FM in Aircraft

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Fracture mechanics in aircraft failure analysis: Uses andlimitations

R.J.H. Wanhill a , , L. Molent b , S.A. Barter ba National Aerospace Laboratory NLR, Amsterdam, The Netherlandsb Defence Science and Technology Organisation (DSTO), Melbourne, Australia

a r t i c l e i n f o

Article history:Available online 16 November 2012

Keywords:Fracture mechanicsFatigue crack growthFailure analysis

a b s t r a c t

Fracture mechanics is generally accepted as essential to the design of metallic aircraftstructures, notably for predicting and validating the fatigue crack growth and residualstrength properties. Applications of Linear Elastic Fracture Mechanics (LEFM) began inthe early 1970s, and would be expected by now to have reached a mature status. Henceone might also expect that LEFM methods can often be used for analysing service fatiguefailures. However, the DSTOs and NLRs experience is that only non-LEFM analyses aredirectly useful for determining service failure fatigue crack growth (FCG) behaviour. LEFManalyses can assist in analysing FCG in full-scale, component and specimen tests, and theseanalyses can be used within limits to predict some aspects of the FCG behaviour of in-service structures and components. The DSTO has developed a hybrid method, the EffectiveBlock Approach (EBA) using both LEFM and non-LEFM analyses, from full-scale testing. TheEBA enables a wide range of FCG predictions for high performance aircraft.

2012 Elsevier Ltd. All rights reserved.

1. Introduction

The NLR and DSTO are the respective principal aerospace research organisations in the Netherlands and Australia. Bothhave more than 40 years experience of failure analysis, covering a variety of aircraft types, numerous structures and com-ponents, and both service and test failures. The present paper draws upon this experience to discuss the uses and limitationsof fracture mechanics, specically Linear Elastic Fracture Mechanics (LEFM), in analysing the failures. The discussion is con-ned to fatigue failures in metallic aircraft structures, since these types of failures predominate and are also the main con-tributors to serious accidents and incidents [15] . This predominance of fatigue reects the dynamic nature of aircraftservice loads and the relatively high design stress levels required to achieve lightweight structures.

2. Potential and actual uses of fracture mechanics in aircraft failure analysis

Fracture mechanics analyses may be considered for service failures, full-scale and component tests, and specimen/coupontests. For full-scale and component tests such analyses can be applied to both interim-detected cracking and cracks detectedduring nal teardown. An example of a full-scale test article is given in Fig. 1 . The complex conguration of this component isevident, and such complexity can and does pose problems for all types of analyses, as will be discussed in Section 5 of thispaper.

The types of analyses and their potential uses are specied here with the aid of Fig. 2 :

1350-6307/$ - see front matter 2012 Elsevier Ltd. All rights reserved.http://dx.doi.org/10.1016/j.engfailanal.2012.10.022

Corresponding author.E-mail address: [email protected] (R.J.H. Wanhill).

Engineering Failure Analysis 35 (2013) 3345

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Engineering Failure Analysis

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(1) Fatigue crack growth (FCG). Life to the maximum/critical crack size a c from an initial discontinuity/crack size a i. Life to the maximum/critical crack size a c from the minimum reliably detectable crack size andi . Feasibility of replacement/repair for remaining service or test life. Effect of aircraft usage severity (individual and eet differences) on life.

The latter two uses require explanation. Failures or detected cracking require assessment of whether replacement orrepair of the failed or damaged components will enable the remaining service or test lives to be reached with or without

Fig. 1. Full-scale fatigue testing of three Boeing F/A-18 centre barrel bulkheads attached to the engine ducts. The bulkheads have highly complexcongurations.

Fig. 2. Schematic of fatigue crack growth froman initial discontinuity/crack size, a i, to thecritical crack size, a c. The minimumreliably detectable crack size,andi , is also indicated.

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further replacements or repairs. FCG analyses based on service or test data canbe used in these assessments. Similaranalysesmay be used to assess the effects of aircraft usage, through its effects on service loads, on the in-service crack growth behav-iour and life.

(2) Fracture toughness. Estimation of the maximum permissible crack size (not necessarily the critical crack size a c) for service and testing.

