| Tutorials home | Decreasing risk exposure | Safety tour | Emergencies | Meteorology | Flight Theory | Navigation | Communications |

Tutorials home page

Builders guide to safe aircraft materials

Reinforcing fibres and composites



Rev. 6 — page content was last changed 27 September 2010
  
Google logo


      Content


Thermosets are low in tensile strength and, as such, have little application as a sole material forming a structural element. However, if thermosets are reinforced with high tensile strength fibres then the resulting composite materials can be formed into ideal aircraft structural elements or even complete monolithic structures. The fibres used in light homebuilt airframes are usually glass, perhaps with some carbon fibre materials purchased as finished tubing or other profiles. There are other fibres in the aromatic polyamide (aramid) group, which may have some airframe application when combined in a hybrid laminate with glass.
17.1 Reinforcing fibres
Glass fibres are small diameter (perhaps 3–8 microns) continuous filaments of glass. They are sometimes produced from glass marbles, in much the same process as that described for polyester fibres, then bundles of 200 to perhaps 800 filaments are formed into twisted or untwisted yarns that are finally woven into fibreglass cloths and tapes.

(Glass fibre is the backbone material of the world wide composites industry; about 90% of all production is based on fibreglass.)

Smaller diameter filaments are better because a larger surface area is provided for the interface bond with the composite matrix. Take, for example, a filament with a diameter of 20 microns and a cross-section area of 314 square microns. That could be replaced with 25 filaments of 4 microns diameter and cross-section area of 12.56 square microns each. The total surface area of 25 such filaments, of one metre length, would be five times that of the single 20 micron filament.

There are several types of basic material used for fibreglass. That commonly used in aircraft applications is E-glass, which was originally developed for single and multi layer 'printed circuit board' applications where sheets of non-conducting epoxy-fibreglass with copper foil laminated to one or both surfaces were (and still are) the structural basis for electrical/electronic component assemblies or packages. Another type, S-glass, was developed particularly for structural applications. It has a tensile strength perhaps 25% greater than E-glass and has been used in such applications as helicopter blades. But in terms of value for money, E-glass doesn't have much application in homebuilt aircraft.

Carbon or graphite fibre is generally manufactured from a polymer (polyacrylonitrile) in complex heating/reheating processes producing light-weight filaments with high strength and high stiffness. The glossy black carbon filaments are generally made in one thickness but formed into yarns, which are classified by the number (1000s) of filaments in the yarn — 3K, 6K, 12K, etc. Rovings contain perhaps 50 000 filaments.

Aromatic polyamide or aramid (or para-aramid) fibres are commonly associated with the DuPont trade name Kevlar and renowned for their light weight and very high impact strength. Carbon and aramid yarns are formed into cloths and tapes in the same manner as fibreglass.

There is great variation in the physical properties, cost and availability of the many types and grades of glass, aramid and carbon reinforcing fibres currently manufactured around the world. The following is a rough indicative summation relating to fibres only; the 'SM carbon' indicated is the 'standard modulus' grade fibre. But be aware that the properties of the composite materials incorporating these fibres are substantially different, as shown in the composite properties table.

If you are unfamiliar with the terms used in the table read stress and strain — definition and application in design in the module 'Properties of metals'; the same mechanical properties apply to fibres and composites.

Simplified comparison of fibres
PropertyE-glassAramidSM carbon
Relative cost of yarn1510
Density [g/cm³]2.51.41.8
Modulus of elasticity [GPa]70100210
Tensile strength [MPa]240030004000
Strain at break point*4.5%2%1.2%
Impact strengthbetterbestfair
Fatigue resistancegoodbetterbest

*Note the very low strain at break point or elongation figure for all fibres compared to that for metals used in aircraft structures. The 1.2% figure for carbon fibre indicates that though very strong, it is also very brittle and will break shortly after the load stress passes yield point. Also carbon fibre aircraft may be prone to impact damage in ground handling incidents so an outer aramid layer or full laminate should perhaps be added to leading edges, wing tips and other areas likely to suffer collision damage.


