Critically Examining The Structural Wellbeing Of Boeing Engineering Essay

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Since the early years of the 20th century, aircraft have had a dramatic structural and performance improvements, including the materials used in the construction of different parts of the aircrafts .The discovery of aluminium in the 1920s provided the first step forward to advanced material technology. Aluminium has remained the primary material for aircrafts; however it is gradually being replaced by advanced composites as the material as the choice for next generation aircraft.

Initially military aircrafts were utilising fibrous composite materials in small quantities during the 1960s and 1970s after which their exploitation in the aerospace industry has dramatically increased. By the commencement of the 1980 era, the composites were being used by civil aircraft manufacturers for a variety of secondary wing and tail components such as rudder and wing trailing edge panels and the first airplanes to fly under their own power were constructed over a century ago (Flower and Soutis, 2003). During the 1980s, engineers became more inclined towards higher strength in the aluminium alloys in an effort to conserve weight. Different aluminium alloys were selected depending on environmental factors like the amount of stress they encounter and the level of fluctuation during the stress period. Therefore, if a component bears higher stress without fluctuation then a high-strength aluminium is a better choice. On the other hand, if there is a lot of fluctuation in stress levels then a tougher alloy is selected because there is an enhanced possibility of fatigue. The problem with aluminium alloys today is that engineers have predicted that they are nearing the end of being able to drastically change aluminium alloys for the better (Teresko, 2007) which indicates that a significant decrease in airframe weight and an improved performance of an airframe can only be achieved by using a different material. The use of composites was then highlighted and engineers, for a while, have known about composites and the drastic impact they could have on the performance and cost of airplanes (Smith, 2003).

Projects objectives:

The paper is intended to highlight the rising trend of exploiting composite materials within the aerospace industry by specifically focussing on Boeing 787 and investigating the related issues. Furthermore, by analyzing the reliability of the composite materials and carefully comprehending the contrast between the designs based on aluminium and composite materials, the prime intent of the study is to examine the unusual structural characteristics of the composite's wing and fuel tank structures of Boeing 787 and thereby assessing their dependability.


The most significant use of composites in commercial transports has been on the Boeing 777 having the structural weight of 10% made by composite materials (Smock, 2007). Figure 1, shows the various composite structural elements used in the B-777.


Cost and reliability are the predominant factors in the case of transport aircraft and the composite applications seem to be levelling off at 20 percent of the structural weight a ceiling lower than for combat aircraft. The barrier in this case is set by the affordability of the airframe since initial acquisition cost plays a major role in airlines' selection of a particular model. Boeing experienced great success with using composites on the 777 and hence they introduced their new airplane anticipated to enter service in 2009, the Boeing 787 with the use of more composites in the design (Smock, 2007). The growing trend of composite within the commercial aircraft industry, it can be predicted that the composites will continue to make a huge impact in the coming future (Smith, 2003). Furthermore, the composite materials are extensively being deployed in primary load carrying structure as for instance, the airbus A380 uses composite materials in its wings, which helps enable a 17% lower fuel use per passenger than comparable aircraft. The expenditures for advanced composite materials in the aerospace industry starting from 1982 till year 2000 are projected in Figure 2. It was in 1983 when Dasa Airbus introduced an all-composite rudder for the A300 and A310, followed by a much more complex vertical tail fin launched two years later.


$ Millions


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1982 1987 1992 2000

Source: P-023N Advanced Polymer Matrix Composites, Business Communication Company, Inc.


Composite materials are a physical combination of two or more compatible materials, those are different in composition or form where the individual constituents retain their separate identities and do not dissolve or merge together. These separate constituents act together to give the necessary mechanical strength or stiffness to the composites system. [1]

Generally consisting of a primary fibre and a binder material in which the binder material forms a matrix to hold the fibres together and fill voids between them thus reinforcing the matrix structure which allows stress transfer between the fibres. Composite laminates are used as structural materials when matrices are layered together to increase their strength and to further provide extra strength and shape to the structure[3]; a core material, i.e. foam, aluminium or Nomex honeycomb is often sandwiched between two sheets of composite laminates as shown in Figure 3. The name of the composite usually identifies what the fibre and matrix materials are as for e.g. glass/phenolic, carbon/epoxy composites (Potti, 2004). Carbon/epoxy and glass/epoxy fibre composites are generally used in primary structures whereas glass/phenolic due to its brittleness is not used in primary structures and the evolution of volatiles, and is employed in secondary structures and cabin furnishings (Potti, 2004).


