EP2512915A2

EP2512915A2 - Method for the production of a composite trailing edge panel for an aircraft element

Info

Publication number: EP2512915A2
Application number: EP10805808A
Authority: EP
Inventors: Denis Millepied; Jean-Luc Pacary; Arnaud Bertrand; Valérian MONTAGNE
Original assignee: Societe Lorraine de Construction Aeronautique SA SLCA
Current assignee: Safran System Aerostructures SAS
Priority date: 2009-12-18
Filing date: 2010-12-14
Publication date: 2012-10-24
Also published as: BR112012011517A2; WO2011073573A2; CN102656085A; RU2012129704A; WO2011073573A3; FR2954269B1; FR2954269A1; CA2782708A1; US20120267479A1

Abstract

The invention relates to a method for producing a structural composite trailing edge panel (1) for an aircraft element having: an upper surface (3), a lower surface (5), and a trailing edge (7) connecting said upper (3) and lower (5) surfaces, the upper surface (3) and the lower surface (5) being connected by transverse stiffeners (9), at least one longitudinal spar (10) being positioned in such a way that the directrix (Δ₁₀) of each longitudinal spar (10) and the directrix (Δ₉) of the transverse stiffeners (9) are not collinear and the structuring panel (1) being made up of a one-piece component forming the upper surface (3), the lower surface (5), the trailing edge (7), the transverse stiffeners (9) and the spar or spars (10).

Description

Composite structuring panel of trailing edge for an element of a

aircraft

The present invention relates to a composite structuring panel for a trailing edge of an element of an aircraft.

The invention also relates to an aircraft element comprising such a panel.

Composite panels are panels frequently used in aerospace because they allow to lighten considerably the aircraft.

Some parts of aircraft require structuring panels ensuring good mechanical strength. Particularly the trailing edges, such as those of aircraft control surfaces.

Sandwich-type composite structuring panels are commonly used, comprising a honeycomb core structure placed between an inner skin and an outer skin.

Typically, the inner skin and the outer skin each consist of one or more fibrous folds pre-impregnated with resin which is then polymerized during a firing step.

Other processes use dry fibrous folds, that is to say not pre-impregnated with resin, the resin being subsequently applied during a cooking step during which it is forced by aspiration to diffuse between the fibrous folds.

A composite sandwich panel may also comprise several central layers, of the same type or of different types, the central layers may themselves be separated by a layer of composite material.

The central layers may, for example, be of cellular type, foam or include one or more fusible inserts.

Composite sandwich panels using a honeycomb core or foam, for example, help reduce the mass of objects while maintaining or increasing their mechanical properties.

However, these types of panels are generally not suitable for the manufacture of trailing edges.

In order to solve this problem, in application FR09 / 02579, a composite structuring panel of trailing edge for an element is proposed. an aircraft having an upper surface, a lower surface and an edge connecting said upper and lower surfaces.

The upper surface and the lower surface are connected by transverse stiffeners.

The structuring panel consists of a single piece forming the upper surface, the lower surface, the trailing edge and the transverse stiffeners.

Despite the advantages such a panel provides, it may be desirable to further limit the buckling of the upper and lower skins by ensuring good flexural and torsional stiffness.

An object of the present invention is therefore to provide a panel for imitating the buckling of the upper and lower skins and to improve the structural mechanical strength while being simple to achieve.

For this purpose, according to a first aspect, the subject of the invention is a composite structuring panel of trailing edge for an element of an aircraft having:

an upper surface,

a lower surface, and

a trailing edge connecting said upper and lower surfaces, the upper surface and the lower surface being connected by transverse stiffeners,

Characterized in that at least one longitudinal spar is arranged so that the director axis of each longitudinal spar and the direction axis of the transverse stiffeners are not collinear and in that the structuring panel consists of a one-piece piece forming the upper surface, the lower surface, the trailing edge, the transverse stiffeners and the longitudinal longitudinal member (s).

"Steering axis" means the axis directing a spar or a transverse stiffener according to the largest dimension of the latter.

