COMPOSITE STRUCTURE, E, METHOD OF PRODUCTION OF A COMPOSITE STRUCTURE

专利摘要:
composite structure, and method of producing a composite structure. a composite structure (200) may include a laminate (204) and a stabilizing element (300). the laminate (204) may have a plurality of composite veneers (214). the composite structure (200) may include a geometric discontinuity (256) that may be associated with the laminate (204). the stabilizing element (300) may be included with the composite veneers (214) and may be positioned close to the geometric discontinuity (256).
公开号:BR112015004962B1
申请号:R112015004962-1
申请日:2013-08-22
公开日:2021-08-31
发明作者:Marc R. Matsen；Mark A. Negley
申请人:The Boeing Company；
IPC主号:

专利说明:

FIELD
[001] The present exposition refers generally to composite materials and methods and, more particularly, to hybrid composite laminates having stabilizing elements. BACKGROUND
[002] Composite materials are used in a wide variety of structures. In aircraft construction, composite materials can be used to form the fuselage, wings, tail section, and other components. For example, an aircraft fuselage can be constructed of composite cladding panels, to which composite structural elements, such as "hat" type spars, can be affixed. "Hat" type stringers can increase the strength and rigidity of cladding panels.
[003] During the manufacture of a composite structure, layers of composite veneers can be laid on a tool or a mold. The tool or mold can be provided in the desired shape of the final composite structure. Composite veneers can comprise a plurality of high modulus or high strength fibers, such as carbon, glass, or other fibers. The fibers can be prepreg with a polymeric matrix material, such as epoxy or thermoplastic resin to form composite prepreg veneers. The fibers in a composite veneer can commonly be aligned or oriented in a single direction (eg, unidirectional) or the fibers in a composite veneer can be woven together in two or more directions in a fabric arrangement. Composite structures can be designed to transmit primary charges along the length of the fibers. In this regard, composite structure formed from unidirectional fibers can have a relatively high tensile strength along one direction by the length of the fibers.
[004] After the prepreg composite veneers are seated on the tool or mold, a curing cycle can be performed on the settling. The curing cycle can comprise the application of heat and compaction pressure to the settlement. The application of heat can reduce resin viscosity, allowing a resin to flow and be intermingled with a resin in adjacent composite veneers. The application of compaction pressure may include installing a vacuum chamber over the seat and/or positioning the seat within an autoclave. Compaction pressure can compact the composite veneers against the tool or mold to minimize or reduce porosity and voids in the final composite structure. In addition, compaction pressure can force seating against the tool or mold to establish the final shape and surface finish of the composite structure.
[005] Although the vacuum chamber can apply substantially uniform pressure to most of the laying of prepreg composite veneers, the reduction in resin viscosity during application of compaction pressure can result in a resin flow in the towards direction. regions of low compaction pressure under the vacuum chamber. Regions of low compaction pressure can occur in locations where there is a geometric discontinuity associated with settlement. Geometric discontinuity can result in out-of-plane fiber movement during cure. For example, a geometric discontinuity can occur at an edge of a structural element (eg a spar, a reinforcement, etc.) that can be assembled or joined (co-cured, bonded, co-consolidated) to a cladding panel formed as a laminate of uncured prepreg composite veneers. Geometric discontinuity at the edge of a gusset can result in bridging the vacuum chamber from the edge of the gusset to a seating surface.
[006] The area under the bridging may comprise a region of low compaction pressure. Resin can flow towards the low compaction pressure region and may cause the fibers in the composite veneers to also shift towards the low compaction pressure region. Fiber movement can cause the fibers to be twisted, resulting in fiber out-of-plane distortion. In resin curing and solidification, distortion of out-of-plane fibers can become permanently solidified in the composite structure. Distortion of out-of-plane fibers can affect the load-bearing capacity of fibers which are typically designed to provide maximum strength when the fibers are oriented in a common direction within the layer or veneer. In this regard, distortion of out-of-plane fibers may have a less than desired effect on the characteristics of the final composite structure.
[007] The document US5685940 discloses the reduction in core crushing and layer wrinkling in the composite honeycomb sandwich structure, preventing slippage of the tie layers in relation to the mandrel and each other during autoclave curing. The method involves applying a film adhesive to the tie layers at the edge of the piece outside the finishing line of the net. During autoclave heating and before high pressure is applied to the composite structure, the film adhesive cures to form a strong bond between the layers and the mandrel. When pressure is applied, the tie layers are locked together and to the mandrel to prevent slippage between any layers of the panel.
[008] Document CN101516613 provides a plastic similar to a fibre-reinforced board having a board and a protrusion rising up on the plate, wherein the plate and protrusion each contain a laminated structure composed of several sheets of fiber reinforcement, each comprising many fibers from 10 to 100 mm in length of fiber arranged in a fixed direction and at least two of the layers constituting the laminated structure are different in the direction of the arrangement of the reinforcing fibers and wherein at least part of the reinforcing fibers they extend continuously from the plate to the protrusion and in at least one of the layers making up the laminated structure of the protrusion has a shape similar to the shape of the protrusion. In addition, Document CN101516613 provides a process for the production of fiber reinforced plastic which comprises cutting a unidirectional pre-impregnated sheet composed of many reinforcing fibers arranged in a fixed direction and a matrix resin into sheets of base material. prepreg with the above fiber length, laminating the prepreg base material sheets with the fiber reinforcement arrangement directions of the base material sheets being different from each other to form a prepreg laminate and heating and pressing this laminate into a mold provided with a recess to form the production.
[009] The document CN1101751 discloses the elimination of resin flow to the honeycomb cells in sandwich structure using an unsupported film adhesive (108), a barrier layer (110) and an adhesive layer with reinforcement support (112) between the composite laminate (102) and the core (106). CN1101751 further reduces core crushing and layer wrinkling in the composite honeycomb sandwich structure by preventing slippage of the tied layers relative to the mandrel and each other during autoclaving. The method disclosed therein involves applying a film adhesive to the tie layers at the edge of the outside of the net finishing line. During autoclave heating and before high pressure is applied to the composite structure, the film adhesive cures to form a strong bond between the layers and the mandrel. When pressure is applied, the lashing plies are locked together and to the mandrel to prevent slippage between any layers of the panel.
[0010] The document US2004065409 discloses a method for applying pressure to the area of a composite part masked by a secondary feature comprising the steps of positioning a pressure-increasing strip in contact with the masked area, securing a pressure transfer wedge on each of the two sides of the secondary feature, so that pressure is transferred to the pressure build-up strip during co-cure/co-bond and the co-cure/co-bond of the composite part and the secondary feature . When the area between the composite part and the secondary feature is quite large, the method can comprise the additional steps of positioning a sine wave spring between the pressure build-up range and the secondary feature, injecting a structural paste adhesive into the voids around the spring sine wave, and curing the structural paste.
[0011] EP1925436 provides a method for manufacturing a fiber reinforced laminate (13,15). The method comprises the steps of: a) constructing a portion (15) of the laminate (13,15) to a determined thickness, b) placing a layer (13) of material on top of the partially completed laminate (15), said layer (13) of material having a stiffness greater than the stiffness of a layer of similar thickness to the uncured laminate, and c) constructing a new part of the laminate (15) with a determined thickness. In case the thickness of the laminate constructed according to steps a) -c) is not as great as the desired thickness of the finished laminate, steps b) and c) are repeated until the thickness of the laminate constructed is equal to the desired thickness of the finished laminate.
[0012] As can be seen, there is a need in the art for a system and method to minimize out-of-plane fiber distortion in composite structures. SUMMARY
[0013] The needs noted above, associated with the distortion of out-of-plane fibers in composite structures, are specifically addressed and mitigated by the present disclosure, which provides a composite structure that may include a laminate and a stabilizing element. The laminate can have a plurality of composite veneers. The composite structure can include a compression discontinuity that can be associated with the laminate. The stabilizing element can be included with the composite veneers and can be positioned close to the compression discontinuity.
[0014] In another embodiment, a composite structure is exposed which may include a laminate and a stabilizing element and wherein the laminate may have a plurality of composite veneers. The composite structure can include a geometric discontinuity that can be associated with the laminate. The stabilizing element can be included with the composite veneers and can be positioned close to the geometric discontinuity.
[0015] A veneer stabilizer is also exposed. The veneer stabilizer may include a stabilizing element for a laminate having a plurality of composite veneers. The laminate may have a compression discontinuity associated with it. The stabilizing element can be included with the composite veneers and can be positioned close to the compression discontinuity.
[0016] A method of producing a composite structure is also exposed. The method may comprise laying a laminate with a plurality of composite veneers. The laminate may have a compression discontinuity or a geometric discontinuity associated with it. The method may further include applying a stabilizing element with the composite veneers and positioning the stabilizing element close to a compression discontinuity or geometric discontinuity.
[0017] The features, functions and advantages that have been discussed can be obtained independently in the various embodiments of the present disclosure or can be combined in still other embodiments, other details of which can be seen with reference to the following description and drawings below. BRIEF DESCRIPTION OF THE DRAWINGS