Source: Rakow & Pettinger 2006


A composite may be broadly defined as a combination of two or more materials; each with distinctive properties which are attractive alternatives to the metals used in general aviation aircraft because they weigh lighter, stronger, stiffer, and have almost no fatigue and corrosion problems (Nangia, 2006). There are several types of composite materials, including particle-reinforced, fibre-reinforced composite materials, etc. Hybrid composite materials containing continuous, fibre-reinforced plies and metal layers are special composite materials because of their high specific strength, high specific modulus, excellent electromagnetic shielding characteristics and very good high-cycle fatigue property. Typical examples of such composites include CARE i.e. carbon fibre reinforced aluminium foams or laminates and ARALLs i.e. Kevlar fibre reinforced aluminium hybrid composites (Mecham, 2005). These composite materials consist of alternating layers of metal sheets and fibre-reinforced epoxy composites. The unique properties of the fibre-reinforced epoxy composites are retained and the materials are immune to environmental attack due to the incorporation of the sandwiched metal layers. The metal layers are also responsible for providing high shear strength thus applications of such hybrid composite materials in aerospace, electronics and automotive industries have been considered (Nangia, 2006). Additionally, composites also allow the option of directional property tailoring and reduced parts count.


When constructing all-composite aircraft, structural components can be joined in one of five general ways as explained below. It is well-established that relatively thin-walled structures are most efficiently joined through adhesives, whereas thick-walled structures lend themselves to mechanical fastening (Jones et al, 2002). As in metallic joints, prime considerations in composite joint design are bearing, tensile, and compressive strengths in the laminate and shear and bending stresses in the fasteners (Jones et al, 2002). Shear and bending stress in the fasteners are typically not the most critical consideration in composite joint design. Due to the weakness of the composite material, the fastener controlled failure modes are usually far less likely than those controlled by the composite material (Krishnamurthy, 2006).


Co-curing is the process of taking two or more composite components that use similar resin systems, and then curing them simultaneously so that the resin flows between the two pieces, and they become one piece (Arthur, 1996). The process is carried out on wet layup parts that have been brought to the handle cure stage; they are then post-cured together. Parts fabricated from pre-impregnated materials that are individually laid up may also be co-cured when they are physically pressed together with vacuum pressure, autoclave pressure, etc., during the cure cycle, often with the addition of a film adhesive.


Secondary bonding with structural adhesive is the combining two or more individual pieces for all but wet layup parts, this is performed after the pieces are fully cured. The surface is prepared for bonding; this generally consists of abrading and solvent cleaning the two surfaces to be bonded (Weaver, 2002). The prepped surface is then sheltered with an adhesive mix of resin system with inert filler materials and the entire assembly is cured by placing the second piece over the adhesive.


Ultrasonic welding, induction bonding, dual resin bonding, resistance heating, and focused infrared energy are few techniques of advanced fusion bonding. All of them, however, are limited to thermoplastic resin systems.


Mechanical fasteners can also be used to join together composite components, although their use generally takes away many of the advantages of composite construction as, for example, the use of a fastener to join together a rib to a wing skin would lessen the effectiveness of an otherwise laminar airfoil (Weaver, 2002). They also break into what would otherwise be continuous fibre reinforcement, thus reducing the composite effectiveness. For the adherence of metal components with composite structures, fasteners such as a bolt or rivet are often used.


Entrapment is the last joining method which consists of embedding mechanical fasteners into the composite structure when the part is cured. The technique of incorporating metal inserts and attachments into composites is expected to provide significant improvements in both part performance and fabrication efficiency and thereby a reduction in cost simultaneous with an increase in reliability over adhesive bonding and direct mechanical fastening (Potti, 2004).