For example, the Royal Australian Air Force (RAAF) has a cracked structure residual strength (RS) requirement of 1.2 Design Limit Load (DLL) without failure [6] . The maximum permissible crack size aRS at the RS P 1.2DLL point isnot necessarily the critical crack size under service loading.

3. FCG analysis methods

In the present context there are basically three types of FCG analyses to consider:

LEFM based analyses. non-LEFM based analyses. LEFM/non-LEFM based analyses: the Effective Block Approach (EBA).

3.1. LEFM based analyses

The starting point for these analyses is the tting of Paris-type equations [7,8] to constant amplitude (CA) fatigue crackgrowth data obtained from laboratory coupon tests in normal air. The basic equations are:

da

dN C D K m

C bD S ffiffiffiffiffiffipap m

1where da /dN is the FCG rate, D K D S

ffiffiffiffiffiffi p ap is the stress intensity factor range, C and m are empirical constants, and b is the

geometry correction factor. These equations describe a t of the central part (Region II) of a generic (simplied) FCG curve,see Fig. 3 , obtained from testing in a normal air environment.

Provided b is constant or approximately constant, Eq. (1) canbe directly integrated to give FCG lifetimes underCA loading,as follows:

For m 2; N f

2

m

2C

bm D S m

pm =2

1

a m 2

=2

0

1

a m 2

=2

f

" # 2

For m 2; N f

1C b2D S

2pLn

a f a 0 3

where a0 and a f are any chosen initial and nal crack sizes, respectively.Eqs. (1)(3) , or similar ones, are often given in textbooks and sometimes used in research paper modelling, e.g. Carpinteri

and Paggi [9] . However, the geometry correction factor b usually varies for real structures and components, and integration

Fig. 3. Schematic da /dN versus D K diagram for long fatigue crack growth in normal air, showing three regions of fatigue crack growth. D K th is the thresholdstress intensity factor range; K Ic and K c are the plane strain and non-plane strain fracture toughnesses, respectively.

R.J.H. Wanhill et al. / Engineering Failure Analysis 35 (2013) 3345 35


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must then be done numerically. Furthermore, numerical integration of CA-based equations is generally necessary becausethe load histories in service or during full-scale and component tests are usually variable amplitude (VA) and often complex.Another complication (a minor one) is that Region II of the FCG curve can have multiple linear segments [10,11] .

Owing to the foregoing complications, all LEFM based FCG prediction models estimate crack growth via a cycle-by-cycleprocess using integration algorithms [12,13] . Fig. 4 surveys the types of models and some of those available. Detailed discus-sions of the types of models are given in [13,14] . One of the most advanced ones, FASTRAN, is illustrated by the ow chart inFig. 5 .

3.2. Non-LEFM based analyses

These analyses require Quantitative Fractography (QF) of fatigue fracture surfaces. This requirement should be regardedas a merit rather than a hindrance, since QF crack growth data represent reality, whether they are from service failures orfull-scale, component and specimen tests. Thereare basically two kinds of QF measurements, depending on whether the loadhistory is nominally CA or VA:

(1) Nominally CA-FCG Measurements of fatigue striation spacings versus crack size, a, to obtain da /dN versus a plots.

Then derive a versus N plots numerically, or analytically by equation tting. Examples of successfully used equations,for Multiple Site fatigue Damage (MSD) in pressure cabin lap splices [15,16] and notched specimens simulating the

blade roots of a gas turbine disc [17] , are:

da

dN AeBa

! N f 1 AB

e Ba 0 e Ba f 4where A and B are empirical constants.

(2) VA-FCG Measurements of frequent or infrequent crack growth progression markers (natural or imposed) versus crack size,

a , to obtain partial a versus N plots. Check for an analytical t to complete the a versus N plots. Measurements of frequent marker spacings, and maybe striation spacings, versus crack size, a , to obtain da /dN ver-

sus a plots. Then derive a versus N plots numerically, or analytically by equation tting.

The choice between these measurement methods depends on the marking characteristics; see e.g. Barter et al. [18,19] .

