Abstract:
A composite panel includes a support layer including a first plurality of prepreg plies wrapped around at least one mandrel; a mid-plane impact layer including a second plurality of prepreg plies, the mid-plane impact layer adjacent to the support layer; a upper skin layer including a third plurality of prepreg plies, the upper skin layer adjacent to the mid-place impact layer; and a lower skin layer including a fourth plurality of prepreg plies, the lower skin layer adjacent to the support layer; whereby each of the first, second, third, and fourth plurality of prepreg plies are co-cured to form the composite panel.

Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     This invention was made with Government support with the United States Navy under Contract No. N00019-06-0081. The Government therefore has certain rights in this invention. 
    
    
     FIELD OF THE INVENTION 
     The subject matter disclosed herein relates generally to the field of composite structures, and more particularly, to an impact resistant composite panel and methods for making such composite panels. 
     DESCRIPTION OF THE RELATED ART 
     Conventional aircraft floor systems incorporate thin metallic or composite prepreg skin materials adhesively bonded to either a metallic or non-metallic honeycomb. Recent improvements in machining technology have led to the use of high speed machined aluminum floor designs for rotary-wing aircraft. High speed machined metallic aircraft floor systems require minimal tooling and have higher tolerances, but may not achieve the weight benefits of composite floor systems. Although composite floor systems are generally lighter in weight than metallic floors, composite floor systems may be more expensive and labor intensive to manufacture compared to metallic floor systems. 
     Conventional typical aircraft composite floor system floor panels utilize a honeycomb core material with pre-cured fiberglass or graphite composite skins (prepreg) bonded thereto in a large heated press or autoclave. Local hard points and edge closeouts are typically accomplished using an epoxy potting compound. These composite floor system floor panels may require relatively complicated and labor intensive process steps to complete fabrication of an individual panel. Moreover, usage of honeycomb core structures in rotary-wing aircraft composite floor systems may suffer inherent moisture absorption in service due to the open cell structure. Such moisture absorption may result in increased weight and resultant performance degradation over a prolonged time period. Accordingly, it is desirable to provide a lightweight aircraft floor system that meets or exceeds design requirement. 
     BRIEF SUMMARY OF THE INVENTION 
     According to one aspect of the invention, a method of forming a composite panel includes applying a support layer to each of a plurality of mandrels to form a plurality of wrapped mandrels; stacking each of the wrapped mandrels in a defined orientation; applying a mid-plane impact layer on top of wrapped mandrels; applying an upper skin layer on top of the mid-place impact layer; applying a lower skin layer to the bottom of the wrapped mandrels; and curing the support layer, the mid-plane layer, upper skin layer, and lower skin layer together at an elevated temperature and pressure; where each of the support layer, the mid-plane impact layer, the upper skin layer, and the lower skin layer includes a plurality of prepreg plies. 
     According to another aspect of the invention, a composite panel includes a support layer including a first plurality of prepreg plies wrapped around at least one mandrel; a mid-plane impact layer including a second plurality of prepreg plies, the mid-plane impact layer adjacent to the support layer; a upper skin layer including a third plurality of prepreg plies, the upper skin layer adjacent to the mid-place impact layer; and a lower skin layer including a fourth plurality of prepreg plies, the lower skin layer adjacent to the support layer; whereby each of the first, second, third, and fourth plurality of prepreg plies are co-cured to form the composite panel. 
     Other aspects, features and techniques of the invention will become more apparent from the following description taken in conjunction with the drawings. 
