• 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:
    The present invention relates to on-site repairs of a composite structure (10) such as an aircraft structure in which a concave recess (14) is machined in the damaged zone, a non-polymerized composite part (16) is set into said recess and part (16) is polymerized under pressure using tooling (18) comprising a stack of plates (32) that increase the pressure applied to the composite part. The present invention thus obtains invisible repairs with mechanical characteristics that are similar to those of the initial structure.
    公开号:US20010008161A1
    申请号:US09/780,403
    申请日:2001-02-12
    公开日:2001-07-19
    发明作者:Fabienne Kociemba；Daniel Soro
    申请人:Airbus Group SAS；
    IPC主号:B29C43-3642
    专利说明:
    [0001] 1. Field of the invention [0001]
    [0002] The present invention relates to tooling designed to repair a composite structure with a damaged zone without the structure having to be displaced to a repair site different from that in which the structure is located. [0002]
    [0003] The term “composite structure” as used here and in the text that follows refers to a structure composed of long fibers, such as carbon or other fibers, arranged in superimposed layers or folds that are embedded in a resin matrix. [0003]
    [0004] The invention also relates to a method for repairing a composite structure using this kind of tooling. [0004]
    [0005] The tooling and method of the invention may be used in any situation where it is not possible or too impractical to send the structure to a specialized workshop or put it through an autoclave. One of the principle applications for the tooling and method of the invention thus relates to repairing composite structures on aircraft. [0005]
    [0006] 2. Background art [0006]
    [0007] Composite materials enable mechanical characteristics to be obtained that are perfectly controlled and comparable to those obtained using metal, while being considerably lighter. Furthermore, the manufacturing techniques of these materials are continually being improved, both in terms of the shapes and measurements of the parts and the length of manufacture time and cost. It is for these various reasons that composite structures are being used more and more frequently in aircraft. [0007]
    [0008] However, compared to metal structures, composite structures have the drawback of being more fragile when they are subjected to impact. When a composite structure is violently hit by an object variable degrees of breaking or cracking occur in the zone of impact. The increase in the number of composite structures used in aircraft has therefore led to significant improvements in the techniques used to repair these structures. [0008]
    [0009] Two criteria should be adopted when the quality of a repair on an aircraft composite structure is evaluated: the appearance of the repaired zone and the mechanical characteristics of the structure in this zone. Furthermore, it is essential for the repair work to be carried out on-site, i.e. in the place where the aircraft is located, in order for costs to be reduced and to limit aircraft downtime to a minimum. In practice this rules out the use of autoclaves. [0009]
    [0010] Appearance is an important criterion given that the damaged zone is usually located on the outside of the aircraft, i.e. in a zone that can be seen. It is therefore preferable for the zone to resemble the composite structure before it was damaged, such that it is practically invisible. [0010]
    [0011] Moreover, the larger the damaged zone, the more crucial it becomes for the mechanical characteristics of the repaired zone to be as close as possible to those initially present in the composite structure before it was damaged. [0011]
    [0012] At present a number of devices exist that enable composite structure repairs to be carried out on-site when the damaged zone is relatively small. These devices include tooling for step machining or scarf machining a recess in the damaged zone of the composite structure and tooling used to polymerize a composite part placed in the recess under pressure. [0012]
    [0013] U.S. Pat. Nos. 5,207,541 and 5,271,145 concern machining tooling. Tooling of this kind is generally used to machine circular- or oval-shaped recesses. However, other shapes are also possible, as disclosed by U.S. Pat. Nos. 4,916,880 and 4,978,404. [0013]
    [0014] The composite part set into a recess as described is of a shape that matches that of the recess. The composite part generally comprises several layers or folds of long fibre fabric and non-polymerized resin. If appropriate, the fabric may previously be impregnated with resin or be initially dry. If the fabric is dry the resin can either be applied as a liquid on each layer of fabric or as solid film interposed between the various layers of dry fabric. [0014]
    [0015] Generally, the tooling used to polymerize and compact the composite part mainly comprises a heating cap that raises the resin to the temperature at which it polymerizes and a bladder fastened around the composite part such that it forms a leaktight seal on the surface of the composite structure. The bladder is connected to an external source of negative pressure such that pressure may be applied to the composite part. [0015]
    [0016] As shown in U.S. Pat. No. 4,554,036, the pressure applied to the composite part in this standard tooling remains very reduced, for example approximately 1.5 bar. [0016]
    [0017] This kind of tooling can therefore be used for “cosmetic” repairs, in other words for reconstituting the original appearance of the composite structure. However, this tooling affects the mechanical characteristics of the repaired zone and can reduce the characteristics by up to 30% compared to the structure in its original condition. Repair work using this technique is, therefore, limited to only small damaged zones. [0017]
