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Airframe metal fatigue revisited
International Journal of Fatigue 131 (2020) 105323
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Airframe metal fatigue revisited☆
T
L. Molent , B. Dixon ⁎
Defence Science and Technology Group, 506 Lorimer Street, Melbourne 3207, Australia
ARTICLE INFO
ABSTRACT
Keywords: Fatigue Defects Cracks Crack growth Airframe airworthiness Structural failures
Good fatigue design and prognostics are essential for safe and sustainable aircraft. A common feature in all metallic fatigue failures is the presence of structural cracks that propagate during the life of the structure. There are fatigue crack growth prediction tools (and certification processes) that have largely ensured flight safety by utilizing crack growth rate data generated from physically large cracks. However, unexpected fatigue cracking and non-critical failures still occur, which increases the cost of aircraft sustainment. This paper attempts to define some of the reasons why these sustainment-related cracking problems can occur and some means of redress. The discussions are limited to production quality metals and fabrication processes.
1. Introduction Failures from metal fatigue and unexpected fatigue cracking that requires active management still occur in airframe components (e.g. [1–3]), despite over 170 years of research into the subject (see [4] for an early example), as reflected in numerous journal papers and conferences (e.g. [5,6]). Metal fatigue encompasses many scientific disciplines and therefore offers a rich variety of phenomena for fundamental and industrial research [7]. With this background in mind, it is hardly surprising that so-called ‘blind predictions’ of airframe fatigue lives may be poor in some instances [8], or that unexpected fatigue problems still occur in certification testing (e.g. the F-35 Joint Strike Fighter [9]) and in service [1–3]. Furthermore, the limitations (accuracy and/or precision) that currently exist with respect to fatigue life predictions mean that conservatism is necessary during design, and therefore aircraft designs may not be optimal for performance or efficiency. Design tools and processes are generally adequate to ensure safety, but are less satisfactory for cost effective sustainment (also referred to as ‘durability’), which may be impacted by unexpected fatigue cracking and non-critical failures. This paper discusses some of the factors that contribute to unexpected fatigue cracking in certification tests and during service. Importantly, the focus is on production airframe components rather than carefully prepared laboratory specimens with highly polished and nominally discontinuity-free surfaces. The scope of the problems posed by these factors can be summarised
as follows, and will be elaborated further in this paper:
• Fatigue phenomena and mechanisms are defined here as localized • • • • • •
material events associated with the creation and/or propagation of cracks; Production materials may have many discontinuities that can nucleate fatigue cracks; Most structural fatigue cracks nucleate and begin to grow early in the life of an airframe component. There are limited cases where discontinuities that occur later in the life of a component may be important: e.g. in-service damage, corrosion, and poor repair; The fatigue cracks of interest are those that can progressively grow from initial sizes to potentially critical sizes (i.e. causing failure or structural impairment) during the life of a component; The fatigue crack growth behaviour depends on the structural fatigue loading (i.e. spectrum), the local structural conditions (stress levels, stress concentration factors, etc.), material, nucleating discontinuity and environment; Upwards of two-thirds of an aircraft component’s total life can be spent growing a crack before it is detectable (> 1 mm long, given the current limitations of non-destructive testing equipment and techniques)1. Thus, by definition, most fatigue cracks in properly designed and managed service aircraft are small; and The size of the largest fatigue crack in a component is generally the only measurable metric of its residual structural integrity. An important exception is widespread fatigue damage (WFD), whereby a
The views expressed in this paper are the authors’ and do not necessarily represent those of DST Group. Corresponding author. E-mail address: [email protected] (L. Molent). 1 The fact that the lower thresholds of most NDI techniques used in service are well above 1mm (see [60] for example) in practice increases the proportion of the fatigue life where the majority of fatigue cracks are not detectable. ☆ ⁎
https://doi.org/10.1016/j.ijfatigue.2019.105323 Received 31 July 2019; Received in revised form 4 October 2019; Accepted 4 October 2019 Available online 09 October 2019 0142-1123/ Crown Copyright © 2019 Published by Elsevier Ltd. All rights reserved.
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local conglomeration of (sub-critical) fatigue cracks may dramatically reduce the residual strength capability below that prevailing in the presence a single (much larger) crack. This phenomenon is mostly confined to ageing transport aircraft [10]. The problems posed by cracks that nucleate due to in-service mechanisms may not be reliably identified via a full-scale fatigue test (FSFT). Such cracks are often found only during in-service inspection and represent a challenge to managing the fleet. Service-induced crack nucleating discontinuities include accidental damage (e.g. scratches, dents, poor fastener extraction and insertion during maintenance), environmentally induced damage (e.g. corrosion pits, exfoliation) and fretting damage. However, cracks that start from manufacturing-induced discontinuities may be more important, even in areas affected by service-induced damage. As an example, while fretting may occur in lap-splice joints [11], the observed fatigue cracking from coupon and in-service fuselage panel lap-joints [12] and a spar splice joint [13] suggested that the largest cracks can start early in the fatigue life and not as the result of later fretting (see also [14]). Any maintenance-induced damage or fretting-induced cracking that occurred during the certification FSFT provides valuable information for the fleet operator. However, a FSFT has limitations in this context, since it cannot replicate the service operating environment and inevitably has local load and stress inaccuracies associated with discrete test fixture load introduction. Also the in-service maintenance practices often deviate from the maintenance needed to continue the FSFT.
Fig. 1. Schematic illustrating the types and sizes of defects/discontinuities present in metallic components (courtesy of J.P. Gallagher).
• Subject to the presence of high stresses, these cracks reveal themselves relatively early in the service life of an aircraft or after a repair/modification (e.g. General Dynamics F-111 and Aermacchi MB326 [2,3], and more recently, the Qantas A380 uncontained engine disc failure [31]). Such cracks are generally not observed in structural certification FSFTs, owing to their rare nature.
b. Early crack growth at structural hot spots that results from design analysis deficiencies:
• Design analysis limitations (e.g. missed loads on components, un•
2. Characteristics of cracks that may lead to airframe failure 2.1. Crack nucleation In the late 19th century it was believed that fatigue was due to the crystallization of material, since the fracture surfaces appeared crystalline at the magnifications available at the time [5,7]. However, since the availability of scanning electron microscopes in the mid-1960s, the general consensus is that the period of fatigue nucleation or initiation (see [6] for definitions) from naturally occurring discontinuities can be a relatively small portion of the total fatigue life of an airframe location (e.g. approximately 0–10%) [15–26]. For the purposes of airframe fatigue life prediction, any period of nucleation can be ignored without undue conservatism, as can a short period of stable tearing that often occurs when a crack is close to its critical depth/size. In fact, while not explicitly stated, these assumptions are inherent to the USAF and FAA damage tolerant design methods (see for example [27,28]). Therefore, in this paper fatigue life is considered to be related only to fatigue cracking.
c. Crack growth from typical manufacturing and fabrication defects and naturally occurring material discontinuities:
• In this category are defects/discontinuities that are generally well • • • •
2.2. Fatigue cracks The following are characteristics of fatigue cracks that could potentially affect the structural integrity of airframes: a. Early crack growth from rare anomalous manufacturing/ fabrication induced defects or material defects:
below the threshold of conventional NDI and could result in short incubation periods; Such defects can be formed during initial fabrication, but also during a repair or modification of the structure; Depending upon their size and how crack-like they are [30,32], these discontinuities could develop cracks with the potential to cause failure if they are not detected before they reach a critical size; The potential for such discontinuities to pose a threat during service may be revealed, and thus potentially rectified/managed, via a thorough FSFT, given that such tests sample large areas of structures under high stresses; and In general, the larger and more crack-like the worst crack-nucleating discontinuity in a component, the shorter is its remaining fatigue life. d. Crack growth from in-service damage:
• Such cracking obviously involves an in-service latency (pre-damage) period before the fatigue-nucleating defect is introduced; • Only those damages that occur in areas of high stresses are likely to
• Rare defects, i.e. those at the upper tail of the discontinuity popu-
• •
representative design spectrum, erroneous stress analyses) result in stresses that are higher than expected; and All aircraft in a fleet have the same design analysis limitations/deficiencies. However, these can be revealed and then rectified via a thorough FSFT that tests a service representative airframe with realistic loading, and for long enough to allow for variations in fatigue performance across a fleet of aircraft.
lation’s size distribution, see Fig. 1, could occur during initial manufacturing and fabrication or during a structural repair. Such defects in service components reflect an unusual lack of quality control and possible limitations in the production and post-repair non-destructive inspections (NDI) designed to prevent such defects from entering service. Some examples of such defects and their causes are given in [30]; Such defects are assumed (and have been shown) to have near zero incubation time before cracks nucleate and start propagating; Such defects do not occur in many airframe locations or in most of the fleet of aircraft; and
• • 2
pose a threat to the structural integrity of the airframe within the service life; Depending on when in the service life the damages occur, how crack-like they are, and the accessibility for inspecting the areas where they occur, cracking from such damages could lead to inservice failures; and The threat posed by such defects may be revealed by inspections, ageing aircraft audits, and teardowns of retired service components.
