Method for producing reinforced structural component made of composite material, and structural component

Abstract

A method for manufacturing a structural component comprising a skin formed from a plurality of layers of composite material, a plurality of stringers made of composite material secured to the skin, and a plurality of ribs made of composite material secured to the skin, the structural component having a plurality of sub-panels joined to each other, and each sub-panel includes a sub-skin defined by a first layer of composite material, at least one stringer and at least one rib secured to an inner surface of the sub-skin; wherein the structural component comprises a continuous sub-wall defined by a complex of outer surfaces (6b) of sub-skins (6) of interconnected sub-panels (5); a second layer of composite material is laminated on the first layer to define a continuous outer upper skin disposed in layers on the sub-wall; the skin is defined by a group of sub-skins and an upper skin, and consists of a first layer and a second layer.

Description

[0001] CROSS-REFERENCE TO RELATED APPLICATIONS

[0002] This patent application claims the priority of Italian Patent Application No. 102023000023190 filed on November 3, 2023, the entire disclosure of which is incorporated herein by reference. Technical Field

[0003] The invention relates to a method for producing a reinforced structural component made of composite material, in particular reinforced structural components with reinforcing stringers and ribs and, if necessary, with other composite reinforcing structures, such as, for example, polymer foam or honeycomb structures known as "honeycomb cores" inserted between two layers of composite material.

[0004] More particularly, this description will explicitly refer to the manufacture of fuselages of aircraft or to the manufacture of tanks subjected to pressure, and without any lack of generality thereby.

[0005] In general, the method according to the invention can be applied to all those closed composite parts having narrow interior spaces and undercuts, ie those parts defining an internal volume.

[0006] The invention also relates to a reinforced structural component made of the composite material. Background Art

[0007] Known structural components are used in the aviation industry (eg fuselages and parts thereof) and are made of composite materials.

[0008] In the prior art there is an aerospace structural component which is made of a light alloy, is reinforced with a metal material and is intended to construct a part of the fuselage of an aircraft.

[0009] As is known, fuselages are designed to ensure adequate protection of the payload (crew, passengers, cargo, etc.), but at the same time they cannot exceed the set weight limits.

[0010] Furthermore, the use of metal parts, although ensuring greater resistance, leads to an increase in the overall costs.

[0011] Therefore, structural components made of composite materials are needed to reduce the overall weight of the aircraft. In fact, the use of composite materials reduces the overall weight of the aircraft and at the same time ensures a very resistant structure.

[0012] Furthermore, the use of metallic elements and their mounting in contact with the structure creates galvanic coupling problems, with the attendant risk of corrosion of the metal and the need for increased inspection levels. This results in an increase in overall costs and structural weight for the manufacturers of the components and, consequently, for the airlines.

[0013] The need to reduce the overall weight of aircraft has therefore given rise to the use of composite materials in order to facilitate assembly operations resulting in a more integrated structure and to eliminate or minimize the corrosion problems affecting aerospace structures, thus increasing their resistance to fatigue.

[0014] Typically, the aforementioned structural components, such as fuselages or parts of fuselages, are manufactured by joining skins made of composite materials with reinforcing subcomponents, such as:

[0015] - a plurality of reinforcing longitudinal beams, also made of composite material and conveniently positioned parallel to the longitudinal extension of the fuselage;

[0016] A plurality of reinforcing ribs, also made of composite material and positioned transversely, in particular orthogonally, to the longitudinal extension of the fuselage.

[0017] In some configurations, the structural components include other strengthening subcomponents, such as, for example, stiffening panels (known as "sandwich cores"), which typically consist of a polymer foam or a structure with a honeycomb core inserted between two layers or sheets of composite material and are positioned between the ribs rather than between the stringers.

[0018] In the most common solution, the composite material used is made of uncured fiber material (eg carbon fibers), which is usually pre-impregnated with a fluid resin according to well-known processes.

[0019] Therefore, a composite material is a material consisting of two phases, a matrix and fibers. In particular, in the case of pre-impregnated materials, each layer of the material is usually composed of a matrix (made of thermosetting, thermoplastic resin, etc.) reinforced by fibers of different properties (such as carbon fibers, aramid fibers, glass fibers, etc.).

[0020] To produce the like, a plurality of layers of the pre-impregnated composite material are laminated to one another.

[0021] Similarly, to make stringers and ribs, multiple layers of pre-impregnated composite material are placed on a suitably shaped forming tool.

[0022] The skin and reinforcement subcomponents (ie stringers and ribs) must then be joined together to obtain the assembly.

[0023] The resulting assembly is then subjected to a curing process at high temperature and pressure to cure the composite material, compressing the aforementioned layers together and bonding the stringers and ribs to the skin ("co-curing").

[0024] According to some known methods, the reinforcement subcomponents (ie the stringers and the ribs) can also be pre-processed and subsequently joined to the skin by means of an adhesive film in order to obtain an assembly.

[0025] The assembly thus formed is subsequently subjected to a curing process at high temperature and pressure in order to solidify the composite material, compress the aforementioned layers together and bond ("co-bond") the stringers and ribs to the skin.