    [0018] When larger zones are damaged the damaged composite structure is usually completely replaced and then sent to be repaired in a specialist workshop equipped with autoclaves that perform high pressure polymerization under appropriate conditions. [0018]
    [0019] In another known technique a composite part is molded in an autoclave into a shape that matches that of the recess previously machined in the damaged zone. The composite part is then polymerized and compacted in an autoclave before being bonded into place in the recess. This technique is discussed in U.S. Pat. No. 5,023,87. [0019]
    [0020] The major drawback of this technique is that the ability to withstand fatigue in the bonded zone is relatively limited, thus constituting a particular drawback for the aeronautics industry. It is for this reason that the fastening of the additional composite part is generally completed by other fastening means, such as bolts or rivets, that require access via the rear surface of the structure. The repair work can be seen from the outside and does not, therefore, meet the above-mentioned requirements concerning appearance. [0020]
    [0021] Furthermore, the polymerized composite part is manufactured in an autoclave, i.e. generally on a different site from that where the aircraft is located. This means that the aircraft will be immobilized for a long and costly period. [0021] DISCLOSURE OF THE INVENTION
    [0022] The invention mainly relates to tooling for on-site repairs of a composite structure with a damaged zone. The tooling restores the original appearance of the composite structure and guarantees mechanical characteristics that are very similar to those of the undamaged structure, thus allowing on-site repairs of relatively large damaged zones. [0022]
    [0023] According to the invention, these results are obtained using on-site repair tooling for composite structures with a concave recess previously machined in a surface of the structure for high-pressure polymerization of a non-polymerized composite part. The shape of the composite part matches that of the recess and the tooling is characterized in that it comprises: [0023]
    [0024] a bladder capable of being fastened around the said composite part such that it forms a leaktight seal on the surface of the composite structure; [0024]
    [0025] means for connecting the bladder to an external source of negative pressure; [0025]
    [0026] means for heating the composite part; and [0026]
    [0027] means for increasing the pressure applied to the composite part, said means being suitable for being interposed between said composite part and the leaktight bladder. [0027]
    [0028] In a preferred embodiment of the invention the means for increasing pressure comprise a stack of at least two plates with surfaces that gradually increase towards the bladder. [0028]
    [0029] For example, the surfaces of the plates the furthest away from the stack are in a ratio of at least 1:2. [0029]
    [0030] To take into account the fact that that the surface of an aircraft composite structure is not generally plane the plates are preferably made of plane leaf metal that can be stretched to take the shape of the composite structure. [0030]
    [0031] Guiding means are provided to hold the plates in a centered position in relation to each other when pressure is applied by creating a vacuum in the bladder. [0031]
    [0032] These guiding means may comprise at least one guiding pin that projects perpendicularly out of at least one of the plates through guiding holes provided in the other plates. [0032]
    [0033] The surface and the shape of the plate that is closest to the composite part are preferably more or less identical to those of said composite part on the surface of the structure. [0033]
    [0034] According to a characteristic known in the art, the heating means comprise a heating cap that is interposed, in this example, between the bladder and the means for increasing pressure. [0034]
    [0035] The invention also relates to a method for on-site repairs of a composite structure with a damaged zone, characterized in that it comprises the following stages: [0035]
    [0036] machining a concave recess in the damaged zone in a surface of the structure; [0036]
    [0037] setting a non-polymerized composite part in the recess, the shape of said composite part matching that of the recess; [0037]
    [0038] polymerizing the composite part in situ using a vacuum created inside a bladder that is sealed onto the said surface, around said composite part, after means for increasing the pressure applied to the composite part have been interposed between the part and the bladder; heating of the composite part. [0038] BRIEF DESCRIPTION OF THE DRAWINGS
    [0039] A preferred embodiment of the invention will now be described, as a non-limitative example, in reference to the attached drawing in which the sole figure is a schematic representation, shown in a partial cross section, of on-site repair tooling according to the invention positioned on a composite structure with a damaged zone. [0039] DETAILED DISCLOSURE OF A PREFERRED EMBODIMENT OF THE INVENTION
    [0040] In the sole figure, reference number [0040] 10 is a general reference to a composite structure, such as an aircraft structure. The outer surface 12 of the composite structure 10 has a concave recess 14 that has been previously machined using known tooling not included in the invention.
    [0041] Recess [0041] 14 has been machined in a damaged zone of composite structure 10 which is to be repaired on site. The recess 14 has an overall circular shape when seen from above but it may, however, have a different shape, such as an oval, depending on the shape of the damaged zone.
    [0042] In addition, recess [0042] 14 is either step machined, as seen in the figure, or scarf machined using known techniques. Recess 14 is preferably scarf machined as this enables mechanical characteristics to be obtained in the damaged zone that are greater than those obtained when the recess is step machined.
