Patent Description:
Composite materials are becoming more widely used in aircraft due to their strength and weight characteristics.

The use of composite materials may also present challenges. For example, in flight, aircraft can be struck by lightning. To safely dissipate electrical energy associated with a lightning strike, aircraft structures are often designed to distribute electrical current throughout the aircraft in a safe manner. Many composite materials are non-conductive or have relatively low electrical conductivity, which can pose challenges for safe electrical current dissipation during a lightning strike.

<CIT> describes an aircraft assembly which includes a skin in which a CFRP layer serves as a main structure, a shear-tie that supports the skin from the inside, and a fastener that couples the skin and the shear-tie. A copper foil and an outside GRFP layer are provided on an outer surface side of the skin, in this order towards the outside, and a copper paint layer, which contains copper powder, is provided on the outside GFRP layer.

<CIT> describes a composite article having a body of at least one layer of glass fibers which includes a pair of opposed outer bearing portions connected with the body, at least one outer bearing portion including ingredients of graphite, carbon or their mixtures, such as in fiber form, to provide improved bearing characteristics. The article is bonded together through a polymer resin. In one form, the article includes a low coefficient of friction material at least as a coating on the bearing portion, especially for very thin, lightweight bearing designs.

<CIT> describes a method of forming a mobile device case. The method according to the present invention includes impregnating at least one layer of fiber fabric with a thermoplastic resin or laminating thermoplastic resin sheets, and nip pressing them. After thinning, optionally, a natural or artificial decorative layer is laminated on the upper surface with an adhesive layer interposed, preheated, and then placed in a heated mold to close the mold. It consists of pressing, cooling the mold, and then opening the mold. <CIT> describes a composite product and forming method. Multiple layers of fiber, prepreg sheets are employed with an internal air channeling sheet to create a composite or laminate product. A further aspect of <CIT> uses compression molding to form composite parts made from prepreg sheets.

The present invention relates to a method of manufacturing an electrical isolation component for insertion between components of an aircraft structural assembly, and an electrical isolation component for insertion between components of an aircraft structural assembly, as defined in the independent claims. Preferred embodiments of the invention are laid down in the dependent claims.

The method of manufacturing an electrical isolation component for insertion between components of an aircraft structural assembly according to the invention comprises: overlaying an outer ply comprising glass fibers on a core preform comprising reinforcing fibers and a thermoplastic matrix material; applying pressure to the outer ply and core preform, thereby causing consolidation of the reinforcing fibers and matrix material; removing the pressure to allow the thermoplastic material to harden, thereby forming a laminate structure from the outer ply and the consolidated reinforcing fibers and matrix material; wherein said reinforcing fibers and said glass fibers are formed of different materials.

Preferably the method includes cutting a washer from said laminate structure.

Preferably said reinforcing fibers comprise carbon fibers.

Preferably said core and said at least one outer ply are pressed between opposing dies.

Preferably said core and said at least one outer ply are pressed between opposing rollers.

Preferably said core comprises a plurality of core plies, each core ply including a sheet of carbon fiber cloth with a thermoplastic material applied thereto.

Preferably the method includes stacking said at least one outer ply and said plurality of core plies and applying pressure to the stack to consolidate said carbon fiber cloth with said thermoplastic material.

The outer ply comprises an outer ply on a top surface of said core and an outer ply on a bottom surface of said core.

Preferably cutting said washer comprises cutting using a water jet.

The electrical isolation component for insertion between components of an aircraft structural assembly according to the invention comprises: a core formed of reinforcing fibers embedded in a thermoplastic matrix material.

The isolation component includes an electrically-insulating outer ply overlaying said core, said outer ply comprising glass fibers, said reinforcing fibers and said glass fibers formed of different materials.

The outer ply comprises first and second electrically-insulating plies formed of glass fiber, said first and second electrically-insulating plies defining top and bottom surfaces of said isolation component.

Preferably said core comprises a plurality of core plies.

Preferably said core comprises at least three of said core plies.

Preferably each of said core plies comprises woven carbon fabric.

Preferably the weave patterns of said core plies of woven carbon fabric are at non-zero angles to one another.

Preferably an epoxy sealant is applied to at least a portion of a surface of said isolation component.

Preferably said isolation component is a shim.