    
    
     
       BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS 
       The subject matter, which is regarded as the invention, is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features, and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which like elements are numbered alike in the several FIGURES: 
         FIG. 1A  is a general perspective view of an exemplary rotary wing aircraft for use with an embodiment of the invention; 
         FIG. 1B  is a partial perspective view of the rotary wing aircraft of  FIG. 1A  which utilizes a composite panel according to an embodiment of the invention; 
         FIG. 2A  is a general perspective view of a composite panel of  FIG. 1B ; 
         FIG. 2B  is a sectional view of a composite panel according to an embodiment of the invention; 
         FIG. 3A  is partial cross-sectional view of a composite panel according to an embodiment of the invention; 
         FIG. 3B  is a chart delineating each layer of the composite panel of  FIG. 3A  according to an embodiment of the invention; 
         FIG. 4A  is partial cross-sectional view of a composite panel according to an embodiment of the invention; 
         FIG. 4B  is a chart delineating each layer of the composite panel of  FIG. 4A  according to an embodiment of the invention; 
         FIG. 5A  is partial cross-sectional view of a composite panel according to an embodiment of the invention; 
         FIG. 5B  is a chart delineating each layer of the composite panel of  FIG. 5A  according to an embodiment of the invention; 
         FIG. 6A  is partial cross-sectional view of a composite panel according to an embodiment of the invention; 
         FIG. 6B  is a chart delineating each layer of the composite panel of  FIG. 6A  according to an embodiment of the invention; and 
         FIG. 7  illustrates a flowchart depicting a method of fabricating a composite panel according to an embodiment of the invention. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to the figures,  FIGS. 1A-1B  schematically illustrates a rotary-wing aircraft  10  incorporating a composite panel  30  ( FIGS. 2A-6B ) according to an embodiment of the invention. As illustrated in  FIG. 1A , rotary-wing aircraft  10  has a main rotor system  12  and includes an airframe  14  having an extending tail  16  which mounts a tail rotor system  18 , such as an anti-torque system. The main rotor system  12  is shown with a multiple of rotor blades  20  mounted to a rotor hub. The main rotor system  12  is driven about an axis of rotation R through a main gearbox by one or more engines  22 . While embodiments of the invention are shown and described with reference to a rotary-wing aircraft  10  and are particularly suited to a rotary-wing aircraft  10 , aspects of this invention can also be used in other configurations and/or machines such as, for example, automotive applications including commercial and military ground vehicles, building structures, construction applications such as infrastructure, cargo applications, oil and gas industrial applications, shipping applications including containers for rail, marine and aircraft, fixed-wing aircraft applications, non-rotary-aircraft applications, high speed compound rotary wing aircraft with supplemental translational thrust systems, dual contra-rotating, coaxial rotor system aircraft, turbo-props, tilt-rotors and tilt-wing aircraft. 
     Referring to  FIG. 1B , the airframe  14  includes an airframe section  14 A. The airframe section  14 A may include a multitude of frame members  24  and a multitude of beam members  26  which support a cabin floor system  28 . The multitude of frame members  24  and beam members  26  may be arranged in a generally rectilinear pattern. The cabin floor system  28  may be formed of a multiple of composite panels  30  (See e.g.,  FIG. 2A ) utilizing a box beam construction for providing the rigidity for rotor-craft applications. It is to be appreciated that additional composite panels  30  may be assembled in other areas of the airframe  14 A and provide the rigidity necessary for high impact rotor-craft applications. 
     Referring to  FIGS. 2A-2B , the composite panel  30  for use with the cabin floor system  28  ( FIG. 1B ) is illustrated according to an exemplary embodiment of the invention. As illustrated, each composite panel  30  may be manufactured using a box beam construction that includes wrapping a multiple of unidirectional or woven fiberglass or graphite fiber plies impregnated with an epoxy resin (fiberglass or graphite prepreg) around a multiple of mandrels (not shown). In an embodiment, the woven plies have at least two threads that are woven together in a plain weave pattern. Also, the prepreg plies may include prepreg tape or prepreg fabric. A plurality of similarly wound mandrels are assembled together and co-cured with additional prepreg plies to form a multi-laminate composite panel. In embodiments, the composite panel  30  may utilize a mandrel (not shown) of various shapes and sizes such as, for example, a rectangular, a triangular, a square, or the like for the specific application. 
