Fatigue and Damage Tolerance of Glare

Applied Composite Materials 10: 223–242, 2003.© 2003 Kluwer Academic Publishers. Printed in the Netherlands.

223

Fatigue and Damage Tolerance of Glare

R. C. ALDERLIESTEN, M. HAGENBEEK, J. J. HOMAN,P. A. HOOIJMEIJER, T. J. DE VRIES and C. A. J. R. VERMEERENFaculty of Aerospace Engineering, Delft University of Technology, Kluyverweg 1, 2629 HS Delft,The Netherlands. e-mail: [email protected]

Abstract. Methods have been developed to describe the fatigue initiation and propagation mecha-nisms in flat panels as well as mechanically fastened joints and to determine the residual strengthof large flat panels. Glare shows excellent crack growth characteristics due to the mechanism ofdelamination and fibre bridging. The fatigue insensitive fibres restrain the crack opening and transferload over the crack in the metal layers. During the initiation phase fibre bridging does not occur andthe behaviour is dominated by the metal initiation properties. Mechanically fastened joints introduceadditional effects such as secondary bending, load transfer and aspects related to the fastener installa-tion. The residual strength of Glare is dependent on the amount of broken fibres and the delaminationsize and can be described with the R-curve approach.

The impact resistance of Glare is related to the aluminium and glass/epoxy properties and issignificantly higher than the impact resistance of monolithic aluminium. The same has been provenfor fire resistance. Depending on the Glare grade and thickness, the outer aluminium layer will meltaway, whereas the other layers will remain intact due to carbonisation of the glass/epoxy layersand delamination of the laminate. The air in the delaminations will act as insulation, keeping thetemperatures at the non-exposed side relatively low.

Key words: fatigue, Fibre Metal Laminates, Glare, impact, joints, burn-through, lightning strike.

Nomenclature

D = Fastener diameter.

γ = Load transfer at the fastener row to be analysed (0 � γ � 1).

Kt,oh = Stress concentration factor of an open hole in a finite width sheet.

Kt,ph = Stress concentration factor of a pin loaded hole in a finite width sheet.

Kt,b = Stress concentration factor of an open hole in a finite width sheet,loaded in out-of-plane bending.

kb = Ratio of the bending stress due to secondary bending and the re-mote gross stress. The ratio follows from the neutral line model. Formore complex structural details the ratio can be determined using FEcalculations.

Sp = Fastener pitch.

224 R. C. ALDERLIESTEN ET AL.

1. Introduction

The fatigue and damage tolerance aspects of Glare have been extensively investi-gated at Delft University of Technology for the last decades. Methods were devel-oped and validated for fatigue initiation, crack propagation and residual strength.Furthermore, impact properties and fire resistance have been determined by meansof tests as well as its behaviour under lightening strike. This paper provides anoverview of the current knowledge of Glare with respect to these properties.

2. Fatigue Initiation

Fatigue crack initiation occurs mainly in the aluminium layers in Glare, analogueto that of monolithic aluminium. A similar stress state in the aluminium layers inGlare and in the monolithic material will have the same initiation life. The fatiguecrack initiation behaviour of the aluminium layers in Glare can, therefore, be com-pared to the behaviour of monolithic aluminium if the actual stress conditions atthe notch edge are equivalent for both materials.

The actual stresses in the aluminium layers in Glare consist of stresses due tothe curing process and stresses due to the external loading, see Figure 1. Internalstresses will occur during the curing process as a result of the different coefficientsof thermal expansion between the layers. This stress system depends on the lay-upand type of Glare. For each Glare laminate, the residual stresses in the aluminiumlayers can be calculated as a result of the curing process.

The actual stresses in the aluminium layers are affected by the different stiff-nesses of the various Glare constituents. This means that the aluminium layers

Figure 1. The stress cycle in the aluminium layers in Glare 3-3/2-0.3 L at room temperatureis a superposition of the stress cycle induced by the applied stress and the curing stress [1].

FATIGUE AND DAMAGE TOLERANCE OF GLARE 225

attract more stress than the fibre layers due to their relatively higher stiffness. Thestresses in the aluminium layers will, therefore, be higher than the applied stresses.

By superposition of the curing stresses and the stresses induced by the appliedfatigue load the actual stresses in the aluminium layers of Glare are obtained.

Using the actual stresses in the aluminium layers, crack initiation can be deter-mined in the same way as for monolithic aluminium. The S −N data of monolithicaluminium, i.e. the number of cycles to failure, are recalculated to S − N i data,which give the number of cycles to crack initiation in Glare based on the actualstresses.