Examples of analytical equations used for many service failures and test cracks in high performance aircraft [20,21]are:

a f a 0 e kN

! N f 1k

Lna f a 0 5

where k is an empirical constant that depends on the load spectrum and material. Eq. (5) represent exponential crackgrowth, which is also the case for Eq. (3) . However, Eq. (3) is for CA loading only and requires b to be constant, whereasEq. (5) do not.

Fig. 4. Overview of LEFM-based fatigue crack growth analysis models. After Iyyer et al. [12] .



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3.3. LEFM/non-LEFM based analyses: Effective Block Approach (EBA)

The EBA is a hybrid of LEFM and non-LEFM analyses that evolved from extensive experience with full-scale testing of highperformance aircraft structures and components, combined with specimen (coupon) tests and QF of fatigue fracture surfaces[22,23] . The EBA requires several conditions/assumptions and inputs:

The load histories consist of relatively short sequences of blocks of VA spectrum loading, such that an FCG curve is pro-duced from at least 30 load blocks [22] .

Each block of VA loading may be treated as if it were a CA cycle, and for each block load history the severity level is quan-tied by an overall reference stress, S ref, usually the peak stress level. Specimen (coupon) tests at different S ref levels to supplement the results from full-scale tests, which are usually done atone S ref level only.

QF measurements of frequent marker spacings (natural or imposed) versus crack size, a , to obtain a versus N B plots, whereN B is the number of VA blocks.

FCG rates can be modelled as power functions of the crack length. Selection of suitable FCG models and determination of their empirical constants from the full-scale and coupon test QFdata.

The models and their empirical constants are then used to predict FCG for different load history conditions pertaining todifferent locations in the aircraft structures and components, and also to (minor) differences in service aircraft usage. Thepredictions are described in more detail by Barter et al. [22] and Molent et al. [23] . Either LEFM or non-LEFM models are

used, depending on the quantied FCG behaviour. The basic equations are given here.

(1) LEFM (Paris-type) equations

da

dN B C B bS ref ffiffiffiffiffiffipap m B

6If b is constant or approximately constant, Eq. (6) can be directly integrated to give FCG lifetimes, as follows:

For m B 2; N B;f 2

m B 2C Bbm BS m B

ref pm B=21

a m B 2=20

1a m B 2=2

f " # 7

For m B 2; N B;f 1

C Bb2

S ref

2p

Lna f a 0 8

Fig. 5. Flow chart of FCG analysis using the strip yield/crack closure model FASTRAN [13] . The module at top left is a special one for calculating the crackopening stress for short cracks.



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As noted in Section 3.1 , b usually varies for real structures and components, and integration of Eq. (6) must then be donenumerically.

(1) Non-LEFM equations (Frost and Dugdale exponential-type) [2224]

For a versus N B plots ; a a 0 exp kBS m Bref

N B ! Lna kBS m Bref

N B Lna 0 9

For da =dN B versus S ref plots ; Ln da

dN B

1a LnkBS

mBref m BLnS ref 10

These (or similar) equations describe FCG for many high-strength structures and components tested under relativelyshort sequence VA loading [2023] . In particular, exponential FCG is often characteristic for lead cracks, i.e. the fastest-growing cracks in particular locations [20,21] .

4. Survey of DSTO and NLR FCG analyses

Table 1 surveys the types of FCG analyses for most of the well-investigated service and test examples from the DSTO andNLR archives and publications. Note that there were no examples where exclusively LEFM based analyses were used. This isdiscussed in Section 6 of the paper.

Tables 2 and 3 list the non-LEFM analysed structures and components. The EBA analysed component was an F/A-18 A/Bwing aft closure rib [23] . All the DSTO non-LEFM analyses used the VA-FCG approach and Eq. (5) , see Section 3.2 of this pa-per. Most of the NLR non-LEFM analyses were numerical, but two used the nominally CA-FCG approach, Eq. (4) and someadditional assumptions foranalysing the FCGbehaviour of pressure cabin longitudinal lap splices in several transport aircraft[16,34,35] .