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critical components2. In many cases, the combination of production process quality and post-production NDI means that an initial defect this large is extremely unlikely to occur at a fracture critical location in any fleet aircraft. Thus while this methodology delivers an appropriate level of safety, it may be excessively conservative and result in unnecessary penalties in aircraft performance and efficiency. A POF–based approach, where an attempt is made to predict the severities of the worst crack nucleating defects, would appear to remedy this potential weakness.
Leaving aside the effect of variable aircraft usage, which can potentially be accounted for with accurate individual aircraft usage tracking (IAT) [33], the distribution of crack sizes present at each critical location in a fleet of aircraft at any time is largely dependent on the sizes and crack-like natures of the discontinuities inherent to the metallic materials or introduced during manufacture/fabrication, plus to some extent, the variability in fatigue crack growth rates for a specific material (see [34–36,29]). Fig. 1 schematically illustrates the expected relative sizes of the different types of defects/discontinuities described above. Considering the overall risk of structural failure posed by such defects/discontinuities, the other key determinant is the frequency of occurrence. This greatly influences the likelihood that each defect/discontinuity type occurs in a structural location where the stress is sufficiently high, and damage tolerance and access for routine inspection is poor enough, that an in-service failure could occur. For example, while abnormal material and manufacturing defects are likely to be considerably more severe than typical discontinuities resulting from manufacturing, they are usually so rare (providing that quality control during manufacture is acceptable) that it is unlikely they will occur in a critical location. On the other hand, the typical discontinuities normally resulting from the manufacturing process are often so numerous as to almost guarantee that they will be present in areas where the airframe is most vulnerable to in-service structural failure.
ii The safe-life design philosophy The safe-life design philosophy is based upon the prediction of the average fatigue life of a component under service-representative loading conditions and usually involves a FSFT or component testing for validation. The average fatigue life is then reduced by appropriate scatter factors to allow for possible underestimation of the average fatigue life with a limited number of test results, variable fatigue performance across a fleet of aircraft and variable fatigue usage when individual aircraft tracking (IAT) is not employed. The safe-life design basis is still used for full airframe design (e.g. U.S. Navy) or a subset of airframe components (e.g. single load path structures for the USAF). In the application of this philosophy, the criterion used to mark the end of a component’s usable fatigue life can vary widely, namely from an engineering crack initiation criterion, to the total available crack growth life at the required residual strength. Particularly noteworthy is the U.S. Navy’s engineering crack initiation basis, which is described in [38]. We have discussed this approach with the author of [39], enabling a concise rationale as follows:
3. Design and certification methods for managing fatigue cracks in aircraft structures Modern aircraft design methods are intended to ensure an acceptable structural Probability of Failure (POF) over the life of type. Some elements of these methods are considered below. i Damage tolerance
• The damage tolerance design philosophy [27,28] was developed for,
• Design shares many of the same production quality considerations as outlined above; • Avoidance of inspections in austere conditions (e.g. during shipborne operations) or at inaccessible locations; • The useful life of an aircraft component is considered the time to
•
•
• • • • • • • •
and has successfully addressed, the potential catastrophic effects of cracks nucleated from rare anomalous defects that could lead to early cracking and failure in structures unpropitiously containing such defects; Implicit consideration of many of the characteristics of high risk fatigue cracks defined above. For example, cracks are assumed to grow from day one of service, as is also the case in typical POF analyses, e.g. [15,16]; Relies on thorough material and process controls, including inspections, in an attempt to preclude rare defects [37]; Design standards require that materials have good resistance to the growth of detectable damage (to facilitate in-service inspection) and that design details are reasonably tolerant to increases in the severity of service load spectra [32]; Fail-safe and structurally redundant multi-element designs are important for improving the chances of detecting damage before structural integrity is compromised; The growth behaviour of large cracks is well understood and has become an integral part of the design process for preventing catastrophic fatigue failures; The crack growth models used to set inspections are calibrated to the results of FSFTs or representative coupon fatigue tests; The NDI sensitivity is a critical part of economically ensuring structural integrity. The ability to detect smaller cracks allows less frequent inspections and allows repairs that have lower cost and availability impacts (e.g. blend-out repairs); The risk of structural failure increases exponentially as the number of detectable cracks increases with time in service [15,16]; and The United States Air Force (USAF) damage tolerance approach is predicated upon the assumption that a surrogate 1.27 mm (0.05 in.) deep crack exists at the beginning of the service life in fracture
grow a discontinuity to a crack depth of 0.254 mm, at which point ‘crack initiation’ is defined to occur. This 0.254 mm crack depth is an engineering definition and corresponds to what is considered to be the lower limit of detection during inspections; and Traditional crack growth models are not reliable for crack depths below 0.254 mm3, and thus, a Pålmgren-Miner type damage accumulation model tends to be used to estimate the fatigue life to this crack depth.
The development of crack growth models that provide accurate predictions in the short crack regime would appear to be a viable alternative to the sometimes unreliable Pålmgren-Miner approach [25]. Crack growth models also provide additional benefits such as crack growth rates applicable (considering the environmental exposure) to the entire fatigue life, which would facilitate the implementation of inspections when operations and location accessibility allow. iii Full-scale fatigue or durability tests
• A thorough full-scale fatigue or durability test using representative
•
spectrum loading has been shown to effectively identify most of the structural locations where an airframe is most vulnerable to fatigue cracking under the expected service spectrum, i.e. fatigue hot spots. The FSFT may show that, due to design oversights, the airframe is vulnerable to fatigue cracking or even failure within the service life. Additional analysis following the FSFT normally ensures that
2 The U.S Federal Aviation Administration (FAA) is less prescriptive, specifying: an initial flaw size of the maximum probable size that could exist as a result of manufacturing or service induced damage [28]. 3 This is often the result of limitations of the constant amplitude crack growth data collected in accordance with ASTM E647 [54], which are used by such models.
3
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L. Molent and B. Dixon
•
•
•
•
•
•
production modifications or in-service maintenance plans address test-determined fatigue cracking problems in a timely fashion. For locations where the applied FSFT load/stress histories do not duplicate in-service usage, the service spectrum needs to be characterized via IAT programmes or via an operational loads measurement (OLM) programmes. An assessment of the FSFT outcomes considering the differences between the FSFT spectrum and fleet fatigue usage then allows appropriate in-service fatigue management. When conducted for long enough to account for fatigue scatter caused by material and build-quality variations, an FSFT identifies when an airframe, or individual component, will be sufficiently cracked that it will no longer ensure structural airworthiness with normal maintenance or inspections and should be retired or the component replaced. The problem of using a single design FSFT to establish the life limits in relation to the onset of WFD (or the Limit of Validity [40]) resulting from typical manufacturing and material discontinuities has not been fully resolved. This problem requires estimation of the crack populations existing in primary airframe components at various times during the aircraft structural life. Combinations of in-service inspections, structural teardowns, FSFTs of late-in-life ex-service components and structural risk assessments provide methods for identifying the onset of WFD and its impact on structural integrity. These methods can also provide validation for the existing fatigue durability assessment and management of the airframe, e.g. [17]; A probabilistic fatigue approach: here a population of fatigue cracks nucleating from representative initial discontinuities is grown forward using realistic service fatigue crack growth predictions, thereby enabling estimations of the likelihood that multi-site damage (MSD) and WFD will occur. In short, such an analysis predicts whether adjacent secondary cracks will be large enough to exert an influence on the crack growth rate and critical crack size of the lead crack late in the service life; and Predicting the initial behaviour of cracks growing from typical manufacturing and material defects (i.e. physically small cracks), and understanding the potential variations in stress state and internal load distributions due to aircraft to aircraft differences in individual component dimensional tolerances and manufacturing variances is an important issue for establishing the time when detectable cracks will be present in the airframe. This prediction is an important consideration for economic airframe sustainment.