The presence of one or more longitudinal longitudinal members disposed substantially perpendicular to the transverse stiffeners makes it possible to limit the buckling of the upper and lower skins and to improve the structural mechanical strength of the panel of the invention in two substantially perpendicular directions of the panel. the invention. In addition, the panel of the invention being entirely made in one piece has a simple manufacture to achieve. Preferably, the director axis of each longitudinal spar and the direction axis of the transverse stiffeners are substantially perpendicular.

Preferably, at least one longitudinal spar is disposed between two transverse stiffeners which locally reinforces the structural strength of the panel of the invention.

Preferably, the skin forming said panel comprises a plurality of pleats, one or more inner plies form the longitudinal or longitudinal spars.

Preferably, the panel of the invention comprises reinforcing plies between the inner plies which reinforces the longitudinal or longitudinal longitudinal members and stiffeners.

According to another aspect, the subject of the invention is a method for manufacturing a panel of the invention, in particular a composite trailing edge panel for an element of an aircraft, characterized in that it comprises:

a first step (A) in which first cores and at least one second core are deposited, each at least partially surrounded by a skin of draping on a base skin, in two non-collinear directions, so that said base skin can be folded back on itself;

a second step (B) in which the base skin is folded over the first and second (s) draped cores;

a third step (C) in which the panel thus obtained is polymerized so as to integrate the folds of the draping into the base skin to form the transverse stiffeners and the longitudinal sill (s); and

a fourth step (D) in which the first cores and the second cores are removed in order to obtain the structuring panel.

Preferably, the director axis of each longitudinal spar and the transverse stiffeners are substantially perpendicular.

Preferably, at least one longitudinal spar is disposed between two transverse stiffeners.

Preferably, the skin forming said panel comprises a plurality of pleats, one or more inner plies form the longitudinal or longitudinal spars. Preferably, the panel comprises reinforcing plies between the inner pleats.

Preferably, the second core or cores have a decreasing height along the cross section of said cores which allows a good aerodynamic line of the panel of the invention.

Preferably, each first and second cores are draped by a monolithic type draping skin having a plurality of plies.

Preferably, in step A, the first cores are arranged before the trailing edge so as to form a space between the trailing edge and the first cores, in which space one or more substantially parallel second cores are installed at the trailing edge. .

According to yet another aspect, the subject of the invention is an element of an aircraft comprising at least one structuring panel of the invention or obtained according to the method of the invention.

Preferably, the element of the invention is an airplane rudder.

The invention will be better understood on reading the nonlimiting description which follows, with reference to the appended figures.

FIG. 1 is a perspective view of a panel of the invention,

FIG. 2 is a perspective bottom view of a variant of the embodiment of the panel of FIG. 1,

FIG. 3 is an enlarged front view of the embodiment of FIG. 1, and

- Figures 4 to 6 are perspective views of the method of manufacturing a panel according to the invention.

As illustrated in FIG. 1, the panel 1 of the invention comprises an upper surface 3, a lower surface 5 and an edge 7 connecting the upper 3 and lower 5 surfaces. The panel 1 of the invention defines a trailing edge 7 It has been obtained from the panel 1 of the invention which simplifies the manufacture of the latter.

The upper surface 3 and the lower surface 5 are connected by transverse stiffeners 9 and at least one or more longitudinal longitudinal members 10, desd it raid isseurs 9 and said longitudinal member (s) 1 0 being integrated therewith. At least one longitudinal spar 10 is arranged such that the director axis Δ10 of each long longitudinal spar 1 0 and the directing axis Δ ₉ transverse stiffeners 9 are not collinear. Advantageously, the panel of the invention 1 has a very good structural strength in two non-parallel directions.

Preferably, the director axis Δ10 of each longitudinal spar 10 and the directing axis Δ ₉ of the transverse stiffeners 9 are substantially perpendicular.

