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Damage Tolerance Engineering Property Evaluations of Aerospace Aluminium Alloys with Emphasis on Fatigue Crack Growth

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... Over the last 50-60 years there have been numerous investigations of fatigue crack initiation in high strength aluminium alloys, e.g. [18][19][20][21][22][23][24][25][26][27][28][29][30]. In some of the earlier work there was an understandable tendency to focus on crack initiation and development along slip bands [18][19][20][21][22], and it was not recognised that fatigue cracks could initiate at large constituent particles unless they were already cracked [18]. ...
... In some of the earlier work there was an understandable tendency to focus on crack initiation and development along slip bands [18][19][20][21][22], and it was not recognised that fatigue cracks could initiate at large constituent particles unless they were already cracked [18]. However, further studies showed that fatigue cracks can nucleate at both cracked and uncracked particles, and that these are the predominant sites of fatigue crack initiation in commercial alloys [19][20][21][25][26][27][28][29][30]. A recent study of fatigue crack initiation in coupons simulating a critical area in a military aircraft found that all the cracks began at cracked particles [31]. ...
... Although large particles are evidently associated with fatigue crack initiation, and there seems to be general agreement that when the particles are uncracked the cracks initially grow along the particle/matrix interfaces [19,21,23,25,28,30], opinions differ as to whether interfacial debonding is involved [21,25,27,30]). Also, it appears that some cracks may start from processing voids between closely-spaced particles [26]. ...
Article
A fatigue lifing framework using a lead crack concept, based on years of detailed inspection and analysis of fatigue cracks in many specimens and airframe components, has been developed by the DSTO for metallic primary airframe components. This framework is an important additional tool for determining aircraft component fatigue lives in the Royal Australian Air Force fleet. Like the original Damage Tolerance concept, developed by the United States Air Force, this framework assumes that fatigue cracking begins as soon as an aircraft enters service. However, there are major and fundamental differences. Instead of assuming initial crack sizes and deriving early crack growth behaviour from back-extrapolation of growth data for long cracks, the framework uses data for real cracks growing from small discontinuities inherent to the material and the production of the component. To this end, this paper examines the types of discontinuities that initiate fatigue cracks in typical metallic airframe structures. These discontinuities and the fatigue cracks that have grown from them are taken from coupon, component and full-scale tests, and also from service aircraft, including commercial transport aircraft and high performance military aircraft.
... These earlier tests revealed complicating issues requiring guidelines for further testing, the theme of this paper. The issues themselves have been discussed and reviewed in Wanhill [1,2]. ...
... The reason for this gap lies in the endeavour to use simple pre-cracked/starter notched sheet specimens for spectrum loading tests. The sizes of feasible pre-cracks and starter notches do not permit stabilised flight simulation fatigue crack growth data to be obtained in the non-inspectable slow crack growth regime [1,2,5]. We proposed further investigation of this problem in 1995 [5], but it has not been given priority. ...
... (2) Semi-random positioning of severe flights to provide characteristic markers on fatigue fracture surfaces is possible, as noted near the beginning of the previous main section of this paper. Figure 4 gives a striking example, whereby the fatigue crack front can be traced back to depths less than 20 µm and crack growth rates less than 10 -10 m/flight [1]. ...
Article
Flight simulation fatigue crack growth tests are necessary for verification of aircraft damage tolerance analyses and crack growth prediction methods, and also for comparing candidate materials for aircraft structural applications. However, such tests involve complicating issues that have emerged from many investigations on aluminium alloys since the late 1960s. These issues are reviewed to provide guidelines for further testing.
... istic K-values as attempts to correlate FCG data obtained under VA loading [18][19][20][60][61][62][63], but precedes further developments by the DST [21,[64][65][66], namely: ...
... These VA load histories may be referred to as 'stationary' or 'quasi-stationary'. Important examples are load histories pertaining to high-strength components in tactical aircraft [62][63][64], in part because training missions with similar load histories are repeated at regular intervals. Quasi-stationary approximations to the service load histories of tactical aircraft are very useful for full-scale, component and specimen tests. ...
... In the mid-1970s to mid-1980s long crack FCG data from flight simulation tests on aluminium alloy specimens were correlated by K max (tactical aircraft) [20,63] and K mf (transport aircraft) [63]: K max was derived from the maximum frequently occurring stress in the tactical aircraft test spectra, and K mf was derived from the so-called mean-stress-in-flight in the transport aircraft test spectrum. Better correla- ...
Chapter
Full-text available
The study of fatigue fractures has a long history [116]. Macroscopic fatigue progression markings on fracture surfaces were mentioned as such by 1926 [117].
... Over the last 50-60 years there have been numerous investigations of fatigue crack initiation in high strength aluminium alloys, e.g. [18][19][20][21][22][23][24][25][26][27][28][29][30]. In some of the earlier work there was an understandable tendency to focus on crack initiation and development along slip bands [18][19][20][21][22], and it was not recognised that fatigue cracks could initiate at large constituent particles unless they were already cracked [18]. ...
... In some of the earlier work there was an understandable tendency to focus on crack initiation and development along slip bands [18][19][20][21][22], and it was not recognised that fatigue cracks could initiate at large constituent particles unless they were already cracked [18]. However, further studies showed that fatigue cracks can nucleate at both cracked and uncracked particles, and that these are the predominant sites of fatigue crack initiation in commercial alloys [19][20][21][25][26][27][28][29][30]. A recent study of fatigue crack initiation in coupons simulating a critical area in a military aircraft found that all the cracks began at cracked particles [31]. ...
... The results mentioned above [31] showed that iron-containing particles were nearly always the sites for fatigue cracking. Although large particles are evidently associated with fatigue crack initiation, and there seems to be general agreement that when the particles are uncracked the cracks initially grow along the particle/matrix interfaces [19,21,23,25,28,30], opinions differ as to whether interfacial debonding is involved [21,25,27,30]). Also, it appears that some cracks may start from processing voids between closely-spaced particles [26]. ...
Chapter
A fatigue lifing framework using a lead crack concept has been developed by the DSTG for metallic primary airframe components. The framework is based on years of detailed inspection and analysis of fatigue cracks in many specimens and airframe components, and is an important additional tool for determining aircraft component fatigue lives in the Royal Australian Air Force (RAAF) fleet. Like the original damage tolerance (DT) concept developed by the United States Air Force (USAF), this framework assumes that fatigue cracking begins as soon as an aircraft enters service. However, there are major and fundamental differences. Instead of assuming initial crack sizes and deriving early crack growth behaviour from back-extrapolation of growth data for long cracks, the Defence Science and Technology (DST) Group framework uses data for real cracks growing from small discontinuities inherent to the material and the production of the component. To this end, this paper examines the types of discontinuities that initiate fatigue cracks in typical metallic airframe structures. These discontinuities and the fatigue cracks that have grown from them are taken from coupon, component and full-scale tests, and also from service aircraft, including commercial transport aircraft and high performance military aircraft.