5. FCG analysis method limitations and problems

5.1. Service and full-scale test FCG aspects

There are a number of actual and potential issues and problems to be considered and/or accounted for when conductingFCG analyses for service aircraft and full-scale tests:

(1) Loads and stresses . The fatigue load histories are obviously important. In most cases the load history will be VA, butsometimes nominally CA, e.g. Table 3 . In both cases the stress ratios, R = S

min/S

max, should be known for computational

modelling. For VA load histories it is necessary also to know or determine the types of spectra, whether the loadsequences are short or long, and the frequency of occurrence and distribution (random or non-random) of peak loadsand underloads. This latter aspect is particularly important for QF measurements of FCG via crack growth progressionmarkers.

(2) Structures and components . These can have complex geometries, e.g. Fig. 1 , resulting in stress concentrations andeccentricities leading to secondary bending. The occurrence of secondary bending during service or simulated-serviceloading is widespread in aircraft structures and is often unrecognized or poorly quantied. Notable examples are pres-sure cabin longitudinal lap splices, which undergo both longitudinal and circumferential secondary bending that sig-nicantly affect FCG from the stress concentrations provided by rivet holes [16,34,35] .Other problems are residual stresses in components and structural assembly (t-up) stresses; and load-shedding,owing to loss of stiffness, from cracked to uncracked components. These complications and their effects on FCG aredifcult or impossible to quantify prior to full-scale testing and QF measurements of FCG.An essential aspect that greatly affects the viability and validity of FCG analyses is the range of initial discontinuitysizes representing the effective pre-crack sizes (EPS) from which crack growth begins [20,21,39,40] . The initial discon-tinuity sizes and EPS values are much smaller than the assumed initial crack sizes commonly used in FCG analyses, e.g.the equivalent initial aw sizes (EIFS) specied by the United States Air Force (USAF) damage tolerance requirements[41] .

Table 1

FCG analysis types used for 21 examples of fatigue cracking and failure.

FCG analysis type Failure category DSTO NLR

LEFM Service/test Non-LEFM Service/test 4/6: Analytical 3/7: Numerical (8) and analytical (2)

EBA Test 1

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(3) Miscellaneous . Environmental inuences on service and full-scale test FCG behaviour are usually well-represented byfatigue in normal air, as evidenced by the similar fractographic appearances and the absence of corrosion damage suchas pitting and oxide formation on the fracture surfaces. However, occasional examples of corrosion-induced damageare observed for service failures [28,34,35] . The importance of an aggressive environment is that it accelerates FCGcompared to normal air, and this acceleration depends on the cycle frequency, see e.g. [42,43] .Fretting is a common initiator of fatigue [1] , notably at faying surfaces and in fastener and pin-loaded holes [44] . Fret-ting causes fatigue cracks to initiate very early in the life [37,45,46] and also increases early FCG rates [46,47] .

Lastly, VA loading can change the fatigue fracture topography compared to CA loading [23,48] . Pronounced changescan affect the FCG rates in unpredictable ways [48] .

The issues and problems mentioned in (1) and (2) are of prime importance to the predictive capability of FCG models. Themiscellaneous aspects in (3) are usually (much) less important, which is just as well, since they are not explicitly consideredfor modelling purposes.


Fig. 5 and Table 4 show that LEFM FCG models have many requirements relating to the loads and stresses, point (1) inSection 5.1 . However, in the rst instance these models take no account of point (2) except to use b -factors to cope withgeometry changes and stress concentrations. In particular, the models do not use actual short crack data. Instead they relyon back-extrapolation of long crack growth data, obtained under CA loading, to describe short crack behaviour.

Two examples will be given here to illustrate the problems in using LEFM based analyses.

5.2.1. Example 1: helicopter simulated fuselage panels under VA loading A Helicopter Damage Tolerance Round Robin Challenge was begun in May 2002 and the results published in 2003

[49,50] . Participants were requested to predict the VA-FCG curves and lives, using LEFM based analyses, for two helicoptersimulated fuselage panels containing small starter aws, see Fig. 6 . The panels were machined from aluminium alloy AA7010-T73651 plate. The participants were given the specimen geometry, Finite Element Method (FEM) stress intensity solu-tions including b -factors, VA load history, and CA long crack growth data. Twelve organisations, from three continents, pro-vided 27 blind (uncalibrated by actual FCG data) predictions using eight different software analysis packages [49] andoptions for load-interaction effects or non-interaction.