3. Fatigue Crack Propagation

Several crack configurations can be distinguished in Glare, see Figure 2. The worstcase is the configuration containing cracks in al aluminium layers of equallength (a), while the least and most common case is the crack in one outer layer (b),induced by a scratch. In the case of secondary bending in joints, a part throughcrack configuration as depicted in Figure 2(c), can occur.

During the crack propagation phase a significant part of the fatigue stresses istransferred over the cracks in the aluminium layers by the intact fibre layers. This

(a)

(b)

(c)

Figure 2. Through-the-thickness crack distribution in a FML: through crack (a), surfacecrack (b) and a part through crack (c).


bridging mechanism restrains the crack opening and results in lower and approx-imately constant values of the stress intensity factor at the crack tip, compared tomonolithic aluminium [2].

The load transfer from the aluminium to the fibre layers induces cyclic shearstresses at the fibre-metal interfaces. These cyclic shear stresses cause shear de-formation and delamination growth between the aluminium and fibre layers. Thecrack opening occurs due to shear deformation of the prepreg layer, correspondingwith the cyclic shear stresses, and due to delamination at the interfaces [3].

As a consequence of the delamination growth, the length over which the fibresare elongated increases, resulting in reduction in fibre bridging stress. The bridgingstress, related to the cyclic shear stress, is on the other hand the driver behindthe delamination growth. In other words, there is a balanced mechanism betweendelamination growth at the fibre-metal interface and the bridging stress.

Thus the crack growth and the delamination growth form a balanced mechanismin which both processes are continuously influencing each other during the fatiguelife.

The effect of variable amplitude loading on the crack growth behaviour of Glarewas investigated by means of (repeated) single overloads and full load spectra.Analogue to the monolithic aluminium behaviour the crack shows short crackgrowth increment after which crack retardation occurs. After certain number ofcycles the crack growth rate returns to the level before the overload was applied.Research [4] revealed that the crack growth retardation is mainly due to plasticdeformation in the aluminium layers.

Between the three crack configurations in Figure 2, several differences in crackgrowth behaviour can be distinguished.

3.1. CRACK GROWTH RATE

Due to the intact layers in surface and part through cracks less stress will be trans-ferred through the net section in the cracked aluminium layers. As a result thecrack growth rate in part through cracks is lower than for through cracks, whilesurface cracks show the lowest crack growth rate in case of comparable loadingand geometry parameters [3, 5].

3.2. STARTER NOTCH EFFECT

Due to the cut fibres in the notch for through cracks and part through cracks, thecrack bridging is reduced, resulting in larger crack growth rates. For surface cracksinduced by a scratch this effect is absent, since no fibres are cut.

3.3. SMALL CRACKS

At the beginning of the crack propagation phase the crack opening is very small,which means that the fibres are not yet fully loaded. A certain amount of crack


length is needed before the fibre bridging of the crack becomes fully effective. Thiseffect is best visible for through cracks and the least visible for surface cracks.

3.4. SECONDARY BENDING

Compared to the through crack configuration, surface cracks and part throughcracks have the additional factor of secondary bending. The neutral line of thelaminate will shift due to the asymmetry created by the broken surface sheet,which induces secondary bending. This effect is relatively small for a surface crackconfiguration, but for part-through cracks this may be important.

3.5. TEMPERATURE AND MOISTURE

Two aspects are especially important for the durability characteristics of Glarewith respect to fatigue crack propagation: influence of temperature and moistureabsorption [6].

The effect of moisture on the material can be divided in three main effects:

• corrosion and more rapid crack growth rates in aluminium;• decrease of adhesive interface strength between fibre layer and aluminium;• plasticising of the matrix.

The decrease of the adhesive interface strength results in an increase of delami-nation growth, while the load transfer will be negatively affected by the weakeningof the matrix as well.

Higher temperatures negatively affect the fatigue crack growth behaviour ofGlare. This is mainly due to two factors.

With increasing temperatures the moisture absorption will also increase. Theelevated temperatures cause voids and microscopic defects to open and attractmore fluid. Another factor is the softening and weakening of the matrix due tothe increase of temperatures.

4. Fatigue in Mechanically Fastened Joints

4.1. CHARACTERISTICS OF MECHANICALLY FASTENED JOINTS

Aircraft structures do have a lot of mechanically fastened joints. The same type ofjoints can and will be used in aircraft fuselage applications with Glare. Such jointsare highly loaded due to local pin loading and load path eccentricity. In pressurizedcabins they are therefore prone to fatigue cracking due to the pressurization cyclein each flight.

