Method for manufacturing reinforced structural parts made of composite materials and structural parts

Abstract

To provide a method of manufacturing structural components that is reliable, cost-effective, and solves the drawbacks associated with known manufacturing methods.SOLUTION: A structural component (1) having a plurality of subpanels (5) joined to each other, each including a subskin defined by a first layer of composite material, at least one stringer, and at least one rib fixed to an inner surface of the subskin and continuous subwalls defined by the bonded outer surfaces of the subskins of the subpanels (5) connected each other, a second layer (10) of composite material is laminated over the first layer to define a continuous external overskin layered over the subwalls, and a skin (4) is defined by a set of the subskin and overskin, the first layer and a second layer (10).SELECTED DRAWING: Figure 8

Description

Technical Field

[0001] Cross - reference to related applications 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.

[0002] The present invention relates to a method for manufacturing reinforced structural parts made of composite materials, in particular reinforced stringers and ribs, and, if necessary, other composite reinforcement structures (such as polymer foams or honeycomb structures known as "honeycomb cores" sandwiched between two layers of composite materials) for reinforcing the structural parts.

[0003] More specifically, without loss of generality, this description explicitly refers to the manufacture of an aircraft fuselage or the manufacture of a tank under pressure.

[0004] Generally, the method according to the present invention can be applied to all these closed composite parts, i.e., parts that define an internal volume with a narrow internal space and undercuts.

[0005] The present invention also relates to a reinforced structural part made of composite material.

Background Art

[0006] Structural parts used in the aircraft industry, such as its fuselage and parts, are known and are made of composite materials.

[0007] In the state of the art, there are aircraft structural parts made of lightweight alloys and thus metallic materials, which are then used to construct part of the aircraft fuselage.

[0008] As is known, the fuselage is designed to ensure proper protection of the payload (crew, passengers, cargo, etc.), but at the same time, it cannot exceed the set weight limit.

[0009] Furthermore, the use of metallic components results in an overall cost increase, despite ensuring a greater resistance.

[0010] Therefore, in order to reduce the overall weight of the aircraft, structural components made of composite materials are required. In fact, the use of composite materials reduces the overall weight of the aircraft while ensuring a very resistant structure.

[0011] Furthermore, the use and installation of metallic components in contact with the structure pose galvanic coupling problems, resulting in a risk of metal corrosion and the need to increase the inspection level. This results in an overall cost increase and an increase in the weight of the structure for the manufacturers of said components and thus for the airline companies.

[0012] Therefore, the use of composite materials results from the need to reduce the overall weight of the aircraft, facilitate the assembly operations that generate a more integrated structure, eliminate or minimize the corrosion problems affecting the aircraft structure, and thus increase its resistance to fatigue.

[0013] Typically, the aforementioned structural components, such as the fuselage or its parts, are manufactured by joining a skin made of composite material with reinforcement sub-components such as the following. - A plurality of reinforcing stringers, also made of composite material and conveniently arranged parallel to the longitudinal extension direction of the fuselage - A plurality of reinforcing ribs, also made of composite material and arranged transversely, in particular orthogonally, with respect to the longitudinal extension direction of the fuselage.

[0014] In some configurations, the structural component includes other reinforcement sub-components such as a reinforcement panel (known as a "sandwich core"), and the reinforcement sub-component typically consists of a polymeric foam or structure having a honeycomb core sandwiched between two layers or sheets of composite material and is arranged between the ribs instead of the stringers.

[0015] In the most common solutions, the composite materials used are made of uncured fiber materials pre-impregnated with a fluid resin, such as carbon fibers, according to well-known processes.

[0016] Thus, the 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 usually consists of a matrix (made of a thermosetting resin, a thermoplastic resin, etc.) reinforced by fibers of different properties such as carbon fibers, aramid fibers, glass fibers.

[0017] To manufacture such a skin, multiple layers of the pre-impregnated composite material are laminated together.

[0018] Similarly, to manufacture stringers and ribs, multiple layers of the pre-impregnated composite material are placed on a properly formed molding tool.

[0019] Then, the skin and the reinforcing sub-components (i.e., stringers and ribs) must be joined together to obtain an assembly.

[0020] The assembly thus formed is then subjected to a curing process by applying high pressure and temperature, the composite material is cured, the aforementioned layers are compressed together, and the stringers and ribs are joined to the skin ("co-curing").

[0021] According to some known methods, the reinforcing sub-components (i.e., stringers and ribs) can also be pre-treated and then joined to the skin by an adhesive film to obtain an assembly.

[0022] The assembly thus formed is then subjected to a curing process by applying high pressure and temperature, the composite material is cured, the aforementioned layers are compressed together, and the stringers and ribs are joined to the skin ("co-bonding").

[0023] In this way, the structural component is manufactured. The manufacturing of the structural component can be carried out in various ways.

[0024] The first mode, known as the "inner mold line" or IML, involves the use of a curing tool, often referred to as a "spindle", which is formed externally to define the structure to be constructed, such as the inner surface of the airframe. The spindle basically has a generally cylindrical shape with respective longitudinal cavities, each designed to accommodate a reinforcing stringer.

[0025] Once the stringers are placed in the aforementioned cavities of the spindle, different types of inserts, known in the industry as "bladder" and "noodle", are inserted into the various cavities formed subsequent to the placement of the stringers on the spindle. These inserts are designed to hold various components in place and prevent them from being crushed due to high pressure during subsequent curing steps.

[0026] At this point, the assembly consisting of the spindle, stringers, and inserts is covered with a corresponding skin that forms the outer surface of the aforementioned airframe (or airframe part).

[0027] Specifically, a plurality of the aforementioned layers of prepreg fiber composite material are laminated onto the assembly consisting of the spindle, stringers, and inserts.

