Aeronautical Materials

AERONAUTICAL MATERIALS

Tristan Burg and Alan Crosky

School of Material Science and Engineering

University of New South Wales 2001

Aeronautical Materials – Teacher Reference 2001 Materials Science and Engineering - UNSW

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1. AIRCRAFT MATERIALS - INTRODUCTION

2. THE AIRFRAME2.1 Aluminium Alloys2.2 Titanium Alloys2.3 Steels2.4 Composites2.5 Polymers

3. THE ENGINE3.1. Creep3.2 Oxidation3.3 Engine Materials3.4 Nickel Based Superalloys3.5 Turbine Blade Design3.6 Turbine Blade Coatings

4. CORROSION4.1 Surface Corrosion4.2 Stress Corrosion Cracking4.3 Galvanic Corrosion4.4 Crevice Corrosion4.5 Pitting Corrosion4.6 Exfoliation

5. NON DESTRUCTIVE TESTING5.1 Liquid Penetrant Inspection5.2 Magnetic Particle Inspection5.3 Ultrasonic Inspection5.4 Radiography5.5 Other Techniques

6. REFERENCES

7. INTERNET LINKS

8. SYMBOLSFront Page: Boeing F/A 18 [1]

In this set of notes we will

a) examine important design constraints involved with aerospace materials

b) provide an introduction to materials selection

c) consider design constraints of a number of aeronautical devices and detail thematerials in use

The notes are directed towards the materials section of the Aeronautical EngineeringModule of the NSW Engineering Studies Stage 6 Syllabus

It is not the aim of these notes to describe every aeronautical material nor everyaeronautical device.


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AIRCRAFT MATERIALS

1. INTRODUCTION

Aircraft come in a variety of shapes and sizes and the materials used depend on theindividual performance requirements and budgets. However, aircraft can be broadlycategorised into four groups:

1) Light Aircraft - An example is the Cessna shown in Figure 1. Materials for theseaircraft have less demanding requirements and their cost must be minimum. Typicalmaterials include steel and aluminium.

2) Business Jets –An example is the Beech Starship shown in Figure 2. Material costis less important for these aircraft and higher performance materials (such as carbonfibre composites) are used.

Figure 1: Cessna Skyhawk: Cost <US$ 500 000, Structure mainly aluminium alloys.Empty weight ~ 1000kg, Baggage ~100kg, Max take-off (TO) weight ~ 1500 kg [2,3]

Figure 2: Beech Starship: Cost US$4.3 million (1995). Majority of the structure isCFRP. Empty weight 4590kg, Max TO weight 6758kg, Max baggage 311kg[4,5]


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3) Civil Transport – These include aircraft such as the Boeing 747 shown in Figure 3.Low weight is critical for these aircraft and cost must be minimised. Take off istypically the most demanding phase of the flight (see Figure 4) and lifetimes extendfrom 50 000 to 100 000 hours. Typical materials include aluminium and composites.

Figure 4: Aircraft operational phases [6]

4) Military Aircraft – These include aircraft such as the Boeing F/A 18 shown inFigure 5. The take off and operational phases have demanding material requirements.Thus, material performance is critical and budgets are large. Lifetimes range from 5000 to 10 000 hours. Composites are used extensively in the airframe.

Figure 5: Boeing F/A 18: Cost ~ US$56 million, Structure mainly consists ofaluminium alloys titanium alloys and CFRP. Empty weight 10 810kg, Max TO weight16 651kg [1,2]

Figure 3: Boeing 747Cost > US$100 million, Structuremainly consists of Al-alloys and CFRP.Empty weight 180 895kg, Max TOweight 362 875kg, Payload ~ 33 00kg[1,2]


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Since the most commonly experienced aircraft are civil transport aircraft, these noteswill focus on this group. A detailed listing of worldwide aircraft, which is updatedannually, is given in Janes All the Worlds Aircraft [4].

The engines used in aircraft also vary considerably, again depending on theperformance requirements of the aircraft and its budget. Light aircraft frequently useair-cooled piston engines, but turbine engines are now used in all other classes ofaircraft. Even turbine engines come in a variety of forms, i.e., turboprop, turbojet andturbofan. In the turboprop engine a gas turbine is used to drive a propeller whereas ina turbojet, propulsion is provided by the jet of exhaust from the rear of the engine. Aturbofan is essentially a combination of the two in which about 90% of the propulsionis provided by a multibladed bypass fan located inside the front of the engine with theremainder coming from the jet thrust. Turbofans are used in all large civil transportsand these engines will be the focus when discussing engine materials.

The materials used for aircraft construction can be split into two broad categories;airframe materials and engine materials, and these will be considered separately.However, selection of materials for both applications is based on the designconstraints. These are defined by the mechanical, chemical and thermal propertyrequirements of each component. Typical design constraints include weight, stiffness,strength, fatigue performance (high/low cycle), corrosion resistance and cost.

Of major consideration is the specific stiffness (stiffness to density ratio E / ρ), andspecific strength (strength to weight ratio σ / ρ). The broad aim for aircraft design isto maximise payload in relation to cost. By increasing specific stiffness and strength,the weight of a given component can be decreased and fuel consumption and runningcosts decreased. For example when a fully loaded aircraft takes off, 20% of the weightis payload, 40% is aircraft structural weight and 40% is fuel. Therefore, a saving inaircraft weight can increase the possible payload or decrease the power required for agiven payload.

2. THE AIRFRAME

The airframe consists of components such as the wing upper, wing lower, fuselage,spars, frames, ribs, landing gear and control surfaces. Essentially, the airframe isrequired to resist applied loads, provide an aerodynamic shape and protect passengers,payload and equipment from the external environmental conditions. Each componenthas different specific constraints, resulting in different material selection criteria foreach component. The constraints on each component are detailed below and shown inFigure 6.

WingsOverall, the wings are subjected to the most complex and highest levels of stress.However, due to the nature of the wing loading the upper and lower wings are loadeddifferently (compression and tension respectively) and thus will be treated separately.