prevention treatments, especially considering the worst aircraft in the fleet, may lead to an unexpected crack-nucleating surface discontinuity that was not previously observed during fatigue certification. Such a scenario is exemplified by the in-service failure of a polished second stage compressor disc of a DC-10 aircraft, where post-failure forensic examination detected significant crack growth early in the life of the disc from a surface-breaking discontinuity [42]. The sources of discontinuities listed in [30] are typical of aircraft metallic materials and airframe production. The occurrence of poorly finished holes and shot-peening defects can be avoided to some extent by improved production quality control or changes to the method of application: e.g. laser shock-peening instead of conventional shot-peening. Such changes will improve the size distribution of the cracknucleating discontinuities (i.e. reduce the mean size and size variability). However, quality control via detection during post-production NDI can have limitations due to the generally small size of the discontinuities (typically a mean size of the order of 10 µm [30,32,43]). Other types of crack-nucleating discontinuities are simply unavoidable owing to production processes or trade-offs in production processes: e.g. etch pits from chemical treatments used as a precursor to corrosion prevention schemes; porosity in castings (rarely used in primary airframe structures) or thick plates; the often already-cracked constituent particles inherent to high-strength aluminium alloys; and constituent particles present even in high quality (triple melted) aircraft steels. The abundance of discontinuities introduced during production almost guarantees that they will occur on the surfaces of airframe details that experience the highest local stresses. This means that crack growth must be expected from early in the service life, and such so-called lead cracks will, as their name implies, tend to lead (in terms of size) any fatigue cracks that form as a result of the in-service environment (e.g. corrosion pitting). All else being equal, the fatigue life varies inversely with the size of the initial discontinuity from which a fatigue crack nucleates. Furthermore, the scatter in the fatigue lives of nominally identical components can be attributed largely to the size variations of the worst nucleating discontinuities present (see [16,35]). Note: a small defect at a critical location (i.e. high stress) can have far more impact than a large defect at a benign, lightly stressed location. 4.2. The importance of the short crack regime Given the current limitations of production and in-service NDI techniques, defects or cracks are not likely to be detected until they are approximately 1 mm long (surface length, in ideal conditions). Hence, owing to the small sizes of typical crack-nucleating discontinuities discussed above (on average approximately 10 μm), upwards of twothirds of the total fatigue life of a component can elapse before any cracking becomes detectable (see [44]). The importance and significance of the physically short crack regime (crack depth < 1 mm) has been known for more than 40 years [45,46,48–51,54]. The problems associated with the use of crack growth rate data generated using long-cracks to predict the growth of short cracks were recognised since the mid-1980s [47,49] and are currently well described in Appendix X3 of the ASTM fatigue test standard E647-13a [53]. One such problem is the fact that short cracks appear to violate the conditions applicable to linear elastic fracture mechanics (LEFM). In addition, while a lower threshold for crack growth is often exhibited by crack growth rate data generated using long cracks, Appendix X3 of the ASTM E647-13a also states [53]: ‘‘It is not clear if a measurable threshold exists for the growth of small fatigue cracks”: see also [49]. The foregoing information is consistent with the observation that short cracks commence growing shortly after being exposed to cyclic loads. This implies that the cyclic threshold stress intensity factor range (ΔK) for lead cracks (i.e. those that will fail first) is low, i.e. close to zero [54]. Therefore it is concluded that growth rate data generated using short cracks should be used to predict growth from the crack nucleating
4. Discussion 4.1. Discontinuities Production aircraft components and structures have many sources of surface or near-surface discontinuities from which fatigue cracking may very quickly develop in highly stressed locations once cyclic loading has commenced, i.e. upon entry of the aircraft into service [15–30]. Also, fatigue cracking in most metallic airframe components is primarily surface-influenced, and therefore greatly depends on small surface discontinuities introduced during component production, as well any surface-connected discontinuities inherent to the material [26]. This is a consequence of the fact that whilst sub-surface discontinuities may be present, any subsurface crack growth occurs in a vacuum, where for metals commonly used in aircraft structures, the fatigue crack growth rate is much slower than in air [41]. Highly polishing a component or applying other surface-fatigue enhancing techniques can remove near-surface discontinuities, potentially causing subsurface defects to become the main threat to structural integrity. However, this can make the cracks harder to find during in-service NDI, owing to a lower probability of detection for sub-surface cracks. Furthermore, any shortfall in the implementation of such fatigue 4
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discontinuities that usually occur in aircraft metallic alloys (see also [52–55]). Here, it is also important to distinguish between the design problem, where a damage tolerance approach may be used with the aim of excluding the possibility of an unexpected early fatigue failure, and the problem faced by an operator, who wants to estimate when the worst aircraft in the fleet is likely to have cracking that causes structural impairment or the failure analyst, or who wants to understand the cause of a test or in-service failure. 4.3. Fatigue scatter Variations in aircraft usage lead to significant differences in fatigue lives, which is why most military agile aircraft are fatigue-usage monitored (see [33]). However, for fixed constant amplitude or variable amplitude spectrum loading, the scatter in the fatigue performance of metallic airframe components is governed by the variability in the metal’s material properties and manufacturing quality. These two aspects of the fatigue problem can be quantified, see [16,54], by gaining an understanding of the variability in the:
• •
1. Initial discontinuities that lead to fatigue cracking and in particular their ability to nucleate and grow fatigue cracks; 2. Stress concentrations leading to inter-aircraft variations in local stress; 3. Variations in local stress state and internal load distributions due to individual component dimensional tolerances; 4. Fit-up (assembly) or residual stresses; 5. Fracture toughness of the material; 6. Crack nucleation and/or initiation period if significant or sub-surface fatigue crack growth occurs; 7. Cyclic threshold stress intensity factor, referred to here as ΔKthr. It is noted that short cracks (as well as cracks grown under spectrum loading) do not show a pronounced dip in the da/dN versus ΔK curve when approaching the threshold region, unlike for long crack data (see [52,59,54]); and 8. Fatigue crack growth rate in the material being examined.
Many of these investigations may benefit from applying machine learning and artificial intelligence methods to the large suites of data that are, or could be, collected as part of the sustainment process. b. Suggested improvements to design, certification and sustainment processes:
• Define
•
Items 1–4 are related to the component manufacturing process and reflect the effect of variable build quality on fatigue life scatter. Items 5–8 define the metal’s material property variability; noting that some aspects of item 1 (e.g. porosity and intermetallic particle discontinuities) may also be related to the production of the material. 5. Recommendations a. Further research to enable improved characterization of the behaviour of short cracks, particularly for enhancing airframe sustainment and failure analyses:
•
• Develop methods to characterize the size distributions of the man-
•
(e.g. 10−10 m/cycle) [52,56]. In our opinion such techniques are a viable path to the generation of crack growth data relevant to the evolution of small cracks in metallic airframe components; – These approaches should support an assessment of the threat posed by WFD within the service life. This analysis should estimate the likelihood that the lead cracking condition that defines the usable life of a component (e.g. per [57], a crack with 0.1% probability of occurrence) will be influenced by adjacent secondary cracks; – Predicting the evolution of the expected crack size distribution in an entire fleet of components should also allow an assessment of the viability of using inspections to mitigate the risk of failure late in the component’s service life; Further research to understand the crack nucleation period (if any) that is applicable for structures such as splice joints; and Perform comprehensive assessments that predict for a given airframe the sites that are considered most likely to be affected by rare (anomalous) defects, in-service defects, design analysis defects and typical manufacturing/material discontinuities (e.g. [17,32]). Case studies assessing the influence of such threats to structural integrity throughout the service lives of similar aircraft types may be useful for such predictions.
