Method for producing a hybrid one-piece composite structure and door produced by said method

Abstract

The invention relates to a method for producing a composite stiffened structure comprising a panel (3) and at least one stiffener (4). The production method comprises the following steps for obtaining a one-piece core (2): - covering the elements of the panel (3) and the stiffeners (4) with fibres; - joining the elements of the stiffeners (4) to the panel (3); - curing the assembly and obtaining a one-piece core (2); - machining the one-piece core (2); - producing additional structural elements intended to stiffen the stiffened structure and support installation parts; - combining the additional structural elements with the machined one-piece core (2).

Description

[0001] DESCRIPTION

[0002] METHOD FOR MANUFACTURING A HYBRID MONOBLOCK COMPOSITE STRUCTURE AND DOOR PRODUCED THEREBY

[0003] FROM THE PROCESS

[0004] TECHNICAL FIELD

[0005] The invention relates to a method of manufacturing a hybrid monobloc composite structure, i.e. a structure of which at least one part is monobloc.

[0006] The invention also relates to a door produced using this manufacturing method. In particular, vehicle doors - aircraft, trains, automobiles, ships - and buildings are intended to be manufactured using this method.

[0007] In the aeronautics industry in particular, aircraft doors allow people and equipment to enter and exit the aircraft cabin. These doors are subject to numerous constraints and carry a variety of equipment: their mass production requires optimization of the manufacturing flow and raw materials.

[0008] STATE OF THE ART

[0009] An aircraft door is manufactured by assembling many separate parts to meet the needs of its complex structure. Indeed, such a door is integrated into the fuselage and creates a partition between the interior of the aircraft, the cabin, and the external environment. During the flight of such an aircraft, the pressure difference between the cabin and the external environment remains significant, subjecting the door to numerous constraints. An aircraft door also carries equipment such as opening and locking mechanisms, safety devices and other equipment that are attached to the door. To meet these requirements, the structure of such a door is generally made from a panel machined and / or shaped to the dimensions of the door frame, this panel being stiffened with added stiffeners and assembled using mechanical fasteners.

[0010] Traditionally, panels and stiffeners are made of metal: metals have mechanical properties that provide robustness to the door elements, allowing them to fulfill their role as an interior-exterior interface and to support the weight of equipment. Conventional metal machining and shaping techniques in the aeronautical field include quality controls and monitoring of the parts produced to ensure the assembly of a door in accordance with current standards. However, these metal doors are heavy and require significant assembly and installation time: they do not correspond to the development of new aircraft with reduced mass.In addition, the assembly of multiple parts, in particular by rivets, locally creates sites of concentration of forces which can lead to static ruptures or propagation of cracks in these parts: an aircraft door maintenance program is then put in place to periodically inspect these sites of concentration of forces.

[0011] Furthermore, patent document US2009078826A1 presents a conventional aircraft door whose panels and stiffeners are made of composite fiber. Each of these structural parts is manufactured individually using a fiber deposition process followed by resin transfer molding. These composite fiber doors are lighter than their metal counterparts. However, assembly and installation times remain significant.

[0012] To reduce the use of assembly rivets and to extend the interval between maintenance inspections, patent document US2004021038A1 presents a metal door whose panels and stiffeners are machined from a single piece. However, the mass of such a door remains close to that of a conventional door.

[0013] Another approach in the manufacture of complex composite material structures uses a so-called monobloc process in which the entire structure is manufactured completely in a single block, i.e. without the need to assemble the constituent elements. However, this manufacturing process requires complex specific tools as well as extended implementation times. DISCLOSURE OF THE INVENTION

[0014] In order to overcome the drawbacks of the state of the art set out above, the main objective of the invention is to improve the manufacture of a composite structure

[0015] - such as an aircraft door - in terms of optimization and adaptation to high production increases.

[0016] To do this, the invention provides for producing an aircraft door from a composite material, the raw form of which, hereinafter referred to as the "core", is produced using a single-piece process, i.e., constituting only a single piece. This core is then machined and additional structural elements are assembled there. This two-stage production process makes it possible to optimize manufacturing times and processes, as well as to adapt the structure of the door to the rest of the structure to which it is then installed.