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Figure 6: Critical requirements of airframe components [7]

Wing upperDuring flight the wing is loaded in bending such that the upper surface is stressed incompression. Thus, to resist the applied stresses and minimise weight, it requires highratios of stiffness to density and yield strength (compression) to density. The upperwing also requires good resistance to stress corrosion cracking fracture. Based on thegeometry of the wing these constraints are given as:

ρ3

1E

,ρ

σ )(compys , ISCCK

Wing lowerThe bending load on the wing causes the lower surface to be stressed in tension. Thusto resist the applied stresses and minimise weight, it requires high stiffness to densityand yield strength to density ratios. The lower wing also requires good resistance tostress corrosion cracking fracture and good corrosion resistance. Due to the tensilenature of the stresses it also requires good fatigue strength and low fatigue crackgrowth rates. Based on the geometry of the wing these constraints are given as:

ρE

,ρ

σ )(tensionys , ISCCK , Corrosion Resistance, FSσ , Na

∂∂

FuselageThe fuselage carries the whole of the payload and is stressed under tension,compression, torsion, bending and pressurisation forces. Most of these forces placethe fuselage under tension. Thus to minimise weight it requires high ratios of stiffnessto density and strength to density. It requires good corrosion resistance and, due to the


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tensile forces, requires high fatigue strength and low fatigue crack growth rates. Basedon the geometry of the fuselage these constraints are given as:

ρ3

1E

,ρ

σ ys , Corrosion Resistance, FSσ , Na

∂∂

Spars, Frames and RibsThese lie under the skin and are used to distribute loads, retain the aerodynamicshape, and increase the buckling strength of the structure. These components areloaded in bending and thus their requirements are similar to those of the wing upperand lower.

ρE

,ρ

σ ys , ISCCK , FSσ , Na

∂∂

Landing GearsThe landing gears, as shown in Figure 7, are subjected to high static and cyclicloading. Thus they must be stiff and strong enough to withstand this loading and haveacceptable fatigue and fracture resistance. The components must also be resistant tostress corrosion cracking failure. In addition there is a volume constraint on the size ofthe components. Thus the constraints are given as:

E , ysσ , Na

∂∂

, ICK , ISCCK

Figure 7: Landing gears [6]


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Control SurfacesThese consist of the rudder, elevators, ailerons and flaps and are, in general, lightlyloaded. Thus they require structural stiffness and light weight.

Skin materialsThese require good corrosion resistance as they are under continuous atmosphericexposure. In supersonic aircraft, skin heating becomes an important factor and is mostpronounced on the nose and leading edges of the aircraft as shown in Figure 8. Table1 shows the degree of skin heating with increasing speed.

Figure 8: Skin heating across the airframe [8]

Table 1: Skin Temperature at 23 000m (-56°C) with Emissivity Factor 0.9

MACH NUMBER SATURATION TEMPERATURE(oC)

2.0 1002.5 1503.0 2003.5 3004.0 370

The materials that best fit the discussed constraints, and are thus used for airframeconstruction, are aluminium, titanium, steels and composites. From Figure 9, it can beseen that while aluminium is the most widely used material for civil transportairframe construction, composites are increasing their role in airframe construction.Military aircraft, with their higher performance requirements and higher budgets,show an even greater use of composite materials in the airframe (see Figure 10).

Figure 9: Civil transport airframe construction materials 1990 & 1995

Ti10%

CFC10%

Misc15%

Al65%

1990 1995

Ti12%

Al55%

CFC21%

Misc12%


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Figure 10: Military aircraft airframe construction materials

2.1 Aluminium Alloys

Aluminium alloys have a low density (2.7g/cm3) and, while their tensile properties arelow compared to steels, they have excellent strength to weight ratios. They haveexcellent thermal and electrical conductivity and have excellent resistance tooxidation and corrosion. The chief limitation of aluminium alloys is their low meltingtemperature (~660oC) that limits their maximum service temperature. However theyremain the major materials for civil airframe construction and are used for manyapplications, see Figure 11.

Figure 11: Materials in airframe construction [8]

Aluminium is alloyed with a range of alloying elements to tailor its mechanical andchemical properties. Thus they have a designation system based on:1) whether they are wrought or cast,2) their main alloying elements, and if wrought3) the applied thermal and mechanical treatmentsThe designations for wrought aluminium alloys are shown in Figure 12.

Ti12%

Al44%

Misc14%

CFC30%

Ti12%

CFC40%

Misc17%

Al31%

1990 1995


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Figure 12: Wrought aluminium alloy designation system [7]

2.1.1. Strengthening and Age Hardening

Aluminium alloys are strengthened in a number of ways including: solid solutionstrengthening, cold working, dispersion strengthening and age-hardening. Agehardening (otherwise known as precipitation hardening) is a process whereby a fineprecipitate structure is formed in the alloy matrix following an ageing heat treatment.For an alloy to be age-hardening it requires:

1) decreasing solid solubility with decreasing temperature2) the ability to suppress the formation of precipitates by quenching from solid

solution3) the formation of metastable coherent precipitates

A typical phase diagram for an age hardening alloy, and the ageing heat treatmentsequence are shown in Figure 13. The ageing process follows three main steps:

1) The solution treatment: The alloy is heated above the solvus temperature todissolve any precipitates and ensure the alloying elements are in solution.

2) Quench: The alloy is rapidly quenched. The alloying elements in solution do nothave time to diffuse and form precipitates. Thus, the alloying elements remain insolution forming what is known as a supersaturated solid solution (SSSS).


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3) Ageing: The alloy is heated to an intermediate temperature below the solvustemperature. The alloying elements are able to diffuse to form coherent precipitateclusters (known as GP zones).

Figure 13: Example age hardening 2XXX series aluminium alloy system [9]

The coherent precipitates increase the strength of the alloy by distorting the lattice andcreating resistance to dislocation motion. The number of precipitates increases withincreasing time thus increasing the strength of the alloy. However, with excessivetime the precipitates become large and incoherent and their strengthening effectdecreases. Thus, during ageing there are four main strengthening:

1) solid solution strengthening in the SSSS2) coherency stress hardening from the coherent precipitates3) precipitation hardening by resistance to dislocation cutting4) hardening through resistance to dislocation bowing between precipitates.