ufacturing and production-induced crack nucleating discontinuities in metallic airframes. Develop innovative testing/analysis techniques to generate the large data samples necessary to accurately predict the probability distributions that describe crack nucleating discontinuity size variability and crack growth rate variability in metallic airframe components; Define approaches for predicting the propagation behaviour of short crack populations from the crack nucleating discontinuity population to detectable crack sizes. Generating good quality short crack growth rate data can be difficult. It cannot be assumed that cracks from two sides of a notch will grow symmetrically, therefore the crack length (a) should be measured from each edge of the notch (i.e. not as 2a, the distance between the two crack tips). Techniques which incorporate fracture surface marker-bands (progression markings) interspersed with constant amplitude cycles can allow the identification of crack growth increments at very low growth rates
•
• 5
procedures that identify the most appropriate in-service timeframes for conducting airframe teardown inspections, component fatigue tests and structural risk assessments. Such procedures should consider the viability of simplified fatigue cycling of an exservice airframe or components, whereby existing cracks are grown to a readily detectable size, and so greatly increase the service condition data gained from an airframe teardown; Continue the certification FSFT for as long as practical in an attempt to (i) maximize damage accrual at under-tested locations, (ii) highlight all areas where the risk of failure would be greater than the acceptable level within the service life and (iii) determine the time to the onset of WFD. The later observation would provide useful information to validate probabilistic fatigue studies that assess the risk of WFD within the service life. Continued testing beyond the point considered necessary to clear the service life should be considered, if feasible. There is value in highlighting as many of the most fatigue sensitive airframe locations as possible (i.e. the next weakest link in the chain), since unexpectedly severe service loadings and fatigue life extension programmes are not uncommon; Decrease the time taken to certify an airframe by increasing cycling rates to the extent possible while maintaining representative stress states, and using improved NDI to find cracks when small and easily repairable; Thorough teardown inspection of the FSFT article, including destructive inspections and quantitative fractography (QF) of the detected cracks provides a valuable quantification of the fatigue damage that can be incurred in service. Accurate QF of end-of-test crack sizes and crack growth rates during the test can allow an estimation of how the existing damage would have grown beyond the end of the test and up to the predicted limits of damage tolerance. Such techniques have been successfully used by DST to predict testdemonstrated fatigue lives beyond the end point of testing e.g. [17]. The data generated can also be used to assess in-service detected cracking and supplant the often conservative assumptions about initial defect sizes; Improved inspection or other crack detection techniques (e.g. Thermoelastic Stress Analysis [58] or contact strain gauges at
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anticipated hot spots) are required to detect damages/cracks at small sizes, in order to facilitate simple repairs and reduce FSFT down-times. The improved programme schedules delivered by such innovations will allow increased damage to be applied (per the recommendation above) within allowable testing timelines and budgets; and Material surface finishes should be assessed and selected for application with the aim to lower the mean size of the crack nucleating discontinuities and to narrow the related scatter in fatigue performance. Note: Surface enhancement treatments (e.g. shot-peening) can increase fatigue life scatter.
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Short crack data derived from the fatigue data of 2024–T3 Al with long cracks: Material, load and geometry effects locked-in by transitional functions. Theor App Fract Mech 2014;71:2–13. [22] Jones R. Fatigue crack growth and damage tolerance. Fat Fract Eng Mat Struct 2014;37:463–83. [23] Sunder R. Unravelling the science of variable amplitude fatigue. ASTM Spec Tech Publ 2012;1546:20–64. [24] MacLean L, Richman A, Hudak M. Failure rates for aging aircraft. Safety 2018;4(1):7. [25] Schutz W. The prediction of fatigue life in the crack initiation and propagation stages-a state of the art survey. Eng Fract Mech 1979;11:405–21. [26] Przybyla C, Prasannavenkatesan R, Salajegheh N, McDowell DL. Microstructuresensitive modeling of high cycle fatigue. Fatigue 2010;32:512–25. [27] Gallagher JP, Babish CA, Malas JC. Damage tolerant risk analysis techniques for evaluating the structural integrity of aircraft structures. In: Proceedings of the Structural Integrity in Transportation Symposium, 11th Conference on Fracture; 2005 March; Turin Italy. [28] Damage Tolerance and Fatigue Evaluation of Structure, AC 25.571-1D [Advisory Circular], 2011, Federal Aviation Administration. [29] Manning SD, Yang JN. Advanced Durability Analysis Vol. 1 – Analytical Methods. AFWAL-TR-86-3027, Air Force Wright Aeronautical Laboratories, Wright Patterson Air Force Base; 1987. [30] Barter SA, Molent L, Wanhill RJH. Typical fatigue-initiating discontinuities in metallic aircraft structures. Int J Fat 2012;41:11–22. [31] In-flight uncontained engine failure Airbus A380-842, VH-OQA.ATSB, Australia. ATSB Transport Safety Report, Aviation Occurrence Investigation AO-2010-089; 2013. [32] Molent L. A review of equivalent pre-crack sizes in aluminium alloy 7050–T7451. Fat Fract Eng Mat Struct 2014;37:1055–74. [33] Molent L, Aktepe B. Review of fatigue monitoring of agile military aircraft. Fat Fract Eng Mat Struct 2000;23:767–85. [34] Virkler DA, Hillberry BM, Goel PK. The statistical nature of fatigue crack propagation. Wright Patterson Air Force Base (OH), AFFDL; 1978. Technical report no. AFFDL-TR-78-43. [35] Lincoln JW. Effect of aircraft failures on USAF structural requirements: In: Proc. 22nd International Congress of Aeronautical Sciences (ICAS 2000), 27 August–1September 2000, Harragote, UK. [36] Dixon B, Molent L, Barter S. A study of fatigue variability in aluminium alloy 7050–T7451. Int J Fatigue 2016;92(1):130–46. [37] Lincoln JW. Structural technology transition to new aircraft, proc. ICAF Symposium, Ottawa, Canada; 1987. [38] Fatigue of Aircraft Structures. Naval Air Systems Command, Department of the
6. Conclusions Despite over 170 years of research into the subject and numerous papers and conferences, in-service failures from cyclic mechanical fatigue still occur in metallic airframe components. Furthermore, blind predictions of the fatigue lives of airframe components subjected to simulated service loading are often poor, and unexpected fatigue problems occur during certification testing and service. It is considered that current fatigue management methods, whether damage tolerance (as it is currently applied) or safe life, could better account for the true, as-produced condition of metallic components, and in particular, the population of discontinuities that could potentially nucleate fatigue cracks. In addition, most current fatigue prediction methodologies cannot accurately predict the growth of the small cracks that nucleate from such discontinuities. To overcome such limitations, fatigue management methods can tend to have undesirable conservatisms, which leads to sub-optimal structural designs. Furthermore, these limitations also mean that the true risk of failure for an airframe is not well understood at any time during the service life, which can lead to fatigue management practices that do not optimally balance maintaining acceptable levels of safety, with maximizing aircraft availability and minimizing the outlay of resources. This paper identifies what the authors believe are the key factors that influence the risk of metallic airframe component failure from cyclic fatigue with the intention of focusing future research and development on the areas that would most productively overcome current limitations. In particular the growth of physically short cracks subject to high stresses and loading spectra, which nucleate upon introduction into service, from naturally occurring material and production-induced discontinuities, is considered an area of crucial importance. Declaration of Competing Interest The author declare that there is no conflict of interest. Acknowledgments The authors are indebted to numerous mentors and practitioners in the field of aircraft structural integrity. These include Dr. R.J.H. Wanhill, Mr. D. Polakovics, Mr R. Ryan, Dr. J.P. Gallagher, Prof. R. Jones, Dr. G.C. Sih, Dr. L.M. Butkus, Dr. D.W. Hoeppner, Mr C. Babish and many colleagues at the Defence Science and Technology Group and the Defence Aviation Safety Authority. Appendix A. Supplementary material Supplementary data to this article can be found online at https:// doi.org/10.1016/j.ijfatigue.2019.105323. References [1] Tiffany CF, Gallagher JP, Babish IV CA. Threats to aircraft structural safety, including a compendium of selected structural accidents/incidents. Wright-Patterson Air Force Base (OH), Aeronautical Systems Centre; 2010. Technical report no. ASC-
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L. Molent and B. Dixon Navy, Report no. NAVAIR 01-1A-13; 1966. [39] Polakovics D. Aging aircraft: the U.S. Navy experience. In: International Conference on Aircraft Damage Assessment and Repair; 1991 August; Melbourne, Australia. [40] Establishing and Implementing Limit of Validity (LOV) to Prevent Wide Spread Fatigue Damage. FAA. Advisory Circular AC No. 120-140; 2011. [41] Wanhill RJH. Fractography of fatigue crack propagation in 2024–T3 and 7075–T6 aluminum alloys in air and vacuum. Met Trans 1975;6:1587–96. [42] United Airlines Flight 232 McDonnell Douglas DC-10—10, Sioux Gateway Airport, Sioux City, Iowa. FAA, 1990. Official Accident Report NSTB/AAR-90/06; 1989. [43] Gallagher JP, Molent L. The equivalence of EPS and EIFS based on the same crack growth life data. Int J Fat 2015;80:162–70. [44] Gallagher JP, Molent L. Effect of load spectra and stress magnitude on crack growth behaviour variability from typical manufacturing defects. Adv Mat Res 2014;891–892:100–5. [45] Pearson S. Initiation of fatigue cracks in commercial aluminium alloys and the subsequent propagation of very short cracks. Eng Fract Mech 1975;7(2):235–40. [46] Suresh S, Ritchie RO. Propagation of short fatigue cracks. Int Met Rev 1984;29(6):445–76. [47] Wanhill RJH. Short cracks in aerospace structures. In: Miller KJ, de los Rios ER, editors. The Behaviour of Short Fatigue Cracks, Mechanical Engineering Publications, London; 1986. p. 27–36. [48] Ritchie RO, Yu W, Blom AF, Holm DK. An analysis of crack tip shielding in aluminium alloy 2124: a comparison of large, small, through-thickness and surface fatigue cracks. Fat Fract Eng Mat Struct 1987;10(5):343–62. [49] Wanhill RJH. Durability analysis using short and long fatigue crack growth data. In: Jones R, Miller NJ, editors. International conference on aircraft damage assessment and repair, institution of engineers, Melbourne 26–28 August 1991, Barton ACT, Australia,1991, pp. 100–104.
