Abstract:
Composite laminates used in structural applications include an interlayer of soft material that provides damping action to reduce noise and vibration. The interlayer may comprise a viscoelastic material which deforms under stress caused by shock, noise or vibration. A reinforcement may be embedded in the viscoelastic material to maintain the mechanical strength and stiffness of the laminate. The reinforcement may include individual or woven fibers or ridged tubes that provide the interlayer with stiffness.

Description:
TECHNICAL FIELD 
     This disclosure generally relates to composite laminates used in structural applications, especially aircraft, and deals more particularly with a composite laminate having a reinforced interlayer that provides structural damping. 
     BACKGROUND 
     Composite materials such as carbon fiber reinforced epoxy resin are used in aircraft applications because of their light weight and high strength, compared to metals such as aluminum. More recently, these composite materials have been used in the fuselage structure which surrounds interior cabins in the aircraft. The use of composite materials in the fuselage structure presents an opportunity to reduce engine and aerodynamic noise, as well as vibration transmission to the interior of the aircraft. 
     In order to reduce noise and vibration, “add-on” parts may be installed on the aircraft which function to at least partially damp vibrations and noise to prevent propagation to the interior cabin. In order to adequately reduce noise and vibration, a relatively large number of these add-on parts may be necessary which are costly both in terms of material and labor installation costs. Moreover, these additional parts add to the weight of the aircraft. 
     Designing aircraft structures such as a fuselage having high inherent damping is particularly challenging when using composite materials. The composite material is typically cured at relatively high temperatures and pressures, in contrast to the operating conditions of the aircraft in which the fuselage skin typically encounters temperatures approaching −60° F. or lower at typical flight altitudes. Thus, engineering a damping material system that performs well at cold temperatures (normally requiring a very soft material) but can survive the heat and pressure when co-cured with the base material, may be particularly difficult. The ideal material that performs well at such cold operating temperatures has a very low glass transition temperature (Tg), such that it is in a soft transition phase at operating temperatures. Further, in order to use thin films of the damping material at these cold temperatures for low-weight applications, the modulus of elasticity of the material will typically be very low compared to the carbon/epoxy composite. Thus, the use of relatively soft materials to provide inherent damping within composite material structures may make it less stiff since the relatively soft damping material is substantially less stiff than the typical plies of carbon fiber reinforced plastics (CFRP), sometimes also referred to as organic composite materials. 
     Accordingly, there is a need for a composite material structure that has relatively high inherent damping qualities without materially reducing the stiffness and other mechanical performance characteristics of the structure. Embodiments of the disclosure are directed towards satisfying this need. 
     SUMMARY 
     An embodiment of the disclosure provides a damped composite laminate, which may include at least first and second layers of a reinforced resin material, and a third layer of damping material co-cured to first and second layers. The third layer of damping material may include a viscoelastic material having a reinforcement medium for stiffening the viscoelastic material. The reinforcement medium may include fibers embedded in the viscoelastic material. The fibers may have a length extending in a direction generally transverse to the planes of the first and second layers. The fibers may be formed of glass or carbon tow or a lightweight synthetic cloth, which are impregnated or coated with the viscoelastic material. The fibers may be formed of a second viscoelastic material, having a glass transition temperature greater than the glass transition temperature of the viscoelastic material in which the fibers are embedded. The third layer may include graphite nano-fibers or nano-tubes (Multi-wall (MWNT) or Single-Wall (SWNT)), or nano or micro sized particles dispersed within the viscoelastic material. The nano-fibers or nano-tubes or particles may be contained in a film of viscoelastic material, such as thermoplastic polyurethane. 
     In accordance with another embodiment, a composite laminate structure is provided, which may include at least first and second layers of a carbon fiber reinforced plastics (CFRP), and a third layer of reinforced viscoelastic material between the first and second layers. The viscoelastic material may be a thermoplastic polyurethane, or other highly damped polymer, such acrylic, or latex rubber. The third layer may not be continuous, but rather may have discontinuities that bridge between the first and second layer. The bridging may be accomplished with a narrow strip of high modulus carbon-organic resin prepreg, or slit-tape. The slit-tape may have a length that runs transverse to the longitudinal stiffeners of the aircraft fuselage. The bridging may also be accomplished by introducing perforations in the viscoelastic material that are filled with resin migrating from the first and second layers during curing. The bridging may be accomplished through the introduction of fiber tow that run perpendicular (Z-Fiber) to the first and second layers, through the thickness of the third layer. The length of these fiber tows may exceed the thickness of the third layer, such that their ends extend into the first and second layers. These fiber tows may consist of carbon or glass fibers and may be pre-impregnated with epoxy or suitable organic resins. The third layer is co-cured with the first and second layers so that the composite laminate is provided with a reinforced interlayer that provides inherent damping of the structure. 