This results in the typical shape of the ageing curve, shown in Figure 14, where themaximum strength (peak aged) is achieved with an optimum dispersion of coherentprecipitates. Increasing ageing temperature allows the peak ageing temperature to bereached in shorter time. Ageing beyond the peak ageing time to form incoherentprecipitates is known as overaging.

Figure 14: A) The effect of ageing temperature and time upon strength [9] B) Agehardening strengthening effects [10]

A) B)


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The most commonly used aluminium alloys for airframe construction are the agehardening alloys in the 2XXX and 7XXX series, as shown in Figure 15. Examples ofthe tempers given to the airframe alloys are shown in Figure 16.

Figure 15: 2XXX and 7XXX series aluminium alloys used in civil transport aircraftairframes

2.1.2. 2XXX Series Aluminium Alloys

The 2XXX series aluminium alloys are alloyed with copper from 1.9 to 6.8% andoften contain additions of manganese, magnesium and zinc. Their precipitationhardening has been widely studied and they are used for applications such as:forgings, extrusions, liquefied gas storage, civil transport and supersonic aircraft.

These alloys have lower crack growth rates (lower Na

∂∂ ) and thus have better

Figure 16: Aluminium airframe alloy tempers [6]

2224


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fatigue performance, than 7XXX series alloys. Therefore, these are used on the lowerwings and fuselage (see Figure 17). The alloys used are 2224, 2324 and 2524 (bothmodified versions of 2224). These alloys are often clad with 99.34% pure aluminiumfor increased corrosion resistance. Compositions of these alloys are included in Table2.

Figure 17: 2XXX series alloys in the airframe

Table 2: Aluminium Airframe Alloy Compositions

Alloy Composition2618 Al – 2.3Cu – 1.6Mg2224 Al – 4.4Cu – 1.5Mg – 0.6Mn7050 Al – 6.2Zn – 2.3Cu – 2.2Mg7075 Al – 5.6Zn – 2.5Mg – 1.6Cu7150 Al – 6.4Zn – 2.4Mg – 2.2Cu

An example heat treatment for 2X24 alloys is shown in Figure 18. This ageingtreatment produces the ageing sequence:

SSSS → GPB(Cu,Mg)→ S'(Al,Cu,Mg) → S(Al2CuMg) coherent semi-coherent incoherent

2224


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Figure 18: Ageing heat treatment for 2x24-Al

As shown in Table 1, the skin temperature of supersonic aircraft increases withincreasing speed. For aluminium clad aircraft the speed must be kept below Mach 2.2to avoid the skin temperature reaching 150oC. This restriction is to avoid thealuminium alloy overaging and its mechanical properties degrading.

Another problem with high temperatures is the phenomena known as creep (seesection 3.1) Special alloys such as 2618 are used to provide adequate creepperformance in current supersonic aircraft.

2.1.3. 7XXX Series Aluminium Alloys

The Al-Zn-Mg system offers the greatest potential for age hardening (out of thealuminium alloys) though copper is often added to improve stress corrosion cracking(SCC) resistance (with the drawback of reducing weldability). SCC resistance(detailed in section 4.2) decreases with increasing Zn:Mg ratio as shown in Figure 19.The SCC problems have been the biggest restriction upon the use of these alloys butthey have still been used in lightweight military bridges, railway carriages, andmilitary and civil aircraft. The use of 7XXX series alloys in the airframe is shown inFigure 20, with the most common alloys being 7075, 7010, 7055 and 7150. All thesealloys are used in the overaged T7 temper as this temper provides the best resistanceto exfoliation corrosion and SCC.

An example T7 heat treatment for 7XXX-Al series alloys is shown in Figure 21. Thesolutionising and quench is followed by an ageing treatment and then a highertemperature overaging treatment. This treatment produces the ageing sequence:

AlSSSS → GP → η'(MgZn2) → η(MgZn2)

T1(Mg32(Al,Zn)49) → T1(Mg32(Al,Zn)49)

0

250

500

0 60 120 180

Tem

pera

ture

(o C)

Time (min)


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Figure 19: Zn:Mg ratio effect of SCC in Al-Zn-Mg alloys [7]

Figure 20: 7XXX series alloys in airframe construction

Figure 21: T7- heat treatment for 7XXX-Al

0

250

500

0 1000 2000 3000

Tem

pera

ture

(o C)

Time (min)


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2.1.4. Aluminium – Lithium alloys

These alloys have been developed especially for the aerospace industry with the focuson the low density of lithium (0.534 g/cm3). These alloys offer the attractive benefitsof being 10% lighter, 10% stronger and 10% stiffer than conventional aluminiumalloys, Figure 22. However, these alloys lack toughness and thus their use is verylimited except in specialty aircraft as shown in Figure 23.

Figure 22: Effect lithium on modulus of Al-alloys [7]

Figure 23: Al-Li alloys in the new Eurofighter (EF2000) [11]

2.2. Titanium Alloys

Titanium alloys are strong, stiff, corrosion resistant and have a low density (density ofpure Ti is 4.5g/cm3). Titanium alloys are stronger and stiffer than aluminium alloysand thus titanium components can be smaller in size than a comparable aluminiumcomponent. Thus, they are used in applications where volume is important, such aslanding gears and attachment points. Titanium alloys can also be used in applications


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where the temperature is too high for aluminium, such as near the engine or in high-speed aircraft. The biggest restrictions on titanium are its higher density than Al, andhigh cost (approximately seven times that of aluminium or steel).

The most common titanium alloy in airframe construction is Ti-6Al-4V, otherwiseknown as Ti-6-4. This is a two phase (α and β) alloy, as is shown in Figure 24, andprovides a good combination of strength, ductility, toughness and creep resistance.The dual phase nature of the alloy allows for a range of processing techniques andmicrostructures to tailor the final properties of the alloy as shown in Figure 25. Thereare a wide range of α− β titanium alloys available with various strengths andtemperature resistance. Other titanium alloys, such as Ti-6Al-4Sn-3.5Zr-0.5Mo-035Si-0.7Nb-0.06C (IMI 834) are used for higher temperature applications. Themaximum operation temperature of titanium alloys is approximately 600oC. Abovethis temperature creep and rapid oxidation occur.