[50] Lincoln JW, Melliere RA. Economic life determination for a military aircraft. Aircraft 1999;36(5):737–42. [51] Ravichandra KS, Richie RO, Murakami Y. Small fatigue cracks: mechanics, mechanisms and applications. Elsevier Science; 1999. [52] Walker KF, Barter SA. The critical importance of correctly characterising fatigue crack growth rates in the threshold regime. In: 26th ICAF Symposium; 2011 1–3 June; Montreal. Netherlands: Springer; 2011. [53] Standard Test Method for Measurement for Fatigue Crack Growth Rates. West Conshohocken (PA): ASTM International. ASTM E647-13a; 2013 < http://www. astm.org > . [54] Molent L, Jones R. The influence of cyclic stress intensity threshold on fatigue life scatter. Int J Fat 2015;82:748–56. [55] Main B, Evans R, Walker K, Yu X, Molent L. Lessons from a fatigue prediction challenge for an aircraft wing shear tie post. J Fatigue 2019;123:53–65. [56] Barter S, Burchill M, Jones M. Improving the prediction of small crack growth in 7XXX aluminium alloys. In: 28th ICAF Symposium; 2015 3–5 June; Helsinki. [57] DEF STAN 00-970: Design and Airworthiness for Service Aircraft, Issue 6; Ministry of Defence, UK; 2010. [58] Rajic N, Brooks C. Automated crack detection and crack growth rate measurement using thermoelasticity. Proc Eng 2017;188:463–70. [59] Burchill M, Barter S, Chan LH. Improving fatigue life predictions with a crack growth rate model based on small crack growth & legacy data. In: 17th Australian International Aerospace Congress (AIAC17), Melbourne, Australia, 26–28 February 2017. [60] Nondestructive Inspection Capability Guidelines for United States Air Force Aircraft Structures, EN-SB-08-012, Rev. B [Structures Bulletin], 2011, Air Force Structures, Wright Patterson Air Force Base.
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International Journal of Fatigue journal homepage: www.elsevier.com/locate/ijfatigue
Airframe metal fatigue revisited☆
T
L. Molent , B. Dixon ⁎
Defence Science and Technology Group, 506 Lorimer Street, Melbourne 3207, Australia
ARTICLE INFO
ABSTRACT
Keywords: Fatigue Defects Cracks Crack growth Airframe airworthiness Structural failures
Good fatigue design and prognostics are essential for safe and sustainable aircraft. A common feature in all metallic fatigue failures is the presence of structural cracks that propagate during the life of the structure. There are fatigue crack growth prediction tools (and certification processes) that have largely ensured flight safety by utilizing crack growth rate data generated from physically large cracks. However, unexpected fatigue cracking and non-critical failures still occur, which increases the cost of aircraft sustainment. This paper attempts to define some of the reasons why these sustainment-related cracking problems can occur and some means of redress. The discussions are limited to production quality metals and fabrication processes.
1. Introduction Failures from metal fatigue and unexpected fatigue cracking that requires active management still occur in airframe components (e.g. [1–3]), despite over 170 years of research into the subject (see [4] for an early example), as reflected in numerous journal papers and conferences (e.g. [5,6]). Metal fatigue encompasses many scientific disciplines and therefore offers a rich variety of phenomena for fundamental and industrial research [7]. With this background in mind, it is hardly surprising that so-called ‘blind predictions’ of airframe fatigue lives may be poor in some instances [8], or that unexpected fatigue problems still occur in certification testing (e.g. the F-35 Joint Strike Fighter [9]) and in service [1–3]. Furthermore, the limitations (accuracy and/or precision) that currently exist with respect to fatigue life predictions mean that conservatism is necessary during design, and therefore aircraft designs may not be optimal for performance or efficiency. Design tools and processes are generally adequate to ensure safety, but are less satisfactory for cost effective sustainment (also referred to as ‘durability’), which may be impacted by unexpected fatigue cracking and non-critical failures. This paper discusses some of the factors that contribute to unexpected fatigue cracking in certification tests and during service. Importantly, the focus is on production airframe components rather than carefully prepared laboratory specimens with highly polished and nominally discontinuity-free surfaces. The scope of the problems posed by these factors can be summarised
as follows, and will be elaborated further in this paper:
• Fatigue phenomena and mechanisms are defined here as localized • • • • • •
material events associated with the creation and/or propagation of cracks; Production materials may have many discontinuities that can nucleate fatigue cracks; Most structural fatigue cracks nucleate and begin to grow early in the life of an airframe component. There are limited cases where discontinuities that occur later in the life of a component may be important: e.g. in-service damage, corrosion, and poor repair; The fatigue cracks of interest are those that can progressively grow from initial sizes to potentially critical sizes (i.e. causing failure or structural impairment) during the life of a component; The fatigue crack growth behaviour depends on the structural fatigue loading (i.e. spectrum), the local structural conditions (stress levels, stress concentration factors, etc.), material, nucleating discontinuity and environment; Upwards of two-thirds of an aircraft component’s total life can be spent growing a crack before it is detectable (> 1 mm long, given the current limitations of non-destructive testing equipment and techniques)1. Thus, by definition, most fatigue cracks in properly designed and managed service aircraft are small; and The size of the largest fatigue crack in a component is generally the only measurable metric of its residual structural integrity. An important exception is widespread fatigue damage (WFD), whereby a
The views expressed in this paper are the authors’ and do not necessarily represent those of DST Group. Corresponding author. E-mail address: [email protected] (L. Molent). 1 The fact that the lower thresholds of most NDI techniques used in service are well above 1mm (see [60] for example) in practice increases the proportion of the fatigue life where the majority of fatigue cracks are not detectable. ☆ ⁎
https://doi.org/10.1016/j.ijfatigue.2019.105323 Received 31 July 2019; Received in revised form 4 October 2019; Accepted 4 October 2019 Available online 09 October 2019 0142-1123/ Crown Copyright © 2019 Published by Elsevier Ltd. All rights reserved.
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local conglomeration of (sub-critical) fatigue cracks may dramatically reduce the residual strength capability below that prevailing in the presence a single (much larger) crack. This phenomenon is mostly confined to ageing transport aircraft [10]. The problems posed by cracks that nucleate due to in-service mechanisms may not be reliably identified via a full-scale fatigue test (FSFT). Such cracks are often found only during in-service inspection and represent a challenge to managing the fleet. Service-induced crack nucleating discontinuities include accidental damage (e.g. scratches, dents, poor fastener extraction and insertion during maintenance), environmentally induced damage (e.g. corrosion pits, exfoliation) and fretting damage. However, cracks that start from manufacturing-induced discontinuities may be more important, even in areas affected by service-induced damage. As an example, while fretting may occur in lap-splice joints [11], the observed fatigue cracking from coupon and in-service fuselage panel lap-joints [12] and a spar splice joint [13] suggested that the largest cracks can start early in the fatigue life and not as the result of later fretting (see also [14]). Any maintenance-induced damage or fretting-induced cracking that occurred during the certification FSFT provides valuable information for the fleet operator. However, a FSFT has limitations in this context, since it cannot replicate the service operating environment and inevitably has local load and stress inaccuracies associated with discrete test fixture load introduction. Also the in-service maintenance practices often deviate from the maintenance needed to continue the FSFT.
