Methods for manufacturing aerospace structural components

Abstract

The present invention provides a method for laminated manufacturing of an integral composite aerospace structural component, the method comprising: (a) providing an additive manufacturing tool including a deposition mold for shaping an aerodynamic surface and at least one head, the at least one head being configured to move on the deposition mold and configured to deposit a fiber material reinforcement and / or a fusible material; (b) depositing a fiber material reinforcement embedded in the fusible material onto the deposition mold, thereby forming at least one layer of a lower aerodynamic panel; (c) depositing the fusible material onto at least a portion of an outer layer of the lower aerodynamic panel, thereby forming at least one layer of a core structure; and (d) depositing a fiber material reinforcement embedded in the fusible material onto at least an outer layer of the core structure, thereby forming at least one layer of an upper aerodynamic panel; steps (b), (c), and (d) are performed using additive manufacturing technology.

Description

Technical Field

[0001] This invention belongs to the field of manufacturing, specifically the field of manufacturing aerospace structural components using additive manufacturing technology. More specifically, this invention has particular applications in the manufacturing of moving parts and control surfaces of aircraft. Background Technology

[0002] Movable components and control surfaces are the primary flight control structural elements through which the aircraft's attitude is controlled during takeoff, flight, and landing. Due to their role in the safe operation of the aircraft, movable components and control surfaces are designed using known manufacturing techniques to ensure compliance with high structural requirements of certification regulations. For example, failure to properly maintain the tolerances, dimensions, or balance of control surfaces can lead to deviations from normal aircraft operation.

[0003] In short, the standard moving parts and control surface design includes: main spars, a set of ribs, upper skin, and lower skin.

[0004] Historically, aircraft components were made of aluminum alloys. In recent decades, with the development of composite manufacturing technologies, these structural components have been manufactured using different techniques, such as co-bonding or co-curing of carbon fiber reinforced plastic (CFRP) components.

[0005] However, all these manufacturing technologies require that each component be manufactured independently at different stages of the operation, and then these different components be assembled together. This is a time-consuming process that predetermines the production rate of the aircraft. Therefore, the final component is achieved through a series of different manufacturing steps, which increases manufacturing costs and time.

[0006] Therefore, manufacturing in the aerospace industry is a labor-intensive, multi-step process that requires rapid, reliable, and efficient production. The combination of multiple steps and the number of final parts to be manufactured (end-use parts) leads to delays, resulting in a major drawback of the aerospace industry. Summary of the Invention

[0007] This invention provides a solution to the aforementioned problems through a method for manufacturing integral composite aerospace structural components using a laminated method and, preferably, for controlling the surface of the aerospace structural components. Preferred embodiments of the invention are defined in the dependent claims.

[0008] In a first aspect, the present invention provides a method for laminated manufacturing of an integral composite aerospace structural component, preferably, the aerospace structural component being a control surface, wherein the method comprises the following steps:

[0009] a) Provide an additive manufacturing tool comprising a deposition die and at least one head, the deposition die forming an aerodynamic surface, the at least one head being configured to move on the deposition die and deposit fibrous material reinforcements and / or fusible materials;

[0010] b) Deposit a fiber reinforcement embedded in a fusible material onto a deposition mold, thereby forming at least one layer of the lower aerodynamic panel;

[0011] c) Depositing a fusible material onto at least a portion of the outer layer of the lower aerodynamic panel, thereby forming at least one layer of the core structure; and

[0012] d) Depositing a fiber reinforcement embedded in a fusible material onto at least the outer layer of the core structure, thereby forming at least one layer of the upper aerodynamic panel;

[0013] Steps b), c), and d) are implemented using additive manufacturing technology.

[0014] Throughout this document, “additive manufacturing technology” (AM) should be understood as those techniques for creating three-dimensional objects by adding layers of material, wherein the material (in the case of reinforcing material, the material is a fusible material or a matrix material) transforms into a liquid when heated and solidifies (or hardens) into a solid when cooled.