... It should be stressed that this approach has only three constants and that unlike crack closure based crack growth models [24], which use significantly more constants one of which changes as the crack growth rate increases [24]Figure 1 where we see quite good agreement between the measured and the computed crack length histories. (As in [25]Figure 1 Pr e d icte d 2 4 5 M PaFigure 1 (case a) Measured and computed crack growth histories for 2048-T851 under FALSTAFF loading, adapted from [25]. The reasons for the different crack growth between the 196 MPa Test 1 and 2 is not discussed in [25]. ...
... It should be stressed that this approach has only three constants and that unlike crack closure based crack growth models [24], which use significantly more constants one of which changes as the crack growth rate increases [24]Figure 1 where we see quite good agreement between the measured and the computed crack length histories. (As in [25]Figure 1 Pr e d icte d 2 4 5 M PaFigure 1 (case a) Measured and computed crack growth histories for 2048-T851 under FALSTAFF loading, adapted from [25]. The reasons for the different crack growth between the 196 MPa Test 1 and 2 is not discussed in [25]. ...
... (As in [25]Figure 1 Pr e d icte d 2 4 5 M PaFigure 1 (case a) Measured and computed crack growth histories for 2048-T851 under FALSTAFF loading, adapted from [25]. The reasons for the different crack growth between the 196 MPa Test 1 and 2 is not discussed in [25]. Case (b) Let us next consider the report by Potter, Gallagher and Stalnaker [26] who presented crack growth data for 0.5 inch (12.7 mm) thick, and 1 inch (25.4 mm) wide 7075-T6511 Aluminum Alloy specimens with a working length of 6.5 inch (165 mm). ...
Article
Full-text available
This paper extends the average block variant of the generalised Frost-Dugdale crack growth law. We first reveal how the generalised Frost-Dugdale crack growth law to cover crack growth from initiation to final failure. We then show how this approach can be used to both represent and predict crack growth in both simple and complex structural geometries under variable amplitude loading
... Over the last 50-60 years there have been numerous investigations of fatigue crack initiation in high strength aluminium alloys, e.g. [18][19][20][21][22][23][24][25][26][27][28][29][30]. In some of the earlier work there was an understandable tendency to focus on crack initiation and development along slip bands [18][19][20][21][22], and it was not recognised that fatigue cracks could initiate at large constituent particles unless they were already cracked [18]. ...
... In some of the earlier work there was an understandable tendency to focus on crack initiation and development along slip bands [18][19][20][21][22], and it was not recognised that fatigue cracks could initiate at large constituent particles unless they were already cracked [18]. However, further studies showed that fatigue cracks can nucleate at both cracked and uncracked particles, and that these are the predominant sites of fatigue crack initiation in commercial alloys [19][20][21][25][26][27][28][29][30]. A recent study of fatigue crack initiation in coupons simulating a critical area in a military aircraft found that all the cracks began at cracked particles [31]. ...
... Although large particles are evidently associated with fatigue crack initiation, and there seems to be general agreement that when the particles are uncracked the cracks initially grow along the particle/matrix interfaces [19,21,23,25,28,30], opinions differ as to whether interfacial debonding is involved [21,25,27,30]). Also, it appears that some cracks may start from processing voids between closely-spaced particles [26]. ...
Chapter
Figure 3.1 illustrates schematically the reported fatigue crack initiation sites for the three main microstructural categories of near-α and α–β alloys (Wells and Sullivan 1969; Stubbington and Bowen 1974; Eylon and Pierce 1976; Eylon and Hall 1977; Postans and Jeal 1977; Ruppen et al. 1979; Bania et al. 1982; Bolingbroke and King 1986; Wojcik et al. 1988; Dowson et al. 1992; Evans and Bache 1994; Demulsant and Mendez 1995; Lütjering et al. 1996; Wagner 1997; Hines and Lütjering 1999).
... The numerical solution is termed further on as the ''exact solution''. Fig. 7. Fatigue crack growth data for 2024-T351 aluminum alloy obtained at stress ratios À2 6 R 6 0.7 [28][29][30]. Fatigue crack growth data for three materials were used to show the stress ratio effect on fatigue crack growth, i.e., aluminum alloy A1 2024-T351, steel alloy St-4340, and titanium alloy Ti-6Al-4V. The cyclic and fatigue properties for all materials are given in Table 1. ...
... The fatigue crack growth data for the Al 2024-T351 aluminum alloy was found in Refs. [28][29][30]. The fatigue crack growth data sets were obtained at various stress ratios, R appl , and are shown in Fig. 7 as a function of the applied stress intensity factor range, DK appl . ...
... Both the ''exact'' FCG curves and the approximate closed form solutions (Eqs. (23), (24), and (30)) are shown as diagrams where the fatigue crack growth rates are plotted as a function of the appropriate driving force Dj. The best results in correlating the FCG under various stress ratios were obtained while using the mixed driving force in the form of K p max;tot DK 0:5 tot . ...
Article
Full-text available
A unified two-parameter fatigue crack growth driving force model was developed to account for the residual stress and subsequently the stress ratio effect on fatigue crack growth. It was found that the driving force should be expressed as a combination of the maximum stress intensity factor, Kmax, and the stress intensity range, ΔK, corrected for the presence of the residual stress. As a result, the effects of residual stresses manifest themselves in changes of the applied maximum stress intensity factor and the applied stress intensity range. A two-parameter function of the maximum total stress intensity factor, Kmax,tot, and the total stress intensity range, ΔKtot, was proposed to model the fatigue crack growth rate data obtained at various R-ratios. Based on the analysis, the unified two-parameter driving force, Δκ=Kmax,totpΔKtot(1-p), was derived accounting for the mean stress or the stress ratio effect on fatigue crack propagation. It was shown that the two-parameter driving force, Δκ=Kmax,totpΔKtot0.5, was capable of correlating fatigue crack growth data obtained under a wide range of load ratios and fatigue crack growth rates spanning from the near threshold to the high growth rate regime.The model was successfully verified using a wide range of fatigue crack growth data obtained for Al 2024-T351 aluminium alloy, St-4340 steel alloy and Ti–6Al–4V titanium alloy with load ratios, R, ranging from −1 to 0.7.