[0018] These and other features of the present exhibition will become more apparent in reference to the drawings, in which the same numbers refer to the same parts throughout the entire description, and in which:
[0019] Figure 1 is a perspective illustration of an aircraft;
[0020] Figure 2 is a perspective illustration of a cylinder section of an aircraft fuselage taken along line 2 of Figure 1;
[0021] Figure 3 is a perspective illustration of a portion of a cylinder section taken along line 3 of Figure 2 and illustrating the cylinder section composed of a panel (e.g., cladding panel) having such structural elements as "hat"-type reinforcements mounted on it;
[0022] Figure 4 is a cross-sectional illustration of a portion of the cylinder section taken along line 4 of Figure 3 and illustrating the "hat" type stiffeners attached to the panel;
[0023] Figure 5 is an exploded cross-sectional illustration of a structural element radius (i.e., a "hat" type reinforcement), a panel, and a layer of adhesive taken along line 5 of Figure 4 and illustrating the attachment of the "hat" type reinforcement to the panel;
[0024] Figure 6 is a cross-sectional illustration of the structural element of Figure 5 being attached to the panel and illustrating the application of pressure by a vacuum chamber, causing a region of low compaction pressure on a structural element edge of the element structural (ie at the edge of the "hat" type reinforcement) and resulting in an out-of-plane distortion of the panel fibers;
[0025] Figure 7 is a cross-sectional illustration of the structural element and the panel of Figure 6 having a stabilization element installed in the panel, close to the edge of the structural element and resulting in minimization of out-of-plane fiber distortion in the panel;
[0026] Figure 8 is a cross-sectional illustration of a radius of the structural element taken along line 8 of Figure 7 and illustrating the application of pressure by a vacuum chamber causing a region of high compaction pressure in a structural element radius and resulting in radius thinning in the structural element radius due to resin flow away from the structural element radius;
[0027] Figure 9 is a cross-sectional illustration of the structural element radius of the structural element (for example, the "hat" type reinforcement) of Figure 8 having a stabilizing element installed close to the structural element radius and resulting in minimizing the radius thinning in the structural element radius;
[0028] Figure 10 is a cross-sectional illustration of a radius fill (i.e., a mass) of the structural element radius taken along line 10 of Figure 7 and illustrating out-of-plane fiber distortion at a nearby location to radius filling;
[0029] Figure 11 is a cross-sectional illustration of the radius fill of the structural element of Figure 10 having a stabilizing element installed close to the radius fill and resulting in minimization of out-of-plane fiber distortion;
[0030] Figure 12 is an end view illustration of the composite cylinder, taken along line 12 of Figure 2 and illustrating the plurality of compression mold plates for mounting on the cylinder section cladding panel;
[0031] Figure 13 is a cross-sectional illustration of a gusset mounted to a panel, taken along line 13 of Figure 12 and illustrating out-of-plane fiber distortion that occurs in an interstice between compression mold plate edges of the compression mold plates;
[0032] Figure 14 is a cross-sectional illustration of the "hat" type reinforcement and the panel of Figure 13 having a stabilization element installed in the panel close to the edges of the compression mold plate;
[0033] Figure 15 is a cross-sectional illustration of a pad on the panel taken along line 15 of Figure 3 and illustrating out-of-plane fiber distortion that occurs at a perimeter edge of the pad;
[0034] Figure 16 is a cross-sectional illustration of the pad in Figure 15 and resulting in the minimization of distortion of out-of-plane fibers in the panel;
[0035] Figure 17 is a cross-sectional illustration of the stabilizing element and composite veneers arranged so that the stabilizing element edges and veneer edges are disposed in opposite relationship to each other;
[0036] Figure 18 is a cross-sectional illustration of the stabilizing element and composite veneers arranged so that the stabilizing element edges and veneer edges are disposed in superimposed relationship with each other;
[0037] Figure 19 is an illustration of a flowchart having one or more operations that can be included in a method of manufacturing a composite structure;
[0038] Figure 20 is an illustration of a block diagram of a composite structure having at least one stabilizing element;
[0039] Figure 21 is an illustration of a flowchart of an aircraft production and service methodology; and
[0040] Figure 22 is a block diagram of an aircraft. DETAILED DESCRIPTION
[0041] Referring now to the drawings, in which what is shown is for purposes of illustrating preferred embodiments and various embodiments of the exhibit, in Figure 1 there is shown a perspective view of a passenger aircraft 100, formed of one or more composite structures 200. For example, aircraft 100 may include a fuselage 102 and a pair of wings 106 extending outwardly from the fuselage 102. The fuselage 102 may be composed of one or more cylinder sections 104 which may be formed, each as a composite structure 200. Each of the wings 106 may also be formed as a composite structure 200. The empennage 108 may include a horizontal stabilizer 110, an elevator 112, a vertical stabilizer 114, and a rudder blade 116, which they can additionally be formed as a composite structure 200. Although the present exhibit is described in the context of a passenger aircraft with fixed wings 100, as illustrated in Figure 1, the modalities exposed may be applied to aircraft of any configuration, without limitation. Also in this regard, the modalities exposed can be implemented in any vehicular or non-vehicular application, without limitation, and are not limited to implementation in an aircraft 100.
[0042] With reference to figure 2, there is shown a perspective view of a portion of a cylinder section 104 of a fuselage 102 (figure 1). Cylinder section 104 may be formed as a composite structure 200 and may include one or more panels 206 (e.g., skin panels). Each panel 206 may be formed as a laminate 204 composed of a plurality of composite veneers 214. The panel 206 may be supported by a plurality of structural elements 400. Each of the structural elements 400 may also be formed as a laminate 204 of a plurality of composite veneers 214. In Figure 2, the structural elements 400 supporting the panel(s) 206 (e.g., cladding panels) may comprise a plurality of longitudinally extending, circumferentially spaced spars 40 or gussets. "hat" 404, and a plurality of circumferentially spaced frames 202. In one embodiment, the "hat" type struts 404 can withstand axial forces (not shown), such as axial stress loads (not shown), bending loads (not shown), and other charges. The frames 202 can maintain the shape of the fuselage 102 and can support circumferential or circular loads (not shown) and other loads. The frames 202 and "hat" type gussets 404 can improve the deformation resistance (not shown) of the fuselage 102 under bending (not shown). Frames 202 and "hat" type gussets 404 may also collectively increase the torsional or flexural rigidity (not shown) of panel(s) 206, among other qualities provided by frames 202 and gusset type of "hat"404. Referring to Figure 3, there is shown a perspective view of a portion of cylinder section 104 (Figure 1) illustrating a panel 206 having a plurality of structural elements 400 (e.g., "hat" type gussets404) mounted on the panel 206 and forming a composite structure 200. In one embodiment, one or more of the structural elements 400 (e.g., "hat" type stiffeners404) may be formed as a laminate 204 of composite veneers 214 as mentioned above. One or more of structural elements 400 can be secured to panel 206 by gluing, co-gluing, or co-curing structural elements 400 to panel 206, as described in greater detail below. When used herein, a structural element 400 may include a "hat" type reinforcement 404, a frame 202 (Figure 2), a spar (not shown), or any other structural element 400 of any configuration or geometry, without limitation, which can be joined to a panel 206. Advantageously, the composite structure 200 can include one or more veneer stabilizers comprising stabilizing elements 300 configured to provide rigidity to the composite veneers 214 and to mitigate or prevent distortion of out-of-plane fibers 244 (figure 6) during compaction, consolidation or curing (not shown) of the composite structure 200 such as during compaction, consolidation, or curing of the composite structure 200.
[0043] Referring to Figure 4, there is shown a cross-sectional view of a cylinder section 104 (Figure 1) illustrating a plurality of structural elements 400 (e.g. "hat" type stiffeners404) mounted on panel 206 Each of the structural elements 400 may represent at least one compression discontinuity 258 associated with the panel 206. In one embodiment, a compression discontinuity 258 may occur at locations where non-uniform compaction pressure 329 is applied to the panel 206. For example , a compression discontinuity 258 may occur at locations where a structural element 400 is mounted on a panel 206. Figure 4 further illustrates stabilizing elements 300 that are advantageously included with the laminated panel 206 204 and positioned close to one or more geometric discontinuities 256 which can be represented by each of the structural elements 400.
[0044] In the embodiment shown in Figure 4, the stabilization element 300 can be interspersed with (e.g., sandwiched between) a pair of the composite veneers 214 of panel 206. However, the stabilization element 300 can be positioned on top (not shown) of an upper surface of laminate 210 and/or on top (not shown) of a lower surface of laminate 212 of laminate 204. ). In this regard, the stabilizing element 300 preferably has a relatively high stiffness or relatively high modulus of elasticity (e.g., high flexural strength) which can cause the composite veneers 214 to remain substantially in-plane (not shown) and prevent distortion. of fibers out of plane in the direction through thickness 241 of laminate 204 during compaction or consolidation of laminate 204, as described in greater detail below.
[0045] In Figure 4, stabilizing element 300 may extend along at least a portion of a length (not shown) of structural element 400. Each of stabilizing elements 300 may have opposing edges of stabilizing element 308 The stabilizing element 300 (Figure 4) may be dimensioned and configured such that one or more of the stabilizing element edges 308 (Figure 4) extend after the structural element edges 418. However, the stabilizing element 300 may be configured such that one or more of the stabilizing element edges 308 does not extend beyond the structural element edges 418.
[0046] With reference to figure 5, an exploded view of a structural element 400 is shown, positioned above a panel 206, without the stabilizing element 300. Although the structural element 400 in figure 5 is shown in a reinforcement of the configuration of the "hat" type 404, structural element 400 may be provided in any of a variety of different sizes, shapes, and configurations, without limitation. In the embodiment shown in Figure 5, structural element 400 (i.e., "hat"-type reinforcement404) may be composed of a plurality of composite veneers 214. However, structural element 400 may be formed of any material including any metallic and/or non-metallic material, without limitation, and is not limited to being formed from composite veneers 214.