[0026] Thereby, the structural component is manufactured. The manufacturing of the structural component can be carried out in different ways.

[0027] The first mode, called "Inner Mould Line" or IML, involves the use of a curing tool, generally called a "spindle", which is externally shaped to define the inner surface of the structure to be built, such as a fuselage. The spindle basically has a generally cylindrical shape with corresponding longitudinal cavities, each of which is designed to accommodate a reinforcing stringer.

[0028] Once the stringers are positioned in the aforementioned cavities of the mandrel, different types of inserts, known in the industry as "bladders" and "noodles", are inserted into the various cavities formed after the stringers are positioned on the mandrel; these inserts are designed to hold the various components in place and prevent them from being crushed due to the high pressure during the subsequent curing step.

[0029] At this point, the assembly consisting of mandrels, stringers and inserts is covered with the relative skin that will constitute the outer surface of the aforementioned fuselage (or part of it).

[0030] In detail, a plurality of the aforementioned pre-impregnated fiber composite material layers are laminated onto the assembly consisting of mandrel, stringers and inserts.

[0031] According to known methods, this lamination, also called "layering", is carried out by an automated machine, generally called AFPM ("Automated Fibre Placement Machine"), which laminates each layer of the composite material constituting the skin onto a mandrel and thus onto the previously applied layer. Preferably, in a known manner, each layer is laminated with a different fiber orientation than the previous and the next layer.

[0032] After curing, a hollow “barrel” or “cylinder” is obtained, consisting of the skin and the stringers attached to the skin in the region of the inner surface of the latter.

[0033] Although the first IML mode described above is functionally effective, the applicant has observed some disadvantages:

[0034] - the use of mandrels is somewhat laborious and complicated, since it requires a long preparation time for the mandrels and requires the mandrels to be adapted and shaped for each application (since the arrangement of the stringers may vary depending on the type of structural component to be manufactured);

[0035] - Since the mandrel is usually axially symmetrical, the longitudinal beams and the inserts must be fixed in the lower cavity, which leads to an increase in the preparation time and the number of necessary parts and the overall cost;

[0036] - the ribs have to be fixed manually to the inside surface of the aforementioned "barrel" or to the finished "cylinder" after it has been manufactured, which can be very complicated, especially if the cylinder has a large axial extension;

[0037] - the fixing of the ribs to the finished cylinder generally requires the use of compensating elements or thicknesses (a technique known as “shimming”), since the skin will no longer be completely resistant, especially for structural parts of large dimensions;

[0038] - the removal (or "slipping-off") of the mandrel is somewhat complicated and laborious and limits the possible configuration of the geometry of the structural component, since this geometry needs to be very simple and without undercuts or excessive variations in cross section;

[0039] - Since a single spindle is used, this type of process requires that the machining cannot be performed in a parallel manner and this implies that the maximum productivity is not high.

[0040] This results in a considerable increase in production time and overall costs.

[0041] A second mode of production of structural components involves individually manufacturing a plurality of "panels" which are then joined together to form the structural component.

[0042] In detail, each panel is manufactured as follows:

[0043] - laminating a plurality of layers of uncured composite material on a forming tool to form a skin;

[0044] - placing a number of longitudinal beams on the previously laminated skin and fixing the cross beams to the skin, for example by means of an adhesive material;

[0045] The assembly thus formed is subjected to the aforementioned curing process in order to firmly connect the stringers to the skins.

[0046] The panels so obtained are then joined together.

[0047] The Applicant observes that this second production mode also has some disadvantages:

[0048] - the necessity to use joining elements between the various panels, such as titanium joints arranged longitudinally along the joining lines between the panels, which joints add considerably to the overall weight of the structural component;

[0049] - the connection line, i.e. the aforementioned joint, defines the critical load point of the structural component, i.e. the area supporting the operational loads;

[0050] The joining lines thus obtained thus define stress zones which are particularly subject to operational stresses and, therefore, potential weak points of the structural component. Summary of the invention

[0051] It is an object of the present invention to provide a method for manufacturing a structural component which is highly reliable and cost-limited and which solves at least some of the above-discussed disadvantages associated with known manufacturing methods.

[0052] According to the invention, this object is achieved by the method according to the invention.

[0053] Furthermore, it is an object of the present invention to provide a reinforced structural component made of composite material which is highly reliable and of limited cost and which solves at least some of the above-discussed drawbacks associated with known structural components.

[0054] According to the invention, this object is achieved by the structural component according to the invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0055] The invention will be best understood upon reading the following description of some preferred and non-limiting embodiments of the invention, which are discussed by way of example only with reference to the accompanying drawings, in which:

[0056] - Figures 1 to 6 is a schematic perspective view showing different and successive steps of a manufacturing method according to the invention, in particular a manufacturing method for manufacturing a subpanel of a structural component according to the invention, with some parts removed for greater clarity;

[0057] - Figures 7 to 10 The following are some other different and sequential steps of the manufacturing method according to the invention - in particular Figures 1 to 6 A schematic perspective view of a step after a step of - with some parts removed for greater clarity;

[0058] - Fig.11 is a schematic perspective view showing a sub-panel of a structural component according to an alternative embodiment of the present invention, with some parts removed for greater clarity;

[0059] - Fig.12 is a partial cross-sectional perspective view of a structural component according to an alternative embodiment of the present invention, with some parts removed for clarity. DETAILED DESCRIPTION

[0060] With reference to the drawings, and in particular with reference to Figure 8 , Fig. 9 , Fig.10 , numeral 1 designates as a whole a structural component made of a reinforced composite material, which has reinforcing subcomponents such as a reinforcing longitudinal beam 2 made of a composite material and a reinforcing rib 3 made of a composite material.