By "longitudinal" is meant a direction substantially co-linear with the steering axis 8 of the trailing edge 7. As illustrated in FIGS. 1 and 2, the steering axis 8 of the trailing edge may be substantially integral with the steering axis Δ10 of each longitudinal spar 10 and / or substantially perpendicular to the directing axis Δ ₉ of the transverse stiffeners 9.

According to a variant not shown, the di recteu r Δ ₉ axis of the transverse stiffeners 9 may be non-collinear with the director 8 direction of the trailing edge without being perpendicular thereto. Likewise, the axis Δ10 of each longitudinal spar 10 may be non-collinear with the directing axis 8 of the fuite edge and also non-collinear with the axis of the irisector Δ ₉ of the transverse stiffeners 9.

A "transverse" direction, a direction substantially perpendicular to the planes formed by the upper surface 3 and the lower surface 5.

The longitudinal longitudinal member (s) 1 0 is typically placed at the end of the transverse stiffeners 9 facing the trailing edge 7. To this end, the transverse stiffeners 9 are placed at a non-zero distance from the trailing edge 7.

The panel 1 of the invention can thus comprise a single longitudinal spar or, conversely, a plurality of longitudinal spars. The use of a plurality of longitudinal members 10, in particular placed between two transverse stiffeners 9 (see FIG. 2), makes it possible to locally limit any buckling of the panel 1 of the invention. Said spar 10 then having a length at most equal to the spacing of the two transverse stiffeners 9 along the direction axis 8.

Typically, the length of a longitudinal spar 1 0 along the director axis Δ10 of the latter may take any value less than or equal to the length of the panel 1 of the invention. In the case where the longitudinal spar 10 does not have its director axis Δ10 substantially parallel to the steering axis 8 of the trailing edge, the length of said spar 10 may be greater than the length of the panel 1 of the invention without said spar 10 protruding from said panel 1 .

Similarly, the length of a transverse stiffener 9 along the director axis Δ ₉ of the latter may take any value less than or equal to the width of the panel 1 of the invention. In the case where the transverse stiffener 9 does not have its directing axis Δ ₉ substantially perpendicular to the directing axis 8 of the trailing edge, the length of said stiffener 9 may be greater than the width of the panel 1 of the invention without said stiffener 9 protrudes from said panel 1.

Furthermore, the panel of the invention 1 consists of a single piece forming the top surface 3, the lower surface 5, the edge 7 and the transverse stiffeners 9 and the longitudinal members 10. To do this, the Panel 1 of the invention may consist of a single monolithic skin.

The monolithic skin may be made of any type of suitable fabrics or fibers known to those skilled in the art which may be impregnated with epoxy resin or the like. For this purpose, mention may be made of carbon, glass or Kevlar® fibers.

As illustrated in FIG. 3, the single monolithic skin is formed of a plurality of plies 18 fused to one another by means of a polymerizable resin, such as the epoxy resin, disposed between the plies 18.

More specifically, the upper part 15 of the skin forming the upper surface 3 and the lower part 17 of the skin forming the lower surface 5 may comprise a plurality of plies 18 whose inner plies 19, 21 are arranged towards the inside of the panel 1 may extend continuously along said panel 1 from a straight section to a second straight section. The fact that the transverse stiffeners 9 and the longitudinal members 10 consist of pl is 1 8 makes it possible to obtain a composite structuring panel 1 that is highly resistant to absorbing a shock that is substantially transverse to the upper surface 3 or lower surface 5. In fact, the panel 1 of the invention is advantageously mechanically reinforced in two non-collinear directions, in particular substantially perpendicular, with respect to the plane formed by the panel 1 of the invention. The inner plies 19 can extend continuously from the lower portion 17 through the panel 1 substantially perpendicularly to the lower surface 5 by forming a portion of the folds of a transverse stiffener 9 or a spar 10 and before extending at the upper surface 3 again along the cross section.

It is the same for the other inner folds 21 of the other cross section.

Thus, the transverse raid 9 and the or longitudinal members 10 are formed by the inner plies 19 and 21 from the straight sections.

