Composite manufacturing method using an assembly of composite modules

ES 2 628 443 T3

DESCRIPTION

Composite manufacturing method using a set of composite modules

Technical field

Disclosure generally refers to aircraft production. The disclosure, more particularly, relates to manufacturing methods and apparatus that use composite modules to manufacture components and structures.

Background

Some modern commercial aircraft components, such as the skin of an aircraft fuselage, for example, can be manufactured using advanced fiber placement (AFP) machines. The benchmark process for the use of AFP machines for aircraft manufacturing processes may require a high capital investment in the purchase of various AFP machines, tools, factory floor space and personnel. At current AFP deposition rates (eg, approximately 6.8 kg / hr), a large number of AFP machines may be required to place the amount of material needed to support acceptable commercial aircraft production rates.

Therefore, a manufacturing method is needed that can have a relatively high rate capacity and low capital investment requirements, and that does not require skilled operators and programmers to implement.

Document WO 2006 118629 refers to a system for manufacturing a composite article that includes an end effector, a robotic vehicle, a sensor and a cutting system.

WO 2004 078641 discloses a link between compounds with incompatible properties and a method for preparing such links. The composites comprise fibers of the first and second type, respectively, as well as resin. The connection comprises a transition zone between the compounds having a layered structure. The transition zone may optionally comprise a transition member and the transition member may optionally be composed of one or more of the compounds.

The present invention is set forth in the Independent claims, with some optional features set forth in the claims dependent thereon.

Add it up

The disclosure is generally directed to a composite manufacturing method. An Illustrative embodiment of the method includes providing a plurality of composite modules; Inspect the composite modules; provide a curing tool; and placing the composite modules on the curing tool.

The method provides a high rate of production and requires only a lower capital investment. On the other hand, the manufacturing method may not require skilled operators and programmers for its implementation.

Optionally, said curing tool comprises an outer mold tool line.

Optionally, said curing tool comprises an Inner line of molding tool.

In another embodiment, an aircraft manufactured by the modular composite manufacturing method is provided.

In yet another embodiment, a method of manufacturing a composite structure is provided, the method comprising:

provide a tool;

provide a plurality of modules; and place the modules in the tool in such a way that:

each of the modules is adjacent to at least one other of the modules; Y

ES 2 628 443 T3 that the adjacent modules are linkable to each other to form a composite structure.

Optionally, said positioning step comprises placing more than one module on the tool at substantially the same time.

Optionally, the placement step comprises placing the modules on the tool in a sequential manner, the method further comprising:

perform a post-placement procedure on one of the modules that has been placed on the tool while a subsequent module is being placed on the tool.

Optionally, said stage of execution includes inspecting one of the modules that has been placed in the tool.

Optionally, said stage of execution includes working in a joint formed between adjacent modules. Optionally, the method further comprises curing the composite structure.

In another embodiment, a method of manufacturing a composite structure is provided, the method comprising:

prepare a module to be stacked on a tool; and placing in the tool a module that has already been prepared in accordance with said preparation step;

said preparation and stacking steps can be executed substantially at the same time.

Optionally, said placement step comprises placing a plurality of modules on the tool such that at least two of the modules can be placed on the tool at substantially the same time.

Optionally, said placement step comprises placing a plurality of the modules on the curing tool in a sequential manner, the method further comprising:

executing a post-placement procedure on at least one of the modules that has been placed on the tool while a subsequent module is being placed on the tool;

said post-placement procedure comprises an inspection procedure;

wherein said stage of execution and said stage of placing a subsequent module can be executed at substantially the same time.

Optionally, said stage of execution includes working on a joint formed between modules positioned adjacently.

In a further embodiment, a method of manufacturing a composite structure using a plurality of prefabricated modules is provided, the method comprising: preparing one of the modules for stacking;

stack one of the prepared modules; and performing a post-stacking procedure on one of the stacked modules;

wherein said preparation, stacking and carrying out steps are performed substantially at the same time.

Optionally, the method further comprises the steps of:

stacking a plurality of modules such that at least two of the modules are stacked substantially at the same time or in a sequential manner; and executing an inspection procedure on at least one of the plurality of modules that have been stacked, or a

ES 2 628 443 T3 method of joining between modules stacked adjacently while stacking a subsequent module in which said stage of execution and said stage of stacking of a subsequent module are executed substantially at the same time.

Brief description of the illustrations

Figure 1 is a schematic top view of a manufacturing system that is suitable for the implementation of an illustrative embodiment of the modular composite airframe skin manufacturing method.

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

Figure 3 is an end view of a SADL (semi-automated doubler locator) machine, a pick and locate machine, and a curing tool.