     In an embodiment, the composite panel  30  is assembled as a sandwich structure having a multiple of layers with a multiple of prepreg plies bonded together and co-cured at the same time through an autoclave process to form a multi-laminate assembly. The composite panel  30  may be manufactured in a single curing process using an autoclave processing but other processing techniques may be utilized. As illustrated in  FIGS. 2A-2B , the composite panel  30  is a multi-laminate system and may be formed from a multiple of prepreg plies forming the uncured layers and generally includes a multiple of rectilinear box members defining a support layer  32 , a multi-ply mid-plane impact layer (shown in  FIGS. 3A-6B ), an upper skin layer  34  and a lower skin layer  36 . These uncured layers are coupled together during assembly and co-cured in an autoclave processing method to form a co-bonded composite panel  30 . As illustrated in  FIG. 2B , each support layer forming the support laminate  32  includes a box construction that may be manufactured by wrapping one or more unidirectional or woven prepreg plies about a generally rectangular mandrel (having a longitudinal length along axis A). Additionally, the wound mandrels (not shown) may be arranged or stacked with a multiple of similar prepreg wound mandrels along axis B. The stacked prepreg wound mandrels are sandwiched between a multiple of prepreg plies that may be draped onto the plurality of mandrels (not shown) and are co-bonded and co-cured using prepreg autoclave processing to yield a lightweight and impact tolerant composite panel  30  having a generally hollow support structure. A trimming operation may be performed after the autoclave curing process in order to provide other feature openings, for example opening  38 , for installation of tie-down pans  40 . In embodiments, the composite panel  30  including a support laminate  32 , upper skin laminate  34 , mid-plane impact laminate (See e.g.,  FIG. 3A ), and lower skin laminate  36  may be manufactured from a thermoset composite matrix material including a multiple of thermoset composite fibers. In an embodiment, the unidirectional or woven graphite and fiberglass plies may be HexPly® Prepreg available from the Hexcel Corporation and may include graphite prepreg as a plain weave IM7 3K with an 8552 resin system, fiberglass prepreg ply as unidirectional fiberglass S-2 with an 8552 resin system, and a woven fiberglass as 7781 8HS fiberglass with an 8552 resin system. 
       FIGS. 3A-3B  illustrate a composite panel  50  to be used with cabin floor system  28  ( FIG. 1B ) that is constructed according to an exemplary embodiment. As illustrated, composite panel  50  is constructed from a multiple of prepreg plies to form the respective laminates described above in  FIG. 2A-2B  and generally includes a multi-ply support layer  52 , a mid-plane impact layer  54 , an upper skin layer  56  and a lower skin layer  58 . The support layer  52  is a four-ply layer and includes, traversing radially outwards from the center of support layer  52  in direction of axis C, a 0 (zero) degree or 90 (ninety) degree woven fiberglass ply, a 0 (zero) degree or 90 (ninety) degree woven graphite ply, a 45 (forty-five) degree woven graphite ply, and a 45 (forty-five) degree woven graphite ply. The mid-plane impact layer  54  is a three-ply layer and includes, traversing radially outwards, a 0 (zero) degree or 90 (ninety) degree woven fiberglass ply, a 0 (zero) degree or 90 (ninety) degree woven fiberglass ply, and a 0 (zero) degree or 90 (ninety) degree woven fiberglass ply. The upper skin layer  56  is a four-ply layer and includes, traversing radially outwards, a 45 (forty-five) degree woven graphite ply, a 45 (forty-five) degree woven graphite ply, a 0 (zero) degree or 90 (ninety) degree woven graphite ply, and a 0 (zero) degree or 90 (ninety) degree woven fiberglass ply. The lower skin layer  58  is a three-ply layer and includes, traversing radially outwards, a 0 (zero) degree or 90 (ninety) degree woven graphite ply, a 45 (forty-five) degree woven graphite ply, and a 0 (zero) degree or 90 (ninety) degree woven graphite ply.  FIG. 3B  is a chart delineating each layer of the composite panel  50 . It is to be appreciated that graphite ply constitutes about 60 percent while fiberglass constitutes about 40 percent of the total number of prepreg plies used in the composite panel  50 , but other percentages may be utilized. It is also to be appreciated that the support layer  52 , mid-plane impact layer  54 , and upper skin layer  56  include prepreg plies that are arranged in a symmetrical order as we traverse radially outwards from the surface of the mandrel to the upper skin layer  56  in order to support thermal expansion of the plies and minimize or reduce warping of the plies during cure. For example, the inner most ply of support layer  52  that is immediately adjacent to the mandrel is the same as the outermost ply of the upper skin layer  56 , which is 0 (zero) degree or 90 (ninety) degree woven fiberglass ply. Similarly, there is a symmetrical arrangement of the other prepreg plies as we traverse radially outwards from the innermost ply of support layer  52 . 