The present invention relates to electrical isolation components comprising composite materials. In various embodiments, washers, shims and other components are disclosed having composite construction including a core of reinforcing fibers embedded in a thermoplastic matrix material and a fiberglass outer layer covering the core. The components disclosed herein may be used to protect against electrical arcing on aircraft, for example, during lightning strikes.

<FIG> depict a wing torque box <NUM> of an aircraft in top schematic view. <FIG> depicts the wing torque box <NUM> with a skin panel <NUM>. <FIG> depicts the wing torque box <NUM> with the skin removed to show internal spars <NUM> and ribs <NUM>. As shown in <FIG>, wing torque box <NUM> defines a fuel tank <NUM>, bounded by skin panel <NUM>, spars <NUM> and ribs <NUM>. <FIG> depicts the wing torque box <NUM> without the fuel tank <NUM> indicated, to show spars <NUM> and ribs <NUM> and with an enlarged portion showing installed fasteners. Wing torque box <NUM> has structural components such as one or more wing skin panels <NUM> (e.g. upper and lower panels), front and rear spars <NUM> and a plurality of ribs <NUM>. One or more fuel tanks <NUM> may be at least partially defined by skin panels <NUM>, spars <NUM> and ribs <NUM>. Other structural or systems components may be partially or fully enclosed within fuel tank <NUM>.

Skin panels <NUM> are fastened to spars <NUM>, ribs <NUM> using fasteners <NUM>. The wing skin may be defined by a single upper panel <NUM> and a single lower panel <NUM>. Alternatively, the wing skin may include multiple panels <NUM> defining each of the upper and lower skin surfaces. Fasteners <NUM> may be installed along each interface between components, and in particular, where skin panels <NUM> overlie spars <NUM> and ribs <NUM>. Some fasteners <NUM> may extend through skin panels <NUM>, spar <NUM> or ribs <NUM> into fuel tank <NUM>. Other fasteners <NUM> may be wholly contained within fuel tank <NUM>. For simplicity, only four fasteners <NUM> are shown in <FIG>. However, any number of fasteners <NUM> may be present on wing <NUM>.

Components of wing <NUM> may be formed from numerous different materials. Some components, including structural components such as ribs <NUM>, may be metallic; e.g. titanium, aluminum or alloys thereof, while structural components such as skin panels <NUM> , spars <NUM> and ribs <NUM> and other aircraft components may be formed from composite materials, such as carbon-fiber reinforced polymers or other fibre-reinforced polymers. Still other wing components may be formed from materials such as titanium, steel, fiberglass, plastics, or the like.

During operation, wing <NUM> may be subjected to lightning strikes. Lightning strikes introduce high-voltage electrical current, which may propagate and be distributed over the wing <NUM>. Conduction by metal components generally promotes safe distribution of electrical current. However, non-conductive components may interfere with propagation of electrical current, creating voltage differentials across components and a consequent risk of sparks or electrical arcing. As will be apparent, such arcing may pose an ignition hazard, particularly in the vicinity of fuel tank <NUM>.

<FIG> depict a cross-sectional view of a structural assembly including first and second structural members, namely a skin panel <NUM> and a rib <NUM>, held together by a fastener <NUM>. Fastener <NUM> includes a metallic shaft <NUM> such as the shaft of a bolt, rivet or the like. Shaft <NUM> comprises a shaft which may be received in a sleeve <NUM>. Shaft <NUM> extends into the skin panel <NUM> and rib <NUM> (individually and collectively referred to as the substrate), and fastener <NUM> includes and is secured by a metallic nut <NUM> receiving the end of shaft <NUM>. An electric isolation component is positioned between nut <NUM> the substrate. As depicted, the electric isolation component is a washer <NUM>.

In the event of a lightning strike, high-voltage electrical current propagates over the wing. Current may, for example, be conducted across the interface between fastener <NUM> and the substrate and may potentially arc between components. Because of the high voltage and current generated by a lightning strike, current may flow through all available conductive paths. For example, electrical current may flow between fastener <NUM> and skin panel <NUM> (e.g. a partially-conductive composite skin panel <NUM>) through the interface of the fastener shaft <NUM> with the skin panel. Likewise, electrical current may flow though the shaft <NUM> to the nut <NUM> and between the fastener <NUM> and the spar <NUM> or rib <NUM> through which it is received. Differences in electrical potential across components may make the assembly prone to arcing, e.g. between nut <NUM> and rib <NUM>.

