Method for manufacturing a structural component made of a composite material reinforced with stiffening stringers and structural component
By rigidly attaching hollow stringers to the skin and using expandable inserts for curing, the problems of lightweighting and durability of composite aerospace structural components are solved, achieving high efficiency in strength and protection.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- LEONARDO SPA
- Filing Date
- 2021-12-22
- Publication Date
- 2026-06-16
AI Technical Summary
Existing technologies struggle to effectively resist the challenges of space debris and impacts while meeting lightweight and strength requirements when manufacturing aerospace structural components made of composite materials.
The composite material manufacturing method using stiffening stringers involves rigidly attaching hollow stringers to the skin and then curing them under high temperature and pressure using expandable inserts to form a closed hollow cross-section structure, thereby enhancing the component's durability.
This approach achieves weight reduction in structural components while improving their resistance to compression and impact, ensuring payload protection and meeting the weight and strength requirements of aerospace vehicles.
Smart Images
Figure CN117203125B_ABST
Abstract
Description
[0001] Cross-references to related applications
[0002] This application claims priority to Italian patent application number 102020000032490, filed on December 28, 2020. Technical Field
[0003] The present invention relates to a method for manufacturing a structural component made of composite material and reinforced with stiffening stringers: particularly structural components of a carrier, such as a space module or a high-speed train.
[0004] The present invention also relates to structural components made of composite materials and reinforced with stiffening stringers, particularly structural components of carriers, such as space modules or high-speed trains.
[0005] In particular, this specification will explicitly refer to the manufacture of the space module shell, where "shell" refers to the tubular external structure that defines the internal environment of a space module (such as a module of a spacecraft), without losing its generality. Background Technology
[0006] Structural components used in the aerospace field (such as aircraft fuselages and parts made of composite materials) are known.
[0007] In existing technologies, some aerospace structural components are made of light alloys, and are therefore manufactured from metallic materials, which are then used to form the fuselage of an aircraft.
[0008] As is well known, the fuselage must provide adequate protection for the payload (crew, passengers, cargo, etc.) but at the same time must not exceed the desired weight limit.
[0009] Therefore, using metal components in exchange for greater strength leads to an increase in overall cost.
[0010] This necessitates structural components made of composite materials to reduce the overall weight of the aircraft. In fact, using composite materials can reduce the overall weight of the aircraft while simultaneously achieving a very robust structure.
[0011] Typically, the aforementioned structural components (such as the fuselage or its parts) are made by connecting a skin made of composite material to a plurality of stiffening stringers, also made of composite material and conveniently positioned parallel to the longitudinal extension direction of the fuselage.
[0012] Specifically, each stringer is typically defined by a thin-walled longitudinal profile, including:
[0013] - The central portion, which is typically concave and has a predetermined cross-section, such as a rectangle, polygon, arc, semi-ellipsoid, semi-ellipse, etc.; and
[0014] - Two longitudinal transverse sections or wings, extending from opposite sides of the central section in a direction relative to the longitudinal direction of the stringer.
[0015] The most commonly used stringers in this field have an Ω-shaped cross-section.
[0016] In the most common solutions, the composite material used consists of fibrous materials (such as uncured or pre-cured carbon fibers) that are typically pre-impregnated with a fluid resin according to a well-known process (e.g., by a method known as "resin transfer molding" or RTM).
[0017] In particular, each layer in the composite material is typically composed of a thermosetting (resin) matrix prepreg that is reinforced with different types of fibers (such as carbon fiber, aramid fiber, glass fiber, etc.).
[0018] In order to produce the skin, multiple layers of the uncured or pre-cured composite material are laminated together.
[0019] Similarly, in order to produce stringers, multiple layers of uncured or pre-cured composite materials are placed on a forming tool of appropriate shape.
[0020] Once the skin and stringers are manufactured, the stringers contact the skin at their respective flanges to form a set of closed profile cavities between the central recess of each stringer and the skin itself.
[0021] The resulting component is then cured by applying high pressure and high temperature to solidify the composite material, compacting the layers together and having struts connected to the skin while retaining the cavity.
[0022] In practice, each stringer is rigidly and integrally applied to one face of the skin, which typically defines the inner wall of the skin, i.e. the face facing the interior of the fuselage during use.
[0023] This results in structural components. The skin and stringers can be connected in a variety of ways.
[0024] The pattern, known as the "internal mold line" or IML, provides a purpose for a curing tool, often referred to as a "spindle." This curing tool is externally shaped to define the inner surface or part of the structure to be achieved, such as a fuselage or a section of the fuselage. In practice, the spindle has corresponding longitudinal cavities, each adapted to accommodate stiffening stringers.
[0025] Once the stringer is positioned in the aforementioned shaft cavity, various types of inserts, referred to in the art as “bladders” and “noodles”, are inserted into the various cavities formed by positioning the stringer on the shaft; these inserts are adapted to hold the various components in place and prevent the various components from being crushed by high pressure during the subsequent curing step.