       FIGS. 4A-4B  illustrate a composite panel  60  to be used with cabin floor system  28  ( FIG. 1B ) that is constructed according to an exemplary embodiment. As illustrated, composite panel  60  generally includes a multi-ply support layer  62 , a mid-plane impact layer  64 , an upper skin layer  66  and a lower skin layer  68 . The support layer  62  is a four-ply layer and includes, traversing radially outwards from the center of support layer  62  in a direction of axis D, a 90 (ninety) degree unidirectional fiberglass S-2 ply, a 0 (zero) degree or 90 (ninety) degree woven graphite ply, a 45 (forty-five) degree woven graphite ply, and a 45 (forty-five) degree woven graphite ply. The mid-plane impact layer  64  is a three-ply layer and includes, traversing radially outwards, a 0 (zero) degree unidirectional fiberglass S-2 ply, a 0 (zero) degree or 90 (ninety) degree woven fiberglass ply, and a 0 (zero) degree unidirectional fiberglass S-2 ply. The upper skin layer  66  is a four-ply layer and includes, traversing radially outwards, a 45 (forty-five) degree woven graphite ply, a 45 (forty-five) degree woven graphite ply, a 0 (zero) degree or 90 (ninety) degree woven graphite ply, and a 90 (ninety) degree unidirectional fiberglass S-2 ply. The lower skin layer  68  is a three-ply layer and includes, traversing radially outwards, a 0 (zero) degree or 90 (ninety) degree woven graphite ply, a 45 (forty-five) degree woven graphite ply, and a 0 (zero) degree or 90 (ninety) degree woven graphite ply.  FIG. 4B  is a chart delineating each layer of the composite panel  60 . It is to be appreciated that graphite ply constitutes about 60 percent while fiberglass constitutes about 40 percent of the total number of prepreg plies used in the composite panel  60 , but other percentages may be utilized. It is also to be appreciated that the support layer  62 , mid-plane impact layer  64 , and upper skin layer  66  include prepreg plies that are arranged in a symmetrical order as we traverse radially outwards from the surface of the mandrel to the upper skin layer  66  in order to support thermal expansion of the plies and minimize or reduce warping of the plies during cure. For example, the inner most ply of support layer  62  that is immediately adjacent to the mandrel is the same as the outermost ply of the upper skin layer  66 , which is 90 (ninety) degree unidirectional fiberglass S-2 ply. Similarly, there is a symmetrical arrangement of the other prepreg plies as we traverse radially outwards from the innermost ply of support layer  62 . 