Fig. 1. Schematic illustrating the types and sizes of defects/discontinuities present in metallic components (courtesy of J.P. Gallagher).
• Subject to the presence of high stresses, these cracks reveal themselves relatively early in the service life of an aircraft or after a repair/modification (e.g. General Dynamics F-111 and Aermacchi MB326 [2,3], and more recently, the Qantas A380 uncontained engine disc failure [31]). Such cracks are generally not observed in structural certification FSFTs, owing to their rare nature.
b. Early crack growth at structural hot spots that results from design analysis deficiencies:
• Design analysis limitations (e.g. missed loads on components, un•
2. Characteristics of cracks that may lead to airframe failure 2.1. Crack nucleation In the late 19th century it was believed that fatigue was due to the crystallization of material, since the fracture surfaces appeared crystalline at the magnifications available at the time [5,7]. However, since the availability of scanning electron microscopes in the mid-1960s, the general consensus is that the period of fatigue nucleation or initiation (see [6] for definitions) from naturally occurring discontinuities can be a relatively small portion of the total fatigue life of an airframe location (e.g. approximately 0–10%) [15–26]. For the purposes of airframe fatigue life prediction, any period of nucleation can be ignored without undue conservatism, as can a short period of stable tearing that often occurs when a crack is close to its critical depth/size. In fact, while not explicitly stated, these assumptions are inherent to the USAF and FAA damage tolerant design methods (see for example [27,28]). Therefore, in this paper fatigue life is considered to be related only to fatigue cracking.
c. Crack growth from typical manufacturing and fabrication defects and naturally occurring material discontinuities:
• In this category are defects/discontinuities that are generally well • • • •
2.2. Fatigue cracks The following are characteristics of fatigue cracks that could potentially affect the structural integrity of airframes: a. Early crack growth from rare anomalous manufacturing/ fabrication induced defects or material defects:
below the threshold of conventional NDI and could result in short incubation periods; Such defects can be formed during initial fabrication, but also during a repair or modification of the structure; Depending upon their size and how crack-like they are [30,32], these discontinuities could develop cracks with the potential to cause failure if they are not detected before they reach a critical size; The potential for such discontinuities to pose a threat during service may be revealed, and thus potentially rectified/managed, via a thorough FSFT, given that such tests sample large areas of structures under high stresses; and In general, the larger and more crack-like the worst crack-nucleating discontinuity in a component, the shorter is its remaining fatigue life. d. Crack growth from in-service damage:
• Such cracking obviously involves an in-service latency (pre-damage) period before the fatigue-nucleating defect is introduced; • Only those damages that occur in areas of high stresses are likely to
• Rare defects, i.e. those at the upper tail of the discontinuity popu-
• •
representative design spectrum, erroneous stress analyses) result in stresses that are higher than expected; and All aircraft in a fleet have the same design analysis limitations/deficiencies. However, these can be revealed and then rectified via a thorough FSFT that tests a service representative airframe with realistic loading, and for long enough to allow for variations in fatigue performance across a fleet of aircraft.
lation’s size distribution, see Fig. 1, could occur during initial manufacturing and fabrication or during a structural repair. Such defects in service components reflect an unusual lack of quality control and possible limitations in the production and post-repair non-destructive inspections (NDI) designed to prevent such defects from entering service. Some examples of such defects and their causes are given in [30]; Such defects are assumed (and have been shown) to have near zero incubation time before cracks nucleate and start propagating; Such defects do not occur in many airframe locations or in most of the fleet of aircraft; and
• • 2
pose a threat to the structural integrity of the airframe within the service life; Depending on when in the service life the damages occur, how crack-like they are, and the accessibility for inspecting the areas where they occur, cracking from such damages could lead to inservice failures; and The threat posed by such defects may be revealed by inspections, ageing aircraft audits, and teardowns of retired service components.
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critical components2. In many cases, the combination of production process quality and post-production NDI means that an initial defect this large is extremely unlikely to occur at a fracture critical location in any fleet aircraft. Thus while this methodology delivers an appropriate level of safety, it may be excessively conservative and result in unnecessary penalties in aircraft performance and efficiency. A POF–based approach, where an attempt is made to predict the severities of the worst crack nucleating defects, would appear to remedy this potential weakness.
Leaving aside the effect of variable aircraft usage, which can potentially be accounted for with accurate individual aircraft usage tracking (IAT) [33], the distribution of crack sizes present at each critical location in a fleet of aircraft at any time is largely dependent on the sizes and crack-like natures of the discontinuities inherent to the metallic materials or introduced during manufacture/fabrication, plus to some extent, the variability in fatigue crack growth rates for a specific material (see [34–36,29]). Fig. 1 schematically illustrates the expected relative sizes of the different types of defects/discontinuities described above. Considering the overall risk of structural failure posed by such defects/discontinuities, the other key determinant is the frequency of occurrence. This greatly influences the likelihood that each defect/discontinuity type occurs in a structural location where the stress is sufficiently high, and damage tolerance and access for routine inspection is poor enough, that an in-service failure could occur. For example, while abnormal material and manufacturing defects are likely to be considerably more severe than typical discontinuities resulting from manufacturing, they are usually so rare (providing that quality control during manufacture is acceptable) that it is unlikely they will occur in a critical location. On the other hand, the typical discontinuities normally resulting from the manufacturing process are often so numerous as to almost guarantee that they will be present in areas where the airframe is most vulnerable to in-service structural failure.
ii The safe-life design philosophy The safe-life design philosophy is based upon the prediction of the average fatigue life of a component under service-representative loading conditions and usually involves a FSFT or component testing for validation. The average fatigue life is then reduced by appropriate scatter factors to allow for possible underestimation of the average fatigue life with a limited number of test results, variable fatigue performance across a fleet of aircraft and variable fatigue usage when individual aircraft tracking (IAT) is not employed. The safe-life design basis is still used for full airframe design (e.g. U.S. Navy) or a subset of airframe components (e.g. single load path structures for the USAF). In the application of this philosophy, the criterion used to mark the end of a component’s usable fatigue life can vary widely, namely from an engineering crack initiation criterion, to the total available crack growth life at the required residual strength. Particularly noteworthy is the U.S. Navy’s engineering crack initiation basis, which is described in [38]. We have discussed this approach with the author of [39], enabling a concise rationale as follows:
3. Design and certification methods for managing fatigue cracks in aircraft structures Modern aircraft design methods are intended to ensure an acceptable structural Probability of Failure (POF) over the life of type. Some elements of these methods are considered below. i Damage tolerance
• The damage tolerance design philosophy [27,28] was developed for,
• Design shares many of the same production quality considerations as outlined above; • Avoidance of inspections in austere conditions (e.g. during shipborne operations) or at inaccessible locations; • The useful life of an aircraft component is considered the time to
•
•
• • • • • • • •
and has successfully addressed, the potential catastrophic effects of cracks nucleated from rare anomalous defects that could lead to early cracking and failure in structures unpropitiously containing such defects; Implicit consideration of many of the characteristics of high risk fatigue cracks defined above. For example, cracks are assumed to grow from day one of service, as is also the case in typical POF analyses, e.g. [15,16]; Relies on thorough material and process controls, including inspections, in an attempt to preclude rare defects [37]; Design standards require that materials have good resistance to the growth of detectable damage (to facilitate in-service inspection) and that design details are reasonably tolerant to increases in the severity of service load spectra [32]; Fail-safe and structurally redundant multi-element designs are important for improving the chances of detecting damage before structural integrity is compromised; The growth behaviour of large cracks is well understood and has become an integral part of the design process for preventing catastrophic fatigue failures; The crack growth models used to set inspections are calibrated to the results of FSFTs or representative coupon fatigue tests; The NDI sensitivity is a critical part of economically ensuring structural integrity. The ability to detect smaller cracks allows less frequent inspections and allows repairs that have lower cost and availability impacts (e.g. blend-out repairs); The risk of structural failure increases exponentially as the number of detectable cracks increases with time in service [15,16]; and The United States Air Force (USAF) damage tolerance approach is predicated upon the assumption that a surrogate 1.27 mm (0.05 in.) deep crack exists at the beginning of the service life in fracture
grow a discontinuity to a crack depth of 0.254 mm, at which point ‘crack initiation’ is defined to occur. This 0.254 mm crack depth is an engineering definition and corresponds to what is considered to be the lower limit of detection during inspections; and Traditional crack growth models are not reliable for crack depths below 0.254 mm3, and thus, a Pålmgren-Miner type damage accumulation model tends to be used to estimate the fatigue life to this crack depth.