17.2 Cloths, tapes and rovings
The manner in which the fabric yarns are woven into cloth commonly includes:
Plain weave fabric Plain weave, which doesn't distort while handling and is suitable for large areas but the high degree of crimp reduces smoothness and mechanical properties compared to twill or satin. Plain weave is more likely to be used in inner surfaces.
Twill weave fabric In twill weave each weft yarn passes alternately over two warp yarns then under the next two, thus having less crimp than plain weave and consequently being a little smoother. Twill weave has better mechanical properties than the plain and is likely to be used in external surfaces.
Satin weave fabric In satin weave a weft yarn passes under one warp then over the next three (or more) producing a loose weave suitable for draping and holding complex shapes such as a wing root fillet. Satin weave fabric, where the fill passes over one warp thread and then under three, is called 'four-harness satin' [4H-satin]. Or if it passes over one then under the next seven it is called 'eight-harness satin' [8H-satin]. Satin weaves have very low crimp and thus high smoothness, drapability and mechanical properties. But the reduction in the interlocking of warp and fill greatly increases the possibility of distortion in handling.


The cloths generally used in light aircraft construction and woven from glass fibre yarns have thicknesses ranging from perhaps 0.1 mm (0.004 inches) to 0.3 mm (0.012 inches) and weigh from 100 g/m² (3 ounces/square yard) to 300 g/m² (10 ounces/square yard).

Reinforcing fibres provide high tensile strength so the lay of the fibres in the cloth — and consequently the laminated composite — is important in achieving the required load carrying capacities without including any 'parasitic' fibres; i.e. fibres that are not called upon to do useful work and just add unnecessary weight. The tensile strength of fabric may be required to be bidirectional (i.e. much the same along the warp [0° orientation] or across the warp [90°]) or unidirectional where the fill yarns (regarded as parasitic) are few and of lighter material; their purpose is just to hold the higher-strength warp straight during the lay-up process.

The bidirectional or biaxial cloth is often referred to as bid and the unidirectional as uni cloth. In a bid cloth weighing 250 g/m², the warp and the fill yarns would each weigh about 125 g/m². In uni cloth the warp yarns would comprise more than 90% of the total cloth weight. Bid is used for skins, shear webs, bulkheads and similar elements. Uni is used for spar caps, flanges, undercarriage legs and other members subject to high bending loads.

Fabrics are also made up as double bias cloth with yarns of equal weight laid at +45° and -45°. There are also quite a few expensive multidirectional combinations produced as multilayer non-woven fabrics — the layers are just stitched or chemically bound together to facilitate handling and there is no crimp within the layers. However there is no economic justification for the use of such fabrics by homebuilders particularly in hand lay-up. See the product range at www.colan.com.au.

 Fibreglass  tow or roving Fibres are also produced as woven unidirectional tapes of various widths and thicknesses, and also as rovings which are flat bundles or tows of perhaps 50 000 filaments or more. Fibreglass filaments are thicker than those used for yarns. Rovings may be woven into a cloth that is much heavier than yarn-woven cloth, but the latter has a better strength/weight ratio.

Reinforcing mat or core mat consists of continuous filament swirls or randomly placed chopped yarn pieces perhaps 50 mm long, lightly held together with a binder and supplied in blanket form. Such mat is usually thicker than woven cloth and may be used as the core of a laminate with outer plies of woven cloth — much the same as plywood fabrication. Polyester is the usual binder so mat material may not be used with epoxy resin.

Lubricants are applied to the yarns in the weaving process to avoid friction damage to the fibres. The lubricant is normally removed at the mill but, if not, must be removed before the cloth is used in a laminate.