[0061] The structural component 1 comprises a skin 4 formed from a plurality of layers of composite material, to which the stringers 2 and the ribs 3 are fixed in the manner described below.

[0062] In particular, without any loss of generality for this reason, the present description will explicitly refer to a structural component 1 employed in the aerospace industry and defining, for example, a fuselage or a portion of a fuselage made of composite material.

[0063] More specifically, the component 1 described and illustrated herein defines a cylindrical modular portion of an aircraft fuselage.

[0064] Alternatively, as better explained below, the structural component 1 may be defined by a box made of composite material or by a portion thereof.

[0065] According to the preferred and non-limiting embodiment described and illustrated herein, component 1 has a generally cylindrical or cylinder-like shape about a central longitudinal axis A (eg, including various changes in radius and / or cross-section).

[0066] Preferably, the stringers 2 are arranged (or oriented) on the skin 4 defined by the axis A parallel to the longitudinal direction of the component 1 .

[0067] In detail, each stringer 2 is of known type and is preferably defined by a spar having a preferably omega-shaped cross section defining a central cavity delimited by the stringer 2 and the skin 4. In principle, each stringer 2 delimits a hollow portion having a closed profile by the skin 4.

[0068] Alternatively, each spar defining a respective longeron 2 may have a different cross-section, such as a rectangular or semicircular, T-shaped, double T-shaped, L-shaped cross-section.

[0069] Each longitudinal beam 2 has an extension in the longitudinal direction which is substantially greater than the extension in the other two directions orthogonal to this longitudinal direction.

[0070] According to a preferred alternative embodiment not shown herein, each longitudinal beam 2 is defined by a multilayer reinforced panel (called "sandwich core") comprising an internal structural core, which can be defined by a polymer foam (called "foam core") or by a honeycomb structure (called "honeycomb core") inserted between two layers or sheets or plies of composite material (of the type described below).

[0071] In the present description and the appended claims, the term "longitudinal beam" refers essentially to a wing spar of the above-mentioned type and a stiffening panel of the above-mentioned core sandwich type.

[0072] The ribs 3 are arranged (or oriented) on the skin 4 transversely, in particular orthogonally, to the longitudinal direction of the component 1 .

[0073] In detail, the rib 3 has a curved shape adapted to the curvature of the skin 4 and defines a circumferential reinforcement for the skin 4 .

[0074] The use of structural components made of composite materials is motivated by the need to reduce the overall weight of the structural component 1 .

[0075] In an embodiment, the composite material consists of a fiber material (eg, carbon fibers) that is not cured or pre-cured in the manner explained below.

[0076] In one embodiment, the material is pre-impregnated with a fluid resin according to a well-known process which is not described in detail.

[0077] In practice, each layer of the composite material generally consists of a prepreg having a thermosetting (resin) matrix reinforced by fibers of different nature (such as carbon fibers, aramid fibers, glass fibers, etc.).

[0078] The invention relates to a method for producing a structural component 1 .

[0079] In particular, the present description will explicitly refer to the manufacture of a hollow cylindrical (ie tubular) component 1 , without thereby losing generality.

[0080] However, the structural and functional features and steps of the method may be considered equally applicable to the manufacture of a structural component 1 having any shape, as long as it defines an internal volume delimited by a skin 4 .

[0081] Therefore, in this case (such as Figure 8 ), the skin 4 has a (substantially) cylindrical shape with a central longitudinal axis corresponding to the axis A. The stringers 2 are arranged with respective longitudinal axes preferably parallel to the axis A, and the ribs 3 are arranged with respective axes transverse to the axis A.

[0082] In an alternative embodiment not shown here, a small non-zero angle may be defined between the axis of the longitudinal beam 2 and the axis A, so that the longitudinal beam 2 is not parallel to the axis A.

[0083] With reference to the accompanying drawings, a process or method for manufacturing a structural component 1 will be described below.

[0084] The structural member 1 comprises a plurality of sub-panels 5 which are joined together (in a manner described below).

[0085] like Figure 6 As shown in , each sub-panel 5 includes: a sub-skin 6; and at least one stringer 2 and one rib 3, preferably a plurality of stringers 2 and ribs 3, such as two stringers 2 and two ribs 3, fixed to the inner surface 6a of the sub-skin 6.

[0086] The sub-skin 6 also has an outer surface 6b opposite to the inner surface 6a.