... The paper "Gust spectrum fatigue crack propagation in candidate skin materials" (Wanhill 1979) addressed a very specific topic, and by itself would not have had much impact in the intervening years. However, most of the paper's content was subsequently included in an extensive report on the Damage Tolerance (DT) properties of aluminium alloys (Wanhill 1994a). This report enabled guidelines for flight simulation fatigue crack growth testing to be formulated (Wanhill 1994b(Wanhill , 2002. ...
... These alloys offered substantial weight savings, owing to decreased density and higher elastic modulus, compared with the AA2000 series DT alloys. Unfortunately, this second generation of Al-Li alloys had several shortcomings, including a tendency to have strongly anisotropic mechanical properties (low short-transverse ductility and fracture toughness), thermal instability (Lynch et al. 2003), problematical stress corrosion resistance (Schra and Wanhill 1999) and inferior flight simulation fatigue crack growth properties (Wanhill 1994a). ...
... Fatigue cracks less than about 0.5 mm in size often grow faster or more erratically than would be predicted from long crack data(Suresh and Ritchie 1984;Wanhill 1986;Wanhill 1994a) This is the so-called "small crack anomaly".NLR-TP-2008-831 ...
Article
IMPACT OF THE PAPER This paper 1 addressed a very specific topic and by itself would not have had much impact in the intervening years. However, most of the paper's content was subsequently included in an extensive report on the Damage Tolerance (DT) properties of aluminium alloys. 2 This report enabled guidelines for flight simulation fatigue crack growth testing to be formulated. 3,4
... Wanhill [62] evaluated the fatigue performance of 160 mm wide centre notched 2048-T851, 2024-T3 and 7075-T6 aluminium alloy sheets, where the thicknesses were 3.3 mm, 2.9 mm, and 4.0 mm, respectively, under FALSTAFF spectrum loading with a maximum stresses of 196 and 245 MPa. The measured and computed results, using Eq. ...
... In all of these cases, noting that the Boeing tests covered a range of aluminium alloys and thicknesses, we see that the materials characterisation program revealed that for the initial growth period, i.e. in the low to medium stress intensity region, crack growth is well represented by a (near) linear relationship between the normalised log of the crack length and the life, i.e. it conforms to Eqs. (17) and (19). Indeed, this relationship can also been in the centre cracked panel tests performed by Wanhill [62] on 204-T851, 7050-T736, and 7075-T6 under both FALSTAFF and MiniTwist flight spectra, as can be seen in Figs. 22 and 23. ...
... Fig.16. Measured and computed crack growth histories for 2024-T3 under FALSTAFF loading, adapted from[62]. ...
Article
This paper summarises recent developments in the formulation and application of the generalised Frost–Dugdale crack growth law. We first reveal the relationship between the generalised Frost–Dugdale crack growth law, dislocation based crack growth laws, the two parameter crack growth model, and fractal fatigue concepts. We then show that a range of aircraft materials characterisation test data are consistent with this law and how it can be used to predict crack growth in a range of full-scale aircraft fatigue tests, and coupon tests including crack growth in aircraft fuselage lap joints.
... he purpose of the work being reported here is to make improvements to current fatigue crack growth models so that better predictions of fatigue life can be made for critical airframe components. Predicting the life of aircraft structural metallic components loaded with VA spectra remains difficult since current methods do not achieve accurate results for crack growth in the small 1 to intermediate size range, which often governs the total fatigue life of a component [4][5][6]. The influence of small crack growth rates on the total fatigue life can be significant for a typical combat aircraft with a highly stressed structure, where the critical crack sizes are often less than 10mm [7] in depth, at least two thirds of the total fatigue life is consumed when the fatigue cracks are small. ...
... In this condition, striations are easily found and the crack growth behaviour at large inclusions (inherent in these materials) are involved in the crack extension and these features influence the crack path. For this reason, the paths taken by fatigue cracks for the two typical regimes; 1) ∆K and K max below ~5MPa√m, and 2) above this value to failure, of crack growth have notable differences 4 . While both references [16] discuss striation formation, this paper investigates crack paths in AA7050-T7451 for ∆Ks <5MPa√m (growth rates of <2x10 -7 m/cycle), with emphasis on ∆Ks less than 3 MPa√m (growth rates <2x10 -8 m/cycle). ...
... The coupons were all cut from a thick plate with a large effective grain size 6 . After 4 After striations become obvious in this material it is often found that the crack becomes rougher with obvious intersections with large second phase particles that tend to be to either side of the main growth plane. This results in a different sort of roughness developing. ...
Article
Full-text available
While it is well known that fatigue crack growth in metals that display confined slip, such as high strength aluminium alloys, develop crack paths that are responsive to the loading direction and the local microstructural orientation, it is less well known that such paths are also responsive to the loading history. In these materials, certain loading sequences can produce highly directional slip bands ahead of the crack tip and by adjusting the sequence of loads, distinct fracture surface features or progression marks, even at very small crack depths can result. Investigating the path a crack selects in fatigue testing when particular combinations of constant and variable amplitude load sequences are applied is providing insight into crack growth. Further, it is possible to design load sequences that allow very small amounts of crack growth to be measured, at very small crack sizes, well below the conventional crack growth threshold in the aluminium alloy discussed here. This paper reports on observations of the crack path phenomenon and a novel test loading method for measuring crack growth rates for very small crack depths in aluminium alloy 7050-T7451 (an important aircraft primary structural material). The aim of this work was to firstly generate short- crack constant amplitude growth data and secondly, through the careful manipulation of the applied loading, to achieve a greater understanding of the mechanisms of fatigue crack growth in the material being investigated. A particular focus of this work is the identification of the possible sources of crack growth retardation and closure in these small cracks. Interpreting these results suggests a possible mechanism for why small fatigue crack growth through this material under variable amplitude loading is faster than predicted from models based on constant amplitude data alone.
... where a f is the final crack size, B f is the associated block, or flight number, and B is the number of blocks at a. This relationship is supported by considering a number of examples in this paper. [31] who evaluated the fatigue performance of 110 mm wide, 5 mm thick centre notch Al 7050-T736 panels under a gust spectrum loading (MINI- TWIST) with an in-flight stress of 55 MPa; and a manoeuvre spectrum loading (FALSTAFF) with a maximum stress of 171.3 MPa. The measured and computed results, using Eq. (10) with the best fit estimates of ...
... (In these problems the left hand side of Eq. (10) is da/d(Flights)). (Case c) Wanhill [31] Mirage III usage and had a peak load equivalent to 7.5g. The experimental and computed crack length histories, using Eq. ...
