Patent Publication Number: US-6666942-B2

Title: Method for manufacturing composite parts

Description:
This application is a continuation of application Ser. No. 09/591,352, filed Jun. 9, 2000, now U.S. Pat. No. 6,406,580, entitled “Method for Manufacturing Composite Parts”. 
    
    
     BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     This invention relates in general to a method of manufacturing composites and in particular to a method for manufacturing composite panels and stiffeners by first preparing a circumferential array of material and then forming and curing the material to the desired shape. 
     2. Description of Related Art 
     Composite structures and parts are necessary parts for industries requiring high strength, lightweight materials. A good example of this is in the aerospace industry, where aircraft and other airborne vehicles require high strength components that weigh as little as possible. 
     Many approaches have been previously developed for forming multiple layers of composite material into a desired shape or shapes. The most common, particularly in the aircraft industry, involves placing individual layers of material onto a form having a desired shape, and then curing the layers. Curing the material through application of heat and pressure fully compacts or debulks the composite material. The cured composite material then has the desired shape and strength. Forming parts in this way does not involve significant reshaping of the composite material during curing and may be very time consuming. 
     Disadvantages inherent in the aforementioned process include the very tedious and time consuming operation of laying individual layers of composite material directly onto a tool to obtain a final non-flat desired shape. The very labor intensive process of placing the layers of material onto a form may require many highly-skilled man-hours for each part, and is, therefore, very expensive. Additionally, the aforementioned process may require stopping after placement of every few layers of material and providing some form of mechanical compaction to the material. This may be necessary to achieve final full compaction of the layers. Failure to achieve full compaction of the material layers prior to curing may result in wrinkles and other anomalies in the final structure, since as individual layers compact, the local path-lengths of the fibers in the layers change. Wrinkles and other anomalies in the cured structure are aesthetically and structurally undesirable. 
     Previously developed methods for forming composite parts also fail to assure uniformity between parts. In many prior art methods, each part is separately made. Each part is formed by the process of placing individual layers onto a form and then curing the layers while on the form. The cured part is removed from the form, allowing the next part to be made by the same process. By this method a number of parts can be formed. Unfortunately, variations in compaction, in resin bleeding from the part, and in fiber “washing” or dislocations from resin bleeding, tend to occur because compaction is occurring three dimensionally, and because of the low viscosity of the resin. These factors may yield parts that lack uniformity. Previously developed methods for building composite parts are, therefore, not compatible with low-cost, high-volume manufacturing methodologies. 
     Composite parts fabricated by previously developed methods often require machining after curing, e.g. routing, grinding, etc., in order to meet final dimension requirements. This machining adds additional time and expense to the process of fabricating the part and can result in damage to the part by delamination of the cured layers. 
     Yet another disadvantage of the previously developed methods for fabricating composite parts is their incompatibility with in-process control (IPC), statistical process control (SPC), and total quality (TQ) methodologies. IPC, SPC, and TQ require repeatable, measurable results to obtain full effectiveness. The custom approach of the prior art to fabricating composite parts is not amenable to obtaining the benefits of IPC, SPC, and TQ, i.e., high quality, high yield, and low cost. 
     Previously developed methods for forming composite parts often do not provide acceptable results when forming complex parts from two or more sub-parts or pre-forms by “co-curing”. In co-curing, two or more sub-parts are made into a single part by placing the sub-parts in the desired orientation and curing the combination. Since the prior art requires the individual layers of a sub-part to be laid-up in their final shape on a form joining two or more individual sub-parts to make a part, e.g., two channels and two plates to form an I-beam, is very difficult. If a foreign material, e.g., backing paper or tape, is accidentally trapped between the layers during layup of a part, there is little likelihood that it will be detected. As a result, high labor costs may be invested in a complex, co-cured part that must be scrapped due to the inclusion. 
     Prior methods for fabricating composite parts often require that the individual composite layers be stored in a freezer prior to layup. This adds additional handling and equipment costs in fabricating a composite part. 
     Therefore a need has arisen for an improved method and system for fabricating parts from composite materials. 
     A need further exists for an improved method and system for reducing the time necessary for fabricating composite parts. 
     A further need exists for a low-cost method and system for fabricating parts from composite materials. 
     Yet another need exists for a method and system for fabricating multiple uniform parts from composite materials that do not require significant amounts of machining after curing. 
     Another need exists for a method for fabricating composite parts compatible with IPC, SPC, and TQ. 
     Yet another need exists for a method and system for fabricating composite parts that eliminate the need for special handling and storage of composite layers. 