FIG. 4 is a flow chart summarizing an illustrative embodiment of the modular composite airframe skin manufacturing method.

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

Figure 6 is a block diagram of an aircraft.

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

Figure 8 illustrates additional embodiments for the fabrication of composite structures.

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

Detailed description

Referring initially to Figures 1-3 of the drawings, it is generally indicated by reference numerals

1 a manufacturing system that is suitable for the implementation of an illustrative embodiment of the modular composite manufacturing method. The manufacturing system 1 is shown in top view in figure 1 and in perspective view in figure 2. The modular composite fabrication method can use simple equipment, of the correct size to allow the placement of composite materials in parallel processes rather than in series as part of the fabrication of composite aircraft fuselages or other parts. Using a parallel process approach to automated lamination of aircraft fuselage skin modules or other parts can dramatically reduce the flow time required to fabricate a single part. This can reduce the capital investment, manufacturing footprint, and support staff required to manufacture composite aircraft fuselage skins and other parts. Furthermore, the method can be used in the manufacture of composite stacked flat parts such as aircraft fuselage skins, for example and without limitation or composite stacked contoured parts such as aircraft wing skins and stabilizers, for example and without limitation. The method can be used to fabricate panels, quarter sections, half fuselage sections, sections of more than half fuselage or whole barrel sections.

As shown in Figures 1 and 2, the manufacturing system 1 may include one or multiple tape flat stacking machines (FTLM) 2 to facilitate the manufacture of composite stacked flat parts such as aircraft fuselage skins, for example. and without limitation. Additionally or alternatively, fabrication system 1 may include one or more Contour Type Layup Machines (CTLMs) (not shown) to facilitate the fabrication of composite stacked contoured parts such as aircraft wing skin. , the nose liner and / or the tail liner, for example and without limitation. The FTLM and CTLM may have a design known to those of skill in the art. Although an exemplary structure and method of operation for the FTLM 2 will be described below, it will be recognized and understood that the same method of operation may be used with respect to one or more CTLMs in addition to or in place of the FTLM 2.

Each FTLM 2 may include a pair of generally elongated frame rail rails 3 parallel and spaced apart. A carriage frame 4 can span and adapt to traverse bi-directionally the rail rails 3 of the frame. The carriage frame 4 may include a pair of generally elongated, parallel, and spaced apart carriage frame members 5. The members 5 of the carriage frame can be oriented in generally perpendicular relation with respect to the rail rails 3 of the frame.

A cutting carriage 6 may be adapted to traverse bi-directionally the frame members 5 of the carriage of the

ES 2 628 443 T3 carriage frame 4. A carriage motor (not shown) can be coupled with the cutting carriage 6 to facilitate movement of the cutting carriage 6 on the carriage frame 4. A cutting device 7 can be provided on cutting carriage 6. In some embodiments, cutting device 7 may be an ultrasonic knife, although alternative cutting instruments known to those skilled in the art and suitable for this purpose may be used.

A rotation rail 10 of the transfer deck, which may be circular or annular, may be provided between the rail rails 3 of the frame and below the carriage frame 4. A transfer deck 14 may be removably provided on the rotation rail 10 of the transfer platform. The transfer platform 14 can generally have a square shape. The corner portions 14a of the transfer platform 14 can slideably or rotatably engage with the rotation rail 10 of the transfer platform according to the knowledge of those skilled in the art, such as through rollers (not shown), for example and without limitation. As will be described below, a carrier sheet 16 can be placed on the transfer platform 14. A composite module 18 can be placed on the carrier sheet 16. The transfer deck 14 can be rotated on the rotation rail 10 of the transfer deck, the carriage frame 4 can be moved along the rail rails 3 of the frame, and the cutting carriage 6 can be moved along along the carriage frame members 5 of the carriage frame 4 to facilitate cutting of the composite module 18 along a selected axis or axes by operating the cutting device 7.

As shown in Figures 1-3, the manufacturing system 1 may also include a SADL (Semi-Automated Doubler Locator) machine 22. The SADL machine 22 may include a pedestal 23 that can be made portable by means of multiple pedestal wheels 24. A module forming platform 25 can be provided on the pedestal 23. The module-forming platform 25 can be adapted to receive and support a carrier sheet 16 on which a composite module 18 is laid for the purposes that will be described below.