... In all of these cases, noting that the Boeing tests covered a range of aluminium alloys and thicknesses, we see that the materials characterization program revealed that for the majority of the life, i.e. in the low to medium stress intensity region, crack growth is well represented by a (near) linear relationship between normalised log of the crack length and the fatigue life, i.e. it conforms to the generalised Frost–Dugdale law. This relationship can also be seen in the centre cracked panel tests performed by Schijve et al. [30], Wanhill [31], and Porter [34] as can be seen in Figs. 22–25, which are non-dimensional representations of the data presented in Figs. ...
Article
This paper builds on a development in the science of fatigue crack growth to present an equivalent spectrum block method for predicting fatigue crack growth under variable amplitude loading. This approach is based on the generalised Frost–Dugdale model and forms an analytical basis for the observation of a near exponential relationship between crack length and fatigue life under variable amplitude loading. A method for extending the Frost and Dugdale model from Region I through to Region III is also presented.
... This means that FCG load history effects would have been small. Also, the chosen material would be expected to show only limited actual (and calculated) FCG load history effects, since it was a high-strength AA 7000 series alloy [51,52]. (5) Panel configuration. ...
... Interpolations are generally acceptable, but extrapolations should be restricted to stress level changes of ±10-20%. These limits are based on VA-FCG life data obtained for gust and manoeuvre spectrum loading [51,52]. (2) Load histories. ...
... In particular, large loads are known to cause variations in constant amplitude crack growth rates, and crack closure effects are often included in the adjustments to the predictive algorithms to account for this, nevertheless, predictions for cracks growing from a small size remain poor [1]. For this reason further work on small and inter-mediate sized cracks with spectrum effects is warranted, so that the appropriate adjustment can be made to predictive models. ...
... Fig. 13 shows a closeup of the striations produced from five successive underloads bounded by high R cycles which is the last part of Sequence 3. The high R cycles do not produce visible striations at this scale. 1 Initially, the first four underload striations appear to be the same size and the fifth (closest to the crack tip) appears larger. However, this interpretation is incorrect. ...
Article
In order to predict variable amplitude crack growth it is necessary to understand the different mechanisms present in variable amplitude and constant amplitude fatigue crack growth. AFM and SEM observations have been made of the fatigue crack fracture surface in AA7050-T7451 alloy, produced by some simple load sequences consisting of periodic underloads (R=−1) in between groups of high stress ratio (R=0.5) loading cycles. These observations have revealed complex fracture surface features that include ridges, depressions and fissures. These features are a result of the slip band formation associated with underloads, which reduces the tendency for a new slip band to occur at the crack tip in the same direction as nearby slip bands. These slip bands change the path of the crack and result in the production of a ridge on the fracture surface. This effect suggests a model of striation formation that also explains the formation of ridges and other associated features, based on the influence of two or more active slip systems combined with the planar slip behaviour of this material.
... Whilst this behaviour may occupy a considerable length of the total growth, it only accounts for a small fraction of the total life of the crack and thus this period of growth may be ignored under the lead crack method. This is demonstrated by the data in Fig. 11, where Wanhill [28] ...
... Fig. 11. Measured crack growth histories for AA2048-T851, AA2024-T3 and AA7075-T6 under FALSTAFF loading, adapted from[28]. ...
Article
Full-text available
Over many years of quantitative fractographic examination of fatigue cracking from in-service and full-scale fatigue tests of metallic airframe components, it has been consistently observed that the largest cracks formed have grown in an approximately exponential manner. These crack growth observations range from the initiation of cracks and their early growth from a few micrometers through to many millimetres in length. It appears that these lead cracks commence growing shortly after the airframe is introduced to the loading environment. Furthermore these cracks usually initiate from production-induced or, less frequently, inherent material discontinuities. Based on these two observations, an aircraft lifing methodology that is based on the results of fatigue testing programs utilising the lead crack concept has been developed and implemented as an additional tool in the determination of aircraft component fatigue lives in several Royal Australian Air Force (RAAF) fleet types. In this paper the lead crack concept is developed and its strengths and weaknesses are discussed. Examples of crack growth behaviour that are considered typical and representative of lead cracks are presented.
... The work being reported here was motivated by the desire to improve current fatigue crack growth models so that better predictions of the fatigue lives of critical airframe components can be made. Predicting the fatigue life of aircraft structural metallic components loaded with VA spectra remains difficult, since current methods do not achieve accurate results for crack growth in the small 2 to intermediate size range; a region that often governs the total fatigue life of a component [4][5][6]. The influence of small crack growth rates on the total fatigue life can be significant for a typical critical component in a combat aircraft since these structures are generally highly stressed. ...
... Progression bands are the result of a series of load cycles; the bands made by these sequences can also be observed at about the same growth increment [3], while evidence of the individual cycles within the band remains unobservable. 4 Thick plate AA7050-T7451, the material of interest here, is commonly used to manufacture major aircraft structures, has a large pre-recrystallization grain size from restricted rolling during manufacture. Although the recrystallization grain size may be small (about 10 lm), the lack of deformation from the rolling leaves these sub-grains with orientations similar to the pre-recrystallization grains from which they were formed. ...
Article
It is accepted that fatigue crack paths in metals that display confined slip, such as high strength aluminium alloys, are responsive to loading direction and local microstructural orientation. It is less well recognised that crack paths in these alloys are also responsive to the loading history since certain loading sequences can produce highly directional slip bands ahead of the crack tip. By adjusting the sequence of loads, distinct fracture surface features or progression marks can result, even at very small crack depths. An investigation into the path a fatigue crack selects as it progresses through a material when it is cyclically loaded with particular combinations of constant and variable amplitude sequences has provided insight into the way these cracks grow. This makes it possible to design load sequences that allow very small advances of crack growth to be measured post-test by Quantitative Fractography, at growth rates well below the conventionally recognised threshold of the aluminium alloy examined here.
... The influence of anodizing on corrosion fatigue performance does not appear to have been investigated systematically. However, several authors have cautioned that anodizing, if not carried out correctly, may cause surface pitting which can act as a fatigue crack initiator (Forsyth (1980), Wanhill (1994b)). Wanhill et al. (1989). ...
... However, under nominally the same conditions, the fatigue performance of lap joints may be either degraded or improved by anodizing. Wanhill (1994b). ...
