Method of fabricating structures using composite modules and structures made thereby

ES 2 638 263 T3

DESCRIPTION

Method of fabricating structures using composite material modules and fabricated structures.

Technical field

This description generally refers to techniques for manufacturing composite material structures, and deals more particularly with a method for manufacturing composite material structures on a large scale by joining composite material modules together.

Background

Large-scale composite material structures such as aircraft fuselage skins can be manufactured using advanced fiber-laying machines (AFPs) capable of disposing composites at relatively high speeds. One way to achieve higher unit production speeds can be obtained by using a greater number of AFP machines, however, the use of additional AFP machines can lead to the need for significant capital investments in terms of machine, tool costs. and factory floor space.

Accordingly, there is a need for a manufacturing method that achieves relatively high production speeds with relatively low investment and operating costs. There is also a need for a manufacturing method that is relatively flexible and relies on equipment that is less complicated than AFP machines.

US 2007/289246 discloses a tongue and groove joint and method. To fabricate a joint in a panel, a slit is generated and the panel is folded along the slit. The panel includes a first skin, a second skin, and a core. The core is sandwiched between the first cladding and the second cladding. The slit passes through the first cladding and at least a part of the core. The groove includes a set of tabs and a corresponding set of grooves. The set of tabs engages the set of grooves in response to bending of the panel along the slit.

Summary

According to one aspect, there is provided a method for manufacturing a composite material structure according to claim 1. According to another aspect, there is provided a composite material structure for aircraft according to claim 3.

According to the disclosed embodiments, a method is provided for manufacturing composite material structures, particularly large scale composite material structures, which provides higher production speeds using lower cost equipment. Production time can be reduced by manufacturing a large-scale structure of modules that are manufactured individually and then joined together and cured together. The individual modules of the structure can be manufactured in parallel using equipment of the correct size which, taken collectively, can be capable of higher material placement speeds compared to conventional AFP machines. The described method also allows the use of equipment capable of handling multiple forms of materials that may be needed to satisfy the loading requirements in particular regions of the structure.

According to a disclosed example, there is provided a method of fabricating a composite material structure comprising: forming a plurality of composite material modules each having an edge; and, joining the modules along their edges. The modules can be joined using a bevel joint between the edges of the modules that can include one or more overlapping ramps. Adjacent modules can be joined together by multiple bevel joints that form a sawtooth pattern in which the modules include unidirectional reinforcing fibers having orientations other than 0 or 90 degrees.

According to another example of the disclosed method, manufacturing a composite material structure comprises: forming a plurality of multilayer composite material modules; assembling the composite material modules together, including forming bevel joints between at least certain modules; and, jointly curing the modules after the bevel joints have been formed. Bevel joints can be formed by abutting similar layers of adjacent modules or by overlapping similar layers.

According to yet another disclosed example, a composite aircraft structure comprises: a plurality of laminated modules of composite material each having edges; and, bevel joints to join the modules along their edges. Each of the modules includes multiple layers. The similar layers of the

Adjacent modules can be either abutted or overlapped at the bevel junctions. In a variation, the bevel joint can be a joint of interlocking wedges and the joints between certain adjacent modules can form a sawtooth pattern.

The embodiments of the disclosure satisfy the need for a method of fabricating composite material structures on a large scale using correctly sized equipment that represents a relatively low capital investment. The described embodiments also satisfy the need for a manufacturing method that is highly flexible and allows multiple modules of the structure to be formed in parallel.

1. Method for manufacturing a composite material structure comprising:

forming a plurality of composite material modules each having an edge; and, joining the modules together along their edges.

two. The method of claim 1, wherein joining the modules includes forming a bevel joint between the respective edges of the modules.

3. The method of claim 1, wherein the modules include prepreg fibers that form an angle of approximately 45 degrees with respect to an axis of orientation, and wherein:

joining the modules together includes forming a joint that defines a sawtooth pattern along the edges of the modules.

Four. The method of claim 3, wherein joining the modules together includes forming a bevel joint along the edges of the modules.

5. A method according to claim 1, wherein each of the modules includes at least two edges, and joining the modules together includes:

forming at least two bevel joints between the modules respectively along at least two edges.

6. The method of claim 1, wherein each of the modules includes a plurality of layers, and joining the modules together along their edges includes:

overlap at least certain layers along the edges.

7. The method of claim 1, wherein each of the modules includes a plurality of layers, and joining the modules together along their edges includes:

put the layers substantially butt along the edges.