The development of crack growth models that provide accurate predictions in the short crack regime would appear to be a viable alternative to the sometimes unreliable Pålmgren-Miner approach [25]. Crack growth models also provide additional benefits such as crack growth rates applicable (considering the environmental exposure) to the entire fatigue life, which would facilitate the implementation of inspections when operations and location accessibility allow. iii Full-scale fatigue or durability tests
• A thorough full-scale fatigue or durability test using representative
•
spectrum loading has been shown to effectively identify most of the structural locations where an airframe is most vulnerable to fatigue cracking under the expected service spectrum, i.e. fatigue hot spots. The FSFT may show that, due to design oversights, the airframe is vulnerable to fatigue cracking or even failure within the service life. Additional analysis following the FSFT normally ensures that
2 The U.S Federal Aviation Administration (FAA) is less prescriptive, specifying: an initial flaw size of the maximum probable size that could exist as a result of manufacturing or service induced damage [28]. 3 This is often the result of limitations of the constant amplitude crack growth data collected in accordance with ASTM E647 [54], which are used by such models.
3
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•
•
•
•
•
•
production modifications or in-service maintenance plans address test-determined fatigue cracking problems in a timely fashion. For locations where the applied FSFT load/stress histories do not duplicate in-service usage, the service spectrum needs to be characterized via IAT programmes or via an operational loads measurement (OLM) programmes. An assessment of the FSFT outcomes considering the differences between the FSFT spectrum and fleet fatigue usage then allows appropriate in-service fatigue management. When conducted for long enough to account for fatigue scatter caused by material and build-quality variations, an FSFT identifies when an airframe, or individual component, will be sufficiently cracked that it will no longer ensure structural airworthiness with normal maintenance or inspections and should be retired or the component replaced. The problem of using a single design FSFT to establish the life limits in relation to the onset of WFD (or the Limit of Validity [40]) resulting from typical manufacturing and material discontinuities has not been fully resolved. This problem requires estimation of the crack populations existing in primary airframe components at various times during the aircraft structural life. Combinations of in-service inspections, structural teardowns, FSFTs of late-in-life ex-service components and structural risk assessments provide methods for identifying the onset of WFD and its impact on structural integrity. These methods can also provide validation for the existing fatigue durability assessment and management of the airframe, e.g. [17]; A probabilistic fatigue approach: here a population of fatigue cracks nucleating from representative initial discontinuities is grown forward using realistic service fatigue crack growth predictions, thereby enabling estimations of the likelihood that multi-site damage (MSD) and WFD will occur. In short, such an analysis predicts whether adjacent secondary cracks will be large enough to exert an influence on the crack growth rate and critical crack size of the lead crack late in the service life; and Predicting the initial behaviour of cracks growing from typical manufacturing and material defects (i.e. physically small cracks), and understanding the potential variations in stress state and internal load distributions due to aircraft to aircraft differences in individual component dimensional tolerances and manufacturing variances is an important issue for establishing the time when detectable cracks will be present in the airframe. This prediction is an important consideration for economic airframe sustainment.
prevention treatments, especially considering the worst aircraft in the fleet, may lead to an unexpected crack-nucleating surface discontinuity that was not previously observed during fatigue certification. Such a scenario is exemplified by the in-service failure of a polished second stage compressor disc of a DC-10 aircraft, where post-failure forensic examination detected significant crack growth early in the life of the disc from a surface-breaking discontinuity [42]. The sources of discontinuities listed in [30] are typical of aircraft metallic materials and airframe production. The occurrence of poorly finished holes and shot-peening defects can be avoided to some extent by improved production quality control or changes to the method of application: e.g. laser shock-peening instead of conventional shot-peening. Such changes will improve the size distribution of the cracknucleating discontinuities (i.e. reduce the mean size and size variability). However, quality control via detection during post-production NDI can have limitations due to the generally small size of the discontinuities (typically a mean size of the order of 10 µm [30,32,43]). Other types of crack-nucleating discontinuities are simply unavoidable owing to production processes or trade-offs in production processes: e.g. etch pits from chemical treatments used as a precursor to corrosion prevention schemes; porosity in castings (rarely used in primary airframe structures) or thick plates; the often already-cracked constituent particles inherent to high-strength aluminium alloys; and constituent particles present even in high quality (triple melted) aircraft steels. The abundance of discontinuities introduced during production almost guarantees that they will occur on the surfaces of airframe details that experience the highest local stresses. This means that crack growth must be expected from early in the service life, and such so-called lead cracks will, as their name implies, tend to lead (in terms of size) any fatigue cracks that form as a result of the in-service environment (e.g. corrosion pitting). All else being equal, the fatigue life varies inversely with the size of the initial discontinuity from which a fatigue crack nucleates. Furthermore, the scatter in the fatigue lives of nominally identical components can be attributed largely to the size variations of the worst nucleating discontinuities present (see [16,35]). Note: a small defect at a critical location (i.e. high stress) can have far more impact than a large defect at a benign, lightly stressed location. 4.2. The importance of the short crack regime Given the current limitations of production and in-service NDI techniques, defects or cracks are not likely to be detected until they are approximately 1 mm long (surface length, in ideal conditions). Hence, owing to the small sizes of typical crack-nucleating discontinuities discussed above (on average approximately 10 μm), upwards of twothirds of the total fatigue life of a component can elapse before any cracking becomes detectable (see [44]). The importance and significance of the physically short crack regime (crack depth < 1 mm) has been known for more than 40 years [45,46,48–51,54]. The problems associated with the use of crack growth rate data generated using long-cracks to predict the growth of short cracks were recognised since the mid-1980s [47,49] and are currently well described in Appendix X3 of the ASTM fatigue test standard E647-13a [53]. One such problem is the fact that short cracks appear to violate the conditions applicable to linear elastic fracture mechanics (LEFM). In addition, while a lower threshold for crack growth is often exhibited by crack growth rate data generated using long cracks, Appendix X3 of the ASTM E647-13a also states [53]: ‘‘It is not clear if a measurable threshold exists for the growth of small fatigue cracks”: see also [49]. The foregoing information is consistent with the observation that short cracks commence growing shortly after being exposed to cyclic loads. This implies that the cyclic threshold stress intensity factor range (ΔK) for lead cracks (i.e. those that will fail first) is low, i.e. close to zero [54]. Therefore it is concluded that growth rate data generated using short cracks should be used to predict growth from the crack nucleating
4. Discussion 4.1. Discontinuities Production aircraft components and structures have many sources of surface or near-surface discontinuities from which fatigue cracking may very quickly develop in highly stressed locations once cyclic loading has commenced, i.e. upon entry of the aircraft into service [15–30]. Also, fatigue cracking in most metallic airframe components is primarily surface-influenced, and therefore greatly depends on small surface discontinuities introduced during component production, as well any surface-connected discontinuities inherent to the material [26]. This is a consequence of the fact that whilst sub-surface discontinuities may be present, any subsurface crack growth occurs in a vacuum, where for metals commonly used in aircraft structures, the fatigue crack growth rate is much slower than in air [41]. Highly polishing a component or applying other surface-fatigue enhancing techniques can remove near-surface discontinuities, potentially causing subsurface defects to become the main threat to structural integrity. However, this can make the cracks harder to find during in-service NDI, owing to a lower probability of detection for sub-surface cracks. Furthermore, any shortfall in the implementation of such fatigue 4
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discontinuities that usually occur in aircraft metallic alloys (see also [52–55]). Here, it is also important to distinguish between the design problem, where a damage tolerance approach may be used with the aim of excluding the possibility of an unexpected early fatigue failure, and the problem faced by an operator, who wants to estimate when the worst aircraft in the fleet is likely to have cracking that causes structural impairment or the failure analyst, or who wants to understand the cause of a test or in-service failure. 4.3. Fatigue scatter Variations in aircraft usage lead to significant differences in fatigue lives, which is why most military agile aircraft are fatigue-usage monitored (see [33]). However, for fixed constant amplitude or variable amplitude spectrum loading, the scatter in the fatigue performance of metallic airframe components is governed by the variability in the metal’s material properties and manufacturing quality. These two aspects of the fatigue problem can be quantified, see [16,54], by gaining an understanding of the variability in the:
• •
1. Initial discontinuities that lead to fatigue cracking and in particular their ability to nucleate and grow fatigue cracks; 2. Stress concentrations leading to inter-aircraft variations in local stress; 3. Variations in local stress state and internal load distributions due to individual component dimensional tolerances; 4. Fit-up (assembly) or residual stresses; 5. Fracture toughness of the material; 6. Crack nucleation and/or initiation period if significant or sub-surface fatigue crack growth occurs; 7. Cyclic threshold stress intensity factor, referred to here as ΔKthr. It is noted that short cracks (as well as cracks grown under spectrum loading) do not show a pronounced dip in the da/dN versus ΔK curve when approaching the threshold region, unlike for long crack data (see [52,59,54]); and 8. Fatigue crack growth rate in the material being examined.