[0087] According to the present invention, reference Figure 1 The subskin 6 of each subpanel 5 is obtained by laminating a first layer 7 of uncured composite material on a forming tool 8 .

[0088] Thereby, an initially uncured subskin 6 made of fiber composite material is obtained.

[0089] Each of the first layers 7 is defined, for example, by a so-called “prepreg”.

[0090] The subskin 6 thus formed has an outer surface 6 b which is in contact with the forming tool 8 .

[0091] In practice, in order to form the sub-skin 6, the method according to the invention comprises the following steps: laminating a first layer 7 of uncured composite material on a forming tool 8, thereby forming the sub-skin 6. The sub-skin 6 has a first surface 6b in contact with the forming tool 8 and a second surface 6a opposite to the first surface 6b ( Figure 1 ).

[0092] Preferably, the first layer 7 is laminated with the fibers of the composite material parallel to a common direction, preferably the longitudinal direction of the structural component 1 .

[0093] The step of laminating the first layer 7 generally comprises:

[0094] - a layer 7 of laminated fibre composite material; and

[0095] The first layers 7 are arranged such that the fibers of all first layers are parallel to a common direction, preferably parallel to the longitudinal direction of the structural component 1 .

[0096] The Applicant has observed that this configuration provides optimal resistance of the subskin 6 to mechanical stresses while simplifying the formation of the subskin 6 .

[0097] Alternatively, each of the first layers 7 is laminated according to its own direction, which may be different from (ie, not parallel to) the direction in which the other first layers 7 are laminated.

[0098] It should be noted that the subskin 6 is a portion of the skin 4 and is defined by a plurality of layers of composite material that are less than the entirety of the layers that make up the skin 4 .

[0099] In particular, as better explained below, the skin 4 is composed of a first layer 7 and a second layer 10 of composite material. In other words, the aforementioned overall layer of the skin 4 only comprises the first layer 7 and the second layer 10 (the latter in Figure 8 (see ).

[0100] In order to complete the formation of the sub-panel 5, the method further comprises the following steps:

[0101] - placing at least one stringer 2 (and a spar or a reinforcing panel, as described above), in particular a plurality of stringers 2 (for example two stringers 2), made of pre-cured composite material, on the inner surface 6a of the sub-skin in a direction parallel to the longitudinal direction of the structural component 1, i.e. the axis A ( Figure 2 );

[0102] - placing at least one rib 3, in particular a plurality of ribs 3 (for example two ribs 3) made of pre-cured composite material on the inner surface 6a in a direction transverse (in particular orthogonal) to the longitudinal direction of the structural component 1 (i.e. the axis A) ( Figure 3 );as well as

[0103] - applying a preset temperature and pressure to the assembly defined by the subskin 6, the stringers 2 and the ribs 3 to cure the composite material and determine the rigidity of the stringers 2 and the ribs 3 fixed integrally to the inner surface 6a ( Figure 5 ).

[0104] Preferably, the stringers 2 and the ribs 3 are placed on the subskin 6 by interposition of an adhesive material which is known per se and is not described in detail.

[0105] The composite material constituting the stringers 2 and ribs 3 is preferably already cured (precured), and a curing step (application of temperature and pressure) is used to cure the material of the subskin 6 and to rigidly fix the stringers 2 and ribs 3 to the subskin 6 .

[0106] Curing is preferably carried out by inserting the assembly comprising the subskin 6, the stringer 2 and the rib 3 into an autoclave 50 which is heated in a pressurized state. Figure 5 is schematically shown in FIG.

[0107] Furthermore, the assembly is preferably enclosed in a vacuum bag 51 of known type and not described in detail, so as to compact the composite materials of the different parts to be joined.

[0108] In this case, the method further comprises the following steps:

[0109] - placing the assembly defined by the subskin 6 , the stringers 2 and the ribs 3 in a vacuum bag 51 ;

[0110] - Make the interior of the vacuum bag 51 evacuated.

[0111] The assembly wrapped in the vacuum bag 51 is then inserted into the autoclave 50 for curing.

[0112] Preferably, according to known methods that are not described in detail, different types of inserts (called "bladders" and "noodles" in the industry) are inserted into the various cavities formed after the positioning of the longerons 2 on the subskin 6; these inserts are designed to hold the longerons 2 in place and prevent them from being crushed due to the high pressures during the curing step.

[0113] In this way, a subpanel 5 made of a cured composite material is obtained.

[0114] This subpanel 5 is reinforced by stringers 2 and ribs 3 and at the same time has a very small thickness defined by the thickness of the subskin 6 comprising only the first layer 7 .

[0115] Thus, the subskin 6 is relatively thin because it has the minimum number of layers of composite material considered structurally indispensable. For example, if the total number of layers of the skin 4 (obtained by the sum of the number of first layers 7 and second layers 10) amounts to ten layers, for example, the first layers 7 are two and therefore the second layers 10 are eight.

[0116] Thus, advantageously, the step of laminating the first layer 7 comprises laminating at most two layers 7, preferably two layers 7, of uncured composite material.

[0117] Repeat the process described above to obtain Figure 6 A plurality of sub-panels 5 of the type shown in FIG. 5 are made of cured composite material.

