Aircraft 

While the utilisation of composites, in tonnage terms, for aircraft components constitutes a relatively small percentage of total use, the materials often find their most sophisticated applications in this industry. In aerospace the demands placed upon materials can be greater than in other areas, often requiring a combination of light weight, high strength, high stiffness and good fatigue resistance. Military aircraft were the first to use composites in significant quantities. The first applications were in radomes and then in secondary structures and internal components. The modulus of glass, however, is low compared with that of metals and it was not until the advent of boron and carbon reinforcements that significant interest in terms of primary structures developed. The situation in the present day, where use of composites is extensive, has been the result of a gradual direct substitution of metal components followed by the development of integrated composite designs as confidence has increased. The Airbus 320 has a whole range of components made from composites, including the fin and tailplane This has led to a weight-saving of 800 kg over its equivalent in aluminium alloy. 

The Harrier AV-8B  is an example of a military aircraft using advanced composite materials. The primary structural applications are the wing torque box and control surfaces, horizontal tail and forward fuselage. Secondary structures are the gun and ammunition packs, strakes, ventral fin, rudder, engine bay doors, nose cones and fairings. Twenty five per cent of the airframe weight is fabricated from composite materials.

For the European Fighter Aircraft (EFA) currently under development, the projected target for composites utilization is 35% involving the main wing, the forward fuselage, and the fin and rudder. For any structure or subassembly it is likely that a combination of materials will be used, each applied so that its individual set of properties can be used to best advantage. As an example a scale test section of a proposed fuselage design is shown below where a number of combinations of materials are employed. 

It incorporates carbon and Kevlar reinforced epoxy, aluminium honeycomb and unidirectional and woven reinforcements. Parts integration is also a key factor. The imaginative design of tooling and the taking advantage of the flexibility available in composites design allows parts lists, compared with the equivalent metal fabrication, to be much reduced. The design of an integrated composite horizontal stabilizer, offering an overall 15% weight-saving is shown below.

Although the basic material cost is more expensive, this is more than offset by reduction in substructure and assembly costs due to the smaller number of component parts. Overall the production cost savings are projected to be 18% over an equivalent all-metal stabilizer. Taking this concept to the limit, designs are now being proposed for an entire filament-wound aircraft fuselage. Two methods of providing the required stiffness characteristics have been studied. The first employs a filament wound isogrid network. Here skins are initially wound at [±45/90] and then split. A geodesic reinforcement is then bonded to the two skin sections. The second method uses a honeycomb stiffened structure with CFRP skins. Lightweight composite tooling is used which is capable of outward expansion and allows consolidation of the part during cure in the female mould. The excellent fatigue performance of composites is used to good effect with propeller designs. Using a combination of a unidirectional carbon spar and glass cloth reinforced skins at [±45°] for torsional resistance produces a very effective component. 

Complex aerodynamic shapes are more easily manufactured with composites than with metal and this provides added design flexibility. A tough polyurethane paint coat is often applied at the end of the production process to protect the structure from debris damage and runway stone impact. In jet engines there is scope for composites, particularly in the cooler regions. Carbon/polyimides have been proposed for both rotors and stators and the early problems associated with turbine blades have been largely overcome, although in the latter case there is strong competition as a result of the good performance of titanium. Ducts of various types are also fabricated from carbon fibres, and Kevlar have application for the ring that surrounds the engine which is required to be designed to contain debris in the event of turbine blade failure. 

Helicopter airframes also provide opportunities for composites. In a Franco-German programme a new aircraft is proposed which includes 80% composite; 24% CFRP 42% CFRP honeycomb and 11 % Kevlar honeycomb. Typical design features are: 

• Frames and beams of Kevlar 49 carbon laminates. 
• Panels of carbon and Kevlar sandwich construction with a Nomex honeycomb core.
• Carbon/ Kevlar sandwich structures for the underfloor for high energy adsorption.
• Landing gear frames of CFRP laminates. 
• Carbon sandwich structures for the tail boom for stiffness and strength. 
• Vertical and horizontal stabilizers of carbon and Kevlar laminates. 

In a similar way to propellers, composites have revolutionized helicopter rotor blade design. As the blade rotates, pitch changes, which are necessary to balance lift forces, cause very high levels of fatigue loading. Composite main rotor blades that utilize unidirectional CFRP in the spar design have virtually unlimited life. 

