• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    "Method of shaping complex composite parts using pre-pleated multi-directional continuous fiber laminate" Fiber prepreg (ud) layers (12, 14) are formed into a pre-pleated multidirectional continuous fiber laminate (10) used as a molding compound to form three-dimensional structures (40). laminate indentations are split and folded along fold lines to provide near-final shaped preforms that can be compression molded to form fiber reinforced composite structures having complex shapes.
    公开号:BR112012003657B1
    申请号:R112012003657-2
    申请日:2010-08-21
    公开日:2019-06-25
    发明作者:Matthew Kweder
    申请人:Hexcel Corporation；
    IPC主号:
    专利说明:

    BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates generally to fiber-reinforced composite structures and to the molding materials which are used to make these composite structures. structures. More particularly, the present invention involves the use of unidirectional pre-impregnated tape to form pre-pleated multidirectional continuous fiber laminates which are suitable for shaping complex three-dimensional fiber reinforced composite structures by near-final shape preforming. 2. Description of Related Art The fiber reinforced composite structures typically include a resin matrix and fibers as the two major components. These structures are well suited for use in demanding environments, such as in the aerospace field, where a combination of high strength and low weight is important. The prepreg composite (prepreg) is widely used in the manufacture of composite parts and structures. The prepreg is a combination of uncured resin matrix and fiber reinforcement which is prepared for molding and curing the final composite part. By pre-impregnating fiber reinforcement with resin, the manufacturer can carefully control the amount and location of resin that is impregnated in the fiber network, and ensure that the resin is distributed in the network as desired. The prepreg is a preferred material for use in the manufacture of load bearing structural parts and particularly of load bearing aircraft parts which are used in wings, fuselages, bulkheads and control surfaces. It is important that these parts have sufficient strength, damage tolerance, and other requirements that are routinely set forth for these parts. Unidirectional tape (UD) is a common form of prepreg. Unidirectional tape fibers are continuous fibers that extend parallel to each other. The fibers are typically in the form of bundles of numerous individual fibers or filaments which are referred to as "tow". Unidirectional fibers are impregnated with a carefully controlled amount of uncured resin. The pre-impregnated UD is typically placed between protective layers to form the final UD tape which is wound up for storage or transportation to the industrial facility. The width of the UD tape typically ranges from less than one inch to one foot or more. The one-way tape is not suitable for use as a molding compound to form three-dimensional structures using compression molding techniques. The parallel orientation and continuous nature of the fibers in the UD tape causes bundling or clogging when the UD tape is forced to conform to the characteristics of the complex part. As a result, the manufacture of complex three-dimensional pieces using UD tape has been limited to a laborious process wherein individual UD tape folds are applied directly to a three dimensional mold which is subsequently processed in an autoclave or other molding apparatus. This laying procedure using UD tape tends to be a long and expensive process.
    Molding compounds which have been found suitable for compression molding complex parts commonly employ randomly oriented short fibers which more readily conform to the characteristics of the part. However, the use of these short fibers, when mounted on a random basis, introduces local weight variation. Weight variation creates a number of problems. For example, it contributes to the complexity of the pleat design, which has to accommodate all possible total weight results by assembling several highly variable pleat layers for a particular part geometry. Local weight variations on the random basis of short fibers also contribute to irregularities during molding because low weight areas are offset by high weight areas. This compensation process unpredictably differs from one molded part to the next, and also differs between different characteristics of a given part. As a result, it is difficult for the designer to predict and determine whether the design of a molding compound will be suitable for forming the desired part.
    In view of the above, there is a continuing need to provide prepreg molding methods which are suitable for use in compression molding of fiber reinforced composite structures having a relatively complex shape. The need for this type of method is especially important in those situations where the strength of the part is a highly relevant factor.
