Source: https://patents.google.com/patent/EP2390514A1/en
Timestamp: 2018-08-20 17:28:14
Document Index: 577683742

Matched Legal Cases: ['arts 308', 'arts 308', 'arts 308', 'arts 308', 'arts 802', 'arts 802']

EP2390514A1 - Method for fastening aircraft components using a composite two-piece fastening system, and so obtained aircraft structure - Google Patents
Method for fastening aircraft components using a composite two-piece fastening system, and so obtained aircraft structure Download PDF
EP2390514A1
EP2390514A1 EP20110178811 EP11178811A EP2390514A1 EP 2390514 A1 EP2390514 A1 EP 2390514A1 EP 20110178811 EP20110178811 EP 20110178811 EP 11178811 A EP11178811 A EP 11178811A EP 2390514 A1 EP2390514 A1 EP 2390514A1
EP20110178811
EP2390514B1 (en )
Steven G Keener
A method for assembling a composite structure for an aircraft. A composite male fastening component (810) is placed through a pair of aircraft parts (802,804). The composite male fastener component has a mechanical locking feature (702). A composite female fastener component (812) is positioned adjacent to and surrounding the mechanical locking feature of the male fastener component. A portion of the composite female fastener component is caused to flow around and form into the mechanical locking feature of the composite male fastener component. The portion of the composite female fastener component that flowed around and formed into the mechanical locking feature of the composite male fastener component is re-solidified and re-consolidated such that the composite female fastener component is securely attached to the composite male fastener component thereby joining the mating aircraft parts.
Composite materials are tough, light-weight materials created by combining two or more dissimilar components to create a component with stronger properties than the original materials. Composite materials are also typically non-metallic materials. In these examples, a composite is a multiphase material, in which the phase distribution and geometry may have been controlled to optimize one or more properties.
The different advantageous embodiments of the present disclosure provide a method and apparatus for attaching parts.
In one embodiment a method is provided for assembling a composite structure for an aircraft. A composite male fastener component is positioned through a pair of aircraft parts, the composite male fastener component having a shaft and a mechanical locking feature. A composite female fastener component is positioned around the shaft and adjacent to the threaded portion of the shaft while the composite male fastener component is positioned through the pair of aircraft parts, wherein the composite female fastener component comprises carbon fiber and resin and has a cylindrical shape. Heat and force is applied to the composite female fastener component to cause a portion of the composite female fastener component to melt and flow into the mechanical locking feature. The heat and force is removed such that the portion of the composite female fastener component that melted collar re-solidifies, wherein the composite female fastener component is attached to the composite male fastener component and the pair of mating aircraft parts are mechanically joined to each other.
Figure 1 is a description of an aircraft manufacturing and service method in which an advantageous embodiment may be implemented;
Figure 2 is a description of an aircraft in which an advantageous embodiment may be implemented;
Figure 3 is a diagram illustrating a system for fastening parts in accordance with an advantageous embodiment;
Figure 4 is a diagram illustrating a fastening system in accordance with an advantageous embodiment;
Figure 5 is a cross-sectional view of a composite female fastener component in accordance with an advantageous embodiment;
Figure 6 is a cross-sectional view of a composite female fastener component in accordance with an advantageous embodiment;
Figure 7 is a cross-sectional view of a composite male fastener component in accordance with an advantageous embodiment;
Figure 8 is a cross-sectional view of a structure assembled using a fastening system in accordance with an advantageous embodiment;
Figure 9 is a flowchart of a process for attaching parts or assembling a structure in accordance with an advantageous embodiment; and
Figure 10 is a flowchart of a process for pre-treating a fastening system in accordance with an advantageous embodiment.
Referring more particularly to the drawings, embodiments of the disclosure may be described in the context of aircraft manufacturing and service method 100 as shown in Figure 1 and aircraft 200 as shown in Figure 2 . During pre-production, aircraft manufacturing and service method 100 in Figure 1 may include specification and design 102 of aircraft 200 in Figure 2 and material procurement 104. During production, component and sub-assembly manufacturing 106 and system integration 108 of aircraft 200 in Figure 2 takes place. Thereafter, aircraft 200 in Figure 2 may go through certification and delivery 110 in order to be placed in service 112. While in service by a customer, aircraft 200 in Figure 2 is scheduled for routine maintenance and service 114, which may include modification, reconfiguration, refurbishment, and other maintenance or service.
