Method for producing Y-stringer, double-Y-cross-section spar and torsion box for aircraft, and use of Y-stringer as reinforcement for panel of aircraft

Abstract

The invention relates to a method for manufacturing a Y-stringer, a double-Y-cross-section spar and a torsion box for an aircraft and to the use of a Y-stringer as a reinforcement for a panel of an aircraft, the Y-stringer (100) comprising: a stringer web (110) having an I-shaped cross section; a lower flange (130a, 130b); a first open triangular cross-sectional structure (120) including a first lower apex (120a) and a second lower apex (120b) joined to the lower flanges (130a, 130b), respectively, and an upper apex (120c) joined to a first end of the stringer web (110), where the first open triangular cross-sectional structure (120) and the stringer web (110) form a Y-shaped cross-section.

Description

Technical Field

[0001] The invention relates to a method for manufacturing longerons and spars made of composite materials, in particular longerons with a Y-shape and spars with a double Y-shaped cross section for use in the aerospace industry.

[0002] The stringer according to the invention can be used as a reinforcement in any part of the airframe of an aircraft.

[0003] The spar according to the invention can be used as a spar for an aircraft frame, in particular as a spar for an aircraft torsion box. Background Art

[0004] CA2619767A1 relates to a stringer made of composite material and used to reinforce composite panels, in particular composite panels used in the aerospace industry. The stringer is formed by a base for connecting the stringer to the panel and a structural element provided at the end opposite to the base with a structural reinforcement made of high modulus unidirectional fibers of the same material as the stringer or of another compatible material.

[0005] Existing "Double T" and "Omega" cross-section stringers have limited moment of inertia and may suffer from buckling and post-buckling behavior as well as torsional behavior. In addition, other manufacturing requirements such as weight reduction and cost savings favor improvements to existing stringers.

[0006] EP3095691A1 relates to a multi-spar torsion box structure, which includes a plurality of spars made of composite materials, wherein the plurality of spars are arranged to form a multi-unit structure, wherein the multi-unit structure has two or more units extending one after another in the span direction in the torsion box, and an upper skin cover and a lower skin cover made of composite materials respectively connected to the upper surface and the lower surface of the multi-unit structure.

[0007] Existing "double T" and "Omega" section stringers used in multi-spar torsion box structures have limited moments of inertia and may suffer from buckling and post-buckling behavior as well as torsional behavior. In addition, other manufacturing requirements such as weight reduction and cost savings favor improvements to existing spars and torsion boxes.

[0008] The present invention meets these needs and addresses the shortcomings of currently existing stringers used as stiffeners and spars used in multi-spar torsion box structures. Summary of the invention

[0009] The present invention proposes a first manufacturing method for obtaining a composite longitudinal beam (carbon fiber or glass fiber with thermosetting or thermoplastic resin), which composite longitudinal beam can be used as a reinforcement for a (preferably composite) panel formed by connecting the base or flange of the longitudinal beam to the panel. The proposed longitudinal beam comprises a structural cross-section with a Y-shape. This shape is designated as Y-shape in this specification and can correspond to all or part of the structural cross-section. The proposed longitudinal beam is denoted as Y-shaped longitudinal beam in this specification. The Y-shaped longitudinal beam according to the present invention comprises two structural elements: an open triangular cross-section structure followed by a longitudinal beam web providing a Y shape. Y-shaped longitudinal beams or Y-section longitudinal beams made of composite materials have various benefits, such as better structural efficiency, less wrinkles during manufacturing due to large angles, no corrosion and better non-destructive testing, NDT processes.

[0010] The Y-shaped stringers may have various shapes and may include vertical wings and caps with different variations. The use of composite materials for this purpose enhances the actual stringer performance as well as inspection and repairability.

[0011] The Y-shaped stringer according to the invention provides an optimization of the "double T" section stringers and "Omega" section stringers existing in the prior art. Advantageously, the proposed Y-shaped stringer increases the moment of inertia, enhances the buckling and post-buckling behavior and the torsion behavior, and provides a new shape that produces the most effective shape. In addition, the Y-shaped section optimizes the arrangement of the stringers due to the wider effective surface of the legs, which means weight reduction, cost savings and higher ROI.

