• 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:
    composite structures having joints of composite material with metal and method of manufacturing them. the present invention relates to a composite structure comprising stacked assemblies of laminated fiber-reinforced resin layers and metal sheets. the edges of the layers of resin and metal sheets are interspersed to form a joint of composite material for metal connecting the layers of resin with the metal sheets.
    公开号:BR112013003588B1
    申请号:R112013003588-9
    申请日:2011-06-22
    公开日:2020-02-11
    发明作者:Kenneth Harlan Griess；Gary E. Georgeson
    申请人:The Boeing Company；
    IPC主号:
    专利说明:

    Descriptive Report of the Invention Patent for STRUCTURE FOR HYBRID COMPOUND METAL AND RESIN AIRCRAFT AND METHOD OF MANUFACTURING A COMPOUND STRUCTURE.
    Technical Field [001] This description generally refers to composite structures, especially fiber-reinforced resin laminates and deals more particularly with a hybrid compound having a joint of metal-composite material.
    Background [002] Bonding techniques are often used to assemble composite structures. In applications where the composite structure also requires fasteners, the local thickness or calibration of the structure surrounding the fastener may need to be increased in order to support the loads transmitted through the fastener joint. As the thickness of the structure increases, the fastener may need to be stretched, thus adding weight to the structure. In addition, the increased local thickness of the structure can increase the eccentricity of the load path through the fastener joint, which can place undesirable layer loads on the fastener.
    [003] A solution to the problems mentioned above consists of fixing metallic accessories to the composite structure in the area of the fasteners. These metallic accessories can be formed of titanium or similar metals that may not react substantially chemically with compounds reinforced with carbon fiber where they come into contact. Titanium fittings, however, can be relatively expensive, particularly when complex shapes need to be formed.
    [004] Accordingly, there is a need for a composite metal resin joint that can be used to connect substanci
    Petition 870190112114, of 11/01/2019, p. 5/29
    2/17 all metal accessories with substantially all composite resin structures, which is relatively inexpensive and easy to manufacture and which can withstand the loads transferred around the fastener connection points. There is also a need for a composite resin joint with metal that substantially prevents chemical reactions between the entire metallic accessory and the entire composite resin structure.
    Summary [005] The described modalities provide a hybrid floodgate structure that includes a metal joint and fiber-reinforced resin compound that can be used to connect a substantially metallic accessory with a structure substantially all consisting of composite resin. The joint provides a transition between the composite and metallic structures that is suitable for use in higher performance applications, such as aerospace vehicles. This transition from a substantially entirely composed material to a substantially all metallic material can reduce or eliminate the possibility of corrosion and / or problems arising from eccentricity. During the laying of the composite structure, metal sheets are replaced by a number of composite layers, and the transition from the composite layers to the metallic sheets takes place in skewed locations in order to provide an adequate load transfer from the composite to the metal part. The skewed transition results in an interleaving between the composite layers and the metal sheets and creates multiple joining lines that can reduce the occurrence and / or propagation of cracks or separations in the joint. An adhesive located between the metallic sheets joins the sheets in an almost solid metallic accessory.
    [006] According to a described embodiment, a composite structure is provided comprising a laminated pile of bed
    Petition 870190112114, of 11/01/2019, p. 6/29
    3/17 of fiber-reinforced resin and a stack of metal sheets. The metal sheets have edges that are interspersed with the edges of the fiber-reinforced resin layers to form a joint of composite material and metal connecting the fiber-reinforced resin layers with the metal sheets.
    [007] According to another embodiment, a metal and hybrid resin structure is provided comprising a composite resin part, a metal part and a transition section between the resin and metal parts. The resin part includes laminated layers of fiber-reinforced resin, and the metal part includes joined sheets of metal. The transition section includes skewed overlaps between the laminated layers and the metal sheets.
    [008] According to another modality, a hybrid composite metallic part is provided. The part comprises a placement of the fiber-reinforced composite material which is enclosed at an interface location. At the interface site, a metallic layer of the same thickness as the composite material continues to the metallic edge of the part, and the placement is repeated with a compound for the metallic interface that is skewed towards the edge of the part of the previous interface site. A layer of structural adhesive is placed between the metal layers, with the next metal interface for composite material skewed away from the edge of the part to produce a nested joint, and the nested tabs produced by stacking the skewed interface continue to its full thickness of the part without any of the composite layers extending fully to the edge of the part.
    [009] According to another additional modality, a method is provided for the manufacture of a composite structure. The method comprises the formation of a multilayered compound having a composite part and a metallic part, and forming a
    Petition 870190112114, of 11/01/2019, p. 7/29
    4/17 metal joint of composite material between the composite part and the metallic part. The method additionally includes compacting and curing the layers.
    [0010] According to another modality, a method is provided for the production of a hybrid metallic part. The method comprises layering at least one fiber-reinforced composite layer that is terminated at an interface location, and layering an adjacent metal layer where the metallic layer is substantially the same thickness as the reinforced composite layer with adjacent fiber. The layering steps of the composite layers and adjacent metallic layers are repeated to form an interface of composite material and metal which is skewed towards said edge of the part from the location of the previous interface. The method further comprises laying a layer of structural adhesive between the metal layers, and repeating the layering of the composite material and metal layer where the next metal interface for composite material is skewed away from the edge of the part for producing a nested joint.
    9. A hybrid composite metal and resin structure, comprising:
    [0011] a part of composite resin including laminated folds of fiber-reinforced resin;
    [0012] a metallic part including joined sheets of metal; and [0013] a transition section between the composite resin part and the metallic part, the transition section including skewed overlaps between the laminated layers and the metal sheets.
    10. Metal structure and hybrid composite resin, according to clause 9, in which:
    [0014] the laminated layers and the metal sheets are arranged
    Petition 870190112114, of 11/01/2019, p. 8/29
    5/17 layered; and [0015] each of the layers includes one of the metal sheets and a plurality of fiber-reinforced resin layer on a substantially edge-to-edge support.
    11. Metal structure and hybrid composite resin, according to clause 10, in which the thickness of the layers in each of the layers is substantially equal to the thickness of the metallic sheet in the layer.