17.3 Structural laminates
The composite materials used in light aircraft structures are glass, carbon or aramid reinforcing fibres encapsulated within a thermosetting polymer matrix. Both fibres and the matrix retain their individual physical and chemical identities but in combination produce structural materials superior in some significant mechanical respects to the traditional airframe materials. The fibres perform the load-carrying task, while the tough, stiff matrix keeps everything in position while transferring and equally distributing the load between fibres — and also protecting them from impact damage, contaminants, moisture, corrosion and elevated temperatures. Epoxy is generally the resin of choice for use with glass, carbon and aramid fibres. The matrix material must bond chemically with the reinforcing fibres, so a coupling agent or sizing compatible with both the fibre and a specified resin(s) is usually applied to the fibres at the mill; a silane agent is generally used to treat fibreglass.

(Note that the term 'advanced composite material' includes any combination of high tensile strength and matrix materials, titanium filaments within copper for example.)

Structural components properly made from such composites have high strength/weight ratios; are resistant to fatigue, deterioration and chemical corrosion; require fewer joints and no weight-adding metal fasteners; and can be fabricated with minimum tooling — but the builder is promoting chemical reactions, so the process may be messy and must be handled with appropriate protection. With thought, great care in design and considerable effort, very smooth, seamless, low-drag airframe surfaces can be achieved for a small flat plate area. That is the big advantage that composite materials offer to scratchbuilders over the traditional materials.

Home-built solid laminates are made up in a wet lay-up process as a series of plies. Each ply starts with a sheet of fibre cloth cut [with scissors] to the required shape and with the required directional lay; for example, the cloth might be cut on the bias. Then just sufficient mixed liquid resin is worked into the cloth with a squeegee or a small paint roller so that all individual filaments are wetted and there is no entrapped air, other voids or excess surface resin. The thickness of the laminate is built up by pressing another layer of cloth onto the wet surface and working in more resin until sufficient plies have been laid to achieve the desired thickness and strength. The lay-up is then left to cure for the required time and temperature after which, if necessary, it can be edge trimmed with a circular saw or band saw. The relative fibre:matrix volume in the lay-up for highest strength and lowest weight (while assuring proper wetting of all fibre surfaces) should be about 60:40; but no less than 50% nor more than 70% of the total volume should be fibre.

(Excessive resin will add a lot of unnecessary weight; excess resin equivalent to one millimetre thickness over the surfaces of an 8 m² wing adds 20 kg to aircraft weight.)

Each ply may use different cloth. For example, the ply that will form the outer face may be twill weave glass for a smoother surface, while the inner plies are unidirectional glass for a better strength/weight ratio — or there could be a mix of carbon and glass plies. The impact resistance of thin aramid-epoxy skins is a bit better than glass-epoxy, but carbon impact resistance is only about 25% of glass. So, in a carbon-epoxy laminate where impact and abrasion resistance is needed, the outer ply might be aramid. The number of plies, the material and the directional lay for each is a matter of balancing loads, weight, surface finish, cost and so on, but just two plies is common for outer skins. However, because of differing mechanical properties, mixing of fibres within the same laminate must be carefully assessed.

In sandwich construction a centre ply is replaced by a thicker, weaker material such as polystyrene foam or a reinforcing mat — the fuselage shell of the Jabiru aircraft is made up of two outer plies and one inner ply, all of bidirectional twill weave glass separated with a centre layer of reinforcing mat.

In the following materials comparison table, the density of the composite laminates is calculated on a basis of 60% fibre to 40% epoxy resin by volume with an epoxy resin density of 1.25g/cm³. The table is predicated on unidirectional cloth so the calculated ratios are more favourable than they would be with bid cloth.

Comparison of composites and traditional airframe materials
PropertyGlass-epoxyAramid-epoxyCarbon-epoxy6061-T6 aluminium4130 N steelHoop pine
Density [g/cm³]
21.41.62.880.5
Modulus of elasticity [GPa]30601307020013
Modulus/density ratio154580252525
Tensile strength [MPa]8001200140030070090
Tensile strength/density ratio40085090011090180


The most significant value in the above table is the high modulus for carbon-epoxy — 65% of that of chrome-molybdenum steel. To make the comparison a little more meaningful I have divided the modulus by the density to produce a 'modulus/density ratio' or specific modulus for all the materials — 6061-T6, 4130 N and hoop pine share much the same value, glass-epoxy is low at 60% of their value, aramid-epoxy is nearly double and carbon-epoxy is three times the traditional materials. This probably indicates why light composite aircraft with strutted wings can use glass-epoxy spars, but unstrutted (cantilever) wings use carbon-epoxy spars to obtain the required rigidity without excessive weight.