    [0043] Once recess [0043] 14 has been machined a composite part 16 is set into the recess, the shape of said composite part matching that of the recess. It should be pointed out that “composite part” does not refer to a pre-polymerized part but to a non-polymerized assembly comprising long fibers, for example carbon fibers, and an adhesive, such as an epoxy resin to which a softener has been added. Consequently, composite part 16 may be made directly on-site without it being necessary to use an autoclave.
    [0044] In practice, and using known techniques, composite part [0044] 16 can be made either by cutting and superimposing layers of fabric that have been pre-impregnated with resin or by cutting and superimposing layers of dry fabric, being careful to impregnate them layer by layer as they are set in place. Another alternative is to cut and superimpose alternate layers of dry fabric and solid films of adhesive. This technique is preferred as it is easy to perform and obtains high quality results.
    [0045] According to the invention, composite part [0045] 16 is set into recess 14 and polymerized under pressure on the site where composite structure 10 is located. Said composite part is polymerized using repair tooling given the overall number 18.
    [0046] Repair tooling [0046] 18 mainly comprises a leaktight bladder 20, means 22 for heating composite part 16 and means 24 for increasing the pressure applied to said composite part.
    [0047] The leaktight bladder [0047] 20 is slightly larger than composite part 16, such that the peripheral edge may be pressed to form a leaktight seal against outer surface 12 of composite structure 10 around composite part 16 using a leaktight putty 26.
    [0048] The leaktight bladder [0048] 20 is equipped with a valve 28 that constitutes means for connecting the bladder to an external source of negative pressure (not shown). No further description of this arrangement will be given as it is standard technology.
    [0049] In this example, means [0049] 22 for heating composite part 16 comprise a heating cap that is connected to an external power supply (not shown) by electrical conductors (not shown). The heating cap that constitutes the means 22 for heating composite part 16 is interposed between bladder 20 and pressure increasing means 24.
    [0050] Pressure increasing means [0050] 24 constitute the essential characteristic of the invention. They are interposed between bladder 20 and composite part 16, such that the load applied to said composite part is increased when the space 30, formed by composite structure 10 and leaktight bladder 20, is put under negative pressure.
    [0051] In the preferred embodiment of the invention shown in the sole figure, pressure increasing means [0051] 24 comprise a stack of plates 32 whose surfaces gradually increase towards bladder 20. More precisely, the surface and the shape of the plate 32 that is closest to composite part 16 are more or less identical to those of said composite part 16 on the outer surface 12 of composite structure 10. In practice, the plate 32 that is closest to part 16 generally has a slightly larger surface area than that of the surface 12 of part 16.
    [0052] Moreover, the plate [0052] 32 that is closest to composite part 16 is centered on said composite part such that it totally covers the part.
    [0053] In addition, the plate [0053] 32 that is closest to bladder 20 has a surface that is, for example, at least twice that of the plate 32 that is closest to composite part 16. Consequently, when a vacuum is created in space 30 the pressure applied to composite part 16 by the plate 32 that is closest to said part is more or less twice the pressure applied by leaktight bladder 20 to the plate 32 that is furthest away from the composite part.
    [0054] Given that the degree of negative pressure created in space [0054] 30 is generally between 0.7 and 0.9 bar for the primary vacuum sources that are most often used in this application, the addition of pressure increasing means 24 enable at least twice the pressure to be applied compared to that applied with standard tooling. Thus, and purely as an example, pressure measuring approximately 3 bar can be applied to composite part 16 with a 0.9 bar vacuum in space 30 and a 3.75 surface ratio.
    [0055] It should be noted that the pressure increase factor depends on the ratio of surfaces between the most distant plates of the stack of plates [0055] 32, such that the ratio may vary by a factor of three and still remain within the scope of the present invention.
    [0056] The number of plates [0056] 32 constituting the stack can be equal to two and these plates can be rigid when outer surface 12 of composite structure 10 is plane or has a continuous curve of a known value.
    [0057] In the example of tooling intended to repair the composite structures of aircraft the outer surface of these structures is not generally plane and may have varying degrees of curvature. In this event pressure increasing means [0057] 24 are preferably used in which the plates 32 are constituted by a relatively large number of leaf metal planes that can be stretched to take the shape of surface 12 of the composite structure to be repaired.
    [0058] Relative guiding means are preferably provided to prevent plates [0058] 32 sliding against each other when space 30 is put under negative pressure.
    [0059] In the embodiment shown in the sole figure these relative guiding means comprise a guiding pin [0059] 34 that projects perpendicularly out of at least one of the plates 32 through guiding holes 36 provided in the other plates 32. In the example shown, guiding pin 34 is fastened to the center of the plate 32 that is closest to bladder 20 and guiding holes 36 are provided in the center of the other plates 32. It should be noted that other arrangements are also possible, such as several guiding pins 34 being fastened to the same plate 32 or to different plates.