[0026] At this point, the assembly consisting of shafts, stringers, and inserts covers the opposite portion of the skin, which will form the outer surface of the aforementioned portion of the fuselage.
[0027] Therefore, in the generated components, the axis defines the innermost part.
[0028] At this point, the entire assembly is brought into an autoclave for high-temperature and high-pressure curing operations, whereby the stringers are securely connected and integrated into the skin.
[0029] Given the above, the outer surface of the component is defined by the smooth outer surface of the skin, while the inner surface of the component (i.e., the inner surface facing the internal environment of the carrier) is defined by the shaped inner surface of the skin, and the stringers are integrally fixed to the shaped inner surface of the skin.
[0030] US2015 / 122413A1, DE102009056978A1, and EP2998228A describe corresponding methods and apparatus for manufacturing aerospace structural components made of composite materials reinforced according to the aforementioned stiffening stringers.
[0031] Composite materials are increasingly used in the aerospace field, and this manual will mention them without losing its general relevance.
[0032] In this field, the use of components made of composite materials is determined by the need to reduce the overall weight of spacecraft and modules.
[0033] In particular, space module shells made of composite materials are required, that is, external tubular structures that define the internal environment of the space module, such as the shells of satellites or spacecraft.
[0034] The enclosure must ensure adequate protection for the payload (sensors, supplies, crew, etc.), but at the same time must not exceed the weight limit imposed by the fact that placing the space module in orbit is very expensive; therefore, any reduction in weight means a significant reduction in cost.
[0035] With particular reference to the aerospace field, it is recognized that there is a need to produce structural components (particularly the shell or portions thereof) of spacecraft or modules made of composite materials, since such modules, when operated in a vacuum, are sufficiently resistant to compression caused by the high stresses (e.g., sudden acceleration) experienced by the module during launch, and are also characterized as being sufficiently resistant to impact bodies (e.g., micrometeorites or space debris generated by human activity).
[0036] This last aspect is becoming increasingly important because in recent years, many of Earth's orbits have been intruded upon by countless pieces of space debris generated by human activities. These debris ranges in size from a few millimeters to quite large objects, all of which can damage the carriers or space modules occupying these orbits. Summary of the Invention
[0037] Therefore, the object of the present invention is to provide a method for manufacturing structural components made of composite materials reinforced with stiffening stringers, which can simply and cost-effectively meet the above requirements and the requirements associated with structural components made of composite materials for use in spacecraft or modules.
[0038] Another object of the present invention is to design a structural component made of composite material reinforced with stiffening stringers, which makes it possible to simply and cost-effectively meet the requirements specified above and related to structural components made of composite materials for use in spacecraft or modules. Attached Figure Description
[0039] To better understand the present invention, some preferred and non-limiting embodiments are described below by way of example and with reference to the accompanying drawings, wherein:
[0040] - Figure 1 This is a perspective view of a carrier, particularly a spacecraft, having an outer shell consisting of two structural components manufactured by the manufacturing method according to the present invention, with parts removed for clarity;
[0041] - Figure 2 yes Figure 1 A perspective view of one of the structural components, with parts removed for clarity;
[0042] - Figure 3 and Figure 4 Two corresponding subsequent steps of the method according to the invention are schematically illustrated in perspective and side view, respectively;
[0043] - Figure 5 and Figure 7a The expandable insert used in the method according to the invention is shown in perspective and side view, respectively.
[0044] - Figure 6 and Figure 7bThe reinforcing element of the structural component 1 according to the present invention is shown in perspective view and side view, respectively;
[0045] - Figure 8 and Figure 9 Two subsequent moments of the steps of the method according to the invention are schematically shown in a three-dimensional diagram;
[0046] - Figure 9a A first embodiment of the steps of the method according to the present invention is shown in a perspective view;
[0047] - Figure 9b A second embodiment of the steps of the method according to the present invention is shown in a perspective view;
[0048] - Figure 9c A third embodiment of the steps of the method according to the present invention is shown in a perspective view;
[0049] - Figure 9d , Figure 9e and Figure 9f A fourth embodiment of the steps of the method according to the invention is shown in a side view;
[0050] - Figure 10 , Figure 11 and Figure 12 The steps of the method according to the present invention are schematically illustrated in a perspective view;
[0051] - Figures 13 to 16 Two subsequent steps of the method according to the invention are schematically illustrated in a perspective view;
[0052] - Figure 17 The steps of the method according to the invention are schematically illustrated in a side view;
[0053] - Figure 18 The steps of the method according to the present invention are illustrated in a perspective view;
[0054] - Figure 19 The method according to the invention is shown in enlarged scale and in partial cross-sectional views. Figure 18 Details of the structural components during the steps;
[0055] - Figure 20 The steps of the method according to the present invention are illustrated in a perspective view;
[0056] - Figure 21 The method according to the invention is shown in enlarged scale and in partial cross-sectional views. Figure 20 Details of the structural components during the steps;
[0057] - Figure 22 The steps of the method according to the invention are schematically illustrated in a perspective view; and
[0058] - Figure 23 and Figure 24 The execution is shown in sectional views. Figure 22 The structural components before and after the steps. Detailed Implementation
[0059] refer to Figure 1 and Figure 2 Reference numeral 1 generally indicates a structural component made of composite material and reinforced by stiffening stringers 2, which in use defines at least a portion of the outer shell 3 of the carrier 4 and includes a composite skin 5 made of composite material and a plurality of composite stringers 2 made of composite material having a hollow cross section and being rigidly and integrally attached to the skin 5.