8. The method of claim 1, wherein joining the modules together includes forming at least one bevel joint along the edges of the modules.

9. Large-scale aircraft structure manufactured by the method of claim 1.

10. Method of manufacturing a composite material structure, comprising:

forming a plurality of multi-layer, composite material modules;

assembling the composite material modules together, including forming bevel joints between at least certain composite material modules; and, jointly curing the composite modules after the bevel joints have been formed.

eleven. The method of claim 10, wherein forming bevel joints includes abutting adjacent layers of at least certain modules joined together.

12. The method of claim 10, wherein forming bevel joints includes substantially overlapping the layers of adjacent layers of at least certain modules joined together.

ES 2 638 263 T3

13. Method according to claim 10, wherein:

forming the multilayer composite modules includes providing fiber-reinforced composite layers having fiber orientations extending transversely to the edges of the modules such that the edges form a sawtooth pattern, and assembling the modules includes adjusting the sawtooth patterns of adjacent modules of modules joined together.

14. The method of claim 13, wherein assembling the modules includes staggering the modules relative to each other such that the saw tooth patterns are offset relative to each other.

fifteen. The method of claim 10, wherein assembling the modules together includes placing at least one of the modules on one of the bevel joints.

16. The method of claim 10, wherein assembling the modules includes positioning the modules relative to one another in a curing tool.

17. Aircraft composite material aircraft structure manufactured by the method of claim 1.

18. Composite material aircraft structure, comprising:

a plurality of composite laminate modules each having edges; and bevel joints to join the modules along their edges.

19. Composite material structure according to claim 18, in which:

each of the modules includes multiple layers, and similar layers of adjacent layers of the modules abut each other at the bevel joint.

twenty. Composite material structure according to claim 18, in which:

Each of the modules includes multiple layers, and similar layers of adjacent layers of the modules overlap at the bevel joint.

twenty-one. Composite material structure according to claim 18, in which the bevel joint is a joint of interlocking wedges.

22. Composite material structure according to claim 18, in which:

at least certain modules having edges joined together by the bevel joints each include at least one layer that includes reinforcing fibers having a common orientation, and the bevel joints joining these certain modules form a sawtooth pattern .

2. 3. Composite structure according to claim 18 further comprising:

at least one layer of composite material that is superimposed and attached to at least one of the bevel joints.

24. Large-scale composite material structure for aircraft, comprising:

a plurality of co-cured composite laminate modules each including multiple layers of fiber reinforced fiber resin, each of the modules including first and second edges extending transversely to each other, the edges of the layers of the adjacent layers of the modules overlap each other to form a bevel joint that joins the adjacent modules together, and at least certain adjacent modules that include layers having a fiber reinforcement oriented transversely to the edges and in which the first and second edges form a sawtooth pattern.

ES 2 638 263 T3

25. Method for manufacturing a large-scale composite structure for aircraft, comprising:

forming a plurality of composite modules in which each includes a plurality of layers of prepreg fibers in which at least the creation of the layers includes orientations of the fibers that extend transversely to the edges of the modules;

positioning the edges of the modules to form a sawtooth profile;

assembling the composite material modules together, including forming bevel joints between the modules and mating the sawtooth profiles of adjacent modules; place the modules in a curing tool; and, jointly curing the modules after the modules have been placed in the curing tool.

Brief description of the drawings

FIG. 1 is a schematic top view of a fabrication system that is suitable for the implementation of an illustrative embodiment of the modular composite fuselage skin fabrication method.

Figure 2 is a perspective view of a fabrication system that is suitable for the implementation of an illustrative embodiment of the modular composite fuselage skin fabrication method.

Figure 3 is an end view of a SADL (semi-automatic folder locator) machine, a pick and place machine, and a curing tool.

Figure 4 is a flow chart summarizing an illustrative embodiment of the modular composite fuselage skin fabrication method.

Figure 5 is a flow chart of an aircraft production and service methodology.

Figure 6 is a block diagram of an aircraft.

Figure 7 is a flow chart illustrating embodiments for fabricating composite material structures.

Figure 8 illustrates additional embodiments for fabricating composite material structures.

Figure 9 is a cross-sectional view of adjacently arranged composite material modules.

Figure 10 is a plan view illustrating a subsection of an aircraft fuselage skin formed from multiple modules in accordance with the disclosed embodiments.

Figure 11 is a perspective view of a fuselage section formed by subsections of the type illustrated in Figure 10.

Figure 12 is a cross-sectional view taken along line 12-12 of Figure 10.