    [0060] The length of the guiding pin(s) [0060] 34 is designed to enable the remaining plates 32 to be guided without, however, the pin(s) projecting beyond the stack of plates. Therefore, the length of the pin(s) is slightly shorter than the accumulated thickness of plates 32 that do not bear any pins.
    [0061] In the sole figure plates [0061] 32 are shown to have spaces between them. This is intended to make the drawing clearer and it is for this reason that pin 34 does not pass through certain plates 32.
    [0062] It should be noted that the relative guiding means of plates [0062] 32 may have a different shape in other modified versions.
    [0063] A guiding liner [0063] 33, positioned around the stack of plates 32, prevents bladder 20 from becoming caught under the stack or between two plates when space 30 is put under negative pressure.
    [0064] In the above-mentioned example, where only two plates [0064] 32 are used, guiding liner 33 that is interposed between composite structure 10 and the bladder around the stack of plates can ensure the relative guiding of plates 32.
    [0065] The repair tooling [0065] 18 of the invention also comprises all the components that generally equip tooling of the prior art. These various components will now be briefly described in reference to the figure.
    [0066] A drainage fabric [0066] 38, the measurements of which are slightly greater than those of heating cap 22, is interposed between said heating cap and leaktight bladder 20. This drainage fabric 38 is used to evacuate air completely when space 30 is put under negative pressure.
    [0067] The remaining components of repair tooling [0067] 18 are positioned between composite part 16 and pressure increasing means 24. Starting from composite part 16, these various components comprise a de-molding fabric 40, a perforated plastic film 42, a drainage fabric 44, a leaktight film 46 and a Mossite film 48.
    [0068] The de-molding fabric [0068] 40 is used to separate tooling 18 from part 16 after polymerization.
    [0069] The perforated plastic film [0069] 42 operates in conjunction with drainage fabric 44 and leaktight film 46 to evacuate air and excess resin towards the periphery when space 30 is put under negative pressure.
    [0070] Finally, Mossite film [0070] 48, which is optional, ensures that the pressure applied to composite part 16 through pressure increasing means 24 is evenly spread.
    [0071] When a composite structure [0071] 10 with a damaged zone is repaired, first of all a concave recess 14 is machined in structure 10 to eliminate all the damaged zones.
    [0072] Should the damage extend through the entire thickness of composite structure [0072] 10, recess 14 must also cross through said thickness. A thin composite plug is then bonded to the base of the recess before the following operations are carried out.
    [0073] The non-polymerized composite part [0073] 16 is then set into recess 14, as described above.
    [0074] The repair tooling [0074] 18, described above, is then installed by connecting valve 28 to the external source of negative pressure and by connecting heating cap 22 to the external power supply.
    [0075] The adhesive contained in composite part [0075] 16 is polymerized under pressure when composite part 16 is heated by heating cap 22 and is put under pressure by space 30 coming under negative pressure. More precisely, pressure, for example of approximately 3 bar, is applied to composite part 16 due to the amplifying action of pressure increasing means 24.
    [0076] When the polymerization has finished the heating stops and space [0076] 30 is reconnected to the outside atmosphere. Repair tooling 18 can then be removed.
    [0077] Tests carried out using tooling of this kind show that the external appearance of surface [0077] 12 is unchanged in the repaired zone and that the mechanical characteristics in said zone are respectively 79% and 89% of those of the initial structure, depending on whether a step machined or a scarf machined recess 14 is used. The mechanical characteristics are therefore considerably greater than those obtained using techniques of the prior art and enable repair work to be carried out on a far larger scale than has previously been possible.
    权利要求:
    Claims (12)
    [1" id="US-20010008161-A1-CLM-00001] 1. Tooling (18) for on-site repairs of a composite structure (10) having a concave recess (14) previously machined in a surface (12) of structure (10), for high pressure polymerization of a non-polymerized composite part (16) with a shape that matches that of recess (14), the tooling being characterized in that it comprises the following:
    a bladder (20) capable of being fastened around said composite part (16) such that it forms a leaktight seal on surface (12) of composite structure (10);
    means (28) for connecting bladder (20) to an external source of negative pressure;
    means (22) for heating composite part (16); and
    means (24) for increasing the pressure applied to composite part (16), said means (24) being suitable for being interposed between said composite part and bladder (20).
    [2" id="US-20010008161-A1-CLM-00002] 2. Tooling of
    claim 1 wherein the means for increasing pressure comprise a stack of at least two plates (32) with surfaces that gradually increase towards the bladder (20).
    [3" id="US-20010008161-A1-CLM-00003] 3. Tooling of
    claim 2 wherein the surfaces of plates (32) the furthest away from the stack are in a ratio of at least 1:2.