[0060] In particular, this specification will explicitly refer to structural components 1 used in the aerospace field, such as the outer shell 3 of a spacecraft or space module, or a portion thereof, as well as methods for manufacturing such shell 3 or a portion thereof, without losing its generality.
[0061] More specifically, component 1 described and illustrated herein defines a cylindrical module portion of the outer shell 3 of the carrier 4, which, in this preferred and non-limiting embodiment, is defined by the spacecraft or spacecraft module (i.e., the portion of the spacecraft used for storing and transporting payload) as expressly referred to in this specification, without losing its generality.
[0062] More precisely, the “outer shell” refers to the protective shell that separates the internal environment of the carrier 4 (i.e., the space module) from the external environment, thereby defining its volume.
[0063] For example, in the case of an aircraft, such a shell 3 is defined by the fuselage. In the case of a train, especially a high-speed train, the shell 3 is defined by an outer shell that defines the main body of each carriage or locomotive.
[0064] It should also be noted that structural component 1 can be defined by a portion or part of such housing 3 (e.g., a panel (not shown)) which precisely includes a portion of skin 5, and stringers 2 are rigidly and integrally fixed to a portion of skin 5 in the manner described below.
[0065] According to the preferred, non-limiting embodiments described and illustrated herein, component 1 has a cylindrical shape around a longitudinal axis A.
[0066] In particular, Figure 1 The housing 3 of the carrier 4 consists of two cylindrical structural components 1 arranged in series and connected together at the axial annular wall.
[0067] Component 1 includes a skin 5 made of composite material and a set of hollow longitudinal stiffening struts 2 (preferably having a closed cross section), which are configured to reinforce and stiffen the panel.
[0068] Each stringer 2 has a much larger range in one longitudinal direction than in the other two directions orthogonal to that longitudinal direction.
[0069] The longitudinal extension direction of the stringer 2 is also parallel to the axis A of the shell 3 formed at least partially by the structural component 1 during use (i.e., under installation / assembly conditions).
[0070] In other words, each stringer 2 has a longitudinal axis B ( Figure 7b ), which is parallel to axis A of the shell 3 (i.e., skin 5) during use.
[0071] The use of composite materials in structural components is driven by the need to reduce the overall weight of spacecraft or space modules.
[0072] In one embodiment, the composite material consists of a fibrous material (e.g., uncured or pre-cured carbon fiber).
[0073] In one embodiment, the material is pre-impregnated with a fluid resin according to a well-known process (e.g., by a method known as "resin transfer molding" or RTM) that is not described in detail.
[0074] In practice, each layer in a composite material typically consists of a thermosetting (resin) matrix prepreg, which is reinforced with different types of fibers (such as carbon fiber, aramid fiber, glass fiber, etc.).
[0075] The present invention relates to a method for manufacturing a structural component 1, which is obtained by rigidly and integrally applying stringers 2 to a skin 5, preferably such that each stringer 2 defines a hollow section with a closed profile with the skin 5 itself.
[0076] In particular, this specification will specifically relate to the manufacture of the hollow cylindrical (i.e. tubular) component 1 without losing its generality.
[0077] However, the structural and functional characteristics and process steps are equally applicable to the manufacture of individual panels that extend along a flat, curved, or rotationally curved lying surface (e.g., along a substantially parabolic, arched, or still (substantially) cylindrical or truncated conical surface).
[0078] Therefore, under the current conditions, skin 5 will have a (basically) cylindrical shape with a central longitudinal axis corresponding to axis A. Stringers 2 will be arranged with their corresponding longitudinal axes B, which are parallel to axis A.
[0079] Furthermore, without loss of generality, this specification will explicitly refer to the manufacturing process known as the “inner mold line” or IML type.
[0080] However, the steps of this process also apply to the manufacture of structural components according to a process known as “outer mold line” or OML type (not shown), which is known in itself but not described in detail.
[0081] refer to Figure 4 Skin 5 is formed by passing the first layer 5a of uncured or precured composite material. Figure 4 The image shows only one layer of the image, which is obtained by arranging it on a curing tool 6, which is commonly referred to as a “shaft”. In the non-limiting example described herein, the shaft has a substantially cylindrical shape and a longitudinal central axis C.
[0082] In practice, to form the skin 5, the manufacturing method includes the step of laminating a first layer 5a of a plurality of uncured or precured composite materials onto the outer surface 6a of the tool 6, such that the first surface 5b of the skin 5, which defines the inner surface of the component 1 during use, is... Figure 2 The outer surface of tool 6 faces the inner surface of component 1, while the inner surface of component 1 faces the internal environment of carrier 4.