Washer <NUM> has a diameter larger than that of nut <NUM>. In some examples, washer <NUM> has an inside diameter of about <NUM> in and an outside diameter of between <NUM> in and <NUM> in and nut <NUM> has a diameter of approximately <NUM> in. Washer <NUM> at least partially insulates nut <NUM> from the substrate (e.g. rib <NUM>). That is, washer <NUM> resists electrical conduction and arcing between rib <NUM> and the face of nut <NUM> that opposes rib <NUM>. Washer <NUM> is sufficiently large to reduce or eliminate the likelihood of electrical arcing between nut <NUM> and shaft <NUM> or spar <NUM>.

Sleeve <NUM> is an expansion sleeve and has an internal bore smaller than the diameter of shaft <NUM>, such that reception of shaft <NUM> into sleeve <NUM> deforms and expands sleeve <NUM> into contact with the substrate, e.g. skin panel <NUM> and rib <NUM>. Such expansion may require a large torque to be applied to shaft <NUM>, thereby generating a large clamping force between shaft <NUM> and nut <NUM>. Nut <NUM> and shaft <NUM> are tightened together to a torque of approximately <NUM> Ibf-in and generate a clamping force of approximately <NUM> lb. In typical aircraft applications fasteners <NUM> are tightened to between <NUM> to <NUM> Ibf-in, however, these loads may vary. Sleeve <NUM> may be omitted and shaft <NUM>, nut <NUM> may directly engage the substrate.

The clamping force causes nut <NUM> to bear against washer <NUM>, transferring the clamping force to the washer <NUM>. In some circumstances, such as in the event of misalignment of shaft <NUM>, nut <NUM> or washer <NUM>, the load on washer <NUM> may be concentrated on a small portion of the washer's surface. For example, <FIG> shows a fastener <NUM> with shaft <NUM> and nut <NUM> misaligned by an angle of <NUM>°.

Shaft <NUM>, sleeve, <NUM> nut <NUM> and rib <NUM> are metallic and may therefore be good electrical conductors. For example, shaft <NUM>, sleeve <NUM> and nut <NUM> may be titanium or an alloy thereof. Rib <NUM> may be aluminum, titanium or an alloy thereof. Skin panel <NUM> is non-metallic and may be a poor electrical conductor. Thus, if an electric charge is applied (e.g. by a lightning strike), electrical arcs may be possible, e.g. between rib <NUM> and nut <NUM>. Washer <NUM> may therefore be constructed to provide resistance against electrical arcing. That is, washer <NUM> may be constructed of an electrically insulating material. However, washer <NUM> may also need to be constructed using materials sufficiently strong to withstand clamping forces associated with installation. Moreover, washer <NUM> may turn or otherwise move during installation and operation, causing washer <NUM> to rub against nut <NUM> and spar <NUM>. Therefore, washer <NUM> may also be constructed to resist abrasion.

Some plastics (e.g. injection-molded thermoplastics) are good electrical insulators. However, such plastics typically lack the requisite strength and toughness to reliably withstand loading by fastener <NUM>. Conversely, metals may be sufficiently strong, but do not provide electrical insulation. Insulating coatings are available, but such coatings are prone to scratching, which may permit conduction of electrical current. Composite materials based on carbon fibers in a thermoset (e.g. epoxy) matrix may be relatively strong, but are very time consuming and expensive to make. For example, such materials typically require extended curing times under specific conditions such as elevated temperature, pressure or the like.

<FIG> depicts a composite washer <NUM> configured for use in aircraft <NUM>. Washer <NUM> has a core <NUM> with one or more core plies <NUM> and a cover with one or more outer plies <NUM>. As described in further detail below, a core ply <NUM> is constructed of a thermoplastic matrix material, with embedded reinforcing fibers. The reinforcing fibers of core plies <NUM> and outer plies <NUM> may be formed of different materials. For example, the composition of core plies <NUM> may provide strength, while outer plies <NUM> may be constructed of an electrically-insulating and abrasion-resistant material such as fiberglass.

<FIG> depicts an example core ply <NUM>. As noted, core ply <NUM> comprises a thermoplastic matrix <NUM> with reinforcing fiber material. The thermoplastic matrix <NUM> is partially removed in <FIG> to show the reinforcing fiber material.