     Another embodiment of the disclosure provides a method for making a damped composite laminate structure. The method may comprise the steps of placing a layer of damping material between first and second layers of carbon fiber reinforced plastic (CFRP) material, and co-curing the layer of damping material with the first and second layers. The co-curing is achieved by compressing the first and second layers with the layer of damping material, and co-curing the first and second layers along with the layer of damping material. The layer of damping material may be attached to the first layer following which the second layer is applied over the layer of damping material. The method may further include introducing reinforcement into the layer of damping material before co-curing is performed. The introduction of reinforcement into the layer of damping material may include providing a reinforcement medium and infusing the reinforcement medium with a viscoelastic material. 
     A further embodiment of the disclosure provides a method of making a composite laminate structure which may comprise the steps of forming first and second pre-pregs; forming a layer of damping material that provides the structure with damped qualities; forming a lay-up by placing the layer of damping material between the first and second pre-pregs; and, co-curing the lay-up. The first and second pre-pregs along with the damping layer are compressed during co-curing. The first and second pre-pregs may be formed by laying up multiple plies of a carbon fiber reinforced plastic material such as carbon epoxy composites. The layer of damping material may be prepared by forming a pre-preg of thermoplastic coated reinforcing fibers comprising either individual fibers or a web of reinforcing fibers. 
     These and further features, aspects and advantages of the embodiments will become better understood with reference to the following illustrations, description and claims. 
    
    
     
       BRIEF DESCRIPTION OF THE ILLUSTRATIONS 
         FIG. 1  is a cross sectional illustration of a composite laminate structure having a damping interlayer according to one embodiment of the disclosure. 
         FIG. 2  is a cross sectional illustration of a composite laminate structure having a damping interlayer according to another embodiment of the disclosure. 
         FIG. 3  is a cross sectional illustration of a composite laminate structure having a damping interlayer according to another embodiment of the disclosure. 
         FIG. 4  is a cross sectional illustration of a composite laminate structure having a damping interlayer according to another embodiment of the disclosure. 
         FIG. 5  is a cross sectional illustration of a wetted reinforcing fiber which may be used in the composite laminate structure shown in  FIG. 4 . 
         FIG. 6  is a cross sectional illustration of a composite laminate structure having a damping interlayer according to another embodiment of the disclosure. 
         FIG. 7  is an enlarged, fragmentary illustration of a portion of the composite laminate structure shown in  FIG. 6 . 
         FIG. 8  is a perspective illustration of a single Z-fiber used in the interlayer shown in  FIGS. 6 and 7 . 
         FIG. 9  is a side elevation illustration of a VEM interlayer having Z-fibers pre-inserted therein. 
         FIG. 10  is a view similar to  FIG. 9 , but showing laminate layers having been pressed onto opposite sides of the VEM interlayer. 
         FIG. 11  is a plan, cross sectional illustration of a composite laminate structure having Z-fibers distributed around the perimeter of a VEM interlayer. 
         FIG. 12  is an illustration similar to  FIG. 11 , but showing Z-fibers uniformly distributed across the VEM interlayer. 
         FIG. 13  is a plan illustration of another embodiment of a composite laminate structure, employing a slit tape reinforcement in the interlayer. 
         FIG. 14  is a sectional illustration taken along the line  14 - 14  in  FIG. 13 . 
         FIG. 15  is a plan illustration of another embodiment of the composite laminate damping structure having a perforated interlayer. 
         FIG. 16  is a sectional illustration taken along the line  16 - 16  in  FIG. 15 . 
         FIGS. 17 a  through 17 c    illustrate examples of perforation geometries that may be employed in the perforated interlayer shown in  FIGS. 15 and 16 . 
         FIG. 18  is a plan illustration of another embodiment of the composite laminate structure having a interlayer reinforced with a net. 
         FIG. 19  is a sectional illustration taken along the line  19 - 19  in  FIG. 18 . 
         FIG. 20  is enlarged, fragmentary illustration of another embodiment of the composite laminate structure in which a damping interlayer is reinforced with particles. 