[0012] Furthermore, the invention proposes another manufacturing method for obtaining a double Y-shaped cross-section spar by connecting two Y-shaped longerons obtained by the first manufacturing method by means of a longeron web or, when the Y-shaped longeron comprises a cap, by connecting the cap of the Y-shaped longeron to the second cap of the second Y-shaped longeron.

[0013] Furthermore, the invention proposes a method for manufacturing a torsion box of an aircraft, for example a wing torsion box comprising the manufactured double Y-shaped cross-section spar. The manufactured torsion box can be used for a vertical tail "VTP", a horizontal tail "HTP", etc. The proposed torsion box manufacturing method allows replacing conventional wing spars with the proposed composite wing spars, which can be lighter than conventional wing spars. Furthermore, when replacing conventional wing spars with double Y-shaped cross-section spar according to the invention to obtain the same mechanical properties, the number of wing spars used in the torsion box can be reduced.

[0014] A first aspect of the present invention relates to a method for manufacturing a Y-shaped longitudinal beam made of a composite material, the Y-shaped longitudinal beam comprising: a longitudinal beam web having an I-shaped cross-section; a lower flange; a first open triangular cross-sectional structure, the first open triangular cross-sectional structure comprising a lower vertex connected to the lower flange and an upper vertex connected to a first end of the longitudinal beam web, wherein the first open triangular cross-sectional structure and the longitudinal beam web form a Y-shaped cross-section, the method comprising: placing a composite material on a first mold and a second mold, the first mold and the second mold each comprising a molding curvature having an angle β=360°-α, wherein α is a working angle formed in a connection portion between the first open triangular cross-sectional structure and the lower flange, wherein α has a value between 100° and 165°; placing the composite material on a third mold; closing the first mold and the second mold using the third mold to obtain a closed mold containing a Y-shaped preform; and curing the Y-shaped preform using an autoclave cycle to obtain the Y-shaped longitudinal beam.

[0015] In one example, the method further includes cutting the Y-shaped stringer.

[0016] In one example, curing the Y-shaped preform using an autoclave cycle includes curing the Y-shaped preform at a temperature of 180 degrees Celsius or less.

[0017] In one example, the method further includes placing a rowing between the first mold and the second mold.

[0018] In one example, the method further comprises obtaining the first mold, the second mold and the third mold by 3D printing.The materials used are thermoplastic materials from the PEI / PEEK / PEKK family because of their good performance at high temperatures and their good resistance.

[0019] In one example, the method further includes establishing a lower flange in a direction perpendicular to the stringer web using shapes of the first mold, the second mold, and the third mold.

[0020] In one example, the method further includes using carbon fibers with thermosetting or thermoplastic resins or using glass fibers with thermosetting or thermoplastic resins as a composite material.

[0021] In one example, the method further includes connecting a cap to the second end of the stringer web, the cap providing the stringer web with a T-shaped, L-shaped, or J-shaped cross-section.

[0022] In one example, the cap comprises a symmetrical laminate consisting of:

[0023] a first layer comprising a composite laminate wrapped around a lower flange, a stringer web, wherein 70% of the total composite laminate from the first layer is oriented at a load angle of + / - 45°, wherein the load angle is the angle of the load relative to a reference axis of the Y-shaped stringer;

[0024] a second layer comprising a composite laminate capable of being built on top of the first layer, wherein 70% of the total composite laminate from the first layer is oriented at a load angle of 0°;

[0025] A third layer comprising a composite laminate capable of being built on top of the second layer, wherein 70% of the total composite laminate from the third layer is oriented at a load angle of + / - 45°.

[0026] In one example, the method further includes creating a vertical wing in a direction parallel to the stringer web using shapes of the first mold, the second mold, and the third mold.

[0027] A second aspect of the invention relates to the use of a Y-shaped stringer according to the first aspect of the invention as a reinforcement for a panel of an aircraft by joining the lower flange to the panel of the aircraft by means of co-curing or co-bonding.

[0028] In a third aspect, the present invention relates to a method for manufacturing a double Y-shaped cross-section spar, wherein the double Y-shaped cross-section spar comprises a spar web, a lower spar flange, an upper spar flange, and a first open triangular cross-section spar structure and a second open triangular cross-section spar structure, the method comprising connecting a second end of the spar web of the Y-shaped longitudinal beam to a second longitudinal beam web of the second Y-shaped longitudinal beam obtained by the method of the first aspect.