[0015] Typically, these techniques utilize computers equipped with 3D modeling software (Computer-Aided Design or CAD), additive manufacturing tools (such as mechanical equipment), and laminated materials. A CAD sketch is a 3D electronic model of the final manufactured 3D object. AM tools are able to read data from CAD files (both cross-sectional geometry and surface patterns) and fabricate 3D objects by laying down or depositing continuous layers of liquids, powders, sheets, etc., in a laminated manner through at least one head.

[0016] Depending on the materials and mechanical techniques used, many technologies are included within additive manufacturing. Among these, selective laser sintering (SLS), stereolithography (SLA), multi-nozzle forming (MJM), 3D printing (3DP), or fused filament fabrication (FFF) should be noted.

[0017] In a preferred embodiment of the method of the present invention, steps b), c) and d) are implemented using fused wire fabrication (FFF).

[0018] Fused wire fabrication (FFF) is a process-oriented manufacturing process involving the use of material injected into a deposition mold through at least one indexing nozzle. The nozzle travels according to a surface pattern, which can be flat or curved, and for each specific layer, the material is hardened before the next layer is applied. This process is repeated until the manufactured article or part is completed. Preferably, the material is a thermoplastic or a fiber reinforcement embedded within a thermoplastic material.

[0019] Throughout this document, the outer layer should be understood as the last layer deposited in each of the steps performed, in accordance with the deposition sequence.

[0020] According to the present invention, the fiber material reinforcement may take the form of, for example, fibers (very short and / or irregular fibers), nanofibers, carbon fillers, short fibers (less than 1 mm in length), or continuous fibers (extending continuously along the entire filament and thus continuously along the entire length / width of the component during manufacturing). Furthermore, the fiber material reinforcement may be glass, carbon, polymer fibers, or any other conventional material used as a reinforcement.

[0021] According to the present invention, the meltable material can be a thermoplastic material, such as PA (polyamide), PEEK (polyetherketone), PAEK (polyaryletherketone), or PEKK (polyetherketone), which can be unreinforced or reinforced using short fibers. In a preferred embodiment, the meltable material is in the form of filaments for better storage and handling.

[0022] In a preferred embodiment, the meltable material is any of the following thermoplastic materials: PEKK, PAEK, or PEEK. More preferably, the meltable material is PAEK or PEEK. And even more preferably, the meltable material is PEEK.

[0023] Lightweight design is achieved by synergistically combining fiber reinforcement and meltable material deposition, as less material is required to meet structural requirements compared to using meltable material alone.

[0024] In a preferred embodiment, the fiber reinforcement is deposited during operation to better adapt to future load requirements. This results in an optimized aerospace structure with structural elements that more precisely match the primary load path. Furthermore, this allows for weight savings due to the optimization of the additional internal reinforcing elements.

[0025] In summary, the laminated aerospace structural component manufactured according to the method of the present invention has the following characteristics:

[0026] ■ High-resolution surfaces, typically with a resolution of 0.1 mm for fusible materials and a resolution of 0.2 mm for fiber reinforcements embedded in fusible materials;

[0027] ■ Once the method is completed, there is no interface between the aerodynamic panel and the core structure, because the fusible material acts as an adhesive to connect the core structure and the panel.

[0028] This final, seamless feature offers additional advantages, such as improved water ingress protection and debonding (unlike conventional composite manufacturing techniques that adhesively bond the core to the panel), because the aerospace structural component is entirely monolithic (the final product is a self-sealing structure). Furthermore, the improved structural performance of the final aerospace structural component enhances operability.

[0029] Regarding additive manufacturing tools, different types and numbers of heads can be used depending on the embedded fibrous material reinforcement. According to the invention, the additive manufacturing tool may have one head configured to deposit a fusible material and another head configured to deposit a fibrous material reinforcement embedded within the fusible material, or it may have only one head configured to switch the material to be deposited. Furthermore, the fibrous material reinforcement may be embedded in the following ways:

[0030] ■ Before entering the head (similar to pre-impregnated composite materials), or

[0031] ■In a suitable position in the head, or

[0032] ■ In a suitable location on the component, the fiber reinforcement and the fusible material are deposited independently, and once deposited, heat is applied to melt the fusible material.