    20. Method of production of a hybrid metallic part, comprising:
    [0016] the placement of at least one composite layer reinforced with fiber that is closed in an interface location;
    [0017] the placement of an adjacent metallic layer where the metallic layer is the same thickness as the composite layer reinforced by adjacent fiber;
    [0018] the repetition of the steps of laying the composite layers and adjacent metallic layers for the formation of a composite material and metal interface that is skewed in the direction of said edge of the part from the previous interface location;
    [0019] the placement of a layer of adhesive structure between the metallic layers; and [0020] the repetition of layering of composite material and metal where the next interface of metal and composite material is skewed away from the edge of part for the production of a nested joint.
    21. Method, in accordance with clause 20, further comprising:
    [0021] the continuation of the skewed interface stacking of composite and metallic layers to produce nested tongues up to the total thickness of the part without any composite layers if esPetição 870190112114, from 11/01/2019, p. 9/29
    6/17 tending fully to the edge of the part.
    22. Method, in accordance with clause 21, further comprising:
    [0022] a vacuum bag processing the part to remove empty air spaces when laying in layers; and [0023] curing the layered part.
    23. Aircraft structure made of metal and hybrid composite resin, comprising:
    [0024] a plurality of laminated layers forming a fully composite part reinforced with fiber, a totally metallic part and an extension joint of metal and hybrid composite material connecting the composite part and the metallic part, [0025] each of the layers including one plurality of layers of composite resin and a titanium metal foil, where the layers and the metal foil are arranged in a support from edge to edge with each other forming a transition point from composite material to metal in the layer, and where the points transition layers are slanted with respect to each other to form the extension joint; and [0026] a layer of adhesive between the metal sheets to unify the metal sheets; and [0027] where the thickness of the layers in each of the layers is substantially equal to the combined thickness of one of the sheets and the adhesive layer.
    24. Method of fabricating a hybrid metal and composite resin aircraft structure, comprising:
    [0028] the formation of layers including a fully composite part reinforced with fiber, a totally metallic part and an extension joint of metal and hybrid composite material connecting the composite part with the metallic part, forming the placement in caPetition 870190112114, from 01 / 11/2019, p. 10/29
    A layer including laying a plurality of layers where each layer is formed by laying a plurality of layers of composite resin in edge-to-edge support with a titanium metal foil forming a transition point of composite material to metal in the layer;
    [0029] placing a layer of adhesive between the metal sheets to unify the metal part;
    [0030] the formation of the joint between the composite part and the metallic part by the bias of the transition points in the layers with respect to each other;
    [0031] the compression of the layers; and [0032] curing the layers.
    12. Metal structure and hybrid composite resin material, according to clause 9, in which the skewed overlays form an extension joint of composite material for metal between the composite resin part and the metal part.
    13. Metal structure and hybrid composite resin, according to clause 9, additionally comprising an adhesive layer between each of the metal sheets to join the sheets and unify the metallic part.
    14. Metal structure and hybrid composite resin, according to clause 9, in which each of the metal sheets is a titanium alloy.
    15. Hybrid composite metal part, comprising:
    [0033] a layering of composite material reinforced with fiber that is enclosed in an interface location, where a metallic layer of the same thickness as the composite material continues until the metallic edge of the part, and the laying of layers is repeated with a composite material to metal interface that is skewed towards the edge of the part from the location of the previous interface and
    Petition 870190112114, of 11/01/2019, p. 11/29
    8/17 includes a layer of structural adhesive between the metal layers, with the next metal interface for composite material skewed away from the edge of the part to produce a nested joint, and the stacking of skewed interface producing nested tabs that continues to thickness of the part without any of the composite layers extending fully to the edge of the part. Brief Description of the Illustrations [0034] Figure 1 is an illustration of a sectional view of a composite structure having a joint of composite material for metal;
    Figure 2 is an illustration of a perspective view of the composite structure including the composite material for metal joint;
    Figure 3 is an illustration of a perspective view of the area designated as Figure 3 in Figure 2;
    Figure 4 is an illustration of a cross-sectional view of the joint, best illustrated interleaving between the composite layers and the metal sheets;
    Figure 5 is an illustration of a cross-sectional view of two separate joint layers illustrated in Figure 4, also illustrating the application of a film adhesive to the metal sheets;
    Figure 6 is an illustration of an enlarged cross-sectional view of a part of the joint formed by two layers illustrated in Figure 5;
    Figure 7 is an illustration of a broad flowchart of a method of creating a composite structure having the composite joint illustrated in Figures 2 to 4;
    Figure 8 is an illustration of a flow chart illustrating additional details of the method illustrated in Figure 7;
    Figure 9 is a flow chart of another method of fabricating a composite structure having the composite joint illustrated in
    Petition 870190112114, of 11/01/2019, p. 12/29
    9/17
    Figures 2 to 4;
    Figure 10 is an illustration of an aircraft production flowchart and service methodology;
    Figure 11 is an illustration of a block diagram of an aircraft.
    Detailed Description [0035] Referring first to Figure 1, a hybrid composite structure 20 includes a composite resin portion 22 joined to a metal portion 24 by a transition section 25 that includes a composite metal material joint 26. In the illustrated example , the composite structure 20 is a substantially flat composite sheet, however, depending on the application, the structure 20 may have one or more curves, contours or other geometric characteristics. For example, the composite structure 20 may comprise an internal and / or external contoured liner of an aircraft (not shown) that is attached to an aircraft frame part 28 by means of a lap joint 30 and fasteners 32 that pass through the liner 20 into the frame 28.
    [0036] The structure 28 can comprise a composite material, metallic or other rigid material, and the metallic part 24 of the structure 20 can serve as a rigid metallic accessory 24 which is suitable for the transfer of a range of loads and types of loads between the structure 28 and the composite part 20. As will be discussed in more detail below, the metallic part 24 may comprise any of several metals such as, without limitation, titanium which is substantially non-reactive and compatible with the composite part 20 and the structure 28. In a practical embodiment, for example, and without limitation, the composite resin part 22 may comprise a carbon fiber reinforced epoxy, the metallic part 24 may comprise a titanium alloy, and the structure 28 may comprise an alloy of
    Petition 870190112114, of 11/01/2019, p. 13/29
    10/17 aluminum or a compound. The transition section 25 and the joint 26 are strong enough to carry the typical range and load types between the composite resin part 22 and the metallic part 24, including, but not limited to tension, bending, twisting and shearing. Although the illustrated transition section 25 and the joint 26 are formed between a fully composite resin part 22 and a fully metallic part 24, it may be possible to use them to join two different composite structures (not shown) or two metallic structures different (not illustrated).