    [4" id="US-20010008161-A1-CLM-00004] 4. Tooling of either of claims 2 and 3 wherein plates (32) are preferably leaf metal planes that can be stretched to take the shape of the composite structure (10).
    [5" id="US-20010008161-A1-CLM-00005] 5. Tooling of any of
    claims 2 to
    4 wherein means (34, 36) for guiding plates relative to one another (32) are provided.
    [6" id="US-20010008161-A1-CLM-00006] 6. Tooling of
    claim 5 wherein the guiding means comprise at least one guiding pin (34) that projects perpendicularly out of at least one of the plates (32) through guiding holes (36) provided in the other plates.
    [7" id="US-20010008161-A1-CLM-00007] 7. Tooling of any of
    claims 2 to
    6 wherein the plate (32) that is closest to composite part (16) has a surface and a shape that are more or less identical to those of said composite part at surface (12) of composite structure (10).
    [8" id="US-20010008161-A1-CLM-00008] 8. Tooling of any of the previous claims wherein the heating means comprise a heating cap (22) interposed between bladder (20) and means (24) for multiplying pressure.
    [9" id="US-20010008161-A1-CLM-00009] 9. Method for on-site repair of a composite structure with a damaged zone characterized in that it comprises the following stages:
    machining a concave recess (14) in the damaged zone in a surface of the structure;
    setting a non-polymerized composite part (16) in the recess (14), the shape of said composite part matching that of the recess;
    polymerizing the said composite part (16) in situ using a vacuum created inside a bladder (20) that is connected around composite part (16) to form a leaktight seal with said surface of the composite structure after means (24) for increasing the pressure applied to the composite part (16) have been interposed between the part and the bladder (20). The composite part (16) is then heated.
    [10" id="US-20010008161-A1-CLM-00010] 10. Method of
    claim 9 wherein means (24) for increasing pressure are interposed, said means comprising a stack of at least two plates (32) with surfaces that gradually increase towards bladder (20).
    [11" id="US-20010008161-A1-CLM-00011] 11. Method of
    claim 10 wherein plates (32) are guided relative to each other during polymerization.
    [12" id="US-20010008161-A1-CLM-00012] 12. Method of any of
    claims 9 to
    11 wherein composite part (16) is heated by a heating cap (22) being interposed between bladder (20) and means (24) for multiplying pressure.
    类似技术:
    公开号 | 公开日 | 专利标题
    US6206067B1|2001-03-27|Equipment for on-site repair of a composite structure with a damaged zone and corresponding method
    US7875141B2|2011-01-25|Repair of a composite sandwich structure having a perforated skin
    US8592024B2|2013-11-26|Repair of perforated acoustic skins by reverse pin transfer molding
    EP1092529B1|2004-04-07|Selectively flexible caul
    US4683018A|1987-07-28|Composite material manufacture by shaping individual sheets followed by consolidating the sheets
    US7883760B2|2011-02-08|Repair of composite sandwich structures with uneven bond surfaces
    CA2245088C|2006-10-24|Method for forming inner mold line tooling without a part model
    US20030098520A1|2003-05-29|High pressure co-cure of lightweight core composite article utilizing a core having a plurality of protruding pins
    US5618606A|1997-04-08|Process for bonding staged composites with a cobonded staged adhesive and article
    US6620369B1|2003-09-16|Net molding of resin composite parts
    US11135824B2|2021-10-05|Multi-stage debulk and compaction of thick composite repair laminates
    GB2238977A|1991-06-19|A composite material
    EP2212100B1|2016-02-17|In-situ, multi-stage debulk, compaction, and single stage curing of thick composite repair laminates
    JP2020515435A|2020-05-28|Tools for manufacturing complex components
    US9302436B2|2016-04-05|In-situ, multi-stage debulk, compaction, and single stage curing of thick composite repair laminates
    同族专利:
    公开号 | 公开日
    FR2763882A1|1998-12-04|
    AU8023398A|1998-12-30|
    AU730746B2|2001-03-15|
    US6206067B1|2001-03-27|
    JP4208262B2|2009-01-14|
    DE69803928D1|2002-03-28|
    CA2291658A1|1998-12-03|
    FR2763882B1|1999-08-20|
    EP0984853A1|2000-03-15|
    ES2173589T3|2002-10-16|
    CN1071626C|2001-09-26|
    EP0984853B1|2002-02-20|