[0033] Advantageously, the method according to the first aspect of the invention allows for simpler tooling compared to conventional composite manufacturing technologies that require autoclaves, vacuum packaging, finishing, etc. Therefore, non-recurring cost savings are achieved. Furthermore, recurring costs are also saved because the buy-to-fly ratio (which is understood as the ratio between the mass of material required to produce the component and the mass of material in the finished aerospace structural component) is reduced compared to existing composite material manufacturing technologies. This is because there is virtually no material waste in the method of the present invention.

[0034] Preferably, in steps b) and d) of the method according to the invention, the head is configured to move on the deposition mold along three translational axes (X, Y, Z) and / or rotational axes (about X, Y, Z). Thus, the upper and lower aerodynamic panels are formed of layers not limited to the horizontal plane XY, and therefore have an external aerodynamic shape. However, in step c), the head preferably moves only in the horizontal direction (X, Y), and therefore the core structure substantially reproduces the external shape of the aerodynamic surface through a set of vertically stacked, generally flat layers (i.e., horizontal layers in the XY plane).

[0035] In a particular embodiment, between step c) and step d), the method further includes covering at least one edge of at least one layer of the core structure by depositing a fiber material reinforcement embedded in a fusible material onto at least a portion of the outer layer of the lower aerodynamic panel and the core structure.

[0036] This step produces at least one structural element, which is a spanwise stiffening element commonly referred to as a spar, which is also integral with the rest of the structural component because the spar is connected to both the panel and the core by a fusible material.

[0037] Advantageously, at least one structural element at at least one edge of at least one layer covering the core structure eliminates the step effect of the core structure (in the case of more than one layer being deposited). The step effect is a result of approaching the surface at a certain angle during manufacturing, where the layer thickness is the height of the step. Therefore, the mechanical properties of the final aerospace structure in the Z direction, i.e., the direction approximately perpendicular to the deposition mold, are also improved.

[0038] In a particular embodiment, in the steps between step c) and step d), the fiber reinforcement comprises continuous fibers. Advantageously, this allows for better load-bearing performance of aerospace structural components due to the high strength and stiffness of the continuous fibers embedded within the fusible material. The fusible material serves as the connecting matrix, thereby allowing the transfer of pressure and shear stress between the fibers.

[0039] Furthermore, since the fiber reinforcement comprises continuous fibers, it may be oriented during the building process to better accommodate future load requirements during operation.

[0040] In a preferred embodiment, in the steps between step c) and step d), all the fiber material reinforcements are continuous fibers.

[0041] In a particular embodiment, the fiber reinforcement deposited in step b) comprises continuous fibers. Advantageously, this improves the load-bearing capacity of the aerospace structural component and reduces warping effects, i.e., reduces deformation of the component due to the shrinkage of the top layer, thereby maintaining its aerodynamic shape.

[0042] In a particular embodiment, the fiber material reinforcement deposited in step d) comprises continuous fibers.

[0043] Similarly, this improves the load-bearing capacity of aerospace structural components and maintains their aerodynamic shape.

[0044] In a particular embodiment, step c) includes depositing a fibrous material reinforcement embedded within a fusible material. Preferably, the fibrous material reinforcement is a short fiber.

[0045] In a particular embodiment, in step c), a hollow space with or without material is formed during the deposition of the fusible material or the fiber reinforcement embedded within the fusible material. This results in a non-uniform core structure. Advantageously, this allows for more precise structures, taking into account future load paths during operation. Furthermore, it allows for weight savings. In this embodiment, step c) can be performed at the beginning and end of the deposition and / or at a non-constant volumetric flow rate.

[0046] In certain embodiments, the density of the non-uniform core structure varies along the wingspan and / or chord length of the aerospace structure.

[0047] In certain embodiments, the core structural component comprises a fragmented geometry or a honeycomb geometry. Advantageously, this allows for high-performance core geometries, thereby enabling improved acoustic and shock-resistant applications for aerospace structural components.

[0048] In a particular embodiment, the structural element produced by the additional step between step c) and step d) is a beam-shaped structural element, which is preferably a wing beam.