Figure 24: Microstructure of Ti-6Al-4V. A) Widmanstatten α, B) Equiaxed α (white)and transformed β (Widmanstatten α, grey), C) Transmission electron microscopehigh magnification image of the structure in B). [7]

a)

b) c)


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Figure 25: Many possible thermal treatments for Ti-6Al-4V [7]

Titanium has been used for compressor blades and discs (Figure 26), fans, spacevehicles, storage tanks, undercarriage components, flap tracks, engine mountings andfasteners, and has the potential to be used for supersonic passenger aircraft skin andstructure (the proposed speed of Mach 2.4 is too high for Al-alloys).

Figure 26: Forged Ti-6Al-4V fan blades [7]


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2.3. Steels

Like titanium, steels are stronger and stiffer than aluminium alloys and thus are usedin applications where volume is important, such as landing gears, attachment points,gears and bearings . They are used in parts where the required tensile strength isgreater than can be supplied by Ti-alloys. The high density of steel is the limitingfactor in their use, which has been declining with time (now down to approximately10% of the structural weight). The most commonly used steels are ultra high strengthlow alloy steels, maraging steels and precipitation hardening (PH) steels.

2.4. Composites

Composites consist of two or more materials combined to give a material withproperties distinct from the original constituents. Composites can be designed toproduce a material with desired combinations of properties such as stiffness, strengthand density. Typically, composites consist of a matrix material and a reinforcingmaterial. The matrix and fibre materials may be metals, ceramics or polymers, but thecomposites used in airframe construction are fibre reinforced polymer matrixcomposites. These have the advantages of:1) high specific strength and stiffness2) tailored directional properties3) non-corroding in salt environments4) excellent fatigue resistance5) dimensional stability5) reduced number of parts required (compared to metal components)But they are susceptible to impact damage, moisture pick-up and lightning strikes,have a relatively high cost, do not yield plastically in regions of high stressconcentration and are subject to random property variation due to the nature ofcomposite manufacturing.

The use of these advanced composites in airframe construction has increasedsubstantially over the past few decades as evidenced in Figures 9 and 10. They areused as floor beams, doors, aerodynamic fairings and for control surfaces, such asrudders, elevators and ailerons, due to their low weight and high stiffness. Theseapplications can be seen in Figures 27 to 29 and Table 3.

Figure 27: Composite applications in the Boeing 767 [12]


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Figure 28: Composite applications in the Boeing 737-600 [13]

Figure 29: Composite applications in the Boeing 757 [13]

Table 3: Composite Applications in Selected Aircraft [14]CompositeComponent

F-14 F-15 F-16 F-18 B1 727 757 767 Lear Fan*(non production)

Doors ü ü ü ü ü üRudder ü ü ü üElevator ü ü ü üVertical Tail ü ü ü ü üHorizontal Tail ü ü ü ü ü ü ü üAileron ü üSpoiler üFlap ü üWing Box ü üBody üMiscellaneous Fairings Speed

BrakeSpeedBrake,

Fairings

Slats,Inlet

Fairings Fairings PropellerBlades


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2.4.1 Composite Properties

The properties of fibre reinforced polymer matrix composites depend on the volumefractions of fibre, length of the fibres and the orientation of the fibres with respect tothe applied load. The fibres provide virtually all the load carrying characteristics ofthe composite, most importantly strength and stiffness. For example the modulus ofelasticity of a unidirectionally reinforced fibre composite in the direction of the fibres(EC) is given by:

fFmfC EfEfE +−= )1(

where ff is the volume fraction of fibres, Ef is the modulus of the fibres and Em is themodulus of the matrix. However, for loading perpendicular to the fibres the modulusof the composite (EC’) is given by:

( )

+

−=

f

f

m

fC

Ef

Ef

E1

1'

The properties of a range of epoxy matrix composites are shown in Figure 30.

Figure 30: Properties of various composites [14]

2.4.2. Fibres

Fibre materials for polymer matrix composites include (in order of increasing cost):E-glass, aramid (eg., Kevlar, see Figure 31), carbon, alumina, silicon carbide andboron. The properties of the fibres are shown in Figure 32. The fibre materials allhave high specific strength and stiffness imparting high strength and stiffness to thecomposite.


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Figure 31: Structure of aramid (Kevlar) [13]

Figure 32: Properties of common reinforcing fibres [13]

The most widely used fibre reinforcement is E-glass, which is used as thereinforcement in fiberglass. Glass reinforced composites are used in aircraft windowsurrounds, storage compartments and flooring panels. However, the most commonfibre reinforcement for polymer composites in airframes is carbon fibre, and theresulting composites are known as carbon fibre reinforced polymers (CFRP). Carbonfibres are used because:1) they have a good combination of strength, stiffness and cost which can be varied

according to the manufacturing route.2) they retain their strength and stiffness at service temperatures3) they show good chemical resistance at room temperature


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4) their manufacturing process is well developed and relatively inexpensive.

Carbon fibres are available in continuous and chopped fibre forms, in diametersranging between 4 and 10µm and are often classified according to their tensilemodulus as: standard, intermediate, high and ultrahigh modulus. However, carbonfibres are difficult to handle and brittle and thus require a matrix material to be usedeffectively. Additionally, due to the fact that they are a high resistance conductor,CFRP parts must be fitted with copper lightning conductors to avoid explosion ifstruck by lightning.

2.4.3. The Matrix

The purpose of the matrix is to support the fibres in the required position, transferload between the fibres, increase the toughness of the composite and protect the fibresfrom damage. While the longitudinal tensile properties are dominated by the fibres,the properties of shear, compression and transverse tension are dominated by thematrix properties. Matrix polymers can be thermoplastics or thermosets.Thermoplastics are fully polymerized materials that are solid at room temperature butmay be melted and shaped at high temperature. Thermosetting resins consist of a baseand a curing agent. When mixed together they react to form a heavily crosslinkedsolid resin that cannot be reshaped once set.