4.2. LOAD TRANSFER

Loads in joints are transferred from one sheet to the other by means of fasteners.Depending on the type of fastener, the load can be transferred either by pin loadingor by friction. In most cases a combination of both will be seen.


It is possible to calculate the load transfer per fastener row in a joint. The cal-culations are based on deflection compatibility and require the fastener flexibilityas an input. The load transfer per row further depends on the stiffness of the sheetmaterial. By assuming that load transfer by friction can be neglected, the modelbecomes similar to the Jarfall model [7]. The choice of fasteners determines thefatigue life to a large extent. The following aspects are of importance:

– Clamping forces (these forces determine the amount of transfer by friction).– Fastener flexibility (fastener flexibility in Glare is subject of on-going research

at the TU Delft).– Installation method. Müller [8] showed that for riveted joints the fatigue life

will be significantly improved by increasing the squeeze force during theriveting process.

4.3. SECONDARY BENDING

Load path eccentricities in joints cause bending moments that lead to out-of-planedeformation of sheets: the so-called secondary bending. This secondary bendingcan locally lead to very high stresses at critical locations in the sheet (i.e. at theholes in the mating surfaces) and are therefore of great importance for the fatiguebehaviour of a joint. The amount of secondary bending is affected by:

– Joint geometry: fastener pitch, fastener row distance, number of fastenersrows, sheet thickness, etc.

– Mechanical properties of the skin material.– Stiffness of the surrounding structure.

By modelling the neutral axis the bending in a joint can be simulated. This hasfirst been done by Schijve [9] for a two-row riveted lap joint and is extended byMüller [8] and De Rijck and Fawaz [10]. Pellenkoft [11] showed that this “NeutralLine” model is also applicable to a large extent for Glare.

4.4. FATIGUE INITIATION IN MECHANICALLY FASTENED JOINTS

As discussed in the previous section, fatigue initiation in Glare can be treated asan aluminium problem. It seems therefore obvious to select an aluminium ap-proach for tackling the fatigue initiation in joints issue. Such an approach hasbeen formulated by Homan and Jongebreur [12]. According to this approach thefatigue initiation life of an arbitrary mechanically fastened joint can be predictedby performing the following procedure:

1. Calculate the stress amplitude and the stress ratio at the hole perimeter. Thisgives the so-called peak stresses at the hole.

2. Start equating with an available reference joint: equal peak stress results inequal fatigue initiation life.


3. Calculate from the peak stresses the remote stress amplitude for the referencejoint.

4. Determine the fatigue initiation life from the S − N curve for reference joint.This fatigue life is then identical to the fatigue life of the analysed joint. Ifnecessary, make a correction for the stress ratio.

The peak stress at a hole can be calculated using the following equation:

σpeak = Kt · σnet = Kt · Sp

Sp − D· σremote.

The stress concentration factor Kt at the hole perimeter is defined as:

Kt = (1 − γ )Kt,oh + γ Kt,ph + kbKt,b.

The method takes the load transfer and the secondary bending very well into ac-count. However, the choice of fasteners is not covered by this method. It is thereforeimportant to select the correct reference joint for the life prediction, i.e. a refer-ence joint with a fastener and fastener installation method close to the joint to beanalysed.

The above method can be extended to Glare under the following conditions:

– Peak stresses can be calculated for each individual aluminium layer.– The fatigue initiation behaviour of an individual layer is the same as for

monolithic aluminium.– The same stress concentration factors as for monolithic aluminium can be

used.– Reference joints for monolithic aluminium can be used.– The bending ratio depends on lay-up and layer to be analysed.

With these conditions only the first step of the prediction method should bemodified. In this step the remote stress must be replaced by the actual stress levelin the aluminium layers according to the procedure as given in the previous section.The stress ratio must be determined specifically for Glare.

4.5. FATIGUE CRACK PROPAGATION IN MECHANICALLY FASTENED JOINTS

As explained before, secondary bending will cause a stress distribution through thethickness with the highest stresses at the mating surfaces. In these surfaces crackinitiation will occur first. Crack propagation will first appear as a surface crack andlater, after initiation in the next layers, as a part through the thickness crack. Bothcrack scenarios are covered in the previous section.

Crack growth rates for the surface crack can be predicted with the appropriatemodel using the stress level in the cracked layer as remote stress. Crack growthbehaviour in the subsequent layers (part through case) are still being investigated.