[0049] In a particular embodiment, the beam-shaped structural element is a Z-shaped wing beam, and:

[0050] ■ The deposition on at least a portion of the outer layer of the lower aerodynamic panel forms the lower leg of the Z-shaped spar, and

[0051] ■ The deposition on the edge of at least one layer of the core structural member forms the web of the Z-shaped wing beam.

[0052] In a particular embodiment, the deposition on the core structure also establishes the upper foot of the Z-shaped spar, which is configured to form a continuous surface together with the outer layer of the core structure.

[0053] In a particular embodiment, the web of the Z-shaped spar is straight and angled to the upper and lower aerodynamic panels, the angle corresponding to the stepped portion of the core structure.

[0054] In a particular embodiment, the connection between the lower leg of the Z-shaped spar and the web and / or between the upper leg of the Z-shaped spar and the web includes a radius of curvature.

[0055] In a second aspect of the invention, the present invention provides an aerospace structural component manufactured using a method according to any embodiment of the first aspect of the invention, the aerospace structural component being preferably a control surface.

[0056] All features and / or steps of the methods described in this specification (including the claims, description and drawings) may be combined in any combination except for combinations of mutually exclusive features and / or steps. Attached Figure Description

[0057] These and other features and advantages of the invention will be clearly understood in light of the detailed description thereof, and will become apparent from the preferred embodiments thereof with reference to the accompanying drawings, which are merely examples and are not intended to limit the invention.

[0058] Figure 1 The figure shows an additive manufacturing tool.

[0059] Figure 2 illustrates a method for manufacturing integral composite aerospace structural components according to the present invention.

[0060] Figure 3 illustrates the method of Figure 2, which is used for the laminated fabrication of a monolithic composite aerospace structural component, which additionally includes internal structural elements.

[0061] Figure 4 Figures 4a to 4b An integral composite aerospace structural component manufactured by the method shown in Figure 2 or Figure 3 is illustrated. This aerospace structural component includes a non-uniform core structural component. Figures 4c to 4g An example of a non-uniform core structure is shown.

[0062] Figure 5 illustrates the different advantages obtained when depositing fiber reinforcements embedded in fusible material. Detailed Implementation

[0063] This invention defines a method for manufacturing a single-piece composite aerospace structural component 10 using a laminated method. The method includes the following steps:

[0064] a) Provide an additive manufacturing tool 20 comprising a deposition mold 23 and at least one head 24, the deposition mold 23 forming an aerodynamic surface, the at least one head 24 being configured to move on the deposition mold 23 and to deposit a fibrous material reinforcement embedded in a fusible material and / or the fusible material; provide the fibrous material reinforcement and provide the fusible material, which is preferably a thermoplastic material or a resin, more preferably a thermoplastic material;

[0065] b) Deposit the fiber material reinforcement embedded in the fusible material onto the deposition mold 23, thereby forming at least one layer of the lower aerodynamic panel 11;

[0066] c) Depositing a fusible material onto at least a portion of the outer layer of the lower aerodynamic panel 11, thereby forming at least one layer of the core structure 12; and

[0067] d) Deposit fiber reinforcement embedded in a fusible material onto at least the outer layer of the core structure 12, thereby forming at least one layer of the upper aerodynamic panel 14;

[0068] Steps b), c), and d) are performed using additive manufacturing technology.

[0069] Figure 1 An embodiment of an additive manufacturing tool 20 that can be used in the method of the present invention is shown. The additive manufacturing tool 20 includes:

[0070] ■Spool 25, which stores fusible material or fiber reinforcement embedded in fusible material, preferably, both fusible material and fiber reinforcement embedded in fusible material are in the form of filaments;

[0071] ■ Deposition mold 23, which shapes the outer aerodynamic surface of the lower aerodynamic panel 11 or the housing; and

[0072] ■ At least one head 24 configured to move on a deposition mold 23 and simultaneously configured to deposit material stored in a roll 25, the at least one head 24 traveling according to the surface pattern of each particular layer, wherein the material is hardened before the application of subsequent layers.