The predicted VA-FCG lives ranged from 240 ight hours to more than 12,500 ight hours, see Fig. 7 . The test lives were420 and 440 ight hours. Many of the predictions were within a factor of 2 of the test lives, which is usually consideredacceptable for blind predictions. However, some results were highly unconservative, even when the analyses used non-inter-action options. Furthermore, none of the analyses gave the actual a versus N curve shapes [49] : accurate predictions of FCGcurve shapes are needed for establishing inspection intervals to enable timely detection of in-service cracking.

Overall, the results were unsatisfactory, though the participants must be given credit for making open-access blind pre-dictions. Several reasons were given for the inadequacies [49,50] :

(1) Modelling curve ts . Software packages use curve ts to CA-FCG data. The curves often t better in one region of crackgrowth rates (see Fig. 3 ) and the t quality can vary for different R values.

(2) Material test data . The predictions relied mainly on CA data in the near-threshold region. This region is characterisedby considerable data scatter, including D K th , and steep da /dN versus D K gradients that may be expected to stronglyinuence the predictions.

(3) Stress intensity factors . The FEM solutions may have been in error, and other errors could have been made by the soft-ware package representations of the solutions.

Some additional aspects have to be considered:

Table 2

DSTO examples of non-LEFM based FCG analyses [2031] .

Aircraft/engine Component and material Service VA-FCG test

Aermacchi MB-326H Wing spar boom: AA 7075-T6 VAWing attachment boom: AISI 4340 steel

Boeing F/A-18 A/B Trailing edge ap lugs: AA 7050-T7451 VAVertical tail root: AA 7050-T7451 Centre barrel bulkhead: AA 7050-T7451 Rudder hinge tting: AA 7075-T73 Horizontal tail spindle: AF1410 steel

GD/LM F-111 Lower wing skin: 2024-T851 VA P&W 125B engine Bearing cage: AISI 4340 steel VA

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(4) Load history effects . The load spectrum was dominated by many low-amplitude high- R cycles interspersed with a fewunderloads rather than peak loads [49] . This means that FCG load history effects would have been small. Also, the cho-sen material would be expected to show only limited actual (and calculated) FCG load history effects, since it was ahigh-strength AA 7000 series alloy [51,52] .

(5) Panel conguration . The simulated fuselage panels were symmetrical on either side of the vertical axis, see Fig. 6 , andthey were also one-piece machined from plate material. Hence they would not have been liable to signicant second-

ary bending or load-shedding (load redistribution to the uncracked sides) except possibly during the later stages of FCG.(6) Starter aw . The fatigue crack starter aw had a 2 mmradius. This aw size is well beyond realistic initial discontinuity

sizes [20,21,39,40] and actually in the long crack regime, i.e. crack sizes more than about 0.5 mm for aluminium alloys[53,54] .

All three of these aspects should have made it easier to obtain (reasonably) accurate predictions of the VA-FCG curves andlives. This emphasizes the problems mentioned in points (1)(3).

5.2.2. Example 2: simple specimens (coupons) under VA loading The DSTO has done many VA-FCG tests on coupons of aluminium alloy AA 7050-T7451, whereby QF was used to deter-

mine the initial discontinuity sizes and FCG rates [1823,27,40] . The results were compared with LEFM based analyses of thecrack growth behaviour. Fig. 8 gives a representative example, with predictions from the widely used software package AF-

GROW [55] . In this example the predictions have been calibrated by selection of a near-arbitrary initial crack depth of 0.05 mm to obtain optimum ts to the nal FCG lives [23] .

The ts are actually poor, especially for lower stress levels. The predicted FCG curve shapes are incorrect, such that earlyFCG rates are severely under-predicted and the later FCG rates are over-predicted. Also, the curves canbe unconservative, seethe set of blue-symbol data and the corresponding predicted curve. There may be several reasons for the poor results [22] .However, the main problem appears to be that LEFM based analyses use long crack CA-FCG data for the predictions and can-not cope with actual short crack growth from realistic initial discontinuity sizes. For example, even small decreases in theselected initial crack depth used for the predictions in Fig. 8 would have drastically increased the life predictions [23] .

