Composite box-shaped monolithic structure for aircraft fuselage and wings and method for manufacturing the structure

Abstract

To provide an efficient and inexpensive method for manufacturing a box-shaped monolithic structure of a composite material.SOLUTION: A method for manufacturing a box-shaped monolithic structure with a cavity by curing a fiber reinforced prepreg material includes a step of using two or more support tools 15 with a reinforcing material and a polymer-based composition suitable for allowing a transition from a rigid state to a flexible elastomeric state and vice versa depending on a morphology and heating / cooling complementary to the cavity being elongated, hollow inside, and manufactured. The support tool is configured to allow direct lamination of the prepreg material on its outer wall in the rigid state and to set a flexible elastomer state at a temperature above 50°C below a curing temperature. During a curing operation, curing pressure is applied both outside the structure to be formed and inside the support tool, making a wall of the support tool flexible and pressing the prepreg material to be cured.SELECTED DRAWING: Figure 5

Description

[Technical Field]

[0001] [CROSS-REFERENCE TO RELATED APPLICATIONS] This patent application claims priority to Italian Patent Application No. 102020000003976, filed February 26, 2020, the entire disclosure of which is incorporated herein by reference.

[0002] The present invention relates to box-shaped monolithic structures in composite material for aircraft fuselages and wings, in particular for airplane horizontal stabilizers, to which the following description makes explicit reference without loss of generality.

[0003] The present invention also relates to a method for manufacturing the box-shaped monolithic structure described above. [Background technology]

[0004] As is known, a horizontal stabilizer is a structural component comprising a fixed surface that is generally part of the tail of an airplane fuselage.

[0005] The stabilizer has a box-shaped monolithic structure of composite material inside, essentially: The upper wall and a lower wall facing the upper wall and spaced apart from the upper wall by a non-zero amount; and a plurality of interconnecting elements or spars extending between and separating the upper and lower walls and having elongated cavities in said upper and lower walls, respectively, extending between front and rear openings disposed in a plane transverse to the walls themselves and the spars.

[0006] The use of composite materials allows the overall weight of the aircraft to be reduced while at the same time obtaining a very strong structure.

[0007] The upper and lower walls are generally parallel, or more precisely, there is a slight convergence between the upper and lower walls towards the front or rear opening.

[0008] The spar is formed by partitions that are substantially perpendicular to the upper and lower walls, or more generally transverse to them.

[0009] The described structure is manufactured by a known method including the steps shown below, starting from a prepreg with a thermosetting matrix (resin) reinforced by fibers of different nature, such as carbon fibers, aramid fibers, glass fibers, etc.

[0010] First, each spar is preformed in a suitable preforming tool by creating, in the prepreg state, two elongated profile bars with a C-section, consisting of a back surface and two wings protruding at right angles from the opposite end of the back surface, and then to form the spar, the two C-shaped profile bars are joined along their respective back surfaces with opposing recesses.

[0011] Specifically, each C-shaped profile bar is laminated to three flat surfaces of a preformed tool having a substantially parallelepiped profile: one surface of the tool is completely coated with prepreg, while the other two surfaces, adjacent to the first surface, perpendicular to the first surface and parallel to each other, are coated only at their extensions.

[0012] After the required vacuum consolidation, at a predetermined temperature, the C-shaped profile bars are joined two by two in the manner described above to form the desired number of spars, which are held in place at a predetermined distance for the subsequent curing step by respective rigid support tools having the form of elongated pins.

[0013] The preforming operation is a vacuum and temperature application process that allows the prepreg material to acquire a solidified shape without undergoing a polymerization process.

[0014] Each support tool is constituted by a substantially parallelepiped, rigid and solid elongated body bounded by planes and having a cross section corresponding to the shape of the elongated cavity of the structure to be produced.