[0073] In one embodiment, the head 24 is configured to move on the deposition mold along three translational axes (X, Y, Z) and / or rotational axes (about X, Y, Z). Alternatively, the head 24 may also be configured to move only in the horizontal directions (X, Y) on the deposition mold, while movement in the vertical Z direction is performed by the deposition mold 23, thereby achieving so-called 2.5D (two-and-a-half-dimensional) manufacturing. The movement of the head 24 and / or the deposition mold can be achieved by actuators and / or servos, with one actuator and / or servo provided on each direction and / or rotational axis.

[0074] Material stored in the roll 25 is guided to the head 24. The head 24 additionally includes an extruder 24.1, which uses a torque and clamping system to feed and retract the fed material to drive the desired amount of material to be deposited. The head 24 also includes a heater assembly 24.2 for heating the molten material to a precise temperature. Once the material is heated, it is forced out of the nozzle 24.3, which has a gradually decreasing diameter, thus allowing for more precise material deposition.

[0075] In this embodiment, the heater component 24.2 is arranged directly in the nozzle 24.3.

[0076] For each type of material to be deposited, even if the fusible material is the same, different temperatures may be required for melting depending on the form or size of the material to be deposited. In other words, different temperatures may be required to melt the fiber reinforcement embedded in the fusible material, or only to melt the fusible material, and therefore, more than one head may be used. For illustrative purposes, only one head 24 is shown.

[0077] Optionally, a head 24 may include two receptacles, each having a nozzle 24.3, one receptacle for driving the molten material and the other receptacle for driving a fiber reinforcement embedded in the molten material, or for driving only the fiber reinforcement.

[0078] Method for laminated manufacturing of integral composite aerospace structural components 10

[0079] Figure 2 schematically illustrates an embodiment of the method for manufacturing an integral composite aerospace structural component 10 according to the present invention.

[0080] Once the additive manufacturing tool 20 and the required materials are provided, a fibrous material reinforcement embedded in the fusible material can be deposited or laid onto the deposition mold 23. This fibrous material reinforcement is preferably a continuous fiber in filament form embedded in the fusible material. This step... Figure 2a It is shown schematically in the middle.

[0081] The process is performed as the head 24 moves along three translational axes (X, Y, Z) and / or rotational axes (each about axes X, Y, Z) following the shape of the aerodynamic surface to be formed. The nozzle 24.3 deposits each layer 2.1 of the fibrous material reinforcement embedded within the fusible material, wherein the material is hardened before applying the next layer 2, and this step is repeated until all the required layers 2.1 are deposited and the lower aerodynamic panel 11 is completely formed. For illustrative purposes, Figures 2a to 2c Only one layer 2.1 of the fiber material reinforcement embedded in the fusible material is shown.

[0082] Arrows are drawn throughout Figures 2 and 3 to indicate the position of the head at each step. This is for illustrative purposes only.

[0083] Once the lower aerodynamic panel 11 is fabricated, fusible material is deposited onto at least a portion of the outer layer of the lower aerodynamic panel 11, thereby forming at least one layer of the core structure 12. Figure 2b This step, schematically illustrated, is performed as the head 24 moves along two translational axes (X, Y), thus performing 2.5D fabrication. The nozzle 24.3 deposits each layer 1 of fusible material, wherein the material is hardened before the next layer 1 is applied, and this process is repeated until all required layers 1 are deposited and the core structure 12 is completely formed. Since the layers 1 are deposited or stacked vertically in step b), a step effect may occur to approximate these angles, where the layer thickness is the height of the step. Therefore, the core structure roughly reproduces the external shape of the aerodynamic surface through a set of vertically stacked, generally flat layers (i.e., horizontal layers along the X, Y directions).

[0084] As will be depicted in more detail in Figure 4, performing this step may leave material-free hollow spaces 12.2 during deposition. Therefore, the resulting core structure 12 may be non-uniform or may include fragmented and / or honeycomb geometries.