[0118] In one embodiment, multiple forming tools 8 may be used to form multiple sub-panels 5 in parallel, thereby reducing the forming time.

[0119] In order to continue to produce the structural component 1 and refer to Figure 7 , the method according to the present invention further comprises the following steps:

[0120] - placing the previously obtained sub-panels 5 on the central support 11 so as to at least partially surround the central support 11 and so that the inner surface 6a of each sub-skin 6 faces the central support 11 and the outer surface 6b of each sub-skin faces the outside (e.g. Figure 7 ); and

[0121] - Pairs of laterally (mutually) adjacent sub-panels 5 are joined together in the region of the corresponding sub-skins 6 (in a manner better described below) to define a prefabricated component 12 having a continuous outer surface 12a defined by the group of outer surfaces 6b of the sub-skins 6 of the joined sub-panels 5.

[0122] More precisely, by the step of placing the sub-panel 5 on the central support 11 , an internal volume is defined which is delimited by the set of internal surfaces 6 a and which (at least partially) houses the central support 11 .

[0123] According to this preferred and non-limiting embodiment, the central support 11 is defined by a spoke system which supports the subpanel 5 in the region of the inner face 6 a of the relative subpanel 6 .

[0124] In the embodiment shown here, in which the structural component 1 is substantially cylindrical, the above-mentioned inner volume is cylindrical.

[0125] Alternatively, the internal volume may be of any shape, determined by the shape of the subpanel 5. In this case, it is sufficient to adapt the central support 11, ie the spoke system, to the different shapes of the subpanel 5 in order to support prefabricated parts 12 of any shape.

[0126] Preferably, the sub-panels 5 are placed on the central support 11 and then joined together so that the respective ribs 3 are aligned circumferentially around the axis A.

[0127] According to an important aspect of the present invention and with reference to Figure 8 The skin 4 is obtained by laminating a second layer 10 of uncured composite material on the outer face 12 a of the prefabricated part, so as to define a continuous outer upper skin 13 on the prefabricated part 12 and to define the skin 4 .

[0128] In detail, as specified above, the skin 4 is composed of a first layer 7 and a second layer 10 .

[0129] In more detail, the skin 4 includes a sub-skin 6 and an upper skin 13 , and in particular consists of the sub-skin 6 and the upper skin 13 .

[0130] In other words, the method according to the invention comprises the following steps:

[0131] Laminating a second layer 10 of uncured composite material onto the outer surface 12 a of the prefabricated part 12 to define a continuous outer upper skin 13 on the prefabricated part 12 and to define the skin 4 composed of the first layer 7 and the second layer 10 .

[0132] Advantageously, the lamination of the second layer 10 on the prefabricated part 12 is carried out by applying continuous fibers 14 of the composite material on the outer surface 12a and subsequently arranging the continuous fibers 14 in layers on the outer surface 12a.

[0133] In other words, the step of laminating the second layer 10 comprises:

[0134] - applying continuous fibers 14 of the composite material on the outer surface 12a; and

[0135] - The continuous fibers 14 are arranged in layers on the outer surface 12a.

[0136] In detail, the aforementioned layered arrangement is obtained by successively applying the continuous fibers 14 around and over the prefabricated component 12, the layered arrangement defining a plurality of complete coatings of the outer surface 12a, wherein each coating after the first covers the previous (lower) coating.

[0137] Each of these coatings defines one of the second layers 10 which as a whole define the outer skin 13 .

[0138] Due to the layered arrangement of the continuous fibers 14 as described above, the component 1 is able to withstand high stresses in operation.

[0139] Thanks to the method according to the invention, the prefabricated part 12 itself defines a tool or a kind of mandrel on which the second layer 10 is laid and the shaping of the skin 4 is completed.

[0140] As a result of the above, the structural component 1 comprises a continuous sub-wall 12a defined by an outer surface 12a, i.e. defined by the union of the outer surfaces 6b of the sub-skins 6 of the mutually connected sub-panels 5. The second layer 10 is laminated on the first layer 7 to define an outer skin 13 layered on the above-mentioned sub-wall 12a.

[0141] In order to complete the production of the structural component 1, the method further comprises the following steps:

[0142] - applying a preset temperature and pressure to the assembly consisting of the prefabricated part 12 and the continuous external upper skin 13 layered thereon, in order to cure the composite material and determine the rigidity and integral fixing of the second layer 10 to the first layer 7 .

[0143] Preferably, if Fig. 9 and Fig.10 As shown in FIG. 5 , this assembly is wrapped in a corresponding vacuum bag 52 , the interior of the vacuum bag 52 is evacuated, and then the vacuum bag 52 is inserted into the autoclave 50 .

[0144] By doing so, a reinforced structural component 1 made of a cured composite material is obtained.

[0145] according to Figure 1 By laminating the first layer 7 on the forming tool 8 , the lateral edges 15 , 16 of the subskin 6 are formed.