5.2.3. SummaryAs mentioned at the beginning of Section 5.2 , LEFM FCG analyses have many requirements relating to the loads and stres-

ses used in their predictions. This means there has to be a trial and error selection of models and model parameters to ob-tain the optimum analyses. Also, there are several actual and potential problems in using these models, notably the inability

to account for (unknown) secondary bending, load-shedding and residual stresses in actual structures and components, andthe application of long crack CA-FCG data to short crack growth predictions, including the use of unrealistic initial discon-tinuity sizes.

Table 3

NLR examples of non-LEFM based FCG analyses [4,15,16,3138] .

Aircraft/engine Component and material Service Test Analysis

Sikorsky S-61 Rotor blade: AA 6061-T6 CA NumericalNorthrop NF-5 Wing root lower skin: AA 7075-T651 VA NumericalFokker F100 Vertical stabilizer rib: AA 7175-T736 VA NumericalFokker F100 Fuselage lap splices: AA 2024-T3 CA AnalyticalBAC 1-11 Fuselage lap splices: AA 2024-T3 CA AnalyticalAirbus Mega Liner Barrel Window frame: AA 7175-T73 VA NumericalLiner barrel Window panel: GLARE 3 a VA Numerical(A380 MLB) Passenger door beam: GLARE 3 a CA NumericalNH industries NH-90 Tail hinge beam: AA 7075-T7351 VA NumericalP&W F100-PW-220 Air supply manifold support rod: In 718 VA Numerical

a GLAss REinforced AA 2024-T3 laminates [37] .

Table 4

LEFM based FCG analysis model requirements.

Basic (CA) FCG curves Crack opening stress, S op Stress ratios R; D K and D K eff Long cracks Curve transitions Short cracks, S tip Thresholds Constraint factors, a , 2D and 3D Short cracks

Load spectrum and loadhistory

Stress intensity factors, K

Cyclecounting, e.g. Rainow b : Geometry, stress concentrations Initial and nal crack sizes

FCG cumulative damage

calculations

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In view of the foregoing remarks, and the examples given in Sections 5.2.1 and 5.2.2 , it is evident that both uncalibrated

and calibrated LEFM based FCG life and curve predictions can be poor and unreliable, and that unconservative results arepossible. This conclusion agrees with previous results, especially with respect to uncalibrated LEFM models [12,56,57] .


Table 5 lists the non-LEFM FCG analysis requirements. The most demanding requirement is QF measurements of FCG,which are needed for both service and test failures. These measurements require considerable expertise and are labour-intensive. Also, for service failures and some full-scale and component tests there is no guarantee that QF is possible: theload history may be too complex and not provide natural markers. This is why there is much interest in providing markersfor tests [18,19] .

If predictions are required, then supplemental FCG specimen (coupon) tests must be done using load spectra and loadhistories representative of the known or expected service load histories, and employing natural or imposed markers[18,19] . The usefulness of these tests may be limited or compromised by an inability to account for geometry changes, sec-ondary bending, load-shedding and residual stresses in actual structures and components. The only way to compensate for

Fig. 6. Schematic diagram of a helicopter simulated fuselage panel (dimensions in mm). The heavy black arrow on Section XX points to the location of a2 mm radius starter aw. After Irving et al. [49] .

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these potential disadvantages is to thoroughly compare the full-scale and component QF results with those from couponsand possibly more complicated specimens [33] . Several examples are given in Molent et al. [20] .

It was mentioned in Section 4 that all the DSTO non-LEFM analyses used the VA-FCG approach and Eq. (5) , while most of the NLR non-LEFM analyses were numerical, with two using the nominally CA-FCG approach, Eq. (4) . The choice of methoddepended on the types of structures and components, the service and test load histories, and individual experience with suc-cessful analyses. For example, Eq. (5) , describing exponential-type FCG, have been successfully used by the DSTO for manyservice and test cracks in high performance aircraft [20,21] .

5.4. Effective Block Approach (EBA)

Table 6 summarises the EBA FCG analysis requirements, most of which are given in Section 3.3 . The load spectrum, loadhistory and specimen test requirements stem from (a) the development of the EBA for analysing the FCG behaviour of highperformance aircraft structures and components, (b) the assumption that blocks of VA loading may be treated as CA cycles,(c) the assumption that the severity level of different block load histories is quantied by a reference stress, S ref, and (d) theneed to supplement the results from full-scale tests, which are usually done at one S ref level only.