     BRIEF SUMMARY OF THE INVENTION 
     The present invention provides a method for manufacturing a composite part having a specified shape, thickness, and density. The composite part is formed in an array of similar parts located around a mandrel having an outer surface and a longitudinal axis. Several tooling members are disposed around the outer surface of the mandrel, the tooling members serving as the molds around which the composite parts will be formed. A plurality of filler members are disposed on the outer surface of the mandrel between the tooling members. An outer surface of each tooling member and an outer surface of each filler member combine to form a generally smooth and rounded application surface that surrounds the mandrel. 
     A plurality of composite layers is positioned on the application surface using an automated positioning technique such as fiber placement. A part formation aid is placed on the composite layers above each tooling member, and a cut is made in the composite layers parallel to the longitudinal axis of the mandrel and between each tooling member. The filler members are then removed from between the tooling members. 
     Finally, the mandrel and composite layers are placed in a vacuum bag, which is then placed in an autoclave. The vacuum bag is an elastomeric material in the form of a tube that is slid over the mandrel. The vacuum bag tube can be pulled out of the way during fiber placement and slid over the mandrel at the time of forming. As the bag is evacuated, each part formation aid deforms toward the mandrel, thereby forming the composite layers around the tooling members. The composite part resulting from the evacuation of the bag is then cured for a specified amount of time to insure that the composite part will maintain its specified shape, thickness, and density. 
     Alternatively, the method according to the present invention is used to fabricate composite parts having more than one component. One application of this method is to produce a flanged panel having a plurality of hat-shaped stiffeners integrally disposed on the panel. 
     The flanged panel is produced using a rectangular mandrel having an outer surface. A filler member having an outer surface is placed on one end of the mandrel, the outer surface of the filler member and the outer surface of the mandrel forming a generally smooth and rounded application surface. 
     A plurality of composite layers is positioned on the application surface using an automated positioning method such as fiber placement. After placement of the layers, the filler member is removed, and the composite layers are subjected to the vacuum bagging technique previously mentioned. This process forms a flanged composite panel. The panel is then partially cured. 
     After partial curing, a plurality of tooling members are disposed on an outer surface of the recently formed panel. Filler members are placed on the outer surface of the panel between the tooling members to form a generally smooth second application surface. Composite layers are applied to the second application surface using fiber placement. The filler members are then removed and the composite layers are cut between the tooling members. 
     The flanged panel, the tooling members, and the newly applied composite layers are placed in a vacuum bag. As the bag is evacuated, the composite layers form into hat-shaped stiffeners around the tooling members. The flanged panels and the stiffeners are finally co-cured or co-bonded to form an integral composite part. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIG. 1 is a partially cut-away perspective view showing a plurality of composite layers positioned around a mandrel, tooling members, and filler members according to the present invention. 
     FIG. 2 is a perspective view showing the mandrel, tooling members, and composite layers of FIG. 1 with a part formation aid located above each tooling member. 
     FIG. 3 is a partially cut-away perspective view of the mandrel and tooling members of FIG. 1 shown installed in a vacuum bag, the vacuum bag being held to the inside of a hard outer tube. 
     FIG. 4 is a sectional view of one of the tooling members, shown during evacuation. 
     FIG. 5 is a perspective of view of a composite part formed using the method of the present invention with the mandrel and tooling members of FIG.  1 . 
     FIG. 6 is a perspective view of a mandrel having a plurality of hat-shaped tooling members for use with the method of the present invention. 
     FIG. 7 is a front view of a mandrel having a plurality of Z-shaped tooling members for use with the method of the present invention. 
     FIG. 8 is a perspective view of a composite part produced by using an alternate embodiment of the method according to the present invention. 
     FIG. 9 is a perspective view of a plurality of composite fibers positioned on a rectangular mandrel and a filling member according to the alternate embodiment of the present invention. 
     FIG. 10 is a perspective view of a plurality of composite fibers positioned on a plurality of tooling members installed on a composite part produced from the composite fibers of FIG.  9 . 
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
     Referring to FIGS. 1,  2 , and  3  in the drawings, a method of manufacturing composite parts according to the present invention is illustrated. A mandrel  11  having a longitudinal axis  13  and an outer surface  15  provides a support base for producing the composite parts according to the present invention. In this embodiment, mandrel  11  is square in cross section, having four flat sides on its outer surface  15 . A plurality of tooling members  17  are disposed on outer surface  15  of mandrel  11 . Tooling members  17  generally have a longitudinal axis that is positioned parallel to longitudinal axis  13  when tooling members  17  are attached to mandrel  11 . Although tooling members  17  are generally made of rubber and are mechanically fastened to mandrel  11 , it is conceivable that tooling members  17  could be an integral part of mandrel  11 . In cross-section, each tooling member is a trapezoid in this embodiment, having sidewalls  18  that converge toward each other and an outer surface  25  that is a portion of a cylinder. 