[0015] Before placing each spar between two support tools, still in the form of preformed prepregs, the support tools are subjected to the following successive steps: - applying a layer of a release agent, for example in the form of a film, to each support tool to facilitate subsequent extraction of the support tool itself from its respective elongated cavity; Attaching the tubular bags thus prepared to each support tool, leaving excess tubular bags on both ends of the support tool itself for subsequent sealing operations; wrapping a ventilating fabric around the outside of each support tool and tubular bag and securing the flaps with a sealant; Attaching the thus prepared tubular separator film to each support tool, again leaving excess tubular separator film on both ends of the support tool itself for subsequent sealing operations; sealing the ends of the tubular bag and the tubular separator film using a sealant; applying a vacuum and waiting for the tubular separator film to shrink the entire dressing onto the relative support tool.

[0016] At this point, the preformed spars, each consisting of two C-shaped profile bars joined together along their respective back surfaces, are placed between the support tools previously subjected to the aforementioned dressing operation. Specifically, each preformed spar: The back surfaces of the opposing pair of C-shaped profile bars that constitute it are inserted between two opposing flat surfaces of two mutually adjacent support tools; The wings of each profile bar rest on the respective planes of the relative support tools and are arranged so that they are parallel to each other and substantially perpendicular to the plane supporting the back surface of the profile bar itself.

[0017] The assembly thus formed and configured with the preformed spar and previously dressed support tool is inserted into a molding die that includes a lower plate, an upper plate, and two opposing side walls connecting the lower and upper plates.

[0018] Specifically, one or more layers of prepreg material are laminated to each of the two upper and lower plates, which layers are set to define upper and lower skins that, after a curing step, make up the upper and lower walls of the box-shaped structure to be manufactured.

[0019] More precisely, the preformed spars, held in place by their respective support tools, are placed on the lower plate of the mould, which supports the lower skin, and then the upper plate of the mould, which supports the upper skin, is closed on the assembly formed by the side walls of the mould itself and the preformed spars and support tools.

[0020] At this point, the separator film, ventilation fabric, and bag film are placed in succession across the mold, and the bag film is sealed with a sealant at the base of the mold.

[0021] A tubular separator film is placed on each support tool, open at both ends, and excess length is cut off to the corresponding support tool.

[0022] The tubular bag on each support tool is unrolled and sealed at both ends.

[0023] Thus, so-called envelope bags are formed by sealing the edges of an outer bag film placed on a forming mold with the ends of a tubular bag on a support tool and by sealing the ends of adjacent tubular bags.

[0024] At this point, a vacuum is applied to the interior of the envelope bag until the material of the outer bag shrinks towards the exterior surface of the mold.

[0025] The ends of the tubular bag are then opened and by continuing to apply a vacuum, the tubular bag itself is detached from its respective support tool, tending to minimize the volume enclosed within the envelope bag.

[0026] At this point, the support tool is withdrawn and the assembly thus formed is transferred to an autoclave where the curing operation is carried out at predetermined pressure and temperature values ​​(for example, for epoxy resins, the curing temperature is about 180°C and the curing pressure is 6-7 bar).

[0027] The method described and the tools used make it possible to correctly position the preforms and maintain their position during the closing operation of the mold to form and produce the envelope bag.

[0028] Extracting the support tool prior to the cure cycle prevents the tool from deforming inappropriately under the pressure and temperature conditions required for cure to ensure uniform pressure is applied to all parts of the composite.

[0029] Instead, the tubular bags can apply pressure evenly to the composite parts they contact. Summary of the Invention [Problem to be solved by the invention]

[0030] However, the applicant has realized that the described method and the box-shaped structures obtained thereby can be improved. In particular, there is a need felt in the field to simplify the method in order to make it more efficient and cheaper. Furthermore, the finish quality of the surfaces, especially the interior of the box-shaped structures, requires further improvement.

[0031] The object of the present invention is to provide a method for manufacturing a box-shaped monolithic structure out of composite material for an aircraft fuselage, which is reliable, has limited costs and is able to meet at least one of the above-mentioned requirements and those associated with methods for manufacturing box-shaped monolithic structures out of composite material of known types.