Traditionally the most common matrices are thermosets with the most commonmatrix being either polyesters or vinyl esters, both used in fiberglass products.However, the most common polymer matrix employed in airframe construction isthermosetting epoxy resin (see Figure 33). Epoxy resin is used because it has goodadhesion to fibres, good resistance to water and high mechanical properties.Additionally, manufacture of epoxy based composites is well developed and wellunderstood. Second generation epoxies such as rubber toughened epoxy resin arecurrently being used as composite matrices to increase matrix toughness and tocompete with advanced thermoplastics such as PEEK (polyetheretherketone, seeFigure 34) PPS (poly-phenylene sulfide) and PEI (polyetherimide) which generallyhave increased temperature resistance.

Figure 33: Structure of epoxy resins A) DGEBPA, B)TGDDM [6]

Figure 34: Structure of PEEK [6]

A)

B)


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2.5. Polymers

Polymers (plastics) are also used in parts of the airframe, namely the windows, lightlystressed parts, interior trim and as electrical insulators. Windows require bothtransparency and acceptable mechanical properties due to the pressurised nature of theaircraft. Usually the window consists of three layers:1) a glass load bearing layer2) a back-up glass load carrying layer in case the first layer cracks3) a plastic shield on the inside to prevent scratching by passengers. These plasticshields are often acrylics (e.g. perspex) although polycarbonates are also used.

3. THE ENGINE

The main components and design of a gas turbine engine are shown in Figure 35.

Figure 35: Gas turbine engine [6]

One of the main differences in material requirements for engine components asopposed to airframe components is the operating temperature. This is due to the hightemperatures encountered inside the engine, as shown in Figure 36. Thus the primaryrequirements of engine materials are:

1) high specific strength (strength / weight ratio) at the relevant operating temperature2) creep resistance3) oxidation/corrosion resistance4) microstructural stability at high temperature5) low density6) high stiffness7) good fabricatability8) acceptable cost9) reproducible performance


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Figure 36: Turbine engine temperature profile [6]

3.1 Creep

Creep is a process whereby the material permanently strains as a function of stresstime and temperature. The creep process begins when the absolute temperature isgreater than 0.3 – 0.5 of the melting temperature measured in Kelvin. Thus, lowmelting point materials such as lead (and also polymers since their glass transitiontemperature is close to room temperature) will often creep at room temperature. Thegeneral shape of creep curves (strain as a function of time) is consistent across metalsceramic and polymers (see Figure 37). In metals creep occurs by dislocation climband grain elongation, both aided by atomic diffusion. To avoid catastrophic failure inthe engine the materials selected must have adequate creep performance at theoperating temperature. This involves the material having:

1) a high melting point2) obstacles to dislocation motion (solid solution and precipitates)3) lattice resistance to creep (e.g. covalently bonded materials)4) large grain size to increase diffusion distances and minimise grain boundary

diffusion5) grain boundary precipitates to avoid grain boundary sliding

Figure 37 : Creep curves [15]


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3.2 Oxidation

Oxidation involves material reacting with oxygen. It is also known as scaling or drycorrosion. Most metals react with oxygen at room temperature. However the rate ofreaction at room temperature is low and the oxide layer that forms on the surface actsas a barrier to further oxidation. For example, from thermodynamic data it is predictedthat aluminium will rapidly bulk oxidise in air. However, the rapid formation of anatural oxide (alumina, Al2O3) film on the surface protects the aluminium fromfurther oxidation in most environments.

However, at high temperature the rate of material oxidation is rapid. The oxide filmgrows at a parabolic rate consuming the base metal. The rapid oxidation is the resultof the increased diffusion rates at increased temperatures and the decreasedeffectiveness of the oxide layer as a barrier. Factors that improve the effectiveness ofthe oxide barrier at high temperatures are a high melting point and high electricalresistivity of the oxide layer. Thus Al2O3 and Cr2O3, which have high melting points(low diffusion rates) and high resistivity, are effective barriers to oxidation.

Another consideration is the volume of the oxide layer with respect to the base metal.If the oxide has less volume than the metal from which it was formed, then it willcrack and expose new metal to the atmosphere. If the volume of the oxide is largerthan the metal from which it formed it will buckle and spall, Figure 38. Both thesesituations lead to constant exposure of the atmosphere to unprotected metal and therates of oxidation will be high. Thus for high temperature oxidation resistance:1) the protective oxide layer must have a high melting point and have low electrical

conductivity2) the oxide volume must be similar to that of the metal from which it formed.

Figure 38: Breakdown of oxide films [16]

The problem with oxidation is that the formation of the oxide decreases the thicknessof the base metal. If oxidation is rapid, then the metal structure will rapidly losethickness, and thus load carrying capability, and will fail under service conditions

3.3 Engine Materials

The materials used in the turbine engine are shown in Figure 39. The differences inmaterial are mainly related to the different operating temperatures. In the forwardsection, where the temperature is low to medium, titanium parts are often used. In thehigh temperature rear combustion areas nickel based superalloys and some ceramicsare used. The outer casing experiences low temperatures and thus aluminium andcomposites are suitable materials. The materials used are summarised in Table 4.


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Figure 39: Turbine engine materials

Table 4: Materials Used in the Gas Turbine Engine [6,17]Component Materials

Fan Blades Superplastically formed titanium alloyskin diffusion bonded to a titaniumhoneycomb core

Compressor Vanes, Discs and Blades With Increasing temperature the materialschange from titanium alloys to 12% Crsteel to nickel based superalloys

Combustor High temperatures means nickel basedsuperalloys are used

Turbine Blades Discs and VanesFigure 40 shows a turbine blade

High temperatures means nickel basedsuperalloys are used (see sections 3.4 and3.5)

Shafts The shafts are made of steelCasings Aluminium alloys and composites

3.4. Nickel Based Superalloys

Nickel based superalloys were developed from Ni-Cr heating elements. Nickel basedsuperalloys possess high strength and toughness at high temperature, creep resistanceup to 1000oC and corrosion resistance making them ideal materials for turbineengines. The strength of various aircraft materials with increasing temperature isshown in Figure 41. The development of superalloy microstructure and compositionover time is shown in Figures 42 & 43. The function of the alloying elements used inthese alloys is shown in Table 5. Superalloy microstructures are shown in Figure 44

.