The other value worthy of note is the 'tensile strength/density ratios' for all three composites, which are roughly four to eight times better than the traditional airframe materials. It is worth noting the better performance of hoop pine compared to 6061-T6 and 4130N, which is one reason many people favour wooden construction — apart from the thinking that a wood and fabric wing imparts a better feel for what is happening in flight.

All laminates are not particularly abrasion-resistant.


17.4 Forming laminated structures
There are three basic methods for producing laminated aircraft structures. The first is the closed male/female mould process, which involves substantial investment in equipment and very accurate tooling — and thus probably only applicable to larger-scale industry.

The second is the open mould process which is usually confined to series-production light aircraft and kits. The construction of open moulds is complex and a little costly as a male master mould (or master pattern or plug) must be made before a female (or cavity) production mould is produced from that. The master mould might be constructed from wood and the female mould from a composite material. The internal surface of the female mould is an exact pattern for the external surface of the component — a fuselage half-shell, for example. So, if the surface of the female mould is made smooth and properly shaped then the parts made from it will also have a smooth, well-formed surface. Only the female mould is used in the laminate production process, the plies being wet laid-up within it. Additional female production moulds are made from the male master as necessary.

Usually a few very thin coats of a pigmented (probably white) polyester resin compound are sprayed onto the surface of the female mould and given time to partly cure before lay-up starts. This thin gel coat becomes an integral part of the finished laminate, and provides some weather and pollutant protection plus improved surface quality ready for a finishing surface coating. The gel coat is the outermost layer of a structural lamination and is consequently subject to high tension and compression loads as the structure bends or flexes. Consequently the formulation of aircraft gel coats must provide a tough material usually at the expense of high gloss. Gel coats add weight, possibly around 20 kg for a composite aircraft in the 544 kg MTOW class. The coating forms a smoother outer surface but the finished structure will still require some further surface preparation if a UV-resistant primer and top coats are required. Thick (0.3 to 0.8 mm) gel coats containing a UV inhibitor, are not usually top coated but are waxed and polished.

In the closed, and sometimes the open moulding techniques, wet lay-up is not used. The lay-up is done using cloth that has been pre-impregnated with resin by the manufacturer to a precise specification, partly cured to the B stage and delivered to the market in refrigerated transport to avoid the curing progressing beyond B stage. The end user of this material — prepreg — must also keep the material refrigerated until use. After dry lay-up it must be finally laminated under pressure, and typically at temperatures between 120° C and 175° C, to achieve final curing.

The open mould construction process and the use of prepreg is obviously unsuitable for building just one aircraft, so simpler mouldless construction methods are used by homebuilders. Although, in the circumstance where there is an existing part available (a damaged wingtip perhaps) a female mould could be built, even a plaster cast, which in itself is a wet composite lay-up of scrim and plaster.

Foamcore construction. In mouldless construction there are several techniques available, the most common being foamcore construction. This involves the creation of a form or plug hot-wire cut from polystyrene (but not polyurethane due to the hazardous emissions) foam blocks on which the wet plies are laid up. For simpler components the form might also be constructed from sections straight cut from foam sheets and built into the final shape by tack-gluing together plus radiusing the internal/external corners to reduce stress field concentrations in the cured plies. In effect, the form is integral to the finished product, and becomes the core material in a sandwich-type construction. The outside surface will lack the quality achievable in a female mould.

Another variation of simple plug construction is to build a light, collapsible internal wooden frame, lay foam core sheets over that, shape, seal with slurry, apply a couple of coats of varnish and seal with wax. The full structural lay-up is made on that. Whatever method is used, all scrapes, dents and hollows in foam surfaces must be slurry sealed, filled and reworked to the required contours before the first ply is laid.