[0083] Conveniently, the outer surface 6a is smooth, i.e. it is unshaped, i.e. it does not have macroscopic surface features suitable for defining grooves or cavities visible to the naked eye.
[0084] Therefore, lamination is performed by arranging layers 5a on the smooth surface 6a of tool 6.
[0085] Advantageously, before laminating layer 5a onto tool 6, the method includes the step of applying (particularly spraying) an adhesive substance 7 (e.g., an adhesive or viscous resin) onto the outer surface 6a, such as... Figure 3 As shown.
[0086] The presence of adhesive 7 on the outer surface 6a of tool 6 also improves the adhesion of composite layer 5a during its lamination.
[0087] Preferably, the method includes the advantageous step of applying a gas barrier material layer before laminating layer 5a onto tool 6 and subsequently applying adhesive 7 onto tool 6.
[0088] In detail, the gas-impermeable or substantially gas-impermeable material layer may be defined by a skin of EVOH (ethylene-vinyl alcohol) or PVDF (polyvinylidene fluoride) or other polymers and / or metallic materials suitable for this purpose.
[0089] Conveniently, according to the advantageous steps (not shown) of the manufacturing method according to the invention, another layer of adhesive 7 can be applied on top of a material layer that is impermeable or substantially impermeable to gases.
[0090] Alternatively, a single-layer material that is impermeable or substantially impermeable to gases can be applied from the outside to a cured (or co-bonded) assembly obtained after a curing process, as described below.
[0091] Preferably, the arrangement of the composite layer 5a on the tool 6 and the resulting lamination are performed by an automated machine 8, which is known in itself and not described in detail, and is commonly referred to as "AFPM" (Automatic Fiber Placement Machine).
[0092] In an alternative embodiment, the arrangement of layer 5a is done manually.
[0093] During the arrangement of each layer 5a, tool 6 is positioned to rotate about its own axis C.
[0094] refer to Figure 5 , Figure 6 , Figure 7a and Figure 7b Each stringer 2 has a longitudinal axis B as described above.
[0095] Specifically, each stringer 2 is made by arranging a second layer 2a of multiple uncured or precured composite materials on the shaped portion of the corresponding hollow expandable insert 10 (commonly referred to as a "bladder").
[0096] More specifically, each insert 10 has an elongated longitudinal shape, opposing axial ends with openings, and a hollow cross section with a closed profile, which, according to the described example, has a substantially trapezoidal shape.
[0097] To form the stringers 2, the method according to the invention includes the step of laminating a plurality of corresponding second layers 2a onto corresponding inserts 10 to form a plurality of stringers 2, each stringer 2 having a corresponding axis B, such that each stringer 2 is wound around the outer surface 10a of the corresponding insert 10 to define a longitudinal cavity 11 for each stringer 2, the longitudinal cavity 11 being engaged by the corresponding insert 10.
[0098] Preferably, layer 2a is arranged to completely surround and cover the outer surface 10a of insert 10.
[0099] Therefore, according to the non-limiting example described herein, each stringer 2 has a trapezoidal closed profile hollow cross section.
[0100] Each insert 10 is made of an elastically deformable material and is expandable during the subsequent curing process of component 1 in order to retain cavity 11, which will be explained in more detail below.
[0101] According to an alternative embodiment not shown, each stringer 2 is formed by partially wrapping the layer 2a around the corresponding insert 10, for example, obtaining a cross-section with an Ω shape, having a central portion and two wings extending laterally from that central portion. In this case, once the stringer 2 is attached to the skin 5, a closed profile hollow cross-section of the stringer 2 is obtained.
[0102] For example, stringer 2 can have an Ω cross section or be semi-circular.
[0103] According to an alternative embodiment not shown, each stringer 2 can be wound around a corresponding insert 10, thus having a T-shaped or L-shaped cross-section. In this case, once positioned on the skin 5, each stringer 2 will define an open profile cross-section, rather than being closed as in the previous case. Furthermore, in this case, a co-bonding or co-curing process is performed to attach the stringers 2 to the skin 5.
[0104] like Figure 6 As shown, a set of reinforcing elements 12 is thus obtained, each reinforcing element 12 being defined by a stringer 2 that accommodates a corresponding insert 10 in its cavity 11, or by a stringer 2 wound around the corresponding insert 10.
[0105] refer to Figure 8 and Figure 9 In order to achieve the structural shape of component 1, the method includes the steps of positioning each reinforcing element 12 on a second surface 5c opposite to the first surface 5b of the skin 5, defining the outer surface of component 1 facing the external environment of the carrier 4 in use, and making the cavity 11 of each stringer 2 face the second surface 5c.
[0106] Conveniently, this positioning step is performed by placing the reinforcing element 12 on the second surface 5c of the skin 5 in the corresponding positions, which are equidistant from each other at an angle around axis C and the corresponding longitudinal axes B of the stringers 2 are parallel to each other and parallel to axis C.