     After installation of tooling members  17 , filler members  21  are placed between tooling members  17  and against outer surface  15  of mandrel  11 . Depending on the spacing between tooling members  17  and the cross-sectional shape of mandrel  11 , more than one filler member  21  may be required in the space between tooling members  17 . In FIG. 1, two filler members  21  are placed in each space between tooling members  17 . Each filler member  21  has a sidewall  22  that diverges from outer surface  15  and abuts one of the sidewalls  18 . Each filler member  21  has an outer surface  27  that is a portion of a cylinder. 
     Filler members  21  are used in conjunction with tooling members  17  to provide a smooth application surface  23 . Typically, filler members  21  are made of tooling foam, which is easily moldable or shapable into application surface  23 . Tooling foam was used because of cost and because this was a one time part. Materials for the filler members  21  are chosen for production with durability as a requirement as well as cost or because of ease of manufacture. Application surface  23  is formed by outer surface  25  of each tooling member and outer surface  27  of each filler member  21 , which are flush. Usually, application surface  23  will be smooth and rounded, and preferably, application surface  23  will have a cylindrical shape as shown in FIG.  1 . 
     After preparing application surface  23 , a plurality of composite layers  37  (not shown in FIG. 3) are positioned on the application surface  23 . Several different methods could be used to position composite layers  37  on application surface  23 . The most basic method would be to hand lay the composite material. However, a major advantage of the method of the present invention is its adaptability towards automation. Although any automated technique for applying composites could be used, the most common automated techniques include filament winding, tape winding, and fiber placement. 
     Filament winding is a process in which a continuous filament, usually constructed of a reinforced fiber impregnated by a matrix material, is wound under tension around a rotating core. Referring to FIG. 1, mandrel  11  is rotated as the filament is applied under tension to application surface  23 . Several rotations of mandrel  11  would be necessary to completely apply the filament, thus constructing the plurality of composite layers  37 . The matrix material is either preimpregnated in the reinforced fiber or is applied to the fiber as the filament is being positioned on application surface  23 . The matrix material serves to adhere the filament to previous layers as it is being applied, and also serves as a curing agent after formation of a final composite part. 
     A similar automated application technique is tape winding. A tow is a bundle of more than a thousand filaments. In a tape winding process, a unidirectional preimpregnated tape is used that consists of several tows that have been formed and spread with resin. The tape is cut to a particular width, common widths being 0.5 inch and 1.0 inch. The tape is positioned on application surface  23  as mandrel  11  is rotated about longitudinal axis  13 . The tape is applied under tension, and no contact is made between the head of the winding machine and application surface  23 . 
     The preferred application technique for the present invention is fiber placement. Unlike filament winding and tape winding, fiber placement does not use rotation of mandrel  11  to pull the composite material onto application surface  23 . Instead, a fiber placement machine having a head makes contact with application surface  23  and is used to apply either a tow or a small width tape to the application surface  23 . As material is placed on application surface  23 , more tows or tapes are pushed to the head of the fiber placement machine. The head of the machine applies a force to the tow or tape as it is applied. The primary advantage of using fiber placement is that the directional placement of the material is not limited by the rotation of mandrel  11 . Instead, material can be placed in any direction. Additionally, scrap is reduced when compared with filament winding and tape winding methods. Finally, fiber placement allows material to be easily added or removed before the material is formed into a final composite part. 
     After applying composite layers  37  to application surface  23 , a part formation aid  39  optionally may be installed on an outer surface of composite layers  37  above each tooling member  17 . Part formation aids  39 , also referred to as “bat-wings,” aid in the formation of a final composite part and are shaped based on the final part to be formed. In this embodiment, part formation aid  39  is generally U-shaped, having a base  40  and two outward extending sidewalls  42 . The part formation aid  39  shown in FIGS. 2 and 3 is used to assist in the formation of a hat-shaped stiffener. Part formation aids  39  are sufficiently flexible to allow the sidewalls  42  to flex inward into an inverted shape to that shown in FIG.  2 . 
     After installation of part formation aids  39 , composite layers  37  are cut parallel to longitudinal axis  13  and between each tooling member  17  along slit  44  (FIG.  2 ). The cutting of the composite material  37  allows individual parts to be formed around each tooling member  17 . Slit  44  results in a plurality of separate composite layer segments, each having side portions extending in opposite directions from each tooling member  17 . For production it is recommended that the cutting operation be automated. The automation can be accomplished by attaching a cutting blade to a multiple axis end effector of the fiber placement machine or winding machine. Alternatively, fiber placement machines have an ability to stop and start material placement on the fly. Therefore, the machine could be programmed to leave a gap in the material between the segments. 