[0047] In Figure 5, the structural element 400 (e.g., "hat" type reinforcement 404) may include a base portion 410 and may have a pair of upwardly extending continuous sections 412 that may be interconnected by a cap 414. The base portion 410 may include flanges 416 at opposite ends of the base portion 410. Each of the flanges 416 may terminate in an edge of structural element 418. In one embodiment, the "hat" type reinforcement "404 may be composed of a plurality of sub-laminates 430. For example, structural element 400 may include a base laminate 434, a primary laminate 432, and a shell laminate 436. The structural element 400 may include a backfill of radius 440 or mass at a joint 438 of sublaminates 430. Radius filler 440 may be composed of unidirectional composite material (not shown) or other alternative materials.
[0048] In Figure 5, in one embodiment, the structural element 400 may be formed as a laminate 204 of cured or precured composite veneers 218. However, the structural element 400 may be provided as a laminate 204 of uncured composite veneers 216. Also, panel 206 may be formed as a laminate 204 of uncured composite veneers 216. However, panel 206 may be provided as a laminate 204 of cured or precured composite veneers 218. In one embodiment, the composite veneers 214 may be composed of fiber reinforced polymeric material 224 including relatively high modulus fibers and high strength fibers 230, such as, without limitation, carbon fibers. However, fibers 230 may be formed from fiber material 232 comprising graphite, glass, carbon, boron, ceramics, aramids, polyolefins, polyethylenes, polymers, tungsten carbide, and/or any other fiber materials 232, without limitation. Fibers 230 of composite veneers 214 can be unidirectional or fibers 230 can be woven or knitted into a fabric arrangement (not shown).
[0049] In Figure 5, composite veneers 214 may be pre-impregnated (e.g., pre-impregnated) with polymeric resin 226. However, the present disclosure is not limited to pre-preg composite veneers 214, but may comprise structures composites 200 formed from dry or substantially dry fiber preforms (not shown), which can be set on a tool (not shown) and infused with liquid resin (not shown). In the present disclosure, a resin 226 may comprise thermosetting resin 226, such as epoxies and polyesters, or a resin 226 may comprise thermoplastic resin, such as polyamides, polyolefins, fluoropolymers, and/or other resin material 228. Fibers 230 may have a fiber stiffness (not shown) in a range of from approximately 1532.17 MPa (32 MSI) (thousands of pounds per square inch) to approximately 4788.03 MPa (100 MSI). However, fibers 230 can be provided with a fiber stiffness that is less than 1532.17 MPa (32 MSI) or greater than 4788.03 MPa (100 MSI).
[0050] Fibers 230 may be provided with an elongation capability of fiber 236 in a range of from approximately 0.1% to approximately 1% or greater than the original fiber length (not shown). However, fibers 230 can be provided in any fiber elongation capacity. Each of the composite veneers 214 can be provided in a veneer thickness 222 (figure 17) in a range from approximately 25.4 µm (1 mil) to approximately 508 µm (20 mils) and more preferably within a thickness of 222 veneer over a range from approximately 101.6 µm (4 mils) to approximately 203.2 µm (8 mils). However, composite veneers 214 may be provided in any thickness of veneer 222, without limitation. Laminate 204 for structural element 400 and/or panel 206 can be formed using conventional laying equipment (not shown), such as a tape laying machine (not shown) or laminate 204 for structural element 400 and/ or panel 206 can be formed by hand laying.
[0051] In Figure 5, in one embodiment, one or more of the structural elements 400 may be secured to the panel 206 by gluing, co-gluing, or co-curing the structural elements 400 to the panel 206, as described in greater detail below. The co-gluing may comprise gluing one or more structural elements 400 formed from cured or precured composite veneers 218 to a panel 206 formed from uncured composite veneers 216, while simultaneously curing the panel 206 during the co-curing process. Co-curing may comprise simultaneously curing one or more structural elements 400 formed from uncured composite veneers 216 and a panel 206 formed from uncured composite veneers 216. The process of co-curing structural element 400 and panel 206 may include a application of heat and pressure to consolidate the uncured composite veneers 216 of structural element 400 and panel 206 and can result in bonding structural element 400 to panel 206.
[0052] Referring to Figure 6, there is shown a structural element 400, such as a "hat" type reinforcement 404, attached to a panel 206 without a stabilizing element 300 so that Figure 6 can illustrate the effect of non-uniform compaction pressure 329 on panel 206. In Figure 6, structural member 400 can be co-attached to panel 206 using a vacuum chamber 326 to apply compaction pressure 324 to form a composite structure 200. Applying pressure Compaction steps 324 may include creating a vacuum (not shown) over a vacuum chamber 326 and/or positioning the composite structure in vacuum chamber 200 within an autoclave (not shown). As indicated above, compaction pressure 324 can be applied during application of heat (not shown) which can result in a reduction in viscosity (not shown) of resin 226. composite veneers 214 drain and be interspersed with resin 226 into adjacent composite veneers 214.
[0053] As shown in Figure 6, the vacuum chamber 326 can result in a compaction pressure 324 that can be applied to the composite veneers 214 of frame element 400 and panel 206. Frame element 400 can include a mandrel 444 that can be temporarily or permanently installed during the application of compaction pressure 324. For example, the mandrel 444 may be foamed (not shown) or the mandrel 444 may comprise an inflatable pocket (not shown) that may be temporarily installed to hold the shape of the structural element 400 during the application of compaction pressure 324. However, the mandrel 444 can be permanently installed in the structural element 400.
[0054] In Figure 6, structural element 400 may represent a compression discontinuity 258 associated with panel 206 on each of the edges of structural element 418. For example, structural element 400 may result in the application of non-uniform compaction pressure 329 to panel 206. In this regard, each of the structural member edges 418 may result in the formation of a low compaction pressure region 330 caused by bridging 328 of the vacuum chamber 326 from the structural member edge 418 to the surface. of laminate 210. The reduced viscosity of resin 226 during application of compaction pressure 324 can result in a resin 226 flowing along a resin flow direction 334 toward the low compaction pressure region 330. of resin 226 can cause fibers 230 to move along resin flow direction 334 which can result in localized stacking of fibers 230 in a d wave. and bow 242 in the low compaction pressure region 330. Bow wave 242 may represent distortion of out-of-plane fibers 244 in fibers 230 of one or more of the composite veneers 214. In the curing and solidification of resin 226, the distortion of out-of-plane fibers 244 can become permanently fixed in the composite structure 200. Distortion of out-of-plane fibers 244 can affect the load-bearing capacity of the composite veneers 214.
[0055] Referring to Figure 7, there is shown an embodiment of a composite structure 200 advantageously having a stabilizing element 300 provided with laminate 204 of panel 206. In the region of low compaction pressure 330, stabilizing element 300 slows down or prevents distortion of out-of-plane fibers 244 (Fig. 6) so that the composite veneers 214 are advantageously held in a fiber-in-plane direction 240. Stabilizing element 300 may be formed of a material having relatively high stiffness element stiffness. stabilization 302 at composite processing temperatures (eg the cure temperature or the set temperature). The relatively high stiffness of stabilizing element 302 of stabilizing element 300 can resist generating a bow wave 242 (Fig. 6) and reduce or mitigate distortion of out-of-plane fibers 244 (Fig. 6). In this regard, the stabilizing element 300 can act as a fiber distortion attenuating element that can improve the load-bearing capacity of the composite structure 200 over the load-bearing capacity of a composite structure having fiber distortion outside. of plan 244.
[0056] In Figure 7, the stabilizing element 300 may be provided in a size, shape, and configuration that extends at least partially through the low compaction pressure regions 330.
[0057] More particularly, stabilizing element 300 may have opposing edges of stabilizing element 308. Stabilizing element 300 may be configured such that at least one of stabilizing element edges 308 extends beyond an element edge 418. In addition, stabilizing element 300 may be configured such that at least a portion of stabilizing element 300 extends through low compaction pressure region 330. For example, stabilizing element 300 may be provided in a width such that at least one of the stabilizing element edges 308 extends beyond a structural element edge 418 by an amount at least approximately equal to a thickness of laminate 246 of laminate 204. Still further, although Figure 7 illustrates the stabilizing element 300 extending through an entirety of the structural element 400, the composite structure 200 may be provided in two parts. is separate stabilizing elements 300 (not shown), wherein each stabilizing element 300 may be positioned proximate to one of the structural member edges 418 and extending through one of the low compaction pressure regions 330.
[0058] In the embodiment shown in Figure 7, the stabilizing element 300 may be positioned close to an upper surface of laminate 210. For example, the stabilizing element 300 may be sandwiched within (e.g., sandwiched between) the composite veneers 214 of laminate 204 and may be positioned at a depth 322 (Figure 17) of no more than approximately ten of the composite veneers 214 below the upper surface of laminate 210. In another embodiment, stabilizing element 300 may preferably be positioned in a depth 322 of no more than approximately three of the composite veneers 214 below the top surface of laminate 210. Although Figure 7 illustrates a single of the stabilizing elements 300 installed in a stack of the composite veneers 214 of laminate 204, any number of stabilizing elements 300 can be installed in a stack of the composite veneers 214. In addition, though the stabilizing element 300 is shown as a thin, relatively flat homogeneous sheet having a relatively constant stabilizing element thickness 306 (Fig. 17), stabilizing element 300 may be provided in alternative configurations including a simply curved shape (not shown - e.g. cylindrical, conical) or complex contour shape (not shown - eg a double curved shape of an aircraft nose) to conform to the simply curved shape (not shown) or complex contour shape (not shown) of composite veneers 214 of stabilizing element 300, and may have a non-uniform thickness (not shown).