[0028] According to known methods, this lamination (also known as "layering") is usually performed by an automated machine, often referred to as an "AFPM" ("Automated Fiber Placement Machine"), which laminates each layer of the composite material that forms the skin on top of the spindle and thus on top of the previously applied layer. Preferably, each layer is laminated with a fiber orientation different from that of the previous and next layers by known methods.

[0029] 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 skin.

[0030] Although the first IML mode described above is functionally effective, the applicant has observed that it has several drawbacks as follows. - The use of a spindle is somewhat laborious and complex because the preparation of the spindle takes a long time and a spindle of a shape suitable for each application is required (the arrangement of the stringers can vary depending on the type of structural part to be manufactured). - Since the spindle is generally axisymmetric, it is necessary to fix the stringers and inserts arranged in the lower cavity, resulting in an increase in the preparation time, the number of parts required, and the total cost. - The ribs have to be manually fixed to the inner surface of the aforementioned "barrel" or to the completed "cylinder" after the cylinder has been manufactured, which can be very complex, especially when the cylinder has a large axial extension. - Since the skin is not completely resistant, especially to large structural parts, it is often necessary to use compensating elements or thicknesses (a technique known as "skim") for fixing the ribs to the completed cylinder. - The removal (or "slipping off") of the spindle is somewhat complex and cumbersome because this geometric shape needs to be very simple and there should be no undercuts or excessive changes in the cross-section, which limits the possible conformations of the geometric shape of the structural part. - In this type of process, it is required that machining cannot be performed in parallel by using a single spindle, which means a not very high maximum production speed.

[0031] This significantly increases the manufacturing time and the total cost.

[0032] The second manufacturing mode of the structural part involves manufacturing a plurality of "panels" separately, which are then joined together to form the structural part.

[0033] Specifically, each panel is manufactured as follows. - To form the skin, a plurality of layers of uncured composite material are laminated on a molding tool. - A predetermined number of stringers are placed on and fixed to the previously laminated skin, for example, by an adhesive material. - The assembly thus formed is subjected to the aforementioned curing process, and the stringer is firmly joined to the skin.

[0034] Then, the panels thus obtained are joined to each other.

[0035] The applicant has observed that this second manufacturing mode also has several drawbacks as follows. - It is necessary to use joining elements between the various panels, such as titanium joints arranged longitudinally along the joining lines between the panels, and said joints significantly increase the total weight of the structural component. - The joining line, i.e., the aforementioned joint, defines the critical load point of the structural component, i.e., the area that supports the operating load. - Therefore, the joining line thus obtained defines a stress area that is particularly subjected to operating stress, and thus defines a potential weakness of the structural component.

Summary of the Invention

Problems to be Solved by the Invention

[0036] An object of the present invention is to provide a method for manufacturing a structural component that is highly reliable, cost-effective, and solves at least some of the above-mentioned drawbacks of known manufacturing methods.

Means for Solving the Problems

[0037] According to the present invention, this object is achieved by the method according to claim 1.

[0038] Furthermore, an object of the present invention is to provide a reinforced structural component made of a composite material that is highly reliable, cost-effective, and solves at least some of the above-mentioned drawbacks of known structural components.

[0039] According to the present invention, this object is achieved by the structural component according to claim 11.

[0040] The present invention will be best understood by reading the following description of some preferred non-limiting embodiments thereof, described by way of example only with reference to the accompanying drawings.

Brief Description of the Drawings

[0041]

Figure 1

Figure 2

Figure 3

Figure 4

Figure 5

Figure 6

Figure 7

Figure 8

Figure 9

Figure 10

Figure 11

Figure 12

DETAILED DESCRIPTION OF THE INVENTION

[0042] Referring to the accompanying drawings, particularly FIGS. 8, 9, and 10, reference numeral 1 generally designates a structural part made of a reinforced composite material having reinforcing sub-parts such as a reinforcing stringer 2 made of a composite material and a reinforcing rib 3 made of a composite material.

[0043] The structural part 1 comprises a skin 4 formed by a plurality of layers of a composite material to which the stringer 2 and the rib 3 are fixed, as will be described below.

[0044] In particular, this description is explicitly related, without loss of generality for this reason, to the structural part 1 used in the aerospace industry and defining, for example, an airframe made of a composite material or a part thereof.

[0045] More specifically, the part 1 described and illustrated herein defines a cylindrical module portion of an aircraft airframe.

[0046] Alternatively, the structural part 1 can be defined by a tank made of a composite material or a part thereof, as will be described in detail below.

[0047] According to the preferred, non-limiting embodiment described and illustrated herein, the part 1 has a generally cylindrical or cylindrical shape (e.g., including various variations in radius and / or cross section) about a central longitudinal axis A.

[0048] Preferably, the stringers 2 are arranged (or oriented) on the skin 4 parallel to the longitudinal direction of the part 1 defined by the axis A.

[0049] In detail, each stringer 2 is of known type and is preferably defined by a spar having preferably an omega-shaped cross section, and preferably defines a central cavity bounded by the stringer 2 and the skin 4. Essentially, each stringer 2 defines, by the skin 4, a hollow portion having a closed contour.

[0050] Alternatively, each wing spar defining a respective stringer 2 may have a different cross-section, for example a rectangular or semicircular, T-shaped, double T-shaped, L-shaped cross-section.

[0051] Each stringer 2 has a longitudinal extension substantially greater than its extension in the other two directions perpendicular to said longitudinal direction.

[0052] According to preferred alternative embodiments not shown here, each stringer 2 is defined by a multi-layered reinforced panel with an internal structural core (known as a "sandwich core"), which may be defined by a polymer foam (known as a "foam core") or by a honeycomb structure (known as a "honeycomb core") sandwiched between two layers or sheets or plies of composite material (of the type described below).