[0085] Once the core structure 12 is fabricated, a fiber reinforcement embedded in the fusible material is deposited on at least the outer layer of the core structure 12. The fiber reinforcement embedded in the fusible material is preferably a continuous fiber in the form of a filament embedded in the fusible material. Figure 2c The step is illustrated schematically as follows: Figure 2aThe process is performed as described, wherein the head 24 moves along three translational axes (X, Y, Z) and / or rotational axes (around X, Y, Z) to deposit a layer 2.2 of fiber material reinforcement embedded in the fusible material onto layer 1 of the core structure 12 to reproduce the shape of the aerodynamic surface of the upper aerodynamic panel 14. For illustrative purposes, only layer 2.2 of the fiber material reinforcement embedded in the fusible material is shown.

[0086] The nozzle deposits each layer 2.2, wherein the material is hardened before the next layer 2.2 is applied, and this step is repeated until all the required layers 2.2 are deposited and the upper aerodynamic panel 14 is completely formed.

[0087] By performing the described steps, an aerospace structural component 10 is thus manufactured. The aerospace structural component preferably controls surfaces comprising two main directions: the spanwise direction, i.e., from the wing root to the wingtip; and the chord direction, i.e., from the leading edge to the trailing edge.

[0088] Advantageously, since the head 24 deposits material according to the shape of the components of the aerospace structural member 10 (rather than depositing horizontally as in 2.5D manufacturing), a staircase effect is avoided in the formation of the upper and lower aerodynamic panels. Furthermore, since the aerospace structural member 10 is manufactured close to its net shape, no final finishing is required to achieve the final dimensions.

[0089] Furthermore, this process enables the manufacture of the final aerospace structural component 10, avoiding repetition at different manufacturing stages. Therefore, this process can be understood as a one-shot process, because the manufactured structural component is final once it is removed from the additive manufacturing tool 20.

[0090] Figure 3 illustrates the method as described with respect to Figure 2, wherein the aerospace structural component 10 additionally includes internal structural elements 13. In this embodiment, respectively in Figure 3a and Figure 3b The construction of the lower aerodynamic panel 11 and core structural member 12, schematically depicted in the diagram, is related to... Figure 2a and Figure 2b The steps described are the same. In this embodiment, an additional step is performed, according to which a fiber material reinforcement embedded in a fusible material is deposited on at least a portion of the outer layer of the lower aerodynamic panel 11 and the core structure 12. The fiber material reinforcement embedded in the fusible material is preferably a continuous fiber in the form of a filament embedded in the fusible material.

[0091] Therefore, at least one edge 12.1 of at least one layer 1 of the core structure 12 is covered by continuous fibers embedded in the fusible material. In the presence of multiple edges 12.1, these edges 12.1 can form the sides of the core structure 12. Figure 3b In the middle, the edge 12.1 of the core structural member 12 is shown with a dashed line.

[0092] The process is performed as described with respect to steps a) and c) in Figure 2. The head 24 moves along three translational axes (X, Y, Z) and / or rotational axes (about X, Y, Z), thereby depositing a layer 2.3 of the fiber reinforcement embedded within the fusible material onto at least a portion of the lower aerodynamic panel 11, and at least partially covering the edge 12.1 of at least one layer 1 of the core structural member 12, thereby forming the structural element 13. This step is schematically depicted in... Figure 3c In the middle. For illustrative reasons, only one layer 2.3 of the fiber material reinforcement embedded in the fusible material is shown.

[0093] The nozzle 24 deposits each layer 2.3, wherein the material is hardened before the next layer 2.3 is applied, and this step is repeated until all the required layers 2.3 are deposited and the structural element 13 is completely formed. Preferably, the formed structural element 13 is a beam-shaped structural element.

[0094] Advantageously, at least one structural element 13 of at least one edge 12.1 of at least one layer covering the core structure 12 eliminates the step effect of the core structure 12.

[0095] Figure 3c The beam-shaped structural element 13 shown is a Z-shaped wing beam, which includes:

[0096] ■ Lower foot 13.1, which is formed by depositing a fiber material reinforcement 2.3 embedded in a fusible material onto at least a portion of the outer layer of the lower aerodynamic panel 11;

[0097] ■ Web 13.2, which is formed by depositing a fiber reinforcement 2.3 embedded in a fusible material onto the edge 12.1 (or side) of at least one layer (or several layers) of the core structure 12; and

[0098] ■ Upper foot 13.3, which is formed by depositing a fiber material reinforcement 2.3 embedded in a fusible material onto a portion of the core structure 12. The upper foot 13.3 forms a continuous surface together with the outer layer of the core structure 12, as in Figure 3c This can be observed in the middle.