As depicted, the reinforcing material is a carbon fibre sheet <NUM>. The sheet includes a plurality of carbon filaments, which may be grouped into bundles or yarns <NUM>. The yarns may be woven together. <FIG> shows an enlarged partial view of carbon fiber sheet <NUM>, showing the weave pattern of yarns <NUM>. Matrix material <NUM> is distributed on, in and around the yarns <NUM> defining sheet <NUM>. As depicted, yarns <NUM> are woven in a lattice pattern, with first yarns 125a at a first orientation and second yarns 125b at a second orientation, at approximately a <NUM> degree angle to yarns 125a. <FIG> shows a cross-sectional view of yarns <NUM> woven in such a pattern. Each yarn 125b is woven through yarns 125a in an alternating over-under pattern. Such fabric may be isotropic or quasi-isotropic. Other arrangements are possible. For example, yarns <NUM> may have the same orientation, such that they are parallel to one another, or yarns <NUM> may have different orientations at angles other than <NUM> degrees to one another, or yarns <NUM> may have random orientations.

As depicted, yarns <NUM> of carbon fabric <NUM> are made up of approximately <NUM> filaments. Such fabric may be referred to as <NUM> fabric. In an example, the filament thickness is approximately <NUM> microns and fabric <NUM> defines a ply thickness of about <NUM>. However, the filament thickness may differ therefrom. Bundles of filaments may define yarns of thickness dependent on bundle shape, weave pattern, etc..

Other fabrics may be used. For example, fabrics with more or fewer filaments per bundle or with differently-sized filaments may be used. Alternatively, carbon filaments may be distributed randomly or in a single bundle, rather than in discrete yarns. The chosen fabric may be selected based on specifications such as strength, thickness and weight in view of the loads to which washer <NUM> is expected to be subjected.

The carbon fibers of core <NUM> are embedded in a thermoplastic matrix material <NUM>. Matrix material <NUM> may be solid at room temperature and atmospheric pressure, but may be flowable when warmed or subjected to high pressure. After reaching a flowable state, matrix <NUM> will cool and harden after removal of the high temperature or pressure.

The matrix material may, for example, be poly etherether ketone (PEEK), poly ether ketone ketone (PEKK), polyphenylene sulfide (PPS), Polyphenylene sulfone (PPSU), polyamides such as polyphthalamide (PPA) or the like. Matrix material <NUM> may be applied to the reinforcing fibers by powder-coating or by forcing a liquid containing the matrix material through the reinforcing fibers. Matrix material may be applied to individual filaments or to individual yarns <NUM> prior to weaving into sheet <NUM>. Alternatively, a woven sheet <NUM> may be power-coated or liquid coated after weaving.

Alternatively, rather than applying matrix material by powder or liquid coating, filaments of matrix material may be commingled with reinforcing fiber filaments. For example, each yarn <NUM> may have approximately <NUM> filaments, of which a proportion are filaments of matrix material.

Matrix material <NUM> and reinforcing fibers are consolidated together to form a composite structure. That is, after matrix material <NUM> is distributed with reinforcing fibers, the matrix material <NUM> is at least partially melted by application of heat, pressure or both, so that the matrix material <NUM> flows together to form a structure with the reinforcing fibers.

Matrix material <NUM> and reinforcing fibers may be present in core ply <NUM> in a volume fraction of about <NUM>% reinforcing fibers. In some examples, reinforcing material may form approximately <NUM>% +/- <NUM>%.

In some embodiments, washer <NUM> may have a core <NUM> with multiple layers, which may be referred to as plies. For example, as depicted in <FIG>, washer <NUM> has three core plies <NUM>-<NUM>, <NUM>-<NUM>, <NUM>-<NUM>, each of which includes a carbon fabric sheet <NUM> and matrix material <NUM>. The core plies <NUM>-<NUM>, <NUM>-<NUM>, <NUM>-<NUM> may have different orientations, such that their yarns <NUM> are positioned at an angle to one another. Alternatively or additionally, core plies <NUM>-<NUM>, <NUM>-<NUM>, <NUM>-<NUM> may have different weave patterns. For example, one ply <NUM>-<NUM> may have yarns <NUM> in a lattice pattern as shown in <FIG>, while another ply <NUM>-<NUM> has yarns <NUM> that are parallel to one another, and third ply <NUM>-<NUM> has filaments at random orientations.