[0053] The term "stringer" in this description and in the appended claims essentially denotes both the aforementioned type of wing spars and the aforementioned core sandwich type of reinforced panels.

[0054] The rib 3 is arranged (or oriented) on the skin 4 in a direction transverse to, in particular perpendicular to, the longitudinal direction of the component 1.

[0055] Specifically, the rib 3 has a curved shape so as to conform to the curvature of the skin 4 and defines a circumferential reinforcement for the skin 4.

[0056] The use of a structural component made of a composite material results from the need to reduce the total weight of the structural component 1.

[0057] In an embodiment, the composite material consists of a fiber material, such as carbon fiber, that is not cured or pre-cured in the manner described below.

[0058] In one embodiment, the material is pre-impregnated with a fluid resin according to a well-known process not described in detail.

[0059] In practice, each layer of the composite material usually consists of a prepreg having a thermosetting (resin) matrix reinforced by fibers of different properties, such as carbon fiber, aramid fiber, glass fiber, etc.

[0060] The present invention relates to a method for manufacturing the structural component 1.

[0061] In particular, without loss of generality, this description explicitly refers to the manufacture of a hollow cylindrical (i.e., tubular) component 1.

[0062] However, it can be considered that the structural and functional features and steps of the method are equally applicable to the manufacture of a structural component 1 having any shape as long as it defines an internal volume delimited by the skin 4.

[0063] Thus, in this case (as shown in FIG. 8), the skin 4 has a (substantially) cylindrical shape with a central longitudinal axis corresponding to the axis A. The stringers 2 are arranged such that their respective longitudinal axes are preferably parallel to the axis A, and the ribs 3 are arranged such that their respective axes cross the axis A.

[0064] In alternative embodiments not shown in this specification, a small non-zero angle can be defined between the axis of the stringer 2 and the axis A, and thus the stringer 2 can be prevented from being parallel to the axis A.

[0065] With reference to the accompanying drawings, a process or method for manufacturing the structural component 1 will be described below.

[0066] The structural component 1 comprises a plurality of sub-panels 5 joined to each other (as will be described later).

[0067] As shown in FIG. 6, 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, for example two stringers 2 and two ribs 3, fixed to the inner surface 6a of the sub-skin 6.

[0068] The sub-skin 6 further has an outer surface 6b on the opposite side of the inner surface 6a.

[0069] According to the present invention, referring to FIG. 1, the sub-skin 6 of each sub-panel 5 is obtained by laminating a first layer 7 of uncured composite material on a molding tool 8.

[0070] Thereby, a sub-skin 6 made of a fiber composite material that is initially uncured is obtained.

[0071] Each of the first layers 7 is defined, for example, by a so-called "prepreg".

[0072] The sub-skin 6 formed in this way has an outer surface 6b that contacts the molding tool 8.

[0073] In practice, to form the sub-skin 6, the method according to the present invention includes the step of laminating a first layer 7 of uncured composite material on a molding tool 8, and thus a sub-skin 6 having a first surface 6b that contacts the molding tool 8 and a second surface 6a on the opposite side of the first surface 6b is formed (FIG. 1).

[0074] Preferably, the first layer 7 is laminated with fibers of a composite material parallel to the common direction of the structural part 1, preferably the longitudinal direction.

[0075] The step of laminating the first layer 7 basically - a step of laminating the fiber composite layer 7, and - a step of arranging the first layer 7 such that the fibers of all the first layers are parallel to the common direction of the structural part 1, preferably the longitudinal direction, is included.

[0076] The applicant has observed that this configuration provides optimal resistance of the sub-skin 6 to mechanical stress and at the same time simplifies the formation of the sub-skin 6.

[0077] Alternatively, each of the first layers 7 may be laminated according to its own direction, which may be different from (i.e., not parallel to) the lamination direction of the other first layers 7.

[0078] It should be pointed out that the sub-skin 6 is part of the skin 4 and is defined by a plurality of layers of composite material smaller than the entire layer constituting the skin 4.

[0079] In particular, the skin 4 consists of a first layer 7 and a second layer 10 of composite material, as will be described in more detail below. In other words, the aforementioned entire layer of the skin 4 includes only the first layer 7 and the second layer 10 (the latter is visible in FIG. 8).

[0080] To complete the formation of the sub-panel 5, the method also - arranging at least one stringer 2 (and thus, as described above, a spar or a stiffening panel), in particular a plurality of stringers 2 made of a pre-cured composite material, for example two stringers 2, on the inner surface 6a of the sub-skin along the longitudinal direction of the structural part 1, i.e., in a direction parallel to the axis A (FIG. 2); - Placing at least one rib 3, in particular a plurality of ribs 3, for example two ribs 3, made of a pre-hardened composite material, on the inner surface 6a along the longitudinal direction of the structural part 1, i.e. in a direction transverse to, in particular orthogonal to, the axis A (Figure 3); - Hardening the composite material and applying a preset temperature and pressure to the assembly defined by the sub-skin 6, the stringer 2 and the rib 3 to determine a rigid and integral fixation to the inner surface 6a of the stringer 2 and the rib 3 (Figure 5); It includes.

[0081] Preferably, the stringer 2 and the rib 3 are placed on the sub-skin 6 by sandwiching an adhesive material known per se and not described in detail.

[0082] The composite material constituting the stringer 2 and the rib 3 is preferably already hardened (pre-hardened), and the hardening step (the step of applying temperature and pressure) is used to harden the material of the sub-skin 6 and firmly fix the stringer 2 and the rib 3 thereto.

[0083] Hardening is preferably carried out by inserting the assembly including the sub-skin 6, the stringer 2, and the rib 3 into the autoclave 50 schematically shown in Figure 5.

[0084] Furthermore, the assembly is preferably wrapped in a vacuum bag 51 of a known type, not described in detail, to facilitate the compression of the composite materials of the different parts to be joined.