5. Impact Properties

An aircraft structure should be able to tolerate a reasonable level of damage and de-fects that might be encountered during production or service. The structure shouldbe ‘damage tolerant’. The literature refers to ‘damage tolerance’ as a property ofthe material or the structure related with damage. However, it is important to realisethat there are four damage aspects [13]:

1. damage resistance (influence of environmental conditions, material choice,design aspects, etc.);

2. damage tolerance (effect of damage on the residual (mechanical) properties);3. inspectability (ease of detection of damage);4. reparability (ease of repair).

Two more aspects that need to be considered in the damage tolerance methodare:

5. damage sources (cause of damage and defects);6. damage types (form of damage).

The structural design and material choice must guarantee the structural integrityafter a realistic impact event. For Glare the impact damage resistance is related tothe aluminium and glass fibre/epoxy properties. Glare is stronger than aluminiumat higher impact velocities due to the high strain hardening effect and strain ratedependent behaviour of the fibres. The damage sources must be examined and theseverity of the impact damage quantified for Glare.

5.1. DAMAGE TYPES AND SOURCES

For increasing impact energy level of the projectile, the following damage can befound in Glare [14]:

1. Dent due to plastic deformation of the aluminium layers.2. Epoxy cracking and minor delamination in the dented area.3. First crack: combined fibre and aluminium cracking at non-impacted side.4. Through crack: also aluminium cracking at impacted side.

Beside the damage in the fibre-prepreg layer a distinction between first andthrough crack is made for Glare in comparison with aluminium.

An important part of the damage tolerance evaluation is the establishment ofimpact damage criteria. The type, probability of occurrence, and damage levelshould therefore be defined and damage sources known.

The impact regime for commercial transport aircraft can be subdivided in lowand high velocity impact. Low velocity impact (<11 m/s) occurs through dam-age from, for example, service trucks, cargo containers and dropped tools duringmaintenance operations. High velocity impact (>11 m/s) can occur when debris


from the runway hit the fuselage during take-off or landing, or when ice from thepropellers strikes the fuselage.

For commercial transport aircraft Davis and Sakata [15] state: To ensure thatthe flight crew is afforded the same level of protection from bird strike providedby the windshield, certain areas of the cockpit structure must be capable of with-standing impact of a 1.8 kg bird. Glare can sustain more damage per unit weightthan aluminium 2024-T3, especially at high velocity, and could well be appliedin the cockpit structure. In the crown section the relative ease of production ofdouble curved Glare panels and the good impact properties are a fine weight-savingcombination.

Hail damage has a low probability, because only few hailstones grow dimen-sions that are big enough to create damage. In case these hailstones do occur, theycan severely damage the aircraft.

At high aircraft speeds small hailstones can also create considerable damage.Hail should be considered separately, both on the ground and in flight. The impactlocations are different.

The application of Glare in the upper fuselage section is therefore prone to hailimpact damage. Figure 3 shows a 3.2 mm thick Glare 4B-4/3-0.5 specimen afteretching, impacted with a 5.7 cm ice ball at a speed of 140 m/s. The frequencyof encounter and impact density of this size of hailstones is extremely small. Thespeed corresponds with the approach of an aircraft. This corresponds to the factthat a hailstorm encounter is most likely to happen during approach, due to theflight altitude and due to the fact that it is more difficult to avoid. A gas gun is usedto accelerate the ice ball to this speed. A high speed camera captures the ice ballduring flight and at impact, and determines the speed and the integrity of the iceball. The Glare specimen showed no damage or delamination after impact [16].

For Glare moisture ingress is possible at the edges or after damage, when thealuminium has cracked. The edges should therefore be sealed to protect the panelfrom moisture and solvents (paint strippers, etc.). Special attention must also be

Figure 3. No damage or delamination can be seen after etching away the outer aluminiumlayers of a Glare 4B-4/3-0.5 specimen (dimensions 200 mm square), impacted with a 5.7 cmice ball at a speed of 140 m/s.


paid to the process control, which can be done by C-scanning the Glare panels.Overheating can degrade the material, but is not likely to occur in the fuselage andcan be detected by the discolouring of the paint.

5.2. DAMAGE QUANTIFICATION

To quantify the severity of impact damage there are four major damage criteria:

1. Maximum permanent deflection.2. Energy restitution coefficient.3. Damage width.4. Damage grade energy level.

The latter is the minimum cracking energy to create a first crack or a throughcrack. The minimum cracking energy is a measure of the ability to withstand im-pact damage. For first crack in fibre metal laminates, the impact energy needed fora visible crack in the specimen is taken. For a through crack the impact energyneeded for a visible crack on both sides of the specimen is determined. Figure 4shows a comparison of damage grade energy level per unit weight for respectivelylow velocity impact. The energy is divided by the specimen’s areal density, i.e.mass per square metre material, for a better comparison. The difference in energylevel for first and through crack is even more evident at high velocity impact test-ing. At both low and high velocity impact all Glare specimens, except the Glare2B-3/2-0.3 specimens, perform better than the aluminium specimens by compari-son with the damage grade energy levels.