[0029] In a fourth aspect, the present invention relates to a method for manufacturing a torsion box for an aircraft, the torsion box comprising a plurality of double Y-shaped cross-section spars and a first panel and a second panel obtained by the method according to the previous aspect, the method comprising: placing at least a plurality of double Y-shaped cross-section spars between a first mold, wherein the first mold is at least associated with the shape of the spar web and the lower spar flange and the upper spar flange; placing a second mold onto the first open triangular cross-section spar structure and the second open triangular cross-section spar structure; placing a composite material at least onto the second mold; closing the first mold and the second mold using a third mold to obtain a closed mold containing a torsion box preform, wherein the third mold is at least associated with the shape of the first panel of the torsion box of the aircraft and the shape of the second panel of the torsion box; and cyclically curing the torsion box preform using an autoclave to obtain the torsion box.

[0030] In an example, the method further comprises cutting the torsion box.

[0031] In one example, curing the torsion box preform using an autoclave cycle includes curing the torsion box preform at a temperature of 180 degrees Celsius or less. BRIEF DESCRIPTION OF THE DRAWINGS

[0032] For a better understanding of the above description and for the purpose of providing examples only, some non-limiting drawings are included which schematically depict practical implementations.

[0033] Figure 1 A first example of a Y-shaped longitudinal beam obtained by a first manufacturing method according to the invention is shown.

[0034] Figure 2 A second example of a Y-shaped longeron including vertical wings obtained by the first manufacturing method according to the invention is shown.

[0035] Figure 3 A third example of a Y-shaped stringer comprising vertical wings and a cap obtained by the first manufacturing method according to the invention is shown.

[0036] Figure 4 An example of a double Y-shaped cross-section spar obtained by the second manufacturing method according to the present invention is shown.

[0037] Figure 5 A mold used in a first manufacturing method for obtaining a Y-shaped longitudinal beam according to the present invention is shown.

[0038] Figure 6 A mould used in a third manufacturing method according to the invention for obtaining a torsion box is shown.

[0039] Figure 7 A Y-shaped stringer 100 is shown for use as a reinforcement for a panel 1010 of an aircraft in accordance with the present invention.

[0040] Figure 8 A torsion box comprising a plurality of double Y-shaped cross-section spars obtained by a third manufacturing method according to the present invention is shown. DETAILED DESCRIPTION

[0041] The Y-shaped stringer obtained by the first manufacturing method according to the present invention comprises three different parts: an open triangular cross-sectional structure which enhances the behavior against torsional loads; a flange connected to the open triangular cross-sectional structure which can be configured to be attached to a panel or skin of an aircraft; and an elongated portion or stringer web established in the normal direction of the flange which increases the moment of inertia and structural stiffness of the Y-shaped stringer according to the present invention.

[0042] The double Y-shaped cross-section spar obtained by the manufacturing method according to the present invention comprises two Y-shaped longerons connected by the web of the longeron.

[0043] Figure 1 A first example of a Y-shaped stringer 100 according to the invention is shown. The Y-shaped stringer 100 may be made of a composite material and may be used as a reinforcement for a panel of an aircraft. The panel 1010 may be made of a composite material. Figure 1 The Y-shaped stringer 100 shown in FIG. 1 includes a stringer web 110 having an I-shaped cross section, and a first open triangular cross-sectional structure 120 including lower vertices 120a, 120b connected or connected to lower flanges 130a, 130b. In this example, the lower flanges 130a, 130b are established in a direction perpendicular to or normal to the stringer web 110. The lower flanges 130a, 130b are configured to be connected to a panel of an aircraft.

[0044] In addition, the first open triangular cross-sectional structure 120 includes an upper vertex 120c connected to the longitudinal beam web 110 having an I-shape. Figure 1 As shown in FIG. 1 , the first open triangular cross-sectional structure 120 is connected to the longitudinal beam web 110 to form a Y shape.