As further shown 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 each other. to the SADL machine 22. As further shown in FIG. 1, in some embodiments a SADL machine 22 may be provided at or generally adjacent to the respective ends of the pick and place machine 30. The pick and place machine 30 may include a gantry crane 31 having a pair of rails 32 generally elongated, parallel and spaced apart. Each rail 32 of the gantry crane 31 may be supported by multiple rail supports 33 spaced apart from each other as shown in Figure 3. At least one placement head carriage 36 may span and slidably engage with the rails 32 gantry crane 31. Each positioning head carriage 36 can be adapted to move bi-directionally on the rails 32, as indicated by the double-headed arrow 8 in Figure 1. A carriage motor (not shown) can be fitted with each head carriage 36 of positioning to facilitate movement of the positioning head carriage 36 on the rails 32.

As shown in Figures 2 and 3, a module placement head 40 can be suspended from each placement head carriage 36. The module placement head 40 may include a head stem 41 and a generally curved or arcuate module engagement member 42. The head stem 41 of the module placement head 40 may be attached to the placement head carriage 36 using any suitable 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 clamp 44 extends from the head stem 41. The module clamp 44 may be connected to at least one head mounting flange 37 by means of a head clamp member 38.

As further shown in Figures 2 and 3, the module engagement member 42 of the module placement head 40 may include a generally convex module-forming surface 43. A scanner 49 of a scan inspection system 46 can be adapted to traverse the module forming surface 43 of the module engagement member 42. Scanner 49 may be attached to module engagement member 42 to the knowledge of those of ordinary skill in the art for this purpose. In some embodiments, a generally elongated and curved scanner slot 47 may be provided in the module engagement member 42 generally adjacent to and along the module-forming surface 43. A scanner clamp 48 may engage with the scanner slot 47 to pass through scanner slot 47. Scanner 49 can be provided on scanner clamp 48. A scanner motor (not shown) can mate with scanner clamp 48 to facilitate selective movement of scanner clamp 48 into scanner slot 47 and scanner 49 along, and adjacent to, the modulus forming surface 43 of the modulus engagement member 42. An inspection and control analysis system (not shown) can be connected to the scanner motor (not shown) and to the scanner 49 to facilitate the scanning movement of the scanner 49 and to retrieve and analyze images received from the scanner 49.

Fabrication system 1 may further include a curing tool, mandrel, or mold 54. Curing tool 54 can be an OML (Outer Mold Line) or curing tool.

ES 2 628 443 T3 an IML (Inner Mold Line), for example and without limitation. As shown in Figures 1, 2, and 3, the curing tool 54 can be positioned generally adjacent to the SADL machine 22 and between the rails 32 of the pick and place machine 30. As shown in Figure 3, in some embodiments the curing tool 54 may include a tool base 55 and spaced apart generally parallel tool sides 56 extending from the tool base 55. A module placement surface 57 Generally curved or semi-circular can be provided on tool base 55 and tool sides 56 and can extend along the length of curing tool 54. However, it will be recognized and understood that curing tool 54 (such as in the case of IML curing tools, for example) need not necessarily have an entirely cylindrical or semi-cylindrical cross section as shown with respect to the placement surface. modulus 57 of curing tool 54. In circumstances where it is desired to use the quarter panel fabrication method, for example and without limitation, curing tools 54 having both an OML and IML configuration could be used. In addition, the curing tool 54 can be configured as a wing or stabilizer mold, a tool, a curing tool, or 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 (Figure 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 prepreg unidirectional carbon fiber; prepreg carbon fiber fabric; fiberglass; KEVLAR® poly (p-phenylene terephthalamide); or other materials. Each module 18 can have at least one layer. A carrier sheet 16, on which a composite module 18 can be laid, can initially be placed on a transfer platform 14. The transfer deck 14 can be positioned on the rotating annular rail 10 of the transfer deck of an FTLM 2. The cutting device 7 on the cutting carriage 6 can be operated to trim or cut the composite module 18 at desired dimensions and shape. During the trimming, cutting, or modulating operation, the composite 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; the movement of the carriage frame 4 along the rail rails 3 of the frame; and / or rotation of the transfer platform 14 on the rotation rail 10 of the transfer platform. Each FTLM 2 can facilitate high-speed modulation of composite modules 18 that are to form the skin 60 of an aircraft fuselage using network clipping technology.

After trimming or cutting of the composite module 18, the transfer deck 14, on which the carrier sheet 16 and the trimmed or cut composite module 18 are laid, can be removed from the rotation rail 10 of the transfer deck. The transfer platform 14 can be transported from the FTLM 2 to one of the SADL machines 22. The removal of the transfer platform 14 from the rotation rack 10 of the transfer platform and / or the transport of the transfer platform 14 can be automated or manual. In the SADL machine 22, the carrier sheet 16 can be removed from the transfer platform 14 and placed on the module-forming platform 25 of the SADL machine 22. Carrier sheet 16 may be provided with multiple indexing / tool openings (not shown) that can be indexed on SADL machine 22 to facilitate proper positioning and positioning of carrier sheet 16 on module-forming platform 25.