Of course, the plies 18 used may be identical in nature or different depending on the desired properties.

As the nature of folds conventionally used, there may be mentioned, inter alia, glass fibers, carbon fibers and kevlar fibers.

In the case where the folds 19, 21 participating in the reinforcements do not present to them alone or have sufficient strength or should be reinforced, it is possible in particular to sew all or part of these folds 19, 21 between them. It is also possible to insert, between the plies 19, 21, folds of reinforcements, such as folds of carbon fibers, for example, which may be present in the transverse stiffeners 9 and / or in the longitudinal members 10.

In addition, according to the invention, the panel 1 of the invention is obtained by the manufacturing method comprising:

a first step (A) in which first cores 11 and at least one second core 12, each of which is at least partially surrounded by a draping skin 15, are deposited on a base skin 13, in two directions colinear Δ10 and Δ ₉ , in particular respectively along a length and a width of said base skin 13, so that the latter can be folded back on itself (see FIG. 4);

- A second step B in which the base skin 13 is folded over the first 1 1 and second (s) 12 draped cores (Figure 5);

a third step C in which the panel thus obtained is polymerized so as to integrate the folds of the draping into the base skin 13 to form the transverse stiffeners 9 and the longitudinal members 10; and

a fourth step D in which the first cores 11 and the one or more of the cores 12 are removed in order to obtain the structuring panel (see FIG. 6). Subsequently, the expressions "at least partially surrounded" and "draped" are synonymous. Thus, the term "draping" refers to at least partially surrounding a core.

With the method of the invention, it is possible to adjust the number of folds between two transverse stiffeners 9 and also at the level of the longitudinal members 10. It is then possible to optimize the mass of the panel 1 of the invention while guaranteeing a significant longitudinal and transverse stiffness.

In addition, thanks to the method of the invention, the panel 1 is formed in one piece by melting of the folded base skin 13 on itself and the skin of the draping.

In addition, the method makes it possible to introduce the desired number of stiffeners and spar (s) as a function of the desired structural strength by increasing or decreasing the number of cores or the dimensions thereof.

E ntree, the procedure is applied by means of co ntrainte for the positioning of the stiffeners and that of the longitudinal member (s). These are placed so as to improve their structural utility.

More particularly, in step A, the first cores 11 are each surrounded at least partially by a drape 15 on the lateral sides of said cores January 1.

The second core or cores 12 are each at least partially surrounded by a drape 15 over at least a portion of a longitudinal side of said cores 12.

The first cores 11 and the second or second cores 1 2 employed have a shape suitable for forming the transverse stiffeners 9 and the longitudinal member or members 10. To do this, they typically have a cross section of substantially triangular, rectangular, square shape. or trapezoidal.

Typically, first cores 1 1 for forming the transverse stiffeners 9 are arranged before the edge 7 so as to form a space in which one or two second cores 12 are installed parallel to the edge 7 making it possible to form the longitudinal member (s) 10 (see Figure 4).

Advantageously, the first cores 11 have a decreasing height along the length of said first cores 11 so as to match the small radius of curvature of the edge 7. In addition, the second core or cores 12 have a cross-section with a decreasing height on the cross-section of said second core (s) 12 so as to match the small radius of curvature of the edge 7. Thus, it is possible to have an excellent aerodynamic profile of the structuring panel 1.

Advantageously, the first 1 1 and second (s) 12 cores are placed on the base skin 13 over a length of the latter appropriate to allow folding the base skin 13 on itself. Thus, the first 1 1 and second (s) 12 cores can be placed a distance less than half the length of said skin 13 which allows to have an upper surface 3 of length substantially equal to that of the lower surface 5.

The draping is done typically before the laying of the first cores 1 1 and the second or second cores 1 2 on the base skin 1 3. The draping is then performed by a monolithic-type draping skin 15 having a plurality of plies. , for example two or three folds so as to obtain optimum draping. Typically, the draping skin 1 5 has a number of folds lower than that of the base skin 13

The base skin 13 may comprise a number of plies greater than 2, equal to 3, 5 or more.