    [0037] With reference to Figures 1 to 4, a layering of layers of composite material 35 is terminated at an interface location 39 referred to here after as a transition point 39, where a metal sheet or layer 37 of substantially the same thickness that the layers of composite material 35 continues to the metallic edge 24a of the metallic part 24, and the layering is repeated with a composite material-to-metal interface 39 which is skewed towards the metallic edge 24a from the interface location above 39 and includes a layer of metal adhesive structure 45 (see Figures 5 and 6) between the metal layers 37, with the next metal interface for composite material 39 slanted away from the edge of part 24a to produce a nested joint 27. This skewed interface stacking, which produces nested tabs 29 (see Figure 3) is continued until the full thickness of the hybrid composite structure 20 without any layers c 35 extending fully to the metal edge 24a of the all metal part 24.
    [0038] With reference now also to Figures 2 to 4, the composite part 22 of the structure 20 comprises a laminated pile 34 of the fiber-reinforced resin layers 35, and the metallic part 24 of the structure 20 comprises a pile 36 of metal sheets or bed
    Petition 870190112114, of 11/01/2019, p. 14/29
    11/17 of the 37 that are joined to form a substantially unified metal structure. As illustrated in Figures 5 and 6, the composite layers 35 and the metal sheets 37 are arranged in layers 38. Each of the layers 38 comprises one or more of the composite layers 35 on a substantially edge-to-edge support with one of the metal sheets 37 Thus, each of the layers 38 passes through a point 39 of a composite material, that is, layers of composite resin 35, to a metal, that is, sheet metal 37.
    [0039] The transition points 39 are skewed with respect to each other according to a predetermined layering schedule so that layers 35 and metal sheets 37 overlap each other in transition section 25 (Figure 1) . The skew of the transition points 39 creates multiple joining lines that can reduce the occurrence and / or propagation of cracks or separation in the joint 26. The skew of the transition points 39 also results in a form of interleaving of the composite layers 35 and metal sheets 37 within the joint 26 which forms a nested joint 27 between the fully composed part 22 and the fully metallic part 24. This nested joint 27 can also be referred to as an extension joint 26, an extension joint 26 or a multiple lap joint steps 26. The adjacent points between the transition points 39 are spaced from each other in the internal plane direction of the structure 20 in order to achieve a joined joint 26 that exhibits optimal performance characteristics, including resistance to loosening and propagation of inconsistencies such as cracks. In the illustrated example, the nested joint 27 forming the joint 26 is a shape of a double extension joint 26 where the transition points 39 are skewed in opposite directions from a generally central point 55 of maximum overlap. However, other configurations
    Petition 870190112114, of 11/01/2019, p. 15/29
    12/17 joint rations are possible including, but not limited to, a joint of singular length in which the multiple transition points 39 are skewed in a single direction.
    [0040] The composite layers 35 may comprise a fiber reinforced resin, such as without limitation, carbon fiber epoxy, which may be in the form of a unidirectional pre-impregnated tape or fabric. Other fiber reinforcements are possible, including fiberglass, and the use of non-impregnated materials may be possible. The composite layers 35 can have predetermined fiber orientations and are placed according to a predefined layer schedule to match the desired performance specifications. As previously mentioned, the bonded sheets 37 can comprise a metal such as titanium which is suitable for the intended application. In the illustrated example, the stack 36 of sheet metal 37 has a total thickness t1 which is generally substantially equal to the thickness t2 of the laminated stack 34 of layers 35. In the illustrated example, however, t2 is slightly larger than ti by a factor of thickness of several overlapping layers 43 on opposite sides of the stack 37.
    [0041] Figures 5 and 6 illustrate details of two adjacent layers 38 of joint 26 shown in Figures 2 to 4. In this example, each layer 38 comprises four layers 35 having a total collective thickness Ti. The individual metallic sheets 37 of the layers adjacent 38 are joined by means of a structural adhesive layer 45, which can comprise a commercial film adhesive or other forms of a suitable adhesive which is placed between the metal sheets 36 during the layering process.
    [0042] The combined thickness of each metal sheet 37 and an adhesive layer 45 represented as T2 in Figure 5 is substantially equal to the thickness T1 of the composite layers 35 in the layer
    Petition 870190112114, of 11/01/2019, p. 16/29
    13/17
    38. Although not shown in the Figures, a thin film of adhesive can be placed between layers 35 to increase resistance to interlaminar bonding. In a practical embodiment, the metal sheets of titanium alloy 37 can be used, each having a thickness of approximately 0.0063 cm, the film adhesive 45 being approximately 0.012 cm. thick, and four layers of composite carbon fiber epoxy 35 which can be used in each layer 38 having a total collective thickness of about 0.76 cm. Depending on the application, the use of metals other than titanium may be possible. The distance between adjacent transition points 39, and thus the length of the overlap between layers 38, in addition to the thickness and number of composite layers 35 and the thickness of the metal sheets 37 will depend on the requirements of the particular application, including the type and magnitude of the loads that must be transmitted through joint 26, and possibly other performance specifications.
    [0043] The different layers 38 of the joint 26 between two different materials of the composite and metallic parts 22, 24 respectively (Figure 1), make the structure 20 well suited for non-destructive assessments of connection quality using built-in or mounted sensors (not shown) ). The ultrasonic structural waves (not shown) can be introduced in the structure 20 at the edge of the metallic part 24, in the composite part 22 or in the transition section 25. These ultrasonic waves travel through what adds a waveguide formed by the metallic sheets 37 and the interfaces (not shown) between the composite layers 35 and the metal sheets 37. MEMS-based sensors (microelectromechanical), thin piezoelectric sensors (not shown) or other transducers located in the structure 20 can be used to receive ultrasonic structural waves for purposes analysis of the condition of the joining lines at the joint 26.
    Petition 870190112114, of 11/01/2019, p. 17/29
    14/17 [0044] Referring now to Figure 7, a method of creating a composite structure 20 comprises the formation of a multilayer composite layering as illustrated in 65. Layering includes the layering of a composite resin part 22 in step 67, layering of a metallic part 24 in 69. Laying step 65 additionally includes the formation of a joint of composite material and metal between the composite resin part and the metallic part of layering, illustrated in 71.