There are significant differences in the behaviour of fibre composites compared to traditional metallic materials e.g. aluminium, steel, and titanium structures when placed under load, or even when failure occurs which often causes composite structures to fall short as compared to metals. For example, a metal structure in tension would fail in tension, whereas an equivalent composite structure in tension might fail in bending (Rakow & Pettinger 2006). The composite is a fibrous matrix with multiple load paths and it is important to note that the plies in a laminate may be oriented differently, can be of varying thickness, or imperfections may exist between the plies such as air bubbles which cause it to behave differently. In general composites are brittle and will therefore fail easily without undergoing deformation, as in the case of metals which are ductile, which serves as a warning that failure is about to occur. These variables make composite materials very unique and directly affect how they fail and behave under load. As a result of which it is inherently more difficult for Transport Safety Investigators (TSIs) to analyse failed composite structures and clearly determine what types of loads were involved since there could be many reasons for the failure of composite materials. Some composites that researchers have tested fatigue lives for essentially follow the same relationship that metals do (Putić et al., 2003). Conversely, some composites are sensitive to fatigue while others appear insensitive (Mecham, 2005). On the other hand, some composites are sensitive to stress concentrations such as notches or holes while other composites do not seem to lose any structural integrity with the presence of notches or holes (Flower and Soutis, 2003). All these conditions have made it difficult for engineers to agree with a common model for composites. Furthermore, cracks in composites are highly complex and a structure does not usually fail from one large crack, like in metals, but from numerous small cracks that coalesce or from delaminating (Kelly, 2008) and numerous small cracks can start in a matrix and be stopped by fibres (Marks, 2005). It has been clearly observed that when the stress is large enough these cracks can, all of a sudden, start growing again and the structure can catastrophically fail without much warning (Marks, 2005) making it tougher for engineers to test composites and estimate the amount of damage done to a structure to derive equations relating loads to fatigue damage. Testing on current composite airplane components has revealed that ninety percent of all stress amplifications or overloads show up as wrinkles on the surface of the composite which are easily detected by eye. However, sixty (60) percent of these surface defects can recover elastically and hide the damage below the surface (Marks, 2005).


Following is the detailed analysis of certain contributing factors that greatly affect on the reliability of composite materials:


The hybrid composite materials are frequently subjected to thermal and mechanical fatigue loading apart from external mechanical loadings; the thermal effect is identified as an essential factor that determines the stress distribution in composite materials. During the curing process, adhesively bonded composite/metal laminate structures are held at elevated temperatures over 120 â-¦C, very high residual stresses could build up because of the difference in coefficients of thermal expansion (CTE) for different materials (Jones et al., 2002). The CTE of aluminium is about 2.36 £ 10−5/â-¦C and for polymers it is higher than 1.05 £ 10−4/â-¦C (Flower and Soutis, 2003). This thermal disparity results in delamination or debonding of hybrid composite materials, which facilitates fatigue crack growth in the polymer/metal interface and the changeability of ambient temperatures is also responsible for thermal cyclic stresses as, for example, the change in environmental temperatures is obvious when an aircraft travels across different continental regions or varying altitudes. For electronic devices, the temperature variation associated with the power on/off can reach as high as several tens of degrees. Therefore, the stress state in a hybrid composite material is not only dependent on service conditions, but also affected by the materials processing parameters (Jones et al., 2002). The overall stress distribution influences the fatigue crack growth behaviour and the durability becomes an increasing concern. Fatigue damages in the interface region account for the majority of failures of materials and this leads to fibre composite structures failing in different ways to metal structures traditionally used in aircraft.


The fibre composite structures that have failed in tension show no common characteristics, which would indicate that a tension load was the cause of the failure on a macroscopic scale. Figure 4 shows a series of carbon fibre reinforced plastic (CFRP) samples that have failed under exactly the same tension force, yet show a huge variety in failure patterns. The samples were split into four groups, each group having the ply fibres oriented in a different direction, where some of the samples splintered upon failure (upper left) and others have snapped or sheared at an angle (upper right and lower left), while in some samples the fracture surface is ripped (lower right). This variety of failures is due to the variation that is inherent in composite structures due to different fibre orientations, and imperfections between plies in the laminate (Cunningham, 2007). This highlights the challenge of analysing composite structures that have failed in tension. Each of these samples share common signs that indicate tension failure, on a microscopic level. In all failures of composite structures under tension, the fracture surface generally has a rough appearance (Rakow & Pettinger 2006).


Source: Rakow & Pettinger 2006

When the fibres are aligned in the direction of the tensile load, fractured fibres are often found sticking out at the fracture surface and this is called fibre pullout, and is a typical indicator of tension failure in composite structures (Figure 5).