[0146] According to another preferred aspect of the present invention, the joining step includes: overlapping the lateral edge 16 of the first sub-panel 5 and the lateral edge 15 of the second sub-panel 5 adjacent to the first sub-panel 5, such as Figure 7 The enlarged figure is shown in detail.

[0147] More precisely, the first lateral edge 15 defines a receiving portion which, in the embodiment shown here, consists of a longitudinal recess which preferably extends along the entire longitudinal direction of the edge 15 (and of the subskin 6 ).

[0148] Similarly, the second lateral edge 16 defines a coupling portion extending along the entire longitudinal direction of the edge 16 (and of the subskin 16 ).

[0149] In order to join two adjacent sub-panels 5, the receiving portion 15 is joined by the coupling portion 16 ( Figure 7 ).

[0150] In detail, these sub-panels 5 are placed on the central support 11 so that the edge 16 engages the recess defined by the edge 15 , or more precisely so that the edge 16 overlaps the edge 15 (defining the recess).

[0151] In more detail, the edge 16 of the subpanel 5 overlaps the edge 15 of the adjacent subpanel, so that the edge 16 is radially larger on the outside compared to the edge 15 .

[0152] in other words:

[0153] - the step of laminating the first layer 7 comprises: forming a receiving portion in the region of the first lateral edge 15 and forming a coupling portion in the region of the second lateral edge 16; and

[0154] The joining step includes engaging the receiving portion 15 of the first sub-panel 5 and the coupling portion 16 of the second sub-panel 5 adjacent to the first sub-panel, and fixing the receiving portion 15 and the coupling portion 16 to each other.

[0155] In fact, the joining between the subpanels 5 is achieved by a simple male-female joint, without the need to provide complex joining elements that would reduce the weight of the structural component 1 and ensure the continuity of the lamination surface for laminating the outer skin 13 .

[0156] Due to this configuration, the joining between the subpanels 5 is greatly simplified compared to the known cases.

[0157] Preferably, an adhesive layer is interposed between the receiving portion 15 and the coupling portion 16 .

[0158] In this case, the joining step further includes fixing the receiving portion 15 and the coupling portion 16 to each other by interposing an adhesive layer (not shown) between the receiving portion 15 and the coupling portion 16 .

[0159] In this way, a secure bond between the sub-panels 5 is ensured.

[0160] Furthermore, the Applicant has observed that, thanks to this type of joining, the resistance to internal stresses (for example variations in the pressure outside component 1 relative to the pressure inside it) is significantly improved.

[0161] In an embodiment, the first layer 7 is laminated on the forming tool 8 by means of a laminating device, for example an automatic machine of the AFPM (“Automatic Fiber Placement Machine”) type, which is known per se and is not shown here.

[0162] Conveniently, during the lamination process, the forming tool 8 is moved towards the lamination device by an offset (not shown) equal to the thickness of the group of second layers 10 constituting the outer skin 13 .

[0163] In other words, the method further comprises the step of moving the forming tool 8 towards the laminating device by an offset equal to the thickness of the group of second layers 10 , ie equal to the thickness of the outer skin 13 .

[0164] In this way, the lamination process is simplified, since the user can avoid programming the lamination equipment to perform a different type of lamination than the classical lamination of the entire skin. Thus, the flexibility of the lamination process is increased.

[0165] Fig.11 and Fig.12 An alternative embodiment of a structural component 1 according to the invention is shown.

[0166] According to this embodiment, the structural component 1 is defined by a box, preferably a cylindrical box, having at least two internal volumes V1 , V2 separated by a partition 17 .

[0167] According to the invention, the partitions 17 are defined by at least some of the ribs 3 of the subpanels 5 joined together.

[0168] More precisely, if Fig.11 As shown in , to manufacture each sub-panel 5 of the part 1 defined by a box, a rib 3 provided with a protruding portion 3 a is placed on the surface 6 a of the sub-skin 6 .

[0169] The projection 3a is shaped so as to extend inwardly at least to the space occupied by the axis A once the component 1 is formed.

[0170] In this regard, the method further comprises the step of placing the ribs 3 of the different subpanels 5 in respective positions adjacent to each other by a step of placing the ribs 3 and a step of joining the subpanels to define a partition 17 of the structural element 1 formed by the group of ribs 3 .

[0171] In other words, each rib 3 defines a sector of the partition 17 .

[0172] according to Fig.12 The partition 17 separates the two volumes V1 and V2 of the box.

[0173] Conveniently, each rib 3 has a transverse extension so that it projects laterally from the associated sub-skin 6 (not shown). In this way, the ribs 3 partially overlap once the sub-panels 5 have been joined. Thus, the bulkhead 17 will be able to withstand high pressures.

[0174] Thanks to the configuration described above, it is possible to manufacture in a simple and economical manner a box of the “common bulked” type, ie a box having two separate internal volumes.

[0175] The features of the method for producing a structural component 1 made of composite material and of the structural component 1 according to the invention reveal significant advantages that can be obtained with them.

[0176] In particular, the use of a formed central mandrel (as in the IML manufacturing method described in the introduction to this specification) is no longer necessary, since the prefabricated part 12 formed by assembling and connecting the individual sub-panels 5 itself defines a tool, a "mandrel", on which the remaining layers of the skin 4, i.e. the second layer 10, are laminated.