    [0045] Figure 8 illustrates additional details of the method illustrated in Figure 7. Starting at step 40, the individual metal sheets 37 are trimmed to a desired size and / or shape. Next in 42, the surfaces of the metal sheets 37 are prepared by suitable processes that can include cleaning the sheets 37 with a solvent, drying them, etc., then in 44, the layering is assembled by laying sheets metallic 36 and composite layers 35 in a sequence that is determined by a predefined layer schedule (not shown) that includes a predetermined bias of the transition points 39 between layers 35 and the metallic sheet 36 on each layer 38.
    [0046] During the layering process, the metal sheets 37 are sequenced as layers in the layering, much like the composite layers are sequenced in a layering in a conventional layering process. As illustrated in step 46, the adhesive can be inserted between the metal sheets 37 in order to join them in a unified metal structure. Similarly, although not shown in Figure 8, a bonding adhesive can be inserted between the individual composite layers 35 in order to increase the bond strength between these layers 35. Next, in 48, the laying in
    Petition 870190112114, of 11/01/2019, p. 18/29
    15/17 layers can be compacted using any of the several known compacting techniques, such as vacuum bagging following which the layers are cured in step 50 using autoclave or non-autoclave curing processes. In step 52, the cured composite structure 20 can be trimmed and / or inspected as needed.
    [0047] Figure 9 illustrates another additional modality of a method of manufacturing a hybrid composite part 20. The method begins at step 73 with the placement of at least one composite layer 35 that is terminated at an interface location 39 on a tool proper placement (not shown). At 75, an adjacent metal layer 37 is placed being substantially the same thickness as the adjacent composite material layer 35. As shown in 77, the laying process is repeated with a composite material and metal interface 39 which is skewed in the direction from the metal edge 24a of part 20 from the previous interface location 39. In 79, a layer 45 of structural adhesive is placed between the metal layers 37. Steps 73-79 are repeated successively for the production of a nested joint 27 and a skewed interface stacking forming the nested tabs 29 to the full thickness of the hybrid part 20, without any composite layers 35 extending fully to the metallic edge 24a of part 20. Although not shown in Figure 9, the completed layering is vacuum bagged, processed to remove voids, and is subsequently cured using any suitable curing method.
    [0048] The description modalities can be used in a variety of potential applications, particularly in the transport industry, including, for example, aerospace, marine and automotive applications. Thus, with reference now to Figures 10 and 11,
    Petition 870190112114, of 11/01/2019, p. 19/29
    16/17 the description modalities can be used in the context of manufacturing an aircraft and service method 60 as illustrated in Figure 10 and an aircraft 62 as illustrated in Figure 11. Aircraft applications of the described modalities may include, for example, a wide variety of structural composite parts and components, especially those that require the use of fasteners during the assembly process. During pre-production, illustrative method 60 may include specification and design 64 for aircraft 62 and material acquisition 66. During production, component and sub-assembly manufacturing 68 and system integration 70 for aircraft 62 takes place. Thereafter, aircraft 62 may undergo certification and distribution 72 in order to be put into service 74. While in service by a customer, aircraft 62 is scheduled for routine maintenance and service 76 (which may also include modification, reconfiguration, reform, and so on).
    [0049] Each of the method 60 processes can be performed by a system integrator, a third party, and / or an operator (for example, a customer). For the 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 sellers, subcontractors and suppliers; and an operator can be an airline, a leasing company, military entity, service organization, and so on.
    [0050] As illustrated in Figure 11, aircraft 62 produced by illustrative method 60 may include an aircraft 78 with a plurality of systems 80 and an interior 82. Examples of high-level systems 82 include one or more among a propulsion system 84 , an electrical system 86, a hydraulic system 88, and an environmental system 90. Any number of other systems can be included. The method
    Petition 870190112114, of 11/01/2019, p. 20/29
    17/17 described can be used to manufacture parts, structures and components used in aerial structure 78 or in interior 82. Although an aerospace example is illustrated, the principles of the description can be applied to other industries, such as marine and automotive industries.
    [0051] The systems and methods embodied here can be employed during any one or more of the production stages and service method 60. For example, parts, structures and components corresponding to the production process 68 can be manufactured in a similar way to parts, structures and components produced while aircraft 62 is in service. In addition, one or more apparatus modalities, method modalities or a combination thereof can be used during production stages 68 and 70, for example, by substantially accelerating assembly or reducing the cost of an aircraft 62. Similarly , one or more apparatus modalities, method modalities, or a combination thereof may be used while aircraft 62 is in service, for example, and without limitation, for maintenance and service 76.
    [0052] Although the modalities of this description have been described in relation to certain illustrative modalities, it must be understood that the specific modalities serve the purpose of illustration and not of limitation, since other variations will occur to those skilled in the art.
    权利要求:
    Claims (5)
    [1]
    1. A hybrid metal and resin composite aircraft structure comprising:
    a plurality of laminated layers (38) forming a reinforced fiber, all composite parts (22), all metallic parts (24), and a hybrid composite material articulation joint (26) connecting the composite part (22) with the metallic part (24);
    characterized by the fact that each of the layers (38) includes a plurality of layers of fiber-reinforced resin (35) and a titanium metal sheet (37);
    wherein the layers and the metal sheet (37) are arranged on edge-to-edge support with one another forming a transition point (39) of composite material and metal in the layer (38);
    wherein the plurality of layers (35) of each of the layers (38) forms a laminated stack (34) of fiber-reinforced resin layers (35);
    wherein the titanium metal sheets (37) of each layer (38) form a stack of metal sheets (37) having edges interspersed with the edges of fiber-reinforced resin layers (35) to form a joint of composite material and metal (26) connecting the fiber-reinforced resin layers (35) with the metal sheets (37), where the transition points (39) in the layers (38) are skewed with respect to each other to form a joint articulation and an adhesive layer (45) between the metallic sheets (37) to unitize the metallic sheets (37);
    wherein the foils (37) are a titanium alloy, and the fiber reinforcement of the layers includes carbon; and where the thickness (T1) of the layers (35) in each of the layers (38) is equal to the combined thickness (T2) of one of the foPetition 870190112114, from 11/01/2019, p. 22/29
    [2]
    2/3 metal sheets (37) and the adhesive layer (45).
    2. Composite structure according to claim 1, characterized by the fact that the metallic sheets (37) form a pile of metallic sheets (37) defining a metallic edge (24a) of all metallic parts (24), in which none of the resin layers (35) fully extend on the metal edge (24a) of all metal parts (24).
