WO2016092238A1

WO2016092238A1 - Method for the repair of an acoustic panel made from composite material

Info

Publication number: WO2016092238A1
Application number: PCT/FR2015/053459
Authority: WO
Inventors: Arnaud DELEHOUZE; Jean-Philippe MUCHERIE; Michel ROGNANT; Carole JULLIEN
Original assignee: Aircelle
Priority date: 2014-12-11
Filing date: 2015-12-11
Publication date: 2016-06-16
Also published as: FR3029831B1; FR3029831A1

Abstract

The invention relates to a method for the repair of a composite panel (10), said panel comprising at least one first composite skin (12) defined by a lower face (18), characterized in that it comprises a cutting step which consists of forming a hole (22) stepped like a staircase in at least the first skin (12), a fitting step which involves arranging a support face of a shaping tool against the lower face (18) of the first skin (12), a reconstitution step of the first skin (12) consisting of closing the hole (22) by means of a repair piece (38) and a fastening step which involves fastening the repair piece (38) on the first skin (12) by mechanical fastening means (42).

Description

Method for repairing an acoustic panel made of material

composite

The invention relates to a method for repairing a composite panel, and more particularly to a method for repairing an acoustic skin of a composite panel for an aircraft nacelle.

As is known per se, an aircraft propulsion unit conventionally comprises a turbojet engine housed inside a nacelle.

The nacelle generally has a tubular structure comprising an air inlet upstream of the turbojet, a median section intended to surround a fan of the turbojet engine and its casing, and a downstream section intended to surround the combustion chamber of the turbojet engine and sheltering where appropriate. thrust reverser means. It can be terminated by an ejection nozzle whose output is located downstream of the turbojet engine.

In order to reduce the noise pollution generated by the turbojet engine, it is common to equip several parts of the nacelle and / or the turbojet engine with acoustic attenuation structures.

These parts of the nacelle and / or the turbojet engine may be, for example, an internal surface of the air circulation duct for a double-flow turbojet engine, a nozzle or ejection cone surface, etc.

An acoustic attenuation structure is generally in the form of a honeycomb-core sandwich type acoustic panel comprising at least one resonator, for example of the honeycomb type, covered with a perforated outer skin (called an acoustic skin). ) and a solid inner skin (called skin support or structuring).

One and / or the other of these skins may be made of composite material comprising a set of fibers (superposition of fibrous folds), embedded in a polymerized resin.

The outer skins are surfaces of the panel which are exposed to the flow of air flowing through the propulsion unit and may be damaged, and in particular undergo tearing, perforation or detachment, for example at the skin / acoustic honeycomb structure interfaces. .

It is then necessary to repair the damaged panel so as to restore the panel its properties, in terms of acoustic performance and structural strength. In the case of damage to a skin of the panel, it is known to repair the skin using a doubler, that is to say a portion of skin, metal or composite for example, fixed in superposition on the skin to repair at the damaged area.

The use of such doublers has the particular disadvantage of forming an extra thickness on the skin of the panel, particularly in the case of repair of an outer skin, disrupts the flow of air flow through the vein delimited by this outer skin.

Also known is a method of repairing a composite panel (of the monolithic type, that is to say not having a sandwich structure) described in document US-2014141 190, which consists of removing the damaged part of the panel forming a frustoconical hole leaving a support partition at the narrower end of the hole, and filling the hole with a repair material.

This type of process does not guarantee sufficient mechanical strength to be applied to a turbojet engine nacelle acoustic panel that is highly solicited.

The present invention aims in particular to solve these drawbacks and relates, for this purpose, to a method for repairing a composite panel, said panel comprising at least a first composite skin which is delimited by a lower wall-forming face of the panel, and a upper face, the method being characterized in that it comprises:

- A cutting step of forming a stepped hole at least in the first skin, the hole having a generally frustoconical shape of decreasing section towards the underside of the first skin, and the hole being delimited by a peripheral edge forming steps in the first skin, each step extending circumferentially around the hole delimiting a housing adapted to receive a fibrous fold,

an apparatusing step which consists in arranging a bearing face of a forming tool against the lower face of the first skin, the bearing surface of the tool having a shape matching the underside of the first skin; , at least on the periphery of the hole,

a step of reconstituting the first skin which consists in closing the hole by means of a repair part, the repair part being formed by fibrous folds deposited, superimposed, in each associated housing formed by the hole, starting with a first fold fibrous on the bearing surface of the forming tool, the first fibrous fold being flush with the underside of the first skin, and

- An attachment step which consists in fixing the repair part on the first skin by mechanical fixing means.