       FIGS. 5A-5B  illustrate a composite panel  70  to be used with cabin floor system  28  ( FIG. 1B ) that is constructed according to an exemplary embodiment. As illustrated, composite panel  70  generally includes a multi-ply support layer  72 , a mid-plane impact layer  74 , an upper skin layer  76  and a lower skin layer  78 . The support layer  72  is a five-ply layer and includes, traversing radially outwards from the center of support layer  72  in direction of axis E, a 45 (forty-five) degree unidirectional fiberglass S-2 ply, a −45 (minus forty-five) degree unidirectional fiberglass S-2 ply, a −45 (minus forty-five) degree unidirectional graphite ply, a 45 (forty-five) degree unidirectional graphite ply, and a 90 (ninety) degree unidirectional graphite ply. The mid-plane impact layer  74  is a three-ply layer and includes, traversing radially outwards, a 0 (zero) degree unidirectional graphite ply, a 90 (ninety) degree unidirectional graphite ply, and a 0 (zero) degree unidirectional graphite ply. The upper skin layer  76  is a six-ply layer and includes, traversing radially outwards, a 0 (zero) degree unidirectional graphite ply, a 90 (ninety) degree unidirectional graphite ply, a 45 (forty-five) degree unidirectional graphite ply, a −45 (minus forty-five) degree unidirectional graphite ply, a −45 (minus forty-five) unidirectional fiberglass S-2 ply, and a 45 (forty-five) degree unidirectional fiberglass S-2 ply. The lower skin layer  78  is a five-ply layer and includes, traversing radially outwards, a 45 (forty-five) degree unidirectional graphite ply, a −45 (minus forty-five) degree unidirectional graphite ply, a −45 (minus forty-five) degree unidirectional graphite ply, a 45 (forty-five) degree unidirectional graphite ply, and a 90 (ninety) degree unidirectional graphite ply.  FIG. 5B  is a chart delineating each layer of the composite panel  70 . It is to be appreciated that graphite ply constitutes about 80 percent while fiberglass constitutes about 20 percent of the total number of prepreg plies used in the composite panel  70 , but other percentages may be utilized. It is also to be appreciated that the support layer  72 , mid-plane impact layer  74 , and upper skin layer  76  include prepreg plies that are arranged in a symmetrical order as we traverse radially outwards from the surface of the mandrel to the upper skin layer  76  in order to support thermal expansion of the plies and minimize or reduce warping of the plies during cure. For example, the inner most ply of support layer  72  that is immediately adjacent to the mandrel is the same as the outermost ply of the upper skin layer  76 , which is a 45 (forty-five) degree unidirectional fiberglass S-2 ply. Similarly, there is a symmetrical arrangement of the other prepreg plies as we traverse radially outwards from the innermost ply of support layer  72 . 
       FIGS. 6A-6B  illustrate a composite panel  80  to be used with cabin floor system  28  ( FIG. 1B ) that is constructed according to an exemplary embodiment. As illustrated, composite panel  80  generally includes a multi-ply support layer  82 , a mid-plane impact layer  84 , an upper skin layer  86  and a lower skin layer  88 . The support layer  82  is a five-ply layer and includes, traversing radially outwards from the center of support layer  82  in direction of axis F, a 45 (forty-five) degree unidirectional fiberglass S-2 ply, a −45 (minus forty-five) degree unidirectional fiberglass S-2 ply, a −45 (minus forty-five) degree unidirectional graphite ply, a 45 (forty-five) degree unidirectional graphite ply, and a 90 (ninety) degree unidirectional graphite ply. The mid-plane impact layer  84  is a three-ply layer and includes, traversing radially outwards, a 0 (zero) degree unidirectional graphite ply, a 90 (ninety) degree unidirectional graphite ply, and a 0 (zero) degree unidirectional graphite ply. The upper skin layer  86  is a six-ply layer and includes, traversing radially outwards, a 0 (zero) degree unidirectional graphite ply, a 90 (ninety) degree unidirectional graphite ply, a 45 (forty-five) degree unidirectional graphite ply, a −45 (minus forty-five) degree unidirectional graphite ply, a −45 (minus forty-five) unidirectional fiberglass S-2 ply, and a 45 (forty-five) degree unidirectional fiberglass S-2 ply. The lower skin layer  88  is a five-ply layer and includes, traversing radially outwards, a 45 (forty-five) degree unidirectional graphite ply, a −45 (minus forty-five) degree unidirectional graphite ply, a −45 (minus forty-five) degree unidirectional graphite ply, a 45 (forty-five) degree unidirectional graphite ply, and a 90 (ninety) degree unidirectional graphite ply.  FIG. 6B  is a chart delineating each layer of the composite panel  80 . It is to be appreciated that graphite ply constitutes about 80 percent while fiberglass constitutes about 20 percent of the total number of prepreg plies used in the composite panel  70 , but other percentages may be utilized. It is also to be appreciated that the support layer  82 , mid-plane impact layer  84 , and upper skin layer  86  include prepreg plies that are arranged in a symmetrical order as we traverse radially outwards from the surface of the mandrel to the upper skin layer  86  in order to support thermal expansion of the plies and minimize or reduce warping of the plies during cure. For example, the inner most ply of support layer  82  that is immediately adjacent to the mandrel is the same as the outermost ply of the upper skin layer  86 , which is a 45 (forty-five) degree unidirectional fiberglass S-2 ply. Similarly, there is a symmetrical arrangement of the other prepreg plies as we traverse radially outwards from the innermost ply of support layer  82 . 