    [3]
    3. Composite structure according to claim 1, characterized by the fact that the resin layers form a laminated pile of fiber-reinforced resin layers (35); and the metal sheets form a stack of metal sheets (37) having edges interspersed with edges of the fiber-reinforced resin layers (35) to form a joint joint of composite material and metal (26) connecting the layers of reinforced resin with fibers with the metal sheets.
    [4]
    4. Composite structure according to claim 3, characterized in that the stacks of fiber-reinforced resin layers are arranged in layers, each layer including a metal sheet and at least one of the resin layers fiber-reinforced.
    [5]
    5. Method of manufacturing a composite structure, characterized by the fact that it comprises:
    forming a multi-layer composite layer (35) having a fiber-reinforced composite resin part (22) and a metallic part (24), including the formation of a composite and metal joint (26) between the composite resin part (22) and the metallic part (24) of the layer, each of the layers (38) including a plurality of layers (35) of fiber-reinforced composite resin and a titanium metal sheet (37), wherein the layers (35 ) and the metallic foil (37) are arranged on an edge-to-edge support one
    Petition 870190112114, of 11/01/2019, p. 23/29
    3/3 with the other forming a transition point (39) of composite material and metal in the layer (38), wherein the plurality of layers (635) of each of the layers (38) forms a laminated pile (34), wherein the titanium metal sheets (37) of each layer (38) form a stack of metal sheets (36), where the composite resin part (22) includes carbon and the metallic part (24) is an alloy titanium, where forming the layer includes forming each of the layers by placing a plurality of layers of fiber-reinforced resin and a metal foil in support edge to edge with each other to form a transition point between the fiber reinforced composite resin and the metal in the layer, where forming the layer includes skewing the transition points in the layers with respect to each other, the method further comprises:
    unitize the metallic part by placing an adhesive piece between the metallic sheets;
    wherein the thickness (Ti) of the layers (35) in each of the layers (38) is equal to the combined thickness (T2) of one of the metal sheets (37) and each of adhesive (45).
    类似技术:
    公开号 | 公开日 | 专利标题
    BR112013003588B1|2020-02-11|STRUCTURE FOR METAL AND RESIN HYBRID COMPOUND AIRCRAFT AND METHOD OF MANUFACTURING A COMPOUND STRUCTURE
    US11084269B2|2021-08-10|Multi-layer metallic structure and composite-to-metal joint methods
    KR102030192B1|2019-10-08|Composite radius fillers and method of forming the same
    KR102067291B1|2020-01-16|Bonded composite airfoil and fabrication method
    RU2654270C2|2018-05-17|Joint for composite wings
    US9302759B2|2016-04-05|Flexible truss frame and method of making the same
    CA2834721C|2017-03-21|Stacked noodle for high capacity pull-off for a composite stringer
    US8939406B2|2015-01-27|Joining composite fuselage sections along window belts
    CA2857783A1|2015-05-07|Laminated composite radius filler with geometric shaped filler element and method of forming the same
    EP2650120B1|2020-06-17|Multi-layer metallic structure
    JP6714363B2|2020-06-24|Apparatus or method for manufacturing a structure
    同族专利:
    公开号 | 公开日
    BR112013003588A2|2018-03-20|
    EP2605902A1|2013-06-26|
    JP5840687B2|2016-01-06|
    EP2605902B1|2021-02-17|
    US8894801B2|2014-11-25|
    US20120045606A1|2012-02-23|
    KR101859783B1|2018-05-18|
    JP2013539428A|2013-10-24|
    CA2805468C|2016-09-13|
    CN103068565A|2013-04-24|
    US8652606B2|2014-02-18|
    KR20130096232A|2013-08-29|
    WO2012024023A1|2012-02-23|
    CA2805468A1|2012-02-23|
    US20130122236A1|2013-05-16|
    CN103068565B|2016-05-18|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US3711934A|1970-09-17|1973-01-23|Monsanto Co|Method of preparing metal foil/graphite fiber/epoxy resin laminates|
    US3780929A|1971-06-11|1973-12-25|Stack Pac Corp|Packaging construction|
    US4072516A|1975-09-15|1978-02-07|Fiber Materials, Inc.|Graphite fiber/metal composites|
    US4229473A|1978-03-24|1980-10-21|The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration|Partial interlaminar separation system for composites|
    US4334124A|1980-03-26|1982-06-08|Intergraph Corporation|Floating coordinate system|
    US4498139A|1981-10-09|1985-02-05|Moskovskoe Nauchno-Proizvodstvennoe Objedinenie Po Stroitelnomu I Dorozhnomu Mashinostroeniju|Apparatus for designing three-dimensional structures|
    US4524620A|1983-02-07|1985-06-25|Hughes Helicopters, Inc.|In-flight monitoring of composite structural components such as helicopter rotor blades|
    US4666546A|1985-10-18|1987-05-19|Treber Gerald C|Method for validating sequential ordering of laminate plies|
    US4849913A|1987-04-10|1989-07-18|The Boeing Company|Method for the design and construction of composite parts|
    US5006990A|1987-04-10|1991-04-09|The Boeing Company|Method and apparatus for the design and construction of composite parts|
    US5031457A|1988-01-22|1991-07-16|Board Of Regents Of The Ou|System for nondestructively determining composite material parameters|
    US4907164A|1988-09-26|1990-03-06|General Electric Company|Automatically optimized NC tool path generation for machining|
    US5038291A|1989-04-03|1991-08-06|General Electric Company|Computerized ply pattern generation|
    US5119309A|1989-04-26|1992-06-02|General Motors Corporation|Feature based method of designing automotive panels|
    US5216612A|1990-07-16|1993-06-01|R. J. Reynolds Tobacco Company|Intelligent computer integrated maintenance system and method|
    US5365996A|1992-06-10|1994-11-22|Amei Technologies Inc.|Method and apparatus for making customized fixation devices|
    US5951800A|1992-11-18|1999-09-14|Mcdonnell Douglas Corp.|Fiber/metal laminate splice|
    US5500272A|1995-01-11|1996-03-19|Northrop Grumman Corporation|High efficiency load transfer in composite structure|