Each of the processes of aircraft manufacturing and service method 100 may be performed or carried out by a system integrator, a third party, and/or an operator as indicated by the "X" in the grid to the right of the flow diagram of Figure 1 . In these examples, the operator may be a customer. For the purposes of this description, a system integrator may include, without limitation, any number of aircraft manufacturers and major-system subcontractors; a third party may include, without limitation, any number of venders, subcontractors, and suppliers; and an operator may be an airline, leasing company, military entity, service organization, and so on.
As shown in Figure 2 , aircraft 200, produced by aircraft manufacturing and service method 100 in Figure 1 , may include airframe 202 with plurality of systems 204 and interior 206. Examples of systems 204 include one or more of propulsion system 208, electrical system 210, hydraulic system 212, environmental system 214, and airframe system 216.
Apparatus and methods embodied herein may be employed during any one or more of the stages of production and aircraft manufacturing and service method 100 in Figure 1 . For example, components or subassemblies provided in sub-assembly manufacturing 106 may be fabricated or manufactured in a manner similar to components or sub-assemblies produced while aircraft 200 is in service. Also, one or more apparatus embodiments, method embodiments, or a combination thereof may be utilized during production stages for component and sub-assembly manufacturing 106 and system integration 108 in Figure 1 , for example, by reducing the weight of aircraft 200. For example, the composite fastener system in the advantageous embodiments may be used in assembling structural and other composite components in aircraft 200. The use of these composite fastening systems reduces the weight of while improving the overall operating performance of aircraft 200. These additional weight savings may aid in increasing fuel savings and increasing payload capacity for aircraft 200.
Turning next to Figure 3 , a diagram illustrating a system for fastening parts is depicted in accordance with an advantageous embodiment. In this example, fastener assembly tool 300 retrieves a composite male fastener component from composite male fastener components 302 and a composite female fastener component from composite female fastener components 304 to form fastening system 306. Fastening system 306 is applied to mating parts 308 as part of the process of mechanically attaching or joining parts 308 to each other. In these examples, parts 308 may be two or more parts, depending on the particular implementation. Further, parts 308 may be composite parts or components used in a product, such as an aircraft.
Turning now to Figure 4 , a diagram illustrating a fastening system is depicted in accordance with an advantageous embodiment. In this example, fastening system 400 is an example of a fastening system, such as fastening system 306 in Figure 3 . Fastening system 400 includes composite male fastener component 402 and composite female fastener component 404, which has a substantially cylindrical shape. Composite male fastener component 402 slides through channels in components 406 and 408, and then into a channel within female fastener component 404. In this example, fastening system 400 is used to join components 406 and 408 to each other.
With reference next to Figure 5 , a cross-sectional view of a composite female fastener component is depicted in accordance with an advantageous embodiment. In this example, composite female fastener component 500 has a cylindrical shape that is viewed in a cross-section. Composite female fastener component 500 includes inner surface 502, which forms channel 504 through which a composite male fastener component may be received. In this example, inner surface 502 is a substantially smooth surface.
Turning next to Figure 6 , a cross-sectional view of a composite female fastener component is depicted in accordance with an advantageous embodiment. In this example, composite female fastener component 600 has inner surface 602, which forms channel 604 through which a composite male fastener component may be received. In this example, inner surface 602 has a threaded section.
Turning now to Figure 7 , a cross-sectional view of a composite male fastener component is depicted in accordance with an advantageous embodiment. In this example, composite male fastener component 700 has an elongate member with threaded section 702, which may be placed into a channel in a composite female fastener component. In this example, section 704 of composite male fastener component 700 may extend through two or more composite parts that are to be joined to each other using the fastening system in the different illustrative examples.
With reference now to Figure 8 , a cross-sectional view of a structure assembled using a fastening system is depicted in accordance with an advantageous embodiment. In this example, structure 800 is a composite structure for an aircraft that is assembled using parts 802 and 804. These parts are attached to each other using fastening system 808.
In these examples, fastening system 808 is comprised of composite male fastener component 810 and composite female fastener component 812. In particular, composite male fastener component 810 is preferably a threaded pin, and composite female fastening component 812 is a swag or 'form' collar. As can be seen, composite male fastener component 810 contains threaded section 814. In the depicted examples, the composite male fastener component may be any elongate number that has a threaded portion. Of course, the threaded portion may only encompass part or the entire elongate number. Composite male fastener component 810 has been placed through hole 806, which is present or drilled through mating parts 802 and 804.