Figure 13 is a schematic illustration showing the layers of the bevel joint shown in Figure 12.

Figure 14 is a cross-sectional view taken along line 14-14 of Figure 10.

Figure 15 is a schematic illustration showing the layers of the bevel joint of Figure 14.

Figure 16 is an enlarged illustration of the surface designated with an "A" in Figure 10.

Figure 17 is a cross-sectional view of another bevel joint according to an alternative embodiment.

Figure 18 is a schematic illustration showing the layers of the bevel joint of Figure 17.

Figure 19 is a perspective view of a composite module having sawtooth edges.

ES 2 638 263 T3

Fig. 20 is a cross-sectional view showing a bevel joint forming another embodiment.

Figure 21 is a schematic illustration illustrating the layers of the bevel joint of Figure 20.

Fig. 22 is a cross-sectional view showing another embodiment of the bevel joint.

Figure 23 is a schematic illustration showing the layers of the bevel joint of Figure 22.

Figure 24 is a plan view illustrating the offsets between the edges of the sawtooth joint between multiple modules.

Figure 25 is a flow chart roughly illustrating the steps of one embodiment of the described method. Detailed description

Referring initially to Figures 1-3 of the drawings, a manufacturing system is generally indicated by reference numeral 1 that is suitable for the implementation of an illustrative embodiment of the modular composite material manufacturing method. The manufacturing system 1 is shown in the top view of Figure 1 and in the perspective view of Figure 2. The modular composite fabrication method can utilize simple, correct-sized equipment that allows composite materials to be placed by parallel rather than serial processes as part of the fabrication of aircraft airframes or other composite parts. . The use of a parallel process technique in the automatic lamination of skin modules for aircraft fuselages or other parts can dramatically reduce the time required to manufacture a single part. This measure can reduce capital investment, factory floor space, and support personnel required to fabricate composite or other part aircraft fuselage skins. Furthermore, the method can be used in the manufacture of flat lay composite parts such as aircraft fuselage skin, for example, and without limitation, or contoured lay composite parts such as wing skin sheets and stabilizers. aircraft, for example, and without limitation, The method can be used to fabricate panels, quarter sections, half fuselage sections, more than half fuselage sections or full cylindrical sections.

As shown in Figures 1 and 2, the fabrication system 1 may include one or multiple flat web laying machines (FTLM) 2 to facilitate the fabrication of flat lay composite parts such as aircraft fuselage skin. , for example, and without limitations. Additionally or alternatively, the fabrication system 1 may include one or more contoured strip placement machines (CTLM) (not shown) to facilitate the fabrication of contoured placement composite parts such as aircraft wing skin, nose skin and / or glue coating, for example, and without limitation. The FTLM and CTLM machines can have a design that is known to those of skill in the art. Although an exemplary structure and method of operation of the FTLM machine 2 will be described hereinafter, it will be obvious to recognize and understand that the same method of operation can be used with respect to one or more CTLM machines in addition to or in place of 2 FTLM machines.

Each FTLM machine 2 may include a pair of separate, parallel, generally elongated frame rails 3. A carriage frame 4 may be extended and adapted to traverse the track rails 3 of the frame bi-directionally. The carriage frame 4 may include a pair of spaced, parallel, generally elongated carriage frame members 5. The carriage frame members 5 may be oriented in a generally perpendicular relationship with respect to the frame track rails 3.

A cutting carriage 6 may be adapted to traverse the carriage frame members 5 of the carriage frame 4 bi-directionally. A carriage motor (not shown) may be applied to the cutting carriage 6 to facilitate movement of the cutting carriage 6 on the carriage frame 4. A cutting device 7 can be arranged on the cutting carriage 6. In some embodiments, the cutting device 7 can be an ultrasonic blade, although alternative cutting implements that are known to those skilled in the art and are suitable for this purpose can be used.

A rotation track 10 of the transfer platform, which may be circular or annular, may be arranged between the frame track rails 3 and below the carriage frame 4. A transfer platform 14 may be removably disposed on the rotation track 10 of the transfer platform. The transfer platform 14 may have a generally square shape. The corner portions 14a of the transfer platform 14 may be slidably or rotatably engaged to the rotation track 10 of the transfer platform in the judgment of those skilled in the art by means of rollers (not shown), for example, and without limitation. . As will be described hereinafter, a transport sheet 16 may be arranged on the transfer platform 14. As used herein, "module" and "composite module" refer to sections of composite material that are joined together to form a

ES 2 638 263 T3 larger structure, and may be, but are not necessarily limited to, a single or multiple set formed from fiber tows or prepreg fabrics. A composite module 18 may be arranged on the transport sheet 16. The transfer platform 14 can be rotated on the transfer platform rotation track 10, the carriage frame 4 can move along the frame track rails 3, and the cutting carriage 6 can move along the Carriage frame members 5 of carriage frame 4 to facilitate cutting of the composite material module 18 along a selected axis or axes by operation of the cutting device 7.