Furthermore, with advances in aerodynamic design it is found that blades of complex form are required for optimum performance. These are now universally made using composites fabrication costs for a similar metal design would be prohibitive. Future blade developments are likely to focus on aeroelastic designs where the blade structure and material properties control motion in order to modify aerodynamic performance, or reduce stresses or vibration. To achieve this in a passive manner is likely to require radially distributed aerofoil sections in combination with highly asymmetric and complex hybrid laminate constructions. 

Conventional rotor hubs, i.e. the structure that connects the blades and the main body of the aircraft, are very complex units which have a multiplicity of bearings, seals and lubricators to allow blade movement whilst ensuring proper load transfer. Novel designs using elastomeric composite materials with high levels of elastic deformation have resulted in concepts which are essentially bearingless. The advantages of this are manifold, including reduced maintenance and drag, reduced parts count, lower weight and improved damage tolerance and lifetimes. The centrifugal load from the elastomeric bearings is carried by a hub plate which is also composite. For all aircraft structures air certification is a major issue and the difference in behaviour between composites and metals has required reconsideration of the prevailing guidelines. The static structural loading cases and the minimum ultimate factor of safety to be considered are the same for conventional metallic structures, i.e. the design ultimate load is 1.5 times the design limit load (the maximum expected service load). However, owing to their anisotropic elastic behaviour, composites can be sensitive to individual design features and therefore regulatory authorities normally require both structural analysis and test results for each design case. 

For analysis, design allowables are established in a similar way to metals, viz.: 

• Failsafe/ redundant designs require a value above which at least 90% of the population of values falls with a confidence of 95%. 
• For a single load path design, the value above which at least 99% of the population is expected to fall with a confidence of 95%. 

A typical route for obtaining the required evidence to substantiate a safe design would be as follows: 

• Structural analysis identifying the key design features. 
• Confirmation of analysis by loading a fully instrumented structure. 
• Determination of allowable values for each significant structural feature with allowance for variability and environmental effects. 
• Test of the complete structure to a minimum of the design ultimate loads and check that allowables are not exceeded. 

For fatigue loading two approaches are used. In the first, a 'safe life' is defined where no significant damage is likely to occur. A component is not allowed to operate beyond this life. The second method is based on a 'failsafe' approach where the component can remain in service accompanied by an inspection regime which will detect any flaw before it becomes of critical importance. The choice of method depends on the aircraft of concern. For large transports a failsafe approach may be taken, whereas helicopters may be based on a safe life methodology. When testing a large structure either statically or in fatigue there are a number of factors that need to be considered. Owing to cost the simplest way is to test at room temperature without imposing environmental conditions and using a factor on loads to cater for the effect of the environment, cycling and variability. Individual factors need to be determined for each case but they are often in the range 1-2. As these terms tend to be dominated by degradation of matrix-based properties, tension members and, indeed, any metal components can be unduly penalised. A second approach is to carry out rigorous testing on sub-assemblies and then carry out a static room temperature test on the completed structure up to ultimate load. If the strain value determined in the structure test exceeds values determined previously, the system is deemed to have failed. In addition to the mechanical design loads and the operating environment, any structure must also be seen to conform to performance standards for irregular loadings, for example impact (bird strike, turbine disc burst, low velocity impacts the results of which are not visible), flammability and lightning strikes.

Design example: sine wave spar

Proposed designs for composite wings often consist of a number of spars with laminated skins. A multispar design means that load on the individual spars is low and buckling tends to be the dominant design criterion. A structure of sine wave configuration is ideal for this application.' z The design of the sine wave configuration itself is influenced by a number of factors:

• The size and spacing of fasteners through the flanges.
• The width of the web.
• The critical buckling load.
• The ease of processing.

Different geometries are possible but it is found that a wave configuration based on arcs which are not tangential, but separated by a small flat region is the optimum of the alternatives. Of the other options tangential arcs pose tooling difficulties and a true sinewave has insufficient buckling stability. A typical spar cross-section is shown below.

The webs of the spars have three layers; two CFRP cloth plies with fibres orientated at [±45°] to transmit shear loads, and one unidirectional CFRP ply in the centre to provide vertical stiffness. The [±45°] layers are folded over to form the flange. Additional reinforcements are applied each side of the web, orientated at 90°, to transmit the load from the fasteners into the web. The structure is completed by capping plies on the tops of each flange. The stability of the web as a function of the number of laminate layers is shown below.

Matched metal tooling can be developed for the sine wave spars, as, in detail, the spars are not symmetrical about a centre-line and the tools must therefore be capable of splitting into several parts.

Return to Top.......