Examples of thermoplastic polymeric resin materials that may be used to include, for examples, without limitation, liquid-crystal polymers (LCP); fluoroplastics, including polytetrafluoroethylene (PTFE), fluorinated ethylene propylene (FEP), perfluoroalkoxy resin (PFA), and polychlorotrifluoroethylene (PCTFE), and polytetrafluoroethylene-perfluoromethylvinylether (MFATM); ketone-based resins, including polyetheretherketone (PEEKTM); polyamides (for example nylon-6/6, 30 percent glass fiber); polyethersulfones (PES); polyamideimides (PAIS), polyethylenes (PE); polyester thermoplastics, including polybutylene terephthalate (PBT), polyethylene terephthalate (PET), and poly(phenylene terephthalates); polysulfones (PSU); poly(phenylene sulfides) (PPS).
A thermoset polymeric resin is a resin that does not readily re-melt or reflow after the initial cure. Still, this type of composite resin material also may be used, depending on the particular implementation. Examples of thermoset polymeric resin materials include, for example, without limitation, allyl polymers, alkyd polyesters, bismaleimides (BMI), epoxies, phenolic resins, polyesters, polyurethanes (PUR), and polyurea-formaldehydes.
Of course, any type of material may be used depending on the various properties and the desired uses. An example of one type of thermoplastic polymeric resin, used in the illustrative examples, is aromatic polyetheretherketone. This type of material is also referred to as a PEEK™ polymer. PEEKTM is a trademark of Victrex, Plc. This type of composite material is desirable for use in the different composite fastening systems illustrated in these examples because this type of material provides extremely good chemical resistance, abrasion resistance, high temperature resistance, hydraulics resistance, flame resistance with low smoke and toxic gasses, along with excellent electrical properties and excellent resistance to gamma rays.
An additional feature that may be implemented in the composite fastening system, in these examples, is to 'pre-treat' or `pre-coat' the composite fastener components. The composite male fastener component and the composite female fastener component may be 'pre-treated' or 'pre-coated' with an organic coating that improves the compatibility of these components with other dissimilar composite and metallic materials. With this improved compatibility, the assembled structures with the pre-coated composite components typically exhibit reduced issues relating to composite delamination, water intrusion, electrical continuity of the component, arcing between components, galvanic corrosion, fuel tightness, and surface lubricity allowing relative movement due to differential expansion of the components in the structure.
With reference now to Figure 9 , a flowchart of a process for attaching or joining parts or assembling a structure is depicted in accordance with an advantageous embodiment. In this example, the process illustrated in Figure 9 may be implemented using fastener assembly tool 300 in Figure 3 .
Thereafter, heat and force are applied to the composite female fastener component (operation 906). In the examples, the fastener assembly tool may be heated to approximately 600°F to 700°F at a minimum, with around 100 pounds to 300 pounds of force being applied for at least around 15 to 30 seconds. These parameters are present for purposes of illustrating one embodiment. The temperature, force, and time used may vary depending on the fastening system being used in other embodiments. Depending on the particular implementation, only heat may be applied to the composite female fastener component. Operation 906 is intended to cause a portion of the composite female fastener component to flow around or form into the threaded portions of the composite male fastener component. Although heat and force are illustrated as the techniques used to cause the portion of the composite female fastener component to flow around or form into the threaded portion(s), other mechanisms may be used. For example, an electron beam or laser process may be used to cause the portion of the composite female fastener component to flow or form.
Turning now to Figure 10 , a flowchart of a process for pre-treating components of a fastening system is depicted in accordance with an advantageous embodiment. The process illustrated in Figure 10 may be applied to components of a composite fastening system prior to the use of the fastening system to mechanically attach or join parts to each other.
A typical sprayable coating solution preferably has about 30 percent by weight ethanol, about 7 percent by weight toluene, about 45 percent by weight methyl ethyl ketone (MEK) as the solvent; and about 2 percent by weight strontium chromate, about 2 percent by weight aluminum powder, with the balance being mixed with phenolic resin and at least one plasticizer. A small amount of polytetrafluoroethylene may optionally be added. Such a product is available commercially under the name, Hi-Kote 1TM from Lisi Aerospace-Hi-Shear Corp., Torrance, California. This product has a standard elevated temperature curing treatment of one hour at +425 degrees, +/- 25 degrees, as recommended by the manufacturer. The Hi-Kote 1TM coating and other similar coatings are described in commonly assigned U.S. Patent Numbers 5,614,037 , 5,858,133 , 5,922,472 , and 5,944,918 , the contents of which are incorporated herein by reference to the extent they do not conflict with the explicit text of this specification.