[0177] This makes it possible to significantly reduce the length of time required for manufacturing the component 1 (since the insertion of the inserts and the longitudinal beams and their fixing, etc., do not require the preparation of a mandrel) as well as the overall dimensions and overall costs.

[0178] In addition, manufactured parts can meet stricter requirements in terms of profile and surface roughness.

[0179] Furthermore, the complicated and laborious insertion of the ribs in a previously formed finished "cylinder" or "barrel" is avoided.

[0180] Furthermore, since the ribs 3 are already fixed inside the prefabricated part 12, it is not necessary to perform the aforementioned shimming operation, since the ribs 3 are already fixed and cured on each sub-panel 5, and therefore no thickness compensation is required.

[0181] In fact, all the reinforcing structures, namely the stringers 2, the ribs 3, possible reinforcing inserts (bladders and needles) and possible adhesives, are already placed and fixed to the subskin 6 and therefore to the prefabricated part 12 during the final lamination of the second layer 10. In other words, both the stringers 2 and the ribs 3 are already cured and fixed to the subskin 6 (jointly bonded), which greatly simplifies the manufacture of the part 1.

[0182] Furthermore, the Applicant has observed that the structural component 1 thus obtained has an improved resistance to stresses: in fact, the second layer 10 , ie the outer skin 13 , bears the operating loads, whereas the subpanels 5 and the joints between them are subjected only to the loads exerted by the internal pressure.

[0183] In this regard, the fact that the second layer 10 is laminated within the continuous fibers 14 allows the manufacturer to further increase resistance to stresses.

[0184] Furthermore, the particular joining between the subpanels 5 by means of the engagement of the coupling portions 16 by means of the receiving portions 15 makes it possible to avoid the use of joining elements (eg titanium joints), which remain an integral part of the structural component. Component 1 will thus be significantly lightened.

[0185] Moreover, since the compression between the receiving part 15 and the connecting part 16 is greater the greater the thrust acting from the inside, and also due to the adhesive that may be interposed between them, this joint configuration allows better resistance to the internal stresses caused by variations in the external pressure of the component 1 relative to its internal pressure (such as in the case of an aircraft fuselage operating at high altitudes or in the case of a pressure tank).

[0186] In view of the above, it is evident that the production advantages of an “integrally” manufactured external component, which is an integrally manufactured load-bearing outer skin 13 , are fully exploited, without the need to insert any reinforcing elements or any subsequent joining elements.

[0187] The invention thus allows to bring about improvements both from a structural point of view and from the point of view of streamlining and simplifying the process of manufacturing component 1 .

[0188] The method and the component 1 disclosed and illustrated herein may be changed and modified without for this reason exceeding the scope of protection set forth in the appended claims.

Claims

1. A method for manufacturing a structural component (1) of composite material, the structural component comprising a skin (4) formed by a plurality of layers (7, 10) of composite material and reinforced with reinforcing stringers (2) and ribs (3) fixed to the skin, the method comprising the following steps: a) laminating a first layer (7) of uncured composite material onto a forming tool (8) to form a subskin (6) having a first surface (6b) in contact with the forming tool (8) and a second surface (6a) opposite the first surface (6b); b) arranging, on the second surface (6a) of the subskin (6), in a direction substantially parallel to the longitudinal direction of the structural component (1), at least one reinforced longitudinal beam (2) made of pre-cured composite material; c) arranging, on the second surface (6a) of the subskin (6), in a transverse direction transverse to the longitudinal direction of the structural component (1), at least one reinforcing rib (3) made of pre-cured composite material; d) applying a preset temperature and pressure to the assembly defined by the skin (6), the at least one stringer (2) and the at least one rib (3) to cure the composite material and determine the rigidity and integral fixation of the stringer (2) and the rib (3) to the second surface (2) of the skin (4), thereby obtaining a subpanel (5) made of cured composite material; e) repeating steps a) to d) to obtain a plurality of sub-panels (5); f) arranging the obtained sub-panels (5) on a central support (11) so as to at least partially surround said central support (11) and so that the second surface (6a) of each sub-skin (6) faces said central support (11) and the first surface (6b) of each sub-skin (6) faces outwards; g) joining together pairs of laterally adjacent subpanels (5) at respective subskins (6) to define a prefabricated component (12) having a continuous outer surface (12a) defined by a set of first surfaces (6b) of the subskins (6) of the joined subpanels (5); h) laminating a second layer (10) of uncured composite material onto the outer surface (12a) of the prefabricated part (12) to define a continuous outer upper skin (13) on the prefabricated part (12) and to define the skin (4), wherein the skin (4) is composed of the first layer (7) and the second layer (10); as well as i) applying a preset temperature and pressure to the assembly comprising the prefabricated part (12) and the continuous outer upper skin (13) to cure the composite material and determine the rigidity and integral fixing of the second layer (10) to the first layer (7).