As shown in Figures 1-3, the manufacturing system 1 may further include a SADL (semi-automatic folder locator) machine 22, The SADL machine 22 may include a pedestal 23 that can be made portable by multiple pedestal wheels 24. The module-forming platform 25 may be arranged on the pedestal 23. The module-forming platform 25 may be adapted to receive and support a carrier sheet 16 on which a composite material module 18 is disposed for purposes to be described hereinafter.

As will be further described in Figures 1-3, the manufacturing system 1 may further include a pick and place machine 30. As shown in Figures 1 and 2, the pick and place machine 30 may be located generally adjacent to the SADL machine 22. As further shown in FIG. 1, in some embodiments a SADL machine 22 may be disposed at or is generally adjacent to respective ends of the pick and place machine 30. The pick and place machine 30 may include a trestle 31 having a pair of spaced, parallel, generally elongated rails 32. The rails 32 of the trestle 31 may each be supported by multiple separate rail supports 33 as shown in Figure 3. At least one carriage 36 with a positioning head may extend and slidably engage the rails 32 of the trestle 31. Each carriage 36 with locating head may extend and slidably engage rails 32 of trestle 31. Each positioning head carriage 36 may be adapted to bi-directional movement along the rails 32, as indicated by the double-headed arrow 8 in Figure 1. A carriage motor (not shown) can be applied to each positioning head carriage 36 to facilitate the movement of the carriage 36 with the positioning head along the rails 32.

As shown in Figures 2 and 3, a module placement head 40 may be suspended from each placement head carriage 36. The module positioning head 40 may include a head shaft 41 and a generally curved or arcuate module engaging member 42. The head shaft 41 of the module placement head 40 may be engaged to the placement head carriage 36 using any appropriate technique known to those skilled in the art. In some embodiments, at least one head mounting flange 37 extends from the positioning head carriage 36. At least one module attachment bracket 44 extends from head shaft 41. The module attachment bracket 44 may be connected to at least one head mounting flange 37 by means of a head attachment member 38.

As further shown in Figures 2 and 3, module engagement module 42 of module placement head 40 may include a generally convex module-forming surface 43. A scanner 49 of an inspection scanning system 46 may be adapted to traverse the module-forming surface 43 of the module application member 42. Scanner 49 may be attached to application member 42 of the module in accordance with the experience of those skilled in the subject art. In some embodiments, a generally elongated, curved scanner slot 47 may be disposed in the module engaging member 42 generally adjacent to and along the module-forming surface 43. A scanner holder 48 may be applied to the scanner slot 47 to traverse the scanner slot 47. The scanner 49 may be arranged on the scanner holder 48. A scanner motor (not shown) may be applied to the scanner holder 48 to facilitate selective movement of the scanner holder 48 in the scanner slot 47 and the scanner 49 along and adjacent to the member module forming surface 43. 42 of module application. An inspection scan and control system (not shown) may be connected to the scanner motor (not shown) and to the scanner 49 to facilitate the scanning movement of the scanner 49 and to collect and analyze images received from the scanner 49.

The manufacturing system 1 may further include a curing tool 54, mandrel or mold. The curing tool 54 may be an OML (outer mold line) or an IML (inner mold line) curing tool, as an example and without limitation. As shown in Figures 1, 2 and 3, the curing tool 54 may be located generally adjacent to the SADL machine 22 and between the rails 32 of the pick and place machine 30. As shown in FIG. 3, in some embodiments the curing tool 54 may include a tool base 55 and separate, generally parallel tool sides 56 extending from the tool base 55. A generally curved or semi-circular module positioning surface 57 may be provided at the base 55 of the tool and on the sides 56 of the tool and may extend along the length of the curing tool 54. However, it will be apparent to recognize and understand that curing tool 54 (such as IML curing tools, for example) does not necessarily have a fully cylindrical or semi-cylindrical cross section as shown with respect to surface 57 module placement of the curing tool 54. In circumstances where

In ES 2 638 263 T3 it is desired to use the quarter panel manufacturing method, for example, and without limitation, curing tools 54 having both an OML configuration and an IML configuration can be used. In addition, the curing tool 54 can be configured as a wing or stabilizer mold, tool, curing tool or in any configuration depending on the part to be manufactured.