[0045] Figure 2 A second example of a Y-shaped stringer 100 according to the present invention is shown. In this example, the Y-shaped stringer 100 includes a vertical wing 140 connected to the lower flanges 130a, 130b and established in a direction parallel to the stringer web 110. The vertical wing 140 can improve the buckling resistance and post-buckling resistance and stiffness of the Y-shaped stringer 100.

[0046] Figure 3 A third example of a Y-shaped stringer 100 according to the invention is shown. Figure 3 The Y-shaped longitudinal beam 100 shown in FIG. Figure 2 All elements of the examples, and in addition, Figure 3 The Y-shaped stringer 100 shown in FIG. 1 includes a cap 150 that provides a T-shape to the stringer web 110 .

[0047] Composite laminates must be oriented to achieve the best possible performance to avoid buckling and prevent warping after manufacturing (laminates must be symmetrical and balanced). The improvement in laminate performance is related to the orientation of the fiber angles. That is, if the load is traveling in the 0° direction according to the Y-shaped stringer reference axis, a larger percentage of laminates will be placed at 0° relative to the Y-shaped stringer reference axis because this is where the laminate will have the best mechanical properties. The total percentage of laminates is divided into laminates at 0°, laminates at 90°, and laminates at + / -45°. Therefore, since different parts of the Y-shaped stringer 100 (lower flanges 130a, 130b, stringer web 110 and cap 150) work in different ways, each part must have a larger amount of laminates in its orientation of interest to obtain the best mechanical response. An example of a symmetrical and balanced composite laminate could be: +45°, -45°, 0°, 90°, 0°, 0°, 90°, 0°, -45°, +45°, which could also be written as (+45, -45, 0, 90, 0)S).

[0048] Graphite fiber reinforced polymer CFRP consists of several layers. Each layer is a "cloth", "tape" or "strip" of carbon fibers (like a piece of cloth). In this "fabric", the fibers can be disordered, woven or oriented mainly in one direction. Usually, those fibers with a main orientation are used, because in this main direction, we know that the fibers have resistance to significant loads. Then, when the layers are stacked, each layer can be oriented as desired to adapt the load resistance of the final piece according to the orientation of each layer of its layer.

[0049] The cap 150 can be divided into three stacked layers:

[0050] a first layer comprising a composite laminate that wraps around the lower flanges 130a, 130b, the stringer web 110, and the cap 150, wherein 70% of the total composite laminates from the first layer are oriented at a load angle of + / - 45°, wherein the load angle is the angle of the load relative to a reference axis of the Y-shaped stringer;

[0051] a second layer comprising a composite laminate that may be built on top of the first layer, wherein 70% of the total composite laminate from the first layer is oriented at a load angle of 0°;

[0052] A third layer comprising a composite laminate may be built on top of the second layer, wherein 70% of the total composite laminate from the third layer is oriented at a load angle of + / - 45°.

[0053] Thus, the first and third layers may include composite laminates oriented at a maximum of + / - 45°, where + / - 45° is the angle of the load relative to the Y-shaped stringer reference axis, and have 70% of the total composite laminates coming from the first and third layers (because it is interesting to have more composite laminates oriented at + / - 45° to avoid local and global buckling and due to design rules and laminate theory).

[0054] The second layer will have a maximum of 70% of the total composite laminate of the second layer oriented at 0°, where 0° is the angle of the load relative to the reference axis of the Y-shaped stringer, because it is important for the cap to have a core with a 0° orientation as this will be the orientation of the primary load (which will be perpendicular to the plane of the section).

[0055] The third example of the Y-shaped stringer 100 is the most complete Y-shaped stringer section, which can improve the behavior of the Y-shaped stringer 100 in the main load direction, thereby enhancing the moment of inertia and the elastic modulus E, increasing the stiffness of the stringer, and thus increasing the stiffness of the panel.

[0056] Figure 4 A first example of a double Y-shaped cross-section spar 300 obtained by the manufacturing method according to the present invention is shown. The double Y-shaped cross-section spar 300 can be made by connecting the Y-shaped stringer 100 to a second Y-shaped stringer 200 including a stringer web 210, and wherein the second end of the stringer web 110 is connected to the second stringer web 210 of the second Y-shaped stringer 200.

[0057] The double Y-shaped cross-section spar 300 includes a spar web obtained by connecting a longeron web 110 and a second longeron web 210 , both webs having an I-shaped cross-section.