[0107] In an alternative embodiment not shown, the reinforcing element 12 and thus the stringers 2 are arranged to be spaced apart unequally around axis C.
[0108] In another embodiment, the reinforcing element 12 and thus the stringer 2 are arranged such that their respective longitudinal axes B are not parallel to each other and not parallel to axis C.
[0109] In view of the foregoing, the reinforcing element 12 and thus the stringer 2 are applied (i.e. attached) to the surface of the skin 5, which in use defines the outer surface of the component 1, i.e. the surface of the component 1 facing the external environment of the carrier 4 in use, rather than the inner surface.
[0110] The method also includes the step of holding (i.e. maintaining) each reinforcing element 12 on the second surface 5c of the skin 5 in a corresponding fixed position relative to the tool 6.
[0111] This holding step is preferably performed by arranging at least one holding element 20, 22, 24, 26 at each reinforcing element 12 in order to hold the reinforcing element 12 itself in its corresponding fixed position relative to the tool 6.
[0112] The hold mode will be described in more detail below.
[0113] To ensure effective adhesion between the stringer 2 and the skin 5, the composite material layers 5a and 2a are compacted together.
[0114] Therefore, refer to Figures 13 to 16 Vacuum bag 13 is wrapped around the components obtained so far. Figure 13 and Figure 14 ), to accommodate tool 6, skin 5 and reinforcing element 12.
[0115] Once the vacuum bag 13 is sealed, a vacuum is applied to the interior of the latter. This compacts the layers 5a and 2a. Figure 15 and Figure 16 ).
[0116] After the compaction step, and because the composite material is pre-impregnated with the aforementioned resin, the reinforcing element 12 (i.e., stringer 2) remains fixed in its corresponding position on the second surface 5c of the skin 5 even after the aforementioned retaining elements 20, 22, 24, 26 are removed.
[0117] refer to Figure 10 , Figure 11 and Figure 12 The method preferably, but not necessarily, includes the step of placing other longitudinal inserts 14 on the second surface 5c of the skin 5 after the step of holding each reinforcing element 12 on the second surface 5c of the skin 5 and before applying (wrapping) the vacuum bag 13 around the components obtained so far.
[0118] In particular, the other inserts 14 are generally referred to as “noodles” and are defined by rods, which preferably have a circular cross-section and are made of composite materials, more specifically adhesive composite materials, and more specifically prepreg composite materials and uncured or pre-cured composite materials.
[0119] As is known in the art, other inserts 14 serve as fillers and are thus configured to engage and fill the cavities and gaps defined between the stringers 2 and the skin 5 under assembly conditions.
[0120] Advantageously, since the insert 14 is made of an adhesive composite material, it does not require any retaining elements to hold it on the skin 5; otherwise, the insert 14 would be fixed in place by adhesive.
[0121] like Figure 10 , Figure 11 and Figure 12 As shown, the step of positioning the other inserts 14 is performed by arranging a pair of other inserts 14 for each reinforcing element 12, each other insert being adjacent to the longitudinal side of the corresponding reinforcing element 12 and parallel to the axis B of the corresponding stringer 2.
[0122] In other words, a pair of other inserts 14 are placed on the side of each stringer 2.
[0123] Therefore, the step of housing the components obtained so far within the vacuum bag 13 also includes housing other inserts 14 within the vacuum bag 13 itself.
[0124] Therefore, the aforementioned compaction of layers 5a and 2a by applying a vacuum inside the vacuum bag 13 also includes compaction of each pair of other inserts 14 at their positions adjacent to the corresponding stringers 2, such that the pair of other inserts 14 are compacted at the longitudinal side of the corresponding reinforcing element 12 to the second layer 2a forming the corresponding stringer 2 and the first layer 5a forming the skin 5.
[0125] Therefore, at the end of the compaction step, the other inserts 14 are compacted together with the first layer 5a and the second layer 2a, as shown below. Figure 17 As shown.
[0126] Preferably, once the above compaction has been performed, the vacuum bag 13 is removed. Figure 17 ).
[0127] like Figures 20 to 24 As shown, the resulting assembly (i.e., the assembly including tool 6, skin 5, reinforcing element 12 and (if applicable) other inserts 14) is then cured by applying high pressure and high temperature to cure the composite material, compacting the aforementioned layers 5a and 2a together and having stringers 2 connecting the skin 5, while maintaining the aforementioned cavity 11.
[0128] Specifically, the curing process involves applying high pressure and high temperature (approximately 6 bar and 180°C) to the aforementioned components.
[0129] In this regard, the method for manufacturing structural component 1 further includes the following steps:
[0130] - After the above compaction is completed, the tool 6, skin 5, reinforcing element 12, and (if applicable) other inserts 14 are housed in other vacuum bags 15. Figure 20 );
[0131] - Seal the vacuum bag 15 to define the vacuum chamber 16 therein; and
[0132] - Preset temperature and pressure (preferably about 6 bar and 180°C) are applied to the outside of vacuum chamber 16 to cure the composite material and determine the rigid and integral attachment of stringer 2 to the second surface 5c of skin 5, and if present, determine the rigid and integral attachment of other inserts 14 to the second surface 5c and the rigid and integral attachment of other inserts 14 to stringer 2.