For purposes of illustration, carbon fabric <NUM> and matrix material <NUM> are shown in <FIG> as discrete layers with exaggerated thickness. However, matrix material <NUM> may be distributed through and around carbon fabric <NUM> to define a single ply of mixed composition.

Referring to <FIG>, outer ply <NUM> is formed of an electrically insulating material. The material of outer ply <NUM> may also be selected to provide abrasion resistance or to limit friction between washer <NUM> and the adjacent fastener or substrate material. As depicted, outer ply <NUM> is formed of woven glass fibres, e.g. glass cloth such as EGlass <NUM> HS 105GSM. Outer ply <NUM> may be thin relative to core <NUM>. In an example, each outer ply is approximately <NUM> thick and each core ply <NUM> is approximately <NUM> thick.

As depicted in <FIG>, washer <NUM> has a cover including two outer plies <NUM>, one on each of the top and bottom faces of the washer. Alternatively, washer <NUM> may have more or fewer outer plies <NUM>. For example, an outer ply <NUM> may be provided on only one of the top and bottom surfaces of washer <NUM>, as shown in <FIG>. Alternatively, multiple outer plies <NUM> may be stacked on the top and bottom surfaces of washer <NUM>, as shown in <FIG>, or outer plies <NUM> may be located between core plies <NUM> as shown in <FIG>.

Each outer ply <NUM> may provide resistance against conduction of electrical current. Thus, conduction or arcing of electricity across washer <NUM> may be prevented, up to a breakdown voltage. That is, arcs may be prevented unless a voltage greater than or equal to the breakdown voltage is applied across the washer <NUM>. The breakdown voltage may depend, for example, on the diameter and width of washer <NUM>, the diameter of nut <NUM> in contact with washer <NUM>, the thickness of each core ply <NUM> and the total thickness of the core, and the thickness and number of outer plies <NUM>. The breakdown voltage may be increased, for example, by increasing the outer diameter of washer <NUM> relative to the bearing surface of nut <NUM>, increasing the number or thickness of outer plies <NUM>, applying non-conductive surface coatings or sealing washer <NUM> in a sealant. Conversely, thinner outer plies <NUM> or a smaller number of outer plies may decrease the breakdown voltage.

Washer <NUM> may be coated with a sealant such as an epoxy lacquer. The sealant may be applied to the edges of washer <NUM> and its top and bottom surfaces.

<FIG> is a flow chart showing an example process <NUM> of manufacturing a washer <NUM>. At block <NUM>, one or more sheets of prepreg material is formed. As noted, the prepreg may be formed by powder coating thermoplastic matrix material onto or forcing liquid matrix material through a sheet of carbon fiber (e.g. a sheet of woven carbon fiber fabric).

At block <NUM> one or more plies of cover material (e.g. fiberglass cloth) is formed.

At block <NUM>, the plies of cover material and prepreg material are stacked. The plies are sheets may be laid flat atop one another as shown in <FIG>. As shown in FIG. , core plies <NUM> may be oriented such that carbon filaments of core plies <NUM> extend at an angle to one another. For example, as shown in <FIG>, core ply <NUM>-<NUM> is oriented at approximately a <NUM> degree angle to core ply <NUM>-<NUM>, and core ply <NUM>-<NUM> is oriented at approximately a <NUM> degree angle to both core plies <NUM>-<NUM> and <NUM>-<NUM>. Likewise, outer plies <NUM> may be oriented such that glass fibers are at an angle to one another and to carbon filaments of the adjacent core ply <NUM>.

At block <NUM>, core plies <NUM> and outer plies <NUM> are pressed together. The plies <NUM> and plies <NUM> may be pressed together in a die press. For example, as shown in <FIG>, the stack of core plies <NUM> and outer plies <NUM> are placed between dies <NUM> of a press <NUM>. Press <NUM> is closed to compress the stack between dies <NUM> as shown in <FIG>. Pressure exerted by the dies causes matrix material <NUM> to at least partially melt into a flowable condition. Once in such a state, matrix material <NUM> flows into and around the reinforcing fibers of core plies <NUM> and adheres to outer plies <NUM>. Press <NUM> may be held in the closed position of <FIG> for a consolidation period, sufficient to allow flowing of matrix material <NUM> to fully consolidate core <NUM>. In some embodiments, pressing may occur at temperature. In other embodiments, plies <NUM>, <NUM> or the pressing device may be heated prior to pressing. Such heating may reduce the amount of pressure required to cause flowing of matrix material <NUM>. In some examples, the consolidation period is between approximately <NUM>-<NUM> minutes at a temperature of about <NUM>-<NUM> and a pressure of about <NUM>-<NUM> psi.