     Filler members  21  (not shown in FIG. 2) are then removed from mandrel  11 , thus forming recesses between the tooling members  17 . Mandrel  11 , which still has tooling members  17 , composite layers  37 , and part formation aids  39  attached, is placed in a vacuum bag  45  as illustrated in FIG.  3 . The vacuum bag tube  45  is placed over the mandrel with the aid of a hard outer tube. The hard outer tube is made in an appropriate shape to allow for easy fit over the mandrel  11 . Vacuum bag  45  is used in conjunction with two sealing end plates  47 , a vacuum fitting  49 , and a vacuum line  51 . After placing vacuum bag  45  over mandrel  11 , the bag  45  is sealed to each end plate  47  using a typical band clamp  53 . Vacuum line  51  is then connected at one end to vacuum fitting  49  and at its other end to a vacuum source (not shown). 
     The formation of a composite part using the present invention is accomplished by creating a vacuum within vacuum bag  45 . After vacuum bag  45  has been sealed, the air from the bag can be evacuated through vacuum line  51  leading from a central passage in mandrel  11 . The evacuation of air causes bag  45  to apply a formation force to each part formation aid  39  that is directed toward the center of mandrel  11 . As the formation force is applied to each part formation aid  39 , the sidewalls  42  of part formation aid  39  deform downward (toward the center of mandrel  11 ), thereby causing the side portions of the segments of composite layers  37  to be forced down and around tooling member  17 . The folding sidewalls  42  cause side portions of composite layers  37  alongside each slit  44  to be pushed into abutment with one of the sidewalls  18 , as shown in FIG.  4 . Tooling members  17  define the general shape that a composite part  61  (FIG. 5) produced by the current process will become. After initial formation of the composite part, the part is held in place by vacuum bag  45  during a curing period. 
     Although not illustrated, the forming and curing processes described above are usually carried out inside an oven or autoclave. An autoclave allows the ambient temperature and pressure surrounding vacuum bag  45  to be raised. The exact temperature and pressure requirements for forming and curing a part depend on the resin that is used with the composite layers  37 . Typically, curing takes place at elevated temperatures and pressures. Some resins actually allow the forming and curing processes to take place at room temperature and atmospheric pressure, which would negate the need for an autoclave. 
     After forming and curing, the composite part  61  (FIG. 5) is in its final shape. In FIGS. 1,  2 , and  3 , the shape of tooling members  17  is designed to produce hat-shape stiffener  61  as shown in FIG.  5 . Stiffener  61  is commonly used with flat panels to provide added rigidity and strength to the panels. The most common application of composite stiffener  61  is in aerospace applications where strong and lightweight materials are needed. 
     Mandrel  11  is configured to produce four hat-shape stiffeners  61  per production run. Referring to FIG. 6, a mandrel  71  having a cylindrical outer surface  73  with a plurality of tooling members  75  is illustrated. Because of the large number of tooling members  75  disposed on mandrel  71 , the production of stiffeners  61  would be greatly increased over the configuration provided by mandrel  11 . The process described above for forming stiffeners  61  using mandrel  11  is the same process that would be used to produce stiffeners  61  using mandrel  71 . 
     Referring to FIG. 7, a mandrel  81  having integrally formed tooling members  83  is illustrated. The shape of tooling members  83  is such that a stiffener  85  having a Z-shaped cross section is formed using the process described above. Foam blocks (not shown) would be initially located between each tooling member  83 , resulting in a cylindrical exterior. Fibers would placed over the cylindrical exterior. The foam blocks would be removed. Formation aids would bend the fibers into the shape as shown in FIG.  7 . Although tooling member  83  are shown as an integral portion of mandrel  81 , tooling members  83  could be separate parts that are mechanically fastened to the mandrel  81 . Again, a material other than foam could be used in place of the foam blocks. 
     Referring to FIGS. 8,  9  and  10 , an alternate method of manufacturing composite parts according to the present invention is illustrated. A composite part  111  (FIG. 8) having a first portion, or flanged panel  113  and a plurality of second portions, or stiffeners  115  is produced, first portion  113  being co-cured to second portions  115 . Composite part  111  has a flanged composite panel (first portion  113 ) with several hat-shaped stiffeners (second portions  115 ) attached to the panel. 