In the typical implementation of the modular composite fabrication method, the method can be used to fabricate an aircraft fuselage skin 60 (FIG. 1) using the multiple composite modules 18. Depending on the application and the part to be manufactured, each module 18 can include any combination of unidirectional carbon fiber prepreg; carbon fiber prepreg fabric; fiberglass; KEVLAR ® poly (p-phenylene terephthalmide); or other materials. Each module 18 can have at least one layer. A carrier sheet 16, on which a composite material module 18 can be deposited, can be initially arranged on a transfer platform 14. The transfer platform 14 may be arranged on the rotation track of the annular transfer platform 10 of an FTLM machine 2. The cutting device 7 of the cutting carriage 6 can be operated to trim or cut the composite material module 18 to the desired dimension and shape. During trimming, cutting, or modulating operation, the composite material module 18 can be positioned in selected orientations relative to the cutting device 7 by moving the cutting carriage 6 along the carriage frame members 5 of the carriage frame. 4; by movement of the carriage frame 4 along the frame track rails 3; and / or by rotation of the transfer platform 14 on the transfer platform rotation track 10. Each FTLM 2 can facilitate high speed modulation of the composite modules 18 that are to form the aircraft fuselage skin 60 using clean cut technology.

After the composite module 18 is trimmed or cut, the transfer deck 14, on which the conveyor sheet 16 and the trimmed or cut composite module 18 are arranged, can be removed from the deck rotation track 10. transfer. The transfer platform 14 can be transported from the FTLM machine 2 to one of the SADL machines 22. The removal of the transfer platform 14 from the transfer platform rotation track 10 and / or the transport of the transfer platform 14 may be automatic or manual. In the SADL machine 22, the transport sheet 16 can be removed from the transfer platform 14 and disposed on the forming platform 25 of the SADL machine 22 module. The transport sheet 16 may have multiple openings for tools / indexes (references), (not shown), which can be indexed to the SADL machine 22 to facilitate proper positioning and positioning of the transport sheet 16 on the forming platform 25. of the module.

The placement head carriage 36 may then be operated to slide along the rails 32 of the trestle 31 of the pick and place machine 30 to arrange the module engaging member 42 of the module placement head 40. directly onto the composite module 18. The module-forming platform 25 of the SADL machine 22 can then be raised against the module-forming surface 43 of the module-engaging member 42 to form or contour the flat composite module 18 following the generally convex contour of the surface 43 modulus former, as indicated by the dashed line in Figure 3, by means of a modulus compaction process. The formation and contouring of the module 18 following the module-forming surface 43 of the module-engaging member 42 may be automatic. Additional composite modules 18 may be transported from FTLM machine 2 to SADL machine 22 module forming platform 25 and formed according to module forming surface 43 of module applying member 42 in a laminated or multilayer manner according to is needed to achieve a desired thickness of the aircraft fuselage skin 60 (FIG. 1). Accordingly, successive composite modules 18 may be arranged one on top of the other to form a laminated module 18 having multiple layers. Depending on the application, the module placement head 40 may arrange continuous fiber layers or continuous fabric layers together with woven layers of the composite modules 18. In some applications, adjacent modules 18 may be applied to each other, as will be discussed in more detail below. It will be apparent to recognize and understand that the composite modules 18 need not always be formed on the module-forming surface 43 of the module engaging member 42 on the module placement head 40. In the case of an IML cure tool 54, modules 18 can be formed directly into cure tool 54 or other coating layers using SADL machine 22.

After the desired number of composite material modules 18 have been arranged in the module placement head 40 of the pick-and-place machine 30 module, the scanner 49 of the inspection scanning system 46 can be operated to traverse the surface. 43 module former 43 module application member for the purpose of inspecting the composite modules 18. The defective composite modules 18 can be removed from the module placement head 40 and replaced with non-defective composite modules 18. Inspection of modules 18 in module placement head 40 may be an automatic process.

The carriage 36 with the placing head of the pick-and-place machine 30 can then be operated to traverse the rails 32 of the trestle 31 of the pick-and-place machine 30 and facilitate placement.

ES 2 638 263 T3 of the stacked, laminated, compressed and inspected composite modules 18 at the desired location on the module placement surface 57 (FIG. 3) of the curing tool 54.