Thus, the method according to the invention allows the repair part to be flush with the first skin of the panel and to retain the aerodynamic properties of the panel. In addition, the method according to the invention ensures a good mechanical strength of the repair part.

In one embodiment, the method comprises an acoustic piercing step which consists in making perforations in the repair part to favor the acoustic absorption characteristics of the panel.

This step of the process makes it possible to preserve the acoustic properties of the panel.

In one embodiment, the fastening means includes metal inserts each extending substantially perpendicular to the first skin and which is adapted to connect the patch to the first skin.

Advantageously, the repair part forms a reinforcing extra thickness, which is arranged opposite the lower face of the first skin, and which is designed to promote the anchoring of the fastening means in the repair part. This feature reinforces the mechanical connection of the repair part on the first skin of the panel.

In another aspect, the hole formed in the panel during the cutting step has a polygonal shape.

The composite panel comprises a honeycomb core, for example an aluminum honeycomb which is cut during the cutting step, the method comprising a step of repairing said cellular core which consists in filling the hole formed in the alveolar soul by a complementary honeycomb core.

In one embodiment, the panel has a second composite skin that extends substantially parallel to the first skin and is cut during the cutting step, the method including a second skin repair step of closing the skin. hole formed in the second skin by a reported closure plate.

In one embodiment, the reconstitution step comprises a first firing step, which consists in baking the repair part. In one embodiment, the support face of the forming tool is designed to rise in temperature so as to bake the repair part. The forming tool comprises for this purpose a heating device, such as an electrical resistance.

In one embodiment, the forming tool is removed at the latest at the end of the repair process according to the invention, and in particular at the end of the reconstitution step.

In one embodiment, the method comprises a second firing step which consists of firing the closure plate, the second firing step being carried out after the first firing step.

The invention also relates to a composite panel, comprising at least a first composite skin which is delimited by a lower wall-forming face of the panel, and by an upper face, the panel being characterized in that it comprises a portion repaired in accordance with a process as defined above.

The invention finally relates to an aircraft turbojet engine nacelle comprising a composite panel as defined above.

Other features and advantages of the invention will appear on reading the detailed description which follows for the understanding of which reference will be made to the appended drawings in which:

- Figure 1 is a schematic cross-sectional view illustrating a repair of an acoustic panel comprising a repair part, according to the method according to the invention;

FIG. 2 is a diagrammatic cross-sectional detail view along the frame II of FIG. 1, which illustrates the deposition of fibrous folds on a forming tool;

- Figure 3 is a schematic detail view in cross section along the frame II of Figure 1, which illustrates the fastening means of the repair part;

- Figure 4 is a schematic bottom view, which illustrates the acoustic panel of Figure 1 repaired according to the method according to the invention.

In the description and the claims, the terms "upper" and "lower" and their derivatives will be used in a nonlimiting manner with reference to the upper part and the lower part respectively of FIGS. 1 to 3.

In addition, to clarify the description and the claims, the longitudinal, vertical and transverse terminology will be adopted without limitation. with reference to the trihedron L, V, T indicated in the figures, the term "vertical" being used without reference to earthly gravity.

For the different embodiments, the same references may be used for identical elements or providing the same function, for the sake of simplification of the description.

FIG. 1 shows a sandwich type composite panel adapted to enter into the manufacture of an acoustic attenuation structure of a turbojet engine nacelle.

The panel 10 has a first composite skin 12, forming the outer skin of the panel 10, also called acoustic skin, and a second composite upper skin 14 forming the inner skin of the panel 10.

The first skin 12 and the second skin 14 extend transversely, perpendicularly to the vertical direction, and define vertically between them a space in which a cellular core 16 forming a resonator is housed.

The honeycomb core 16 is here of the honeycomb type, and preferably a honeycomb 16 of aluminum, adapted to absorb the acoustic waves.

The first lower skin 12 is delimited by a lower face 18 forming the wall of the panel 10, and an upper face 20 arranged opposite the cellular core 16.

The panel 10 shown in FIGS. 1 to 4 has undergone a repaired deterioration by the implementation of a repair method according to the invention, the method described hereinafter.

The method comprises a preliminary step of cutting which consists in forming a hole 22 in the panel 10 in order to remove the damaged part of the panel 10, the hole 22 passing through the panel 10 from one side to the other.

As can be seen in FIG. 2, the first lower skin 12 is formed of several superimposed fibrous folds 24, the hole 22 being made in these folds 24 so as to give the hole 22 a generally frustoconical cross-section of decreasing section from the upper face 20, to the lower face 18 of the first skin 12.