[0017] More specifically, the present invention relates to a method for manufacturing a composite stiffened structure comprising a panel and at least one stiffener. The manufacturing takes place according to the following steps for obtaining a single-piece core:

[0018] - fiber draping of the panel elements and stiffeners;

[0019] - assembly of said stiffener elements on the panel;

[0020] - cooking of the assembly and obtaining the single-block core; which are followed by stages of finalization of the stiffened structure from the single-block core into a hybrid single-block structure, namely:

[0021] - machining of the single-block core;

[0022] - manufacture of additional structural elements intended to stiffen the stiffened structure and to support installation parts;

[0023] - combination of additional structural elements to the machined single-block core.

[0024] Advantageously, the production of a single-piece core consisting of the panel and stiffeners makes it possible to reduce the assembly phases and the number of mechanical fasteners due to the curing of the single-piece core. In addition, the structure of the core is adaptable by modifying the draping and the position of the panel elements and stiffeners. Also advantageously, the manufacture and assembly of additional structural elements makes it possible to simplify the production of the single-piece core. Indeed, the core comprising fewer elements, it is then faster and easier to manufacture. In addition, the additional structural elements complete the structure of the core allowing it to be adapted to different needs. The same core can thus be suitable for different uses, the additional structural elements providing the specificities necessary for these different needs and uses.This core therefore constitutes a base that is easily adaptable during its design but also during its assembly with the addition of additional structural elements. Part management is thus facilitated by limiting the number of part references.

[0025] According to certain preferred forms of implementation taken alone or in combination:

[0026] - the draping step is automated;

[0027] - draping superimposes structural fiber plies with sacrificial fiber plies which will be partially eliminated during the core machining step;

[0028] - the machining of the single-piece core includes steps of trimming the panel and cutting at least one end of the stiffeners to predetermined dimensions;

[0029] - the machining of the single-piece core is carried out at the level of the sacrificial fiber folds;

[0030] - the draping of the stiffener elements is carried out directly according to a predetermined final geometry;

[0031] - additional structural elements are made of composite material and draped in fiber;

[0032] - cooking is carried out in an autoclave-type oven, and

[0033] - additional structural elements and the single-piece core are fired simultaneously.

[0034] Advantageously, the automation of draping allows for productivity gains in terms of time and raw materials.

[0035] Also advantageously, draping the stiffeners according to a predetermined final geometry eliminates the need for traditional flat draping followed by hot forming. Curing the assembly in an autoclave saves time and equipment because a single curing tool is sufficient for all the structural elements. In addition, depending on the size of the autoclave, several cores can be cured at the same time, this process making it possible to meet high-speed production needs while limiting the required tooling.

[0036] The invention also relates to a hybrid single-piece composite aircraft door with a stiffened structure produced using the preceding method, this door comprising:

[0037] - a single-piece core made of composite material constituting a panel and at least one stiffener, each stiffener having a web and a heel;

[0038] - additional structural elements assembled in combination with this single-piece core to form a functional whole.

[0039] Advantageously, such a hybrid monobloc composite structure door makes it possible to combine the structural benefits of a monobloc structure with the flexibility of assembled structures. Indeed, the monobloc core provides automation of manufacturing as well as assembly of a part of the structure by co-consolidation during curing, eliminating the use of mechanical fasteners, which can harm the mechanical properties of the material. Furthermore, a purely monobloc structure has the disadvantages of its lack of adaptability, particularly during the molding phases and the necessary tools, which rules out their use for mass production. The production of the monobloc core therefore makes it possible to simplify the structure to be molded as well as the necessary tools, the rest of the door structure being assembled in a traditional manner by conventional fasteners.

[0040] According to preferred embodiments:

[0041] - the sacrificial fiber plies of the monobloc core are made of fiberglass;

[0042] - additional structural elements are selected from edge frames installed on the panel perpendicular to the stiffeners, stops, joints, functional element supports and / or door fasteners.