Figure 4. Damage grade energy level per unit areal density for a crack in aluminium 2024-T3and for a first and a through crack in Glare, at low velocity impact (∼10 m/s, drop weight).


5.3. RESIDUAL PROPERTIES

Glare 4-5/4-0.5 (3.2 mm thickness) and aluminium 2024-T3 (3.8 mm thickness)show a similar compression strength after impact, i.e. 2% strength reduction due toa dent, and 10% strength reduction due to a through crack [17]. Global buckling ofthe specimens was initiated by local buckling in the dented area. No delaminationbuckling (critical for composite materials) was observed in the Glare specimens.The Glare specimen only shows crack growth in the outer aluminium layer at theimpacted side and no failure after 200.000 cycles (CA, 120 MPa, R = 0.05). Thecrack initiation time in aluminium 2024-T3 is much longer, but the crack growsthrough the thickness very fast resulting in final failure after approximately 130.000cycles. Due to a bending effect in the dent area, the crack starts to grow at thehigher loaded impacted side. This is favourable since the impacted side is usuallythe inspectable, outer side of the structure.

5.4. INSPECTION AND REPAIR

Accurate non-destructive testing (NDT) methods are needed to detect damage incomposites. The aircraft should be able to sustain detectable damage for a certainperiod before its detection. It is difficult to C-scan Glare specimens after impact,because not only the delamination but also the plastically deformed dent in thealuminium layers will reflect the ultrasounds since it is in general not perpendicularto the signal. However, Glare has a visible dent after impact and the damage size isalways smaller than the dent size. The effect of a dent on the residual strengthand fatigue properties is limited. Thus, only the surface scratches in the outeraluminium layer need to be considered for Glare. Inspection can be limited to thedented spots and scratches where the dent depth and respectively scratch length anddepth are the decisive factors for eventual repair. The repair of a Glare structure isessentially the same as for aluminium.

5.5. CONCLUSION

Glare is stronger than aluminium at higher impact velocities due to the strain ratedependent behaviour of the fibres. The dent depth after impact is comparable to alu-minium with the same thickness. The residual properties for Glare are comparableto (e.g., compressive strength), or better than (e.g., fatigue) monolithic aluminium.The inspection of a Glare structure is no more complicated than for aluminium,since Glare has the same indentation and because the damage size due to impact issmaller than the visible dent size.

6. Burn-through and Lightning Strike

For civil passenger aircraft a significant number of fatalities, in otherwise surviv-able accidents, are associated with the effects of fire. The number of fatalities due


Figure 5. Typical damage due to lightning strike on an unprotected carbon component [19].

to fire in these accidents is about 19 per accident [18]. Although the frequencyof accidents involving fires has been reduced, improvements in aircraft fire safetyhave become increasingly more important for future aircraft design.

Aluminium has relatively short burn-through times and the structural integrityof a fuselage rapidly diminishes during the melting of the skin and other structuralmembers. Therefore, new materials such as composites and Glare are subjected toburn-through testing.

A comparable subject for Glare is the strike of an aircraft by lightning. Allaircraft in service will suffer from lightning strike.

Metallic structures have a good electrical conductivity and will therefore onlysuffer from very local and minor damage due to a lightning strike. Compositematerials on the other hand, have a much lower conductivity and will suffer fromsevere structural damage if the material is not protected properly, see Figure 5.Glare is a hybrid material combining both behaviours.

6.1. GLARE EXPOSED TO FIRE

Fibre Metal Laminates have been studied in their burn-through behaviour.The studies were performed to find a way of improving the burn-through resis-

tance of aircraft fuselages.

6.2. REGULATIONS AND TESTS

The available Federal Airworthiness Regulations (FAR) do not impose burn-through requirements for fuselage skin structures. The requirements only treat thefire resistance of interior materials such as seats and cushions. Glare was developed


Figure 6. Burn through test set up.

as a material for the primary structure of a fuselage and consequently burn-throughtests on Glare were mostly performed using the test setup for cargo compartmentliners.