[0099] The web 13.2 of the Z-shaped spar 13 can be straight and angled to the upper aerodynamic panel 14 and the lower aerodynamic panel 11, the angle corresponding to the stepped portion of the core structure 12, or the web 13.2 can include a radius of curvature between the lower leg 13.1 and the web 13.2 and / or in the connection between the upper leg 13.3 and the web 13.2.

[0100] When the fiber reinforcement embedded in the fusible material is deposited, the radius of curvature should be desirable to avoid collision between the head 24 and certain components of the aerospace structural component 10.

[0101] Additionally, by placing continuous fibers 2.1, 2.2 embedded in a fusible material to form the lower aerodynamic surface 11 and upper aerodynamic surface 14 of the aerospace structural component 10, an aerospace structural component 10 with smooth curvature can be produced, thereby improving the aerodynamic surface of the aerospace structural component 10.

[0102] Figure 4 shows the comparison with Figure 2 (corresponding to Figure 2). Figure 4a ) or Figure 3 (corresponding to Figure 4b The manufacturing method of the aerospace structure 10 corresponds to different integral composite material aerospace structural components 10, which additionally include non-uniform core structural components 12.

[0103] As in Figure 4a and Figure 4b As can be observed in both, the core structure 12 in these embodiments is manufactured with a hollow space 12.2. An additional example of this core structure 12 is... Figures 4c to 4g It is shown in the middle.

[0104] exist Figure 4c In the core structure 12, a fragmented geometry is included. Figure 4d In the core structure 12, a honeycomb geometry is included. Figures 4e to 4g In the core structure 12, there is a non-uniform core structure with varying density. Specifically, the density increases as it gets closer to the upper and lower aerodynamic panels 14 of the aerospace structure 10.

[0105] Figure 5 illustrates the advantages of combining the deposition of fusible material alone with the deposition of fiber reinforcement embedded within the fusible material.

[0106] The main advantages of this combination are as follows. For illustrative purposes, a comparison is shown in each figure between deposited fusible material and deposited fiber reinforcement embedded within the fusible material, wherein the lower portion of the figure shows only the deposition of fusible material, and the upper portion of the figure shows the deposition of fiber reinforcement embedded within the fusible material:

[0107] Figure 5a : Reduce warpage.

[0108] Because additive manufacturing utilizes the physical effects of melting and hardening through temperature changes, thermal expansion and contraction generate internal stresses in the manufactured parts. In the case of large, flat structural components manufactured through material deposition, this can lead to component deformation due to the shrinkage of layer 1. However, when fiber reinforcements with good coefficients of thermal expansion, such as carbon fibers, are embedded, the warpage of layer 2.1 of the fiber reinforcement embedded within the fusible material can be significantly reduced.

[0109] Figure 5b : Bridging gap.

[0110] In some cases, additive manufacturing can produce overhangs with specific angles without the need for supporting structural members. However, in the absence of a supporting structure, a completely flat overhang is only possible, and the overhang has narrow gaps; otherwise, the molten fusible material 1 wires may elongate downwards. However, by embedding fiber material reinforcement 2.2, this elongation is reduced, making it possible to bridge larger gaps 12.2 without a supporting structure.

[0111] This advantage is particularly beneficial during the fabrication of a core structure 12 having a hollow space 12.2 (as shown in FIG4), and therefore also particularly beneficial during the fabrication of a core structure 12 having a non-uniform core or including a fragmented geometry or including a honeycomb geometry.

[0112] Figure 5c : Reduce the step effect and enhance it in the Z direction.