[0032] A further object of the present invention is to manufacture a box-shaped monolithic structure of composite material for an aircraft fuselage, which has a high quality finish and is capable of meeting the above needs. [Means for solving the problem]

[0033] According to the invention, the first object is achieved by a method for manufacturing a box-shaped monolithic structure from composite material for an aircraft fuselage as set forth in claim 1.

[0034] According to the present invention, the second object is achieved by a monolithic box structure of composite material for an aircraft fuselage as set forth in claim 9. [Brief explanation of the drawings]

[0035] [Figure 1] 1 is a perspective view of an aircraft tail section incorporating therein a monolithic box structure of composite material manufactured in accordance with the teachings of the present invention; [Figure 2] 2 shows a perspective view of the box-shaped monolithic structure of FIG. 1 on an enlarged scale with parts removed for clarity; [Figure 3] 3A and 3B are perspective views of a support tool at successive steps during the lamination of a portion of the spar of the structure of FIGS. 1 and 2; [Figure 4] 4 is a reduced-scale exploded perspective view of the series of support tools of FIG. 3 during assembly of a plurality of spars of the structure of FIGS. 1 and 2; FIG. [Figure 5] 4 is an exploded perspective view of a mold used to manufacture the structure of FIGS. 1 and 2, and into which the series of support tools of FIG. 3 are inserted for this purpose; FIG. [Figure 6] 6A and 6B are perspective views, enlarged in scale, of the mold of FIG. 5 during successive steps of the method of manufacturing the structure of FIGS. 1 and 2. [Figure 7] 6A and 6B are perspective views, enlarged in scale, of the mold of FIG. 5 during successive steps of the method of manufacturing the structure of FIGS. 1 and 2. DETAILED DESCRIPTION OF THE INVENTION

[0036] In order that the invention may be better understood, preferred, non-limiting embodiments thereof will now be described, purely by way of example, with the aid of the accompanying drawings, in which: With reference to Figures 1 and 2, the reference number 1 indicates, in general, a box-shaped monolithic structure of composite material for an aircraft fuselage, in particular for a horizontal stabilizer 2 of a fuselage 3 of an airplane, for example an airplane 4.

[0037] Structure 1 (Figure 2) is an upper wall 5; a lower wall (6) facing the surface of the upper wall (5) having the greater extension along its surface with the greater extension and spaced apart from the upper wall (5) by a non-zero amount; and a plurality of interconnecting elements or spars 7 extending between and separating the upper and lower walls 5, 6 and having a series of elongated cavities 8 in said upper and lower walls 5, 6, respectively, extending between respective front and rear openings 9 (only the front openings are visible in Figure 2) arranged in a plane transverse to the walls 5, 6 themselves and to the spars 7.

[0038] In the example shown, the upper wall 5 and the lower wall 6 are substantially parallel to one another. According to a possible alternative not shown, the upper wall 5 and the lower wall 6 may also slightly converge between them towards the front or rear opening 9.

[0039] The spar 7 is constituted by a longitudinally elongated partition that is substantially perpendicular to the upper and lower walls 5 and 6, or more generally transverse to and substantially parallel to the upper and lower walls 5 and 6.

[0040] The structure 1 is made from a prepreg having a polymer matrix, for example made of thermosetting resin, reinforced by fibres which may have different properties, for example carbon fibres and / or aramid fibres and / or glass fibres.

[0041] Alternatively, the structure 1 can be made from a fiber-reinforced thermoplastic matrix prepreg of the type described above.

[0042] In both cases, the upper and lower walls 5, 6 are made from respective skins 5a, 6a (FIG. 5) of the aforementioned fiber-reinforced prepreg laminated on their respective planes, as will be explained in more detail below.