    SUMMARY OF THE INVENTION
    According to the present invention, it has been found that unidirectional tape (UD) can be formed in a pre-pleated multidirectional laminate which can be processed to be suitable for molding. The pre-pleated laminate is formed by taking individual layers of UD and placing them on top of each other so that the layers or pleats of the laminate contain fibers extending in different directions. The pre-pleated laminate of the present invention does not have any significant weight variations, which could cause variations during molding that could affect structural performance of the three-dimensional structures. In addition, it has been found that the multidirectional nature of the pre-pleated laminate allows the parts designer to cut slits in the laminate and then bend the laminate in a shape close to the final shape of the three-dimensional piece. Formation in the near-final or "preforming" shape of the workpiece to be cured avoids bending and sealing of fiber which has previously presented problems when UD tape was used directly as a molding compound. The present invention is directed to methods for forming fiber reinforced composite structures wherein a multidirectional pre-pleated laminate is folded into a preform, which is then cured to form the composite structure. The pre-pleated laminate includes at least a first prepreg layer and a second prepreg layer, wherein each of the prepreg layers comprises an uncured resin matrix and unidirectional fibers. As one aspect of the present invention, the directions of the unidirectional fibers in the two prepreg layers are different. The folding of the laminate to form a preform provides a person with the ability to control fiber orientation during molding and thereby produces very strong three-dimensional pieces which are particularly well suited for use in aircraft and other aerospace applications. The present invention is also directed to the fiber reinforced composite structures which are produced using the laminate. In addition, the preforms which are formed using cutouts from laminate are encompassed by the invention. In addition, aerospace assemblies, such as an aircraft, which include fiber reinforced composite structures that are produced using the pre-pleated laminate are encompassed by the invention.
    As one aspect of the present invention, the laminate cutouts may be folded along at least two fold lines intersecting one another to form the three-dimensional piece. Further, the laminate is split at the intersection of the folding lines in order to increase the folding process and prevent fiber fouling. The inclusion of slits also allows the designer to form folded preforms, where the fibers are oriented, as desired, to meet the reinforcement design requirements. The foregoing description and many other aspects and concomitant advantages of the present invention will be better understood by reference to the detailed description below when taken in conjunction with the accompanying drawings.
    BRIEF DESCRIPTION OF THE DRAWINGS Fig. 1 is a diagrammatic representation showing how individual layers of unidirectional prepreg or tape (UD) are oriented to form a four-layer quasi-isotropic laminate which is suitable for use in the manufacture of a preferred laminate according to the present invention. FIG. 2 is a diagrammatic representation showing the fiber orientation that results when two quasi-isotropic four-layer laminates, as shown in Fig. 1, are combined to form a preferred exemplary eight layer symmetrical quasi-isotropic laminate according to the present invention. FIG. 3 shows an example eight-layer symmetrical quasi-isotropic laminate according to the present invention. FIG. 4 shows cutouts in the laminate which are made to form an exemplary flanged support structure. FIG. 5 is a preform of an example flanged support structure which has been formed by folding and matching the cutouts shown in Fig. 4. Fig. 6 is a flanged example support structure which is formed when the preform of Fig. 5 is cured. FIG. 7 is an exemplary cutout of the laminate which is used to make an aircraft clip structure which is used to join two pieces of aircraft of primary structure. FIG. 8 is a preform of an exemplary aircraft clip structure which is formed by folding the cutout shown in Fig. 7. Fig. 9 is an exemplary aircraft clip structure which is formed when the preform of Fig. 8 is cured.
    DETAILED DESCRIPTION OF THE INVENTION The present invention involves the formation and use of multidirectional pre-pleated laminates which are used in compression molding processes to form three-dimensional pieces. The invention is particularly well suited for making three dimensional parts having complex shapes and which are designed to withstand heavy loads. Although the present invention is primarily used in the aerospace industry, the process may also be used in accordance with the present invention to make three-dimensional structures which are suitable for a wide variety of applications where high strength and light weight are desired. The following detailed description is directed primarily to the production of aircraft parts as preferred types of structures made using the present invention. It will be understood that the invention also has wide application in the production of other types of high strength parts, such as any number of complex three-dimensional parts which are used in the automotive, railway, marine, energy, and sports industries.
    FIGS. 1-3 depict how a pre-pleated symmetrical quasi-isotropic laminate of example eight layers 10 is made in accordance with the present invention. As shown in Fig. 1, a first layer of prepreg 12 and second layer of prepreg 14 are provided. Each of the prepreg layers 12 and 14 includes an uncured resin matrix and unidirectional fibers 16 and 18, respectively. The unidirectional fibers 16 in prepreg 12 are oriented in the 0ø direction. The unidirectional fibers 18 in the prepreg 14 are oriented in the direction of 45ø. As further shown in Fig. 1, a second set of two layers of prepregs 12a and 14a are provided. The two layers 12a and 14a are identical to the layers 12 and 14, except that they have been rotated by 90 ° so that the unidirectional fibers 16a are oriented at 90 ° and the unidirectional fibers 18a are oriented at -45 °. The four layers 12, 14, 12a and 14a are combined, as shown in Fig. 1 to form a four layer quasi-isotropic laminate 20.