[0059] In Figure 7, the stabilizing element 300 may advantageously be formed of a stabilizing element material having a stabilizing element 302 stiffness (e.g. stabilizing element modulus of elasticity) at the processing or curing temperature of composite veneers 214 which is greater than the stiffness of composite laminate 234 at the processing or curing temperature of composite veneers 214. For composite veneers 214 formed from thermosetting material, stabilizing element 300 may be formed from a stabilizing element material having a 302 stabilizing element hardness at a cure temperature of approximately 25°F (Fahrenheit) to 35°F or greater. For composite veneers 214 formed from thermoplastic material, stabilizing element 300 may be formed from stabilizing element material having a stabilizing element 302 stiffness at a processing (e.g., setting) temperature of approximately 600°F to 720 °F or higher. Stabilizing element 300 may be formed of a stabilizing element material that has a stabilizing element 302 stiffness in a range of from approximately 718.20 MPa (15 MSI) to approximately 3830.42 MPa (80 MSI), as indicated above, although stabilizing element 300 may be formed of any stabilizing element material having a stabilizing element 302 stiffness that is greater or less than in the range of 718.20 - 3830.42 MPa (15 - 80 MSI) . In one embodiment, stabilizing element 300 may be formed from molybdenum having a stabilizing element 302 stiffness of approximately 2250.37 MPa (47 MSI) at a cure temperature of approximately 177°C (350°F), typically associated with carbon epoxy materials. Advantageously, the stabilizing element 300 is also preferably a relatively inert material that exhibits minimal galvanic corrosion in the presence of graphite, epoxy or other composite materials.
[0060] Still referring to Figure 7, the stabilizing element 300 may be formed of stabilizing element material having a coefficient of thermal expansion (CTE) 304 that is comparable to that of the flat laminate CTE 238 of the composite laminate 204. For example, as indicated above, stabilizing element 300 may be formed from molybdenum which may have a CTE of stabilizing element 304 in a range of approximately 2.5 x 10-6 to 3.5 x 10-6 inch/inch/ °F (degree Fahrenheit) at a composite cure temperature of 350 °F and which can compare favorably with the 238 laminate CTE which can be in the range of from approximately 0.5 x 10-6 to 6.0 x 10- 6 inch/inch/°F. However, depending on the stabilizing element material, the stabilizing element 300 may have a stabilizing element 304 CTE that is greater or less than the range of 2.5 x 10-6 to 3.5 x 10-6 inch/ inch/°F. In one embodiment, stabilization element 300 may have a stabilization element CTE 304 that is substantially equivalent to laminate CTE 238. For example, stabilization element 300 may have a stabilization element CTE 304 that is within pile. minus ten percent of the CTE of laminate 238 at the curing temperature (eg, processing, setting) to minimize distortion or minimal stress (not shown) that might otherwise occur in laminate 204 during the curing and/or curing process. consolidation.
[0061] The stabilizing element 300 may be formed from stabilizing element material comprising a metallic material, a non-metallic material, or any other relatively high modulus material at composite processing temperatures (e.g., cure temperature, temperature solidification, etc.).
[0062] For example, the metallic material may comprise molybdenum, iron, and/or titanium, or any alloy thereof or other materials (eg, Invar, steel). Stabilizing element 300 may also be formed from a non-metallic material, such as a cured composite material and/or a ceramic material. In this regard, the stabilizing element 300 can be formed from a material having relatively high stiffness, a relatively low coefficient of thermal expansion, minimal galvanic corrosion in the presence of composite materials, and which maintains its mechanical properties at the curing temperatures associated with the laminate 204. The stabilizing element 300 may also preferably have a relatively high thermal conductivity to improve heat flux through the laminate 204 during curing, to assist in the uniform distribution of heat during the curing of the laminate 204.
[0063] Referring to Figure 8, there is shown an example of a geometric discontinuity 256 that can be associated with the structural element 400 that can be formed as a laminate 204 of uncured composite veneers 216. The geometric discontinuity 256 can comprise a change of cross-section shape 408 in the form of a structural element radius 420 at the intersection of the continuous section 412 and the cap 414 of the "hat" type reinforcement 404. Structural element radius 420 can result in a compression discontinuity 258 in structural element 400. For example, a region of high compaction pressure 332 can at a positive radius 446 of mandrel 444 relative to compaction pressure 324 occurring in the element structural 400 locations outside the radius of structural element 420 and resulting in differential pressure with respect to the high compaction pressure region 332. The high compaction pressure region 332 may occur during operation in the element's vacuum chamber and/or autoclave The localized region of high compaction pressure 332 can result in a thinning of radius 428 in the radius of structural element 420 relative to the nominal thickness of structural element 426 of the structural element 400.
[0064] In Figure 8, radius thinning 428 may occur due to the flow (not shown) of resin 226 away from structural element radius 420. Radius thinning 428 can have an undesirable effect on the fit of structural element 400 with conjugated components (not shown). In addition, radius thinning 428 may have an effect on the outward pull capability (not shown) of structural element 400 and/or the flexural load capacity (not shown) of structural element 400. stabilizing element 300 may advantageously be included at any location in any laminate 204 (eg a panel 206, a structural element 400) having a simply curved shape (not shown - eg simple cylindrical or conical shape) and/or in any laminate 204 having a complex contour shape (not shown - eg aircraft nose shape, mid-wing-to-fuselage fairing shape, etc.).
[0065] Referring to Figure 9, a stabilizing element 300 is shown positioned close to the radius of structural element 420. Advantageously, the stabilizing element 300 has a relatively high rigidity which can result in distribution of compaction pressure 324 that is applied. through the vacuum chamber 326 to the structural element 400. The stabilizing element 300 can minimize or eliminate the high pressure region of compaction 332 (figure 8) which otherwise causes resin 226 to flow out (not shown) and which may otherwise cause a thinning of radius 428 (figure 8) in the radius of structural element 420.
[0066] In the embodiment shown in Figure 9, the stabilizing element 300 can be positioned close to an outer surface 424 of the structural element radius 420. However, the stabilizing element 300 can be positioned anywhere within the laminate 204 of the element For example, stabilizing element 300 may be positioned on top of an outer surface 424 of structural element radius 420 or elsewhere within composite veneers 214. Although Figure 9 illustrates a single of the stabilizing elements 300 Installed within composite veneers 214 on structural element radius 420, any number of stabilizing elements 300 may be installed within composite veneers 214. Stabilizing element 300 may be sized and configured such that stabilizing element edges 308 are extend beyond the tangent points of structural element 422. However, stabilizing element 300 may only is dimensioned and configured such that both of the stabilizing element edges 308 are within the structural element tangent points 422, or such that only one of the stabilizing element edges 308 is between the structural element tangent points 422. Although Figure 9 illustrates stabilizing element 300 positioned at a positive radius 446, the embodiments discussed include installing a stabilizing element 300 close to a negative radius (not shown) of a structural element 400.
[0067] With reference to figure 10, another example of a geometric discontinuity 256 that may be formed in a structural element 400 at a junction 438 of two or more sublaminates 430 is shown. In figure 10, the geometric discontinuity 256 comprises a mass or filler of radius 440 positioned at the junction 438 of the base laminate 434, the primary laminate 432, and the sheath laminate 436 which constitute laminate 204 of the "hat" type reinforcement laminate 404. The radius fill 440 can result in distortion of out-of-plane fibers 244 in the composite veneers 214 positioned adjacent to the radius fill 440. The distortion of out-of-plane fibers 244 can occur during curing of structural element 400 and during application of pressure from compaction 324 to structural element 400. Referring to Figure 11, there is shown a stabilizing element 300 positioned close to infill of radius 440 and installed within base laminate 434 of structural element 400 close to infill of radius 440. of stabilization 300 can minimize or prevent distortion of out-of-plane fibers 244 (FIG. 10) in the composite veneers 214 (FIG. 10). In this regard, the stabilizing element 300 can improve the characteristic strength and rigidity of the structural element 400. In addition, by minimizing the distortion of out-of-plane fibers 244 in the composite veneers 214 adjacent to the radius fill 440, the tensile capacity for outside (not shown) of the "hat" type reinforcement 404 or the outward pull capability (not shown) of other types of spars 402 or structural elements 400 can be improved. In the embodiment shown, stabilizing element 300 may be sized and configured such that stabilizing element edges 308 extend past the fill tangent points of radius 442. However, stabilizing element 300 may be provided at any width, which can slow down or minimize distortion of out-of-plane fibers 244.
[0068] Referring to Figure 12, an example of a compression discontinuity 258 that occurs as a result of applying compression mold plates 500 to panel 206 of cylinder section 104 is shown. cylinder 104, multiple compression mold plates 500 may be required. Figure 12 illustrates three of the compression mold plates 500, removably positionable against panel 206, to provide a surface against which panel 206 can be compacted under compaction pressure 324 (Figure 11) applied by a vacuum chamber. 326 (figure 11) (not shown) on an opposite side of panel 206. Compression mold plate 500 may be formed of relatively rigid material and may be provided as an aid in controlling an external mold line (not shown) and surface finish of the final composite structure 200. To accommodate thermal expansion of the compression mold plates 500 during heating of the cylinder section 104, the compression mold plates can be sized and configured to provide interstices for compression mold plates 504 between the compression mold plate edges 502 of the compression mold plates 500.