         FIG. 21  is a diagrammatic illustration of apparatus for transferring a reinforced film onto a pre-preg used in fabricating composite laminate structures having damping interlayers. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  illustrates a damped composite laminate structure  10  comprising first and second layers  12 ,  14  respectively, and an interlayer  16  disposed between and co-cured to the first and second layers  12 ,  14 . Layers  12 ,  14  may each comprise a plurality of plies of a reinforced synthetic material, such as a carbon fiber reinforced epoxy resin and carbon fiber reinforced plastic material. The interlayer  16  may include a reinforcement  17 . The reinforcement  17  may be a woven or a knitted fabric comprising continuous fibrous strands in the form of yarn, tow, roving, tape or resin, impregnated with a viscoelastic material. The reinforcement  17  may also comprise a second viscoelastic material, in fiber form. The reinforcing fibers forming reinforcement  17  may have a direction of orientation in which all of the fibers in an individual layer extend parallel to each other, and the direction of orientation of adjacent layers have differing angles so as to improve the mechanical characteristics, and particularly the stiffness of the laminate structure  10 . 
     The interlayer  16  may be formed of a material that is relatively soft, compared to the first and second layers  12 ,  14 , such as, without limitation, a viscoelastic material (VEM). VEMs encompass a variety of material classified as thermoplastics, thermoplastic elastomers or thermosets. The VEM should have a high loss tangent, or ratio of loss modulus to storage modulus, in order to provide the laminate structure  10  with damping properties. The glass transition temperature (Tg) of the VEM material should be below the operating temperature, such that the VEM is operating in its soft transition phase. Tg is the approximate midpoint of the temperature range of which glass transition takes place, and is the temperature at which increase molecular mobility results in significant changes in the property of a cured resin system. Generally, polymers may be less than usefully ductile or soft below their glass transition temperature, but can undergo large elastic/plastic deformation above this temperature. 
     The VEM may have a modulus that is approximately 2 or more orders of magnitude less than the modulus of the resin used in the plies of the first and second layers  12 ,  14 . As a result of the relative softness of the VEM forming the interlayer  16 , the interlayer  16  may be made relatively thin, but yet remains effective at very cold temperatures, resulting in a weight-efficient design. More particularly, the relative softness of the interlayer  16  allows the first and second layers  12 ,  14  to move relative to each other in their respective planes, which strains the VEM in the interlayer  16  in shear. The shear strain in the VEM within the interlayer  16 , along with its high loss tangent property, allows the laminate structure  10  to dissipate energy from shock, vibration and acoustic excitation. The reinforcement  17  reinforces the interlayer  16  so that mechanical properties, such as stiffness, of the laminate structure  10  are not diminished by the presence of the relatively soft VEM in the interlayer  16 . 
     The damping action of the laminate structure  10  arises from a phase lag between the applied stress and strain response of the VEM. The damping or loss tangent is the phase angle between the stress and strain, which is an inherent material property. The phase lag is a result of the relaxation of the long chain-like molecules. Damping or relaxation decreases with higher pre-load (static) but increases with larger (dynamic) alternating stress. In designing the laminate structure  10 , it is desirable to increase the strain in the VEM within the interlayer  16 . The shear strain in the VEM may be optimized based on its location in the carbon epoxy laminate structure  10 . The strain can also be increased using local inclusions such as, without limitation, particles or chopped carbon fibers. These inclusions increase the strain in the polymer interlayer  16 , thereby increasing the energy dissipation action within the laminate structure  10 . 
     Another embodiment of the laminate structure  10   a  is shown in  FIG. 2 , which has an interlayer  16  that may be formed of an open weave net  19  or cloth of VEM fibers or strips having a glass transition temperature Tg that provides sufficient stiffness at the full range of operating temperatures of the aircraft, yet which provides high damping when placed in shear 
     The VEM  19  net is impregnated with a VEM resin having a relatively low Tg so that the VEM matrix surrounding the VEM net  19  remains relatively soft at the full range of the aircraft&#39;s operating temperatures. The VEM matrix may comprise, for example, without limitation, a thermoplastic or thermoplastic elastomer with a low Tg and high loss tangent, and the VEM net  19  may comprise a thermoplastic polyurethane or other synthetic fiber cloth that is impregnated with the VEM. 
     In the embodiment shown in  FIG. 2 , optional barrier layers  20 ,  22  are formed, respectively between the interlayer  16 , and the first and second layers  12 ,  14 . The barrier layers  20 ,  22  may comprise a material such as, without limitation, another thermoplastic, or nylon fabric (Cerex)) that is chemically and thermally compatible with the epoxy resin. The barrier layers  20 ,  22  function to limit the migration of VEM in the interlayer  16  and epoxy resin in layers  12 ,  14  so that these two materials are separated and prevented from mixing together. Mixing the VEM and epoxy resin may reduce the damping properties of the interlayer  16 . In one embodiment providing satisfactory results, the barrier layers  20 ,  22  may be between 0.0005 inches to 0.002 inches thick. The barrier layers  20 ,  22  may also function to make the VEM film more suitable to be dispensed using an automated tape laying machine. Each of the barrier layers  20 ,  22  is relatively stiff so as to allow VEM film to be peeled off of a roll when used in automated fiber placement manufacturing using a Multi-Head Tape Layer (MHTL) Machine. 