[0043] Advantageously, glass fibers are neutral and insulating, providing a barrier against electrochemical reactions that can lead to galvanic corrosion of metal elements. Glass fibers also allow for rapid visual inspection of the machining to verify that the structural folds have not been damaged by this machining. PRESENTATION OF FIGURES

[0044] Other characteristics and advantages of the present invention will emerge from the following reading of a detailed example of embodiment without limiting its scope, with reference to the appended figures which represent, respectively:

[0045] - figure 1, a perspective view of a hybrid single-block door according to the invention;

[0046] - figure 2, an exploded perspective view of the elements of the panel and the stiffeners of the door according to figure 1;

[0047] - figure 3, a perspective view of the single-block core after firing of this door;

[0048] - figure 4, a perspective view of the single-block core after cutting out said door, and

[0049] - figure 5, an exploded perspective view of the single-piece core according to figure 4 and of additional structural elements of the door according to the previous figures.

[0050] DETAILED DESCRIPTION

[0051] In the figures, identical reference signs refer to the same element as well as to the corresponding passages of the description.

[0052] Figure 1 shows a hybrid monoblock composite aircraft door 1 with a stiffened structure. This door 1 comprises:

[0053] - a single-piece core 2 made of composite material constituting a panel 3 and six stiffeners 4, each stiffener 4 having a web 4a and a heel 4b;

[0054] - edge frames 5a installed on the panel 3 perpendicular to the stiffeners 4;

[0055] - peripheral joints 5b;

[0056] - stops 5c for holding door 1 in the fuselage of the aircraft (not shown);

[0057] - 5d closing sheets which limit the deformation of the ends of panel 3 during pressurization of the aircraft;

[0058] - door fasteners (not shown) on this fuselage;

[0059] - functional element supports (see figure 2), the frames 5a, the joints 5b, the stops 5c, the fasteners and the supports being assembled in combination with the single-piece core 2 to form a functional and structural assembly.

[0060] The door structure 1 thus obtained is structurally complete and sized for its installation in the aircraft after fixing its functional elements, in particular the handles, fasteners and door opening systems.

[0061] Figure 2 illustrates the bottom element 3c of the panel 3 and the intermediate elements 4c, end elements 4d and upper elements 4e of the stiffeners 4 used to manufacture the single-piece core 2 of the aircraft door 1, the latter forming a composite stiffened structure. The method of manufacturing such a door 2, which in this embodiment comprises a panel 3 and six stiffeners 4, takes place according to the following steps for obtaining the single-piece core:

[0062] - fiber draping of the bottom element 3c of the panel 3 as well as the intermediate elements 4c, extreme 4d and upper 4e of each stiffener 4;

[0063] - assembly of the intermediate elements 4c, extreme 4d and upper 4e of the stiffeners 4 on the bottom element 3c of the panel 3;

[0064] - cooking the assembly and obtaining the single-piece core.

[0065] The stiffened structure is then finalized from the single-block core 2 into a hybrid single-block door structure 1 obtained according to the following steps:

[0066] - machining of the single-block core 2;

[0067] - manufacture of additional structural elements (not shown) intended to stiffen the stiffened structure and to support installation parts;

[0068] - combination of additional structural elements to the machined single-piece core 2. More precisely, the bottom element 3c of the panel 3 and the intermediate elements 4c, extreme 4d and upper 4e of the stiffeners 4 are draped in carbon fibers and superimposed on three levels:

[0069] - the first level 6a is composed of the bottom element 3c forming the external surface of the panel 3 - that is to say the skin 3a;

[0070] - the intermediate level 6b is composed of the intermediate elements 4c and extreme elements 4d which form the internal face of the panel 3 as well as the core 4a (see figures 1, 4) and the support 4f of the heels 4b (see figures 1, 4) of the stiffeners 4;

[0071] - the last level 6c is composed of the upper elements 4e forming the said heels 4b of the stiffeners 4.