[0058] The double Y-shaped cross-section spar 300 includes two parts connected by a spar web, ie, a lower part corresponding to the Y-shaped longeron 100 and an upper part corresponding to the second Y-shaped longeron 200 .

[0059] The lower portion corresponding to the Y-shaped longitudinal beam 100 includes lower flanges 130a, 130b, which can be configured to be connected to the first panel 1010 (such as the torsion box of the aircraft) of the aircraft. Figure 8 a first open triangular cross-sectional structure 120, the first open triangular cross-sectional structure 120 includes a first lower vertex 120a and a second lower vertex 120b connected to the lower flanges 130a, 130b, respectively, and an upper vertex 120c connected to the first end of the longitudinal beam web 110.

[0060] The upper portion corresponding to the second Y-shaped longitudinal beam 200 includes: upper flanges 230a, 230b, which are configured to be connected to the second panel 1020 of the torsion box of the aircraft; and a second open triangular cross-sectional structure 220, which is connected to the upper flanges 230a, 230b and to the second longitudinal beam web 210.

[0061] The first open triangular cross-sectional structure 120 , the second open triangular cross-sectional structure 220 , and the stringer web 110 and the second stringer web 210 form a double Y-shaped cross-sectional shape.

[0062] The double Y-shaped cross-section spar 300 is the most suitable example for production according to the invention as this will still have the best structural behaviour and can be easily mass produced using 3D printing technology with rapid prototyping moulds.

[0063] Figure 5 A mold used in the method for manufacturing the Y-shaped longitudinal beam 100 according to the present invention is shown.

[0064] Figure 5 A first mold A and a second mold B are shown, which include a molding curvature X with an angle β=360°-α, where α is a working angle formed in the connection between the first open triangular cross-sectional structure 120 of the Y-shaped longitudinal beam 100 and the lower flanges 130a, 130b, where α has a value between 100° and 165°.

[0065] The angles β and α may be adjusted to optimize load transfer, but to ensure that wrinkles are not formed during the manufacturing process of the Y-shaped stringer 100. The optimum working angle may be, for example, 135°.

[0066] also, Figure 5 A third mold C is shown, which closes the first mold A and the second mold B to obtain a Y-shaped preform 100a. The Y-shaped preform 100a is composed of laminates, and strips D2 may be placed between the laminates to ensure that no voids are created in the joining portions of the laminates.

[0067] The manufacturing process of the Y-shaped longitudinal beam 100 may include: placing a composite material on a first mold A and placing the composite material on a second mold B, the first mold A and the second mold B including a molding curvature X having an angle β=360°-α; and placing the composite material on a third mold C; and closing the first mold A and the second mold B using the third mold C to obtain a closed mold containing a Y-shaped preform 100a; and curing the Y-shaped preform 100a using an autoclave cycle to obtain the Y-shaped longitudinal beam 100.

[0068] Additionally, the method may include cutting the Y-shaped stringer 100 .

[0069] In one example, the cheapest manufacturing option for the Y-shaped stringer 100 and the most suitable option for production would be: a stringer without vertical wings 140 and a stringer web 110 with an I-shape without a cap 150. This Y-shaped stringer 100 may have the best structural behavior and may be easily mass-produced using 3D printing technology with rapid prototyping molds.

[0070] In a specific example, the manufacture of the Y-shaped longitudinal beam 100 includes: manufacturing a first mold A, a second mold B, and a third mold C using 3D printing; performing a fiber placement process of pre-impregnated fibers or dry fibers that can be infused with resin above the first mold A and the second mold B; closing the first mold A and the second mold B using the third mold C to conform the Y-shaped preform 100a; using an autoclave cycle at 180° or lower to optimize the use temperature of the Y-shaped longitudinal beam 100, and a final additional step of cutting the Y-shaped longitudinal beam 100.

[0071] Figure 6 A torsion box 2000 for an aircraft is shown. Figure 8 The torsion box comprises a plurality of double Y-shaped cross-section spars 300 and a first panel 1010 and a second panel 1020 obtained by the method according to the previous aspect.