CompositeAluminiumTitaniumSteelNickelKevlar


29

Figure 40: A single crystal turbine blade [6]

Figure 41: Strength of aircraft materials with increasing temperatures [6]

Figure 42: Superalloy alloying elements over time [17]


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Figure 43: Microstructure of Ni based superalloys over time [17]

Table 5: Function of alloying elements in Ni based superalloysAlloying Elements FunctionCr Oxidation resistanceAl, Ti, Nb Strengthening from ordered γ’ precipitatesCr, Mo, W, V Carbide strengtheningCr, Mo, W Solid solution strengtheningHf Grain boundary pinning

Figure 44: Ni based superalloy microstructures. A) Carbides at the grain boundaryand γ’ precipitates in the matrix B) Superalloy aged at two temperatures to producelarge and small cubic γ’ precipitates [9]

The combination of strengthening techniques from all the alloying elements gives thesuperalloys their outstanding high temperature properties. The solid solutionstrengthening lasts to high temperatures and the large atomic species used meansdiffusion is slow and creep is retarded. The carbides pin both dislocations andboundaries, strengthening the alloy and preventing creep. The γ’ precipitates that formare coherent and of the form Ni3Al or Ni3Ti. The low surface energy of the


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precipitates minimises overaging and thus provides good strength and creep resistanceat high temperatures. By varying ageing temperatures, combinations of small andlarge precipitates can be formed giving higher fractions of precipitate and increasingstrength.

3.5 Turbine Blade Design

Turbine blades operate at 1400oC and are required to have a service life of 10 000-20000 hours. The techniques used to meet these requirements are:1) the use of nickel based superalloys (containing many precipitates, solid solution

atoms and having high oxidation resistance)2) the blades are cast a single crystal. This means that there are no grain boundaries

within the structure and thus minimises creep see Fig 453) the blades are internally cooled to allow increased operating temperatures. Typical

cooling design features are shown in Figure 464) the blades are coated to increase the oxidation resistance as shown in Figure 47.

A timeline of techniques to raise engine operating temperatures is shown in Figures48 & 49

Figure 45: Turbine blades cast to promote different grain structures [15]

Figure 46: Blade cooling techniques [16]


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Figure 47: History of turbine blade coating systems [17]

Figure 48: Developments to increase engine operating temperatures [16]

Figure 49: Developments to increase turbine metal capability [9]


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3.6 Turbine Blade Coatings

There are two types of coating for turbine blades. The first are metallic coatings suchas chromium or aluminium or complex NiCoCrAlY alloys. These provide corrosionand oxidation protection by forming protective Cr2O3 and Al2O3 surface oxide layers.The second type of coating is the thermal barrier coating (TBC) as shown in Figure50. The superalloy substrate provides the structural strength. The ceramic coatingprovides a thermal barrier between the superalloy and the hot combustion gas whilethe metallic coating provides oxidation and corrosion protection and provides asurface for the ceramic layer to adhere to.

Figure 50: Thermal barrier coatings

4. CORROSION

Corrosion control is of extreme importance to the aircraft industry because of itspotential impact on human safety and expensive aircraft structures. Thus, corrosionresistance is a vitally important factor to consider when choosing aircraft materials .Corrosion is experienced as aircraft may be in service for up to 30 years in exposureto many varied environments: sub zero temperatures, high humidity, tropicalconditions, rain, salt spray, ice, UV radiation, atmospheric oxygen and pollutants. Forthe airframe, corrosion in the form of stress corrosion cracking, pit, and crevicecorrosion are the most important factors. For engine components, high temperatureoxidation is the most important consideration due to the high temperature operatingconditions.

Corrosion resistance is often improved by adding a coating to the material to protectthe surface. For example; the anodising of aluminium promotes a thick aluminacoating on the surface of the material. This both protects the aluminium from theatmosphere and corrosive chemicals. Chromate conversion coating of aluminium is asimilar technique widely used in aircraft structures, again to prevent corrosion. Themain types of aircraft corrosion are detailed below.

TBC System

Ceramic Top Coat:Provides - thermal insulation

Metallic Bond Coat :Provides - oxidation / corrosionprotection - surface for ceramic to

Superalloy Substrate:Provides - strength

CoolingAir


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4.1. Surface Corrosion

Surface corrosion is a superficial, uniform corrosion that is only objectionable froman aesthetic viewpoint. Dulling of a bright surface or discolouration are examples ofthis. However, surface corrosion can be an indication of a protective barrier break-down and thus should be examined closely to avoid the development of more seriouscorrosion.

4.2. Stress corrosion cracking (SCC)

This is a process whereby a typically ductile alloy fails in a brittle manner whensubject to the simultaneous effects of a tensile surface stress and a corrosiveenvironment – neither of which would cause major damage if acting alone. Wherealternating stresses are involved, it is known as corrosion fatigue. The cracks growfrom stress induced flaws (therefore a threshold stress is required) but may haveorigins in features such as corrosion pits. The stress does not need to be appliedexternally, but may be the result of residual manufacturing stresses or thermalstresses. SCC is the major source of corrosion failure in thick aircraft structures suchas thick plate and forgings. In aluminium alloys, the cracking is intergranular and onlyoccurs in alloys when appreciable amounts of solute elements such as Cu, Mg, Si, Znand Li are present after certain heat treatments are applied. The factors thought tocontribute to the SCC effect in Al-alloys include:

1) precipitate free zones at the grain boundaries2) peak aged microstructures (due to Guinier Preston / GP zones)3) dispersion of precipitates at the grain boundaries4) differences in solute concentrations at boundaries5) hydrogen embrittlement at the grain boundaries6) chemisorption of atomic species at the crack tip

Titanium was considered immune to SCC for a long time though it has been shown tobe susceptible in specific environments. However, titanium SCC failures in serviceare rare. Nevertheless, care must be taken during manufacture and overhaul oftitanium parts as some alloys are embrittled by common degreasing solvents (such asorganic chlorides).