Regulations prescribe a burn-through test (Figure 6) for cargo compartment lin-ers consisting of a 930 ◦C flame with a heat flux of 91 kW/m2. In comparison withthe characteristics of a large post-crash fuel fire – a temperature of 1150 ◦C and aheat flux of 160 kW/m2 [20] – these values are rather low. However, the presenttest is very useful for comparison purposes between the behaviour of traditionalmaterials and Glare.

6.3. BURN-THROUGH SYSTEM

The aluminium layers in Glare show behaviour identical to that of a monolithicaluminium sheet when exposed to fire and will burn through within only sec-onds when exposed to fire. After burn-through of the outer aluminium layer, Fig-ure 7, the glass-fibre prepreg layer is exposed to the flames. The melting tem-perature of the glass fibres is higher than the flame temperature, therefore theglass fibres in combination with the epoxy matrix are able to keep the fire frompassing the layer and thus create a fire barrier. The remaining aluminium andfibre layers in the laminate are protected by the fire barrier and remain intact.In this way, the non-exposed side of the laminate is insulated from the fire. Thetemperature on the non-exposed side will remain relatively low. The excellentburn-through properties of Glare were verified in 1994. The Boeing Companyperformed fireproof tests according to FAA standards. The results of these testsshowed that Glare can be qualified as fireproof and can be used for firewall appli-cations.


Figure 7. Typical damage on Glare panel after 10 minutes of exposure to 1100 ◦C flames(courtesy Boeing) [21].

6.4. GLARE EXPOSED TO LIGHTNING

6.4.1. Lightning Strike Testing

Aircraft can be hit by a lightning strike during flight. Lightning strike tests havebeen performed on Glare by KEMA (a Dutch research institute for electrical com-ponents and electrical power supplies), using the standard lightning strike classes,described in the MIL handbook (MIL-STD-1757a).

Since Glare is not a monolithic material and also consists of non-conductivematerials, it is interesting to investigate the influence of a lightning strike on Glare.Improper protection of full composites has shown that a lightning strike can causesevere damage to the structure.

6.4.2. Glare Behaviour

The behaviour of Glare when hit by lightning can be explained by the layeredcomposition of the material. The fibre layers in the laminate are non-conducting,therefore the current must find its way through only the aluminium layers.

When a lightning strike hits the surface, a lot of heat is created in the material.Similar to Glare behaviour when exposed to fire, the outer aluminium layer meltsand vaporises locally. The underlying fibre layer carbonises and, depending on theseverity of the hit, the fibre layer will be partially damaged.

In all cases, the underlying layers of the laminate will not be damaged. Fig-ure 8 shows a detail of damage caused by a Direct Hit. As can be clearly seen,the damage to the laminate is concentrated on the outer aluminium layer and the


Figure 8. Details of damage due to lightning. The outer aluminium layer is melted and thefirst fibre layer is damaged after a class 1B Direct Hit.

first fibre layer only. The damage will not penetrate through the thickness of thesheet. Performing identical tests on aluminium will cause melting of the materialthroughout the thickness.

From the lightning strike tests, it can be concluded that the resistance of Glareagainst lightning strike is better than for aluminium. The combination of conduc-tive aluminium layers and isolating fibre layers results in lightning damage thatonly affects the outer two layers (one aluminium and one prepreg layer), leavingthe underlying layers undamaged.

The damage due to the lightning strike directly influences the fatigue life of thealuminium specimens. Fatigue cracks will start from the damage with fatigue crackpropagation extending over the thickness of the panel. Since in Glare only the outerlayers are damaged, the intact fibre layers bridge the lightning damage, resulting inan outstanding fatigue behaviour and residual strength.

7. Residual Strength

The residual strength of a material is defined as the remaining static strength of thematerial in the presence of any damage that can occur during the service life ofan aircraft. This is an imprecise definition since the term damage is not properlydefined and can vary between a smooth dent and a sharp crack. For design verifi-cation it is conservative to represent damages in Glare as sharp cracks. However,a distinction must be made between two different kinds of damage in the Glarelaminate:

1. Cracks that occur in one or more aluminium layers while the fibres in theprepreg layers that bridge the crack are still intact.

These cracks are mainly caused by fatigue loading. In this case one or moreof the aluminium layers are broken but the fibre layer remains intact. Fibresbridge the crack and reduce the stress intensity at the crack tip. Other causes of


“part-through cracks” in Glare may be scratches, corrosion damage or lightningstrikes.

2. Cracks that occur simultaneously in the fibre prepreg layers and aluminiumlayers.

These cracks occur due to severe Discrete Source Damage (DSD) whenan object penetrates the structure. This damage creates “through-the-thicknesscracks” with broken fibre and aluminium layers. Examples are engine non-containment, bird-strike and runway debris.