[0085] In this case, the method further includes - Placing the assembly defined by the sub-skin 6, the stringer 2, and the rib 3 into the vacuum bag 51; - Applying a vacuum inside the vacuum bag 51. It includes.

[0086] Then, the assembly wrapped in the vacuum bag 51 is inserted into the autoclave 50 for hardening.

[0087] Preferably, according to known processes not described in detail, different types of inserts known in the industry as "bladder" and "noodle" are inserted into various cavities formed following the positioning of the stringer 2 on the sub-skin 6. These inserts are designed to hold the stringer 2 in place and prevent the stringer 2 from being crushed due to the high pressure applied during the curing step.

[0088] Thus, a sub-panel 5 made of cured composite material is obtained.

[0089] This sub-panel 5 is reinforced by the stringer 2 and the rib 3 and at the same time has a very thin thickness defined by the thickness of the sub-skin 6 which only contains the first layer 7.

[0090] Therefore, the sub-skin 6 is relatively thin because the number of layers of the composite material considered structurally essential is minimal. For example, if the total number of layers of the skin 4 (the sum of the first layer 7 and the second layer 10) is 10 layers, the first layer 7 is, for example, 2 layers, and thus the second layer 10 is, for example, 8 layers.

[0091] Therefore, advantageously, the step of laminating the first layer 7 includes laminating a maximum of two layers 7, preferably two layers 7, of uncured composite material.

[0092] The above procedure is repeated to obtain a plurality of sub-panels 5 made of cured composite material of the type shown in FIG. 6.

[0093] In one embodiment, a plurality of forming tools 8 can be used to form a plurality of sub-panels 5 in parallel, thereby shortening the forming time.

[0094] To proceed with the manufacture of the structural component 1, referring to FIG. 7, the method according to the invention also - Place the previously obtained sub-panels 5 on the central support 11 so as to at least partially surround the central support 11, and such 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 outward (as shown in FIG. 7). - In the region of each sub-skin 6, join pairs of laterally (mutually) adjacent sub-panels 5 to each other (in a manner described in more detail below) so as to define a pre-configured component 12 having a continuous outer surface 12a defined by the set of outer surfaces 6b of the joined sub-panels 5. including.

[0095] More precisely, by placing the sub-panels 5 on the central support 11, an internal volume is defined that is delimited by the set of inner surfaces 6a and that (at least partially) accommodates the central support 11.

[0096] According to this preferred non-limiting embodiment, the central support 11 is defined by a kind of spoke system that supports the sub-panels 5 within the region of the inner surface 6a of the corresponding sub-panels 5.

[0097] In the embodiments shown herein, the structural component 1 is substantially of a generally cylindrical shape, and the aforementioned internal volume is cylindrical.

[0098] Alternatively, the internal volume can take on any shape defined by the shape of the sub-panels 5. In this case, it is sufficient to adapt the central support 11, i.e., the spoke system, to the different shapes of the sub-panels 5 in order to support the pre-configured component 12 having any shape.

[0099] Preferably, the sub-panels 5 are placed on the central support 11 and then joined to each other, whereby each rib 3 is circumferentially aligned about the axis A.

[0100] According to an important aspect of the present invention, referring to FIG. 8, the skin 4 is obtained by laminating a second layer 10 of uncured composite material onto the outer surface 12a of the preconfigured part to define a continuous external overskin 13 on the preconfigured part 12 and by defining the skin 4.

[0101] Specifically, the skin 4 consists of a first layer 7 and a second layer 10 as described above.

[0102] More specifically, the skin 4 includes and in particular consists of a subskin 6 and an overskin 13.

[0103] In other words, the method according to the present invention - includes the step of laminating a second layer 10 of uncured composite material onto the outer surface 12a of the preconfigured part 12 so as to define a continuous external overskin 13 on the preconfigured part 12 and so as to define the skin 4 consisting of the first layer 7 and the second layer 10.

[0104] Advantageously, the lamination of the second layer 10 onto the preconfigured part 12 is performed by applying continuous fibers 14 of the composite material onto the outer surface 12a and then laminating the continuous fibers 14 onto the outer surface 12a.

[0105] In other words, the step of laminating the second layer 10 - includes the step of applying continuous fibers 14 of the composite material onto the outer surface 12a and - the step of laminating the continuous fibers 14 onto the outer surface 12a. and includes.

[0106] Specifically, the aforementioned lamination is obtained by continuously applying continuous fibers 14 around and on the preconfigured part 12, thereby defining a plurality of complete coatings of the outer surface 12a, and each coating following the first coating covers the previous (lower) coating.

[0107] Each of these coatings defines one of the second layers 10 that as a whole defines the overskin 13.

[0108] Thanks to the layering of the continuous fibers 14 as described above, the component 1 can withstand the high stresses during operation.

[0109] Thanks to the method according to the invention, the pre-configured component 12 itself defines a tool or a kind of spindle for laying up the second layer 10 thereon to complete the formation of the skin 4.

[0110] As described above, the structural component 1 comprises a continuous sub-wall 12a defined by the outer surface 12a, that is, by the joining of the outer surfaces 6b of the sub-skins 6 of the sub-panels 5 joined to each other. The second layer 10 is laminated on the first layer 7 so as to define an over-skin 13 laminated on the aforementioned sub-wall 12a.

[0111] To complete the manufacture of the structural component 1, the method further - includes the step of applying a preset temperature and pressure to an assembly comprising the pre-configured component 12 and the continuous external over-skin 13 laminated thereon to cure the composite material and determine a firm and integral fixation of the second layer 10 to the first layer 7.

[0112] Preferably, as shown in FIGS. 9 and 10, this assembly is wrapped in a corresponding vacuum bag 52 to which a vacuum is applied inside and then inserted into an autoclave 50.