The coating material is preferably provided in solution so as to be evenly applied. The coating material preferably is a formulation that is primarily of an organic nature but may also contain additives to improve the properties of the final coating. By way of example, a TeflonTM compound may be added to improve coating lubricity, which allows for reduced insertion force requirements, better hole fill, and improved sheet take-up. By way of another example, aluminum powder pigment may be added to improve overall coating material integrity. The coating is desirably dissolved initially in a carrier liquid to facilitate various methods of deposition on a substrate.
Optionally, before curing in operation 1004, the majority of solvent in the coating solution may be removed from the as-applied coating by drying or "flash cure" 1003, either at room temperature or slightly elevated temperature, so that the coated article or component is dry to the touch. Flash cure or drying may be achieved at about 200 degrees for about one to two minutes and accomplishes evaporation of the majority of solvent, allowing the coated article or component to be handled without altering or damaging the coating layer.
To fully cure the coating, in these examples, the coating is heated to and maintained above the curing temperature of the coating material as described in operation 1004. With this recommended "full-up" or full-curing process, cure time will vary with the coating material used and the associated cure temperature selected. Typical cure temperatures range from about +250 degrees to about +450°F, and typical cure times range from about one hour to about four hours, not respective, and more typically from about +400 degrees to about +450 degrees for about 1 to about 1.5 hours, not respective. It is understood that the term "pre-coated" or "pre-coating" refers to the coating process of the fastener component prior to installation and assembly in its final use.
In the case of the Hi-Kote 1™ coating, a flash cure temperature of +200 degrees for one to two minutes or full cure temperature of +425 degrees for approximately one hour are exemplary. The majority of the solvent portion of the coating is removed by a "flash-cure" drying process at a slightly elevated temperature of +200 degrees for one to two minutes. Flash curing volatizes the volatile portion of the coating solvent and allows handling of the coated fastener prior to full curing.
Aspects and features of the present disclosure are set out in the following numbered clauses which contain the subject-matter of the claims of the parent European patent application as originally filed:
1. A method for assembling a composite structure for an aircraft, the method comprising:
positioning a composite male fastener component through a pair of aircraft parts, the composite male fastener component having a shaft and a mechanical locking feature;
positioning a composite female fastener component around the shaft and adjacent to the threaded portion of the shaft while the composite male fastener component is positioned through the pair of aircraft parts, wherein the composite female fastener component comprises carbon fiber and resin and has a cylindrical shape;
applying heat and force to the composite female fastener component to cause a portion of the composite female fastener component to melt and flow into the mechanical locking feature; and
removing the heat and force such that the portion of the composite female fastener component that melted collar re-solidifies, wherein the composite female fastener component is attached to the composite male fastener component and the pair of mating aircraft parts are mechanically joined to each other.
2. The method of clause 1, wherein the composite female fastener component has an externally cylindrical shape and a threaded section on an inside portion of the composite female fastener component designed to mechanically interact and engage with the threaded portion of the composite male fastener component.
3. The method of clause 1, wherein the resin is selected from one of a thermoplastic polymeric resin or a thermoset polymeric resin.
4. The method of clause 1 further comprising:
applying a coating of a material to the composite male and female fastener components; and
thermally-treating the composite male and female fastener components to cure the coating prior to positioning the composite male fastener component through the pair of aircraft parts.
5. The method of clause 1, wherein the first part and the second part are composite parts.
6. The method of clause 5, wherein the composite parts are aircraft parts.
7. The method of clause 1, wherein the mechanical locking feature is one of a threaded or concentric grooves.
8. A method for attaching parts, the method comprising:
placing a composite male fastener component through a first part and a second part, the composite male fastener component having a mechanical locking feature;
positioning a composite female fastener component adjacent to the threaded portion;
causing a portion of the composite female fastener component to flow around the mechanical locking feature of the composite male fastener component; and
solidifying the portion of the composite female fastener component that flowed around the threaded portion of the composite male fastener component such that the composite female fastener component is attached to the composite male fastener component.
9. The method of clause 8, wherein the causing step comprises:
applying heat to the composite female fastener component to cause the portion of the composite female fastener component to flow around the threaded portion of the composite male fastener component.
10. The method of clause 8, wherein the solidifying step comprises:
cooling the composite female fastener component to cause the composite female fastener component to solidify such that the composite female fastener component is attached to the composite male fastener component.
11. The method of clause 10, wherein the cooling step comprises:
ceasing application of heat to the composite female fastener component.
12. The method of clause 9, wherein the causing step further comprises:
applying pressure to the composite female fastener component while applying the heat to the composite female fastener composite component.
13. The method of clause 8, wherein the composite female fastening component has a threaded section on an inside portion of the composite female fastener component designed to mechanically interact with the threaded portion of the composite male fastener component.