Some basically rectangular parts, a spar for example, may be simply built-up on a flat surface as a solid laminate.

A peel ply of light polyester cloth may be rolled onto the wet outer surface of the lay-up to help smooth out the surface without shifting the outer ply, and perhaps to wick up patches of excess resin. The peel ply also protects the surface from contamination after lay-up and usually remains attached until the part is ready for surface coating. At this point it is peeled off, leaving a uniform, slightly roughened surface that only requires a light sanding prior to priming and painting. Peel ply is also used to protect surfaces that will subsequently be bonded to another part, or to provide support to resin/cloth contact in difficult positions. Of course, the peel ply must be removed before bonding or pre-painting surface preparation, so a colour-striped cloth is often used as a visual reminder.

Peel ply cloth is available in various weights and porosities, and is treated with a release agent to allow easy detachment. Black peel ply is an aid to post-curing by solar radiation.

A vacuum bagging process, often associated with both open-moulded and foamcore construction, uses the ambient atmospheric pressure to squeeze the plies together and into/onto the form. This removies air and solvent vapours and also squeezes excess resin from the lay-up. In sandwich and foamcore construction the process also aids bonding of the core to the outer layers. The airtight bag (or a sheet sealed around the edges) used may be a clear, flexible plastic film such as PVA or polyethylene laid over the work piece — and the mould or a backing plate — and connected to some form of pump that will draw down a near-vacuum within the bag. To prevent the bag being epoxied to the work piece it is probably necessary to first wrap the work piece in release film, which is a thin, stretchable plastic that is treated so that it won't bond to a particular resin.

Schematic of butt joint and toggled joint In the latter stages of airframe assembly the formed structural components must be bonded together; for example, wing ribs to spars, then wing skins to the rib and spar caps. Most joints tend to be butt joints — as in wood/plywood airframes — so the mating parts must be carefully trimmed to minimise joint gaps. Inside corner fillets with an appropriate radius are formed with suitable filler. The joints are then bonded with an overlay of strips or tapes of reinforcing cloth impregnated with the normal laminating resin; these plies distribute the flight loads.

Butt joints are inappropriate in circumstances such as joining the two halves of a fuselage shell where accurate, fully formed, self-aligning joints extending the length of the fuselage are necessary. Such joints or joggles provide larger surfaces for bonding with structural adhesive, plus the option of temporary clamping to hold the surfaces together while the structural epoxy adhesive is curing. Temporary clamping can be accomplished by drilling through the joggles and inserting clecos. After curing, the clamps are removed and the holes filled.

The sketch shows a simple joggle joint. But a builder might take a 'belt and braces' approach by combining a more complex double joggle with additional external lay-ups of laminating resin and cloth strips or tapes overlapping the seams. The additional reinforcing is laid within a shallow recess formed to allow the external contour to be maintained.

For more information on composites and laminating techniques go to www.fibreglast.com and view the material contained in their 'Learning Center'.


 Deperdussin The first successful 'composite' aircraft was manufactured 100 years ago. This was the Deperdussin Racer which, in 1913, won the 5th James Gordon Bennett trophy race, comprising 20 laps of a 10 km circuit in France, at an average speed of 204 km/h [110 knots]. Each half of the streamlined fuselage shell was formed in a mould from strips of steamed tulipwood glued together to form three cross-ply layers. The half-shells were joined together, covered with fabric and varnished to form the very light monocoque (i.e. single shell with no stringers or longerons) fuselage. This was the first aircraft with a 'stressed' or 'load-bearing' skin, preceding other stressed-skin aircraft by 10 years or so. The all-wood De Havilland Mosquito of World War 2 fame might be considered the ultimate development of Deperdussin's concept.


The next module in this fabrics, composites and coatings group is 'Surface coatings and finishes'

Back to top



Builders guide to aircraft materials – fabrics, composites and coatings modules

| Guide contents | Aircraft fabric covering systems | Plastics and thermosets |

| [Reinforcing fibres and composites] | Surface coatings and finishes |


Copyright © 2006–2010 John Brandon     [contact information]