4.3. Galvanic corrosion

This occurs when two different metals of different galvanic potential are in contact inthe presence of water (electrolyte). The more anodic material will corrode at anaccelerated rate resulting in a build up of corrosion product near the contact area.Galvanic corrosion can be a problem for aluminium alloys as it is anodic to mostother structural metals. However, the occurrence of galvanic corrosion depends onfactors other than just the electrode potential. Impurity elements and alloyingelements can pose problems in aluminium alloys. At localised regions of high or lowalloying element concentration, localised regions of high and low corrosion resistance(electrode potential) form. This leads to pitting type corrosion in the areas where thecorrosion resistance is lowest. In addition, segregation of certain elements can lead tointergranular corrosion (where the corrosion follows the grain boundaries) andexfoliation where corrosion products force surface layers and grains to delaminate.


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4.4. Crevice corrosion

This is where local differences in electrolyte ionic concentrations on the metal surfacecause corrosion to occur. In the crevice or gap in a structure the electrolyte is deprivedof oxygen and thus oxidation reactions occur within the crevice. An example of this isshown in Figure 51. Crevice corrosion often occurs at rivets and metal joints. Crevicecorrosion is the major corrosion problem with titanium alloys.

Figure 51: Crevice corrosion [15]

4.5. Pitting Corrosion

Pitting corrosion is a form of localised corrosion in which small pits form in thematerial surface, see Figure 52. Pitting can occur due to chemical variations withinthe base material or due to a crevice corrosion type effect in preexisting flaws. Anexample of pitting corrosion occurs when a water droplet shields the underlying metalfrom oxygen. It is most common in marine environments as Cl- ions locally attackprotective oxide layers. Titanium is stable in most corrosive environments but issubject to pitting corrosion in halide containing aqueous solutions at high temperature.

4.6. Exfoliation

Exfoliation corrosion, shown in Figure 53, occurs in high strength aluminium alloyswhere the grains have become elongated and flattened during processing. Thecorrosion is intergranular (occurs along grain boundaries) and proceeds along planesparallel to the surface. The corrosion products, which have greater volume than themetal from which they formed, then cause delamination of the surface metal. This isthe main corrosion problem in airframe sheet materials.


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Figure 52: Pitting corrosion [16]

Figure 53: Exfoliation corrosion [7]

5. NON DESTRUCTIVE TESTING

The critical safety requirement of aircraft components means that they must beregularly tested for flaws. To avoid damaging the components during testingspecialised non-destructive testing (NDT) techniques are used. The most commonnon-destructive testing techniques are:

5.1 Liquid Penetrant Inspection

This is a simple method of testing for surface defects and may be used on complexshapes. The piece to be tested is first cleaned and dried. Then the penetrant is appliedto the surface of the material. The penetrant is a highly wetting liquid capable of beingdrawn into small cracks and is often brightly coloured or fluorescent to enable easydetection. After a period of time to allow the liquid to be drawn into cracks, the excesssurface penetrant is removed. A thin film of absorbent material, the developer, is thenadded to the surface. The developer draws the penetrant from defects back to thesurface where it can be seen and detected by visual inspection (Figure 54). However,the technique does not detect sub-surface flaws and coatings can often preventdetection.


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Figure 54: Liquid penetrant inspection: a) the penetrant is applied, b) excesspenetrant is removed leaving penetrant in crack, c) developer applied drawingpenetrant out and delineating the crack [18]

5.2 Magnetic Particle Inspection

This testing is based on the principle that the magnetic field around a magnetisedferromagnetic material (such as steel) will be distorted in the vicinity of defects(Figure 55A). To test the component it is first cleaned and then a magnetic field isapplied. Once magnetised, magnetic particles are applied to the component surface.These are attracted to sites where the magnetic field is distorted and thus delineate theflaws. Both surface and sub-surface flaws can be detected. The surface is then re-cleaned and de-magnetised. As flaws parallel to the magnetic field will not bedetected it is important to test the material in a range of magnetic field orientations(Figure 55B).

5.3 Ultrasonic Inspection

Ultrasonic inspection involves sending high frequency sound waves through amaterial. The transmitted or reflected sound waves are monitored and interpreted(Figure 56). The surface of the component is coated with a coupling agent to allowtransmission of the ultrasonic waves from a sending transducer. The sendingtransducer emits the ultrasonic vibrations into the part while a receiving transducermeasures the reflected or transmitted vibrations, which are then displayed on anoscilloscope. The received vibrations from a defect will appear different to those froman unflawed part. If the receiver is on the same side of the component as thetransmitter then it measures the vibration echoes and is known as the pulse-echotechnique. If the transmitter and receiver are on different sides of the component thereceiver measures the transmitted vibrations and is known as the through transmissiontechnique. While ultrasonic inspection has high sensitivity to flaws and is a rapidtesting procedure, it can be difficult to use with complex shapes and extensiveoperator training may be required.


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Figure 55: A) Effect of defects on the induced magnetic field. B) Crack detection indifferent magnetic field orientations [18]

Figure 56: Ultrasonic testing; A) transmission and echo, B) oscilloscope output [9]

5.4 Radiography

Radiography covers X-ray, gamma ray and neutron beam inspection. It involvesmeasuring the differential penetration and absorption of radiation through a material.The radiation penetration is recorded on photographic film. The extent of exposure ofthe film is proportional to the amount of radiation passing through the component(Figure 57). By examining the exposure of the film it is possible to detect thecomposition and thickness of the component as well as the presence of flaws (egcracks do not absorb radiation and thus in these regions the radiation has highpenetration). This is the most expensive NDT method, due to the expensive radiation

A) B)

A) B)


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sources, film and processing. In addition it is difficult to apply to complex shapes andsafety precautions are necessary when dealing with radiation. However the techniqueis good for examining the internal regions of a material.