[0113] Because additive manufacturing deposition along two horizontal axes (X, Y) produces a step effect due to the approaching angles, the threshold is eliminated by depositing an additional layer to cover the edges in a non-planar manner. Additionally, since continuous fibers have the highest mechanical properties along their longitudinal axis, and layer 1 of the fusible material is deposited horizontally, it has a relatively weak load-bearing capacity in the Z direction (approximately perpendicular to the deposition mold). However, this load-bearing capacity is increased by embedding fiber material reinforcements into layer 2.3, resulting in improved mechanical properties of the manufactured part 10 in all directions.

Claims

1. A method for laminated manufacturing of an integral composite aerospace structural component (10), wherein, The method includes the following steps: a) Provide an additive manufacturing tool (20) comprising a deposition mold (23) and at least one head (24), the deposition mold (23) forming an aerodynamic surface, the at least one head (24) being configured to move on the deposition mold (23) and to deposit fibrous material reinforcements and / or fusible materials; b) Deposit the fiber material reinforcement embedded in the fusible material onto the deposition mold (23), thereby forming at least one layer of the lower aerodynamic panel (11); c) Depositing a fusible material onto at least a portion of the outer layer of the lower aerodynamic panel (11), thereby forming at least one layer of the core structure (12); and d) Deposit a fiber material reinforcement embedded in a fusible material onto at least the outer layer of the core structure (12), thereby forming at least one layer of the upper aerodynamic panel (14); Steps b), c), and d) are performed using additive manufacturing technology.

2. The method according to claim 1, wherein, Between step c) and step d), the method further includes: covering at least one edge (12.1) of at least one layer of the core structure by depositing a fiber material reinforcement embedded in a fusible material onto at least a portion of the outer layer of the lower aerodynamic panel (11) and the core structure (12), thereby creating at least one structural element (13).

3. The method according to claim 2, wherein, In the steps between step c) and step d), the fiber material reinforcement is continuous fiber.

4. The method according to claim 1 or 2, wherein, The fiber material reinforcement deposited in step b) comprises continuous fibers.

5. The method according to claim 1 or 2, wherein, The fiber material reinforcement deposited in step d) comprises continuous fibers.

6. The method according to claim 1 or 2, wherein, Step c) includes depositing a fibrous material reinforcement embedded in a fusible material, wherein the fibrous material reinforcement is a short fiber.

7. The method according to any one of claims 1 to 3, wherein, In step c), during the deposition of the fusible material or the fiber reinforcement embedded in the fusible material, a hollow space with or without material is formed (12.2), thereby producing a non-uniform core structure (12) with different densities.

8. The method according to claim 7, wherein, The density of the non-uniform core structure (12) varies along the wingspan and / or chord length of the aerospace structure (10).

9. The method according to claim 2, wherein, The structural element (13) is a beam-shaped structural element, which is a wing beam.

10. The method according to claim 9, wherein, The beam-shaped structural element is a Z-shaped wing beam, and wherein: The deposition on at least a portion of the outer layer of the lower aerodynamic panel (11) forms the lower foot (13.1) of the Z-shaped spar, and The web (13.2) of the Z-shaped spar is formed by deposition on the edge of at least one layer of the core structural member (12).

11. The method according to claim 10, wherein, The deposition on the core structure (12) also forms the upper foot (13.3) of the Z-shaped spar, the upper foot (13.3) being configured to form a continuous surface together with the outer layer of the core structure (12).

12. The method according to claim 10 or 11, wherein, The web (13.2) of the Z-shaped wing beam is straight, and the web (13.2) is at an angle to the upper aerodynamic panel (14) and the lower aerodynamic panel (11), the angle corresponding to the stepped portion of the core structure (12).

13. The method according to claim 11, wherein, The connection between the lower leg (13.1) and the web (13.2) of the Z-shaped spar and / or between the upper leg (13.3) and the web (13.2) of the Z-shaped spar includes a radius of curvature.

14. The method according to any one of claims 1 to 3, wherein, The core structural component (12) includes a fragmented geometry or a honeycomb geometry.

15. An aerospace structural component (10) manufactured using the method according to any one of claims 1 to 14.

16. The aerospace structural component (10) according to claim 15, wherein, The aerospace structural component (10) is a control surface.

Images