     Referring more specifically to FIG. 9, a mandrel  117  having an outer surface  119  is provided to aid in the formation of first portion  113 . Although the actual shape of first portion  113  and thus the shape of the mandrel could vary depending on the application, the flanged composite panel  113  of part  111  is produced using a rectangular mandrel  117 . 
     Before forming flanged panel  113 , a filler member  121  is attached to outer surface  119  of mandrel  117  in any place needed to form a rounded or smooth corner for the formation of first portion  113 . An outer surface of filler member  121  combines with outer surface  119  to form an application surface  123 . 
     A plurality of composite layers  131  are positioned on application surface  123  using fiber placement, tape winding, filament winding, or hand-laying methods. A preferred method of applying composite layers  131  is fiber placement. Fiber placement of the composite material  131  would prevent having to wrap composite layers  131  around the entire circumference of mandrel  117 . Instead the composite layers  131  can be placed wherever material is needed on application surface  123 . 
     After applying composite material  131 , the material  131  is cut along slit  133  to trim and remove any excess material. Filler member  121  is then removed from the mandrel  117 . Mandrel  117  and composite material  131  is placed in a vacuum bag (not shown) as described previously. The vacuum bag is sealed and evacuated of air, causing the bag to apply a force to the composite layers  131 , thereby forming the flanged composite panel  113  of composite part  111 . 
     Depending on the curing requirements of the composite material  131  used, the evacuation process may or may not take place inside an autoclave. Regardless of whether an autoclave is used, flanged panel  113  is only to be partially cured. This allows stiffeners  115  to be co-cured to flanged panel  113 . 
     Referring more specifically to FIG. 10, a plurality of tooling members  141 , each having a center channel  142 , is disposed on a surface of flanged panel  113  opposite mandrel  117 . The shape of tooling members  141  determine the final shape of stiffeners  115 . In FIG. 8, tooling members  141  are trapezoidal in cross-section to produce a hat-shaped stiffener similar to stiffener  61 . Filler members  143  are placed on flanged panel  113  between tooling members  141  to form a smooth flat application surface  145 . 
     A plurality of composite layers  151  are positioned on application surface  145 . The preferred method of applying composite material  151  is fiber placement, although any method of positioning composites could be used. Part formation aids (not shown but similar to part formation aids  39 ) are placed on composite layers  151  above each tooling member  141 . Composite layers  151  are cut between each tooling member  141  on a line parallel to the lengthwise axis of each tooling member  141 . Filler members  143  are removed from between tooling members  141 . 
     Stiffeners  115  are formed using the evacuation technique previously described. Mandrel  117 , first portion  113 , tooling members  141 , composite layers  151 , and the part formation aids are placed inside a vacuum bag (not shown). The vacuum bag is sealed and evacuated, thereby causing the bag to exert a force on the part formation aids. The part formation aids deform toward flanged panel  113 , which causes composite layers  151  to conform to the shape of tooling members  141 . 
     After forming stiffeners  115 , both the flanged panel  113  and the stiffeners  115  are co-cured to obtain the composite part  111 . Depending on the requirements of the resin used in composite layers  131 ,  151 , the forming and curing processes will likely be performed inside an autoclave at elevated temperatures and pressures. 
     After forming composite part  111 , tooling members  141  are removed by creating a vacuum within center channel  142 . Since tooling members  141  are generally made from rubber, the tooling members  141  will compress inwardly and are easily removed from within stiffeners  115 . 
     The scope of the alternate method described is not limited to forming a part having only a first portion and a second portion. Instead, a part having a plurality of portions may be formed by similar steps, each portion being co-cured to another in one of the final steps. Finally, it is conceivable that a similar method be used to form and attach a composite part to a preexisting composite part formed by a process outside the scope of the present invention. 
     One advantage of the present invention is that it provides an automated method for manufacturing composite parts using a circular array of tooling members. By providing a smooth and partially rounded application surface, composite material can be quickly and efficiently positioned by an automated positioning technique (i.e. fiber placement, tape winding, filament winding). Additionally, the circular configuration of tooling members allows rapid, large-scale production of complex composite shapes that normally require extensive manufacturing time. 
     Another advantage of the present invention is that it provides a method of automating the manufacture of two or more composite components that will form a final composite part. By preparing smooth application surfaces, automated composite positioning techniques can be used to create any number of different composite components. The various components can then be joined to form the final part during a co-curing or co-bonding process. 
     It should be apparent from the foregoing that an invention having significant advantages has been provided. While the invention is shown in only a few of its forms, it is not just limited but is susceptible to various changes and modifications without departing from the spirit thereof.