Source: Rakow & Pettinger 2006

Fibre pullout is caused by the individual fibres breaking and being pulled out of the matrix which results in holes in the matrix, which is another indication of tension failure. In some tensile failures where the matrix itself fails, the fibres do not break which is referred to as fibre bridging. The length of pulled-out fibres can indicate the environmental and loading conditions that the composite was exposed to at the time of failure, such as exposure to moisture, temperature and rate of loading (Cunningham, 2007). When the fibres are not aligned in the direction of the tensile load common with multiple-ply laminates, failure often occurs in the matrix rather than the fibres (Rakow & Pettinger 2006). Tension matrix failures generally occur between fibres and these types of matrix failures usually cause hackles, which are rough features on the fracture surface as illustrated in Figure, 6.


Source: Rakow & Pettinger 2006


The growth of cracks between different plies in a laminate, called Delamination, is the most common failure mode for fibre composite structures. This occurs when shear loads are applied between plies in the laminate. Since the fibres are significantly stronger in tension than the matrix, the matrix cracks and delamination occurs (Brimhall 2007). Delamination can propagate throughout the composite structure upon repeated loading, causing catastrophic failure if left undetected (Rakow & Pettinger 2006). Delamination failures are characterised as one or a combination of three modes including: Opening (Mode I); and/or Sliding-shear (Mode II); or Tearing-shear (Mode III).

Previously, only Modes I and II were considered when analysing the tolerance of composite structures to damage, however a new edge crack torsion test has allowed better analysis of toughness against Mode III failures (Glaessgen & Schoeppner 2006).


Source: Werfelman, 2007


Improvement of the reliability of hybrid composite materials relies on the enhancement of polymer/metal interface bonding. Various surface treatments including alkaline etching and acid pickling applied separately or in combination with phosphoric acid anodizing, plasma processing, ion beam irradiation (Khosravi et al, 2008) and coupling agent treatment have been explored to examine the effect of pre-treatment on the adhesive bonding between metals and polymers. It has been observed that the presence of oxide and small molecules such as water in the interface region is responsible for the degradation of bonded joints (Jones et al., 2002). Numerous studies reveal that the chemical bonding at metal-polymer interface plays an important role in adhesion, thus, the interfacial bonding and subsequent adhesion are directly influenced by the way that the interface is formed. Furthermore, it has been estimated that aluminium/polymer joint where a thin and uniform metal sodium layer was coated on the polymer surface. The nature of the bond formation at the metal/polymer interface was investigated in view of compound formation and charge transfer between sodium and the polymer (David et al, 2004). A bonded joint was tested in terms of its strength, thermal resistance and tightness to show the interfacial properties. Underhill and Rider (2005) investigated hydrated oxide film formation on aluminium alloys immersed in warm water. Porous oxide structure was found due to the growth of hydrated oxide films on 2024 and 7075 aluminium alloys immersed in de-ionised water, at the temperatures of 40»50 â-¦C for periods up to a couple of hours (Cunningham, 2007). In contrast with film growth studies reported for pure aluminium, the alloy systems do not appear to show an incubation period prior to hydrated oxide growth. Various characterisation techniques were applied to study the properties of the oxide structure including Fourier Transform Infrared Spectroscopy (FTIR), weight gain measurements, high resolution Scanning Electron Microscopy (SEM) and Atomic Force Microscopy (AFM) and it has been submitted that the films formed at 50 â-¦C are much thicker than those formed at 40 â-¦C (Underhill and Rider, 2005). However, the porosity of the films appears to be comparable at both temperatures. The research has suggested that a porous oxide structure is likely to be very suitable for adhesive bonding because of the increase in interface area of nano-porous structure, which results in the high shear loading capability (Xie and Wong, 2003). However, the interface nanostructure remains to be revealed by further systematic study. Conclusively, reinforcement/matrix interface plays the key role in determining the performance of advanced composite materials. To enhance the interfacial bonding, nanostructures are introduced into composite materials. Formation of nano-pores on metal surface can increase the bonding strength of the metal/polymer interface. Surface treated carbon nanotubes are used in preparing nano-reinforced matrices. The nano-fibre reinforced epoxies containing reactive graphitised carbon nanotubes as new adhesives can help to alleviate the residual stress problem because they are more ductile than the conventionally used pure epoxy adhesives (David et al, 2004). Finally, the progressive damage of interfaces in composites can be evaluated by nonlinear models due to the complexity of the deformation and failure processes.