    US5866272A|1996-01-11|1999-02-02|The Boeing Company|Titanium-polymer hybrid laminates|
    US6030483A|1996-09-10|2000-02-29|Wilson; Graeme Paul|Method of forming laminates using a tessellated core|
    US6253218B1|1996-12-26|2001-06-26|Atsushi Aoki|Three dimensional data display method utilizing view point tracing and reduced document images|
    US5984511A|1997-05-12|1999-11-16|Mcdonnell Douglas Corporation|Knowledge driven composite design optimization process and system therefor|
    US7010472B1|1997-05-12|2006-03-07|Mcdonnell Douglas Corporation|Knowledge driven composite design optimization process and system therefor|
    EP0983141B1|1997-05-28|2002-10-23|Structural Laminates Company|Method for making a laminate and laminate obtainable by said method|
    US6041132A|1997-07-29|2000-03-21|General Electric Company|Computed tomography inspection of composite ply structure|
    US6052631A|1997-08-08|2000-04-18|Management Systems Data Service, Inc. |Method and system for facilitating vehicle inspection to detect previous damage and repairs|
    US6445390B1|1997-12-29|2002-09-03|The United States Of America As Represented By The Adminstrator Of The National Aeronautics And Space Administration|Triangle geometry processing for surface modeling and cartesian grid generation|
    US6278457B1|1998-01-15|2001-08-21|International Business Machines Corporation|Methods and apparatus for performing sampling based synthesis of three-dimensional geometric models|
    US6138056A|1998-03-02|2000-10-24|Therwood Corporation|System and method for maintenance and repair of CNC machines|
    JP2000182081A|1998-12-14|2000-06-30|Suzuki Motor Corp|Method and device for generating analysis model, and storage medium stored with analysis model generating program or analysis model data|
    DE59901643D1|1999-03-20|2002-07-11|Dlr Ev|COMPOSITE MATERIAL WITH A REINFORCED CONNECTION AREA|
    US6356437B1|1999-03-29|2002-03-12|Siemens Dematic Postal Automation, L.P.|System, apparatus and method for providing a portable customizable maintenance support instruction system|
    DE69910174T3|1999-04-07|2012-03-08|Mcdonnell Douglas Corp.|Knowledge-based design optimization process for composites and system for it|
    JP2000298734A|1999-04-12|2000-10-24|Mitsubishi Electric Corp|Cross section generating device for three-dimensional model|
    US7424543B2|1999-09-08|2008-09-09|Rice Iii James L|System and method of permissive data flow and application transfer|
    US6959235B1|1999-10-28|2005-10-25|General Electric Company|Diagnosis and repair system and method|
    US6629302B2|1999-12-22|2003-09-30|Shinji Miura|Circuit board design aiding apparatus, design aiding method, and storage medium storing design aiding program|
    US6484776B1|2000-04-07|2002-11-26|Northrop Grumman Corporation|System for constructing a laminate|
    JP2002036177A|2000-05-26|2002-02-05|Gerber Technol Inc|Device and method for cutting out combination of sheet members|
    US6982711B1|2000-09-18|2006-01-03|Hitachi, Ltd.|Solid shape describing method and device therefor and solid shape design support system using them|
    US6690159B2|2000-09-28|2004-02-10|Eldec Corporation|Position indicating system|
    US6799081B1|2000-11-15|2004-09-28|Mcdonnell Douglas Corporation|Fiber placement and fiber steering systems and corresponding software for composite structures|
    JP2002183432A|2000-12-14|2002-06-28|Ibm Japan Ltd|Data extracting method, data operating method, credit information extracting method, database system, credit commercializing device, storage medium, and computer program|
    US6625618B1|2000-12-15|2003-09-23|Tsunehiko Arai|Maintenance manual interface system and medium containing a computer program product thereof|
    US7079996B2|2001-05-30|2006-07-18|Ford Global Technologies, Llc|System and method for design of experiments using direct surface manipulation of a mesh model|
    US7366643B2|2003-03-20|2008-04-29|Delphi Technologies, Inc.|System, method, and storage medium for determining a packaging design for a container|
    US20030055812A1|2001-09-14|2003-03-20|Xccelerator Technologies, Inc.|Vehicle parts monitoring system and associated method|
    DE10148232B4|2001-09-21|2004-04-15|Vaillant Gmbh|Procedure for troubleshooting, preferably in heating devices|
    US7324103B2|2001-10-30|2008-01-29|Ford Motor Company|System and method of direct mesh manipulation|
    US7171344B2|2001-12-21|2007-01-30|Caterpillar Inc|Method and system for providing end-user visualization|
    US20030145017A1|2002-01-31|2003-07-31|Patton Thadd Clark|Method and application for removing material from documents for external sources|
    GB0208088D0|2002-04-09|2002-05-22|Delcam Plc|Method and system for the generation of a computer model|
    GB0209543D0|2002-04-26|2002-06-05|Rolls Royce Plc|The automation and optimisation of the design of a component|
    US6843565B2|2002-08-02|2005-01-18|General Electric Company|Laser projection system to facilitate layup of complex composite shapes|
    US20040024661A1|2002-08-02|2004-02-05|Girish Shirhatti|Systems and methods for inventory management|
    US7244230B2|2002-11-08|2007-07-17|Siemens Medical Solutions Usa, Inc.|Computer aided diagnostic assistance for medical imaging|
    CA2411898A1|2002-11-15|2004-05-15|Idelix Software Inc.|A method and system for controlling access to detail-in-context presentations|
    US6907313B2|2003-01-17|2005-06-14|Dp Technology Corp.|Method and system for computer aided manufacturing|
    JP4364520B2|2003-01-29|2009-11-18|富士通株式会社|Three-dimensional sheet metal model creation method and computer program|
    JP2004264993A|2003-02-28|2004-09-24|Honda Motor Co Ltd|System and method for supporting integration of component for repair|
    US7809679B2|2003-03-03|2010-10-05|Fisher-Rosemount Systems, Inc.|Distributed data access methods and apparatus for process control systems|
    US20040236561A1|2003-05-21|2004-11-25|Gary Smith|Component management engine|