       FIG. 3  depicts another embodiment of the laminate structure  10   b  in which the interlayer  16  is formed from a woven or knitted cloth  21  of carbon fibers where the fiber strands are alternately arranged in a cross-ply (i.e. 0/90°) or angle-ply (+θ/−θ) configuration. The carbon fiber cloth  21  is impregnated with a low Tg VEM. The VEM may comprise a film of material such as thermoplastic polyurethane or other resin matrix which is hot pressed onto the carbon fiber cloth  21 . 
     A further embodiment of the laminate structure  10   c  is shown in  FIGS. 4 and 5  in which the interlayer  16  is formed of unidirectional carbon fiber tows  30  which are coated with a VEM  32 . As shown in  FIG. 6 , the carbon fibers within the tow  30  may be completely wetted with the VEM  32 . Glass fibers may be substituted for the carbon fiber tows  30 , depending on the application. In the embodiment shown in  FIGS. 4 and 5 , the carbon or glass fibers  30  provide the required mechanical stiffness and strength for the interlayer  16 , while the VEM coating  32  on the fibers  30  provides the desired damping. Because the damping mechanism provided by the VEM material  32  is largely from extension, rather than shear in the embodiment of  FIGS. 4 and 5 , the interlayer  16  may be placed at various locations within the laminate structure  10   c . For example, where the layers  12 ,  14  each comprise multiple plies of composite material, the interlayer  16  may be disposed between any of the plies in either the layers  12  or the layers  14 , or both. More than one interlayer  16  be used, depending on the application, and these multiple interlayers  16  be positioned next to each other or between any of the plies within layers  12 ,  14 . 
     A further embodiment  10   d  is shown in  FIGS. 6-12 , in which the interlayer  16  is formed by a plurality of Z-fibers  34  (thru the thickness fibers) held within a VEM matrix  43 . Fibers  34  are referred to as “Z” fibers due to their inserted orientation in what is conventionally the geometrical Z-direction, perpendicular to the plane of the layers  12 ,  14 . Each of the Z-fibers  34  comprises a tow  37  of reinforcing fibers such as glass or carbon fibers, having ends  39 ,  41  that fan out as individual fibers oriented perpendicular to the main body of the tow  37 . As can be seen in  FIG. 7 , the tow body  37  extends generally perpendicular to layers  12 ,  14 , and the individual fiber strands on the ends  39 ,  41  are respectively co-cured with laminate layers  12 ,  14 . 
     The Z-fibers  34  are introduced into the VEM matrix  43 , which can be a film, with known insertion methods such that their ends  39 ,  41  extend beyond both sides of the VEM  43 . As best seen on  FIG. 7 , the ends  39 ,  41  of the fiber tows  37  anchor the fibers  34  to and/or within the stiffer materials of the layers  12 ,  14  on both sides of the VEM  43  in order to transfer loads through the “Z” direction  40   a . Thus, the space between the Z-fibers  34  is occupied with VEM material  43  which provides the interlayer  16  with the necessary damping qualities. The Z-fibers  34  effectively mechanically connect laminate layers  12 ,  14 , thereby providing the interlayer  16  with the necessary rigidity, and increasing the bending stiffness of the interlayer  16 . 
     As shown in  FIG. 9 , the interlayer  16  may be prepared by inserting the Z-fibers  34  into a film  43  of the VEM, using conventional inserting equipment. With the Z-fibers  34  having been pre-inserted into the film  43 , the film  43  is then placed in a lay-up  45 , between the layers  12 ,  14 , as shown in  FIG. 10 . The lay-up  45  is then compacted and cured at elevated temperature using conventional techniques. 
     The Z-fibers  34  can be arranged in various lay-outs within the interlayer  16 . For example,  FIG. 11  shows an aircraft skin section  44  which includes an interlayer  16  of VEM  43 . Z-fibers  34  are inserted into the VEM layer  44 , around the perimeter of the VEM film  43 . The Z-fibers  34  may also be inserted in a uniform pattern over the interlayer  16 , as illustrated by the matrix lay-out of Z-fibers  34  shown in  FIG. 12 . 