[0133] Specifically, after sealing the vacuum bag 15, the entire assembly is inserted into the autoclave 17, and the aforementioned temperature and pressure conditions are applied inside the autoclave 17. Figure 22 ).
[0134] Inside the autoclave 17, the insert 10 expands due to the applied high temperature, thereby maintaining the cavity 11 and preventing the cavity 11 from being crushed by the applied high pressure.
[0135] Through the curing process, the composite material is cured, and we see that the skin 5, stringer 2, and insert 14 are mutually attached and compacted parts. Figure 23 The conditions of ) are transformed into the conditions of these parts being limited to a single structure 18 made of cured composite material ( Figure 24 ).
[0136] This yields structural component 1. At this point, the vacuum bag 15 is removed, and preferably, the insert 10 is removed, while the other inserts 14 remain (as described above) included in the structure 18 of component 1.
[0137] It should be noted that after the insert 10 is removed following the curing process, each reinforcing element 12 is defined only by the associated stringer 2.
[0138] Conveniently, the vacuum bag 13, which is compacted by it, is removed before the above-described steps of the curing process.
[0139] Preferably, in order to improve the curing process, especially after the above-mentioned compaction, the corresponding pressure bag 19 is inserted into the cavity 11 defined by each hollow insert 10.
[0140] Therefore, bag 19 is also contained within vacuum chamber 16 defined by vacuum bag 15.
[0141] The method also includes the step of placing each bag 19, which is pre-reserved in the cavity of each insert 10, into fluid communication with the interior of the autoclave 17, i.e., fluid communication with the curing environment outside the vacuum chamber 16.
[0142] Therefore, during the curing process, the aforementioned preset temperature and pressure are also applied inside each bag 19.
[0143] Therefore, during the curing process in autoclave 17, bag 19 expands ( Figure 23 and Figure 24 This further helps to maintain the cavity 11 of each stringer 2.
[0144] According to the present invention, the above-described steps of holding each reinforcing element 12 of the second surface 5c of the skin 5 in a corresponding fixed position opposite to the tool 6 can be performed in various ways, some of which will be described below.
[0145] According to the first embodiment, holding the reinforcing element 12 in its respective fixed position includes using a strap element 20 that defines the aforementioned constraint element, such as... Figure 9b As shown.
[0146] Specifically, the method includes the following steps:
[0147] - Insert the element 20 through the cavity of each insert 10;
[0148] - Attach the corresponding opposite ends of the element 20 to the tool 6; and
[0149] - Each reinforcing element 12 is secured to the tool by inserting it into the cavity of the corresponding insert 10 and fixing it to the tool 6 by the strap element 20.
[0150] Conveniently, the tool 6 includes cavities 21, each cavity 21 being adapted to engage with a corresponding belt element 20, particularly with the end of the corresponding belt element 20, and within the cavity 21, the ends are attached to the tool 6 or to each other to determine the engagement of the corresponding reinforcing element 12 with the tool 6 itself.
[0151] According to another embodiment, holding the reinforcing element 12 in its respective fixed position includes using a magnetic element 22 ( Figure 9a ).
[0152] Specifically, in this case, tool 6 includes a ferromagnetic portion made of a ferromagnetic material, and the method includes the following steps:
[0153] - At least one magnetic element 22 is arranged at each reinforcing element 12; and
[0154] - The reinforcing element 12 is sandwiched between the corresponding magnetic element 22 and the tool 6 so that the reinforcing element 12 is held in place by the magnetic interaction between the magnetic element 22 and the ferromagnetic portion of the tool 6.
[0155] In detail, each magnetic element 22 is defined by a permanent magnet.
[0156] More specifically, a magnetic strip 23 (i.e., a strip of polymer or composite material carrying the magnetic element 22) is inserted into a cavity defined by each insert 10. The magnetic interaction between the magnetic element 22 of the strip 23 and the ferromagnetic portion of the tool 6 holds the associated reinforcing element 12 in place.
[0157] According to another embodiment, the reinforcing element 12 is connected to terminal 24 ( Figure 9c Each terminal 24 is fixed to the corresponding insert 10 and tool 6 to fix the corresponding reinforcing element 12 to the tool 6.
[0158] In view of the above, the step of housing the components in the vacuum bag 13 further includes housing the retaining elements in the vacuum bag, whether these retaining elements are defined by the tape element 20, the magnetic element 22 or the terminal 24.
[0159] Once the above compaction step has been performed with vacuum bag 13, the retaining elements 20, 22, and 24 are removed before the curing process steps, particularly before wrapping the assembly with another vacuum bag 15.
[0160] Due to compaction, the reinforcing element 12 remains in its fixed position even after the retaining elements 20, 22, and 24 are removed.