The placement head carriage 36 may then be operated to slide along the rails 32 on the gantry crane 31 of the pick and place machine 30 to position the module engagement member 42 of the placement head. module 40 directly onto 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 engagement member 42 to form or contour the composite module 18 flat to the generally convex contour of the module-forming surface. 43, as indicated by the dotted line in Figure 3, by a modulus compaction process. Forming and contouring of the module 18 on the module-forming surface 43 of the module engagement member 42 can be automated. Additional composite modules 18 can be transported from the FTLM machine 2 to the module-forming platform 25 of the SADL machine 22 and formed on the module-forming surface 43 of the module engagement member 42 in a laminated or multilayer fashion when you need to achieve a desired thickness of the skin 60 of the fuselage of an aircraft (Figure 1). Accordingly, successive composite modules 18 can be placed on top of each other to form a laminated module 18 having multiple layers. Depending on the application, the module placement head 40 can place continuous fiber layers or continuous fabric layers in conjunction with woven layers of the composite modules 18. In some applications, adjacent modules 18 can be coupled to each other according to knowledge. from experts in the field. It will be recognized and understood that the composite modules 18 need not always be formed on the module-forming surface 43 of the module engagement member 42 on the module placement head 40. In the case of an IML curing tool 54, the modules 18 can be formed directly on curing tool 54 or other coating layers using SADL machine 22.

After the desired number of composite modules 18 have been arranged on the module placement head 40 of the pick and place machine 30, the scanner 49 of the scan inspection system 46

ES 2 628 443 T3 can be operated to traverse the module forming surface 43 of the module engagement member 43 for the purpose of inspecting the composite modules 18. The defective composite modules 18 can be removed from the module placement head 40 and can be replaced with 18 non-defective composite modules. Inspection of modules 18 on module placement head 40 may be an automated process.

The head carriage 36 of the pick and place machine 30 can then be operated to traverse the rails 32 on the gantry crane 31 of the pick and place machine 30 and facilitate the precise positioning of the stacked composite modules 18 , rolled, compressed, and inspected at the desired location on module placement surface 57 (FIG. 3) of curing tool 54. Additional stacked, laminated, and compressed modules 18 can be similarly formed and positioned at desired locations on module placement surface 57 to form the skin 60 of an aircraft fuselage. Module 19 edges of adjacent modules 18 may be ramping and overlapping by a bevel joint (not shown) or ramp splice (not shown) until the skin 60 of an aircraft fuselage has been fully stacked.

Referring now to flow chart 300 of FIG. 4, an illustrative embodiment of a modular composite manufacturing method is summarized. The method can be used to manufacture an aircraft fuselage skin having a desired thickness, for example and without limitation. At block 302, a composite module can be stacked. 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 stacked using an FTLM (tape stacking machine flat), for example and without limitation. Additionally or alternatively, the module can be used to fabricate a structure such as an aircraft wing skin, nose skin or tail skin, for example and without limitation, in which case the module can be stacked using a CTLM (Contour Tape Stacking Machine). At block 304, the module can be transferred to a SADL (Semi-Automated Folder Locator) machine. At block 306, the module can be formed to contour to an aircraft fuselage skin or other structure. In block 308, the steps carried out in blocks 302, 304 and 306 can be repeated to form a desired thickness of the skin of the fuselage of an aircraft or other structure. At block 310, the modules can be inspected. At block 312, the stacked or laminated modules can be stacked on top of a curing tool. The curing tool can be an OML (external mold line) or IML (internal mold line) curing tool, for example and without limitation. In block 314, the steps carried out in blocks 302, 304, 306, 308, 310 and 312 can be repeated as many times as necessary to fully stack the skin of an aircraft fuselage or other structure.

Referring now to Figures 5 and 6, embodiments of the disclosure may be used in the context of an aircraft production and service method 78 as shown in Figure 5 and an aircraft 94 as shown in Figure 6 During pre-production, exemplary method 78 may include specification and design 80 of aircraft 94 and procurement of material 82. During production, manufacturing 84 of components and sub-assemblies and an integration 86 of the aircraft system 94 take place. The aircraft 94 can then undergo certification and delivery 88 in order to be put into service 90. While in service In service by a customer, aircraft 94 can be scheduled for routine maintenance and service 92 (which may also include modification, reconfiguration, refurbishment, and so on).