More particularly, the hole 22 is delimited by a peripheral edge 26 (see FIG. 1) which forms steps 28 (see FIGS. 2 and 3) in the first skin 12, each step 28 extending circumferentially around the hole 22 each defining a transverse housing adapted to receive an associated fibrous fold.

The housing 30 thus formed are superimposed vertically, each housing 30 completely covering and overflowing the housing 30 directly below.

For this purpose, the first skin 12 is cut fold by fold, from top to bottom, starting with the upper face 20 of the first skin 12, so that each fold 24, except the lower fold, leaves exposed a strip 32 globally peripheral.

The hole 22 formed in the panel 10 during the cutting step has a polygonal shape.

Following the cutting step, the method comprises an apparatusing step which consists in arranging a bearing face 34 of a forming tool 36 against the lower face 18 of the first skin 12.

The bearing face 34 of the tool 36 has a shape substantially identical to the shape of the portion removed from the lower face 18 of the first skin 12, so that the bearing face 34 of the tool 36 is support without play on the lower face 18 of the first skin 12, in particular at the periphery of the hole 22.

The fitting step is followed by a step of reconstituting the first skin 12 which consists, in a first step, of closing the hole 22 by means of a repair part 38.

The repair part 38 is formed by repair folds 40 which are deposited successively in each housing 30 of the hole in a superimposed manner, starting by depositing a first fold 40 on the bearing face 34 of the forming tool 36.

As can be seen in FIG. 2, the first repair fold 40, here the lower fold 40, is flush with the lower face 18 of the first skin 12, thus allowing the first skin 12 to regain its original aerodynamic characteristics. .

Each repair fold 40 has a geometric shape complementary to the shape of the associated housing 30.

With reference to FIG. 3, each fold 40 of repair is supported both on the fold 40 of directly lower repair and on the strip 32 of the fold 24 of directly lower origin, such an arrangement promoting cohesion between the folds 24 of origin of the first skin 12 and the folds 40 reported from the repair part 38.

Preferably, at the end of the reconstitution step, the tool 36 is removed. The tool 36 does not remain in place in any case, and is removed at the latest after the implementation of the repair process.

The reconstitution step includes a consolidation step, during which the repair part 38 is polymerized. This consolidation step corresponds to a first cooking step which will be described later.

Once the repair part 38 is consolidated, the method comprises a fixing step which consists in fixing the repair part 38 on the lower skin 12 by mechanical fixing means 42, represented in FIGS. 3 and 4.

By means of mechanical fastening 42 is meant any fastening means other than a fastening of the type by gluing.

Preferably, the mechanical fixing means 42 are inserts made of metal designed to withstand high mechanical stresses, and preferably rivets 42, the rivets having the advantage of not, or little, projecting from the face lower 18 of the first skin 12. Alternatively, we can use screws, including countersunk screws.

In the example of Figure 4, the fastening means comprise two rows of rivets 42 which are arranged on a peripheral strip of the repair part 38, each rivet 42 extending vertically so as to cross and connect the repair part 38 on the lower skin 12. Alternatively, it may of course provide a single row of fastening means, or more than two rows of fixing means, for example three or four rows.

Advantageously, the upper part of the repair part 38 has an extra thickness 44 of reinforcement illustrated in Figure 3, which is formed of fibrous folds for example, and which is designed to promote the anchoring of the rivets 42 in the repair part 38 The extra thickness 44 is for example of the order of one millimeter thick, the cumulative thickness of the plies 40 of the repair and the extra thickness 44 being of the order of two millimeters. In the example, the method according to the invention also comprises a step of repairing the honeycomb 16 which consists in filling the hole formed in the honeycomb 16 by a core 46 of honeycomb, visible at the figure 1 .

For this purpose, the core 46 is arranged on the upper face of the repair part 38.

Also, the method comprises a step of repairing the second upper skin 14 of the panel 10, which consists in closing the hole formed in the second skin 14 by an attached closure plate 48.

As can be seen in FIG. 1, the plate 48, which is for example formed of fibrous folds, extends transversely over the hole formed in the second skin 14, the plate 48 comprising a peripheral band 50 which extends on the second skin 14.

The peripheral band 50 is fixed on the second skin 14 by fastening means, for example rivets 52.

Preferably, the original honeycomb 16 is equipped with a plurality of inserts 54 which extend vertically and which are designed to promote the anchoring of rivets 52 associated.

As mentioned above, the process according to the invention, and more particularly the reconstitution step, comprises a first firing step which consists in cooking the repair part 38. For this purpose, the plies constituting the repair part 38 are impregnated with a resin, in particular a polymer resin. The first firing step, which consolidates the repair part 38, is performed before the fixing step.