Additional stacked, laminated, and compressed modules 18 can be similarly formed and arranged at desired locations on the module placement surface 57 to form the aircraft fuselage skin 60. Module edges 19 of adjacent modules 18 may be joined and overlapped with a bevel (not shown) or joined by a splice (not shown) until the aircraft fuselage skin 60 is fully disposed.

Referring now to the flow chart 300 of FIG. 4, an illustrative embodiment of a composite material manufacturing method is summarized therein. The method can be used to fabricate an aircraft fuselage skin having a desired thickness, for example, and without limitation. In block 302 a composite material module can be provided. The module can be an aircraft fuselage skin module that can be used to fabricate a structure such as an aircraft fuselage skin, for example, and without limitation, and can be arranged using an FTLM (flat web laying machine). For example, and without limitation, additionally or alternatively, the module can be used to fabricate a structure such as a wing skin, nose liner or aircraft tail liner, for example and without limitation, in which case the module can be arranged using a CTLM (Contoured Strip Laying Machine). At block 304, the module can be transferred to a SADL (semi-automatic folder locator) machine. At block 306, the module can be shaped to contour to an aircraft fuselage skin or other structure. In block 308, the steps taken in blocks 302, 304, and 306 can be repeated to form a desired thickness of the fuselage skin for aircraft or other structure. At block 310, the modules can be inspected. At block 312, the stacked or laminated modules can be stacked on a curing tool. The curing tool can be an OML (outer mold line) or IML (inner mold line) curing tool, for example, and without limitation. In block 314, the steps taken in blocks 302, 304, 306, 308, 310 and 312 can be repeated as necessary to fully lay out the skin of the fuselage for aircraft or other structure.

Referring now to Figures 5 and 6, embodiments of the description may be used in the context of a method 78 of manufacturing and servicing an aircraft as shown in Figure 5 and an aircraft 94 as shown in FIG. 6. During pre-production, sample method 78 may include specification and design 80 of aircraft 94 and acquisition of material 82. During production, the manufacturing of the components and sub-assemblies 84 and the integration of the systems 86 of the aircraft 94 takes place. From this moment, the aircraft 94 can pass certification and delivery 88 to be put into service 90. While in service In service with a buyer, routine maintenance and service 92 (which may also include modification, reconfiguration, renovation, and so on) of aircraft 94 can be scheduled.

Each of the processes in method 78 can be performed or carried out by a systems integrator, a third party, and / or an operator (eg, a buyer). For the purposes of this description, a systems integrator may include without limitation any number of aircraft manufacturers and major systems subcontractors; a third party may include without limitation any number of vendors, subcontractors and suppliers; and an operator can be an airline, a leasing company, a military entity, a service organization, and others.

As shown in Figure 6, aircraft 94 produced by method example 78 may include an aircraft cell 98 with a plurality of systems 96 and an interior 100. Examples of high-level systems 96 include one or more systems 102 propulsion system, an electrical system 104, a hydraulic system 106, and an environmental system 108. Any number of other systems can be included. Although an example from the aerospace industry is shown, the principles of the invention can be applied to other industries, such as the automotive industry.

The apparatus incorporated herein may be employed during any one or more stages of the production and service method 78. For example, the components or sub-assemblies corresponding to the production process 84 may be manufactured or manufactured in a similar manner to the components or sub-assemblies produced while the aircraft 94 is in service. One or more fixture additions may also be performed during production steps 84 and 86, for example, substantially expediting assembly, or reducing cost, of an aircraft 94. Similarly, one or more fixture incorporations may be performed. apparatus while aircraft 94 is in service, for example, and without maintenance and service limitations 92.

As mentioned above, the modular composite fabrication method can utilize equipment of the correct size to allow the placement of composites in parallel rather than serial processes as part of the fabrication of aircraft airframes or other parts of aircraft. composite material. In this regard, reference is made to Figures 7 and 8. According to a number of embodiments, a method 700 for manufacturing a

ES 2 638 263 T3 composite material structure may include arranging 702 a plurality of modules 18 in tool 54 such that each of modules 18 is adjacent to at least one other of modules 18, as shown in the figure 9 and so that adjacent modules 18 can be joined together, such as at a joint 704, to form a composite material structure.