[0161] According to an optional embodiment, holding the reinforcing element 12 in its respective fixed position includes using a vacuum bag 25, which comprises a plurality of adjacent longitudinally recessed pockets 26 tightly sealed together with fluid. Figure 9d , Figure 9e , Figure 9f ).
[0162] Specifically, in this case, the method includes the step of wrapping the bag 25, which includes the concave bag 26, around the skin 5 laminated on the tool 6, particularly before positioning the reinforcing element 12 on the same skin 5.
[0163] Subsequently, the reinforcing element 12 is positioned by sequentially arranging each reinforcing element 12 within the corresponding concave bag 26. Figure 9d ).
[0164] Therefore, according to this embodiment, the step of housing the component in the vacuum bag is performed by the steps of wrapping the bag 25 and arranging the reinforcing element 12 in the concave bag 26.
[0165] Therefore, in this case, vacuum bag 13 is defined by bag 25.
[0166] Then, each concave bag 26 is fluid-tightly resealed (e.g., with a regular bag sealant), and a vacuum is applied to its interior.
[0167] Then, the sequential compaction of layers 5a and 2a, as well as other inserts 14 (if present), is performed by sequentially applying vacuum within each concave bag 26.
[0168] Therefore, the step of holding the reinforcing element 12 on the second surface 5c of the skin 5 is performed by the above-described compaction.
[0169] Preferably, once all the concave bags 26 have been compacted, and before the curing process described above, the bags 25 are removed.
[0170] In this case, the retaining element is defined by the concave bag 26.
[0171] Finally, through the above manufacturing method, a structural component 1 (preferably a cylindrical tubular shell 3 of a carrier 4 defined by a space module) is obtained. The structural component 1 includes a skin 5 and a plurality of stringers 2 rigidly and integrally fixed to a second surface 5c of the skin 5. The second surface 5c defines an outer surface of the structural component 1 suitable for facing the environment outside the carrier 4 during use.
[0172] Therefore, the inner surface of structural component 1 does not have any stringers 2, but is smooth.
[0173] When examining the manufacturing method and features of the structural component 1 described above, the advantages they allow for are obvious.
[0174] Specifically, because the stringers 2 are arranged on the second surface 5c of the skin 5 and thus define the outer surface of the structural member 1, and preferably the outer surface of the shell 3, the structural member 1 can be provided with improved resistance to compression and increased resistance to impacts (e.g., micrometeorites or space debris generated by human activity), thereby ensuring adequate protection of the payload contained within the carrier 4. This increase in strength is preferably achieved through another aluminum protective layer (not shown) arranged to cover the stringers 2, i.e., the actual outer shell of the carrier 4.
[0175] At the same time, structural component 1, made of composite materials, is more likely to meet the desired weight constraints and therefore has a greater capacity to carry additional payload. This is particularly advantageous when the carrier 4 is limited by a space module or spacecraft.
[0176] In this case, the specific external construction of the structural component 1 with the stringers 2 arranged on the outside ensures greater tolerance to the stresses experienced by the carrier 4 during launch.
[0177] The manufacturing method according to the invention makes it possible to obtain such a structural component 1, because it advantageously demonstrates the process of obtaining a structural component 1 with external stringers 2 by means of IML technology.
[0178] Obviously, changes and variations can be made to the manufacturing method of the structural component 1 described and shown herein without departing from the scope of protection defined by the claims.
[0179] In particular, the described manufacturing method is also applicable to the case where the structural component 1 is defined by the panel (i.e., the flat portion) of the housing 3.
Claims
1. A method for manufacturing a structural member (1) made of composite material and reinforced by stiffening stringers (2), said structural member (1) defining at least a portion of an outer shell (3) of a box (4) in use and comprising a skin (5) and a plurality of stringers (2) having a hollow section rigidly and integrally fixed to said skin (5), said method comprising the steps of: a) Laminating a first layer (5a) of a plurality of uncured or precured composite materials onto the outer surface (6a) of a curing tool (6) to form the skin (5), such that the first surface (5b) of the skin (5) defining the inner surface of the structural component (1) in use faces the outer surface (6a) of the tool (6), and the inner surface of the structural component (1) faces the internal environment of the housing (4); b) A second layer (2a) of a plurality of corresponding composite materials is laminated on a longitudinally hollow expandable insert (10) to form a plurality of said stringers (2), each stringer (2) having a longitudinal axis (B) such that each stringer (2) is at least partially wound around the outer surface (10a) of a corresponding hollow insert (10), thereby defining a longitudinal cavity (11) of each stringer (2) joined by the associated hollow insert (10), and thus obtaining a plurality of reinforcing elements (12), each reinforcing element (12) being defined by one of said stringers (2) wound around a hollow insert (10); c) Position each reinforcing element (12) on the second surface (5c) of the skin (5) opposite to the first surface (5b), and define the outer surface of the structural member (1) in use, the outer surface facing the external environment of the housing (4), and such that the cavity (11) of each stringer (2) faces the second surface (5c). d) Hold each reinforcing element (12) on the second surface (5c) of the skin (5) in a corresponding fixed position relative to the tool (6); e) The tool (6), the skin (5), and the reinforcing element (12) are housed in a vacuum bag; f) By applying a vacuum inside the vacuum bag, the first layer (5a) forming the skin (5) and the second layer (2a) forming the stringer (2) are compacted together; g) After step f), the tool (6), skin (5) and reinforcing element (12) are housed in other vacuum bags; h) Seal the other vacuum bags to define a vacuum chamber (16) therein; i) Apply a preset temperature and pressure to the outside of the vacuum chamber (16) to cure the composite material and determine the rigidity and integral attachment of the stringer (2) to the second surface (5c) of the skin (5); The method further includes the following steps: n) Prior to step c), the vacuum bag is wrapped around the skin (5) laminated on the tool (6), the vacuum bag comprising a plurality of adjacent longitudinal concave bags (26) capable of being fluid-tightly sealed to each other; The positioning step c) includes placing each reinforcing element (12) into the corresponding concave bag (26) in sequence; The containing step e) is performed through the wrapping step n) and the positioning step c). The compaction step f) is performed by sequentially applying a vacuum within each concave bag (26) containing the corresponding reinforcing element (12); Furthermore, step d) of maintaining the compaction mentioned in step f) is performed.