       FIG. 7  illustrates a method  90  of fabricating a composite floor  30  ( FIGS. 2A-2B ) according to an embodiment of the invention. With continued reference to  FIGS. 2A-6B , at  91 , prepreg plies for support layer  52 ,  62 ,  72 ,  82  are compiled. In an embodiment, a plurality of rectangular shaped mandrels are selected as the core and a rectangular pattern for each prepreg ply comprising support layer  52 ,  62 ,  72 , or  82  is cut to a predetermined length. One full wrap of each prepreg ply is wrapped around an external surface of a mandrel along its radial width. The wrapped prepreg ply is trimmed or cut to provide one full wrap of an external circumferential diameter of the mandrel. This wrapping process is repeated for the other prepreg plies that form the support layer  52 ,  62 ,  72  or  82 . In  92 , additional mandrels are wrapped according to the step depicted in  92 . Further, the wrapped mandrels are ganged together along their longitudinal length according to, in one non-limiting example, the pattern depicted in  FIG. 2B . 
     In  93 , the mid-plane impact layer  54 ,  64 ,  74 ,  84  is applied onto the top surface of the support layer  52 ,  62 ,  72  or  82 . Particularly, prepreg plies forming the mid-plane impact layer  54 ,  64 ,  74 ,  84  are selected and cut according to the shape and size of the composite panel  30  being fabricated. Each prepreg ply forming the mid-plane impact layer  54 ,  64 ,  74 ,  84  is applied to the top surface of the support layer  52 ,  62 ,  72  or  82 . Similarly, in  94 , prepreg plies forming the upper skin layer  56 ,  66 ,  76  or  86  are selected and cut according to the shape and size of the composite panel  30  being fabricated. Each prepreg ply is applied to the top surface of the mid-plane impact layer  54 ,  64 ,  74 ,  84 . In  95 , prepreg plies forming the lower skin layer  58 ,  68 ,  78  or  88  are selected and cut according to the shape and size of the composite panel  30  being fabricated. Next, the assembled group of uncured prepreg plies forming the support layer  52 ,  62 ,  72 ,  82 , the mid-plane impact layer  54 ,  64 ,  74 ,  84 , and the upper skin layer  56 ,  66 ,  76 , or  86  is placed on each prepreg ply forming the lower skin layer  58 ,  68 ,  78 , or  88  to form an uncured prepreg assembly. In an embodiment, the uncured prepreg assembly is laid onto a flat tooling surface and held together with caul plates on its edges for transmitting normal pressure to the finished layers during curing. At  96 , the prepreg assembly is subjected to a co-curing process at an elevated temperature (e.g., in excess of about 250 degree Fahrenheit or 394 degree Kelvin) and at the same time, pressure may be applied (e.g., about 1 bar (about 100,000 Pa) to about 10 bar (about 1,000,000 Pa)) in order to activate the epoxy resin. In  97 , the mandrels are removed from the cured laminate (i.e., cured prepreg assembly) and a trimming operation may be performed to the cured laminate in order to provide other features such as, for example, openings for tie-down pans or attachment of further components to complete the assembly of the composite panel  30 . While a specific process  90  is described above, it is understood that other temperatures, pressures and environments may be used according to the specific application, and that process  90  need not be used in all aspects of the invention such as where openings or features are not needed. 
     The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the invention. While the description of the present invention has been presented for purposes of illustration and description, it is not intended to be exhaustive or limited to the invention in the form disclosed. Many modifications, variations, alterations, substitutions or equivalent arrangement not hereto described will be apparent to those of ordinary skill in the art without departing from the scope and spirit of the invention. Additionally, while the various embodiment of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.