[0069] With reference to Figure 13, there is shown a portion of the panel 206 of a cylinder section 104 in a compression mold plate interstice 504 between adjacent compression mold plates 500 and in which the "hat" type reinforcement "404 (figure 12) is omitted for clarity. A vacuum chamber 326 can be applied to an opposite side of panel 206 to apply compaction pressure 324 to panel 206 for consolidation thereof. The compression mold plate interstice 504 between the compression mold plate edges 502 can result in a low compaction pressure region 330. The low compaction pressure region 330 can cause distortion of out-of-plane fibers 244 in the veneers composites 214. Referring to Fig. 14, there is shown a stabilizing element 300 positioned near the interstice 504 between the edges of compression mold plate 502. Advantageously, stabilizing element 300 can be installed within laminate 204 of composite veneers 214 Due to the rigidity of stabilizing member 302 of stabilizing member 300, stabilizing member 300 can cause composite veneers 214 to remain substantially in-plane during application of compaction pressure 330 (FIG. 13). In this way, the stabilizing element 300 can prevent distortion of out-of-plane fibers 244 (FIG. 13) during operation in the vacuum chamber and/or autoclave. In addition, the stabilization element 300 can minimize or prevent the occurrence of visible marks (not shown).
[0070] Referring to Figure 15, there is shown an example of a geometric discontinuity 256 in the form of a pad 250 that may be formed with the panel 206. The pad 250 may comprise a local increase in the amount of composite veneers 214 of the panel. 206. For example, laminate 204 may be formed at a substantially constant thickness and may have a pad 250 comprising a located composite veneer 214 formed over laminate 204. A pad 250 may be provided in the areas of a panel 206 around joints (not shown), holes (not shown), cutouts (not shown), and other features that may constitute tension enhancers (not shown) in laminate 204. In this regard, a pad 250 may be included with panel 206 to reinforce locally the panel 206 to accommodate mounting or mating of components (not shown) to the laminate, or to increase the local stiffness or strength of the laminate 204.
[0071] In Figure 15, although pad 250 is shown as a gradual or staggered increase or formation in the amount of composite veneers 214, pad 250 can comprise any thickness variation in laminate thickness 246. For example, pad 250 can be provided as an abrupt increase in the thickness of laminate 246 or a change in the cross-sectional profile of the composite structure 200. It should also be noted that, although panel 206 in Figure 15 is illustrated as having a flat configuration, panel 206 can be formed in a contoured or curved configuration (not shown), or as a combination of a flat configuration and a contoured or curved configuration.
[0072] With reference to Figure 16, a stabilization element 300 positioned close to a perimeter edge 252 of pad 250 is shown (Figure 15). Stabilizing element 300 may be installed within laminate 204 (FIG. 15) of composite veneers 214. For example, stabilizing element 300 may be installed near a perimeter edge 252 of pad 250. Stabilizing element 300 may be configured such that the stabilizing element edges 308 extend beyond the perimeter edge 252. In the embodiment shown, the stabilizing element 300 can be configured such that each of the stabilizing element edges 308 extends beyond the edge of perimeter 252 of pad 250. Although a single stabilizing element 300 is shown, one or more stabilizing elements 300 may be installed on one or more of the perimeter edges 252 of pad 250.
[0073] With reference to Figure 17, there is shown a cross-sectional illustration of a laminate 204 having a stabilizing element 300 and a composite veneer 214 positioned in a common plane 316 and in which the stabilizing element edges 308 and the edges of veneer 220 are arranged in opposite relationship 320 to each other. In one embodiment, stabilizing element 300 may be provided in a stabilizing element 306 thickness that is approximately equivalent to a multiple of a veneer thickness 222 of the composite veneers 214 positioned immediately adjacent to stabilizing element 300. In one embodiment, the thickness of stabilizing element 306 may be approximately equivalent to a thickness of veneer 222. In another embodiment, the thickness of stabilizing element 306 may be approximately twice or more of the thicknesses of veneers 222. The thickness of veneer 222 may be measured after compaction of the composite veneers 214. As indicated above, the composite veneers 214 can have a veneer thickness 222 in a range from approximately 25.4 µm (1 mil) to approximately 508 µm (20 mils) or greater. However, 222 veneer thickness can be provided in a range from approximately 101.6 µm (4 mils) to approximately 203.2 µm (8 mils). Stabilizing element 300 may have a stabilizing element thickness 306 in a range of from approximately 1 mil to approximately 508 µm (20 mils) although the thickness of stabilizing element 306 may be greater than 508 µm (20 mils).
[0074] With reference to Figure 18, there is shown a cross-sectional illustration of a laminate 204 having a stabilizing element 300 and wherein at least one of the composite veneers 214 is arranged such that the stabilizing element edges 308 and the veneer edges 220 are disposed in superimposed relationship 318 to each other. In this regard, the laminate 204 is configured such that at least one of the composite veneers 214 in a common plane 316 with the stabilizing element 300 is extended up and over the stabilizing element edges 308 and overlapping the stabilizing element edges. stabilization 308. However, panel 206 may be arranged in any of a variety of combinations of overlapping relationship 318 and/or opposite relationships 320 of the stabilizer element edges 308 and the veneer edges 220.
[0075] In an embodiment shown in Figures 17-18, the stabilizing element 300 may be bonded to one or more of the composite veneers 214. For example, a layer of adhesive 314 may be included in the laminate 204 between the stabilizing element 300 and at least one of the composite veneers 214. The adhesive layer 314 may comprise an adhesive material, such as a thermosetting epoxy resin or a thermoplastic resin. The adhesive material may also comprise polyimide resin, bismaleimide resin, polyurethane adhesive, acrylic resin, or any other suitable resin, without limitation. In one embodiment, the adhesive layer 314 can have a thickness in a range from approximately 12.7 µm (0.5 mil) to 50.8 µm (2.0 mils) or greater. Adhesive layer 314 may advantageously facilitate gluing of stabilizing element 300 with one or more immediately adjacent composite veneers 214. A surface treatment 312 may be applied to one or more of stabilizing element surfaces 310 of stabilizing element 300 to improve bonding between the stabilizing element 300 and at least one of the composite veneers 214.
[0076] Referring to figure 19, an illustration of a flowchart of an embodiment of a method 600 of manufacturing a composite structure 200 is shown (figure 20). Step 602 of method 600 may comprise laying a laminate 204 (Figure 20) with a plurality of composite veneers 214 (Figure 20) wherein the laminate 204 may have a compression discontinuity 256 (Figure 20) and/or a geometric discontinuity 256 (figure 20) associated with it. The laying process of laminate 204 can be performed using conventional laying equipment such as a tape laying machine (not shown), and/or laminate 204 can be laid by hand. Laminate 204 may comprise a structural element 400 (Figure 20) formed from a plurality of cured or precured composite veneers 218 (Figure 20) and configured in a desired cross-sectional shape, such as "hat" type reinforcement. 404 illustrated in Figure 5, or in any other cross-sectional format, without limitation.
[0077] Alternatively, laminate 204 (figure 20) may be formed as a structural element 400 (figure 20) comprising a plurality of uncured composite veneers 216 (figure 20). In another embodiment, laminate 204 may be formed as a panel 206 (FIG. 20) comprising a plurality of uncured composite veneers 216, which may be co-cured with one or more structural elements 400. Panel 206 may be provided in a generally flat configuration and/or in a curved configuration, such as in the cylinder section 104 shown in Figure 2. A composite structure 200 (Figure 20) may also be formed by co-gluing one or more structural elements 400 (Figure 20) formed from cured or precured composite veneers 218 (FIG. 20) to a panel 206 formed from uncured composite veneers 216, while simultaneously curing the panel 206 during a co-gluing process.
[0078] Step 604 of method 600 of Figure 19 may include applying at least one stabilizing element 300 (Figure 20) with the composite veneers 214 (Figure 20). For example, one or more of the stabilizing elements 300 may be installed with the composite veneers 214 of laminate 204 (Figure 20), as shown in Figures 7, 9, 11, 14, and 16. In one embodiment, the stabilizing element 300 can be positioned at a depth 322 (figure 17) of no more than approximately ten of the composite veneers 214 below a laminate top surface 210 (figure 17) or laminate bottom surface 212 (figure 17). More preferably, the stabilizing element 300 may be positioned at a depth 322 of no more than approximately two or three of the composite veneers 214 below the upper surface of laminate 210 or lower surface of laminate 212. Alternatively, the method may include applying the element of stabilization 300 at the top (not shown) of the upper laminate surface 210 and/or at the top (not shown) of the lower laminate surface 212.
[0079] Step 606 of method 600 of Figure 19 may include positioning the stabilizing element 300 (Figure 7) near a compression discontinuity 256 (Figure 20) and/or a geometric discontinuity 256 (Figure 20) associated with the laminate 204 (figure 7). For example, stabilizing element 300 can be positioned proximate to a structural element edge 418 of a structural element 400 (figure 7) that can be mounted to a panel 206 (figure 7). In this regard, the edge of structural element 418 may result in the occurrence of a compression discontinuity 258 associated with the panel 206. The method may include positioning the stabilizing element 300 relative to the structural element 400 (FIG. 7) such that a stabilizing element edge 308 (FIG. 7) extends beyond structural element edge 418 (FIG. 7) of structural element 400. For example, stabilizing element 300 can be positioned such that stabilizing element 300 extends across of a low compaction pressure region 330, as may be caused by bridging 328 (Fig. 7) of a vacuum chamber 326 (Fig. 7), as described above.
[0080] Step 606 of method 600 may also include positioning the stabilizing element 300 (figure 9) next to one or more other types of compression discontinuities 258 (figure 20) and/or geometric discontinuities 256 (figure 9), which can be associated with a structural element 400 (figure 9). For example, one or more stabilizing elements 300 may be positioned close to a cross-sectional shape change 408 (Figure 9) in a structural element 406 (Figure 9) cross-section of a structural element 400. Figure 9 illustrates one stabilizing element 300 positioned proximate to a structural element radius 420 (figure 9) of a structural element cross section 406. The stabilizing element 300 may be positioned proximate an outer surface 424 (figure 9) of the structural element radius 420 to promote an even distribution of compaction pressure 324 (figure 9) across the entire laminate 204 (figure 9). Stabilizing element 300 can thus minimize or prevent the occurrence of a high compaction pressure region 332 (FIG. 8) which may otherwise result in differential pressure with respect to the relatively lower compaction pressure 324 on structural element 400 at locations adjacent to structural element radius 420. Such regions of high compaction pressure 332 (figure 8) that may otherwise cause resin 226 to flow (not shown) away from structural element radius 420 and may result at radius thinning 428 (Fig. 8) to structural element radius 420. As shown in Fig. 11, in one embodiment, stabilizing element 300 may also be positioned near a geometric discontinuity 256 comprising a filler of radius 440 at a joint 438 of a plurality of sublaminates 430 of structural element 400, as described above.