The most demanding requirement of the EBA is the large number of QF measurements. This is offset by the advantage of an extensive data base that can be used to predict FCG for different load history conditions pertaining to different locations,and also to minor differences in service aircraft usage, as mentioned in Section 3.3 .

Like the non-LEFM analyses discussed in Section 5.3 , EBA analyses may be partially compromised by an inability to ac-count for geometry changes, secondary bending, load-shedding and residual stresses. However, this limitation applies onlyto locations for which no QF data are available. Supplemental specimen tests, FEM analysis and strain gauge measurements(if not already available from full-scale and component testing) may be needed to cope with these locations.

6. Selection and use of FCG analysis methods


Uncalibrated (blind) LEFM based FCG analyses are unsuitable for service failures, since the predictions are too unreliable.This situation is unlikely to change, owing to the many uncertainties and problems illustrated and discussed in Section 5.2 .

Calibrated LEFM models tted to service and test data, e.g. Fig. 8 , may be used within limits for FCG life predictions toindicate the effects of usage severity for service aircraft. These predictions should be limited to minor stress level changesand differences in load histories:

(1) Stress levels . Interpolations are generally acceptable, but extrapolations should be restricted to stress level changes of 1020%. These limits are based on VA-FCG life data obtained for gust and manoeuvre spectrum loading [51,52] .

(2) Load histories . Life predictions for minor load spectrum variations could be considered, notably for short sequence loadhistories with frequent peak loads and underloads typical of high performance aircraft. However, there are no well-dened guidelines, only suggestions and opinions based on altering the load histories to provide crack growth pro-gression markers for QF [18,19,58] .

For high performance aircraft the model calibrations and predictions would probably benet from being included in EBAanalyses [22,23] , which, however, require extensive data bases.

Fig. 7. VAFCG life predictions and test lives for helicopter simulated fuselage panels. After Irving et al. [49] .

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Non-LEFM based FCG analyses are the only ones that the DSTO and NLR have found to be directly applicable to servicefailures, provided that QF measurements of FCG were possible. These analyses are also useful for full-scale, componentand specimen tests, again only if QF is possible.

Calibrated analyses derived from service and test data can be used for FCG life predictions, a versus N curves, and da /dnversus a behaviour. Most usefully, these analyses can provide estimates of the inspectable lives and inspection intervals forservice aircraft, and repair or replacement options for failed or damaged components, i.e. whether the components can berepaired or must be replaced in order to achieve the remaining service (or test) lives. Similar analyses may be used to assessthe effects of service usage severity via stress level interpolations of test data and limited extrapolation of test data for minorvariations in stress levels and load histories, see points (1) and (2) in Section 6.1 .

For high performance aircraft the non-LEFM analysis of test results are known to benet from being included in EBA anal-yses [22,23] .

6.3. Effective Block Approach (EBA)

The EBA is inherently restricted to the analysis of test data and therefore not directly applicable to service failures. How-ever, the EBAs exibility in using an extensive data base and LEFM and non-LEFM analyses enables the prediction of FCGbehaviour for many locations in high performance service aircraft that experience broadly similar load histories. In turn,these predictions enable estimates of the inspectable lives and inspection intervals; repair or replacement options for failedor damaged components; and the effects of usage severity, subject to the limitations already mentioned in Sections 6.1 and6.2 .

Fig. 8. Comparison of calibrated VAFCG predictions and QF-determined VAFCG for coupons of aluminium alloy AA 7050-T7451 tested with a tacticalaircraft spectrum. The characteristic stress is a measure of the VA load history severity, in this case the peak stress level. After Molent et al. [23] .

Table 5

Non-LEFM based FCG analysis requirements.

Load spectrum and load history Numerical or preferably analytical ts to QF data Natural or imposed markers

QF measurements of FCG Predictions also require: CA loading: striation spacings FCG specimen (coupon) tests VA loading: marker spacings Initial a and nal crack sizes

a Derived from back-extrapolation of QF data to obtain the effective pre-crack sizes (EPS).

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