    JP2005056090A|2003-08-01|2005-03-03|Tsunehiko Arai|Interface system for maintenance manual and program recording medium of same|
    US7159112B1|2003-08-26|2007-01-02|Nvidia Corporation|Decryption of graphics data in a graphics processing pipeline|
    US7115323B2|2003-08-28|2006-10-03|The Boeing Company|Titanium foil ply replacement in layup of composite skin|
    US8262823B2|2003-09-04|2012-09-11|The Boeing Company|Window skin panel and method of making same|
    US7300693B2|2003-09-04|2007-11-27|The Boeing Company|Resin infused transparent skin panel and method of making same|
    US7099725B2|2003-12-02|2006-08-29|The Boeing Company|Alternate ply representation for composite design and manufacturing|
    US6819966B1|2003-12-06|2004-11-16|Paul E. Haeberli|Fabrication of free form structures from planar materials|
    US7243055B2|2004-01-28|2007-07-10|The Boeing Company|Composite stacking sequence optimization for multi-zoned composites|
    US7076323B2|2004-02-06|2006-07-11|The Boeing Company|Apparatus and method for creating sketch-based eggcrate substructures for composite parts|
    CA2563784A1|2004-04-21|2005-11-10|Ingersoll Machine Tools, Inc.|Performing high-speed events "on-the-fly" during fabrication of a composite structure by automated fiber placement|
    CA2563663A1|2004-04-21|2005-11-10|Ingersoll Machine Tools, Inc.|Automated fiber placement using multiple placement heads, replaceable creels, and replaceable placement heads|
    WO2005106317A1|2004-04-29|2005-11-10|Federico Martinez-Cepeda|Stiffened channel|
    US7621435B2|2004-06-17|2009-11-24|The Regents Of The University Of California|Designs and fabrication of structural armor|
    TW200629058A|2004-07-19|2006-08-16|United Technologies Corp|System and method for fault code driven maintenance system|
    US7526418B2|2004-08-12|2009-04-28|Saudi Arabian Oil Company|Highly-parallel, implicit compositional reservoir simulator for multi-million-cell models|
    WO2006023877A2|2004-08-21|2006-03-02|Co-Exprise, Inc.|Methods, systems, and apparatuses for extended enterprise commerce|
    US7365747B2|2004-12-07|2008-04-29|The Boeing Company|Methods and systems for controlling an image generator to define, generate, and view geometric images of an object|
    US20060265198A1|2005-03-01|2006-11-23|Satoshi Kanai|Apparatus and a method of feature edge extraction from a triangular mesh model|
    JP2006350499A|2005-06-14|2006-12-28|Shin Caterpillar Mitsubishi Ltd|Maintenance management device for work machine|
    JP5017909B2|2005-06-22|2012-09-05|コニカミノルタエムジー株式会社|Region extraction apparatus, region extraction method, and program|
    US7423523B2|2005-08-03|2008-09-09|The Boeing Company|Composite ply layup using electronically identifiable tags|
    US7643970B2|2005-11-09|2010-01-05|The Boeing Company|Tape course generation method and apparatus for programming a composite tape lamination machine|
    US7376480B2|2005-11-09|2008-05-20|The Boeing Company|Multihead composite material application machine programming method and apparatus for manufacturing composite structures|
    US7769481B2|2005-12-23|2010-08-03|The Boeing Company|Head assignment system and method|
    US7747421B2|2005-12-23|2010-06-29|The Boeing Company|Head assignment modeling and simulation|
    US7561996B2|2006-01-13|2009-07-14|Chrysler Llc|Automated dimensional drawing generating apparatus|
    US7761183B2|2006-02-13|2010-07-20|Sullivan Douglas G|Methods and systems for producing numerical control program files for controlling machine tools|
    EP1840775A1|2006-03-31|2007-10-03|Airbus Espana, S.L.|Computer-aided method of obtaining a ply model of composite component|
    US7545274B2|2006-04-10|2009-06-09|The Boeing Company|RFID data management system|
    US20080052281A1|2006-08-23|2008-02-28|Lockheed Martin Corporation|Database insertion and retrieval system and method|
    DE102007063608B4|2007-05-23|2011-04-21|Airbus Operations Gmbh|Composite and fuselage cell section with such a composite|
    US7809531B2|2007-06-15|2010-10-05|The Boeing Company|Methods and systems for explicit representation of composite structures|
    US7912602B2|2007-06-29|2011-03-22|Caterpillar Inc.|Visual diagnostic system and subscription service|
    US7555404B2|2007-08-09|2009-06-30|The Boeing Company|Methods and systems for automated ply boundary and orientation inspection|
    US7809454B2|2007-09-17|2010-10-05|The Boeing Company|Method and apparatus for simulating composite panel manufacturing|
    US7720561B2|2007-09-21|2010-05-18|The Boeing Company|Optimizing non-productive part motion in an automated tape laydown machine|
    US8442804B2|2007-10-25|2013-05-14|The Boeing Company|Method and apparatus for composite part data extraction|
    US8321180B2|2007-10-25|2012-11-27|The Boeing Company|Method and apparatus for composite part data extraction|
    US8285407B2|2007-10-25|2012-10-09|The Boeing Company|Method and apparatus for composite part data extraction|
    US20090138139A1|2007-11-26|2009-05-28|Tsai Ta C|Technical data navigation system and method|
    US7983809B2|2007-12-21|2011-07-19|Sikorsky Aircraft Corporation|Aircraft integrated support system |
    US8165703B2|2008-08-18|2012-04-24|Airbus Operations S.L.|Computer assisted method for the advanced design of bent parts of composite material|
    US8209838B2|2008-12-19|2012-07-03|The Boeing Company|Repairing composite structures|
    US8108058B2|2009-02-09|2012-01-31|The Boeing Company|Method of analyzing composite structures|
    JP4952763B2|2009-10-07|2012-06-13|富士通株式会社|Cooperation analysis simulation apparatus, cooperation analysis simulation method, and cooperation analysis simulation program|
    US8620627B2|2009-10-13|2013-12-31|The Boeing Company|Composite information display for a part|
    US8652606B2|2010-08-17|2014-02-18|The Boeing Company|Composite structures having composite-to-metal joints and method for making the same|
    US8333345B2|2010-08-26|2012-12-18|The Boeing Company|Composite aircraft joint|US9586699B1|1999-08-16|2017-03-07|Smart Drilling And Completion, Inc.|Methods and apparatus for monitoring and fixing holes in composite aircraft|