    WO1998053982A1|1998-12-03|
    DE69803928T2|2002-08-22|
    JP2002511030A|2002-04-09|
    CN1258247A|2000-06-28|
    CA2291658C|2008-05-20|
    US6468372B2|2002-10-22|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    WO2003055747A1|2001-12-22|2003-07-10|Bae Systems Plc|Method of forming and indirect testing of a bond on or in an aircraft component|
    US6656299B1|2001-12-19|2003-12-02|Lockheed Martin Corporation|Method and apparatus for structural repair|
    US20060059828A1|2004-07-29|2006-03-23|Stevenson James F|Repair method for noise suppression structure|
    WO2008104265A1|2007-02-26|2008-09-04|Eads Deutschland Gmbh|Method for repairing a component|
    WO2009080038A1|2007-12-21|2009-07-02|Vestas Wind Systems A/S|A method of repairing a fibre composite solid member|
    US20090293711A1|2008-06-03|2009-12-03|Triton Systems, Inc.|Armor repair kit and methods related thereto|
    US20100136232A1|2007-03-30|2010-06-03|Laurence John Walker| curing of polymer composites|
    US20100227117A1|2009-03-09|2010-09-09|The Boeing Company|Tapered patch for predictable bonded rework of composite structures|
    US20100227105A1|2009-03-09|2010-09-09|The Boeing Company|Predictable bonded rework of composite structures|
    US20100227106A1|2009-03-09|2010-09-09|The Boeing Company|Predictable bonded rework of composite structures using tailored patches|
    US20100233424A1|2009-03-10|2010-09-16|The Boeing Company|Composite structures employing quasi-isotropic laminates|
    GB2498866A|2012-01-26|2013-07-31|Gen Electric|Method for repairing a laminated article having a damaged area|
    US8524356B1|2009-03-09|2013-09-03|The Boeing Company|Bonded patch having multiple zones of fracture toughness|
    US8617694B1|2009-03-09|2013-12-31|The Boeing Company|Discretely tailored multi-zone bondline for fail-safe structural repair|
    US20140141190A1|2012-11-20|2014-05-22|Mitsubishi Aircraft Corporation|Method of repairing composite material and composite material|
    GB2531600A|2014-10-24|2016-04-27|Short Brothers Plc|Apparatus and methods for manufacturing and repairing fibre-reinforced composite materials|
    US9492975B2|2009-03-09|2016-11-15|The Boeing Company|Structural bonded patch with tapered adhesive design|
    EP3566844A4|2017-03-10|2020-01-08|Mitsubishi Heavy Industries, Ltd.|Composite material pressurizing device and composite material forming method|
    EP3815888A1|2019-10-30|2021-05-05|The Boeing Company|Vacuum bag-less composite repair systems and repairs|US4554036A|1984-05-07|1985-11-19|Newsom Cosby M|Portable vacuum device and method of using same|
    EP0180865A3|1984-10-29|1986-12-30|The Boeing Company|Blind side panel repair patch|
    US4916880A|1986-07-21|1990-04-17|The Boeing Company|Apparatus for repairing a hole in a structural wall of composite material|
    US4978404A|1986-07-21|1990-12-18|The Boeing Company|Method for repairing a hole in a structural wall of composite material|
    US4808253A|1987-11-06|1989-02-28|Grumman Aerospace Corporation|Method and apparatus for performing a repair on a contoured section of a composite structure|
    US5207541A|1988-12-13|1993-05-04|The Boeing Company|Scarfing apparatus|
    US5023987A|1989-08-28|1991-06-18|The Boeing Company|Strato streak flush patch|
    US5442156A|1991-04-09|1995-08-15|The Boeing Company|Heating apparatus for composite structure repair|
    US5271145A|1991-06-03|1993-12-21|The Boeing Company|Rope saw cutting apparatus and method for scarfing composites|
    FR2705914B1|1993-06-02|1995-08-18|Aerospatiale|Pressure application device, possibly heating, on any surface of any shape, in particular for repairing damaged structures.|
    BR9510294A|1995-01-27|1997-11-11|Sikorsky Aircraft Corp|Method for producing composite core articles|
    FR2744054B1|1996-01-29|1998-04-30|Aerospatiale|METHOD FOR PRODUCING PARTS OF HIGH DIMENSIONAL PRECISION COMPOSITE MATERIAL USING IONIZATION POLYMERIZATION AND PARTS OBTAINED THEREBY|
    US5738741A|1996-03-26|1998-04-14|The University Of Dayton|Pre-fabricated vacuum bag and vacuum bag process to externally reinforce structural members with advanced composites|
    JPH10315257A|1997-05-22|1998-12-02|Meiki Co Ltd|Apparatus and method for vacuum laminating|GB9901860D0|1999-01-29|1999-03-17|Davey Terence J|Treating products moulded from fibre reinforced plastics|
