Method for manufacturing a monolithic structure made of composite materials for the wings or tail of an aircraft.

Abstract

The invention relates to a monolithic structure (1) made of composite material, manufactured starting from a fiber-reinforced prepreg material. The structure comprises two walls (6, 7) facing each other and at least one interconnecting element (8) extending transversely between the walls (6, 7) and connected to them, with which each elongated cavity (9) is defined. The walls (6, 7) extend symmetrically on both sides of a direction (B). The interconnecting element (8) is a rib (10) extending transversely to said direction (B), at least one wall (6, 7) having a sandwich structure and comprising two panels (11, 12) facing each other and at least one spar element (13) extending transversely between the panels (11, 12) and connected to them, with which each elongated cavity (14) is defined, extending transversely to the rib (10).

Description

Technical Field

[0001] Cross - reference to related applications This patent application claims priority based on Italian Patent Application No. 102021000000044 filed on January 4, 2021. The entire disclosure of this Italian patent application is incorporated herein by reference.

[0002] Technical Field The present invention relates to a method for manufacturing a monolithic structure made of composite materials for the wings or empennages of aircraft, particularly airplanes. In the following description, although airplanes are specifically referred to, generality is not lost.

Background Art

[0003] As is well known, wings or empennages are structural components having fixed surfaces, which are part of an aircraft and incorporate internally a structure made of composite materials having cavities suitable for weight reduction.

[0004] Specifically, the term "wing" means a component of an aircraft or airplane, and the surface of this wing is arranged according to a specific attitude with respect to the fluid flow flowing over its surface, and is capable of generating a series of hydrodynamic actions (as a result of forces and moments) caused by complex physical actions related to local differences in velocity, pressure, and viscosity actions acting on the surface itself.

[0005] A tail fin is a part of an airplane or aircraft that has a stabilizing function and comprises one or more horizontal surfaces and one or more vertical surfaces. The horizontal surfaces often consist of a fixed part, which is a stabilizer, and a movable part, which is a balancer, hinged to the rear of the stabilizer. In some examples, the horizontal surface is formed by a single surface that is a stabilizer (a shortened form of stabilizer and balancer). Furthermore, the vertical surface consists of two surfaces, one of which is a fixed tail vertical stabilizer and the other is a movable rudder, although in some examples of high-performance aircraft (military and aerobatic), a fully movable vertical tail fin exists.

[0006] Furthermore, the wings and tail sections of an aircraft often have a tapered pattern from the base, i.e., the area connected to the fuselage, to the free end.

[0007] In other words, by identifying the direction connecting the central zone of the base to the central zone of the free end as the longitudinal extension direction of the wing or tail, and by identifying the virtual segment extending in the transverse direction relative to the aforementioned longitudinal direction and joining the wind leading edge of the wing or tail itself to the wind trailing edge as the chord, it becomes possible to understand how the length of the aforementioned chord decreases as it moves away from the fuselage toward the free end of the wing or tail.

[0008] As clearly stated above, both known types of wings and tails have internal structures made of composite materials, and these structures are The first wall section, A lower wall portion facing the first wall portion, the lower wall portion being positioned at a non-zero distance from the first wall portion, A plurality of interconnecting elements extending between a first wall and a second wall, and defining a boundary between these walls, wherein each of the walls extends through a cavity, and It is mainly composed of the following.

[0009] The use of composite materials makes it possible to reduce the overall weight of an aircraft while simultaneously achieving a very strong structure.

[0010] The first and second walls of known types of structures are nearly parallel and, more precisely, slightly converge toward each other towards the free end of the wing or tail.

[0011] These interconnecting elements are defined by spar members, which extend transversely with respect to the chord of the wing or tail wing, and join their bases to the free ends of the wing or tail wing itself in each direction.

[0012] More precisely, spar members are generally composed of partitions that are substantially perpendicular to the wall of the structure, or more generally, in the transverse direction.

[0013] The aforementioned structures are manufactured using known methods that include the following steps, starting with a prepreg containing a thermosetting matrix (resin) reinforced with various natural fibers such as carbon fibers, aramid fibers, and glass fibers.

[0014] First, in the prepreg state, each spar member is preformed on a suitable preforming tool to form two elongated profile elements having a C-shaped cross-section and consisting of two flanges on the back that project perpendicularly from both ends of the back edge. To form the spar member, these two C-shaped profile elements are joined together along their respective backs so that they form recesses on opposite sides.

[0015] Specifically, each C-shaped profile element is laminated onto three flat surfaces of a preform molding tool that has a substantially parallelepiped profile. The first surface of this tool is completely covered with the prepreg, while the other two surfaces, which are perpendicular to and adjacent to the first surface and parallel to each other, are covered only over a certain length.

[0016] After the necessary vacuum consolidation and at a predetermined temperature, the C-shaped profile elements are joined in pairs in the manner described above to form the desired number of spar members. These spar members are held in place and at predetermined distances from each other by their respective rigid support tools having an elongated pin shape for a subsequent curing step.

[0017] Preform molding is a process that applies vacuum and temperature, allowing prepreg materials to acquire a solidified shape without causing a polymerization process.