Figure 57: X-ray radiography [9]

5.5 Other Techniques

There are a range of other NDT techniques used in the aerospace industry, such as:

5.5.1: Eddy Current InspectionThis involves measuring the surface currents produced in conductive material whenbrought near an alternating current coil. It is possible to detect surface and nearsurface flaws with high speed and low cost. However the response is sensitive to anumber of variables so interpretation can be difficult and extensive training isrequired.

5.5.2 Acoustic Emission MonitoringThis involves measuring the high frequency sounds emitted when a material deformsor cracks. In critical components constant surveillance is possible to give warning ofimpending danger. However, the technique can only detect growing flaws.

5.5.3 ThermographyThis involves monitoring the temperature of a material as it undergoes a temperaturechange. In the vicinity of flaws the rate of heat flow will be altered and thus the flawcan be detected. Similar techniques are used to detect the presence of unwantedwater/ice in aircraft composite structures.


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6. REFERENCES

1. Boeing Aircraft Company Website : http://www.boeing.com, 2001

2. Janes All the Worlds Aircraft: 1999-2000 Edition, Ed. P.Jackson, Surrey, UK,1999

3. Cessna Aircraft Company Website : http://www.cessna.com, 2001

4. Janes All the Worlds Aircraft: 95-96 Edition, Ed. P.Jackson, Surrey, UK, 1995

5. Airliners Website : http://www.airliners.net, 2001

6. Flower, H.M. ed., High Performance Materials in Aerospace, Chapman &Hall, London, 1995

7. Polmear, I.J., Light Alloys: Metallurgy of the Light Metals, Arnold, London,1995

8. Charles, J.A., Crane, F.A.A. & Furness, J.A.G., Selection and Use ofEngineering Materials: Third Edition, Butterworth Heinemann, Oxford, 1997

9. Askeland, D.R., The Science and Engineering of Materials: Third SI Edition,Chapman and Hall, London, 1996

10. Ashby, M.F. & Jones, D.R.H., Engineering Materials 2: An Introduction toMicrostructures, Processing and Design, Pergammon Press, Oxford, 1994

11. Interavia, December 1997 ‘Eurofighter Gets the Go’

12. Niu, C-Y. M, Airframe Structural Design: Practical Design Information andData on Aircraft Structures, Conmilit Press Ltd, Hong Kong, 1988

13. Middleton, D.H. ed., Composite Materials in Aircraft Structures, LongmanScientific & Technical, UK, 1990

14. Engineered Materials Handbook Volume 1: Composites, Ed. T.J. Reinhart,ASM International, Ohio, 1987

15. Callister, W.D., Jr, Materials Science and Engineering an Introduction: 5th

Edition, John Wiley and Sons Inc, New York, 2000

16. Ashby, M.F. & Jones, D.R.H., Engineering Materials 1: An Introduction totheir Properties and Applications, 2nd Edition, Butterworth Heinemann,Oxford, 1997

17. Sims, C.T., Stoloff, N.S. & Hagel, W.C. eds., Superalloys II, John Wiley &Sons, New York, 1987


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18. Degarmo, E.P., Materials and Processes in Manufacturing, MacmillianPublishing Company, New York, 1988

19. Schlenker, B.R., Introduction to Materials Science: SI Edition, John Wiley &Sons, Milton, 1974

20. Cutler, J., Understanding Aircraft Structures: Third Edition, BlackwellScience, Oxford, 1999

21. Megson, T.H.G., Aircraft Structures for Engineering Students: Third Edition,Arnold, London, 1999

22. ASM Handbook Volume 20: Materials Selection and Design, Ed. G.E. Dieter,ASM International, Ohio, 1997

23. Metals Handbook Ninth Edition: Volume 13: Corrosion, Eds. L.J. Korb &D.L. Olson, ASM International, Ohio, 1987

7. INTERNET LINKS

7.1 General aerospacehttp://www.howstuffworks.com/sc-aviation-transportation.htm

http://www.matweb.com/

http://www.howstuffworks.com/airplane.htm

http://www.howstuffworks.com/search/index.htm?words=materials+engineering

http://smc.larc.nasa.gov/coe/

http://as.wm.edu/Nondestructive.html

http://www.grc.nasa.gov/WWW/HSR/EPMAirf.html

http://www.grc.nasa.gov/WWW/HSR/index.html

http://www.boeing.com/commercial/aeromagazine/aero_07/corrosn.html

http://www.boeing.com/commercial/747family/background.html

http://www.boeing.com/commercial/747family/index.html

http://www.boeing.com/news/feature/new747/passenger.html

7.2 Aluminiumhttp://www.aluminum.org/default.cfm/0/2

http://www.capral-aluminium.com.au7.3 Titaniumhttp://www.ife.no/media/446_FactsTiAlloy.pdf

http://www.ife.no/media/485_Faktaark_oversikt21.pdf


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http://www.ramcast.com/whytitan.htm

7.4 Compositeshttp://www.compositesone.com/product.html

http://www.aero.usyd.edu.au/wwwcomp/cssgmain.html

http://www.dupont.com/cgi-bin/corp/proddbx.cgi

http://www.composite-solutions.com/FAQinfo.htm

http://www.advancedcomposites.com

http://www.compositetek.com/benefits.htm

7.5 Turbines and materialshttp://www.gas-turbines.com/begin/index.htm#HISTORY

http://www.howstuffworks.com/turbine.htm

7.6 Corrosionhttp://nace.org/nace/index.asp

http://www.corrosionsource.com

8. SYMBOLS

E Youngs Modulus

Ef Youngs Modulus, Fibres

Em Youngs Modulus, Matrix

Ff Volume Fraction of Fibres

ISCCK Stress Corrosion Cracking Fracture Toughness

ρ Density

σy Yield Stress

FSσ Fatigue Strength

Na

∂∂

Fatigue Crack Growth Rate, Crack growth (a) per cycle (N)

Documents

Aeronautical Materials