A transient charge from a strike of light is projected by a higher conductivity of aluminium, causing metal body discharges with current from the discharge being distributed relatively evenly over the body. It has been observed that a minor damage to aircraft components is caused by typical lightning strike to a metal aircraft. On the other hand, the carbon fibre composites generally have a higher strength to weight ratio than aluminium hence they are increasingly replacing aluminium structural components. Unfortunately, the Carbon Fibre Reinforced plastic is approximately 2000 times more resistant than aluminium. Insecure components that are either entrenched or affixed to CRFP aircraft skin that is 35-40% composed of resin, if stroked by radiance, does not disperse as readily as in case of metal (Brimhall, 2007). Considerable amount of damage can occur due to a strike of lightning on CRFP aircraft skin that results in obliteration of vital portions in the aircraft. Damage is amplified by the temperature and the carbon fibres becoming extremely hot due to discharged current through the composite skin's resistance which dissipates a burst of energy. Another result of which is that the skin temperature of the CFRP structure becomes much hotter in comparison with the conventional aluminium structure. The rise in temperature results in CRFP resins to vaporise turning a part of it from solid to gaseous state. After which, there is a possibility of entrapment of a small amount of gas inside the CRFP skin layer (Dodge, 2007). The internal gas pressure from the vaporised resin can damage the structure causing delamination and possibly puncture the underlying systems or structure. As the vaporised resin escapes explosively from the skin, hot particles (sparks) eject from fastener's interfaces and composite joints from the CRFP. The lightning strike which has less effect on aluminium structure may cause disastrous effect on CFRP. For adequate protection for a composite wing structure, the exterior CFRP structure must withstand not only the initial lightning strike but also at least 100 kilo amperes (100kA) of discharged current without adverse effects or impact to the safety of the flight (Croft, 2007). Furthermore, a direct strike to an exposed surface makes skin fasteners most vulnerable as it may cause sparking at structural joints and, more importantly, in the fuel tank (Dodge, 2007). In order to minimise any chance of sparking the composite structure of the aircraft must have some protection especially at exposed skin fasteners and at fuel and hydraulic couplings in the fuel tank. The economic feasibility for this protection is highly solicited during the initial application of the composite-based structure followed by its enhanced effectiveness in order to avoid any probable damage. The crash worthiness and potential counter flammability of the aircraft determines its positioning as a protected airliner thus abating concerns about composite materials.


In accordance with the applicable airworthiness regulations and special condition requirement by the FAA, the 787 needs to comply with certain limitations associated with noise, fuel vent and exhaust emission. With the help of an integrated fuel tank, 787 anticipate to control fuel tank flammability. The NGS fuel tank is cleverly designed to considerably reduce the fuel tank flammability of the aircraft wing below the maximum perimeters as positioned by the FAA regarding the flammability of the wing fuel tank. (FAA, 2006). The atypical design characteristic of the Boeing 787 wing fuel tank proclaims a higher level of performance and distinction in the structural composition of the aircraft. Concerned with human life security issues, FAA requires the aviation company to prove the crash worthiness and counter-flammability capabilities of the aircraft. There is an increased apprehension about the ignition source in the fuel tank system of 787 and the company is obligated to provide sufficient evidence about the sanctuary measures that have been undertaken with intent to circumvent malfunctioning of the wing structure and the fuel tank resulting in collapse or flammability of the airliner due to any single or a combination of factors. Furthermore, the company is forced to highlight all the relevant and concealed technical aspects to surmount over any improbable concerns. This requirement defines three types of scenarios (FAA, 2006) that must be addressed:

Despite of probable episode of structural malfunctioning in the aircraft, there should be no chance of ignition due to any single failure,

The chance of failure due to an obscure failure condition not shown to be at least extremely remote, despite of any possible likelihood, must not cause an ignition source;

Due to any combined sources that might cause any structural damage in the wing or fuel tank of the all-composite aircraft, any chance of ignition must be diminished.

Most importantly the main intention of the Boeing Company is to focus on preparing an authentic report satisfying all the regulatory requirements as set by the FAA for the purpose of securing human lives by integrating every possible damage prevention measure and ensuring the absolute crash worthiness and counter-flammability of the Boeing 787. In addition to the documentation, Boeing must comply with the FAA guidelines regarding the security concerns associated with the wing and fuel tank structure of the all-composite aircraft.