[0018] Each support tool consists of a rigid, solid, elongated body that is substantially parallelepiped, and has a cross-section that corresponds to the shape of the elongated cavity of the structure to be defined and manufactured by a flat surface.

[0019] Before positioning each spar member, which is still in the form of a preformed prepreg, between two support tools, these support tools are used in the following series of steps: For example, by applying a separation agent layer in the form of a film to each support tool, the step of facilitating the subsequent removal of the support tool itself from each elongated cavity, The steps include attaching a tubular bag to each of the support tools prepared in this manner, leaving excess portions of the tubular bag at each end of the support tool itself for the subsequent sealing process, The steps include wrapping a breathable cloth around each support tool and the outside of the tubular bag, and securing the flap of this breathable cloth with sealant, The tubular separation film is attached to each of the support tools prepared in this manner, and in this example as well, the excess portion of the tubular separation film is left at each end of the support tool itself for the subsequent sealing process, The steps include sealing the ends of the tubular bag and the ends of the tubular separation film with sealant, The steps include applying a vacuum and waiting for the entire dressing to shrink on each support tool by the tubular separation film, and undergoes a dressing operation consisting of

[0020] At this point, the preformed spar members, each consisting of two C-shaped profile elements joined together along their respective backs, are positioned between the support tools that have previously undergone the above-described dressing operation. Specifically, each preformed spar the backs of each pair of C-shaped profile elements constituting the spar are disposed between two opposing flat surfaces of two mutually adjacent support tools, the wings of each profile element are placed on respective flat surfaces of each support tool that are mutually parallel and substantially orthogonal to the flat surface supporting the back of the profile element itself so as to be arranged.

[0021] The thus-formed assembly, consisting of the preformed spar members and the previously dressed support tools, is inserted into a mold. This mold consists of a lower plate, an upper plate, and two opposing side wall portions connecting the lower plate and the upper plate.

[0022] Specifically, one or more layers of prepreg material are laminated on each of the upper plate and the lower plate, and these material layers are for defining a first skin and a second skin. These skins will constitute the first wall portion and the second wall portion of the structure to be manufactured after the curing step.

[0023] More precisely, the preformed spar members held in position by each support tool are positioned on the lower plate of the mold that carries the first skin. Thereafter, the upper plate of the mold that carries the second skin is closed on the side wall portions of the mold itself and on the assembly consisting of the preformed spar members and the support tools.

[0024] At this point, the separation film, the ventilation cloth, and the bag film are continuously arranged over the entire molding die. The bag film is sealed by a sealant at the bottom of the molding die.

[0025] The tubular separation film disposed on each support tool is opened at its end. A portion that becomes an excess in length with respect to each support tool is cut off.

[0026] The tubular bags on each support tool are opened and sealed at their ends.

[0027] The so-called envelope bag is formed in such a manner by sealing the edge of the outer bag film disposed on the molding die to the end of the tubular bag of the support tool and further by sealing the ends of adjacent tubular bags to each other.

[0028] At this point, a vacuum is applied inside the envelope bag until the material of the outer bag contracts toward the outer surface of the molding die.

[0029] Next, the ends of the tubular bags are opened, and the tubular bags themselves are detached from their respective support tools by continuing the application of the vacuum, and the volume surrounded inside the envelope bag tends to be minimized.

[0030] At this point, the support tools are taken out, and the assembly thus formed is processed inside an autoclave so as to perform a curing operation at a given pressure value and temperature value (for example, in the case of an epoxy resin, the curing temperature is about 180 °C, and the curing pressure is between 6 and 7 bar).

[0031] By the method and tools described above, it becomes possible to accurately position the preforms and maintain their positions during the closing operation of the mold for the formation and manufacture of the envelope bag.

[0032] By removing the support tool before the solidification cycle, improper deformation of the tool under the pressure and temperature conditions required for solidification is prevented, thereby ensuring uniform pressure application to all parts made of composite material.

[0033] Alternatively, the tubular bags allow for the application of uniform pressure to the composite material parts that the tubular bags come into contact with.

[0034] However, the applicant has noticed that there is room for improvement in the aforementioned structure and the method for realizing that structure.

[0035] Specifically, in the known types of multi-spar structures analyzed in detail so far, external loads are distributed between the two skins defining the first and second walls, as well as between the various spar members. Wings or tails having this type of structure perform perfectly for their end use, while exhibiting technically superior performance with respect to bending compared to torsion.

[0036] Therefore, there is an industry need to manufacture airfoils and tailplanes that exhibit the same behavior as known airfoils and tailplanes under bending loads, while showing a better response to torsional loads during operation.

[0037] Furthermore, there is an industry need to simplify the manufacturing process for wing or tail structures to make it more efficient and cost-effective. Finally, the quality of surface finishes, specifically the quality of the interior surface finishes of the structures, requires further improvement.