Article
Full-text available
Increasing economic pressure has encouraged extension of the service lives of many civil and military aircraft fleets beyond their original design goals. Consequently, since the incidence of corrosion tends to increase with aircraft age, its importance as a life limiting form of degradation has increased in these fleets. While the proportion of aircraft accidents and incidents attributed directly to the presence of corrosion is relatively small, the potential of corrosion damage and corrosive environments to cause or accelerate structural failure in aircraft will need to be incorporated into RAAF fleet structural integrity management approaches. In several recent cases the presence of corrosion has raised uncertainties over the continued airworthiness of some RAAF aircraft and, while these cases were resolved, structural integrity concerns associated with the detection of corrosion can lead to reduced aircraft availability and substantial increases in maintenance and support costs. Recent examples of this have included the discovery of stress corrosion cracking in P-3C Orion wing rear spar caps and Macchi MB326H tailplane spar caps. The absence of suitable methods of analysing the effect of this type of corrosion damage on the static strength and fatigue performance of these components made it impossible to guarantee the long-term airworthiness of these aircraft. It was therefore ultimately necessary to replace the components at some considerable cost. The main aims of this report have been: .To review the literature concerned with the effects of prior corrosion and corrosive environments on aircraft static strength and fatigue performance. .To identify and review current research programs in this area. .To identify potential research areas which could most effectively assist the RAAF in the management of structural integrity issues associated with corrosion in ageing aircraft fleets. It has been concluded from the review that the engineering assessment of corrosion in aircraft structures is a complex problem and a multidisciplinary approach is required to address critical issues in a number of key areas. These include: .Incorporating corrosion damage and environmental effects into conventional structural life management approaches. .Characterising aircraft operating environments. .Determining realistic upper bounds of corrosion progression. .Determining the effectiveness of corrosion preventive compounds in retarding the growth of corrosion while at the same time preserving airframe structural performance. .Modelling the initiation of fatigue cracking from corrosion damage. Future research aimed at resolving these and other issues will benefit from a range of activities at three different levels: 1. Data surveys, information exchange, and the evaluation of existing corrosion management programs. 2. Applied research, aimed at addressing specific life management issues. 3. Strategic research, aimed at addressing developing capability in key areas. Potential areas of research are discussed on the basis of their usefulness for the support of RAAF aircraft, the risk associated with obtaining useful outcomes, and the possibility of obtaining benefit from overseas collaboration. A potential research program is presented in the final chapter of the report. This discusses discusses the influence of corrosion on aircraft structural integrity. Brief introductions to corrosion in aircraft and aircraft structural integrity are provided and the literature concerned with the effect of prior corrosion and corrosive environments on static strength and fatigue performance is reviewed. RAAF and overseas experience with structural integrity issues associated with corrosion in aircraft is described, with emphasis placed on corrosion in airframes and structural structural elements. The discusses discusses the difficulties associated with incorporating the effects of corrosion into conventional life management approaches and the contribution of corrosion control programs to continuing airworthiness. Where possible, current research programs throughout the world are reviewed. Finally, potential areas of research are identified, primarily on the basis of their potential usefulness for the future support of RAAF aircraft and opportunities for collaborative research. RAAF DTA-LSA
... The inclusions in figure 5a are FeNiAl 9 , and are characteristic of this alloy, though they need not always be cracked [11]. [11,25] The sizes of the inclusions in figure 5 are typical. FeNiAl 9 is specific to iron-and nickel-containing alloys, but there are several kinds of large inclusions that commonly occur in commercial alloys, including Al 7 Cu 2 (Fe,Cr), (Fe,Mn)Al 6 , Mg 2 Si and Al 2 CuMg [23,24,26,27]. ...
... (2) The assumed defects are all surface-connected, for example table 4. When designing for safety the assumed defects are all larger than 0.5 mm. When designing for durability the assumed defects are generally larger than 0.1 mm [73,74] and unlikely to be smaller than 0.03 mm [25]. ...
Technical Report
Full-text available
Fatigue crack initiation ( nucleation ) in aerospace aluminium alloys, components and structures is discussed with particular reference to the concept of self-healing of fatigue cracks in aluminium alloys. If self-healing occurs it is limited to internal slip band cracks ~ 1 micrometer. Self-healing should be considered inapplicable in practice to aerospace aluminium alloys, components and structures.
... an overview on the small coupon fatigue and crack growth properties of friction stir welded butt joints is given. the investigations were carried out on 2024 (Wanhill 1990), 2024A (Warner et al. 1999), 6013 (Cieslak 1987) and 6056 (Blanc and Mankowski 1998) aluminum alloys in the thickness range of 1.6 mm to 6 mm. AA 6013 was welded also in T-joint form (Erbslöh et al. 2003) and fatigue tested. ...
Article
The present chapter covers the microstructural development of friction stir welds (FSW), their corrosion behaviour and mechanical properties. The section on microstructural aspects of friction stir welds presents a general description of the different weld zones and their main characteristics. The sub-structural evolution of an AlMgSc alloy is described in detail due to its industrial significance. For the same reason, the metallurgy of dissimilar welds is described based on two case studies: Al-high strength steel and Al-Mg joints. The section on corrosion starts with a general description of corrosion phenomena in Al alloys. This is followed by a specific description of the corrosion behaviour of friction stir welds and their corrosion fatigue properties. This section is concluded with an analysis of the influence of the base material temper on the corrosion resistance of the friction stir welded joints. The chapter is concluded with a comprehensive analysis of the mechanical behaviour of friction stir welded joints focused on fatigue and fatigue crack propagation. In both cases relevant factors controlling joint performance are discussed including thickness and notch effects, loading condition as well as environmental and residual stress effects. In general terms, this chapter addresses high performance applications of friction stir welding such as those found in the transportation industry and other load carrying engineering structures. However the information provided is certainly also applicable to less stringent industrial applications of the process.
... The resultant non-dimensional crack growth history for specimens KDE10, KD1R23, KD1P24, and KS1G29 are shown inFigure 10. Wanhill [85] also presented data on 7050-T736 tested under FALSTAFF (generic fighter wing root spectrum) and MiniTWIST (generic transport wing root spectrum) flight spectra and the resultant non-dimensional plots are also shown inFigure 10, where they are referred to as FALSTAFF and MiniTWIST respectively.Figure 10 also shows the crack growth data presented by Hsu, Chan and Yu [86], which is labelled 7050 OL, for cracking in a 76.2 mm wide, 356 mm long, and 6 mm thick centre cracked 7050-T76 panel tested with overloads. Here the specimen was tested under repeated block loading where each block consisted of 2500 cycles of constant amplitude loading, with a peak stress of 62.7 MPa and R = 0.05, followed by a single 150% overload (OL). ...