[0113] By doing so, a reinforced structural component 1 made of a cured composite material is obtained.

[0114] According to FIG. 1, by laminating the first layer 7 on the molding tool 8, the side edges 15, 16 of the sub-skin 6 are formed.

[0115] According to a further preferred embodiment of the invention, the joining step includes overlapping the side edge 16 of the first sub-panel 5 and the side edge 15 of the second sub-panel 5 adjacent to the first sub-panel 5, as shown in detail in the enlarged view of FIG. 7.

[0116] More precisely, in the embodiment shown herein, the first side edge portion 15 defines a receiving portion that preferably consists of a longitudinal recess that extends preferably along the entire longitudinal extension of the edge portion 15 and thus of the sub-skin 6.

[0117] Similarly, the second side edge portion 16 defines a connecting portion that extends longitudinally along the edge portion 16 and thus along the entire longitudinal extension of the sub-skin 6.

[0118] To join two adjacent sub-panels 5, the receiving portion 15 is engaged by the connecting portion 16 (Figure 7).

[0119] Specifically, these sub-panels 5 are arranged on the central support 11 such that the edge portion 16 engages the recess defined by the edge portion 15, or rather, the edge portion 16 overlaps the edge portion 15 (defining the recess).

[0120] More specifically, the edge portion 16 of the sub-panel 5 overlaps the edge portion 15 of the adjacent sub-panel, and thus the edge portion 16 is radially outside the edge portion 15.

[0121] In other words, - The step of laminating the first layer 7 includes the step of forming a receiving portion in the region of the first side edge portion 15 and the step of forming a connecting portion in the region of the second side edge portion 16, - The step of joining includes engaging the receiving portion 15 of the first sub-panel 5 with the connecting portion 16 of the second sub-panel 5 adjacent to the first sub-panel and fixing the receiving portion 15 and the connecting portion 16 to each other.

[0122] In practice, the joining between the sub-panels 5 is performed only by a kind of male-female connection without the need to provide complex joining elements that can ensure the continuity of the lamination surface that weighs down the structural part 1 and results in the lamination of the over-skin 13.

[0123] Thanks to this configuration, the joining between the sub-panels 5 is significantly simplified compared to known cases.

[0124] It is preferable that an adhesive layer is sandwiched between the receiving portion 15 and the coupling portion 16.

[0125] In this case, in the joining step, further, a step of sandwiching an adhesive layer (not shown) between the receiving portion 15 and the coupling portion 16 and fixing the two to each other is included.

[0126] Thereby, a strong joining between the sub-panels 5 is ensured.

[0127] Furthermore, the applicant has observed that thanks to this type of joining, the resistance to internal stresses, for example the change in the pressure outside the component 1 with respect to the pressure inside the component 1, has been significantly improved.

[0128] In one embodiment, the lamination of the first layer 7 on the molding tool 8 is performed by a laminating device, for example, an automatic machine of the "AFPM" ("Automated Fiber Placement Machine") type, which is known per se and not shown herein.

[0129] Advantageously, during lamination, the molding tool 8 is moved towards the laminating device by an offset (not shown) equal to the thickness of the set of second layers 10 that form the overskin 13.

[0130] In other words, the method further includes a step of moving the molding tool 8 towards the laminating device by an offset equal to the thickness of the set of second layers 10, that is, equal to the thickness of the overskin 13.

[0131] In this way, the user can avoid programming the laminating device to perform various types of lamination different from the conventional lamination of the entire skin, so that the lamination process is simplified. Therefore, the flexibility of the lamination process is improved.

[0132] Figures 11 and 12 show alternative embodiments of the structural component 1 according to the present invention.

[0133] According to this embodiment, the structural part 1 is defined by a tank, preferably a cylindrical tank, having at least two internal volumes V1, V2 separated by a partition wall 17.

[0134] According to the invention, the partition wall 17 is defined by at least some of the ribs 3 of the sub-panels 5 joined to each other.

[0135] More precisely, as shown in FIG. 11, in order to manufacture each sub-panel 5 of the part 1 defined by the tank, ribs 3 provided with protruding portions 3a are arranged on the surface 6a of the sub-skin 6.

[0136] The protruding portion 3a has a shape that extends inwardly at least to the space occupied by the axis A when the part 1 is formed.

[0137] In this regard, the method further includes arranging the ribs 3 and joining the sub-panels, and arranging the ribs 3 of different sub-panels 5 at respective adjacent positions to each other, and defining the partition wall 17 of the structural part 1 formed by the set of ribs 3.

[0138] In other words, each rib 3 defines a sector of the partition wall 17.

[0139] According to FIG. 12, the partition wall 17 separates the two volumes V1, V2 of the tank.

[0140] Advantageously, each rib 3 has a lateral extension so as to protrude laterally from the corresponding sub-skin 6 (not shown). In this way, when the sub-panels 5 are joined, the ribs 3 partially overlap. Therefore, the partition wall 17 can withstand high pressure.

[0141] Thanks to the above configuration, it is possible to simply and economically manufacture a "common bulky" type of tank, that is, a tank having two separate internal volumes.

[0142] The manufacturing method of the structural part 1 made of the composite material according to the present invention and the characteristics of the structural part 1 clarify the obvious advantages that can be obtained by using them.

[0143] In particular, the pre-formed part 12 itself formed by assembling and joining various sub-panels 5 defines a tool, i.e., a kind of "spindle", for laminating the remaining layers of the skin 4, i.e., the second layer 10. Therefore, the use of a formed central spindle (such as the IML manufacturing method described at the beginning of the description) is no longer necessary.

[0144] This can not only significantly reduce the time required to manufacture the part 1 (because it is not necessary to prepare the spindle involving the insertion and fixing of inserts and stringers, etc.), but also significantly reduce the overall dimensions and total cost.