The residual strength of a fatigue crack in Glare is superior to the residualstrength of a fatigue crack in aluminium or a through-the-thickness crack in Glare.Extensive investigation was done after the Glare fracture mechanism [22] andresidual strength calculation [23, 24].

7.1. FRACTURE MECHANISM

The fracture mechanism is discussed for Glare materials with through-the-thicknesscracks and part-through cracks respectively.

7.1.1. Through-the-thickness Cracks

Glare material with a through-the-thickness crack that is loaded to fracture exhibitsa capability for slow stable crack growth prior to rapid failure which occurs whenthe specimen is loaded under force control. The fracture behaviour is similar tothe response of monolithic aluminium sheet material. Figure 9 shows a typicalcurve of the gross stress versus the physical crack extension for a Glare 2 laminate,obtained with a uniaxially loaded Centre Cracked Tension (CCT) specimen. Thecrack extension is the extension measured in the visible outer aluminium layers.

Figure 9. Typical crack extension versus applied load curve for Glare 2.


7.1.2. Fatigue Pre-cracking

During fatigue cycling, Figure 10(a), the fibres stay intact and bridge the crackin the aluminium layers. As a result of the fatigue pre-loading, debonding occursbetween the fibre prepreg and the aluminium layers. Thus fibres are still intact andcrack length’s in fibre and aluminium layers are different.

7.1.3. Static Loading

During static loading, the fatigue crack in the aluminium layers opens, Figure 10(b).At a certain load, the fatigue crack in the aluminium layer starts to propagate. Atapproximately 90% of the residual strength of the specimen, the first fibre bundlein the visible prepreg layer fails at the tip of the fatigue crack, illustrated in Fig-ure 10(c). During further load increase, random failure of fibre bundles is observedbetween the tip of the saw cut and the tip of the fatigue crack in the aluminiumlayer.

Figure 10. Crack growth sequence for a Glare 2 CCT-specimen.


Figure 11. Schematic overview of fracture process within Glare 2 and Glare 3, originatingfrom a fatigue starter crack.

Extensive research showed that crack extensions in prepreg layers and alu-minium layers become approximately equal during the remaining fracture process.However, this is not visible in Figure 10 due to the slanted crack faces of the outeraluminium layer.

7.1.4. Part-through Cracks

Compared to the extensive stable crack growth in statically loaded specimens withthrough-the-thickness cracks, specimens with part-through cracks do not show anystable crack growth at all when loaded up to failure.

Figure 11 shows differences between fracture mechanisms of Glare 2 andGlare 3. In Glare 2 material, static delamination occurs between the prepreg andthe aluminium layer at loads close to fracture. This delamination grows in fibredirection and results in larger crack opening of the aluminium layers. Final fracturewas initiated as fibres fail near the fatigue crack tip while fibres in the wake of thefatigue crack remain intact. Fracture in Glare 3 however was initiated in the fibresnear the hole at the centre of the fatigue crack, see Figure 11. This results in alarger crack opening and subsequent fibre failure from the centre towards the tipsof the fatigue crack. After all the bridging fibres failed, an almost sudden fractureoccurred.


Figure 12. KR curves for Glare 3-3/2-0.3 for different W and a0, based on compliance plasticzone correction.

7.2. R-CURVE AS A MATERIAL PARAMETER

An energy approach like the R-curve method was found to be a way to predictthe fracture behaviour and residual strength of Glare. The method allows for thephenomenon of stable crack extension and a limited amount of plasticity. Figure 12shows an R-curve result for Glare 3. It is demonstrated that the R-curve is fairlyspecimen size independent.

References

1. Kieboom, O. T., ‘Fatigue Crack Initiation and Early Crack Growth in Glare at DifferentTemperatures,’ Master thesis, Structures and Materials Laboratory, Faculty of AerospaceEngineering, Delft University of Technology, November 2000.

2. Alderliesten, R. C., ‘Development of an Empirical Fatigue Crack Growth Prediction Model forthe Fibre Metal Laminate Glare,’ Master thesis, Structures and Materials Laboratory, Facultyof Aerospace Engineering, Delft University of Technology, 14 December 1999.

3. Marissen, R., ‘Fatigue Crack Growth in ARALL, a Hybrid Aluminium-Aramid CompositeMaterial, Crack Growth Mechanisms and Quantitative Predictions of the Crack Growth Rate,’PhD thesis, June 1988, Report LR-574, Faculty of Aerospace Engineering, Delft University ofTechnology.