[0072] The torsion box 2000 comprises a plurality of double Y-shaped cross-section spars 300 obtained by the manufacturing method according to the present invention. The double Y-shaped cross-section spars 300 comprise a spar web, a lower spar flange, an upper spar flange, and a first open triangular cross-section spars structure and a second open triangular cross-section spars structure.

[0073] The manufacture of the torsion box 2000 may include placing a plurality of double Y-shaped cross-section spars 300 between a first mold E, wherein the first mold E is associated with at least the shape of the spar web and the lower and upper spar flanges.

[0074] The manufacture of the torsion box 2000 may also include placing a second mold F onto the first open triangular cross-section spar structure and the second open triangular cross-section spar structure, and placing a composite material onto at least the second mold F. Placing the composite material may include a fiber placement process of pre-impregnated fibers or dry fibers over the mold, which may be later infused with resin.

[0075] The manufacture of the torsion box 2000 may also include using a third mold G to close the first mold E and the second mold F to obtain a closed mold containing the torsion box preform 300a, wherein the third mold G is at least associated with the shape of the first panel 1010 of the torsion box of the aircraft and the shape of the second panel 1020 of the torsion box, and the torsion box preform is cured using an autoclave cycle.

[0076] The method further comprises cutting the torsion box 2000 .

[0077] In the proposed method, curing the torsion box preform 300a using an autoclave cycle includes curing the torsion box preform at a temperature of 180 degrees Celsius or less.

[0078] In one example, α is equal to 135°.

[0079] The first mold E and the second mold F conform the spar web. The second mold F and the third mold G conform the open triangular cross-section spar structure to the first panel 1010 of the torsion box 2000 and the second panel 1020 of the torsion box 2000.

[0080] α is variable and can be adjusted to optimize load transfer but ensure that wrinkles are not formed during manufacturing.

[0081] Figure 7 The Y-shaped longitudinal beam 100 according to the present invention is shown, which is formed by connecting the lower flanges 130a, 130b of the Y-shaped longitudinal beam 100 to the Figure 7 The composite panel 1010 of the aircraft shown in FIG. 1 is used as a reinforcement for the panel 1010 of the aircraft. The panel 1010 may be a composite panel.

[0082] The joining of the lower flanges 130a, 130b of the Y-shaped stringer 100 to the panel 1010 of the aircraft may include co-curing in a first alternative or co-bonding in a second alternative.

[0083] It is important to note that this shape still allows the use of ratholes 160 without affecting the structural behavior while reducing weight. Ratholes 160 are holes made in the ribs of the aircraft structure to help locate and assemble the wing structure. These holes also allow for the interconnection of the fuel tanks.

[0084] Figure 8 An example torsion box 2000 including a double Y-shaped cross-section spar 300 according to the present invention is shown. Figure 8 A possible arrangement of a wing torsion box is shown showing first and second panels 1010 , 1020 of the torsion box, front and rear spars and a number of double Y-cross section spars 300 which will vary depending on the size of the torsion box 2000 of the aircraft.

Claims

1. A method for manufacturing a Y-shaped longitudinal beam (100), wherein the Y-shaped longitudinal beam (100) is made of a composite material, wherein the Y-shaped longitudinal beam (100) comprises: A longitudinal beam web (110), wherein the longitudinal beam web (110) has an I-shaped cross section; a lower flange (130a, 130b); a first open triangular cross-sectional structure (120), the first open triangular cross-sectional structure (120) comprising a first lower vertex (120a) and a second lower vertex (120b) respectively connected to the lower flanges (130a, 130b) and an upper vertex (120c) connected to the first end of the longitudinal beam web (110), wherein the first open triangular cross-sectional structure (120) and the longitudinal beam web (110) form a Y-shaped cross-section, the method comprising: Placing the composite material onto a first mold (A) and a second mold (B), each of which comprises a molding curvature (X) having an angle β=360°-α, wherein α is a working angle formed in the connection between the first open triangular cross-sectional structure (120) and the lower flange (130a, 130b), wherein α has a value between 100° and 165°; placing the composite material onto a third mold (C); closing the first mold (A) and the second mold (B) using the third mold (C) to obtain a closed mold containing a Y-shaped preform (100a); and The Y-shaped preform (100a) is cured by autoclave cycle to obtain the Y-shaped longitudinal beam (100).