Article
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The objective of this report was to critically evaluate the current state of knowledge related to the modelling of crack growth in F/A-18 Hornet structural materials, with a particular focus on the status of the Generalised Frost-Dugdale model and its equivalent block variant. This was achieved through: 1. A brief review of the published literature with respect to the use of the Generalised Frost-Dugdale model to model crack growth in F/A-18 structural materials. 2. A brief review of the published literature with respect to the equivalent block variant of the Generalised Frost-Dugdale model. 3. Comparing predicted crack length histories using the Generalised Frost- Dugdale model with those obtained experimentally at DSTO and those reported in the open literature, particularly for short cracks. 4. Evaluating the need for further research. 5. Proposing future work to meet the observed shortcomings, if any. This report reveals that the so-called short crack effect associated with 7050-T7451 aluminium alloy arises as a consequence of attempting to relate crack length per cycle (da/dN) to the range of the stress intensity factor (ΔK) and that that cracking in both 7050 series aluminium alloys and Mil Annealed Ti-6Al-4V conforms to the Generalised Frost-Dugdale model. The report recommends how to best determine the constants used in the Generalised Frost- Dugdale model. Furthermore, when determining these constants it recommends using crack growth data obtained from simple surface flaw specimens subjected to the loading spectrum of interest, where the initial flaws are allowed to develop naturally. The results of this work will assist in the development of robust fatigue assessment tools in support of maintaining airworthiness in the RAAF fleets. The F/A-18 Hornet is one of the ADF’s premier defence assets. It utilises a highly optimised metallic structure to assist in achieving its performance objectives. Thus the structure (particularly the fracture critical wing attachment bulkheads or the “centre barrel”) is uniformly highly stressed and thus susceptible to fatigue cracking. In order to help assess the fatigue critical regions of the aircraft it is essential that the tools used to predict/assess fatigue crack growth are consistent with the known fatigue behaviour of the F/A-18 materials. This report provides a critical review of the current state of knowledge with respect to the use of the Generalised Frost-Dugdale model for assessing fatigue crack growth in F/A-18 structural metallic materials and evaluates the need for further research in this area. DGTA
... The fatigue performance of 110mm wide, 5mm thick centre notch Al7050T736 panels under either a gust loading spectrum (MINITWIST) with an in-flight stress of 55 MPa; or manoeuvre spectrum loading (FALSTAFF) with a maximum stress of 171.3 MPa are considered. These data are drawn from Reference [18]. The measured and computed results, using Equation The next set of data considered was chosen specifically because it had been generated for a modern aircraft design and included a full range of spectrum types likely to be experienced in the aircraft's lifetime [19]. ...
Article
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A review of experimental data show that for many lead fatigue cracks in service components loaded with service spectra, exponential growth (i.e. log crack depth versus cycles or hours) applies for the majority of the life. This behaviour is shown to extend from the micro to macro range of crack sizes in a variety of metals. As a consequence of this, it will also be shown that the crack growth rate is directly proportional to the crack depth. By combining these observations with traditional fracture mechanics approaches to crack growth modelling, a model that is a function of the stress intensity factor (K) with a fixed crack depth influence (non-similitude for the K parameter alone) is proposed. It will then be shown that this model allows for Region I to be smoothly integrated with Region II of the constant amplitude da/dN data. Further, it will be shown that for variable amplitude crack growth data, crack growth ranging from microns to many millimetres can be modelled using this single model. This modelling approach is of particular importance in structural integrity analysis where fatigue cracking cannot always be avoided and the majority of the fatigue life of highly stressed, nominally gross defect free structure is spent growing physically small cracks from initiating discontinuities (i.e. loads in Region I for constant amplitude loading growth rates) up to the point of loss in acceptable strength.
... Through the forensic examination of fracture surfaces of fatigue cracks, valuable information regarding both fatigue crack origin and progression [5] can be derived. Growth rates of fatigue cracks in many aerospace metallic alloys are often observed to grow in a stable but exponential manner [6], [7], [8]. A clear consequence of the exponential behaviour is that most of the fatigue life of a component will be spent while the crack is relatively small [9]. ...
Conference Paper
Previous literature has shown that during fatigue crack growth, application of variable amplitude load cycles will produce distinct physical features on the fracture surface of metals related to the loading variation. Some of these particular topographic features are predictable and coincide with measurable changes in the path of the fatigue crack. These crack path changes are often utilised to assist quantitative fractography and the measurement of crack growth rates. However, a widely accepted accurate physical model to explain these features and the associated crack growth has yet to be found. In order to develop a better understanding of the physical phenomena occurring during crack path changes, accurate three dimensional crack path measurement data, correlated to the loading is desirable. A comparison of measurements taken from the fracture surface of a typical aircraft structural aluminium alloy, via three techniques is presented. These techniques were: scanning electron microscopy stereo pairing, optical interferometry, and atomic force microscopy. A discussion of each technique is presented along with recommendations to achieve useful results. Finally, a discussion on the accuracy of each technique is presented, highlighting the applicability of each method for the material and fracture surfaces chosen in this study.
... However, since the scatter band is broad, particularly for AA2024-T3, here the size effect cannot be excluded completely. We did not find any conclusive evidence on possible effect(s) of specimen width on fatigue crack propagation of M(T) specimens or recommendations in this context in the literature [18] which provides the largest single source laboratory data on aerospace Al-alloys. Small width speci-mens have been used in the literature [19,20], but how these fatigue crack propagation data correlate with those of large width specimens has not been commented. ...
Article
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Accuracy of indirect fatigue crack length measurement by potential drop method or by compliance technique may be affected at low load ratio due to fracture surface contact, crack closure or mixed mode fracture. As an alternative, the maximum value of crack opening displacement, CODmax, from a clip gauge was utilized. Middle crack tension M(T) specimens were used to obtain conservative data at a low load ratio (R = 0.1). Thin sheet specimens (B = 3.2 mm) with different widths (100 mm s W s 400 mm) of AA6056-T4 were investigated in the mid-regime (Paris regime), which is of interest for damage tolerance analysis. The use of CODmax is found to provide crack lengths equivalent to those measured optically. Hence, the method is very suitable for indirect crack length measurement. Furthermore, small width specimens provided data equivalent to large width specimens. Insofar, the size effect is found to be absent, and fatigue crack propagation data can be acquired on small width specimens when material availability is limited.
... The fatigue crack growth data for the Al 2024-T351 aluminum alloy was found in [127,128,129]. The fatigue crack growth data sets were obtained at various stress ratios, R appl , and are shown in Figure 4-13 as a function of the applied stress intensity factor range, ΔK appl . ...
... The TTCI distribution is physically observable and can be obtain by experiments and tests results. Fatigue crack initiation and early crack growth in a SENT specimen tested with the Fokker 100 Reduced Basic (RB) gust spectrum [7] is shown in Fig. 4. The spacings of the bands on the fracture surface above the fatigue origin correspond to blocks of 5000 flights. fatigue origin In region II stable fatigue crack growth conditions prevail and the fatigue crack growth rate (FCGR) is given by the well-known Paris-Erdogan relation [8][9][10]. ...