4. Woerden, H. J. M., ‘Variable Amplitude Fatigue of Glare® 3, Influence of Peak Loadsand Flight Spectrum Development,’ Preliminary thesis, Structures and Materials Laboratory,Faculty of Aerospace Engineering, Delft University of Technology, May 1998.

5. Gonesh, K. A. M., ‘Development of a Fatigue Crack Growth Model for Surface Cracks inGLARE,’ Master thesis, Structures and Materials Laboratory, Faculty of Aerospace Engineer-ing, Delft University of Technology, August 2000.

6. van der Hoeven, W., ‘Final Report Glare Durability Program, Results of Tests Carried out atNLR,’ NLR-CR-2000-237, National Aerospace Laboratory NLR, 2000.


7. Jarfall, L., ‘Shear Loaded Fastener Installations,’ Saab-Scania Report KHR-3360, April 1983.8. Müller, R. P. G., ‘An Experimental and Analytical Investigation on the Fatigue Behaviour of

Fuselage Riveted Lap Joints,’ Doctor thesis, Delft University of Technology, 1995.9. Schijve, J., ‘Some Elementary Calculations on Secondary Bending in Simple Lap Joints,’ NLR-

TR-72036.10. De Rijck, J. J. M. and Fawaz, S. A., ‘A Simplified Approach for Stress Analysis of Mechanically

Fastened Joints,’ 4th Joint DoD/FAA/NASA Conference on Aging Aircraft, St. Louis, May 2000.11. Pellenkoft, F., ‘Secondary Bending in Glare – Investigation on the Elastic Behaviour of and

Crack Propagation in Glare Loaded by Combined Tension and Bending,’ Masters thesis, TUDelft, 2000.

12. Homan, J. J. and Jongebreur, A. A., ‘Calculation Method for Predicting the Fatigue Life ofRiveted Joints,’ Proceedings of International Conference on Aircraft Fatigue, Stockholm, 1993,pp. 175–190.

13. Vlot, A., ‘Low-velocity Impact Loading – on Fibre Reinforced Aluminium Laminates (Arall)and Other Aircraft Sheet Materials,’ PhD thesis, Delft University of Technology, 1991.

14. Vlot, A., ‘Impact Loading on Fibre Metal Laminate,’ International Journal of Impact Engi-neering 18(3), 1996, 291–307.

15. Davis, G. W. and Sakata, I. F., ‘Design Considerations for Composite Fuselage Structure ofCommercial Transport Aircraft,’ NASA CR 159296, 1981.

16. Hagenbeek, M., ‘Impact of Ice, Impact Properties of Glare,’ Report B2v-98-15, Delft Univer-sity of Technology, Delft, February 1999.

17. Homan, J. J., ‘Compression after Impact,’ Report B2v-98-11, Delft University of Technology,Delft, December 1998.

18. van der Kevie, G., ‘The Modification of Glare to Develop a more Fire-resistant and Non-delaminating Fuselage Skin for the Airbus A3XX,’ Master thesis, Delft University ofTechnology, Faculty of Aerospace Engineering, 1997.

19. Reid, G. W., ‘Mechanical Damage to Aircraft Structures from Lightning Strikes,’ Journal ofAerospace Engineering, Part G 207, 1993, 1–14.

20. Dodd, D. C., Hall, C. T. M., Pollard, J. and Snell, M. A., ‘Burn through Resistance of Fuselages:Initial Findings,’ Civil Aviation Authority CAA Paper 94002, London, July 1994.

21. Roebroeks, G. H. J. J., ‘Glare; a Structural Material for Fire Resistant Fuselages,’ Aircraft FireSafety, AGARD Conference Proceedings 587, October 1996, pp. 26-1–26-13.

22. Vermeeren, C. A. J. R., ‘The Residual Strength of Fibre Metal Laminates,’ PhD thesis, Facultyof Aerospace Engineering, Delft University of Technology, The Netherlands, 1995.

23. de Vries, T. J. and Pacchione, M., ‘The Application of the R-curve Concept for the Predictionof the Residual Strength of Fiber Metal Laminates,’ Proceedings of the 1995 USAF StructuralIntegrity Program Conference, WL-TR-96-4093, United States, 1996.

24. de Vries, T. J., ‘Various Parameters Concerning the Residual Strength of Aircraft Materi-als,’ Master thesis, Faculty of Aerospace Engineering, Delft University of Technology, TheNetherlands, 1994.

Documents

Fatigue and Damage Tolerance of Glare