[0038] U.S. Patent Application Publication 2017 / 0174313A1 describes a method for manufacturing a composite structure for an aircraft wing. This method uses two skins of uncured composite material wrapped in a ring around an assembly of support tools positioned side-by-side and parallel to each other. The uncured composite material, which will define each rib of the wing, is positioned between each pair of support tools. The method also provides the use of a preformed reinforcing structure, which is inserted at specific positions between the two skins and has a sufficient degree of crosslinking to avoid the use of support elements during the final curing step. Therefore, the components of the preformed reinforcing structure are manufactured at a different time than the skins and ribs, making this method lengthy and complex. Adhesives or other fastening systems must be used to join the preformed and at least partially crosslinked reinforcing structure to the uncured skins. Thus, the realized structure cannot be considered monolithic, i.e., it is not realized in a single step using prepreg materials that are equivalent in each part and have the same conditions as before the curing step. [Prior art documents] [Patent Documents]

[0039] [Patent Document 1] U.S. Patent Application Publication No. 2017 / 0174313A1 [Overview of the project] [Problems that the invention aims to solve]

[0040] The object of the present invention is to provide a method for manufacturing a monolithic structure made of composite materials for an aircraft wing or tail section that is highly reliable, has limited cost, and can satisfy at least one of the requirements specifically described above and the requirements related to a method for manufacturing a monolithic structure made of known types of composite materials. [Means for solving the problem]

[0041] According to the present invention, this objective is achieved by a method for manufacturing a monolithic structure made of composite material for an aircraft wing or tail section, as claimed in claim 1.

[0042] Hereinafter, in order to better understand the present invention, a preferred, non-limiting embodiment will be described entirely by example and with the aid of the accompanying drawings. [Brief explanation of the drawing]

[0043] [Figure 1] This is a plan view of an aircraft incorporating a monolithic structure made of composite material manufactured by the method of the present invention into its wing section, with several parts removed for clarity. [Figure 2] This is an enlarged perspective view of the monolithic structure in Figure 1, with several parts removed for clarity. [Figure 3] Figure 2 is an enlarged partial cross-sectional perspective view of the details of the monolithic structure. [Figure 4] Figure 3 is a further enlarged partial cross-sectional perspective view of the details. [Figure 5] Figure 3 is a further enlarged partial cross-sectional perspective view of the details. [Figure 6] This is a partial cross-sectional perspective view of the first support tool during a series of steps in the process of forming and laminating each wall section of the monolithic structure of Figure 2 using prepreg composite material. [Figure 7] This is a partial cross-sectional perspective view of the first support tool during a series of steps in the process of forming and laminating each wall section of the monolithic structure of Figure 2 using prepreg composite material. [Figure 8] This is a partial cross-sectional perspective view of the first support tool during a series of steps in the process of forming and laminating each wall section of the monolithic structure of Figure 2 using prepreg composite material. [Figure 9]This is a partial cross-sectional perspective view of the first support tool during a series of steps in the process of forming and laminating each wall section of the monolithic structure of Figure 2 using prepreg composite material. [Figure 10] This is a partial cross-sectional perspective view of the second support tool during a series of steps in the process of forming and laminating each interconnecting element of the monolithic structure of Figure 2 using prepreg composite material. [Figure 11] This is a partial cross-sectional perspective view of the second support tool during a series of steps in the process of forming and laminating each interconnecting element of the monolithic structure of Figure 2 using prepreg composite material. [Figure 12] This is a partial cross-sectional perspective view of the second support tool during a series of steps in the process of forming and laminating each interconnecting element of the monolithic structure of Figure 2 using prepreg composite material. [Figure 13] Figures 6 to 12 show partial cross-sectional perspective views of the monolithic structure of Figure 2 in a series of manufacturing steps using the first and second support tools. [Figure 14] Figures 6 to 12 show partial cross-sectional perspective views of the monolithic structure of Figure 2 in a series of manufacturing steps using the first and second support tools. [Figure 15] Figures 6 to 12 show partial cross-sectional perspective views of the monolithic structure of Figure 2 in a series of manufacturing steps using the first and second support tools. [Figure 16] This figure, similar to Figure 5, shows a detail of one possible modification of a monolithic box-shaped structure made of composite materials, in a partial cross-sectional perspective view. [Figure 17] This is a perspective view of two partially sectioned second support tools during the process of forming and laminating interconnecting elements of a monolithic structure using a modified prepreg composite material as shown in Figure 16. [Modes for carrying out the invention]

[0044] Referring to Figures 1 and 2, 1 shows as a whole a monolithic structure made of composite materials for an aircraft, i.e., for a vehicle capable of flying at least over short distances.

[0045] In the example shown in Figure 1, the aircraft is shown as Airplane 2, which comprises a fuselage 3, two wing sections 4, and multiple tail fins 5 in a known configuration. Among these, the tail vertical stabilizer and two stabilizers are partially visible.

[0046] Aircraft 2 has an axis A that is the same as the axis of the fuselage 3 and the forward direction of aircraft 2 itself. The wing sections 4 and tail section 5 originate from the fuselage 3 and extend in a cantilever manner.

[0047] More specifically, each wing section 4 has a base section 4a connected to the fuselage 3 and a free end section 4b located on the opposite side of the base section 4a.