The arsenal of engineering materials available for aircrafts is now comprised of composites which are formulated by inserting chopped-fibres and unidirectional fibres into a matrix made of another material (Potti, 2004), usually a type of thermosetting polymer or resin which, with the addition of heat and pressure, hardens around the fibres creating the composite (Smith, 2003). For the purpose of acquiring a desired material structural property, the composites can be further complicated by stacking several laminates on top of each other in different orientations. Composites have become a critical facet in many engineering applications today because of their incomparable and impressive strength-to-weight ratio and by being stronger and lighter than the aluminium alloys used in the aircraft industry (Peter, 2008). Most composites anticipated to be used in aircraft are polymer composites where fibres of particular orientation and direction are suspended in a thermosetting polymer (Flower and Soutis, 2003). For a proper application, the direction of fibre is of particular importance. The most common fibre direction used in composites is the 0°/90° fibre orientation giving maximum strength of the composite in the 0° and 90° directions (Peter, 2008). Composite wings (Figure, 8) are highly optimised, given different loading conditions, by changing fibre orientation, fibre material, matrix material, laminate orientation, and laminate stacking sequence.


Composite materials are being used extensively in load-bearing applications throughout the

Aerospace industry and Boeing 787 has been introduced as an all-composite aircraft. The technical analysis of the aircraft reveals the potential weakness of a composite material which is its poor out-of-plane load transfer capability. To improve the out-of-plane performance of composite laminates and joints, transverse reinforcement is used through stitching, fibre insertion, or Z-fibre pinning. Fibre yarn or tow is used to conduct transverse stitching which is a two-sided continuous stitching process. On the other hand, z-directional reinforcement is provided by the z-fibres which are the discrete pins placed through the thickness of a laminate or joint. Analytical and numerical evaluations of single lap joints with transverse stitching shows a reduction of the critical peel stress when stitches were located near the specimen edge; this increases the stitch modulus which is shown to reduce critical peel stress. The effects of stitching on the stiffness, strength and failure mechanisms of composite laminates have been evaluated (Mouritz and Cox, 2000). The mechanics of z-fibre reinforcements are studied and it was found that the presence of z-fibres at the interface reduced the crack tip nodal stress (Stickler, 2001).


The mechanical behaviour and failure mechanisms of T-joints constructed using a fibre-insertion process are investigated under flexure, tensile and shear loading (Stickler, 2001). The T-joints with transverse stitching (Figure, 8) are shown to fail in a series of steps with increased load. The composite materials with transverse reinforcement include identifying damage mechanisms, damage modelling, and static and fatigue performance testing. The NASA/Boeing Advanced Subsonic Transport (AST) program developed a composite wing using dry fibre performs, transverse stitching, and resin film infusion (RFI) with the goal of reducing the cost and weight of composite wing structures for commercial transports by 20 percent and the airline operating cost by 4 percent over that of conventional aluminium designs (Karal, 2001). The wing cover panel during the stitching process (Figure, 8) and a completed stitched/RFI wing cover panel as illustrated in Figure 9 shows that the structural wing box was modelled using nonlinear finite element analysis and good agreement was shown for the global behaviour (Jegley, 2001).


Boeing 787 relies on laying down the fibres by hand into the matrix and then adding the heat and pressure required to harden the composite which is a very expensive and labour-intensive process that has made the composite structures more expensive. Composite structures were once only used when weight savings was critical and price was not (Khosravi et al, 2008). In the case of the Boeing 787, the process of combining the fibres and matrix together has been made more efficient with the introduction of (ATLs) i.e. automatic tape layers (Smith, 2003). Once the mixture is completed, the fibres and unhardened matrix go into an autoclave which adds the heat and pressure required to harden the matrix (Flower and Soutis, 2003). Composites are then formed into larger, more complex shapes to cut down on joints and fasteners which give them an added benefit. The biggest attributes that composites have over aluminium alloys in building, for example, a fuselage, is the fact that composites can be formed into a continuous cylinder without any longitudinal joints (Smith, 2003). According to any strength of materials text, the stresses that act tangential in a pressurized cylinder are approximately twice as high as the stresses that act in the longitudinal direction and, therefore, it can be said that not only do aluminium alloys have the inherent disadvantage of possessing less strength than composites, but they also must have a weaker joint somewhere along the cylinder's longitudinal axis where two aluminium alloys panels come together. Composite skins naturally have a great advantage over aluminium in that they can be made thinner due their higher strength and lack of weak joints which leads to a manufacturing problem. Composite parts which are as large as fuselage sections must also have ATLs and autoclaves large enough to make and cure the composite. To address this, Boeing developed the world's largest autoclave that weighs in excess of 500 tons to help construct the massive one-piece fuselage sections (Flower and Soutis, 2003) that exemplifies how the switch from aluminium alloys to composites can have an extreme initial cost. It has been anticipated that the manufacturing process for composite construction will decline and eventually it will be cheaper to build composite airplanes versus ones made primarily out of aluminium alloys (Smith, 2007). Furthermore, there is another characteristic which makes polymer composites superior to aluminium alloys and that is the inability to corrode which is why engineers have wanted to make the move from metal to composite airframes.