    As shown in Fig. 2, a second four-layer quasi-isotropic laminate 20a is provided, which is the same laminate 20, except that it has been turned to provide unidirectional fiber orientations which are a mirror image of the laminate 20. The two four layers 20 and 20a are combined as shown in reference numeral 30 in Figs. 2 and 3 to form the preferred eight-layer pre-pleated symmetrical quasi-isotropic laminate. Laminates having more or less 8 unidirectional prepreg layers are possible as long as they can be split and folded to forming preforms which do not undergo fiber bundling or sealing when the preform is molded / cured by compression. For example, the laminates may be made so that the number of unidirectional prepreg layers is up to 16 or more, or only 2. The term "prepunched" as used herein means that the various layers of UD are combined to form a multiple pleated laminate before the laminate is cut and / or folded to form a near-finished shaped preform. As shown for preferred laminate 10, adjacent layers of the laminate may have the same direction, provided that at least two of the layers have UD fibers that are oriented in a different direction. It is also important to note that the UD fiber directions of 0 °, ± 45 ° and 90 ° are preferred fiber fiber orientations. A quasi-isotropic orientation, as shown in Figs. 1-3, is preferred. However, a wide variety of other fiber orientations are possible. The particular combination of UD fiber layers and UD fiber orientations is chosen by the part designer depending on the desired fiber orientation for the composite structure that is formed after the laminate has been split, if necessary, and folded to form the three-dimensional shape of the final piece. Unidirectional tape (UD) is the preferred type of prepreg layer that is used to form the pre-pleated laminate of the present invention. The one-way tape is available from commercial sources, or it may be manufactured using known prepreg forming processes. The dimensions of the UD tape can vary greatly depending on the number and size of the cutouts that are required to form the desired three-dimensional composite structure. For example, the width of the UD tape (the dimension perpendicular to the UD fibers) may range from 0.5 inch to a few feet or more depending on the size and number of the desired cutouts. The tape will typically be 0.004 to 0.012 inches (0.01 to 0.03 cm) thick, the length of the UD tape (the size parallel to the UD fibers) may range from 0.5 inch (1.3 cm) to a few feet (one meter) or more depending on the size and number of cutouts you want. The UD or prepreg tape may contain from 25 to 45 weight percent of an uncured resin matrix. Preferably, the amount of uncured resin matrix will be from 30 to 40 weight percent. The resin matrix may be any of epoxy resins, bismaleimide resins, polyimide resins, polyester resins, vinyl ester resins, cyanate ester resins, phenolic resins or thermoplastic resins. Exemplary thermoplastic resins include polyphenylene sulfide (PPS), polysulfone (PS), polyether-teretone (PEEK), polyetherketone ketone (PEKK), polyether sulfone (PES), polyetherimide (PEI), polyamide-imide (PA1). Epoxy resins which are cured with a thermoplastic are preferred resin matrices. Resins which are typically present in UD tape of the type used in the aerospace industry are preferred. The UD fibers may be carbon, glass, aramid or any other fiber material which is typically used in the manufacture of composite parts that are used in high stress environments. The fibers may contain from a few hundred filaments to 12,000 or more filaments. Preferred UD fibers are carbon fibers.