[0043] 3 and 4, each spar 7 is obtained from two aforementioned longitudinally elongated profile bars 11, which have a C-shaped cross section and are each constituted by two end appendages 13 projecting laterally in substantially perpendicular directions from the back surface 12 and from both end edges of the back surface 12 itself, and which are located on the same side thereof. In particular, to form the spar 7, the two profile bars 11 are joined along their respective back surfaces 12 so as to have opposing recesses.

[0044] Advantageously, the profile bars 11 of said prepreg material are stacked on said respective longitudinally elongated support tools 15, which are then set to maintain the spars 7 in position within the structure 1 during the manufacturing steps, and which have a composition based on reinforcing materials and polymers suitable for enabling them to transition from a rigid state to a flexible elastomeric state and vice versa upon heating and respectively cooling, i.e. in response to a temperature stimulus.

[0045] The polymer comprising the support tool 15 is advantageously a thermosetting or thermoplastic polymer with shape memory of a known type, which may be, for example, an epoxy polymer with shape memory, a cyanate ester polymer with shape memory, a polyurethane polymer with shape memory, a vinyl polymer with shape memory, a polyimide polymer with shape memory, a maleimide polymer with shape memory, or combinations thereof, including copolymers.

[0046] Due to the polymer's properties with shape memory, the support tool 15 is able to regain its original rigid shape even after extensive repeated use and numerous heating and cooling cycles.

[0047] The reinforcement material of the support tool 15 includes one or more elastic fibers.

[0048] Specifically, the reinforcement material may also 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, regular weft fibers, twill fibers, or other types of fibers and combinations thereof. A suitable commercially available example of nylon fiber is nylon manufactured by Invista (Wichita, Kansas).

[0049] In some embodiments, the support tool 15 may include two or more different types of reinforcements.

[0050] 3-7, each support tool 15 has a parallelepiped shape that is elongated in the aforementioned longitudinal direction and hollow in the same direction. Specifically, each support tool 15 has a cross-section with an external polygonal outline that is complementary to the outline of each cavity 8 formed in the structure 1. In the illustrated example, each support tool 15 is bounded by four side walls 16, two of which are parallel and each of which is perpendicular to two adjacent walls 16.

[0051] The walls 16 are connected to one another by chamfered edges 17 and delimit a longitudinal through-cavity 18 of each support tool 15 .

[0052] Each profile bar 11 is laminated in a rigid state to the exterior of three adjacent walls 16 of a relative support tool 15 after application of a resin-based adhesive (known as a "tackifier").

[0053] Specifically, the wall 16 of the support tool 15 in its rigid state is completely coated with the prepreg of the relative profile bar 11, while the other two walls 16 adjacent to it are coated only to a certain extent.

[0054] Thus, each support tool 15 can support two prepreg C-shaped profile bars 11 on either side.

[0055] Due to their rigid state, it is possible to place one or more prepreg layers on each support tool 15 and, furthermore, due to their construction, it is not necessary to dress the support tools 15 before laminating the profile bars 11.

[0056] After the profile bars 11 have been placed on the respective walls 16 of each support tool 15, the walls 16 are enclosed and sealed in an outer bag (known per se and not shown) to which a vacuum is applied in a known manner in order to obtain compression of the profile bars 11 themselves.

[0057] At this point, by placing each compressed prepreg profile bar 11 (FIG. 4) laterally adjacent or alongside one another on an externally carrying support tool 15, the profile bars 11 themselves can be joined two by two along their respective back surfaces 12 to form spars 7 that are held at a desired distance from one another by the support tool 15.

[0058] Specifically, each prepreg spar 7 is The back surfaces 12 of the opposing pairs of profile bars 11 that make up the profile bar 11 are inserted between two opposing walls 16 of two adjacent support tools 15, The end appendages 13 of each profile bar 11 rest on respective walls 16 of the relative support tool 15 and are substantially parallel to each other and substantially perpendicular to the walls 16 supporting the back surface 12 of the profile bar 11 itself.