[0048] Therefore, each wing section 4 extends along a longitudinal direction B, which is identified as a joint connecting the central zone of the base section 4a to the central zone of the free end section 4b. The longitudinal extension direction B of each wing section 4 is located in the transverse direction with respect to the axis A of the aircraft 2, and in the example where it is inclined, it forms an obtuse angle with the axis A itself toward the forward portion of the fuselage 3. Alternatively, direction B may also form an acute or right angle with respect to the axis A.

[0049] Each wing section 4 is defined in the transverse axis direction with respect to direction B by the wind leading edge 4c and the wind trailing edge 4d. In each wing section 4, the leading edge 4c is located directly in front of the trailing edge 4d with respect to the forward direction of movement of the aircraft 2.

[0050] In the illustrated example, each wing section 4 originates from the base portion 4a and tapers toward the free end portion 4b, forming a tapered pattern along direction B. In practice, if we identify a virtual segment extending in the transverse axis direction relative to direction B and joining the trailing edge 4d to the leading edge 4c of each wing section 4 as the chord, then in the illustrated example, it can be pointed out that the length of the aforementioned chord decreases as it moves away from the fuselage 3 toward the free end portion 4b of the wing section 4 itself.

[0051] As can be seen in the example shown in Figure 1, structure 1 constitutes the internal part of each wing section 4.

[0052] It should be noted that structure 1 can also be used for each tail fin 5 of aircraft 2, or for the wings or tail fins of the aircraft in general.

[0053] Referring to Figures 1 to 5, Structure 1 is, The first wall portion 6, A second wall portion 7, which is positioned along the maximum surface of the second wall portion 7 itself, facing the maximum surface of the wall portion 6, and spaced apart from the wall portion 6 by a non-zero amount, Multiple interconnecting elements 8, which extend in the horizontal direction between wall portion 6 and wall portion 7, are connected to wall portion 6 and wall portion 7, and together with wall portion 6 and wall portion 7 themselves define each elongated cavity portion 9, and It is equipped with.

[0054] In the illustrated example, wall 6 forms the lower wall of the structure 1 at the position where each wing 4 is used, and wall 7 forms the upper wall of the structure 1 itself.

[0055] Wall sections 6 and 7 extend symmetrically on both sides of the longitudinal direction B and are substantially parallel to direction B itself.

[0056] The interconnecting element 8 is composed of ribs 10 that extend in the lateral direction with respect to the longitudinal direction B.

[0057] Each wall section 6, 7 has a "sandwich" structure, and this sandwich structure is A first panel 11, which, along the maximum surface of the first panel 11 itself, faces the maximum surface of the other wall portions 7, 6, A second panel 12, which, along the maximum surface of the second panel 12 itself, faces the maximum surface of panel 11, Multiple spar members 13, which extend in the transverse direction between panel 11 and panel 12, are connected to panels 11 and 12, define each elongated cavity 14 together with panels 11 and 12 themselves, and extend in the transverse direction with respect to the rib 10, and It consists of.

[0058] In practice, the panels 11 of wall sections 6 and 7 face each other, and are positioned between the panels 12 in the horizontal direction relative to the panels 11 and 12 themselves, or are positioned further inward.

[0059] Panel 11, and similarly panel 12, extend symmetrically on both sides of the longitudinal direction B and are approximately parallel to direction B itself.

[0060] According to possible alternative examples not shown, only one of wall 6 and wall 7 may have a sandwich structure, while the other is formed by a single panel.

[0061] In the illustrated example, the distance between panels 11 and 12 of each wall section 6 and 7 is shorter than the distance between panels 11 themselves in the horizontal direction relative to panels 11 and 12.

[0062] In an alternative example where only one of wall sections 6 and 7 has the aforementioned sandwich structure and the other is formed by a single panel, the distance between panels 11 and 12 of wall sections 6 and 7 having the sandwich structure is shorter than the distance between panels 11 and wall sections 6 and 7 formed by a single panel.

[0063] Panels 11 and 12, and more generally walls 6 and 7, converge slightly toward each other toward the free end portion 4b of the wing 4 in the illustrated example. According to one possible alternative example not shown, panels 11 and 12, and more generally walls 6 and 7, may be parallel to each other.

[0064] In either case, panels 11 and 12, and more generally, wall sections 6 and 7, can be considered to be substantially parallel to one another.

[0065] The rib 10 and the spar member 13 can have an open cross-section or a closed cross-section.

[0066] In the solutions shown in Figures 2, 3, and 5, the rib 10 has a C-shaped cross-section. According to other possible alternatives not shown, the rib 10 may also have a cross-section corresponding to an I-shape, T-shape, double T-shape, double C-shape, Z-shape, or other shape.

[0067] In the illustrated example, the spar members 13 of each wall section 6 and 7 have oblique I-shaped cross-sections that converge in pairs toward the outermost panel 12. In this example as well, the cross-sections of the spar members 13 can individually correspond to, for example, a vertical I, a C, a Z, a T, a double T, etc.