    One commercially available one-way pre-impregnant is Hex-Pleated® 8552, which is available from Hexcel Corporation (Dublin, California). Hex-Pleated®8552 is marketed in a variety of unidirectional ribbon configurations containing an amine cured cured epoxy resin matrix in amounts ranging from 34 to 38 weight percent, and UD carbon or glass fibers having from 3,000 to 12,000 filaments. The fibers typically account for 60 percent volume of UD tape. The formation of an aircraft flange support structure 40 according to the present invention is shown in Figs. 4-6. The flange support structure 40 is designed to connect two pieces of aircraft together. The two aircraft parts 42 and 44 are shown dashed in Fig. 6. The aircraft parts 42 and 44 are elements of the aircraft's primary load-bearing structure. The term "load bearing" means that the part is designed to have sufficient strength and rigidity to withstand a given stress or load without failure. Typical loads that are supported by these parts are on the order of 1000 pounds or more. Loads in the order of 6000 pounds or higher are not uncommon in parts made with this material or process. The flanged support structure 40 must be capable of supporting the same or larger loads as the aircraft primary parts 42 and 44 in order to prevent flange failure 40 which connects the two load bearing pieces together. The aircraft parts 42 and 44 may be affixed to the flange support 40 by means of securing through holes 46 or by adhesive attachment. The flange support structure 40 is made using six cut-outs (A-F), as shown in Fig. 4. The cutouts are cut from the preferred symmetrical quasi-isotropic laminate 10. Linear slits are formed in cutout "A" as shown in reference numeral 48. Cutout A is folded along fold lines 50, 52 and 54, as shown in dashed lines. The slots 48 are located at the intersection of fold lines 50 and 54 and at the intersection of fold lines 52 and 54. The slots 48 extend from the inside of the cutout A at the intersection of the fold line to the edge of the cutout. The slots 48 are collinear with fold lines 50 and 52. The cutout B is folded in dashed lines 56 and 58. Since the two fold lines do not intersect, no crevices are required. Cutouts C and D are mirror images of one another with fold lines being shown dashed. The slits are located at the intersections of the folding lines in the cutouts C and D in the same manner as the cutout A. The cutouts E and F are mirror images of each other with the folding lines also being shown dashed. The slits are located at the intersections of the folding lines in the cut-outs E and F in the same manner as the cut-out A. The slits function as relief cuts which allow several cut-off sections to be folded in position without deformation of the fiber orientation, clogging the cutouts or flowing resin matrix where desired.
    The cutouts A-F are folded and pooled to form a preform 60, as shown in Fig. 5. The fold lines may be slightly heated, if desired, to reduce the viscosity of the resin matrix and make it easier to fold the cutouts. However, the heating should be kept to a minimum, to prevent inadvertent premature curing of the cut. In general, the folding line should be heated just enough to decrease the viscosity of the resin matrix without initiating cure of the entire cut. The temperature will vary depending on the type of resin matrix used. The folding lines should be heated only for a time sufficient to allow the laminate to be curved in the desired shape. The folding lines should be heated for as short a time as possible, and preferably in the order of no more than a few minutes, to avoid localized curing of the cutout.
    An advantage of the present invention is that the preform 60 can be shaped close to the shape of the end piece. The preform is preferably "formed in near-finished shape". The near-end formate means that the dimensions of the preform are within at least 3 mm of the molded dimensions of the cured fiber reinforced composite structure. Preforms that are undersized by more than 3 mm of the molded dimensions of the cured fiber reinforced composite structure are possible. For example, a preform may be designed to be undersized by 25 mm or more of the molded dimensions of the cured piece depending on the size, geometry, and expected structural performance of the final cured piece. The preform 60 may be partially cured in order to increase the viscosity of the resin matrix, in order to ensure that the preform retains the desired near-final shape. The resin matrix may be partially cured (advanced) by any of the known partial curing processes, so long as the viscosity of the resin is adjusted, so that the shape of the preform 60 is retained. A process commonly known as "Stage B" is a preferred process for advancing the resin matrix prior to compression molding or other curing process. Curing of the preform 60 may be accomplished by any of the molding or curing protocols commonly used in the aerospace industry Compression molding is a preferred process for converting the preform 60 into the final aircraft flange support 40. This process involves applying pressure to the preform 60 by means of a mold closed in a press The curing pressures, temperatures and times will vary depending on a number of variables, including the resin matrix type and the preform size. Such curing variables are known to those skilled in the art. various types of resin systems, including epoxy resins, bismaleimide resins, polyimide resins, polyester resins, vinylester resins, cyan esters resins The preforms formed in near-final shape are formed and partially cured, if necessary, to control resin and fiber flow and to form a three-dimensional piece formed in near-final shape. The formation of an aircraft clip structure 70 according to the present invention is shown in Figs. 7-9. The clip structure 70 is designed to connect two pieces of primary structure aircraft together. The two primary frame aircraft parts 72 and 74 are shown dashed in Fig. 9. The clip structure 70 is made using one or more cutouts as shown in reference numeral 78 of Fig. 7. The cutouts are cut from the preferred symmetrical quasi-isotropic laminate 10. A linear slit is formed in the cutout as shown in reference numeral 80. The cutout 78 is folded along the fold lines 82 and 84 as shown in dashed lines . The slot 80 is located at the intersection of fold lines 82 and 84. The slot 80 extends from the inside of the cutout at the intersection of fold line to the cut edge. The slit 80 is collinear with the fold line 82. The cutout 78 is folded to form a clip preform 90 as shown in Fig. 8. The clip preform 90 is formed in the near-ended shape so that it is close to the shape of the end clip 70. The orientation of the fibers in the preform 90 is shown in simplified form as lines 92 to demonstrate that the orientation of the fibers changes as the cutout 78 is folded to form the preform 90. The redirection of the fibers UD occurring during the bending step needs to be taken into consideration by the designer in order to provide the desired fiber orientation in the pre- 90 and in the final clip 70.