2. The method according to claim 1, further comprising the following step: j) Before lamination step a), an adhesive substance (7) is applied to the outer surface (6a) of the tool (6) to improve the adhesion between the first layer (5a) and the outer surface (6a) during step a).
3. The method according to claim 2, further comprising the following step: q) Before lamination step a) and subsequent application step j), a layer of gas-barrier material that is gas-impermeable or substantially gas-impermeable is applied to the outer surface (6a) on which the adhesive material (7) was previously applied.
4. The method according to claim 1, further comprising the following steps: k) Insert a corresponding pressure bag (19) into the cavity defined by each hollow insert (10); Step g) further includes housing the pressurization bag (19) inside the vacuum chamber (16); Step h) further includes arranging each pressurized bag (19) pre-reserved in the cavity of each insert (10) in fluid communication with the curing environment (17) outside the vacuum chamber (16); Furthermore, step i) also includes applying a preset temperature and pressure inside each pressure bag (19).
5. The method according to claim 1, wherein, Step d) of maintaining includes the following steps: l) At least one retaining element is arranged at each reinforcing element (12) to hold the reinforcing element (12) in its corresponding fixed position relative to the tool (6).
6. The method according to claim 5, wherein, Step 1) includes the following steps in sequence: - Insert the belt element (20) into the cavity passing through each insert (10); - Attach the corresponding opposite ends of the strap element (20) to the tool (6); and - Each reinforcing element (12) is secured to the tool (6) using the band element (20) which is inserted into the cavity of the corresponding insert (10) and fixed to the tool (6).
7. The method according to claim 5, wherein, The tool (6) includes a ferromagnetic portion made of a ferromagnetic material, and wherein step 1) includes: - At least one magnetic element (22) is arranged at each reinforcing element (12); and - The reinforcing element (12) is sandwiched between the corresponding magnetic element (22) and the tool (6) so that the reinforcing element (12) is held in place by the magnetic interaction between the magnetic element (22) and the ferromagnetic portion.
8. The method according to claim 5, wherein, Step e) of the containment process also includes containing the retaining element within the vacuum bag; Furthermore, the method further includes the following steps: m) Once the compaction mentioned in step f) is complete and before performing steps g), h) and i), the retaining element is removed.
9. The method according to claim 1, further comprising the following step: o) Once the compaction mentioned in step f) is complete and before performing steps g), h) and i), the vacuum bag is removed.
10. The method according to claim 1, further comprising the following step: p) After step d) and before step e), position the other longitudinal inserts (14) on the second surface (5c) of the skin (5); Step p) includes arranging a pair of the other inserts (14) for each reinforcing element (12), each of the other inserts (14) being adjacent to the longitudinal side of the corresponding reinforcing element (12) and parallel to the longitudinal axis (B) of the corresponding stringer (2).
11. The method according to claim 10, wherein, The other inserts (14) are made of an adhesive composite material; The accommodating step e) further includes accommodating the other insert (14) in the vacuum bag; And wherein, the compaction step f) includes compacting each pair of other inserts (14) adjacent to each reinforcing element (12) such that the pair of other inserts (14) are compacted at the longitudinal side of the respective reinforcing element (12) to form the second layer (2a) of the respective stringer (2) and the first layer (5a) of the skin (5).
12. The method according to claim 1, wherein, The lamination step a) is performed by placing the first layer (5a) on the smooth outer surface (6a) of the curing tool (6).
13. The method according to claim 1, wherein, The tool (6) has a generally cylindrical shape around a central axis (C); The positioning step c) is performed by positioning the reinforcing element (12) on the second surface (5c) of the skin at positions that are angularly spaced around the central axis (C), and the longitudinal axes (B) of the corresponding stringers (2) are parallel to each other and parallel to the central axis (C).