[0081] Step 608 of method 600 of Figure 19 may include gluing the stabilizing element 300 to at least one of the composite veneers 214 (Figure 17) using a layer of adhesive 314 (Figure 17). The bond between the stabilizing element 300 (FIG. 17) and the composite veneers 214 can be improved by applying a surface treatment 312 (FIG. 17) to one or more of the stabilizing element surfaces 310 (FIG. 17) of the mating element. stabilization 300. In one embodiment, surface treatment 312 may comprise chemically treating surfaces of stabilization element 310, such as by applying a sol-gel surface treatment (not shown), chemical cleaning, chemical causticizing, and scrubbing. with solvent, or by mechanically treating the surfaces of stabilizing element 310 by abrasive blasting, sanding, sandblasting, abrasion, laser ablation, or any of a variety of other surface treatments 312. Step 608 can including applying a stabilizing element 300 with the composite veneers 214 of the laminate 204 so that the stabilizing element 300 and one of the composite veneers 214 are am positioned on a common plane 316 and a stabilizing element edge 308 and a veneer edge 220 are in generally opposite relationship 320 to each other, as shown in Figure 17. Alternatively, step 608 may include interleaving a stabilizing element 300 ( Figure 18) within the composite veneers 214 (Figure 18) such that at least one of the composite veneers 214 of the laminate 204 extends up and over one or more of the edges of stabilizing element 308 (Figure 18) in superimposed relation 318 to the stabilizing element 300 as shown in figure 18.
[0082] Step 610 of method 600 of Figure 19 may include applying compaction pressure 324 to laminate 204 such as during operation in the vacuum chamber and/or autoclave. Figure 7 illustrates the co-gluing of structural element 400 to panel 206. Structural element 400 may comprise cured or precured composite veneers 218 or non-composite material. Panel 206 may comprise uncured composite veneers 216. Vacuum chamber 326 may be extended over structural member 400 and panel 206 to apply compaction pressure 324 to consolidate and/or cure composite structure 200. The curing process may optionally be performed in an autoclave (not shown) to provide controlled curing conditions including control of vacuum pressure magnitude (not shown), control of heating rate (not shown) of composite veneers 214, control of temperature of cure (not shown), control hold time (not shown), and/or control other cure parameters. During curing, the composite veneers 214 can be heated to reduce the viscosity of the resin 226 (figure 7) and allow a resin 226 to flow and be intermingled with a resin 226 in the adjacent composite veneers 214 (figure 7). Heating the composite veneers 214 can also initiate a crosslinking reaction to cure the composite veneers 214 formed of thermosetting material. Composite veneers 214 formed of thermoplastic material can be heated to a temperature that exceeds the glass transition temperature to reduce the viscosity of resin 226 to promote interweaving of resin 226.
[0083] Step 612 of method 600 of Figure 19 may include slowing the distortion of fibers 244 (Figure 7) in the composite veneers 214 (Figure 7) of the composite structure 200 (Figure 7) using one or more stabilizing elements 300 (Figure 7) 7) which can be positioned in one or more compression discontinuities 258 (figure 20) and/or geometric discontinuities 256 (figure 9), which can be associated with a laminate 204 (figure 7). Compression discontinuities 258 or geometric discontinuities 256 may occur as a result of non-uniform compaction pressure 329, differential coefficients of thermal expansion (CTE) of laminate 204 (e.g., CTE in-plane vs. CTE through-thickness), and/or as a result of differences in the CTE of composite veneers 214 relative to the CTE of other components (not shown). Compression discontinuities 258 and/or geometric discontinuities 256 may also occur at locations that may be susceptible to curing shrinkage (not shown) of resin material 228 (figure 7) in composite veneers 214, in regions of low compaction pressure 330 ( figure 6), in regions of high compaction pressure 332 (figure 8), and/or in places where there is variation in the thickness of laminate 246 (figure 8) such as in pads 250 (figure 15) in a panel 206 (figure 15). However, such compression discontinuities 258 or geometric discontinuities 256 may occur as a result of any factor that may result in a deviation of the fibers 230 (Figure 8) from the desired orientation (not shown) in the composite veneers 214.
[0084] Referring to Figure 20, an illustration of a block diagram of a composite structure 200 having one or more stabilizing elements 300 included with the composite veneers 214 is shown. The composite structure 200 can be made of a laminate 204 to form a structural element 400, a panel 206, or any of a variety of other composite structures 200, without limitation. Laminate 204 may be made from composite veneers 214. Each of the composite veneers 214 may be formed from polymeric material reinforced with fibers 224 and including resin 226 and fibers 230. The fibers 230 in each of the composite veneers 214 may be commonly aligned ( eg, unidirectional) or fibers 230 can be woven in one or more directions to form a fabric (not shown).
[0085] In Figure 20, each of the composite veneers 214 may have a laminate thermal expansion coefficient (CTE) 238. One or more compression discontinuities 258 or geometric discontinuities 256 may be associated with laminate 204. As described above, a compression discontinuity 258 may occur at a location of non-uniform compaction pressure 329 applied to a panel 206 and/or a structural element 400. For example, a compression discontinuity 258 may comprise a structural element edge 418 which may be disposed in a panel 206 and which can generate a low compaction pressure region 330 (FIG. 6) due to the vacuum chamber bridging 328 (FIG. 7) as described above. A geometric discontinuity 256 may comprise a cross-sectional shape change 408 that may be associated with laminate 204, or geometric discontinuity 256 may be a result of other factors. For example, geometric discontinuity 256 may comprise a curvature change 248, such as in a structural element 400. Geometric discontinuity 256 may also comprise a pad 250 or a local increase in the amount of veneers in laminate 204 that forms a panel 206. Geometric discontinuity 256 may also comprise a filler of radius 440 which may be incorporated into a structural element 400.
[0086] Still referring to Figure 20, the composite structure 200 may further include a stabilizing element 300 that can be installed with the composite veneers 214 or applied on top of the composite veneers 214. The stabilizing element 300 may be adhesively bonded to one or more of the composite veneers 214 using a layer of adhesive 314. The stabilizing element 300 may preferably have a relatively high rigidity of stabilizing element 302 at the curing temperature or processing temperature of the composite veneers 214, so that the stabilizing element stabilization 300 can mitigate or minimize the occurrence of out-of-plane fiber distortion 244 (FIG. 13), which may otherwise occur in a laminate 204 due to compression discontinuities 258 or geometric discontinuities 256 associated therewith. In addition, stabilizing element 300 preferably has a stabilizing element 304 CTE that may be substantially similar to laminate CTE 238 in order to minimize the generation of minimal stress in composite structure 200 during the curing process. Stabilizing element 300 may have stabilizing element edges 308 that may preferably be positioned to extend beyond the locations of a compression discontinuity 258 or geometric discontinuity 256 in order to cause the composite veneers 214 to remain substantially in-plane (not shown ) and preventing distortion of out-of-plane fibers 244 during compaction and/or consolidation of the composite veneers 214.
[0087] With reference to figures 21-22, modalities of the exhibition can be described in the context of a method of manufacturing and servicing an aircraft 700, as shown in figure 21, and an aircraft 702, as shown in figure 22. During pre -production, method 700, for example, may include specification and design 704 of the 702 aircraft and acquisition of 706 material. During production, the fabrication of 708 components and subassemblies and the 710 systems integration of the 702 aircraft takes place. The 702 aircraft may then pass certification and provision 712 in order to be placed into service 714. While in service by a customer, the 702 aircraft is scheduled for routine maintenance and service 716 (which may also include modification, reconfiguration , remodeling and others).
[0088] Each of the 700 method processes can be performed or performed by a system integrator, a third party, and/or an operator (eg, a customer). For purposes of this description, a system integrator may include, without limitation, any number of aircraft manufacturers and main system subcontractors; a third party may include without limitation any number of vendors, subcontractors and suppliers; and an operator can be an airline, leasing company, military organization, service organization, and others.
[0089] As shown in Figure 22, the 702 aircraft produced by the example 700 method may include a 718 fuselage with a plurality of 720 systems and a 722 interior. Examples of high-level 720 systems include one or more of a propulsion system 724, an electrical system 726, a hydraulic system 728, and an environmental system 730. Any number of other systems can be included. Although an aerospace example is shown, the principles of the exposed modalities can be applied to other industries, such as the automotive industry.
[0090] The stabilization element 300 (figure 17) and methods incorporated herein may be employed during any one or more of the stages of the production and service method 700. For example, components or subassemblies corresponding to the production process 708 may be manufactured or performed in a manner similar to components or subassemblies produced while the 702 aircraft is in service. Also, one or more of the stabilizing element 300 modalities, method modalities, or a combination thereof may be used during production stages 708 and 710, for example, by substantially speeding up assembly or reducing the cost of an aircraft 702 Similarly, one or more of apparatus modes, method modes, or a combination thereof may be used while aircraft 702 is in service, for example, and without limitation, for maintenance and service 716.
[0091] Further modifications and improvements to the present exhibit may be apparent to those of common knowledge in the art. Thus, the particular combination of parts described and illustrated here is intended to represent only certain modalities of the present exhibition and is not intended to serve as limitations on alternative modalities or devices within the spirit and scope of the exhibition.