[0072] The bottom elements 3c and upper elements 4e - respectively of the first level 6a and the last level 6c - have a surface topology dimensioned to the size of the panel 3 and the heels 4b of the stiffeners 4. The base elements 4c and 4d of the intermediate level 6b have generally “U” or “L” shaped sections including:

[0073] - a flat sole 4g ​​opposite the bottom element 3c, the set of soles 4g extending to cover the bottom element 3c;

[0074] - one (for the extreme elements 4d located at the ends of the panel 3) or two (for the intermediate elements 4c) half-webs 4h of the stiffeners 4 perpendicular to each flange 4g and at the end 4i of this flange 4g, and

[0075] - a 4f heel support perpendicular to each 4h half-soul and at the end of this 4h half-soul.

[0076] The intermediate elements 4c and extreme elements 4d of the intermediate level 6b are arranged on the skin 3a via the flanges 4g and juxtaposed with each other by bringing the half-webs 4h into contact, two half-webs 4h joined together forming a web 4a of stiffener 4. In this embodiment example, seven intermediate elements 4c and extreme elements 4d are present in the intermediate level 6b to produce six stiffeners 4. To do this, the two extreme elements 4d then have a single half-web 4h with an “L” section while the intermediate elements 4c have two half-webs 4h with a “U” section.

[0077] The draping of the base 3c, intermediate 4c, extreme 4d, and upper 4e elements of the panel 3 and of the stiffeners 4 is automated and carried out directly according to their predetermined final geometry, in particular according to the “L” and “U” sections of the base elements 4c, 4d of the stiffeners 4 of the intermediate level 6b. Alternatively, these intermediate 4c and extreme 4d elements can be draped flat and then preformed in order to obtain their final L and U geometry.

[0078] The single-piece core 2 after curing in an autoclave-type oven is illustrated in Figure 3: after assembly of the base 3c, spacers 4c, end 4d, and upper 4e elements of the panel 3 and the stiffeners 4, resin is injected into the assembly and then the assembly is placed in the oven for curing. This resin can be of the thermosetting or thermoplastic type. Alternatively, the draping can be carried out with pre-impregnated fibers: curing is then carried out directly without resin injection. Curing outside the autoclave oven is also possible, in particular by a known process of resin infusion into the assembly.

[0079] The draping superimposes structural fiber plies 7a with sacrificial fiber plies 7b which will be partially eliminated during the machining step of the single-piece core 2. The structural fiber plies 7a are responsible for the strength of the structure. The draped sacrificial plies 7b make it possible to facilitate the assembly of the single-piece core 2 by using elements with simple geometry, then a portion of these sacrificial plies 7b is eliminated by machining in order to obtain the final dimensions of the single-piece core 2. Thus, these sacrificial plies 7b are located at the edge 3b of the panel 3 and at the ends 4j of the stiffeners 4. In this embodiment, the structural plies 7a are made of carbon fibers for mechanical strength and the sacrificial plies 7b are made of glass fibers to facilitate their machining.

[0080] Figure 4 shows the single-piece core 2 after machining, this being carried out at the level of the sacrificial fiber plies 7b. This machining includes steps of trimming the panel 3 on its edge 3b and cutting the end 4j of the stiffeners 4 to predetermined dimensions. Thus the contour of the panel 3 is advantageously rounded and its dimensions are adjusted to the opening that this panel 3 of door 1 completes. In addition, the stiffeners 4 are cut at their web 4a and their heel 4b in order to meet the dimensional constraints relating to the installation of said door 1. Figure 5 illustrates the additional structural elements which have been manufactured in order to be fixed on the single-piece core 2.These additional structural elements are intended to consolidate the stiffened structure of the single-piece core 2 so as to be able to support the addition of installation parts: these additional structural elements are chosen from edge frames 5a installed on the panel perpendicular to the stiffeners 4, stops 5c, closing sheets 5d, system element supports, joints 5b and / or door fasteners (not shown). In particular, the edge frames 5a may be made of composite material and draped in fibers, the curing of the edge frames 5a and the single-piece core 2 being advantageously carried out simultaneously.

[0081] These additional structural elements are assembled using traditional methods using rivet-type fasteners or screw / nut assemblies, or epoxy glue or equivalent. The resulting door is therefore a hybrid door in which structural elements, composite or metallic, are assembled on a single-piece composite core.