For nearly eighty years, Aluminium alloys have been the preferable choice for aircraft materials due to its high strength-to-weight ratios and being readily available and less susceptible to corrosion compared to any other metal. Metallic materials have been exceedingly predictable in their behaviour given the inputs of metal parameters, component shape, and loading conditions such as impact loading, cyclic loading, and static loading (Potti, 2004). The weight of an airplane is one of the largest contributors to fuel efficiency and cargo capacity therefore the airframe must possess maximum strength-to-weight ratio of the structure. Aluminium has very desirable material properties for use in airplane structures as compared with other metals and much effort has been put forth in optimizing the extraction and purification, alloying, and the manufacturability of aluminium. Companies such as Boeing have turned to computer simulations to optimise these manufacturing processes over the past couple decades (Meyendorf et al., 2002) and there is not much room left to improve the process of manufacturing airplanes out of aluminium (Potti, 2004). Pure aluminium has the distinction of possessing excellent corrosion resistance yet poor strength and engineers routinely add pure sheets of thin aluminium on the surfaces of aluminium alloys, called cladding, or place special coatings on the surface to improve the alloy's corrosion resistance (Klassen and Roberge, 2008). Despite the drastic effect cladding and special coating has on aluminium alloys' corrosion resistance, corrosion inevitably becomes a major concern when using aluminium alloys for aircraft structures (Meyendorf et al., 2002). The life of a structural component is endangered by corrosion leading to corrosion pits that create stress concentrations which, in turn, results in fatigue failures and further providing an opportunity for cracks to amplify. Corrosion causes much damage in terms of cost and weight than any other material hence billions of dollars are spent by the aviation industry just on corrosion repairs. The exceeding corrosive nature of the metallic components have given rise to the adoptability of composite materials within the commercial airline industry as a significant maintenance cost can be saved by utilising a better alternative, structural material.


Unlike metals, composites are inhomogeneous and are non-isotropic. There is no general model that predicts the mechanical behaviour of composites like there is for metals (Khosravi et al, 2008). Composites have the tendency to fatigue which is considered as comparatively a much more complicated process (Edmonds and Hickman, 2000) causing a blend of four damage modes under fatigue loading with composites i.e. matrix cracking, fibre-matrix debonding, delamination, and fibre fracture (Peter, 2008). Therefore, regular inspections of the composite must be administered so that these defects can be observed and fixed before they disappear. One of the most promising benefits of composite designs over its rival aluminium is its ability to be formed into large and highly complex shapes without the use of rivets and bolts (Dodge, 2007). This means that larger structures can be made instead of several smaller ones hence reducing manufacturing cost of putting together several smaller sections and worrying about the numerous fasteners required to hold the structure together. Moreover, this also signifies that there is no need to account for numerous rivets and bolts fatiguing in addition to allowing the structure to be much more aerodynamic (Edmonds and Hickman, 2000) and ensures crash worthiness of the airliner. The flammability of fibre composites used in aircraft is regulated by the burn tests performed on exterior and engine compartment structures made of composites to ensure that they have the same or better fire resistance than equivalent aluminium structures. In the case of the all-composite fuselage of the Boeing 787 Dreamliner, the FAA has stated that the fuselage cannot be assumed to have the fire resistance previously afforded by aluminium (FAA, 2006). This is partly a precautionary measure, due to regulatory inexperience with large scale applications of fibre composite in aircraft (Croft 2007). Fibre composite materials have different flammability characteristics than traditional aircraft materials such as aluminium, with the key area of difference between the flammability of metal versus composite structures is the chemicals used to bond fibres together. When composites are exposed to high temperatures (300-400 °C and above) the bonding matrix decomposes, releasing heat, soot, smoke and toxic gases. The reinforcing fibres (such as aramid or carbon) may also decompose, creating fibrous dust and adding to the heat and toxic smoke (Mouritz 2006). The burn point for composite is at 350 degrees which is far lesser compared to aluminium which is at 600 degrees. Boeing is using graphite epoxies on the 787 where the design and construction optimises the characteristics so the fibres carry the loads and the resin bonding the epoxies gives the composite stiffness and strength. The fuselage, whether composite or aluminium, is the least of the problem as the fuel serves as a driver as for instance, wool suits that men wear gives off enough cyanide to kill everybody on board hence, it's not the fuselage, and it's the fuel, the interior and what people bring on board (Hamilton, 2007).