2. The method of claim 1, wherein: The step (h) of laminating the second layer comprises: applying continuous fibers (14) of a composite material onto the outer surface (12a) of the prefabricated component (12); and The continuous fibers (14) are layered on the outer surface (12a) of the prefabricated component (12).

3. The method according to claim 1 or 2, wherein: Step a) of laminating the first layer comprises: laminating up to two layers (7) of uncured composite material, and preferably includes: The two layers of uncured composite material are laminated (7).

4. A method as claimed in any one of the preceding claims, wherein: Step a) of laminating the first layer comprises: laminating a layer of fiber composite material (7); The first layers (7) are arranged such that the fibers of all first layers are parallel to a common direction, preferably parallel to the longitudinal direction of the structural component (1).

5. A method as claimed in any one of the preceding claims, wherein: The step a) of laminating the first layer comprises: forming lateral edges (15, 16) on the sub-skin (6); And wherein the linking step (g) comprises: The lateral edge (16) of the first sub-panel (5) and the lateral edge (15) of the second sub-panel (5) adjacent to the first sub-panel (5) are overlapped.

6. The method of claim 5, wherein: Step a) of laminating the first layer comprises: forming a receiving portion (15) at a first lateral edge of the lateral edges (15, 16), and forming a coupling portion (16) at a second lateral edge of the lateral edges (15, 16); And wherein the linking step (g) comprises: engaging a receiving portion (15) of a first sub-panel (5) with a coupling portion (16) of a second sub-panel (5) adjacent to the first sub-panel (5); and The receiving portion (15) and the coupling portion (16) are fixed to each other.

7. The method of claim 6, wherein: The joining step (g) further includes fixing the receiving portion (15) and the coupling portion (16) to each other by interposing an adhesive layer between the receiving portion (15) and the coupling portion (16).

8. A method as claimed in any one of the preceding claims, wherein: Step d) of applying temperature and pressure comprises: placing the assembly defined by the subskin (6), the at least one stringer (2) and the at least one rib (3) in a vacuum bag (51), Making the interior of the vacuum bag (51) evacuated; and / or Wherein, step i) of applying temperature and pressure comprises: placing the assembly comprising the prefabricated part (12) and the continuous outer upper skin (13) in a vacuum bag (52), A vacuum is applied inside the vacuum bag (52).

9. The method according to any one of the preceding claims, further comprising the steps of: l) arranging the ribs (3) of different sub-panels (5) adjacent to each other in respective positions by means of an arrangement step f) and a connection step g) to define a partition (17) of the structural component (1) formed by the group of the ribs (3); Therein, the structural component (1) is defined by a box having at least two internal volumes (V1, V2) separated by the partition (17).

10. A method as claimed in any one of the preceding claims, wherein: performing a step of laminating a first layer (7) by a laminating device; And wherein, the method further comprises the following steps: m) moving the forming tool (8) towards the laminating device by an offset equal to the thickness of the group of the second layers (10) constituting the upper skin (13).

11. A method as claimed in any one of the preceding claims, wherein: Each longitudinal beam (2) is defined by: spar; or A multi-layer reinforced panel defined by a polymer foam or honeycomb structure including an internal structural core sandwiched between two layers or sheets of the composite material.

12. A structural component (1) made of a composite material, comprising: a skin (4) formed from a plurality of layers (7, 10) of composite material; a plurality of reinforced stringers (2) made of composite material, fixed to the skin (4) and oriented substantially parallel to the longitudinal direction of the structural component (1); and a plurality of reinforcing ribs (3) made of composite material, fixed to the skin (4) and oriented transversely to the longitudinal direction of the structural component (1); The structural component (1) has a plurality of sub-panels (5) which are connected to each other, and each of the plurality of sub-panels (5) comprises a sub-skin (6) defined by a first layer (7) of the plurality of layers of composite material, and at least one stringer (2) and at least one rib (3) fixed to an inner surface (6a) of the sub-skin (6); Wherein, each sub-skin (6) has an outer surface (6b) opposite to the inner surface (6a); wherein the structural component (1) comprises a continuous sub-wall (12a), the sub-wall (12a) being defined by a union of the outer surfaces (6b) of the sub-skins (6) of the sub-panels (5) connected to each other; wherein a second layer (10) of the plurality of layers of composite material is laminated to the first layer (7) to define a continuous outer upper skin (13) arranged in a divided manner on the sub-wall (12a); and The skin (4) is defined by a group of the sub-skin (6) and the upper skin (13), and the skin (4) is composed of the first layer (7) and the second layer (10).

13. The structural component of claim 12, wherein: The second layer (10) is defined by continuous fibers (14) of a composite material laminated and layered on the sub-wall (12a); and / or The first layer (7) comprises at most two layers, preferably two layers, of composite material; and / or Each of the first layers (7) comprises a layer of fiber composite material, the fibers of all first layers being parallel to a common direction, preferably parallel to the longitudinal direction of the structural component (1).

14. A structural component according to claim 12 or 13, wherein: Each longitudinal beam (2) is defined by: spar; or A multi-layer reinforced panel defined by a polymer foam or honeycomb structure including an internal structural core sandwiched between two layers or sheets of the composite material.

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