In some of the embodiments, more than one module 18 may be disposed in tool 54 at substantially the same time; In such embodiments, a plurality of transfer platforms 14 may be provided. In addition, in the formation of the composite material structure, the modules 18 can be arranged in the tool 54 in a substantially sequential manner, with a procedure 706 being performed after the placement in one of the modules 18 that has already been arranged in the tool. 54 (as indicated by the n-2 notation of Figure 8) while the next module 18 is being arranged in the tool 54 (as indicated by the n-1 notation of Figure 8). In many embodiments, the placement step 702 and the post-placement procedure step 706 may be performed while a succeeding module 18 is being prepared 708 for placement (as indicated by the notation n of FIG. 8).

With respect to performing a post-setup procedure 706, this procedure may include inspecting 710 one of the modules 18 that has already been disposed in the tool 54 (as indicated by the n-2 notation of FIG. 8). Embodiment 706 of a post-placement procedure may also include working on the joint 704 formed between adjacent modules 18. In some of the embodiments, such as those shown in Figures 7 and 8, the manufacturing method 700 may include performing, for example, the inspection steps 710 and post-module placement procedure step 706 18 differently. arranged substantially at the same time.

In many embodiments, positioning 702 of a module 18 in a tool 54 may include loading 712 a module 18 onto a transfer platform 14, indexing 714 the loaded module in an appropriate position, and / or then disposing 716 the module in a tool 54. After all modules 18 have been arranged in tool 54 and any subsequent post-placement processing has been performed, the framework can be cured.

Referring now to Figures 10 and 11, composite modules 208 of the general type previously described can be joined together along their mutual edges 210 to form skin 202 of a large-scale structure such as a fuselage section. of an aircraft with a cylindrical shape 204. As described in more detail below, each of the modules 208 may comprise one or more laminated layers of unidirectional or bi-directional fiber reinforced composite material and may include trim (not shown) and / or reinforcements (not shown) and / or profiles (not shown) used in forming features, such as a door 206 in the fuselage skin 202 shown in Figure 11.

Referring now to Figures 12-15, Adjacent modules of modules 208 may be joined along their mutual edges 210 by bevel joints, examples of these joints are designated by numbers 213a, 213b. As used herein, "bevel joint" refers to a joint between two pieces of material made by beveling or chamfering their ends, edges, or sides so that when the parts are arranged together they have overlapping edges that substantially form a member. continuous. As shown in Figure 13, each of two adjacent modules 208a, 208c joined by a bevel junction 213a may comprise multiple layers 212 of unidirectional or bi-directional prepreg fibers sandwiched between top and bottom surface sheets 216, 218 respectively. Each of the outer sheets 216, 218 may comprise fabric or other sheet materials. The bevel junction 213a is formed by arranging the layers 212 of the two modules 208a, 208c such that the outer ends 214 of the similar layers of the modules 208a, 208c overlap each other. The bevel joint 213a represents a simple bevel with a ramp with a slope of 40: 1 (ie, a length / height ratio) comprising a total of 12 overlapping layers 212; ramps with other ratios are possible, depending on the application.

A particular cladding subsection 200 (Figure 10) may comprise modules 208 joined together along their mutual edges 210 by more than one type of bevel joints 213, and in fact a particular module 208 may be joined along its edges 210 to adjacent modules 208 by different types of bevel joints. For example, one edge 210a of module 208a shown in Figure 10 may be attached to module 208c via the bevel joint 213a shown in Figures 12 and 13, while another edge 210b of module 208a may be attached to module 208b by another form of bevel joint 213b shown in Figures 14 and 15. As shown in Figures 14 and 15, the bevel joint 213b is a simple bevel configuration comprising 12 layers 212 arranged with a ramp with an 80: 1 slope in which the outer edges 214 of similar layers 212 they are arranged substantially in the same plane, and abut each other; Other ramp relationships are possible, depending on the application.

Referring now to Figure 10 and Figures 16-18, some of the modules 208 such as modules 208d and 208e may comprise layers 212 in which the angle of orientation of the reinforcing fibers extends transversely to edges 210. In the illustrated example, the direction of orientation of the fibers is

ES 2 638 263 T3 designated by the number 215 in figure 16 and comprises 45 degrees. To form a bevel joint 213c between adjacent modules 208d and 208e, edges 210 are formed in a zigzag or sawtooth pattern 220 that is symmetrical about an axis 221 of the joint (FIG. 16). The sawtooth pattern 220 is produced by beveling the edges 210 of the modules 208d, 208e in two orthogonal directions. In the illustrated example, as best seen in Figure 18, the bevel junction 213c comprises 12 layers 212 arranged in a ramp with an 80: 1 slope in which similar layers abut, rather than overlap each other. other. Ratios other than 80: 1 are possible, depending on the application. Figure 19 illustrates a typical module 208 having sawtooth edges 220 in which the sawtooth patterns are offset from layer to layer.