Article
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Failure analysis and prevention are important to all of the engineering disciplines, especially for the aerospace industry. Aircraft accidents are remembered by the public because of the unusually high loss of life and broad extent of damage. In this paper, the artificial neural network (ANN) technique for the data processing of on-line fatigue crack growth monitoring is proposed after analyzing the general technique for fatigue crack growth data. A model for predicting the fatigue crack growth by ANN is presented, which does not need all kinds of materials and environment parameters, and only needs to measure the relation between a (length of crack) and N (cyclic times of loading) in-service. The feasibility of this model was verified by some examples. It makes up the inadequacy of data processing for current technique and on-line monitoring. Hence it has definite realistic meaning for engineering application.
... The use of zirconium decreases quench sensitivity, which means that uniform properties are more easily achieved in thick sections. A consequence of this is a greater sensitivity to corrosion pitting when compared to 7075-T73, Wanhill, 1995, which necessitates a good surface corrosion protection scheme if corrosion pitting is not to adversely affect fatigue life. ...
... To answer this question we will use the crack growth data given by Wanhill (1995) where a 0 is the initial flaw size, which in this case is ~ 5 mm. This prediction is shown below together with the experimental data. ...
Research
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Updated version of the MAE4408 Lecture Notes covering topic in Damage Tolerance, Fatigue crack growth, composite repairs, cubic rule, fracture mechanics, etc
... Wanhill showed that a single overload had to be 140% of the CA load to mark the fracture surface. 20 Thus, it is not surprising that the 130% overload spectrum used for two of the specimens did not adequately mark the fracture surface since the change in the effective stress intensity factor during the overload cycles was insufficient. The second spectrum created marker bands through final fracture and to crack lengths as small as 9 mm. ...
... Such variations in crack growth can be a retardation or acceleration of the growth rate in comparison to the rates predicted via long crack CA growth rate data [9][10][11][12]. Attempts to predict these transient 1 growth rates have been made [13][14][15] for single or multiple overloads (or underloads) within a CA load sequence. An understanding of the underlying physics occurring at the crack tip could inform the corrections that should be applied, a priori to any one cycle in a VA loading spectrum. ...
Article
Whilst crack growth retardation following the application of an overload during variable amplitude loading is a well-known spectrum effect, the order of smaller amplitude cycles can also affect crack growth. One method of accounting for such spectrum effects in crack growth prediction algorithms is by correcting the estimated effectiveness of each cycle or group of cycles. Such predictions are usually verified by comparison to demonstrated total fatigue lives. These corrections have been shown to be suitable for cracks that are at least several millimetres long. However, for small cracks: here defined as less than a few mm in depth, such predictions are often far less accurate. In the case of aluminium alloy aircraft structures, where cracks typically grow approximately exponentially from small discontinuities, these predictive algorithms should also be evaluated for their ability to predict small crack growth rates, rather than total life alone.
... Previous studies by the Defence Science and Technology (DST) Group and other researchers have investigated discontinuities associated with the etched, anodised and peened surface finishes applied to aluminium alloy (AA) components, as well as different fastener hole finishes (e.g. reamed) [5][6][7][8][9][10][11][12][13][14]. These studies focused on the size, ability to grow fatigue cracks and/or effect on fatigue life of each discontinuity type. ...
Article
Porosity is a well-known cause of fatigue cracking in aluminium alloy components when they are subjected to cyclic loading. This is well understood for many situations where castings are used in critical structure. However, fatigue cracking from porosity can also occur in thick wrought plates of high strength aluminium alloys. Such plates are used in aircraft components and can have significant amounts of shrinkage porosity present due to the lack of sufficient rolling prior to machining. An example of such behaviour has been observed in early production F/A-18 A-D Hornet components, where fatigue cracks nucleated from porosity during several structural durability tests. This demonstrated the potential for porosity to cause fatigue cracks in these airframe components during service. The present study quantifies the fatigue severity of the porosity present in the Aluminium Alloy (AA) 7050-T7451 thick plate that is used for some F/A-18 A-D Hornet critical structural components. To this end, over one hundred polished AA7050-T7451 coupons were fatigue tested to failure and the crack growth from the porosity that precipitated the fatigue failure of each coupon was measured using quantitative fractography. The data measured for each porosity discontinuity was then used to determine the fatigue crack depth that would have produced equivalent crack growth if it had existed at the start of the fatigue life. This crack depth, denoted the equivalent pre-crack size of the porosity, is considered to represent the fatigue severity for each porosity discontinuity examined here and the collated data were used to estimate the distribution of porosity fatigue severities in AA7050-T7451 thick plate. The authors propose that using such a metric to describe the initial condition of a structure facilitates the use of fatigue crack growth prediction models to make deterministic and probabilistic fatigue life predictions for realistic service loading spectra. Moreover, it allows the fatigue severity of porosity to be compared to those of the other discontinuity types that cause fatigue cracks in similarly manufactured AA7050-T7451 components.
Technical Report
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An investigation was made of the effect of starter notch geometry on flight simulation fatigue crack growth in three damage tolerant aluminium sheet alloys, 2024-T3, 2091-T84 and 8090-T81. Changing the starter notch geometry resulted in significant differences in initial non-stabilized fatigue crack growth behaviour. However, there was only a slight effect on the rankings of the alloys for the in-service inspectable fatigue crack growth regime. Based on the results, a proposal for further investigation was made. This proposal should enable determining the influence of several important factors on flight simulation fatigue crack growth in damage tolerant aluminium sheet alloys. These factors are starter notch geometry, alloy yield strength, spectrum clipping level, and the specimen or sheet thickness. The proposal also provides an opportunity to try to bridge the gap that currently exists between short and long fatigue crack growth under flight simulation loading. DATE 950315 PP ref 27 11
Article
Combined strengthening-toughening technologies of several high property titanium alloys, such as TC4-DT, TC6, TC18, TC21 for aviation uses, have been studied via purification, quasi-b heat treatment, quasi-b forging and grain refinement. The effects of microstructure parameters of lamellar structure, basket-weave structure, refined grain structure etc. on the comprehensive mechanical properties of titanium alloys have been analyzed. The results have shown that, to acquire highly comprehensive static mechanical properties and excellent damage tolerance properties, quasi-b treatment and purification processing should be used for medium strength titanium alloys to get high ductility lamellar structure, while quasi-b forging processing be utilized for high strength titanium alloys to obtain high ductility basket-weave structure. Grain refinement processing is very necessary for both the strength levels of titanium alloys.