    Like the flanged support preform 60, the clip preform 90 may be partially cured in order to increase the viscosity of the resin matrix in order to ensure that the preform forms properly in the cured dimensions of the desired part shape. The "B-stage" is the preferred process for advancing the resin matrix of the clip preform 90 prior to compression molding or other curing process. The clip preform 90 may be cured using any of the usual curing or molding protocols in the same manner as described above for curing the flanged support preform 60. Compression molding is a preferred process for converting the preform shaped 90 in the final aircraft clip 70.
    Having thus described exemplary embodiments of the present invention, it is important to be observed by those of skill in the art that the disclosures contained herein are exemplary only, and that various other alternatives, adaptations and modifications may be made within the scope of this invention. invention. Therefore, the present invention is not limited by the above-described embodiments, but is limited only by the following claims.
    权利要求:
    Claims (20)
    [1]
    A method for forming a fiber reinforced composite structure, comprising the steps of providing a multidirectional pre-pleated laminate comprising an interior and an edge, said laminate comprising at least a first layer of prepreg and a second layer of prepreg wherein each of said prepreg layers comprises an uncured resin matrix and unidirectional fibers, and wherein the direction of said unidirectional fibers in said first prepreg layer is different from the direction of said unidirectional fibers in said second layer of prepreg and wherein said laminate comprises a first flap portion and a second flap portion; folding said laminate to form a preform, wherein said first flap portion and said second flap portion overlap and are pushed together to form a portion of the preform; and curing said preform to form said fiber reinforced composite structure, wherein the portion of said preform that is formed by said first and second flap portions is cured to form a cured portion of said composite structure.
    [2]
    A method for forming a fiber reinforced composite structure according to claim 1, characterized in that said laminate comprises at least two folding lines along which said laminate is pleated during formation of said preform, and said fold lines intersecting each other.
    [3]
    A method for forming a fiber reinforced composite structure according to claim 1, characterized in that said laminate comprises at least one linear slit extending from the interior of said laminate to the edge of said laminate, said linear slit defining an edge of said first flap portion and said second flap portion.
    [4]
    A method for forming a fiber reinforced composite structure according to claim 2, characterized in that said laminate comprises at least one linear slit extending from the interior of said laminate to the edge of said laminate, said linear slit defining a the edge of said first flap portion and said second flap portion and wherein said linear slit is collinear with at least one of said folding lines.
    [5]
    A method for forming a fiber reinforced composite structure according to claim 4, characterized in that said linear slot extends from the intersection of said fold lines to the edge of said laminate.
    [6]
    A method for forming a fiber reinforced composite structure according to claim 1, characterized in that said laminate comprises at least one fold line, and wherein said fold line is heated prior to folding said laminate to form said fold line preform.
    [7]
    A fiber-reinforced composite structure that is made according to a method CHARACTERIZED by comprising the steps of providing a multidirectional pre-pleated laminate comprising an interior and an edge, said laminate comprising at least a first layer of prepreg and a second layer of prepreg, wherein each of said prepreg layers comprises an uncured resin matrix and unidirectional fibers, and wherein the direction of said unidirectional fibers in said first prepreg layer is different from the direction of said unidirectional fibers in said second layer of prepreg and wherein said laminate comprises a first flap portion and a second flap portion; folding said laminate to form a preform, wherein said first flap portion and said second flap portion overlap and are pushed together to form a portion of the preform; and curing said preform to form said composite structure, wherein the portion of said preform that is formed by said first and second flap portions is cured to form a cured portion of said composite structure.
    [8]
    The fiber reinforced composite structure of claim 7, characterized in that said laminate comprises at least two folding lines along which said laminate is pleated during formation of said preform, folding intersect with each other.
    [9]
    A fiber reinforced composite structure as claimed in claim 7, characterized in that said laminate comprises at least one linear slit extending from the interior of said laminate to the edge of said laminate, said linear slit defining an edge of the laminate. said first flap portion and said second flap portion.