[0082] This approach allows the automation of the layup and therefore the orientation of the fibers of the composite parts of the door, such automation leading to an improvement in the precision and speed of manufacturing these composite parts. The resulting single-piece core also reduces the number of mechanical fasteners that can harm the mechanical properties of the material: in fact, each drill hole creates a zone of local over-stressing from which the door structure is likely to crack.

[0083] These mechanical fasteners also remain particularly advantageous regarding the adaptability of the structure with regard to the means developed in equipment, research and costs during possible modifications or developments of the structure. A hybrid structure with a single-piece core, the subject of the present invention, therefore combines the advantages of a single-piece structure and the benefits of limited use of mechanical fasteners. The invention is not limited to the examples of embodiment and implementation described and shown. Thus, the draping of the panel and stiffeners can be carried out manually.

[0084] In addition, materials other than composites can be used to produce additional structural elements, in particular plastic or metal such as aluminum, steel, nickel or titanium alloys: these elements are then produced using traditional machining and shaping methods.

[0085] Also other types of fibers can be used for draping: in particular carbon fibers for all structural and sacrificial plies, or any other fiber offering suitable mechanical characteristics with the use of the obtained hybrid monoblock structure.

Claims

DEMANDS 1. Method for manufacturing a stiffened composite structure comprising a panel (3) and at least one stiffener (4), the manufacture taking place according to the following steps to obtain a monobloc core (2): - fibre draping of the elements (3c, 4c, 4d, 4e) of the panel (3) and of the stiffeners (4); - assembly of said elements (3c, 4c, 4d, 4e) of the stiffeners (4) on the panel (3); - baking of the assembly and obtaining the monobloc core (2); the process being characterized in that the baking is carried out in an autoclave-type oven and in that the steps of obtaining the monobloc core (2) are followed by subsequent steps of finalizing the stiffened structure from the monobloc core (2) into a hybrid monobloc structure: - machining of the one-piece core (2); - manufacturing of additional structural elements intended to consolidate the stiffened structure and to support installation parts; - combination of additional structural elements to the machined monobloc core (2).

2. A manufacturing process according to the preceding claim, wherein the draping step is automated.

3. A manufacturing method according to any one of claims 1 to 2, wherein the draping step superimposes structural fiber plies (7a) with sacrificial fiber plies (7b).

4. A manufacturing method according to any one of claims 1 to 3, wherein the machining of the one-piece core (2) comprises steps of trimming the panel (3) and cutting at least one end (4j) of the stiffeners (4) to predetermined dimensions.

5. A manufacturing method according to any one of claims 1 to 4 in combination with claim 3, wherein the machining of the one-piece core (2) is carried out at the level of the sacrificial fiber plies (7b).

6. Manufacturing method according to any one of claims 1 to 5, wherein the draping of the elements (4c, 4d, 4e) of the stiffeners (4) is carried out directly according to a predetermined final geometry.

7. A manufacturing method according to any one of claims 1 to 6, wherein additional structural elements are made of composite material and draped in fiber.

8. A manufacturing method according to any one of claims 1 to 7 in combination with claim 7, wherein additional structural elements and the one-piece core (2) are baked simultaneously.

9. Hybrid one-piece composite aircraft door (1) with a stiffened structure produced by the process according to any one of claims 1 to 8, characterized in that this door (1) comprises: - the one-piece core (2) made of composite material constituting a panel (3) and at least one stiffener (4), each stiffener (4) having a web (4a) and a heel (4b); - additional structural elements (5a, 5b, 5c) assembled in combination with this monobloc core (2).

10. Hybrid one-piece composite door (1) according to claim 9 wherein the sacrificial fibre plies (7b) of the one-piece core (2) are made of glass fibre.

11. Hybrid one-piece composite door (1) according to any one of claims 9 to 10 wherein the additional structural elements are selected from edge frames (5a) installed on the panel (3) perpendicular to the stiffeners (4), stops (5c), seals (5b), functional element supports and / or door fasteners.

Images