[0145] Furthermore, the manufactured part can comply with more stringent requirements regarding the profile and surface roughness.

[0146] Furthermore, the complicated and cumbersome insertion of ribs into a pre-formed finished "cylinder" or "barrel" is avoided.

[0147] Furthermore, since the rib 3 is already fixed inside the pre-formed part 12, it is not necessary to perform the aforementioned simming operation. Since the rib 3 is already fixed and cured on each sub-panel 5, therefore, it does not require thickness compensation.

[0148] In fact, all the reinforcement structures (i.e., the stringer 2, the rib 3, possible reinforcing inserts (bladder and needle), and possible adhesive layers) are already arranged and fixed on the sub-skin 6, and thus on the pre-formed part 12 at the time of the final lamination of the second layer 10. In other words, both the stringer 2 and the rib 3 are already cured and fixed to the sub-skin 6 (co-bonding), which greatly simplifies the manufacturing of the part 1.

[0149] In addition, the Applicant observed that the structural part 1 thus obtained has improved resistance to stress. In fact, the second layer 10, i.e., the overskin 13, supports the operating load, while the sub-panels 5 and the joints between them support only the load exerted by the internal pressure.

[0150] In this regard, the fact that the second layer 10 is laminated within the continuous fibers 14 enables the manufacturer to further enhance the resistance to stress.

[0151] Furthermore, thanks to the particular joining of the sub-panels 5 by the engagement of the receiving part 15 by the joining part 16, the use of joining elements (e.g., titanium joints) that remain as an integral part of the structural part can be avoided. Thus, the part 1 is clearly lightweighted.

[0152] Furthermore, this joining configuration enables better resistance to internal stress resulting from the change in the external pressure of the part 1 with respect to the internal pressure of the part 1 (as in the case of an aircraft fuselage operating at high altitude or a pressure tank), since the greater the thrust acting from the inside, the greater the compression between the receiving part 15 and the joining part 16, which is also increased by the adhesive layer that can be sandwiched between them.

[0153] In light of the above, it is clear that the manufacturing advantages of manufacturing an external part "as one piece" (the overskin 13 that supports the load manufactured as one piece) can be fully utilized without the need to insert any reinforcing elements or any subsequent joining elements.

[0154] Therefore, the present invention can bring improvements from both a structural point of view and a point of view of rationalizing and simplifying the process for manufacturing the part 1.

[0155] The methods and parts 1 disclosed and illustrated in this specification can be subject to changes and modifications without exceeding the scope of protection described in the appended claims for this reason.