    EP1162058A1|1999-02-16|2001-12-12|Toray Industries, Inc.|Frp structure body and production method therefor|
    CN1224495C|2000-03-03|2005-10-26|快速科技有限公司|Production, forming, bonding, joining and repair systems for composite and metal components|
    KR100910653B1|2001-01-25|2009-08-05|퀵스텝 테크놀로지즈 피티와이 리미티드|Composite and metal component production, forming and bonding system|
    US7100656B2|2002-04-15|2006-09-05|Northrop Grumman Corporation|System and method for curing composite material|
    DE102004062064B4|2004-12-23|2008-07-17|Röder Präzision GmbH|Process for producing and repairing composite fiber composite metal components and apparatus for carrying out the method|
    EP1683627A1|2005-01-25|2006-07-26|Saab Ab|Method and apparatus for repairing a composite article|
    US8790485B2|2005-02-25|2014-07-29|Bell Helicopter Textron Inc.|Single vacuum debulk composite panel repair|
    JP4874326B2|2005-03-22|2012-02-15|クイックステップ、テクノロジーズ、プロプライエタリ、リミテッド|Manufacture of composite tube|
    US7686905B2|2005-09-06|2010-03-30|The Boeing Company|Copper grid repair technique for lightning strike protection|
    FR2891192B1|2005-09-29|2007-10-26|Airbus France Sas|PROCESS FOR REPAIRING EXTERNAL SKIN OF COMPOSITE STRUCTURE|
    US7419556B2|2005-10-11|2008-09-02|United Technologies Corporation|Method of repair for inlet caps of turbine engines|
    GB0619667D0|2006-10-05|2006-11-15|Short Brothers Ltd|Apparatus and method for close tolerance forming of areas in fibre reinforced resin composite components|
    US8936695B2|2007-07-28|2015-01-20|The Boeing Company|Method for forming and applying composite layups having complex geometries|
    US7849729B2|2006-12-22|2010-12-14|The Boeing Company|Leak detection in vacuum bags|
    US9770871B2|2007-05-22|2017-09-26|The Boeing Company|Method and apparatus for layup placement|
    US8568551B2|2007-05-22|2013-10-29|The Boeing Company|Pre-patterned layup kit and method of manufacture|
    GB2452298B|2007-08-30|2010-01-13|Gkn Aerospace Services Ltd|Composite structure|
    US8916010B2|2007-12-07|2014-12-23|The Boeing Company|Composite manufacturing method|
    US8752293B2|2007-12-07|2014-06-17|The Boeing Company|Method of fabricating structures using composite modules and structures made thereby|
    US8371009B2|2007-12-12|2013-02-12|General Electric Company|Methods for repairing composite containment casings|
    US8403624B2|2007-12-12|2013-03-26|General Electric Company|Composite containment casings having an integral fragment catcher|
    US8333864B2|2008-09-30|2012-12-18|The Boeing Company|Compaction of prepreg plies on composite laminate structures|
    US8202056B2|2008-10-16|2012-06-19|Rolls-Royce Corporation|Morphable composite structure|
    US20100098896A1|2008-10-16|2010-04-22|Edward Claude Rice|Patch|
    US8778487B2|2008-10-16|2014-07-15|Rolls-Royce Corporation|Tape|
    DE102009002501A1|2009-04-20|2010-10-28|Wobben, Aloys|Rotor blade element and manufacturing process|
    US8524020B2|2009-06-16|2013-09-03|The Boeing Company|Method of restoring a composite airframe|
    US7927077B2|2009-07-09|2011-04-19|General Electric Company|Wind blade spar cap laminate repair|
    CH701994A1|2009-10-08|2011-04-15|Joerg-Ulrich Zowalla|Wet lamination and foil tube for wet lamination.|
    US8444127B2|2009-12-14|2013-05-21|The Boeing Company|High temperature composite patch tool|
    JP5598842B2|2010-02-24|2014-10-01|三菱航空機株式会社|Repair method for members made of fiber reinforced resin|
    US8707766B2|2010-04-21|2014-04-29|The Boeing Company|Leak detection in vacuum bags|
    JP6218323B2|2010-08-13|2017-10-25|ヘクセル コーポレイション|Machinable composite materials|
    US8986479B2|2010-09-30|2015-03-24|The Boeing Company|Systems and methods for on-aircraft composite repair using double vacuum debulking|
    US10155366B2|2011-01-11|2018-12-18|Textron Innovations Inc.|Single stage debulk and cure of a prepreg material|
    DE102011004775B4|2011-02-25|2012-10-25|Airbus Operations Gmbh|Method of making a connection, connection and aircraft or spacecraft|
    US8945321B2|2011-05-30|2015-02-03|The Boeing Company|Method and apparatus for reworking structures using resin infusion of fiber preforms|
    DE102011086453A1|2011-11-16|2013-05-16|Wobben Properties Gmbh|Heating device for repair or production of components of a wind turbine and parts thereof and wind turbine|
