, and guarantee that the repaired area will have fire resistance equivalent to the original panel. Indeed, the rivet fastening maintains the cohesion of all the fiber plies 200, 24, 26, 28 (i.e., the sound fiber plies 200 and the repair plies 24-28) when the resin is destroyed in the event of a fire.

[0042] Figure 6 illustrates the repair area viewed from above, and more specifically shows an example of the distribution of the fixing rivets 32. Advantageously, the fixing rivets 32 are arranged approximately equidistant from the center of the opening 20a. If the surface of the panel 2 at the repair area 2b is flat, the fixing rivets 32 are thus arranged along a circle of diameter D7. This diameter D7 is smaller than the diameter DI of the opening 22a in the metallized layer 22, but larger than the largest diameter of the opening 20a in the fibrous layer 20. As can be seen in Figure 5, this ensures that each fixing rivet 32 ​​passes not only through the outer 28 and inner 26 cover plies, but also through all the fibrous plies 200 of the fibrous layer 20. Advantageously, the fixing rivets 32 are angularly and regularly spaced.The number of fixing rivets 32 will of course depend on the value of the diameter D7. In order to ensure satisfactory fire resistance, the spacing E between two adjacent fixing rivets 32 will preferably be less than or equal to 70 mm, and for example between 15 and 70 mm.

[0043] Figures 7 and 8 illustrate two types of mounting of the fixing rivets 32. Advantageously, at least one out of every two fastening rivets 32 is configured to be electrically conductive, in order to facilitate the passage of an electric current in the event of a lightning strike. Figure 7 shows a rivet of this first type, while Figure 8 shows a rivet of the second type.

[0044] To ensure electrical conductivity, the fixing rivet 32 ​​of Figure 7 is mounted without the addition of any insulating material (such as sealant). Furthermore, the fixing rivet 32 ​​is coated with a layer 320 of varnish. Advantageously, fixing rivets 32 of the type shown in Figure 7 are identified, for example, by applying a colored varnish. This configuration allows the rivet type to be identified during assembly.

[0045] Figure 8 shows that a layer 322 of an insulating material has been deposited on the surface of the fixing rivet 32, and in particular on the head 32a of the fixing rivet. In this example, it is a layer of sealant.

[0046] Regardless of the type of mounting of the fixing rivets 32, the fixing shank 32b of the fixing rivets 32 should be embedded in an insulating material 324, such as silicone. This arrangement prevents the generation of sparks in the event of high-intensity electrical currents, for example during a lightning strike.

[0047] Advantageously, in order to preserve the aerodynamic qualities of the panel, the fixing rivets 32 have milled heads 32a, as seen in figures 7 and 8.

Claims

Demands

1. A method for repairing (100) a panel (2) made of aircraft composite material, the panel (2) being formed by a fibrous layer (20) having fibrous plies (200) and covered by a metallized layer (22), the method comprising: - a step (102) of removing the damaged parts of the panel, the damaged parts being removed so as to provide within the fibrous layer (20) a through opening (20a); - a step (104) of filling the opening (20a) with fibrous repair folds (24-28), the fibrous repair folds (24-28) comprising replacement folds (24) arranged within the opening (20a) made in the fibrous layer (20), and comprising external cover folds (28) and internal cover folds (26) arranged on either side of the replacement folds (24), the external cover folds (28) being covered by a metallic protective element (30); - a step (106) of impregnation and hardening of a resin within the fibrous repair folds (24-28) and the metallic protective element (30); - a consolidation step (108) using fixing rivets (32), each fixing rivet (32) passing through the metal protection element (30), all the cover folds (26-28) and all the fibrous folds (200) of the fibrous layer (20).

2. A repair method (100) according to the preceding claim, wherein the fastening rivets (32) are arranged around the opening (20a) formed in the fibrous layer (20) so as not to penetrate the replacement folds (24).

3. A repair method (100) according to claim 1 or 2, wherein the fastening rivets (32) are arranged equidistant from the opening (20a) formed in the fibrous layer (20).

4. Repair method (100) according to the preceding claim, wherein the fixing rivets (32) are arranged along a circle.

5. Repair method (100) according to the preceding claim, wherein the fixing rivets (32) are angularly regularly distributed around the opening (20a) formed in the fibrous layer (20).

6. Repair method (100) according to any one of the preceding claims, wherein the spacing between two adjacent fixing rivets (32) is between 15 and 70 mm.

7. Repair method (100) according to any one of claims 1 to 6, wherein a part of the fixing rivets (32), for example at least one fixing rivet (32) out of two, is mounted without the addition of an insulating material.

8. Repair method (100) according to the preceding claim, wherein the fixing rivets (32) which are mounted without the addition of an insulating material are covered with a layer (320) of a varnish.

9. A repair method (100) according to any one of the preceding claims, wherein each fixing rivet (32) has a fixing shank (32b), the fixing shanks (32b) of the fixing rivets (32) being embedded in a layer of an electrically insulating material (324), such as silicone.

10. A repair method (100) according to any one of the preceding claims, wherein each fastening rivet (32) has a countersunk head (32a).

11. A panel (2) made of aircraft composite material, the panel (2) having a repair made by means of the repair method (100) according to any one of the preceding claims.

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