[0156] [Additional Clause 1] A method for manufacturing a structural element (1) in a composite material formed by a plurality of layers (7, 10) of composite material and reinforced with reinforcing stringers (2), comprising a skin (4) and ribs (3) fixed to said skin (4). a) Laying a first layer (7) of uncured composite material on a forming tool (8), thereby forming a sub-skin (6) having a first surface (6b) in contact with the forming tool (8) and a second surface (6a) opposite to said first surface (6b). b) Placing at least one reinforcing stringer (2) made of pre-cured composite material on said second surface (6a) of said sub-skin (6) along a direction substantially parallel to the longitudinal direction of the structural component (1). c) Placing at least one reinforcing rib (3) made of pre-cured composite material on said second surface (6a) of said sub-skin (6) along a transverse direction which is transverse to said longitudinal direction of said structural component (1). d) Applying a preset temperature and pressure to the assembly defined by said sub-skin (6), at least one of said stringers (2), and at least one of said ribs (3) to cure said composite material and determine a rigid and integral fixation of said stringers (2) and said ribs (3) to said second surface (6a) of said sub-skin (6), thus obtaining a sub-panel (5) made of cured composite material. e) Repeating steps a) to d) to obtain a plurality of said sub-panels (5). f) Placing the obtained sub-panels (5) on a central support (11) such that the central support (11) is at least partially surrounded, the second surface (6a) of each sub-skin (6) faces the central support (11), and the first surface (6b) of each sub-skin (6) faces outward. g) Joining pairs of the laterally adjacent sub-panels (5) to each other with their respective sub-skins (6) so as to define a pre-configured part (12) having a continuous outer surface (12a) defined by a set of the first surfaces (6b) of the sub-skins (6) of the joined sub-panels (5); h) Laying a second layer (10) of uncured composite material on the outer surface (12a) of the pre-configured part (12) so as to define a continuous external over-skin (13) on the pre-configured part (12) and so as to define the skin (4), the skin (4) consisting of the first layer (7) and the second layer (10); i) Curing the composite material and applying a preset temperature and pressure to the assembly including the pre-configured part (12) and the continuous external over-skin (13) to determine a rigid and integral fixation of the second layer (10) to the first layer (7); A method comprising the steps above. [Appendix 2] The step h) of laying the second layer includes - applying continuous fibers (14) of a composite material to the outer surface (12a) of the pre-configured part (12); - layering the continuous fibers (14) on the outer surface (12a) of the pre-configured part (12). The method according to Appendix 1, including the steps above. [Appendix 3] The step a) of laying the first layer includes laying up to two layers (7) of uncured composite material, preferably including laying two layers (7) of uncured composite material. The method according to Appendix 1 or 2. [Appendix 4] The step a) of laying the first layer includes - laying a layer (7) of a fiber composite material; - arranging the first layer (7) such that the fibers of all the first layers are parallel to a common direction of the structural part (1), preferably the longitudinal direction. The method according to any one of Appendices 1 to 3, including the steps above. [Additional item 5] The step a) of laminating the first layer includes a step of forming side edges (15, 16) on the sub-skin (6), The step g) of joining includes a step of overlapping the side edge (16) of the first sub-panel (5) with the side edge (15) of the second sub-panel (5) adjacent to the first sub-panel (5), according to the method of any one of claims 1 to 4. [Additional item 6] The step a) of laminating the first layer includes a step of forming a receiving portion (15) at a first side edge of the side edges (15, 16) and a step of forming a connecting portion (16) at a second side edge of the side edges (15, 16), The step g) of joining is - a step of engaging the receiving portion (15) of the first sub-panel (5) with the connecting portion (16) of the second sub-panel (5) adjacent to the first sub-panel (5); - a step of fixing the receiving portion (15) and the connecting portion (16) to each other, including the method according to claim 5. [Additional item 7] The step g) of joining further includes a step of fixing the receiving portion (15) and the connecting portion (16) to each other by sandwiching an adhesive layer therebetween, according to the method according to claim 6. [Additional item 8] The step d) of applying temperature and pressure is - a step of placing the assembly defined by the sub-skin (6), at least one of the stringers (2), and at least one of the ribs (3) in a vacuum bag (51); - a step of applying a vacuum inside the vacuum bag (51); including and / or The step i) of applying temperature and pressure is - a step of placing the assembly including the pre-configured component (12) and the continuous external over-skin (13) in the vacuum bag (52); - applying a vacuum inside the vacuum bag (52); The method according to any one of claims 1 to 7, comprising: [Claim 9] l) further comprising arranging the ribs (3) of the different sub-panels (5) in respective adjacent positions by the steps f) of arranging and g) of joining so as to define a partition wall (17) of the structural part (1) formed by the set of the ribs (3); The method according to any one of claims 1 to 8, wherein the structural part (1) is defined by a tank having at least two internal volumes (V1, V2) separated by the partition wall (17). [Claim 10] The step a) of laminating the first layer (7) is performed by a laminating device, The method further comprises m) moving the forming tool (8) towards the laminating device by an offset equal to the thickness of the set of the second layers (10) constituting the overskin (13). The method according to any one of claims 1 to 9. [Claim 11] Each of the stringers (2) is - a spar, or - a multi-layer reinforced panel including an internal structural core defined by a polymer foam or honeycomb structure sandwiched between two layers or sheets of the composite material The method according to any one of claims 1 to 10, defined by: [Claim 12] A structural part (1) made of a composite material, - a skin (4) formed by a plurality of layers (7, 10) of the composite material, and - a plurality of reinforcing stringers (2) made of the composite material, fixed to the skin (4) and oriented substantially parallel to the longitudinal direction of the structural part (1). - A plurality of reinforcing ribs (3) made of a composite material, fixed to the skin (4) and oriented in a transverse direction with respect to the longitudinal direction of the structural component (1); comprising the structural component (1) having a plurality of sub - panels (5) joined to each other, each having a sub - skin (6) defined by a first layer (7) of a plurality of said composite material layers, and at least one stringer (2) and at least one rib (3) fixed to the inner surface (6a) of the sub - skin (6); each sub - skin (6) having an outer surface (6b) opposite to the inner surface (6a); the structural component (1) comprising a continuous sub - wall (12a) defined by the joint of the outer surfaces (6b) of the sub - skins (6) of the sub - panels (5) joined to each other; a second layer (10) of the plurality of layers of the composite material being laminated on the first layer (7) so as to define a continuous external over - skin (13) laminated on the sub - wall (12a); the skin (4) being defined by a set of the sub - skin (6) and the over - skin (13), and the structural component (1) consisting of the first layer (7) and the second layer (10). [Appended Claim 13] - the second layer (10) being defined by continuous fibers (14) of a laminated composite material laminated on the sub - wall (12a), and / or - the first layer (7) comprising at most two, preferably two, composite material layers, and / or - each of the first layers (7) comprising a layer of a fiber composite material, and the fibers of all the first layers being parallel to a common direction of the structural component (1), preferably the longitudinal direction. The structural component according to appended claim 12. [Appended Claim 14] Each of the stringers (2) - a spar, or - a multi - layer reinforced panel comprising an internal structural core defined by a polymer foam or honeycomb structure sandwiched between two layers or sheets of the composite material. The structural component according to appended claim 12 or 13, defined by

Explanation of reference signs

[0157] 1 Structural component, 2 Stringer, 3 Rib, 3a Protrusion, 4 Skin, 5 Sub-panel, 6 Sub-skin, 6a Inner surface, second surface, 6b Outer surface, first surface, 7 First layer, 8 Molding tool, 10 Second layer, 11 Central support, 12 Pre-configured component, 12a Outer surface, sub-wall, 13 Over-skin, 14 Continuous fiber, 15 First side edge portion, receiving portion, 16 Second side edge portion, coupling portion, 17 Partition wall, 50 Autoclave, 51 Vacuum bag, 52 Vacuum bag