    DE102012207468B4|2012-05-04|2021-07-22|Bayerische Motoren Werke Aktiengesellschaft|Method for repairing a component made of a fiber composite material with a thermoplastic matrix and device for carrying out the method|
    US10265915B2|2014-01-17|2019-04-23|Sikorsky Aircraft Corporation|Composite bonded repair method|
    US9434139B2|2014-05-01|2016-09-06|The Boeing Company|Structural integrated repair method and system tool device for repairing delaminated composite structures|
    US10093068B2|2014-07-08|2018-10-09|Polycorp Ltd.|Method and system for providing a rubber lining|
    FR3029820B1|2014-12-16|2016-12-30|Snecma|PROCESS FOR REPAIRING A BLOWER HOUSING|
    US10222353B2|2015-03-18|2019-03-05|The Boeing Company|Method and assembly for inspecting a partially cured repair patch prior to installation|
    CN105196575B|2015-08-20|2017-08-25|国电联合动力技术有限公司|A kind of method for maintaining for wind generator set blade|
    FR3042779B1|2015-10-27|2018-07-13|Airbus Operations|METHOD FOR REPAIRING AN AIRCRAFT STRUCTURE FROM DEFORMABLE PLATES|
    US9994009B2|2015-11-09|2018-06-12|The Boeing Company|Inverted double dome double vacuum debulk tool|
    US10695993B2|2016-01-15|2020-06-30|GM Global Technology Operations LLC|In-situ polymerization of polyamides for composite part repair|
    US10544676B2|2016-02-03|2020-01-28|General Electric Company|Situ gas turbine prevention of crack growth progression|
    US20170218762A1|2016-02-03|2017-08-03|General Electric Company|Situ Gas Turbine Prevention of Crack Growth Progression|
    US10094221B2|2016-02-03|2018-10-09|General Electric Company|In situ gas turbine prevention of crack growth progression|
    US10247002B2|2016-02-03|2019-04-02|General Electric Company|In situ gas turbine prevention of crack growth progression|
    US10443385B2|2016-02-03|2019-10-15|General Electric Company|In situ gas turbine prevention of crack growth progression via laser welding|
    US10024163B2|2016-03-01|2018-07-17|General Electric Company|In situ tip repair of an airfoil tip in a gas turbine engine via frictional welding|
    US9970326B2|2016-03-01|2018-05-15|General Electric Company|Gas turbine in situ inflatable bladders for on-wing repair|
    US10589477B2|2016-05-02|2020-03-17|GM Global Technology Operations LLC|Cosmetic repair of a thermoplastic carbon fiber composite|
    US10753277B2|2016-07-27|2020-08-25|Raytheon Technologies Corporation|Inlet cap of an engine|
    FR3056141A1|2016-09-20|2018-03-23|Airbus Operations |PROCESS FOR MANUFACTURING A PIECE OF COMPOSITE MATERIAL|
    KR101863276B1|2017-01-12|2018-05-31|한국과학기술연구원|Method and Apparatus For Repairing Compsoite Material Using Solvation Process|
    US10611104B2|2017-06-15|2020-04-07|GM Global Technology Operations LLC|Heating elements for repair of molding defects for carbon fiber thermoplastic composites|
    FR3067643B1|2017-06-20|2020-08-14|Airbus Operations |INFLATABLE FRAME REPAIR DEVICE AND PROCEDURE|
    CN107756830A|2017-09-28|2018-03-06|国营芜湖机械厂|One kind is applied to outfield radome damage technique for rapidly repairing method|
    US10935460B2|2018-07-17|2021-03-02|General Electric Company|Ultrasonic tank for a turbomachine|
    CN109676974A|2018-12-12|2019-04-26|核工业第八研究所|A method of repairing carbon fiber cross bar|
    法律状态:
    2001-06-07| AS| Assignment|Owner name: AEROSPATIALE SOCIETE NATIONALE INDUSTRIELLE, FRANC Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:KOCIEMBA, FABIENNE;SORO, DANIEL;REEL/FRAME:011878/0694 Effective date: 19991122 |
    2002-10-03| STCF| Information on status: patent grant|Free format text: PATENTED CASE |
    2006-04-07| FPAY| Fee payment|Year of fee payment: 4 |
    2010-04-15| FPAY| Fee payment|Year of fee payment: 8 |
    2014-04-17| FPAY| Fee payment|Year of fee payment: 12 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR9706602A|FR2763882B1|1997-05-29|1997-05-29|ON-SITE REPAIR TOOLS OF A COMPOSITE STRUCTURE HAVING A DAMAGED AREA AND CORRESPONDING METHOD|
    FR9706602||1997-05-29||
    US09/423,208|US6206067B1|1997-05-29|1998-05-28|Equipment for on-site repair of a composite structure with a damaged zone and corresponding method|
    US09/780,403|US6468372B2|1997-05-29|2001-02-12|On-site tooling and method for repairing a damaged zone in a composite structure|US09/780,403| US6468372B2|1997-05-29|2001-02-12|On-site tooling and method for repairing a damaged zone in a composite structure|
    [返回顶部]