The skin removal may comprise a number of plies greater than 2, equal to 3, 5 or more.

The folds of the base skin 13 and the draping skin 15 are impregnated with polymerizable resin such as the epoxy resin.

In step B, the base skin 13 is folded back on itself by any means known to those skilled in the art so as to form an edge 7, an upper surface 3 and a lower surface 5. This makes it possible to achieve performing in a single operation and in a simple manner a structuring panel of trailing edge in which the trailing edge is made in the same operation as the upper and lower surfaces and ensuring structural continuity between these three elements.

The production of a structuring panel of trailing edge in which it will then be necessary to proceed to the structural connection between the upper and lower surfaces by a reported trailing edge complicates the manufacturing process of a structuring panel.

Typically, the polymerization of step C is carried out by heating to a baking temperature. The cooking temperature depends on the type of resin used to make the monoblock panel 1 of the invention. For example, if the base skin 13 and / or drape 15 is (are) made with epoxy resin, the firing temperature is between 60 ° C and 200 ° C.

This step is typically carried out in an autoclave or any heating means.

Typically, the base skin 13 and the draping skin 15 comprise fiber-based plies such as glass fibers, carbon fibers and Kevlar fibers, said fibers being impregnated with polymerizable resin when the material is baked. .

In step D, the first cores 11 and the second core or cores 12 of the panel thus formed are removed by any means known to those skilled in the art, in particular by extractors manipulated manually or automatically. The removal of the cores 11 and 12 is typically carried out according to a substantially col inear e d irection d irection q ue take the transverse stiffeners 9 or the longitudinal members or 10, if appropriate.

The panel 1 of the invention can be advantageously used in an element of an aircraft, such as an airplane rudder.

Claims

1. A method of manufacturing a composite trailing edge panel (1) for an element of an aircraft, characterized in that it comprises:

a first step (A) in which first cores (1 1) and at least one second core (12) are deposited, each at least partially surrounded by a draping skin (1 5) on a base skin (1 3); ), in two non-collinear directions (Δι ₀ , Δ ₉ ), so that said base skin 13 can be folded on itself;

a second step (B) in which the base skin (13) is folded over the first (1 1) and second (12) draped cores;

a third step (C) in which the panel thus obtained is polymerized so as to integrate the folds of the draping into the base skin (13) to form the transverse stiffeners (9) and the longitudinal sill (s) (10); and

a fourth step (D) in which the first cores (1 1) and the second corona or cores (1 2) are removed in order to obtain the structuring panel (1).

2. Method according to the preceding claim, wherein the director axis (Δ10) of each longitudinal spar (10) and the director axis (Δ ₉ ) transverse stiffeners (9) are substantially perpendicular.

3. A method according to any preceding claim, wherein at least one longitudinal spar (10) is disposed between two transverse stiffeners (9).

4. Method according to any one of the preceding claims, characterized in that the skin forming said panel (1) comprises a plurality of plies (18), one or more inner plies (19, 21) form the longitudinal longitudinal or longitudinal ( 10).

5. Method according to the preceding claim, characterized in that the panel (1) comprises reinforcing plies between the inner plies (19, 21).

6. Method according to any one of the preceding claims, characterized in that the second core or cores (12) have a decreasing height according to the cross section of said cores (12).

7. Process according to any one of the preceding claims, characterized in that each first (1 1) and second (12) cores are draped by a monolithic-type draping skin (15) having a plurality of pleats.

8. Process according to any one of the preceding claims, characterized in that, in step A, the first cores (1 1) are arranged before the trailing edge (7) so as to form a space between the trailing edge (7) and the first cores (1 1), space in which one or more second cores (12) are installed substantially parallel to the trailing edge (7).

9. Element of an aircraft comprising at least one structuring panel (1) obtained according to any one of the preceding claims.

10. Element according to the preceding claim being an airplane rudder.