[0068] Structure 1 is manufactured starting from a prepreg having a polymer matrix, such as one made from a thermosetting resin, which is reinforced with fibers that may have various properties, such as carbon fibers and / or aramid fibers and / or glass fibers.

[0069] Alternatively, structure 1 may be manufactured starting from a matrix prepreg made from a thermoplastic resin reinforced with the type of fibers described above.

[0070] In both of these examples, panels 11 and 12 are manufactured starting from skins 11a and 12a (Figures 8, 9, 13, 14, and 15). Each skin is formed by one or more layers of the aforementioned fiber-reinforced prepregs laminated on each flat surface, as will be described in more detail later.

[0071] Referring to Figures 6 to 9, the spar member 13 is realized by laminating the prepreg material defined above onto each elongated support tool 15 in the longitudinal direction of the spar member 13 itself.

[0072] Specifically, the support tool 15 is designed to hold the spar member 13 in a predetermined position within the structure 1 to be manufactured, and has a reinforcing material and polymer-based composition suitable for enabling transitions from a rigid state to a flexible elastomer state and vice versa in response to heating and cooling, respectively, i.e., in response to thermal stimuli.

[0073] Advantageously, the polymer constituting the support tool 15 is a known type of shape-memory thermosetting polymer or thermoplastic polymer. This polymer includes, for example, shape-memory epoxy polymers, shape-memory cyanate ester polymers, shape-memory polyurethane polymers, shape-memory vinyl polymers, shape-memory polyimide polymers, shape-memory maleimide polymers, or combinations thereof, copolymers.

[0074] Due to the shape-memory properties of the polymer, the support tool 15 can recover its original rigid shape, even after long-term repeated use and numerous heating and cooling cycles, and in some cases by being inserted into a dedicated reshaping tool.

[0075] The reinforcing material of the support tool 15 may include one or more elastic fibers.

[0076] Specifically, reinforcing materials may include nylon fibers, Lycra fibers, polyester fibers, carbon fibers, glass fibers, aramid fibers, boron fibers, basalt fibers, polymer fibers, chopped fibers, mesh, three-dimensional fiber preforms, standard weft fabrics, twill fabrics, or other types of fabrics, and combinations thereof. A suitable commercially available example of nylon fiber is nylon manufactured by Ivista (Wichita, Kans).

[0077] In some embodiments, the support tool 15 may include two or more different types of reinforcing materials.

[0078] Each support tool 15 has an elongated prism-shaped portion in the direction of extension of the spar member 13 to be formed, and a hollow-shaped portion in the same direction. Specifically, each support tool 15 has a cross-section with a polygonal external profile that is complementary to the profile of each cavity 14 to be formed in each wall portion 6, 7 of the structure 1. In the illustrated example, each support tool 15 has an isosceles trapezoidal cross-section and is defined by four side wall portions 16 that are joined together by oblique edges 17. These wall portions 16 define the longitudinally penetrating cavity 18 of each support tool 15.

[0079] According to one possible alternative example not shown, the support tool 15 may also have a parallelepiped section with oblique edges.

[0080] In the illustrated example, the prepreg material is externally laminated onto all wall portions 16 of each rigid support tool 15 after application of a resin-based adhesive (known as a "tackifier"), depending on the tackiness of the prepreg material as needed.

[0081] Depending on the cross-sectional configuration of the spar members 13 to be formed, the covering with prepreg material can be limited to some of the wall portions 16 of each support tool 15. Furthermore, depending on the cross-sectional configuration of the spar members 13 to be formed, the support tool 15 can have a cross-section that is always complementary to various, but in any case, shapes of cavities 14 that will be created between the spar members 13.

[0082] Because each support tool 15 is in a rigid state, it is possible to place one or more prepreg layers on each support tool 15. Furthermore, depending on the composition of the prepreg layer, it may be unnecessary to dress the support tool 15 before laminating the prepreg material.

[0083] After the prepreg material is placed on the wall portion 16 of each support tool 15, the support tool 15 is enclosed and sealed in an external bag (not shown as it is well known), and then a vacuum is applied in a known manner within this external bag to achieve compression of the various layers of the prepreg material itself.

[0084] At this point, by bringing the support tools 15, each having compressed laminated prepreg material on its exterior, closer together in a horizontal direction, or by removing the external bags described in the previous paragraph (Figures 7 and 8), it becomes possible to join two prepreg material portions in a horizontal direction together, covering the wall portions 16 that define the diagonal sides of the trapezoidal cross-section of the support tool 15 itself.

[0085] The flat wall portions 16 of each support tool 15 are joined together by the beveled edges 17, so that the prepreg material has the same external shape as the wall portion 16 on which the prepreg material is placed, and the same external shape as the beveled edges 17 that join these wall portions 16 together.

[0086] Therefore, recesses 19 are formed between each pair of support tools 15, which are arranged side by side and externally coated with prepreg material. Recesses 19 have a substantially V-shaped cross-section. Recesses 19 are filled by prepreg filling beads 20 or noodles (Figure 8) having a profile complementary to the profile of recess 19 itself.