[0059] The flat walls 16 of each support tool 15 are joined by a chamfered edge 17, so that the profile bars 11 take on the same contour as the walls 16 on which they rest and the chamfered edge 17 that joins them.

[0060] Thus, a depression 19 is formed between each pair of adjacent coplanar end appendages 13 having a generally V-shaped cross section. The depressions 19 are filled with respective prepreg filler beads 20 or noodles (FIGS. 4 and 5) having a contour complementary to that of the depressions 19 themselves.

[0061] The thus formed assembly consisting of the prepreg spars 7, the support tool 15 that holds them in place, and the filler beads 20 is inserted into a molding die 21 that includes a bottom plate 22, a top plate 23, and two opposing side plates 24 that connect the bottom plate 22 and the top plate 23 (FIG. 5).

[0062] Specifically, the molding die 21 is hollow inside and therefore has respective end openings 25 at the front and rear (only the front is visible in Figure 5) bounded by front and rear surfaces 27 and 28 of the respective frames defined by the opposite head edges of the lower plate 22, upper plate 23, and side plate 24.

[0063] In particular, on each of the two lower and upper plates 22 and 23, one or more layers (called "plies") of fiber reinforced prepreg material of the type described above are laminated in a manner configured to define upper and lower skins 5a and 6a which, after a curing step, constitute the upper and lower walls 5 and 6 of the structure 1 to be manufactured.

[0064] More precisely, each of the lower plate 22 and the upper plate 23 is pre-applied with a resin-based adhesive (known as a "tackifier") and then the various layers of fiber-reinforced prepreg material are laminated to it, with the adhesive being applied to at least the portions of the lower plate 22 and the upper plate 23 that are configured to receive the fiber-reinforced prepreg material.

[0065] Preferably, after laminating a predetermined number of layers, for example every four layers, a vacuum compaction step is carried out by wrapping the corresponding lower 22 or upper plate 23 and the prepreg in a sealed bag (known per se and not shown), inside which a vacuum is applied in a known manner.

[0066] At this point, the preformed spars 7 and filler beads 20, held in place by their respective support tools 15, are placed onto the lower plate 22 of the mould 21 which supports the lower skin 6a.

[0067] The top plate 23 of the mould 21, which supports the top skin 5a, is then closed on the side plates 24 of the mould 21 itself and on the assembly formed by the preformed spars 7, the filling beads 20 and the support tool 15.

[0068] In this state (Figures 6 and 7), the support tool 15 protrudes from the front and rear openings 25 of the molding die 21 at its respective front and rear ends 15a (only the front is visible in Figures 6 and 7).

[0069] Next, a ventilation fabric (known per se and not shown) and a tubular bag 30 protruding from the support tool 15 itself at the front and rear are inserted inside each support tool 15 itself.

[0070] At each of the front frame surface 27 and the rear frame surface 28, the axial end 31 of a further external tubular bag 32 suitable for placing itself around the respective front or rear end 15a of the support tool 15 during use is sealed by a sealant.

[0071] At this point, the ends of the tubular bags 30 protruding forward from their respective support tools 15 are sealed with a sealant to the axial end 33 of the corresponding outer tubular bag 32 opposite the axial end 31, and the same sealing operation is performed between the corresponding outer tubular bag 32 sealed at the rear of the molding die 21 and the ends of the tubular bags 30 protruding rearward from their respective support tools 15.

[0072] All ends of adjacent tubular bags 30 are then closed and sealed together.

[0073] In this way, a so-called envelope bag is formed, the front and rear parts of which are respectively indicated by reference numeral 35.

[0074] The mold 21 thus prepared is transferred to an autoclave together with the support tool 15, the upper and lower skins 5a and 6a, the spars 7, the filling beads 20 and the envelope bag 35, and the curing operation is carried out at predetermined pressure and temperature values ​​(for example, in the case of epoxy resin, the curing temperature is about 180°C and the curing pressure is 6-7 bar).