Figures 20 and 21 illustrate a set of modules 219 in which a bevel joint 213d is used to join two modules 228a, 228b, each consisting of multiple layers 212 in which similar layers overlap at their edges 217. A Third module 226 comprises multiple layers 227 that are disposed on bevel junction 213d. The embodiment of Figures 20 and 21 illustrates that the bevel junction 213d may be staggered through subsection 200 (Figure 10) and may not be continuous through the entire thickness of subsection 200. Reference is now made to the Figures 22 and 23 illustrating another embodiment of a set of modules 229 employing a bevel joint of the interlocking wedge joint 213e type formed by multiple bevels 221. The bevel joint 213e extends through the full thickness of the module assembly 229 and joins multiple adjacent modules 230-236. In this example, the set of modules 229 comprises twelve layers 212 including the top and bottom surface sheets 216, 218, in which each of the bevels 221 has a ramp with a slope of 80: 1; ramps with other ratios are possible, depending on the application. In this example too, similar layers 212 of adjacent modules 230-236 are abutted, rather than overlapping each other.

Reference is now made to FIG. 24 which illustrates a laminated structure 238 made up of multiple modules 208 that are joined together along bevel joints (not shown) that form sawtooth patterns 240. In this example, the patterns Sawtooth 240 are arranged in two groups 240a, 240b. The sawtooth patterns 240 are staggered relative to each other so that the phases of the two groups 240a, 240b are offset by a distance "x" so that adjacent sawtooth patterns 240 are not aligned with each other. Yes. This staggering of patterns 240 on adjacent saw teeth can improve the structural properties of the laminated structure 238.

Reference is now made to FIG. 25 which outlines the broad steps of a method for fabricating structures using modules 208 previously described. Beginning at step 246, modules 208 are formed which may comprise one or multiple layers of composite material. Next at 248, modules 208 are assembled using one of several types of bevel joints 213 discussed above. Following assembly, modules 208 may be arranged on or in a cure tool in step 250 which may comprise an inner mold line or outer mold line cure tool. As previously described in connection with Figures 1-3, modules 208 can be assembled by sequentially arranging them in the cure tool or by assembling groups of modules 208 and then arranging the groups in the cure tool. Finally, at step 252, the assembled modules 208 are cured together on the curing tools, causing the flow of resin through the bevel joints 213 to form a substantially homogeneous consolidated structure.

Although the embodiments of this disclosure have been described with respect to certain exemplary embodiments, it will be apparent that the specific embodiments are for illustration and not limitation, as are other variations that may occur to those skilled in the art. technique.

Also disclosed is a method of fabricating an exemplary composite structure, comprising forming a plurality of composite modules each having an edge, and joining the modules together along their edges.

Optionally, joining the modules includes forming a bevel joint between the respective edges of the modules.

Optionally, the modules include prepreg fibers that form an angle of approximately 45 degrees with respect to an axis of orientation, and joining the modules together includes forming a joint that defines a sawtooth pattern along the edges of the modules. .

Optionally, joining the modules together includes forming a bevel joint along the edges of the modules.

Optionally, each of the modules includes at least two edges, and joining the modules together includes forming at least two bevel joints between the modules respectively along the at least two edges.

Optionally, each of the modules includes a plurality of layers, and joining the modules together along their edges includes overlapping at least certain of the layers along the edges.

ES 2 638 263 T3

Optionally, each of the modules includes a plurality of layers, and joining the modules together along their edges includes substantially abutting the layers along the edges.

Optionally, joining the modules together includes forming at least one bevel joint along the edges of the modules.

A large-scale aircraft structure manufactured by the exemplary method is also disclosed.

A composite material structure for aircraft is also disclosed, comprising a plurality of laminated composite material modules each having edges, and bevel joints for joining the modules along their edges.

Optionally, each of the modules includes multiple layers, and like layers in the adjacent layers of the 10 modules abut each other at the bevel joint.

Optionally, each of the modules includes multiple layers, and similar layers in adjacent layers of modules overlap at the bevel joint.

Optionally, the bevel joint is an interlocking wedge joint.

Optionally, at least certain of the modules having edges joined together by the bevel joints each include at least one layer that includes reinforcing fibers having a common orientation, and the bevel joints joining the certain modules form a sawtooth pattern.

Optionally, there is at least one layer of composite material that overlaps and is attached to at least one of the bevel joints.