[0087] The assembly thus formed, consisting of side-by-side support tools 15, a prepreg material covering the support tools 15, and a filling bead 20, is placed and positioned between skin 11a and skin 12a, and these skins form panels 11 and 12 of each wall section 6 or 7.

[0088] Please note that Figures 8 and 9 show the formation of only wall 6, and it is understood that the formation of the other wall 7 is carried out in exactly the same manner.

[0089] In the illustrated example, the skin 12a of the wall portion 6 is laminated onto a molded surface, which is not shown as it is well known. More precisely, a resin-based adhesive (known as a "tackifier") is pre-applied to this molded surface, and then layers of various fiber-reinforced prepreg materials that form the skin 12a are laminated onto it.

[0090] Preferably, after stacking a predetermined number of layers, such as every four layers, a vacuum consolidation step is performed.

[0091] At this point, the preformed spar members 13 and filling beads 20, held in place by each support tool 15, are positioned on the skin.

[0092] Subsequently, skin 11a (Figure 9), which is laminated and compressed in the same manner as seen with skin 12a, is applied to the thus formed assembly.

[0093] Referring to Figures 10 to 12, the rib 10 is realized by laminating the prepreg material defined above onto a further support tool 21 that extends in the main direction of extension of the rib 10 itself and has the same composition and properties as the support tool 15.

[0094] Specifically, each rib 10 is realized starting from each elongated profile element 22. This elongated profile element 22 has a C-shaped cross-section and consists of a main portion 23 that is flat and substantially perpendicular to the skins 11a, 12a and panels 11, 12, and two end appendages 24 that project transversely in a substantially perpendicular direction from both end edges of the main portion 23 itself and are located on the same side of the main portion 23.

[0095] While not mandatory, preferably, each support tool 21 has an elongated parallelepiped portion in the longitudinal direction of each rib 10 and a hollow portion in the same direction. Specifically, each support tool 21 has a cross-section having a polygonal external profile complementary to the profile of each cavity 9 that will be formed in the structure 1. In the illustrated example, each support tool 21 is defined by four sidewalls 25, these four sidewalls being parallel in pairs and perpendicular to each other in directions of two adjacent wall portions 25.

[0096] The wall portion 25 defines the longitudinally penetrating cavity portion 26 of each support tool 21.

[0097] Each profile element 22 is externally laminated onto three adjacent wall portions 25 of each rigid support tool 21 after the application of a resin-based adhesive (known as a "tackifier").

[0098] Specifically, the wall portion 25 of the rigid support tool 21 is completely covered with the prepreg of each profile element 22, while two other wall portions 25 adjacent to this wall portion 25 are covered only over a predetermined length.

[0099] Therefore, each support tool 21 is able to support each prepreg C-shaped profile element 22 on one of its sides.

[0100] Because the support tools 21 are in a rigid state, it is possible to place one or more prepreg layers on each support tool 21. Furthermore, depending on the composition of the prepreg layers, it may be unnecessary to dress the support tools 21 before laminating the profile elements 22.

[0101] After the profile elements 22 are placed on the walls 25 of each support tool 21, the support tools 21 are enclosed and sealed in an external bag (not shown as it is well known), and then a vacuum is applied in a known manner within this external bag to achieve compression of the profile elements 22 themselves.

[0102] At this point, by arranging or positioning the support tools 21, each having a compressed prepreg profile element 22 (Figure 12) on its exterior, side by side, the support tools 21 themselves can maintain these profile elements 22 at a desired distance from one another. Furthermore, appropriate prepreg filling beads 29 or noodles are inserted into the areas between the skins 11a and the connecting areas between the main portion 23 and each end appendage 24.

[0103] Specifically, each prepreg profile element 22 is The main part 23 is positioned between two opposing wall portions 25 of two mutually adjacent support tools 21. The end appendages 24 are placed on each wall portion 25 of each support tool 21, which are substantially parallel to each other and substantially perpendicular to the wall portions 25 that support the main portion 23. They are arranged in this manner.

[0104] The assembly thus formed, consisting of a prepreg profile element 22 and a support tool 21 for holding the prepreg profile element 22 in place, is placed on the previously formed skin 11a and on the assembly that will form the wall portion 6.

[0105] Similar to what was seen with respect to the formation of wall 6, skins 11a and 12a, which will form wall 7, are applied to the assembly defined by the profile elements 22 and the support tools 21, with interpositions of preformed spar members 13 and filling beads 20 held in place by each support tool 15 between the skins 11a and 12a themselves (Figures 13 and 14).

[0106] This entire assembly is inserted into a mold. The mold is a well-known type and is not shown.

[0107] As can be seen in Figures 13 to 15, when the support tool 15 is attached to define the structure 1 to be formed, it has a main extending direction that is lateral to the main extending direction of the support tool 21.

[0108] At this point, the breathable cloth (not shown as it is well known) and the tubular bags 27 that protrude forward and backward from the support tools 15 and 21 themselves are inserted into the interior of each support tool 15 and 21 (Figure 15).

[0109] According to possible alternative examples not shown, the breathable cloth and tubular bag 27 can be positioned around each support tool 15, 21.