14. The method of clause 8, wherein the composite female fastener component includes a resin that is selected from one of a thermoplastic polymeric resin or a thermoset polymeric resin.
15. The method of clause 8, wherein the first part and the second part are composite parts.
16. The method of clause 15, wherein the composite components are aircraft parts.
17. The method of clause 15, wherein the composite parts are vehicle parts.
18. The method of clause 8, wherein the mechanical locking feature is one threads or concentric grooves.
a composite female fastener component attached to the composite male fastener component through processing composite female fastener component in a manner that caused a portion of the composite female fastener component to flow around the mechanical locking feature of the composite male fastener component and re-solidify around the threaded section to join the composite female fastener component to the composite male fastener component.
20. The composite aircraft structure of clause 19, wherein the first composite aircraft part and the second composite aircraft part are part of a wing for an aircraft.
21. The composite aircraft structure of clause 19, wherein the first composite aircraft part and the second composite aircraft part are structural components in an aircraft.
A method for assembling a composite structure (800) for an aircraft, the method comprising:
positioning a composite male fastener component (810) through a pair of aircraft parts (802, 804), the composite male fastener component (810) having a shaft and a mechanical locking feature (702);
positioning a composite female fastener component (812) around the shaft and adjacent to the mechanical locking feature (702) while the composite male fastener component (810) is positioned through the pair of aircraft parts (802, 804), wherein the composite female fastener component (812) comprises carbon fiber and resin and has a cylindrical shape;
applying heat and force to the composite female fastener component (812) to cause a portion of the composite female fastener component (812) to melt and flow into the mechanical locking feature (702); and
removing the heat and force such that the portion of the composite female fastener component (812) that melted collar re-solidifies, wherein the composite female fastener component (812) is attached to the composite male fastener component (810) and the pair of mating aircraft parts (802, 804) are mechanically joined to each other.
The method of claim 1, wherein the composite female fastener component (812) has an externally cylindrical shape and a threaded section on an inside portion of the composite female fastener component (812) designed to mechanically interact and engage with a threaded portion (814) of the composite male fastener component (810).
The method of claim 1, wherein the resin is selected from one of a thermoplastic polymeric resin or a thermoset polymeric resin.
applying a coating of a material to the composite male and female fastener components (810, 812); and
thermally-treating the composite male and female fastener components (810, 812) to cure the coating prior to positioning the composite male fastener component (810) through the pair of aircraft parts (802, 804).
The method of claim 1, wherein the first part (802) and the second part (804) are composite parts.
The method of claim 5, wherein the composite parts are aircraft parts.
The method of claim 1, wherein the mechanical locking feature (702) is one of threads or concentric grooves.
A composite aircraft structure (800) comprising:
a first composite aircraft part (802);
a second composite aircraft part (804);
a composite male fastener component (810) placed through the first composite aircraft part (802) and the second composite aircraft part (804), wherein the composite male fastener (810) has a mechanical locking feature (702); and
a composite female fastener component (812) attached to the composite male fastener component (810) through processing composite female fastener component (812) in a manner that caused a portion of the composite female fastener component (812) to flow around the mechanical locking feature (702) of the composite male fastener component (810) and re-solidify around the mechanical locking feature (702) to join the composite female fastener component (812) to the composite male fastener component (810).
The composite aircraft structure of claim 8, wherein the first composite aircraft part (802) and the second composite aircraft part (804) are part of a wing for an aircraft.
The composite aircraft structure of claim 8, wherein the first composite aircraft part (802) and the second composite aircraft part (804) are structural components in an aircraft.
EP20110178811 2007-08-14 2008-08-12 Method for fastening aircraft components using a composite two-piece fastening system, and so obtained aircraft structure Active EP2390514B1 (en)
EP20080162257 EP2025954B1 (en) 2007-08-14 2008-08-12 Method for fastening components using a composite two-piece fastening system, and the related assembled part
EP08162257.3 Division 2008-08-12
EP20080162257 Division EP2025954B1 (en) 2007-08-14 2008-08-12 Method for fastening components using a composite two-piece fastening system, and the related assembled part
EP2390514A1 true true EP2390514A1 (en) 2011-11-30
EP2390514B1 EP2390514B1 (en) 2015-12-16
EP20080162257 Active EP2025954B1 (en) 2007-08-14 2008-08-12 Method for fastening components using a composite two-piece fastening system, and the related assembled part
EP20110178811 Active EP2390514B1 (en) 2007-08-14 2008-08-12 Method for fastening aircraft components using a composite two-piece fastening system, and so obtained aircraft structure
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