Many of these investigations may benefit from applying machine learning and artificial intelligence methods to the large suites of data that are, or could be, collected as part of the sustainment process. b. Suggested improvements to design, certification and sustainment processes:
• Define
•
Items 1–4 are related to the component manufacturing process and reflect the effect of variable build quality on fatigue life scatter. Items 5–8 define the metal’s material property variability; noting that some aspects of item 1 (e.g. porosity and intermetallic particle discontinuities) may also be related to the production of the material. 5. Recommendations a. Further research to enable improved characterization of the behaviour of short cracks, particularly for enhancing airframe sustainment and failure analyses:
•
• Develop methods to characterize the size distributions of the man-
•
(e.g. 10−10 m/cycle) [52,56]. In our opinion such techniques are a viable path to the generation of crack growth data relevant to the evolution of small cracks in metallic airframe components; – These approaches should support an assessment of the threat posed by WFD within the service life. This analysis should estimate the likelihood that the lead cracking condition that defines the usable life of a component (e.g. per [57], a crack with 0.1% probability of occurrence) will be influenced by adjacent secondary cracks; – Predicting the evolution of the expected crack size distribution in an entire fleet of components should also allow an assessment of the viability of using inspections to mitigate the risk of failure late in the component’s service life; Further research to understand the crack nucleation period (if any) that is applicable for structures such as splice joints; and Perform comprehensive assessments that predict for a given airframe the sites that are considered most likely to be affected by rare (anomalous) defects, in-service defects, design analysis defects and typical manufacturing/material discontinuities (e.g. [17,32]). Case studies assessing the influence of such threats to structural integrity throughout the service lives of similar aircraft types may be useful for such predictions.
ufacturing and production-induced crack nucleating discontinuities in metallic airframes. Develop innovative testing/analysis techniques to generate the large data samples necessary to accurately predict the probability distributions that describe crack nucleating discontinuity size variability and crack growth rate variability in metallic airframe components; Define approaches for predicting the propagation behaviour of short crack populations from the crack nucleating discontinuity population to detectable crack sizes. Generating good quality short crack growth rate data can be difficult. It cannot be assumed that cracks from two sides of a notch will grow symmetrically, therefore the crack length (a) should be measured from each edge of the notch (i.e. not as 2a, the distance between the two crack tips). Techniques which incorporate fracture surface marker-bands (progression markings) interspersed with constant amplitude cycles can allow the identification of crack growth increments at very low growth rates
•
• 5
procedures that identify the most appropriate in-service timeframes for conducting airframe teardown inspections, component fatigue tests and structural risk assessments. Such procedures should consider the viability of simplified fatigue cycling of an exservice airframe or components, whereby existing cracks are grown to a readily detectable size, and so greatly increase the service condition data gained from an airframe teardown; Continue the certification FSFT for as long as practical in an attempt to (i) maximize damage accrual at under-tested locations, (ii) highlight all areas where the risk of failure would be greater than the acceptable level within the service life and (iii) determine the time to the onset of WFD. The later observation would provide useful information to validate probabilistic fatigue studies that assess the risk of WFD within the service life. Continued testing beyond the point considered necessary to clear the service life should be considered, if feasible. There is value in highlighting as many of the most fatigue sensitive airframe locations as possible (i.e. the next weakest link in the chain), since unexpectedly severe service loadings and fatigue life extension programmes are not uncommon; Decrease the time taken to certify an airframe by increasing cycling rates to the extent possible while maintaining representative stress states, and using improved NDI to find cracks when small and easily repairable; Thorough teardown inspection of the FSFT article, including destructive inspections and quantitative fractography (QF) of the detected cracks provides a valuable quantification of the fatigue damage that can be incurred in service. Accurate QF of end-of-test crack sizes and crack growth rates during the test can allow an estimation of how the existing damage would have grown beyond the end of the test and up to the predicted limits of damage tolerance. Such techniques have been successfully used by DST to predict testdemonstrated fatigue lives beyond the end point of testing e.g. [17]. The data generated can also be used to assess in-service detected cracking and supplant the often conservative assumptions about initial defect sizes; Improved inspection or other crack detection techniques (e.g. Thermoelastic Stress Analysis [58] or contact strain gauges at
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anticipated hot spots) are required to detect damages/cracks at small sizes, in order to facilitate simple repairs and reduce FSFT down-times. The improved programme schedules delivered by such innovations will allow increased damage to be applied (per the recommendation above) within allowable testing timelines and budgets; and Material surface finishes should be assessed and selected for application with the aim to lower the mean size of the crack nucleating discontinuities and to narrow the related scatter in fatigue performance. Note: Surface enhancement treatments (e.g. shot-peening) can increase fatigue life scatter.
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6. Conclusions Despite over 170 years of research into the subject and numerous papers and conferences, in-service failures from cyclic mechanical fatigue still occur in metallic airframe components. Furthermore, blind predictions of the fatigue lives of airframe components subjected to simulated service loading are often poor, and unexpected fatigue problems occur during certification testing and service. It is considered that current fatigue management methods, whether damage tolerance (as it is currently applied) or safe life, could better account for the true, as-produced condition of metallic components, and in particular, the population of discontinuities that could potentially nucleate fatigue cracks. In addition, most current fatigue prediction methodologies cannot accurately predict the growth of the small cracks that nucleate from such discontinuities. To overcome such limitations, fatigue management methods can tend to have undesirable conservatisms, which leads to sub-optimal structural designs. Furthermore, these limitations also mean that the true risk of failure for an airframe is not well understood at any time during the service life, which can lead to fatigue management practices that do not optimally balance maintaining acceptable levels of safety, with maximizing aircraft availability and minimizing the outlay of resources. This paper identifies what the authors believe are the key factors that influence the risk of metallic airframe component failure from cyclic fatigue with the intention of focusing future research and development on the areas that would most productively overcome current limitations. In particular the growth of physically short cracks subject to high stresses and loading spectra, which nucleate upon introduction into service, from naturally occurring material and production-induced discontinuities, is considered an area of crucial importance. Declaration of Competing Interest The author declare that there is no conflict of interest. Acknowledgments The authors are indebted to numerous mentors and practitioners in the field of aircraft structural integrity. These include Dr. R.J.H. Wanhill, Mr. D. Polakovics, Mr R. Ryan, Dr. J.P. Gallagher, Prof. R. Jones, Dr. G.C. Sih, Dr. L.M. Butkus, Dr. D.W. Hoeppner, Mr C. Babish and many colleagues at the Defence Science and Technology Group and the Defence Aviation Safety Authority. Appendix A. Supplementary material Supplementary data to this article can be found online at https:// doi.org/10.1016/j.ijfatigue.2019.105323. References [1] Tiffany CF, Gallagher JP, Babish IV CA. Threats to aircraft structural safety, including a compendium of selected structural accidents/incidents. Wright-Patterson Air Force Base (OH), Aeronautical Systems Centre; 2010. Technical report no. ASC-
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