[0110] An additional outer bag 28, having two visible upper flaps 28a and lower flaps 28b in Figure 15, is positioned around a pre-fabricated structure 1, which will be formed and welded to the protruding end of the tubular bag 26 when in use.

[0111] The assembly thus formed, comprising a support tool 15, a support tool 21, a filling bead 20, prepreg material positioned around the support tool 15, a profile element 22, and skins 11a and 12a, is placed in an autoclave in a known (not shown) manner and solidified at predetermined pressure and temperature values ​​(for example, in the case of epoxy resin, the solidification temperature is approximately 180°C and the solidification pressure is between 6 and 7 bar).

[0112] During the solidification step, thermal stimulation causes the walls 16 and 25 of the support tools 15 and 21 to transition from a rigid state to a flexible elastomer state. In fact, the support tools 15 and 21 are configured to become flexible elastomers at temperatures below the solidification temperature and above 50°C. Solidification pressure acts both outside the structure 1 to be formed and inside the support tools 15 and 21, as well as inside the cavities 9 and 14 of the structure 1 itself through the tubular bag 26, so the tubular bag 26 compresses the walls 16 and 25 that have become flexible due to the state change. As a result of the solidification pressure, the walls 16 and 25 then uniformly compress the prepreg material to be polymerized.

[0113] Once the solidification step is complete, the support tools 15 and 21 are heated again to transition to a flexible elastomer state, which allows them to be removed from the cavities 9 and 14 of the molded structure 1.

[0114] A modified example of Figure 16 shows structure 1. In this structure 1, the rib 10 has an I-shaped cross-section and is realized by joining two profile elements 22 to arrange the main parts 23 in a face-to-face contact state, and having the end appendages 24 protrude from opposite directions (Figure 17).

[0115] From the examination of the features of structure 1 and the features of the method for manufacturing structure 1 as described above, the advantages that can be realized are clear.

[0116] Specifically, the wing section 4 and tail section 5, which have a structure 1 having both spar members 13 and ribs 10, exhibit high performance not only in bending but also in torsion.

[0117] In other words, structure 1 behaves the same as known types of structures under flexible loads, but is capable of a better response to torsional loads during operation.

[0118] Furthermore, the method described above is simpler and faster than known methods because the prepreg material is laminated directly onto the support tools 15 and 21, and not onto a suitable preforming tool, but then transferred to the support tool.

[0119] Finally, it becomes unnecessary to remove the support tools 15 and 21 before the solidification step, because these tools uniformly compress the prepreg material, causing it to exhibit a flexible elastic state during the solidification process.

[0120] It is clear that modifications and alterations can be made to the structure 1 and related manufacturing methods described and illustrated herein without departing from the scope of protection defined by the claims.

[0121] Specifically, the support tools 15 and 21 can be made from polystyrene or other low-melting-point thermoplastic materials, so that these tools melt and "disappear" at the solidification temperature. The solidification pressure is applied through the tubular bag 27.

[0122] In this example, by using the low-melting-point material of the type described above for the support tools 15 and 21, the only operation that will be performed after solidification is the removal of the tubular bag 27.

[0123] Further alternatives could include support tools that are soluble in water or other liquids.

[0124] In this example of a support tool, at the end of the solidification process, the removal of the support tools 15 and 21 is achieved by dissolving them in water or a dedicated solvent. [Explanation of Symbols]

[0125] 1 structure 2. Airplane 3 Torso 4 Wings 4a Base part 4b Free end part 4c Wind leading edge 4d wind trailing edge 5 tail fin 6. First wall section 7. Second wall section 8 Interconnected elements 9 Elongated cavity 10 Ribs 11. Panel 1 11a Skin 12. Second panel, outermost panel 12a Skin 13 Spar members 14 Elongated cavity 15 Support Tools 16 Side walls, walls, flat walls 17 Beveled Edges 18 Longitudinal through cavity 19 Recess 20 prepreg filling beads 21 Support Tools 22 Elongated profile elements, prepreg C-shaped profile elements, prepreg profile elements 23 Main parts 24 End attachments 25 Wall 26 Longitudinal through-cavity, tubular bag 27 Tubular bags 28 External Bags 28a Upper flap 28b Downward flap 29 Prepreg-filled bead