权利要求:
Claims (13)
[0001]
1. A composite structure (200), comprising: a laminate (204) having a plurality of composite veneers (214); a compression discontinuity (258) associated with the laminate (204); and a stabilizing element (300) included with the composite veneers (214) and being positioned close to the compression discontinuity (258), characterized by the fact that the laminate (204) comprises a panel (206) formed from the plurality of the composite veneers ( 214); the compression discontinuity (258) comprises a low compaction pressure region (330) positioned proximate a structural element edge (418) of a structural element (400) mounted to the panel (206); and the stabilizing element (300) being included with the composite veneers (214) and being positioned close to the structural element edge (418).
[0002]
2. Composite structure (200), according to claim 1, characterized in that: the laminate (204) comprises a panel (206) formed from the plurality of composite veneers (214); the compression discontinuity (258) comprises a compression mold plate interstice (504) positioned between a pair of compression mold plates (500) positioned removably against the panel (206); and the stabilizing element (300) being included with the composite veneers and being positioned proximate the compression mold plate interstice (504).
[0003]
3. Composite structure (200) according to any one of claims 1 to 2, characterized in that: the stabilizing element (300) has a stabilizing element thermal expansion coefficient (304) than an expansion coefficient thermal laminate (238).
[0004]
4. Composite structure (200) according to any one of claims 1 to 3, characterized in that: the stabilizing element (300) has a stabilizing element rigidity (300) in a range from 718.20 MPa (15 MSI) up to 3830.42 MPa (80 MSI).
[0005]
5. Composite structure (200) according to any one of claims 1 to 4, characterized in that: the stabilizing element (300) is formed of stabilizing element material (300) comprising at least one of composite material cured, ceramic material, and metallic material.
[0006]
6. Composite structure (200) according to any one of claims 1 to 5, characterized in that: the stabilizing element (300) has a stabilizing element thickness (300) that is equivalent to one of the following: a veneer thickness (222) of a composite veneer (214), a multiple of the veneer thickness (222).
[0007]
A method of producing a composite structure (200) as defined in claim 1, comprising the steps of: laying a laminate (204) with a plurality of composite veneers (214), the laminate (204) having at least one of a compression discontinuity (258) and a geometric discontinuity (256) associated with the laminate (204); applying a stabilizing element (300) with the composite veneers (214); and positioning the stabilizing element (300) close to at least one of the compression discontinuity (258) and the geometric discontinuity (256), characterized in that the method further comprises the steps of: laying the laminate (204) as a panel (206) formed from the plurality of composite veneers (214) and having a structural element (400) mounted thereto; and positioning the stabilizing element (300) proximate an edge of the structural element (418).
[0008]
8. Method according to claim 7, characterized in that it further comprises the steps of: applying compaction pressure (330) to the laminate (204); generating a compression discontinuity (258) comprising a region of low compression pressure (330) associated with the structural member edge (418); and slow down, using the stabilizing element (300), fiber distortion in the composite veneers (214).
[0009]
9. Method according to any one of claims 7 to 8, characterized in that it further comprises the step of: laying the laminate (204) as a structural element (400) formed from the plurality of composite veneers (214).
[0010]
10. Method according to claim 9, characterized in that it further comprises the step of: positioning the stabilizing element (300) close to a geometric discontinuity (256) in a cross-section of a structural element (400).
[0011]
11. Method according to any one of claims 9 to 10, characterized in that it further comprises the step of: positioning the stabilizing element (300) close to the geometric discontinuity (256) comprising a change in cross-sectional shape ( 408) of a structural element cross section (400).
[0012]
12. Method according to any one of claims 9 to 11, characterized in that it further comprises the step of: positioning the stabilizing element (300) close to the geometric discontinuity (256) comprising a structural element radius (400) of a cross section of structural element (400).
[0013]
13. Method according to any one of claims 9 to 12, characterized in that it further comprises the step of: positioning the stabilizing element (300) close to the geometric discontinuity (256) comprising a radius filling (440) at a joining of a plurality of underlaminates (430) of the structural element (400).