Article
The equivalent initial flaw size (EIFS) concept was developed nearly 30 years ago in an attempt to account for the initial quality, both manufacturing and material properties, of a structural detail prone to fatigue cracking. Widespread use of this concept has been limited due to the large amount of test data required to develop a reliable EIFS distribution. In this effort, an EIFS distribution was determined for four types of flat, production like transport aircraft fuselage skin joints loaded by remote tension. Two crack growth prediction codes, AFGROW and FASTRAN, were used to not only develop the EIFS but also to compare the crack growth algorithms in each code. The EIFS calculations are prone to compounding errors in the crack growth analysis due to the changing stress intensity factor solutions and stress fields as the crack gets longer. Thus, only including EIFS calculations for mechanically small cracks, crack lengths less than 1.27 mm, results in a mean EIFS of 18.0 μm with a standard deviation of 3.78 μm.
Article
A fatigue lifing framework using a lead crack concept has been developed by the DSTO for metallic primary airframe components The framework is based on years of detailed inspection and analysis of fatigue cracks in many specimens and airframe components and is an important additional tool for determining aircraft component fatigue lives in the Royal Australian Air Force (RAAF) fleet Like the original Damage Tolerance (DT) concept developed by the United States Air Force (USAF) this framework assumes that fatigue cracking begins as soon as an aircraft enters service However there are major and fundamental differences Instead of assuming initial crack sizes and deriving early crack growth behaviour from back-extrapolation of growth data for long cracks the DSTO framework uses data for real cracks growing from small discontinuities inherent to the material and the production of the component Furthermore these data particularly for lead cracks are characterized by exponential crack growth behaviour Because of this common characteristic the DSTO framework can use lead crack growth data to provide reasonable (i e not overly conservative) lower-bound estimates of typical crack growth lives of components starting from small natural discontinuities and continuing up to crack sizes (thus encompassing short-to-long crack growth) that just meet the residual strength requirements Scatter factors based on engineering judgement are then applied to these estimates to determine the maximum allow able service life (safe life limit) The aim of the paper is to present the framework of assumptions and observations used in conjunction with a unique measure of the initiating discontinuity and a simple crack growth law to predict a lower bound fatigue life estimate Crown Copyright (C) 2010 Published by Elsevier Ltd All rights reserved
Chapter
The different microstructures were obtained by three different hot processes for Ti-5.5Al-4Mo-6V-2Nb-1Fe, a new high strength and toughness titanium alloy. The typical microstructure characteristics were observed and analyzed by scanning electron microscope, and the influence of different microstructure on tensile, impact toughness, fracture toughness and crack growth were studied. The results show that all the three different microstructures present good mechanical properties. Especially, the fine lamellar microstructure, obtained by the new beta annealing process, exhibits substantially superior properties relative to the basket microstructure and the rough lamellar microstructure, high strength, high Young’s modulus, very high fracture toughness and very low crack growth rate. This makes the new alloy a promising material for advanced aeronautical application.
Chapter
Standard data on ambient temperature mechanical and environmental properties, including yield and tensile strengths, fatigue and fatigue crack growth, fracture toughness, corrosion and stress corrosion, are essential—indeed mandatory—for the qualification and certification of aerospace structural materials and the design of actual structures and components. This chapter discusses the determination of important ambient temperature mechanical and environmental properties of aerospace alloys at the basic level of specimen and coupon testing.
Article
Analysis of the risk of failure of airframes has seen increased international activity over the past decade. It provides an alternative to the airworthiness standards for maintaining safe operation of ageing aircraft. DSTO has been developing a risk capability in recent years, and seeks to learn as much as possible from international developments and applications. This report conducts a review of the methods and assumptions made in structural risk assessments. Its aim is to provide a guide for conducting a structural risk assessment of an airframe, and as such is limited to techniques generally used in the aircraft industry. The literature was reviewed to investigate how the risk of fatigue failure for an aircraft has previously been determined. Factors include the initial defects that trigger fatigue cracks, the growth of fatigue cracks to failure, the determination of residual strength and the variability in the spectrum loads and the maximum loads that result in failure. Towards the end of the life of an aircraft, the overwhelming risk or probability of failure is caused by the growth of fatigue cracks in the aircraft structure from the loads sustained through usage of the aircraft throughout its lifetime. A summary of the fatigue crack growth laws used in probabilistic risk assessments is given. Guidance is given on what distributions should be used for modeling the various parameters in a probabilistic risk assessment. Data and a knowledge of the mechanisms that produce the variability in each parameter will define the probabilistic distributions that should be used. The most common error found in many of the risk assessments examined is to assume the risk of failure is simply the summation of the individual risks of failure of all the elements that can fail in the aircraft. This is incorrect. The most critical crack will fail first and the risk of failure is simply the probability of the largest crack failing. The remaining cracks in the aircraft do not directly contribute to the risk of failure. This difference is significant when there is a large number of possible fasteners that can initiate failure, such as occur in the wing of transport aircraft. The number of cracks present contribute indirectly by changing the probability distribution of the size of the largest crack. Summaries of a number of risk assessments on different aircraft are provided, such as for the KC135, F-16, C-141, B1-B Bomber, F/A-18 and the B707. These assessments have been used to support keeping the associated aircraft in service. DSTO has learned from these risk assessments in order to apply the capability to ADF aircraft, thereby evaluating and maintaining their safety of flight as they age. This report looks at the published literature on methods and assumptions made in performing structural risk assessments on aircraft. Because the major contributor to the risk of structural failure is fatigue, most methods of risk assessment involve modelling the effect of fatigue growth by some probabilistic method. Many risk assessments use the equivalent initial flaw size approach to allow for the variability in fatigue crack growth. Common errors in the formulation are made in many risk assessments, which can be significant and are described in this report. It is found that the standard approach can produce an acceptable assessment of the probability of failure of an aircraft if care is taken in understanding what is being modelled and the assumptions on which the analysis is based. A number of case studies of risk assessment’s performed on different aircraft are summarised. ASI
Article
Researchers define damage-tolerant structural systems as those systems which not only have adeqate intact strength to withstand initial failure but also adequate residual strength to minimize the possibility of, and hence the consequences of, further failure. The incorporation of damage tolerance cannot be done in total isolation of the function being required of the system and the costs associated with obtaining improved damage tolerance. The approach, therefore, is to formulate multiple-objective, multi-level decision support problems (DSP), the solutions of which represent a compromise between higher costs and higher damage tolerance. Mulitple-objective decision support problems are easily solved in the linear domain. These formulations, however, include both linear and nonlinear constraints and goals, which in the past, have not been considered due to the resulting complexity. Here, researchers: (1) present a complete discussion and description of decision support problems; (2) identify what further research needs to be done in order to obtain information that is required but not known for solving problems using these models; and (3) identify what needs to be done to implement this prototype method in practice.
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