The method according to the invention may further comprise a second firing step, which consists in cooking the plate 48, in order to polymerize it. For this purpose, the plate 48 is impregnated with a resin, in particular a polymer resin. The second cooking step is performed after the first cooking step, and in particular after the fixing step.

Advantageously, the support face 34 of the forming tool 36 is designed to rise in temperature so as to bake the repair part 38 during the first firing step (or consolidation step). For this purpose, the tool 36 includes heating means, such as an electrical resistor for example. In this case, the first firing step is performed before the fixing step of the repair part 38. The second firing step is performed after this fixing step. In another aspect, the method comprises an acoustic piercing step which consists in making perforations 56 in the repair part 38 to favor the acoustic absorption characteristics of the panel, as can be seen in FIG. 4.

The present description of the invention is given by way of non-limiting example.

Claims

1. A method for repairing a composite panel (10), said panel comprising at least a first composite skin (12) which is delimited by a bottom wall-forming face (18) of the panel (10), and an upper face (20) ), the method being characterized in that it comprises:

- a cutting step of forming a step (22) stepped at least in the first skin (12), the hole (22) having a generally frustoconical shape of decreasing section towards the underside (18) of the first skin ( 12), and the hole (22) being delimited by a peripheral edge (26) forming steps (28) in the first skin (12), each step (28) extending circumferentially around the hole (22). delimiting a housing (30) adapted to receive a fibrous fold,

an equipment step consisting in arranging a bearing face (34) of a forming tool (36) against the lower face (18) of the first skin (12), the bearing face (34) of the tool (36) having a shape conforming to the underside (18) of the first skin (12), at least at the periphery of the hole (22),

- A step of reconstituting the first skin (12) which consists of closing the hole (22) by means of a repair part (38), the repair part (38) being formed by fibrous folds (40) deposited superimposedly in each associated housing (30) formed by the hole (22), starting by depositing a first fibrous fold (40) on the bearing face (34) of the forming tool (36), the first fibrous fold being flush with the underside (18) of the first skin (12),

- An attachment step of fixing the repair part (38) on the first skin (12) by mechanical fixing means (42).

2. Method according to claim 1, characterized in that it comprises an acoustic piercing step which consists in making perforations in the repair part (38) to promote the acoustic absorption characteristics of the panel (10).

3. Method according to any one of the preceding claims, characterized in that the fastening means (42) comprise metal inserts which each extend substantially perpendicularly to the first skin (12) and which are designed to connect the piece of repair (38) on the first skin (12).

4. Method according to any one of the preceding claims, characterized in that the repair part (38) forms an extra thickness (44) of reinforcement, which is arranged opposite the lower face (18) of the first skin (12), and which is designed to promote the anchoring of the fastening means (42) in the repair part (38).

5. Method according to any one of the preceding claims, characterized in that the hole (22) formed in the panel (10) during the cutting step has a polygonal shape.

6. Method according to any one of the preceding claims, characterized in that the composite panel (10) comprises a honeycomb core (16), for example an aluminum honeycomb, which is cut during the step of cutting, the method comprising a step of repairing said cellular core (16) which consists in filling the hole formed in the alveolar core (16) by a complementary honeycomb core (46).

7. Method according to any one of the preceding claims, characterized in that the panel (10) comprises a second skin (14) composite which extends substantially parallel to the first skin (12) and which is cut during the cutting step, the method comprising a step of repairing the second skin (14) which comprises closing the hole formed in the second skin (12) by a plate (48) reported closure.

8. Method according to one of claims 1 to 7, characterized in that the reconstitution step comprises a first firing step, which consists in baking the repair part (38).

9. Method according to the preceding claim, characterized in that the bearing face (34) of the forming tool (36) is designed to rise in temperature so as to bake the repair part (38), the tool Forming apparatus (36) comprising a heating device, such as an electrical resistor.

10. Method according to one of the preceding claims, wherein the forming tool (36) is removed at the latest after the repair process, and in particular at the end of the reconstitution step.

1 1. Process according to one of claims 8 to 10, characterized in that it comprises a second cooking step which consists in cooking the closing plate (48), the second cooking step being implemented after the first cooking step .

A composite panel (10) having at least a first composite skin (12) which is delimited by a bottom wall-forming face (18) of the panel (10), and a top face (20), the panel being characterized by it comprises a portion repaired according to a method according to one of the preceding claims.

14. Aircraft turbojet engine nacelle comprising a panel according to the preceding claim.