The overall weight of Boeing 787 is evidently reduced as compared to an all-aluminium structure by 10,000 lb which is approximately equivalent to 4536kg and the observation asserts for about 3-4 percentage points of the projected 20% fuel savings per passenger with the Boeing 787 over similar-sized aircraft in service as for instance, Boeing 767 and the Airbus A330-200 (Werfelman, 2006). Other factors affecting fuel efficiency include new engines in development by GE and Rolls-Royce, which will generate half of the savings; an optimized aircraft design that minimizes drag (3% to 4% improvement in fuel efficiency); and more efficient operational systems (3% to 4%). The integration of higher quantity of composites by Boeing 787 offer savings in other areas that includes reduction in airport landing fees which are based on the overall weight of the aircraft. Furthermore, Boeing estimates at least 20% lower airframe maintenance costs than on other comparable aircrafts. Boeing then introduced an inert system in 2005, subjected to in-service evaluation involving both Boeing 737 and 747 and was subsequently installed on both in 2007 followed by the installation on other aircrafts in 2008. The noteworthy aspect of this inert system was that it diverts engine bleed air into an air segregating module which consequently separates the nitrogen and forwards it into centre-wing fuel tank. Due to its appealing and advantageous properties, the same inerting system has also been integrated into the fuel tanks of new composite-based Boeing 787 model which is anticipated to enhance the fuel efficiency together with other benefits associated with the airliner.

A number of studies support the fact that fatigue and corrosion results in longer intervals between scheduled maintenance, higher system reliability, and eliminating the non-routine maintenance in the airframe which are the serious areas of concerns (Gillespie, 2009). Serving as a rescuer to the situation Boeing 787 claims to be available for revenue service in comparison to any other commercial aircraft with a service life of approximately 40 to 50 years which is about double to the contemporary airframes (Dodge, 2007). The advanced composites used in 787 not only perform equal to aluminium in impact resistance but has the potential to outperform it in some cases. Boeing has conducted extensive tests on the composite material in the past 10 years which evidently demonstrates a record of safety and performance in civilian as well as military aircrafts. Further adding to its benefits, Boeing 787 is all set to make a positive impression on passengers by resisting corrosion and fatigue thereby causing highly pressurised and more humid cabin environment resulting in travel comfort. This allows maintenance costs to be drastically cut since thorough checks around the structure would not have to include searching for oxidation damage and also serve the objective to obtain crash worthiness and to combat flammability.


To sum up the overall study, it can be said that the use of composites by the civil aviation industry and particularly by Boeing is shrewd and competitive as the advantages of adopting the composite technology extends beyond weight savings. Furthermore, increasingly practical ways can be observed in order to integrate functions into a single system as the use of composites is not anticipated to stop here because it's not just about the structural benefits. With the superior strength of the composite fuselage, higher pressurisation of the passenger cabin becomes feasible for controlled temperature, humidity and aeration making it an ideal technology that may add value to the concept of enhanced customer experience. While critically examining the composite materials, it has been constituted that they are much more durable than aluminium because of corrosion and fatigue benefits and also providing dramatic reduction in fasteners which makes the structure of the Boeing 787 to be a giant macromolecule in which everything is fixed firmly through cross-linked chemical bonds toughened with carbon fibre. By examining the increasing concerns about the composite technology that questioned the reliability of unusual structural characteristics, especially regarding the Composites Wing and Fuel Tank Structures of Boeing 787, the study explored that the aircraft manufacturing and materials technology is becoming advanced with a continuous thought process being involved to attain better performing results from the aircrafts. This paper has dealt with investigating the backdrop of issues pertaining to Boeing 787, and consequently explored the notion of carbon fibre materials which has been extensively incorporated in the manufacture of the aircraft. Followed by detailed comparison between the old designs of wing and fuel tanks with the new one i.e. composites and aluminium, this paper has maintained its focus to examine every possible aspect that may enhance the understanding of the benefits associated with the use of composite materials.