     A further embodiment of the composite laminate structure  10   e  is shown in  FIGS. 13 and 14 . An aircraft skin section  46  includes an interlayer  16  patch comprising a strip of slit tape  50  of reinforcing material, such as carbon fiber reinforced epoxy. The tape  50  is disposed within a VEM matrix  48 . The interlayer  16  is referred to as a “patch” because the width of the interlayer  16  is less than the width of the skin section  46 , and the length of the interlayer  16  is less than the length of the skin section  46 . The interlayer  16  is wholly disposed between a plurality of plies  54 . The outer surfaces of the plies  54  are covered with a layer  52  of carbon fiber reinforced epoxy impregnated cloth. 
     Attention is now directed to  FIGS. 15 and 16  which illustrate another embodiment of a composite laminate structure  10   f , such as a fuselage skin section  64 , in which the interlayer  16  is formed by a film  60  of a suitable VEM in which a plurality of perforations  58  are formed that extend between laminate layers  12 ,  14 . The film  60  may comprise, for example a viscoelastic rubber such as that identified by the trade name SMACTANE® available from SMAC in Toulon, France. The number and size of the perforations  58  will vary depending upon the particular application. The perforations  58 , which pass completely through the interlayer  16 , allow the migration of resin between the layers  12 ,  14  which, when cured, form rigid connections between layers  12 ,  14  that are surrounded by the VEM film matrix  60 . The direct connection between layers  12 ,  14  provided by the resin that fills the perforations  58  reduces the possibility that laminate structure  10   f  may behave as a split laminate when the interlayer  16  is too soft. 
     The perforations  58  may be laid out randomly or in a uniform pattern across the interlayer  16 . The perforations  58  may have any of a variety of cross sectional geometries. For example, the cross sectional shape of the perforations  58  may be round as shown in  FIG. 17 a   , elongate as shown in  FIG. 17 b    or square as shown in  FIG. 17 c   , or a combination of one or more of these or other geometries. 
       FIGS. 18 and 19  illustrate another embodiment of the composite laminate structure  10   g , comprising a skin section  66 . The skin section  66  includes an interlayer  16  comprising a single layer VEM net  68  impregnated with a VEM resin  70 , generally similar to the laminate structure  10   a  in  FIG. 2 . The glass transition temperature Tg of the VEM net  68  is higher than that of the VEM resin  70  so that, over the full operating range of the aircraft, the VEM net  68  provides adequate stiffness and the VEM resin  70  remains relatively soft. In this embodiment, the interlayer  16  is wholly surrounded by the layers  12 ,  14  of laminate plies so as to be encapsulated, and therefore form a damping patch within the skin section  66 . 
     In the case of each of the laminate structures  10 - 10   g  described above, the interlayer  16  is assembled in a lay-up with the first and second layers  12 ,  14 , and are co-cured using conventional techniques, such as vacuum bagging or autoclaving, so the interlayer  16  becomes co-cured to the first and second layers  16 ,  18 , producing a consolidated laminated structure  10 - 10   g.    
     Other variations of the damped laminate structures discussed above are possible. For example, as shown in  FIG. 20 , the interlayer  16  containing VEM matrix material  43  may be reinforced by mixing relatively stiff material into the VEM material  43  This reinforcing material may be micro (meter) sized particles  77  of chopped carbon or ceramic micro-balloons. Also, the particles  77  can be nano (meter) sized using multi-walled and single-walled nano-tubes or nano-fibers. These particles  77  or inclusions may be mixed into the damping polymer when it is still in its aqueous phase (before being formed into a thin film.) The micro-meter sized particles  77  are much stiffer than the VEM  43  and when dispersed into the VEM  43 , the combination of the two materials (thru a Rule of Mixtures) is stiffer and stronger than the neat VEM  43 , i.e., a VEM  43  not containing any reinforcing materials. The nano-sized particles  77  function largely on the atomic level of the molecules, and help increase the strength of ionic bond between molecules which increases the strength of the bond between the VEM  43  and carbon epoxy layers  12 ,  14 . 
       FIG. 21  illustrates an apparatus for forming a pre-preg of a fiber reinforced epoxy resin matrix  78  and a VEM film  74 . The VEM film  74  is fed from a continuous roll  76  along with a pre-preg  78  of a fiber reinforced epoxy resin material to a heating element  80 . The heating element  80  preheats the pre-preg  78  and film  74  which are then passed through consolidating rollers  82  that bond the film  74  to the pre-preg  78 . Release paper  84  is fed from a continuous roll  86  onto the surface of the pre-preg  78 , and the resulting, final pre-preg  88  is accumulated on a roll  90 . 
     Although the embodiments of this disclosure have been described with respect to certain exemplary embodiments, it is to be understood that the specific embodiments are for purposes of illustration and not limitation, as other variations will occur to those of skill in the art.