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

As shown in FIG. 6, the aircraft 94 produced by the exemplary method 78 may include an aircraft frame 98 with a plurality of systems 96 and an interior 100. Examples of high-level systems 96 include one or more than one propulsion system 102, an electrical system 104, a hydraulic system106, and an environmental system 108. Any number of other systems can be included. While an aerospace example is shown, the principles of the invention can be applied to other industries, such as the automotive industry.

The apparatus made herein may be used during any one or more of the phases of the production and service method 78. For example, the components or sub-assemblies corresponding to production may be manufactured similarly to the components or sub-assemblies produced while the aircraft 94 is in operation. service. In addition, one or more embodiments of the apparatus may be used during production steps 84 and 86, for example, substantially speeding up the assembly of an aircraft 94 or reducing its cost. In a similar manner, one or more embodiments of the apparatus may be used while the aircraft 94 is in service, for example and without limitation, for maintenance and service 92.

As mentioned above, the modular composite manufacturing method can use equipment of the appropriate size to allow the composites to be stacked in parallel processes rather than in series as

ES 2 628 443 T3 part of the manufacture of composite aircraft fuselages or other parts. In this regard, reference is made to Figures 7 and 8. According to a number of embodiments, a method 700 of manufacturing a composite structure may include positioning 702 a plurality of modules 18 on tool 54 such that each of modules 18 is adjacent to at least one other of modules 18, such as shown in Figure 9 and such that adjacent modules 18 can be linked together, such as at a joint 704, to form a composite structure.

In some of the embodiments, more than one module 18 may be placed on the 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 structure, modules 18 can be placed on tool 54 in a substantially sequential fashion, with a post-placement procedure 706 performed on one of the modules 18 that has already been placed on the tool 18 (as indicated by note n-2 in Figure 8) while a subsequent module 18 is placed on tool 564 (as indicated by note n-1 in Figure 8). In various embodiments, the placement step 702 and the post-placement process step 706 may be performed while 708 preparing a subsequent module 18 for placement (as indicated by the note n in FIG. 8).

Regarding the execution of a post-placement procedure 706, this may include the inspection 710 of one of the modules 18 that has already been placed on the tool 18 (as indicated by note n-2 in figure 8 ). Performing a post-placement procedure 706 may further 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 executing, for example, the inspection steps 710 and the post-placement process step 706 on different modules. 18 placed substantially at the same time

In various embodiments, positioning 702 a module 18 on a tool 54 may include loading 712 of a module 18 on a transfer platform 14, indexing 714 of the loaded module in a suitable position, and / or then stacking 716 of the module on tool 54. After all modules 18 have been positioned on tool 54 and any subsequent setting process 706 has been carried out, then the framework can be cured.

Although the embodiments of this disclosure have been described with respect to certain exemplary embodiments, it is to be understood that the specific embodiments are for illustrative and non-limiting purposes, as other variations will be found by those skilled in the art.

The aspects and characteristics of the present disclosure are set out in the following numbered clauses that contain the subject matter of the claims of the relative European patent application as it was initially filed:

1. A modular composite manufacturing method, comprising:

providing a plurality of composite modules;

inspect said composite modules;

provide a curing tool; and positioning said composite modules on said curing tool.

two. The method of clause 1 which further comprises providing a selected contour for each of said compound modules.

3. The method of clause 2 further comprising cutting said composite modules prior to said facilitation of a selected contour for each of said composite modules.

Four. The method of clause 1 wherein each of said plurality of composite modules comprises at least one layer.

5. The method of clause 1 which further comprises the coupling of adjacent modules of said plurality of modules composed to each other.

6. The method of clause 1 further comprising placing successive modules of said plurality of composite modules one on top of the other to form a plurality of layers.

ES 2 628 443 T3

7. The method of clause 1 wherein said composite modules comprise ramp module edges arranged in overlapping relationship with respect to each other on said curing tool.

8. The method of clause 1 wherein each of said composite modules includes any combination of a prepreg unidirectional carbon fiber, a carbon fiber fabric, glass fiber, a glass fiber fabric, poly (p-phenylene terephthalamide) or a poly (p-phenylene terephthalamide) fabric.

9. The method of clause 1 wherein said inspection of said composite modules comprises providing a scan inspection system and scanning said composite modules by operating said scan inspection system.

10. The method of clause 1 wherein said curing tool comprises a separate tool base and tool sides 10 extending from said tool base and a generally concave module placement surface provided on said tool sides and said base of tool and wherein said placing said composite modules on said curing tool comprises placing said composite modules on said placement surface of module.