    [10]
    The fiber reinforced composite structure of claim 8, characterized in that said laminate comprises at least one linear slit extending from the interior of said laminate to the edge of said laminate, said linear slit defining an edge of said laminate first flap portion and said second flap portion and wherein said linear slit is collinear with at least one of said fold lines.
    [11]
    A fiber reinforced composite structure according to claim 10, characterized in that said linear slot extends from the intersection of said fold lines to the edge of said laminate.
    [12]
    The fiber reinforced composite structure of claim 7, wherein said composite laminate comprises at least one folding line, and wherein said folding line is heated prior to folding said laminate to form said folded pre- form.
    [13]
    A preform for use in forming a fiber reinforced composite structure, said preform being made according to a method, comprising the steps of providing a multidirectional pre-pleated laminate comprising an interior and an edge, said laminate comprising at least a first layer of prepreg and a second layer of prepreg, wherein each of said layers of prepreg comprises a matrix of uncured resin and unidirectional fibers, and wherein the direction of said fibers unidirectional in said first layer of prepreg is different from the direction of said one-way fibers in said second layer of prepreg and wherein said laminate comprises a first flap portion and a second flap portion; and folding said laminate to form said preform, wherein said first flap portion and said second flap portion overlap and are pushed together to form a portion of the preform.
    [14]
    A preform according to claim 13, characterized in that said laminate comprises at least two folding lines along which said laminate is pleated during formation of said preform, and wherein said fold lines intersect between each other.
    [15]
    A preform according to claim 13, characterized in that said laminate comprises at least one linear slit extending from the interior of said laminate to the edge of said laminate, said linear slit defining an edge of said first laminate portion of flap and said second flap portion.
    [16]
    A preform according to claim 14, characterized in that said laminate comprises at least one linear slit extending from the interior of said laminate to the edge of said laminate, said linear slit defining an edge of said first laminate the flap portion and said second flap portion and wherein said linear slit is coli-inlaid with at least one of said fold lines.
    [17]
    A preform according to claim 16, characterized in that said linear slit extends from the intersection of said fold lines to the edge of said laminate.
    [18]
    A preform according to claim 13, characterized in that said composite laminate comprises at least one folding line, and wherein said folding line is heated prior to folding said laminate to form said preform.
    [19]
    Aerospace assembly, CHARACTERIZED in that it comprises a first piece of aircraft; a second piece of aircraft; a fiber reinforced composite structure according to claim 7 connecting said first aircraft part to said second aircraft part.
    [20]
    Aerospace assembly according to claim 19, characterized in that said first and second aircraft parts are load bearing elements of said aircraft.
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    法律状态:
    2018-04-10| B06F| Objections, documents and/or translations needed after an examination request according [chapter 6.6 patent gazette]|
    2019-02-12| B06T| Formal requirements before examination [chapter 6.20 patent gazette]|
    2019-06-04| B09A| Decision: intention to grant [chapter 9.1 patent gazette]|
    2019-06-25| 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 21/08/2010, OBSERVADAS AS CONDICOES LEGAIS. (CO) 20 (VINTE) ANOS CONTADOS A PARTIR DE 21/08/2010, OBSERVADAS AS CONDICOES LEGAIS |
    2021-06-22| B21F| Lapse acc. art. 78, item iv - on non-payment of the annual fees in time|Free format text: REFERENTE A 11A ANUIDADE. |
    2021-10-13| B24J| Lapse because of non-payment of annual fees (definitively: art 78 iv lpi, resolution 113/2013 art. 12)|Free format text: EM VIRTUDE DA EXTINCAO PUBLICADA NA RPI 2633 DE 22-06-2021 E CONSIDERANDO AUSENCIA DE MANIFESTACAO DENTRO DOS PRAZOS LEGAIS, INFORMO QUE CABE SER MANTIDA A EXTINCAO DA PATENTE E SEUS CERTIFICADOS, CONFORME O DISPOSTO NO ARTIGO 12, DA RESOLUCAO 113/2013. |
    优先权:
    申请号 | 申请日 | 专利标题
    US12/561,492|US8263205B2|2009-09-17|2009-09-17|Method of molding complex composite parts using pre-plied multi-directional continuous fiber laminate|
    US12/561.492|2009-09-17|
    PCT/US2010/046253|WO2011034684A1|2009-09-17|2010-08-21|Method of molding complex composite parts using pre-plied multi-directional continuous fiber laminate|
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