2. The method of claim 1, further comprising cutting the Y-shaped stringer (100).

3. The method according to claim 1 or 2, wherein: Curing the Y-shaped preform (100a) using an autoclave cycle includes curing the Y-shaped preform (100a) at a temperature of 180 degrees Celsius or less.

4. The method according to claims 1 to 3, further comprising placing a strip (D1, D2) between the first mould (A) and the second mould (B).

5. The method according to claims 1 to 4, further comprising: The first mold (A), the second mold (B) and the third mold (C) are obtained by 3D printing.

6. The method according to claim further comprises establishing the lower flange (130a, 130b) in a direction perpendicular to the longitudinal beam web (110) using shapes of the first mold (A), the second mold (B) and the third mold (C).

7. The method according to any one of the preceding claims, further comprising: Carbon fibers and thermosetting or thermoplastic resins or glass fibers and thermosetting or thermoplastic resins are used as composite materials.

8. The method according to the preceding claim further comprises using the shapes of the first mold (A), the second mold (B) and the third mold (C) to create a vertical wing (140) in a direction parallel to the longitudinal beam web (110).

9. The method according to claims 1 to 7, further comprising connecting a cap (150) to the second end of the stringer web (110), the cap (150) providing the stringer web (110) with a T-shaped or L-shaped or J-shaped cross-section.

10. The method according to the preceding claim, wherein: The cap (150) comprises: a first layer comprising a composite laminate wrapping the lower flange (130a, 130b), the stringer web (110), wherein 70% of the total composite laminate from the first layer is oriented at a load angle of + / -45°, wherein the load angle is the angle of the load relative to a reference axis of the Y-shaped stringer; a second layer comprising a composite laminate capable of being built on top of the first layer, wherein 70% of the total composite laminate from the first layer is oriented at a load angle of 0°; A third layer comprising a composite laminate capable of being built on top of the second layer, wherein 70% of the total composite laminate from the third layer is oriented at a load angle of + / - 45°.

11. Use of a Y-shaped stringer (100) according to any one of the preceding claims as a reinforcement for a panel (1010) of an aircraft, the Y-shaped stringer being used as the reinforcement by joining the lower flange (130a, 130b) to the panel (1010) of the aircraft by means of co-curing or co-bonding.

12. A method for manufacturing a double Y-shaped cross-section spar (300), the double Y-shaped cross-section spar (300) comprising at least a spar web, a lower spar flange, an upper spar flange, and a first open triangular cross-section spar structure and a second open triangular cross-section spar structure, the method comprising: connecting a second end of a longitudinal beam web (110) of a Y-shaped longitudinal beam (100) to a second end of a second longitudinal beam web (210) of a second Y-shaped longitudinal beam (200), the Y-shaped longitudinal beam (100) and the second Y-shaped longitudinal beam (200) each being obtained by a method according to claims 1 to 8; Or the cap (150) of the Y-shaped stringer (100) is connected to a second cap of a second Y-shaped stringer (200), the Y-shaped stringer (100) and the second Y-shaped stringer (200) each being obtained by a method according to claims 8 to 9.

13. A method for manufacturing a torsion box (2000) for an aircraft, the torsion box (2000) comprising a plurality of double Y-shaped cross-section spars (300) obtained by the method according to the preceding claim and a first panel (1010) and a second panel (1020), the method comprising: placing at least the plurality of double Y-shaped cross-section spars (300) between a first mold (E), wherein the first mold (E) is associated with at least the shape of the spar web and the lower and upper spar flanges; placing a second mold (F) onto the first open triangular cross-section spar structure and the second open triangular cross-section spar structure; placing the composite material on at least the second mold (F); closing the first mold (E) and the second mold (F) by means of a third mold (G) to obtain a closed mold containing a torsion box preform (300a), wherein the third mold (G) is associated with at least the shape of the first panel (1010) of the torsion box (2000) of the aircraft and the shape of the second panel (1020) of the torsion box (2000), and The torsion box preform (300a) is cured by an autoclave cycle to obtain the torsion box (2000).

14. The method of claim 13, further comprising cutting the torsion box (2000).

15. The method according to claim 13 or 14, wherein: Curing the torsion box preform using an autoclave cycle includes curing the torsion box preform at a temperature of 180 degrees Celsius or less.

Images