Claims

1. A method for manufacturing a monolithic structure (1) made of composite material for a wing (4) or tail (5) of an aircraft (2), wherein the structure (1) is manufactured starting from a prepreg material containing a fiber-reinforced polymer matrix, The first wall section (6), A second wall portion (7) is positioned along the maximum surface of the second wall portion (7) and facing the maximum surface of the first wall portion (6), and is spaced a non-zero amount away from the first wall portion (6) itself. At least one interconnecting element (8) extends in the transverse direction between the first wall portion (6) and the second wall portion (7), is connected to the first wall portion (6) and the second wall portion (7), and defines each first elongated cavity portion (9) by the first wall portion (6) and the second wall portion (7) themselves. Equipped with, The first wall portion (6) and the second wall portion (7) extend symmetrically on both sides in a direction (B) that coincides with the direction of extension of the wing portion (4) or tail wing (5) during use, from the base portion (4a) connected to the fuselage (3) of the aircraft (2) to the free end portion (4b) of the wing portion (4) or tail wing (5) itself. The interconnecting element (8) is a rib (10) that extends in the transverse direction with respect to the direction (B), At least one of the first wall portion (6) and the second wall portion (7) (6, 7) has a sandwich structure, A first panel (11) facing the other (7, 6) of the first wall (6) and the second wall (7), A second panel (12) facing the first panel (11), At least one spar member (13) extending in the transverse direction between the first panel (11) and the second panel (12), connected to the first panel (11) and the second panel (12), defining each second elongated cavity (14) by the first panel (11) and the second panel (12) themselves, and extending in the transverse direction with respect to the rib (10), and In a method that provides, a) A step of preparing at least two elongated first support tools (21), each of which has a cross-section having a polygonal external profile complementary to the profile of the first elongated cavity (9) to be formed within the structure (1), b) The step of arranging the first support tools (21) side by side in a lateral direction so that one or more layers of the prepreg material that will form the ribs (10) are placed between the first support tools (21), c) A step of preparing at least two second elongated support tools (15), each of which has a cross-section having a polygonal external profile complementary to the profile of the second elongated cavity (14) to be formed within the structure (1), d) The step of arranging the second support tools (15) side by side in the lateral direction so that one or more layers of the prepreg material that will form the spar member (13) are placed between the second support tools (15), e) The step of forming the skins (11a, 12a) by laminating one or more layers of the prepreg material for at least three separate skins (11a, 12a), f) The step of forming the first wall portion (6) in the first wall portion (6) and the second wall portion (7) by inserting the assembly formed of the prepreg material positioned between the first and second skins of the at least three separate skins (12a, 11a) using the second support tool (15), g) Inserting an assembly formed of the first support tool (21) and the prepreg material positioned between the first support tools (21) between the second skin (11a) of the at least three separate skins (11a) and the third skin (11a) of the at least three separate skins (11a, 12a), h) The step of placing the group formed in steps a) to g) in an autoclave at a predetermined solidification temperature and pressure. Includes, The first support tool (21) and the second support tool (15) are hollow inside and have a composition such that they become rigid at room temperature. The first support tool (21) has a main extending direction that is the transverse axis with respect to the main extending direction of the second support tool (15), Steps b) and d) are carried out by directly laminating the prepreg material onto the outer surfaces of the walls (25, 16) of the first support tool (21) and the second support tool (15) in a rigid state. method.

2. The first support tool (21) and the second support tool (15) have a composition based on reinforcing materials and polymers suitable for transitioning from a rigid state to a flexible elastomer state and vice versa in response to heating and cooling, respectively, and the first support tool (21) and the second support tool (15) are configured to exhibit the flexible elastomer state at temperatures below the solidification temperature and above 50°C, the solidification pressure is applied both inside the autoclave and inside the first support tool (21) and the second support tool (15) during step h), the walls (25, 16) of the first support tool (21) and the second support tool (15) are made flexible by the transition from the rigid state to the flexible elastomer state and are therefore pressed by the solidification pressure itself to adhere to the adjacent prepreg material. The method according to claim 1.

3. The polymer of the first support tool (21) and the second support tool (15) is a shape memory thermosetting polymer or a thermoplastic polymer, and the reinforcing material of the first support tool (21) and the second support tool (15) includes one or more elastic fibers. The method according to claim 2.

4. The first support tool (21) and / or the second support tool (15) are made from a low-melting-point thermoplastic material, particularly polystyrene, and are externally covered with the tubular bag (27) so that during step h) the first support tool (21) and / or the second support tool (15) melt and the solidification pressure is applied through the tubular bag (27). The method according to claim 1.

5. The first support tool (21) and / or the second support tool (15) are made from a material that is soluble in a given solvent, such as water. The method according to claim 1.

6. The method is such that the other (7, 6) of the first wall portion (6) and the second wall portion (7) of the structure (1) also has the same sandwich structure as the one of the first wall portion (6) and the second wall portion (7), i) The step of forming a fourth skin (12a) by laminating one or more layers of the prepreg material, l) The step of inserting a further assembly formed of the prepreg material between the third skin (11a) and the fourth skin (12a) by the second support tool (15) and between the second support tools (15) themselves, thereby also forming the other (7, 6) of the first wall (6) and the second wall (7) in the sandwich structure. Further including, The method according to any one of claims 1 to 5.

7. The distance between the first panel (11) and the second panel (12) of one of the first wall portion (6) and the second wall portion (7) is shorter than the distance between the first panel (11) and the other of the first wall portion (6) and the second wall portion (7) (7, 6). The method according to any one of claims 1 to 5.

8. The distance between the first panel (11) and the second panel (12) of the first wall (6) and the second wall (7) is shorter than the distance between the two opposing first panels (11). The method according to any one of claims 1 to 7.

9. The rib (10) is a profile element having an open or closed cross-section. The method according to any one of claims 1 to 8.

10. The spar member (13) has an open cross section or a closed cross section. The method according to any one of claims 1 to 9.

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