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US20060060726A1|2006-03-23|Polyimide resin and carbon fiber molded tube clamp

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AU2016247045B2|2020-05-21|Composite structures with stiffeners and method of making the same

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CA2714175A1|2009-08-13|Method for manufacturing a fibre-composite component, fibre-composite component and fibre-composite fuselage component of an aircraft

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EP2977192B1|2019-07-10|Fabric jacketed unidirectional noodle and its method of construction

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CA2858829A1|2013-06-20|An aircraft structure with structural non-fiber reinforcing bonding resin layer

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WO2017081456A1|2017-05-18|Methods and patches for repairing composite laminates

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JP2003071864A|2003-03-12|Method for manufacturing composite material reinforced panel

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JP6732626B2|2020-07-29|Corner tension fitting

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US20160009045A1|2016-01-14|Composite Filler

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US10005267B1|2018-06-26|Formation of complex composite structures using laminate templates

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US20210339843A1|2021-11-04|Stringer assemblies and methods of forming thereof

同族专利:

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公开号 | 公开日

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EP2903801B1|2020-12-30|

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BR112015004962A2|2017-07-04|

---

KR102197337B1|2021-01-04|

---

KR20200008035A|2020-01-22|

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CA2881308C|2017-07-11|

---

WO2014055168A1|2014-04-10|

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EP2903801A1|2015-08-12|

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CN109808199B|2021-11-23|

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CN104703781A|2015-06-10|

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US20140099477A1|2014-04-10|

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CN109808199A|2019-05-28|

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JP2016500586A|2016-01-14|

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KR20150065669A|2015-06-15|

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KR102069926B1|2020-01-23|

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JP6240200B2|2017-11-29|

---

CN104703781B|2019-02-15|

---

CA2881308A1|2014-04-10|

---

US9517594B2|2016-12-13|

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法律状态:
2018-03-06| B06F| Objections, documents and/or translations needed after an examination request according [chapter 6.6 patent gazette]|

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2018-03-13| B06F| Objections, documents and/or translations needed after an examination request according [chapter 6.6 patent gazette]|

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2018-03-20| B06I| Publication of requirement cancelled [chapter 6.9 patent gazette]|Free format text: ANULADA A PUBLICACAO CODIGO 6.6.1 NA RPI NO 2462 DE 13/03/2018 POR TER SIDO INDEVIDA. |

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2019-12-10| B06U| Preliminary requirement: requests with searches performed by other patent offices: procedure suspended [chapter 6.21 patent gazette]|

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2021-01-05| B06A| Patent application procedure suspended [chapter 6.1 patent gazette]|

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2021-07-06| B09A| Decision: intention to grant [chapter 9.1 patent gazette]|

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2021-08-31| B16A| Patent or certificate of addition of invention granted [chapter 16.1 patent gazette]|Free format text: PRAZO DE VALIDADE: 20 (VINTE) ANOS CONTADOS A PARTIR DE 22/08/2013, OBSERVADAS AS CONDICOES LEGAIS. |

优先权:

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申请号 | 申请日 | 专利标题

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US13/644,628|2012-10-04|

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US13/644,628|US9517594B2|2012-10-04|2012-10-04|Composite structure having a stabilizing element|

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PCT/US2013/056255|WO2014055168A1|2012-10-04|2013-08-22|Composite structure having a stabilizing element|

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