Claims

1. A method for manufacturing a structural element (1) in composite material, comprising a skin (4) formed by several layers (7, 10) of composite material and reinforced with reinforcing stringers (2), and ribs (3) fixed to said skin (4), comprising: a) laminating a first layer (7) of uncured composite material onto a moulding tool (8) thereby forming a subskin (6) having a first surface (6b) in contact with the moulding tool (8) and a second surface (6a) opposite said first surface (6b); b) arranging on said second surface (6a) of said subskin (6) at least one reinforcing stringer (2) made of a pre-cured composite material along a direction substantially parallel to the longitudinal direction of the structural component (1); c) arranging on said second surface (6a) of said subskin (6) at least one reinforcing rib (3) made of pre-cured composite material along a transverse direction transverse to said longitudinal direction of said structural component (1); d) subjecting the assembly defined by said subskin (6), at least one said stringer (2) and at least one said rib (3) to a predetermined temperature and pressure to harden said composite material and determine a rigid and integral fixation of said stringer (2) and said rib (3) to said second surface (6a) of said subskin (6), thus obtaining a subpanel (5) made of hardened composite material; e) repeating steps a) to d) to obtain a plurality of said sub-panels (5); f) placing the resulting sub-panels (5) on the central support (11) so as to at least partially surround the central support (11) and so that the second surface (6a) of each sub-skin (6) faces towards the central support (11) and the first surface (6b) of each of the sub-skins (6) faces outward; g) joining pairs of laterally adjacent sub-panels (5) together with their respective sub-skins (6) to define a pre-constructed part (12) having a continuous outer surface (12a) defined by a set of first surfaces (6b) of the sub-skins (6) of the joined sub-panels (5); h) laminating a second layer (10) of uncured composite material on said outer surface (12a) of said pre-component (12) so as to define a continuous outer overskin (13) on said pre-component (12) and so as to define said skin (4), said skin (4) consisting of said first layer (7) and said second layer (10); i) subjecting said assembly including said pre-component (12) and said continuous outer overskin (13) to a preset temperature and pressure in order to cure said composite material and determine a rigid and integral fixation of said second layer (10) to said first layer (7); The method includes:

2. The step h) of depositing the second layer comprises: - applying continuous fibres (14) of a composite material to said external surface (12a) of said pre-component (12); - layering said continuous fibres (14) onto said outer surface (12a) of said pre-component (12); The method of claim 1 , comprising:

3. 2. The method according to claim 1, wherein the step a) of laminating the first layer comprises laminating at most two layers (7) of uncured composite material, preferably comprising laminating two layers (7) of uncured composite material.

4. The step a) of depositing the first layer comprises: - laying up a layer (7) of fibre composite material; - arranging said first layers (7) so that the fibres of all said first layers are parallel to a common direction, preferably the longitudinal direction, of said structural part (1); The method of claim 1 , comprising:

5. said step a) of laminating said first layer comprises forming side edges (15, 16) on said subskin (6); 2. The method of claim 1, wherein the joining step g) comprises overlapping a side edge (16) of a first of the sub-panel (5) with a side edge (15) of a second of the sub-panel (5) adjacent to the first sub-panel (5).

6. said step a) of laminating said first layer comprises the steps of forming a receiving portion (15) at a first one of said side edges (15, 16) and forming a bonding portion (16) at a second one of said side edges (15, 16); The joining step g) further comprises: - engaging the receiving portion (15) of the first sub-panel (5) with the joining portion (16) of the second sub-panel (5) adjacent to the first sub-panel (5); - fixing said receiving part (15) and said connecting part (16) to each other; The method of claim 5 , comprising:

7. 7. The method of claim 6, wherein said joining step g) further comprises the step of fixing said receiving portion (15) and said mating portion (16) to each other by sandwiching an adhesive layer between them.

8. said step d) of applying temperature and pressure - placing said assembly defined by said subskin (6), at least one said stringer (2) and at least one said rib (3) in a vacuum bag (51); - applying a vacuum inside said vacuum bag (51); Including, and / or The step i) of applying temperature and pressure comprises: - placing said assembly including said pre-component (12) and said continuous outer overskin (13) in said vacuum bag (52); - applying a vacuum to the interior of said vacuum bag (52); The method of claim 1 , comprising:

9. l) positioning the ribs (3) of different sub-panels (5) in respective positions adjacent to one another by positioning step f) and joining step g) so as to define a bulkhead (17) of the structural component (1) formed by said set of ribs (3); 2. The method according to claim 1, wherein the structural component (1) is defined by a tank having at least two internal volumes (V1, V2) separated by the partition (17).

10. The step a) of laminating the first layer (7) is carried out by a lamination device, The method further comprises: m) moving said forming tool (8) towards said lamination device by an offset equal to the thickness of said second set of layers (10) constituting said overskin (13), The method of claim 1.

11. Each of the stringers (2) - Wing spars, or - a multi-layer reinforced panel comprising an internal structural core defined by a polymer foam or honeycomb structure sandwiched between two layers or sheets of said composite material; The method of claim 1 , wherein the distance is defined by:

12. A structural part (1) made of composite material, a skin (4) formed by several layers (7, 10) of composite material; - a number of reinforcing stringers (2) made of composite material fixed to said skin (4) and oriented substantially parallel to the longitudinal direction of said structural component (1); a number of reinforcing ribs (3) made of composite material fixed to said skin (4) and oriented transversely to the longitudinal direction of said structural part (1); Including, The structural component (1) comprises a plurality of sub-panels (5) joined together, each comprising a subskin (6) defined by a first layer (7) of a plurality of the composite plies, at least one stringer (2) and at least one rib (3) fixed to an inner surface (6a) of the subskin (6), Each subskin (6) has an outer surface (6b) opposite to the inner surface (6a), the structural component (1) comprises a continuous sub-wall (12a) defined by the joints of the outer faces (6b) of the sub-skins (6) of the sub-panels (5) joined together, a second ply (10) of the plurality of plies of composite material is laminated onto the first ply (7) to define a continuous outer overskin (13) layered on the sub-wall (12a); A structural component (1), wherein the skin (4) is defined by a set of the subskin (6) and the overskin (13), and comprises the first layer (7) and the second layer (10).

13. - said second layer (10) is defined by continuous fibres (14) of a composite material laminated and stacked on said sub-walls (12a); and / or - said first layer (7) comprises at most two, preferably two, composite layers; and / or - each of said first layers (7) comprises a layer of fibre composite material, said fibres of all said first layers being parallel to a common direction, preferably to the longitudinal direction, of said structural part (1); The structural component of claim 12.

14. Each of the stringers (2) - Wing spars, or - a multi-layer reinforced panel comprising an internal structural core defined by a polymer foam or honeycomb structure sandwiched between two layers or sheets of said composite material; The structural component of claim 12 defined by:

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