    US9625361B1|2001-08-19|2017-04-18|Smart Drilling And Completion, Inc.|Methods and apparatus to prevent failures of fiber-reinforced composite materials under compressive stresses caused by fluids and gases invading microfractures in the materials|
    US8442804B2|2007-10-25|2013-05-14|The Boeing Company|Method and apparatus for composite part data extraction|
    US8620627B2|2009-10-13|2013-12-31|The Boeing Company|Composite information display for a part|
    US8993084B2|2010-08-17|2015-03-31|The Boeing Company|Multi-layer metallic structure and composite-to-metal joint methods|
    US9522512B2|2010-08-17|2016-12-20|The Boeing Company|Methods for making composite structures having composite-to-metal joints|
    US8652606B2|2010-08-17|2014-02-18|The Boeing Company|Composite structures having composite-to-metal joints and method for making the same|
    JP6690910B2|2014-10-09|2020-04-28|ザ・ボーイング・カンパニーＴｈｅ Ｂｏｅｉｎｇ Ｃｏｍｐａｎｙ|COMPOSITE MATERIAL STRUCTURE HAVING JOINT OF COMPOSITE MATERIAL AND METAL, AND MANUFACTURING METHOD THEREOF|
    EP2650120B1|2012-04-10|2020-06-17|The Boeing Company|Multi-layer metallic structure|
    GB201016279D0|2010-09-28|2010-11-10|Airbus Operations Ltd|Stiffener run-out|
    US8875931B2|2010-11-19|2014-11-04|The Boeing Company|Composite sandwich shell edge joint|
    US8784596B2|2010-11-19|2014-07-22|The Boeing Company|Method for making and joining composite sandwich shell edge joint|
    KR20130059153A|2011-11-28|2013-06-05|현대자동차주식회사|A method for preparing joints of metallic sheets and polymer composites by using adhesive|
    FR2989723A1|2012-04-20|2013-10-25|Astrium Sas|ASSEMBLING TRUNCTIONS OF STRUCTURABLE PARTS|
    US8770134B2|2012-04-30|2014-07-08|Carl E. Baugh|Structurally-interlaminated marine vessel|
    US10655328B2|2012-05-24|2020-05-19|The University Of Kentucky Research Foundation|Structural reinforcement, reinforced structural member and related method|
    GB201209437D0|2012-05-28|2012-07-11|Kitchener Renato|Power supply and battery charger|
    CN102941703A|2012-10-21|2013-02-27|刘君才|Super composite material laminated plate|
    EP2948367B1|2013-01-24|2018-08-01|Saab AB|An aerial skin article with reinforced fastening holes|
    US9040138B2|2013-04-29|2015-05-26|General Electric Company|Composite article including composite to metal interlock and method of fabrication|
    WO2014189946A1|2013-05-21|2014-11-27|Cutting Dynamics, Inc.|Reinforced thermoplastic structural component and process to make the same|
    US9227718B2|2013-06-07|2016-01-05|The Boeing Company|Lower joints between outboard wing boxes and center wing sections of aircraft wing assemblies|
    RU2548435C2|2013-07-18|2015-04-20|Открытое акционерное общество "Национальный институт авиационных технологий" |Connection of sheets of metal composites and method to this end|
    US9524356B2|2013-10-16|2016-12-20|General Electric Company|System and methods of generating a computer model of composite component|
    DE102013221168A1|2013-10-18|2015-05-07|Bayerische Motoren Werke Aktiengesellschaft|component arrangement|
    US9403350B2|2013-12-16|2016-08-02|The Nordam Group, Inc.|Flash control metal bonding|
    JP5972854B2|2013-12-17|2016-08-17|三菱重工業株式会社|Joining method and joined body of fiber reinforced composite member, and manufacturing method of wind turbine blade|
    WO2015105439A1|2014-01-13|2015-07-16|Kjell Lindskog|Method and arrangement for manufacture of a product or completion of a product|
    NL2012458B1|2014-03-17|2016-01-08|Gtm-Advanced Products B V|Laminate of mutually bonded adhesive layers and metal sheets, and method to obtain such laminate.|
    US9719536B2|2014-07-03|2017-08-01|The Boeing Company|Assemblies including shape memory alloy fittings and composite structural members|
    JP6681071B2|2015-02-13|2020-04-15|国立大学法人 東京大学|Adhesive structure|
    DE102016104187A1|2016-03-08|2017-09-14|Profil Verbindungstechnik Gmbh & Co. Kg|functional element|
    KR101846648B1|2016-03-15|2018-04-09|현대자동차주식회사|Composite product with inosculation structure|
    CN107453018B|2016-05-31|2021-03-09|洛阳尖端技术研究院|Metamaterial antenna and forming method thereof|
    US10603873B2|2016-08-29|2020-03-31|Northrop Grumman Innovation Systems, Inc.|Hybrid metal composite structures, rocket cases, and related methods|
    JP6789047B2|2016-09-23|2020-11-25|矢崎総業株式会社|Vehicle ground structure|
    KR101836716B1|2016-10-12|2018-03-08|현대자동차주식회사|Composite parts with inserted metal plate and manufacturing method of the same|
    NL2017849B1|2016-11-23|2018-05-28|Gtm Advanced Products B V|Laminate of mutually bonded adhesive layers and spliced metal sheets|
    JP6810593B2|2016-12-14|2021-01-06|三菱重工業株式会社|Composite material molding method and composite material|
    US10661838B2|2017-05-31|2020-05-26|Honda Motor Co., Ltd.|Multi-material vehicle roof stiffener|
    US10994856B2|2017-10-16|2021-05-04|Rohr, Inc.|Structural panel with splice joint between adjacent core structures|
    US11155056B2|2019-04-08|2021-10-26|The Boeing Company|Methods of making laminated metallic structures|
    法律状态:
    2018-04-10| B06F| Objections, documents and/or translations needed after an examination request according [chapter 6.6 patent gazette]|
    2019-08-06| B06U| Preliminary requirement: requests with searches performed by other patent offices: procedure suspended [chapter 6.21 patent gazette]|
    2019-12-10| B09A| Decision: intention to grant [chapter 9.1 patent gazette]|
    2020-02-11| B16A| Patent or certificate of addition of invention granted|Free format text: PRAZO DE VALIDADE: 20 (VINTE) ANOS CONTADOS A PARTIR DE 22/06/2011, OBSERVADAS AS CONDICOES LEGAIS. |
    优先权:
    申请号 | 申请日 | 专利标题
    US12/857,835|2010-08-17|
    US12/857,835|US8652606B2|2010-08-17|2010-08-17|Composite structures having composite-to-metal joints and method for making the same|
    PCT/US2011/041519|WO2012024023A1|2010-08-17|2011-06-22|Composite structures having composite-to-metal joints and method for making the same|
    [返回顶部]