[0075] During the curing step, due to the temperature stimulus, the walls 16 of the support tool 15 transition from a rigid state to a flexible elastomeric state. In fact, the support tool 15 is configured to set the flexible elastomeric state at a temperature lower than the curing temperature but higher than 50°C. The curing pressure acts both on the outside of the structure 1 to be formed and on the inside of the support tool 15, and therefore on the inside of the cavity 8 of the structure 1 itself via the tubular bag 30, thus pressing against the walls 16, which have been made flexible by the change in state. Then, as a result of the curing pressure, the walls 16 uniformly press against the polymerizing prepreg material.

[0076] Once the curing step is complete, the support tool 15 is again heated to transition to a flexible elastomeric state so that it can be withdrawn from the cavity 8 of the just-formed structure 1 .

[0077] As a result of the above method, and in particular the use of a support tool 15 that transitions from a rigid state to a flexible elastomeric state during the curing operation, it is possible to obtain a structure 1 having an average surface roughness of less than 2 microns on all surfaces included between the upper wall 5 and the lower wall 6.

[0078] This result is particularly important as it makes it possible to obtain a smooth surface on the structure 1 which provides an accurate and stable joint for the connection of devices or other structures in the fuselage 3 of the airplane 4.

[0079] The described method is simpler and faster than known methods, since the profile bar 11 is not laid up in a special preforming tool but directly on the support tool 15 and then transferred to the support tool. Furthermore, there is no need to extract the support tool 15 before the curing step, since these tools set up a flexible elastomeric state during the curing operation and therefore press uniformly onto the prepreg material.

[0080] Obviously, modifications can be made to the method and structure 1 as described and illustrated herein without departing from the scope of protection defined by the claims.

Claims

1. A method for manufacturing a box-shaped monolithic structure (1) for a fuselage (3a) and wings (3b) of an aircraft (4) using a fiber-reinforced polymer matrix prepreg material, comprising: The structure (1) is an upper wall (5); a lower wall (6) facing said upper wall (5) and spaced from said upper wall (5) by a non-zero amount; at least one interconnecting element (7) extending laterally between and connected to said upper wall (5) and lower wall (6) and defining respective elongated cavities (8) therein; Equipped with The method comprises: a) providing at least two elongated support tools (15), each having a cross section with an external polygonal contour complementary to the contour of a cavity (8) to be formed in said structure (1); b) externally placing, on three adjacent walls (16) of each support tool (15), elongated profile bars (11) made of said prepreg material, with a C-shaped cross section and comprising a back surface (12) and two end appendages (13) projecting laterally from said back surface (12) and located on the same side of said back surface (12) itself; c) joining a plurality of said profile bars (11) along their respective back surfaces (12) so that the profile bars (11) themselves have opposing concave surfaces, and arranging a plurality of said support tools (15) side by side to form an assembly; d) laminating one or more layers of said prepreg material onto the lower plate (22) of the mould (21) to form a lower skin (6a) adapted to constitute said lower wall (6) of said structure (1); e) inserting the assembly of step c) onto the lower skin (6a) supported by the lower plate (6) of the mould (21) and between the respective side plates (24) delimiting the mould (21) itself, the profile bars (11) being arranged with their respective end appendages (13) in the same plane in contact with the lower skin (6a); f) laminating one or more layers of said prepreg material onto the top plate (23) of said mould (21) to form an upper skin (5a) adapted to constitute said top wall (5) of said structure (1); g) mounting the upper plate (23) on the side plate (24) of the mould (21) in a position facing the lower plate (22) so that the upper skin (5a) is placed in contact with the end appendix (13) of the profile bar (11) opposite the end appendix (13) in contact with the lower skin (6a); h) transferring the mould (21) thus prepared and filled with the upper and lower skins (5a) and the profile bar (11) to an autoclave to carry out the curing operation at a predetermined curing temperature and pressure; Including, the support tool (15) is hollow and comprises a composition based on a reinforcement material and a polymer suitable for transitioning from a rigid state to a flexible elastomeric state upon heating and from a flexible elastomeric state to a rigid state upon cooling; The support tool (15) is configured to set the flexible elastomeric state at a temperature below the curing temperature and above 50°C; Step b) is carried out by directly laminating the prepreg material onto the outer surface of the wall (16) of the support tool (15) in a rigid state, During step h), a curing pressure is applied both inside the mould (21) and inside said support tool (15), whose walls (16) are made flexible by transitioning from said rigid state to said flexible elastomeric state and therefore, pressed by the curing pressure itself, adhere to said profile bar (11) and to said lower skin (6a) and upper skin (5a); Step b) is preceded only by step n) of applying a first resin-based adhesive to at least a portion of the wall (16) of the rigid support tool (15) configured to receive the prepreg material; A method characterized by:

2. A method for manufacturing a box-shaped monolithic structure (1) for a fuselage (3a) and wings (3b) of an aircraft (4) using a fiber-reinforced polymer matrix prepreg material, comprising: The structure (1) is an upper wall (5); a lower wall (6) facing said upper wall (5) and spaced from said upper wall (5) by a non-zero amount; at least one interconnecting element (7) extending laterally between and connected to said upper wall (5) and lower wall (6) and defining respective elongated cavities (8) therein; Equipped with The method comprises: a) providing at least two elongated support tools (15), each having a cross section with an external polygonal contour complementary to the contour of a cavity (8) to be formed in said structure (1); b) externally placing, on three adjacent walls (16) of each support tool (15), elongated profile bars (11) made of said prepreg material, with a C-shaped cross section and comprising a back surface (12) and two end appendages (13) projecting laterally from said back surface (12) and located on the same side of said back surface (12) itself; c) joining a plurality of said profile bars (11) along their respective back surfaces (12) so that the profile bars (11) themselves have opposing concave surfaces, and arranging a plurality of said support tools (15) side by side to form an assembly; d) laminating one or more layers of said prepreg material onto the lower plate (22) of the mould (21) to form a lower skin (6a) adapted to constitute said lower wall (6) of said structure (1); e) inserting the assembly of step c) onto the lower skin (6a) supported by the lower plate (6) of the mould (21) and between the respective side plates (24) delimiting the mould (21) itself, the profile bars (11) being arranged with their respective end appendages (13) in the same plane in contact with the lower skin (6a); f) laminating one or more layers of said prepreg material onto the top plate (23) of said mould (21) to form an upper skin (5a) adapted to constitute said top wall (5) of said structure (1); g) mounting the upper plate (23) on the side plate (24) of the mould (21) in a position facing the lower plate (22) so that the upper skin (5a) is placed in contact with the end appendix (13) of the profile bar (11) opposite the end appendix (13) in contact with the lower skin (6a); h) transferring the mould (21) thus prepared and filled with the upper and lower skins (5a) and the profile bar (11) to an autoclave to carry out the curing operation at a predetermined curing temperature and pressure; Including, the support tool (15) is hollow and comprises a composition based on a reinforcement material and a polymer suitable for transitioning from a rigid state to a flexible elastomeric state upon heating and from a flexible elastomeric state to a rigid state upon cooling; The support tool (15) is configured to set the flexible elastomeric state at a temperature below the curing temperature and above 50°C; Step b) is carried out by directly laminating the prepreg material onto the outer surface of the wall (16) of the support tool (15) in a rigid state, During step h), a curing pressure is applied both inside the mould (21) and inside said support tool (15), whose walls (16) are made flexible by transitioning from said rigid state to said flexible elastomeric state and therefore, pressed by the curing pressure itself, adhere to said profile bar (11) and to said lower skin (6a) and upper skin (5a); Steps d) and f) are preceded only by step o) of applying a second resin-based adhesive to at least a portion of the lower (22) and upper (23) plates of the mould (21) configured to receive the prepreg material. A method characterized by:

3. A method for manufacturing a box-shaped monolithic structure (1) for a fuselage (3a) and wings (3b) of an aircraft (4) using a fiber-reinforced polymer matrix prepreg material, comprising: The structure (1) is an upper wall (5); a lower wall (6) facing said upper wall (5) and spaced from said upper wall (5) by a non-zero amount; at least one interconnecting element (7) extending laterally between and connected to said upper wall (5) and lower wall (6) and defining respective elongated cavities (8) therein; Equipped with The method comprises: a) providing at least two elongated support tools (15), each having a cross section with an external polygonal contour complementary to the contour of a cavity (8) to be formed in said structure (1); b) externally placing, on three adjacent walls (16) of each support tool (15), elongated profile bars (11) made of said prepreg material, with a C-shaped cross section and comprising a back surface (12) and two end appendages (13) projecting laterally from said back surface (12) and located on the same side of said back surface (12) itself; c) joining a plurality of said profile bars (11) along their respective back surfaces (12) so that the profile bars (11) themselves have opposing concave surfaces, and arranging a plurality of said support tools (15) side by side to form an assembly; d) laminating one or more layers of said prepreg material onto the lower plate (22) of the mould (21) to form a lower skin (6a) adapted to constitute said lower wall (6) of said structure (1); e) inserting the assembly of step c) onto the lower skin (6a) supported by the lower plate (6) of the mould (21) and between the respective side plates (24) delimiting the mould (21) itself, the profile bars (11) being arranged with their respective end appendages (13) in the same plane in contact with the lower skin (6a); f) laminating one or more layers of said prepreg material onto the top plate (23) of said mould (21) to form an upper skin (5a) adapted to constitute said top wall (5) of said structure (1); g) mounting the upper plate (23) on the side plate (24) of the mould (21) in a position facing the lower plate (22) so that the upper skin (5a) is placed in contact with the end appendix (13) of the profile bar (11) opposite the end appendix (13) in contact with the lower skin (6a); h) transferring the mould (21) thus prepared and filled with the upper and lower skins (5a) and the profile bar (11) to an autoclave to carry out the curing operation at a predetermined curing temperature and pressure; Including, the support tool (15) is hollow and comprises a composition based on a reinforcement material and a polymer suitable for transitioning from a rigid state to a flexible elastomeric state upon heating and from a flexible elastomeric state to a rigid state upon cooling; The support tool (15) is configured to set the flexible elastomeric state at a temperature below the curing temperature and above 50°C; Step b) is carried out by directly laminating the prepreg material onto the outer surface of the wall (16) of the support tool (15) in a rigid state, During step h), a curing pressure is applied both inside the mould (21) and inside said support tool (15), whose walls (16) are made flexible by transitioning from said rigid state to said flexible elastomeric state and therefore, pressed by the curing pressure itself, adhere to said profile bar (11) and to said lower skin (6a) and upper skin (5a); the polymer of the support tool (15) is a thermosetting or thermoplastic polymer with shape memory; A method characterized by:

4. the reinforcement of the support tool (15) comprises one or more elastic fibers; 4. The method according to any one of claims 1 to 3.

5. At least two interconnection elements (7) are formed, and in step a) at least three elongated support tools (15) are provided, and in step b) two profile bars (11) are stacked directly on either side of at least one wall (16) of the support tools (15).

5. The method according to any one of claims 1 to 4.

6. After step h), it further comprises a step p) of heating said support tools (15) to bring them into a flexible elastomeric state and extracting them from the respective cavities (8) of the just formed structure (1), 6. The method according to any one of claims 1 to 5.

7. 7. A method according to any one of claims 1 to 6, characterized in that it provides an average surface roughness of less than 2 microns on all surfaces included between the upper wall (5) and the lower wall (6). A box-shaped monolithic structure (1) for the fuselage (3a) and wings (3b) of an aircraft (4).

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