At block <NUM>, after pressing of plies <NUM>, <NUM> and consequent flowing of thermoplastic matrix material, pressure may be partially or fully released. For example, as shown in <FIG>, press <NUM> is opened to release pressure. Plies <NUM>, <NUM>, including matrix material <NUM> may cool, causing the matrix material to return to a solid state. The solidified matrix material <NUM> holds core plies <NUM> together to define a washer core <NUM> and binds outer plies <NUM> to the core <NUM>. Plies <NUM>, <NUM> then define an insulation sheet <NUM>. In some examples, plies <NUM>, <NUM> are cooled to room temperature at a rate of between <NUM>-<NUM> per minute.

At block <NUM>, insulation sheet <NUM> is removed from the pressing device (e.g. by separating die halves and lifting the sheet <NUM>). A plurality of washers <NUM> are cut from insulation sheet <NUM>. The washers may be cut, for example, using water jets or another suitable cutting technique.

At block <NUM>, the cut parts are finished by application of a sealant such as an epoxy glaze. The sealant may be electrically insulating and may be applied at least to the edges of washers <NUM>. In some embodiments, sealant may also be applied to the top and bottom surfaces of washers <NUM>.

In some embodiments, rather than being pressed between dies of a press, prepreg and cover material may be dispensed from rolls <NUM> and compressed in a stack between rollers <NUM> as shown in <FIG>. Matrix material <NUM> may at least partially melt and flow during compression by rollers <NUM>. After exiting rollers <NUM>, matrix material <NUM> cools and hardens, and the consolidated material may then be cut into insulation sheets <NUM>.

As described above, electrically isolating washers are formed of composite material including one or more core plies of carbon fiber embedded in a thermoplastic matrix, and insulative glass fiber outer plies. However, in other embodiments, such material may be used to form other aircraft components. For example, such material may be used for relatively low-cost and lightweight washers or other fastener components that are not installed for electrical isolation. Additionally, such material may be used for bushings, sleeves, or the like. Moreover, such material may be used for other components which occupy space between aircraft structures for electrical isolation of the aircraft structure. For example, material as disclosed herein may be used to form shims to occupy space between and electrically isolate aircraft skin panels and structural members such as a wing spar or rib, or between adjacent structural members. <FIG> shows an example of such an embodiment. As depicted, a shim <NUM>' is interposed between a skin panel <NUM> and spar <NUM> to occupy space and to electrically isolate skin panel <NUM>-<NUM> from spar <NUM>. Shim <NUM>' has a composite structure similar to that of washer <NUM>. Specifically, shim <NUM>' is a laminate including two outer layers <NUM>' of a material comprising glass fibers, and a core comprising core plies <NUM>' of reinforcing fibers, such as carbon fibers, embedded in a thermoplastic matrix. Shim <NUM>' may be used, for example, to provide electrical isolation between structural components proximate or associated with fuel tank <NUM>, particularly between skin panels <NUM> and spars <NUM> or ribs <NUM> and between spars <NUM> and ribs <NUM>.

Claim 1:
A method of manufacturing an electrical isolation component (<NUM>, <NUM>') for insertion between components of an aircraft structural assembly, comprising:
overlaying an electrically-insulating outer ply (<NUM>) comprising glass fibers on a core preform (<NUM>) comprising reinforcing fibers (<NUM>) and a thermoplastic matrix material (<NUM>), wherein said outer ply (<NUM>) comprises an outer ply (<NUM>) on a top surface of said core (<NUM>) and an outer ply (<NUM>) on a bottom surface of said core (<NUM>);
applying pressure to said outer ply (<NUM>) and core preform (<NUM>), thereby causing consolidation of said reinforcing fibers and matrix material;
removing said pressure to allow said thermoplastic material to harden, thereby forming a laminate structure (<NUM>) from the outer ply (<NUM>) and the consolidated reinforcing fibers (<NUM>) and matrix material (<NUM>);
wherein said reinforcing fibers (<NUM>) and said glass fibers are formed of different materials.